Certification of Transport Category Rotorcraft (Changes 1 - 9 incorporated)

CONTENTS Page No.

CHAPTER 1 - PART 21 CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS AC 21.16 Special Conditions .......................................................................... FAR 21-1 AC 21.31 Type Design .................................................................................... FAR 21-2 AC 21.33 Inspection and Tests ....................................................................... FAR 21-2 AC 21.35 Flight Tests ..................................................................................... FAR 21-2 AC 21.39 Flight Test Instrument Calibration and Correction Report ............... FAR 21-6 CHAPTER 2 - PART 29 AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT SUBPART A - GENERAL GENERAL AC 29.1 Applicability ........................................................................................... A-1 AC 29.1A Applicability ........................................................................................... A-2 AC 29.2 Special Retroactive Requirements ........................................................ A-3 SUBPART B - FLIGHT GENERAL AC 29.21 Proof of Compliance ......................................................................... B-1 AC 29.21A Proof of Compliance ......................................................................... B-3 AC 29.25 Weight Limits .................................................................................... B-3 AC 29.25A Weight Limits .................................................................................... B-4 AC 29.27 Center of Gravity Limits ..................................................................... B-5 AC 29.29 Empty Weight and Corresponding Center of Gravity ........................ B-6 AC 29.31 Removable Ballast ............................................................................ B-8 AC 29.33 Main Rotor Speed and Pitch Limits ................................................... B-8 PERFORMANCE AC 29.45 Performance - General ...................................................................... B-10 AC 29.45A Performance - General ...................................................................... B-20 AC 29.45B Performance - General ...................................................................... B-20 AC 29.49 Performance at Minimum Operating Speed . ..................................... B-26.2  AC 29-2C, Chg 5 6/13/2014 Page No.

SUBPART B - FLIGHT (continued) PERFORMANCE (continued) AC 29.51 Takeoff Data - General ...................................................................... B-29 AC 29.51A Takeoff Data - General ...................................................................... B-30 AC 29.53 Takeoff: Category A ......................................................................... B-31 AC 29.53A Takeoff: Category A ......................................................................... B-34 AC 29.55 Takeoff Decision Point: Category A ................................................. B-36 AC 29.59 Takeoff Path: Category A ................................................................. B-36 AC 29.59A Takeoff Path, Distance and Rejected Takeoff, Ground Level and ... B-46 Elevated Heliport: Category A AC 29.63 Takeoff: Category B ......................................................................... B-57 AC 29.65 Climb: (All Engines Operating) ......................................................... B-60 AC 29.65A Climb: (General and All Engines Operating) .................................... B-61 AC 29.67 Climb: One Engine Inoperative ........................................................ B-63 AC 29.67A Climb: One Engine Inoperative ........................................................ B-65 AC 29.71 Rotorcraft Angle of Glide: Category B .............................................. B-66 AC 29.73 Performance at Minimum Operating Speed ...................................... B-69 AC 29.75 Landing ............................................................................................. B-71 AC 29.75A Landing ............................................................................................. B-80 AC 29.77 Balked Landing: Category A ............................................................ B-91 AC 29.79 Limiting Height - Speed Envelope ..................................................... B-92 AC 29.85 Balked Landing: Category A ............................................................ B-100 AC 29.87 Limiting Height - Speed Envelope ..................................................... B-101 FLIGHT CHARACTERISTICS AC 29.141 Flight Characteristics – General ........................................................ B-102 AC 29.143 Controllability and Maneuverability ................................................... B-104 AC 29.143A Controllability and Maneuverability ................................................... B-106 AC 29.151 Flight Controls ................................................................................... B-108 AC 29.161 Trim Control ...................................................................................... B-109 AC 29.161A Trim Control ...................................................................................... B-110 AC 29.171 Stability: General .............................................................................. B-110 AC 29.173 Static Longitudinal Stability ............................................................... B-111 AC 29.173A Static Longitudinal Stability ............................................................... B-111 AC 29.175 Demonstration of Static Longitudinal Stability ................................... B-112.2 AC 29.175A Demonstration of Static Longitudinal Stability ................................... B-116 AC 29.177 Static Directional Stability .................................................................. B-  AC 29.177A Static Directional Stability.................................................................. B-116.5 AC 29.181 Dynamic Stability: Category A Rotorcraft ......................................... B-117 Page  6/13/2014 AC 29-2C, Chg 5 Page No.

SUBPART B - FLIGHT (continued) GROUND AND WATER HANDLING CHARACTERISTICS AC 29.231 General (Ground and Water Handling Characteristics) ..................... B-118 AC 29.235 Taxiing Condition .............................................................................. B-118 AC 29.239 Spray Characteristics ........................................................................ B-119 AC 29.241 Ground Resonance ........................................................................... B-120 MISCELLANEOUS FLIGHT REQUIREMENTS AC 29.251 Vibration ............................................................................................ B-125 SUBPART C - STRENGTH REQUIREMENTS GENERAL AC 29.301 Loads ................................................................................................ C-1 AC 29.303 Factor of Safety ................................................................................. C-1 AC 29.305 Strength and Deformation ................................................................. C-2 AC 29.307 Proof of Structure .............................................................................. C-2 AC 29.307A Proof of Structure .............................................................................. C-3 AC 29.309 Design Limitations ............................................................................. C-4 FLIGHT LOADS AC 29.321 General - Flight Loads ....................................................................... C-5 AC 29.337 Limit Maneuvering Load Factor ......................................................... C-5 AC 29.339 Resultant Limit Maneuvering Loads .................................................. C-8 AC 29.341 Gust Loads ........................................................................................ C-8 AC 29.351 Yawing Conditions ............................................................................ C-10 AC 29.351A Yawing Conditions ............................................................................ C-13 AC 29.351B Yawing Conditions ............................................................................ C-15 AC 29.361 Engine Torque .................................................................................. C-18 CONTROL SURFACE AND SYSTEM LOADS AC 29.391 Control Surface and System Loads - General ................................... C-20 AC 29.391A Control Surface and System Loads - General ................................... C-20 AC 29.395 Control System ................................................................................. C-20 AC 29.395A Control System ................................................................................. C-22 AC 29.397 Limit Pilot Forces and Torques ......................................................... C-22 AC 29.399 Dual Control System ......................................................................... C-23 AC 29.401 Auxiliary Rotor Assemblies .............................................................. C-23 Page  AC 29-2C, Chg 5 6/13/2014 Page No.

SUBPART C - STRENGTH REQUIREMENTS (continued) CONTROL SURFACE AND SYSTEM LOADS (continued) AC 29.401A Auxiliary Rotor Assemblies ............................................................... C-24 AC 29.403 Auxiliary Rotor Attachment Structure ................................................ C-25 AC 29.403A Auxiliary Rotor Attachment Structure ................................................ C-25 AC 29.411 Ground Clearance: Tail Rotor Guard ............................................... C-26 AC 29.413 Stabilizing and Control Surfaces ....................................................... C-27 AC 29.413A Stabilizing and Control Surfaces ....................................................... C-29 AC 29.427 Unsymmetrical Loads ....................................................................... C-29 GROUND LOADS AC 29.471 Ground Loads - General ................................................................... C-33 AC 29.473 Ground Loading Conditions and Assumptions .................................. C-33 AC 29.475 Tires and Shock Absorbers ............................................................... C-34 AC 29.477 Landing Gear Arrangement ............................................................... C-34 AC 29.479 Level Landing Conditions .................................................................. C-34 AC 29.481 Tail-Down Landing Conditions .......................................................... C-36 AC 29.483 One-Wheel Landing Conditions ........................................................ C-36 AC 29.485 Lateral Drift Landing Conditions ........................................................ C-36 AC 29.493 Braked Roll Conditions ...................................................................... C-37 AC 29.497 Ground Loading Conditions: Landing Gear With Tail Wheels ........... C-37 AC 29.501 Ground Loading Conditions: Landing Gear With Skids ..................... C-39 AC 29.501A Ground Loading Conditions: Landing Gear With Skids ..................... C-40 AC 29.505 Ski Landing Conditions ..................................................................... C-41 AC 29.511 Ground Load: Unsymmetrical Loads on Multiple-Wheel Units ......... C-41 WATER LOADS AC 29.519 Hull Type Rotorcraft: Water-Based, Amphibian ............................... C-43 AC 29.521 Float Landing Conditions .................................................................. C-45 MAIN COMPONENT REQUIREMENTS AC 29.547 Main Rotor Structure ......................................................................... C-46 AC 29.547A Main Rotor and Tail Rotor Structure ................................................. C-46 AC 29.549 Fuselage and Rotor Pylon ................................................................. C-50 AC 29.551 Auxiliary Lifting Surfaces ................................................................... C-50 Page  6/13/2014 AC 29-2C, Chg 5 Page No.

SUBPART C - STRENGTH REQUIREMENTS (continued) EMERGENCY LANDING CONDITIONS AC 29.561 Emergency Landing Conditions - General ........................................ C-51 AC 29.561A Emergency Landing Conditions - General ........................................ C-53 AC 29.561B Emergency Landing Conditions - General ........................................ C-54 AC 29.562 Emergency Landing Dynamic Conditions ......................................... C-55 AC 29.563 Structural Ditching Provisions ........................................................... C-92 AC 29.563A Structural Ditching Provisions ........................................................... C-92 FATIGUE EVALUATION AC 29.571 Fatigue Evaluation of Flight Structure ............................................... C-94 AC 29.571A Fatigue Evaluation of Flight Structure ............................................... C-97 AC 29.571B Fatigue Tolerance Evaluation of Metallic Structure ........................... C-98 AC 29.573 Damage Tolerance and Fatigue Evaluation of Composite ............... C-130 Rotorcraft Structures SUBPART D - DESIGN AND CONSTRUCTION GENERAL AC 29.601 Design ............................................................................................... D-1 AC 29.602 Critical Parts ...................................................................................... D-2 AC 29.603 Materials ........................................................................................... D-4 AC 29.605 Fabrication Methods ......................................................................... D-5 AC 29.607 Fasteners .......................................................................................... D-5 AC 29.609 Protection of Structure ...................................................................... D-6 AC 29.610 Lightning and Static Electricity Protection ......................................... D-8 AC 29.611 Inspection Provisions ........................................................................ D-12 AC 29.613 Material Strength Properties and Design Values .............................. D-13 AC 29.613A Material Strength Properties and Design Values .............................. D-14 AC 29.619 Special Factors ................................................................................. D-17 AC 29.621 Casting Factors ................................................................................. D-17 AC 29.623 Bearing Factors ................................................................................. D-27 AC 29.625 Fitting Factors ................................................................................... D-28.1 AC 29.625A Fitting Factors .................................................................................. D-28.1 AC 29.629 Flutter ................................................................................................ D-29 AC 29.629A Flutter and Divergence ...................................................................... D-29 AC 29.631 Bird Strike ......................................................................................... D-31 Page  AC 29-2C, Chg 5 6/13/2014 Page No.

SUBPART D - DESIGN AND CONSTRUCTION (continued) ROTORS AC 29.653 Pressure Venting and Drainage of Rotor Blades .............................. D-32 AC 29.659 Mass Balance ................................................................................... D-32 AC 29.661 Rotor Blade Clearance ...................................................................... D-32 AC 29.663 Ground Resonance Prevention Means ............................................. D-34 AC 29.663A Ground Resonance Prevention Means ............................................. D-36 CONTROL SYSTEMS AC 29.671 Control Systems - General ................................................................ D-37 AC 29.672 Stability Augmentation, Automatic, and Power-Operated Systems ... D-37 AC 29.673 Primary Flight Controls ...................................................................... D-38 AC 29.674 Interconnected Controls .................................................................... D-39 AC 29.675 Stops ................................................................................................. D-39 AC 29.679 Control System Locks ....................................................................... D-41 AC 29.681 Limit Load Static Tests ...................................................................... D-42 AC 29.683 Operation Tests ................................................................................ D-42 AC 29.685 Control System Details ...................................................................... D-43 AC 29.687 Spring Devices .................................................................................. D-45 AC 29.691 Autorotation Control Mechanism ....................................................... D-45 AC 29.695 Power Boost and Power-Operated Control System .......................... D-46 LANDING GEAR AC 29.723 Shock Absorption Tests .................................................................... D-49 AC 29.725 Limit Drop Test .................................................................................. D-52 AC 29.727 Reserve Energy Absorption Drop Test ............................................. D-52 AC 29.727A Reserve Energy Absorption Drop Test ............................................. D-53 AC 29.729 Retracting Mechanism, Landing Gear ............................................... D-53 AC 29.731 Wheels .............................................................................................. D-55 AC 29.733 Tires .................................................................................................. D-56 AC 29.735 Brakes ............................................................................................... D-58 AC 29.737 Skis ................................................................................................... D-59 FLOATS AND HULLS AC 29.751 Main Float Buoyancy ........................................................................ D-61 AC 29.753 Main Float Design ............................................................................. D-61 AC 29.755 Hull Buoyancy ................................................................................... D-64 AC 29.757 Hull and Auxiliary Float Strength ....................................................... D-65 Page  5/1/2014 AC 29-2C, Chg 4 Page No.

SUBPART D - DESIGN AND CONSTRUCTION (continued) PERSONNEL AND CARGO ACCOMMODATIONS AC 29.771 Pilot Compartment ............................................................................ D-67 AC 29.773 Pilot Compartment View .................................................................... D-68 AC 29.775 Windshields and Windows ................................................................ D-74 AC 29.775A Windshields and Windows ................................................................ D-74 AC 29.775B Windshields and Windows ................................................................ D-74 AC 29.777 Cockpit Controls ................................................................................ D-75 AC 29.779 Motion and Effect of Cockpit Controls ............................................... D-76 AC 29.783 Doors ................................................................................................ D-78 AC 29.783A Doors ................................................................................................ D-82 AC 29.785 Seats, Safety Belts, and Harnesses .................................................. D-83 AC 29.785A Seats, Berths, Belts, Safety Belts, and Harnesses ............................ D-85 AC 29.785B Seats, Berths, Litters, Safety Belts, and Harnesses ......................... D-87 AC 29.787 Cargo and Baggage Compartments .................................................. D-88 AC 29.787A Cargo and Baggage Compartments .................................................. D-88 AC 29.801 Ditching ............................................................................................. D-89 AC 29.803 Emergency Evacuation ..................................................................... D-97 AC 29.803A Emergency Evacuation ..................................................................... D-97 AC 29.805 Flightcrew Emergency Exits .............................................................. D-98 AC 29.805A Flightcrew Emergency Exits .............................................................. D-98 AC 29.807 Passenger Emergency Exits ............................................................. D-99 AC 29.807A Emergency Exits ............................................................................... D-100 AC 29.809 Emergency Exit Arrangement ........................................................... D-101 AC 29.809A Emergency Exit Arrangement ........................................................... D-102 AC 29.811 Emergency Exit Marking ................................................................... D-103 AC 29.811A Emergency Exit Marking ................................................................... D-105 AC 29.812 Emergency Lighting .......................................................................... D-106 AC 29.813 Emergency Exit Access .................................................................... D-106 AC 29.815 Main Aisle Width ............................................................................... D-108 AC 29.831 Ventilation ......................................................................................... D-109 AC 29.833 Heaters ............................................................................................. D-110 FIRE PROTECTION AC 29.851 Fire Extinguishers ............................................................................. D-112 AC 29.853 Compartment Interiors ...................................................................... D-113 AC 29.855 Cargo and Baggage Compartments .................................................. D-115 AC 29.855A Cargo and Baggage Compartments .................................................. D-117 AC 29.859 Combustion Heater Fire Protection ................................................... D-117 AC 29.861 Fire Protection of Structure, Controls, and Other Parts ..................... D-121 AC 29.861A Fire Protection of Structure, Controls, and Other Parts ..................... D-124 Page ix Page No.

SUBPART D - DESIGN AND CONSTRUCTION (continued) FIRE PROTECTION (continued) AC 29.863 Flammable Fluid Fire Protection ....................................................... D-124 EXTERNAL LOAD ATTACHING MEANS AC 29.865 External Load Attaching Means ........................................................ D-129 AC 29.865A External Load Attaching Means ........................................................ D-134 AC 29.865B External Load Attaching Means (SEE MG 12) .................................. D-140 MISCELLANEOUS (DESIGN AND CONSTRUCTION) AC 29.871 Leveling Marks .................................................................................. D-165 AC 29.873 Ballast Provisions .............................................................................. D-165 AC 29.877 Ice Protection .................................................................................... D-166 SUBPART E - POWERPLANT GENERAL AC 29.901 Installation ......................................................................................... E-1 AC 29.901A Installation ......................................................................................... E-8 GENERAL (continued) AC 29.901B Installation ......................................................................................... E-8 AC 29.903 Engines ............................................................................................. E-9 AC 29.903A Engines ............................................................................................. E-11 AC 29.903B Engines ............................................................................................. E-12 AC 29.903C Turbine Engine Installation ................................................................ E-13 AC 29.907 Engine Vibration ................................................................................ E-29 AC 29.908 Cooling Fans ..................................................................................... E-29 AC 29.908A Cooling Fans ..................................................................................... E-30 ROTOR DRIVE SYSTEM AC 29.917 Design ............................................................................................... E-33 AC 29.917A Design ............................................................................................... E-34 AC 29.921 Rotor Brake ....................................................................................... E-37.2 AC 29.923 Rotor Drive System and Control Mechanism Tests ........................... E-39 AC 29.923A Rotor Drive System and Control Mechanism Tests ........................... E-45 AC 29.923B Rotor Drive System and Control Mechanism Tests ........................... E-52 Page x 6/23/23 AC 29-2C, Chg 9 Page No.

SUBPART E - POWERPLANT (continued) ROTOR DRIVE SYSTEM (continued) AC 29.923C Rotor Drive System and Control Mechanism Tests .......................... E-53 AC 29.927 Additional Tests ................................................................................ E-54.1 AC 29.927A Additional Tests ................................................................................ E-56 AC 29.931 Shafting Critical Speeds .................................................................... E-58 AC 29.935 Shafting Joints .................................................................................. E-63 AC 29.939 Turbine Engine Operating Characteristics ........................................ E-63 FUEL SYSTEM AC 29.951 Fuel System - General ...................................................................... E-66 AC 29.952 Fuel System Crash Resistance ......................................................... E-68 AC 29.953 Fuel System Independence .............................................................. E-93 AC 29.954 Fuel System Lightning Protection ..................................................... E-94 AC 29.955 Fuel Flow .......................................................................................... E-95 AC 29.955A Fuel Flow .......................................................................................... E-99 Fuel Flow .......................................................................................... E- 102.1 I AC 29.955 B AC 29.957 Flow Between Interconnected Tanks ................................................ E-103 AC 29.959 Unusable Fuel Supply ....................................................................... E-103 AC 29.961 Fuel System Hot Weather Operation ................................................ E-104 AC 29.961A Fuel System Hot Weather Operation ................................................ E-106 AC 29.963 Fuel Tanks: General ........................................................................ E-107 AC 29.963A Fuel Tanks: General ........................................................................ E-107 AC 29.963B Fuel Tanks: General ........................................................................ E-108 AC 29.965 Fuel Tank Tests ................................................................................ E-109 AC 29.967 Fuel Tank Installation ........................................................................ E-111 AC 29.967A Fuel Tank Installation ........................................................................ E-113 AC 29.969 Fuel Tank Expansion Space ............................................................. E-113 AC 29.969A Fuel Tank Expansion Space ............................................................. E-114 AC 29.971 Fuel Tank Sump ................................................................................ E-114 AC 29.971A Fuel Tank Sump ................................................................................ E-116 AC 29.973 Fuel Tank Filler Connection .............................................................. E-116 AC 29.973A Fuel Tank Filler Connection .............................................................. E-117 AC 29.975 Fuel Tank Vents and Carburetor Vapor Vents .................................. E-118 AC 29.975A Fuel Tank Vents and Carburetor Vapor Vents .................................. E-119 AC 29.975B Fuel Tank Vents and Carburetor Vapor Vents .................................. E-119 AC 29.977 Fuel Tank Outlet ............................................................................... E-120 AC 29.979 Pressure Refueling and Fueling Provisions Below Fuel Level .......... E-121 Page  6/23/23 AC 29-2C, Chg 9 Page No.

SUBPART E - POWERPLANT (continued) FUEL SYSTEM COMPONENTS AC 29.991 Fuel Pumps ....................................................................................... E-123 AC 29.991A Fuel Pumps ....................................................................................... E-124 AC 29.993 Fuel System Lines and Fittings ......................................................... E-124 AC 29.995 Fuel Valves ....................................................................................... E-125 AC 29.997 Fuel Strainer or Filter ........................................................................ E-125 AC 29.997A Fuel Strainer or Filter ........................................................................ E-126 AC 29.997 B Fuel Strainer or Filter ........................................................................ E-12 7.1 AC 29.999 Fuel System Drains........................................................................... E-127 .2 AC 29.999A Fuel System Drains........................................................................... E-127 .2 AC 29.1001 Fuel Jettisoning ................................................................................. E-128 OIL SYSTEM AC 29.1011 Oil System - General ......................................................................... E-130 AC 29.1013 Oil Tanks ........................................................................................... E-131 AC 29.1015 Oil Tank Tests ................................................................................... E-131 AC 29.1017 Oil Lines and Fittings ........................................................................ E-132 AC 29.1019 Oil Strainer or Filter ........................................................................... E-132 AC 29.1019A Oil Strainer or Filter ........................................................................... E-133 AC 29.1019 B Oil Strainer or Filter........................................................................... E-13 4.1 AC 29.1021 Oil System Drains ............................................................................. E-134 .2 AC 29.1023 Oil Radiators ..................................................................................... E-134 AC 29.1025 Oil Valves.......................................................................................... E-134 .2 AC 29.1027 Transmission and Gearboxes: General ............................................ E-135 COOLING AC 29.1041 Cooling - General .............................................................................. E-137 AC 29.1043 Cooling Tests .................................................................................... E-140 AC 29.1043A Cooling Tests .................................................................................... E-142 AC 29.1045 Cooling Test Procedures ................................................................... E-143 AC 29.1047 Takeoff Cooling Test Procedures ...................................................... E-147 AC 29.1049 Hovering Cooling Test Procedures ................................................... E-149 INDUCTION SYSTEM AC 29.1091 Air Induction ...................................................................................... E-151 AC 29.1093 Induction System Icing Protection ..................................................... E-152 AC 29.1093A Induction System Icing Protection..................................................... E-160.2 AC 29.1101 Carburetor Air Preheater Design ....................................................... E-161 AC 29.1103 Induction System Ducts and Air Duct Systems ................................. E-161 AC 29.1105 Induction System Screens ................................................................ E-164 Page  5/1/2014 AC 29-2C, Chg 4 Page No.

SUBPART E - POWERPLANT (continued) INDUCTION SYSTEM (continued) AC 29.1107 Intercoolers and After-Coolers .......................................................... E-165 AC 29.1109 Carburetor Air Cooling ...................................................................... E-165 EXHAUST SYSTEM AC 29.1121 General ............................................................................................. E-166 AC 29.1123 Exhaust Piping .................................................................................. E-167 AC 29.1125 Exhaust Heat Exchangers ................................................................. E-169 POWERPLANT CONTROLS AND ACCESSORIES AC 29.1141 Powerplant Controls: General .......................................................... E-173 AC 29.1142 Auxiliary Power Unit Controls ............................................................ E-174 AC 29.1143 Engine Controls ................................................................................. E-175 AC 29.1143A Engine Controls ................................................................................. E-176 AC 29.1143B Engine Controls ................................................................................. E-176 AC 29.1145 Ignition Switches ............................................................................... E-176 AC 29.1147 Mixture Controls ................................................................................ E-177 AC 29.1151 Rotor Brake Controls ......................................................................... E-178 AC 29.1157 Carburetor Air Temperature Controls ................................................ E-179 AC 29.1159 Supercharger Controls ...................................................................... E-179 AC 29.1163 Powerplant Accessories .................................................................... E-180 AC 29.1165 Engine Ignition Systems .................................................................... E-181 POWERPLANT FIRE PROTECTION AC 29.1181 Designated Fire Zones: Regions Included ....................................... E-182 AC 29.1183 Lines, Fittings, and Components ....................................................... E-183 AC 29.1185 Flammable Fluids .............................................................................. E-184 AC 29.1187 Drainage and Ventilation of Fire Zones ............................................. E-186 AC 29.1189 Shutoff Means ................................................................................... E-186 AC 29.1191 Firewalls ............................................................................................ E-187 AC 29.1193 Cowling and Engine Compartment Covering .................................... E-191 AC 29.1193A Cowling and Engine Compartment Covering .................................... E-192 AC 29.1194 Other Surfaces .................................................................................. E-193 AC 29.1195 Fire Extinguishing Systems ............................................................... E-194 AC 29.1197 Fire Extinguishing Agents ................................................................. E-197 AC 29.1199 Extinguishing Agent Containers ........................................................ E-200 AC 29.1201 Fire Extinguishing System Materials ................................................. E-201 AC 29.1203 Fire Detector Systems ....................................................................... E-201 Page xiii 6/23/23 AC 29-2C, Chg 9 Page No.

SUBPART F - EQUIPMENT GENERAL AC 29.1301 Function and Installation ................................................................... F-1 AC 29.1303 Flight and Navigation Instruments ..................................................... F-1 AC 29.1305 Powerplant Instruments .................................................................... F-7 AC 29.1305A Powerplant Instruments .................................................................... F-8 AC 29.1305B Powerplant Instruments .................................................................... F-9 AC 29.1305C Powerplant Instruments .................................................................... F-12 AC 29.1305 D Powerplant Instruments .................................................................... F-12 .1 AC 29.1307 Miscellaneous Equipment ................................................................. F-12 .3 AC 29.1309 Equipment, Systems, and Installations ............................................. F-13 AC 29.1309A Equipment, systems, and Installations . ............................................. F-40 AC 29.1309 B Equipment, systems, and Installations. ............................................. F-40 .1 AC 29.1316 Electrical and Electronic System Lightning Protection ...................... F-41 AC 29.1317 High-Intensity Radiated Fields (HIRF) Protection .............................. F-42 INSTRUMENTS - INSTALLATION AC 29.1321 Arrangement and Visibility ................................................................. F-43 AC 29.1322 Warning, Caution, and Advisory Lights ............................................. F-44.2 AC 29.1323 Airspeed Indicating System ............................................................... F-48 AC 29.1323A Airspeed Indicating System ............................................................... F-49 AC 29.1325 Static Pressure Systems ................................................................... F-50 AC 29.1327 Magnetic Direction Indicator .............................................................. F-51 AC 29.1329 Automatic Pilot Systems ................................................................... F-51 AC 29.1329A Automatic Pilot System. .................................................................... F-60.1 AC 29.1329 B Automatic Pilot and Flight Guidance System. ................................... F-60. 2a AC 29.1331 Instruments Using a Power Supply ................................................... F-60.3 AC 29.1333 Instrument Systems .......................................................................... F-61 AC 29.1333 A Instrument Systems .......................................................................... F-61 .1 AC 29.1335 Flight Director Systems ..................................................................... F-62 AC 29.1335 A Flight Director Systems ..................................................................... F-6 3.1 AC 29.1337 Powerplant Instruments - Fuel Gauge Calibration............................ F-63 .2 AC 29.1337A Powerplant Instruments .................................................................... F-64 ELECTRICAL SYSTEMS AND EQUIPMENT AC 29.1351 Electrical Systems and Equipment - General .................................... F-66 AC 29.1353 Electrical Equipment and Installations ............................................... F-74 AC 29.1353 A Energy Storage Systems ............................................. .................... .. F-7 8.1 AC 29.1355 Distribution System ........................................................................... F-79 AC 29.1355A Distribution System ........................................................................... F-79 AC 29.1357 Circuit Protective Devices ................................................................. F-80 AC 29.1357A Circuit Protective Devices ................................................................. F-82 AC 29.1359 Electrical System Fire and Smoke Protection ................................... F-82.1 AC 29.1363 Electrical Systems Tests ................................................................... F-83 Page xiv AC 29-2C, Chg 5 Page No.

SUBPART F - EQUIPMENT (continued) LIGHTS AC 29.1381 Instrument Lights .............................................................................. F-84 AC 29.1383 Landing Lights ................................................................................... F-85 AC 29.1385 Position Light System Installation ..................................................... F-85 AC 29.1387 Position Light System Dihedral Angles ............................................. F-85 AC 29.1389 Position Light Distribution and Intensities .......................................... F-85 AC 29.1391 Minimum Intensities in the Horizontal Plane of Forward and Rear Position Light ................................................................................. F-85 AC 29.1393 Minimum Intensities in Any Vertical Plane of Forward and Rear Position Lights ................................................................................ F-85 AC 29.1395 Maximum Intensities In Overlapping Beams of Forward and Rear Position Lights ................................................................................ F-85 AC 29.1397 Color Specifications .......................................................................... F-85 AC 29.1399 Riding Light ....................................................................................... F-86 AC 29.1401 Anticollision Light System ................................................................. F-87 SAFETY EQUIPMENT AC 29.1411 Safety Equipment - General .............................................................. F-88 AC 29.1413 Safety Belts: Passenger Warning Device ......................................... F-90 AC 29.1415 Ditching Equipment ........................................................................... F-91 AC 29.1419 Ice Protection .................................................................................... F-93 MISCELLANEOUS EQUIPMENT AC 29.1431 Electronic Equipment ........................................................................ F-114 AC 29.1433 Vacuum Systems .............................................................................. F-114 AC 29.1435 Hydraulic Systems ............................................................................ F-115 AC 29.1439 Protective Breathing Equipment ........................................................ F-118 AC 29.1457 Cockpit Voice Recorder .................................................................... F-119 AC 29.1459 Flight Recorders ................................................................................ F-120 AC 29.1461 Equipment Containing High Energy Rotors ....................................... F-123 SUBPART G - OPERATING LIMITATIONS AND INFORMATION AC 29.1501 Operating Limitations - General ........................................................ G-1 OPERATING LIMITATIONS AC 29.1503 Airspeed Limitations: General .......................................................... G-1 AC 29.1505 Never-Exceed Speed ........................................................................ G-3 Page  6/23/23 AC 29-2C, Chg 9 Page No.

SUBPART G - OPERATING LIMITATIONS AND INFORMATION (continued) OPERATING LIMITATIONS (continued) AC 29.1509 Rotor Speed ...................................................................................... G-6 AC 29.1517 Limiting Height-Speed Envelope ....................................................... G-8 AC 29.1517 A Limiting Height- Velocity Envelope................................................ . .... G- 9.1 AC 29.1519 Weight and Center of Gravity............................................................ G-9 .2 AC 29.1521 Powerplant Limitations...................................................................... G-9 .2 AC 29.1521A Powerplant Limitations ...................................................................... G-10 AC 29.1521B Powerplant Limitations ...................................................................... G-11 AC 29.1522 Auxiliary Power Unit Limitations ........................................................ G-12 AC 29.1523 Minimum Flightcrew .......................................................................... G-12 AC 29.1525 Kinds of Operations .......................................................................... G-16 AC 29.1527 Maximum Operating Altitude ............................................................. G-17 AC 29.1529 Instructions for Continued Airworthiness (Maintenance Manual) ...... G-17 AC 29.1529A Instructions for Continued Airworthiness (Maintenance Manual) ...... G-22 MARKINGS AND PLACARDS AC 29.1541 Marking and Placards - General ....................................................... G-23 AC 29.1543 Instrument Markings: General.......................................................... G-23 AC 29.1543 A Instrument Markings: General .......................................................... G-2 7.1 AC 29.1545 Airspeed Indicator ............................................................................. G-27 . 6 AC 29.1545 A Airspeed Indicator ............................................................................. G-27 . 7 AC 29.1547 Magnetic Direction Indicator.............................................................. G-27 . 9 AC 29.1549 Powerplant Instruments .................................................................... G-27 . 9 AC 29.1549A Powerplant Instruments .................................................................... G-27 . 9 AC 29.1549 B Powerplant Instruments .................................................................... G-27 . 10 AC 29.1551 Oil Quantity Indicators ....................................................................... G-28 AC 29.1553 Fuel Quantity Indicator ...................................................................... G-28 AC 29.1555 Control Markings ............................................................................... G-29 AC 29.1555 A Control Markings............................................................................... G -29.2 AC 29.1557 Miscellaneous Markings and Placards .............................................. G-30 AC 29.1559 Limitations Placard ............................................. ............................... G-31 AC 29.1561 Safety Equipment .............................................................................. G-32 AC 29.1565 Tail Rotor .......................................................................................... G-33 ROTORCRAFT FLIGHT MANUAL AC 29.1581 Rotorcraft Flight Manual - General .................................................... G-34 AC 29.1583 Operating Limitations ........................................................................ G-42 AC 29.1583A Operating Limitations ........................................................................ G-45 AC 29.1585 Operating Procedures ....................................................................... G-45 AC 29.1587 Performance Information ................................................................... G-47 AC 29.1587A Performance Information ................................................................... G-51 AC 29.1587B Performance Information ................................................................... G-51 AC 29.1587 C Performance Information.................................................................. G-5 4.1 AC 29.1589 Loading Information .......................................................................... G-51 Page  2/4/2016 AC 29-2C, Chg 7 Page No.

CHAPTER 3 - MISCELLANEOUS GUIDANCE (MG), TRANSPORT CATEGORY ROTORCRAFT AC 29 MG 1 Certification Procedure for Rotorcraft Avionics Equipment ............. MG 1-1 AC 29 MG 2 Standardized Test Procedure for Rotorcraft DC Electrical ............. MG 2-1 System Tests AC 29 MG 3 Rotorcraft and Systems Certification for Category II Operations .... MG 3-1 AC 29 MG 4 Full Authority Digital Electronic Controls (FADEC) ......................... MG 4-1 AC 29 MG 5 Agricultural Dispensing Equipment Installation ............................... MG 5-1 AC 29 MG 6 Emergency Medical Service (EMS) Systems, Installations, .......... MG 6-1 Interior Arrangements, and Equipment AC 29 MG 7 Structural Condition Indicators ....................................................... MG 7-1 AC 29 MG 8 Substantiation of Composite Rotorcraft Structure .......................... MG 8-1 AC 29 MG 9 Rotorcraft One-Engine-Inoperative (OEI) Power Assurance .......... MG 9-1 AC 29 MG 10 Advisory Material for Substantiation of Emergency ........................ MG 10-1 Flotation System AC 29 MG 11 Fatigue Tolerance Evaluation of Transport Category ..................... MG 11-1 Rotorcraft Metallic Structure (Including Flaw Tolerance) AC 29 MG 12 External Loads ............................................................................... MG 12-1 AC 29 MG 13 Systems Certification Considerations ............................................ MG 13-1 AC 29 MG 14 Certification Procedure for Installation of Vapor Cycle .................. MG 14-1 Air Conditioning Systems AC 29 MG 15 Airworthiness Approval of Rotorcraft Health Usage Monitoring ..... MG 15-1 Systems (HUMS) AC 29 MG 16 Certification Procedure for Rotorcraft Night Vision Imaging .......... MG 16-1 Systems (NVIS) Equipment AC 29 MG 17 Guidance on Analyzing an Advanced Flight Controls (AdFC) ........ MG 17-1 System AC 29 MG 18 Helicopter Terrain Awareness and Warning System (HTAWS)...... MG 18-1 AC 29 MG 19 Guidance on Electronic Display Systems (EDS) for Rotorcraft ...... MG 19-1 Installations AC 29 MG 20 Human Factors (HF) ....................................................................... MG 20-1 AC 29 MG 21 Guidance on Creating a System Level Functional Hazard ............. MG 21-1 Assessment (FHA) AC 29 MG 22 Rotorcraft One Engine Inoperative (OEI) Training Mode ................ MG 22-1 AC 29 MG 23 Automatic Flight Guidance and Control Systems (AFGCS) ............ MG 23-1 Installation in Part 29 Rotorcraft Page xvii AC 29-2C, Chg 9 6/23/23 APPENDICES AC 29 Appendix A – Instructions for Continued Airworthiness.................................. Apdx A-1 AC 29 Appendix B - Airworthiness Guidance for Rotorcraft Instrument Flight ........... Apdx B-1 AC 29 Appendix C - ( reserved) (Icing Certification ) .................................................. Apdx C-1 AC 29 Appendix D - (reserved) (Criteria for Demonstration of Emergency ) ............ Apdx D-1 Evacuation Procedures Under § 29.803 Page xviii 2/4/2016 AC 29-2C, Chg 7 FIGURES The first digits in the figure number are the same as the related paragraph number. The digits after the hyphen are the number within the paragraph. Most figures are located at the end of the related paragraph.

Figure No. Title Page No.

AC 29.45-1 H-V Requirements. B-22 AC 29.45-2 IGE Controllability Requirements. B-23 AC 29.45-3 Shaft Horsepower vs. Turbine Outlet Temperature - B-24 Sea Level Standard Day.

AC 29.45-4 Uninstalled Takeoff Power Available. B-25 AC 29.45-5 Installed Takeoff Power Available, Anti-Ice Off, 400 RPM. B-26 AC 29.45-6 Power Assurance Check Chart. B-26.1 AC 29.53-1 Takeoff Performance Category A. B-33 AC 29.53A-1 Takeoff Performance Category A. B-35 AC 29.59-1 Category A Vertical Takeoff Profile Ground Level Heliport. B-44 AC 29.59-2 Category A Vertical Takeoff Profile Pinnacle. B-45 AC 29.59A-1 Category A Vertical Takeoff Profile Ground Level Heliport. B-55 AC 29.59A-2 Category A Vertical Takeoff Profile Elevated Heliport. B-56 AC 29.63-1 Conventional Takeoff Profile Category B. B-59 AC 29.71-1 Autorotational Characteristics – Typical. B-68 AC 29.75-1 Category A Conventional Landing - Clear Heliport. B-78 AC 29.75-2 Category A Vertical Landing. B-79 AC 29.75A-1 Category A Conventional Landing - Clear Heliport. B-89 AC 29.75A-2 Category A Vertical Landing. B-90 AC 29.79-1 Height-Velocity (HV) Diagram. B-98 AC 29.79-2 Altitude/Weight Accountability. B-99 AC 29.175-1 Static Longitudinal Stability. B-115 AC 29.175A-1 Static Longitudinal Stability. B-116.3 AC 29.251-1 Demonstration Points. B-127 AC 29.337-1 Load Factor Gross Weight Curve. C-7 AC 29.351-1 Sample Yaw/Forward Speed Diagram. C-12 Page xix AC 29-2C, Chg 7 2/4/2016 Figure No. Title Page No.

AC 29.351A-1 Sample Yaw/Forward Speed Diagram. C-14 AC 29.351B-1 Yaw/Forward Speed Diagram. C-17 AC 29.351B-2 Safety Factors for Probability of Failure. C-18 AC 29.427-1 (View Looking Forward). C-31 AC 29.427-2 (View Looking Forward). C-32 AC 29.562-1 Seat Restraint System Dynamic Tests. C-58 AC 29.562-2 Schematic Floor Deformation Fixture; Seat Legs Attached C-65 at Floor Level.

AC 29.562-3 Schematic Test Fixture; Side Wall Mounted Seat. C-69 AC 29.562-4 Installation of Pelvic Lumbar Spine Load Cell in Part 572B C-73 Anthropomorphic Dummy.

AC 29.562-5 Measurement to be Taken Over Full Width of Seat C-85 Bottom Cushion.

AC 29.562-6 Post Test Condition. C-85 AC 29.562-7 Impulse Shape. C-91 AC 29.562-8 Change in Velocity. C-91 AC 29.571-1 Comparison Percent of Time in Level Flight. C-96 AC 29.571B-1 Seven Fatigue Evaluation Methods discussed C-103 in this guidance.

AC 29.571B-2 No Growth. C-118 AC 29.571B-3 Slow Growth of Undetectable Crack. C-118 AC 29.571B-4 Slow Growth of Detectable Crack. C-119 AC 29.571B-5 Decreasing Residual Life in Secondary Load Path C-122 for Multiple Element Crack Growth with Inspections.

AC 29.571B-6 Multiple Load Path Structure Analytical Evaluation To C-123 Support Inspection for a Failed Load Path.

AC 29.571B-7 Multiple Load Path Structure Evaluation by Test to C-125 Support Inspection for a Failed Load Path.

AC 29.571B-8 Multiple Load Path Structure Analytical Evaluation to C-126 Support Inspection for Less than a Failed Load Path.

Page xx AC 29-2C, Chg 7 2/4/2016 Figure No. Title Page No.

AC 29.573-1 Schematic Diagram of Building Block Tests. C-144 AC 29.573-2 Characterization of Impact Damage. C-150 AC 29.573-3 Residual Strength vs. Time. C-154 AC 29.573-4 Illustration of Residual Strength and Damage Size. C-155 Relationships for Fail-Safe Substantiation.

AC 29.621-1 Example of Casting Load Sheet Retract Actuator D-24 Support-Landing Gear.

AC 29.621-2 Castings to be Static Tested for Model XYZ. D-26 AC 29.773-1 Cockpit Visibility. D-73 AC 29.801-1 Sea State Code - World Meteorological Organization. D-96 AC 29.863-1 Flammable Fluids Fire Protection Flow Chart. D-128 AC 29.865-1 Summary of Part 133 Rotorcraft Load Combinations D-136 Certifiable Under § 29.865.

AC 29.865B-1 U.S. Operational (Part 133) Rotorcraft-Load D-146 Combination vs. Applicable Cargo Type Data And Definition Summary.

AC 29.877-1 Continuous Icing-Temperature vs. Altitude Limits. D-181 AC 29.877-2 Intermittent Icing-Temperature vs. Altitude Limits. D-182 AC 29.877-3 Mean Effective Drop Diameter-Microns. D-183 AC 29.877-4 Intermittent Icing-Liquid Water Content vs. Drop Diameter. D-184 AC 29.877-5 Supercooled Clouds. D-185 AC 29.903C-1 Estimated Path of Fragments. E-23 AC 29.903C -2 Single One-Third Disc Fragment. E-24 AC 29.903C -3 Intermediate and Small Pieces of Debris. E-25 AC 29.903C -4 Blade Fragment Definition. E-26 AC 29.903C -5 Distribution of Translational and Rotational Kinetic. E-27 Energy of Rotor-Component Fragments as a Function of Fragment Size.

AC 29.903C -6 Cross Section through Aircraft at Plane of Rotation of E-28 the Engine Disk Fragment.

AC 29.923A-1 Rotor Drive System Endurance. E-48 Page xxi AC 29-2C, Chg 7 2/4/2016 Figure No. Title Page No.

AC 29.923A-2 Takeoff Run Portion of Rotor Drive System Endurance .......... E-49 Test If 2 ½ Minute OEI Rating is Requested.

AC 29.923A-3 Takeoff Run Portion of Rotor Drive System Endurance .......... E-50 Test If 30 Second/2 Minute OEI Rating is Requested.

AC 29.931-1 Cantilevered Shaft, First Critical Speed. ................................. E-62 AC 29.931-2 Shaft Between Support Bearings, First Critical Speed . ........... E-62 AC 29.952-1 Static Tension and Shear Loading Modes. ............................. E-87 AC 29.952-2 Hose Assembly Tests. ............................................................ E-88 AC 29.952-3 Minimum Average and Individual Loads for Hose and ............ E-89 Hose-End Fitting Combinations.

AC 29.952-4 Typical Method of Breakaway Fuel Fitting Load Calculations. E-90 AC 29.952-5 Typical Methods of Frangible or Deformable Attachment ....... E-91 Load Calculations (Frangible Bolts).

AC 29.952-6 Typical Methods of Frangible or Deformable Attachment ....... E-92 Load Calculations (Frangible Baffle).

AC 29.955A-1 Fuel Flow. E-102 AC 29.1041-1 Additional Cooling Test Point. ................................................. E-139 AC 29.1045-1 All-Engine-Critical Altitude Additional Cooling Test Point ........ E-146 AC 29.1191-1 Table of Materials and Gages Acceptable for Fireproof .......... E-190 Protective Devices with Flat Surface Geometries.

AC 29.1309-1 Relationship Between Probability and Severity of Failure .......F-17 / F-40.5 Condition Effects.

AC 29.1309-2 Failure Condition Categories and Probability Definitions. .......F-18 / F-40.6 AC 29.1309-3 Failure Condition Conversion .................................................. F-19 AC 29.1309-4 Safety Assessment Process Overview ....................................F-24 / F-40.10 AC 29.1309-5 Depth of Analysis Flowchart ....................................................F-25 / F-40.11 AC 29.1321-1 Primary Fields of View ............................................................ F-43 AC 29.1321-1 Secondary Fields of View ........................................................ F-44 AC 29.1329-1 Height Loss Method. ............................................................... F-54 AC 29.1329-2 Deviation Profile . ..................................................................... F-58 AC 29.1329-3 Deviation Profile . ..................................................................... F-60 AC 29.1419-1 Continuous Icing-Temperature vs. Altitude Limits . .................. F-110 Page xxii AC 29-2C, Chg 7 2/4/2016 Figure No. Title Page No.

AC 29.1419-2 Intermittent Icing-Temperature vs. Altitude Limits. .................. F-111 AC 29.1419-3 Continuous Icing-Liquid Water Content vs Drop Diameter ...... F-112 AC 29.1419-4 Intermittent Icing-Liquid Water Content vs Drop Diameter. ..... F-113 AC 29.MG 3-1 Category II Approach Evaluation .............................................MG 3-5 AC 29.MG 3-2 Category II Approach Evaluation. (cont.) ................................MG 3-6 AC 29.MG 5-1 Acceptable Ultimate Load Factor for Agricultural ....................MG 5-3 Dispensing Equipment Design.

AC 29.MG 5-2 Sketch of Tank Pressure Test .................................................MG 5-7 AC 29.MG 6-1 Typical Liquid Oxygen System ................................................MG 6-22 AC 29.MG 6-2 Typical Liquid Oxygen System - Using Combination Valve .....MG 6-23 AC 29 MG 8-1 Schematic Diagram of Building Block Tests ............................MG 8-16 AC 29 MG 8-2 Characterization of Impact Damage ........................................MG 8-20 AC 29 MG 8-3 Residual Strength vs. Time. ....................................................MG 8-23 AC 29 MG 8-4 Residual Strength and Damage Size Relationships for ...........MG 8-24 Fail-Safe Substantiation.

AC 29 MG 9-1 Installed Engine Fuel Flow Capacity Check ............................MG 9-8 AC 29 MG 11-1 Crack Growth for Single Element Structure ............................MG 11-14 AC 29 MG 11-2 Crack Growth for Remaining Elements of Multiple..................MG 11-16 Element Structure.

AC 29 MG 11-3 No Crack Growth Scenario. ....................................................MG 11-18 AC 29 MG 11-4 Slow Crack Growth Scenario. .................................................MG 11-18 AC 29 MG 11-5 Detectable Crack Growth Scenario. ........................................MG 11-19 AC 29 MG 14-1 Refrigeration Cycle ..................................................................MG 14-7 AC 29 MG 15-1 Flow Chart for Use of Mitigating Actions Applied for ...............MG 15-8 Reduction of Software Level.

AC 29 MG 15-2 Flow Chart for Application of HUMS Specific Software ...........MG 15-11 with a Ground Base that uses COTS Software for Operational Software AC 29 MG 17-1 Verification Methodology .........................................................MG 17-4 AC 29 MG 19-1 Primary Field of View ..............................................................MG 19-15 AC 29 MG 19-2 Secondary Field of View .........................................................MG 29-16 AC 29 MG 20-1 Primary and Secondary Horizontal Fields of View ..................MG 20-18 Page xxiii AC 29-2C, Chg 7 2/4/2016 AC 29 MG 20-2 Primary and Secondary Vertical Field of View ........................MG 20-19 AC 29 MG 20-3 Sample HF Compliance Matrix ................................................MG 20-27 AC 29 MG 21-1 Generic Examples by ATA ......................................................MG 21-5 AC 29.APX B-1 Rotorcraft Dynamic Stability Requirements for IFR............... Apdx B-27 Page xxiv 2/4/2016 AC 29-2C, Chg 7 Acronyms AC advisory circular EFIS electronic flight instrument ACO aircraft certification office System [Replaced by EDS.]

ADF automatic direction finding EFP engine failure point ADI attitude direction indicator EHE exhaust heat exchanger AEH airborne electronic hardware ELOS equivalent level of safety AEO all engines operating ELT emergency locator transmitter AdFC advanced flight control EMC electromagnetic compatibility AFCS automatic flight control EMI electromagnetic interference systems EMS emergency medical service AFGCS automatic flight guidance and EOL end of life control systems EUROCAE European Organization for AGL above ground level Civil Aviation Equipment AHRS attitude heading reference EVS enhanced vision system system FAA Federal Aviation Amdt. Amendment Administration APU auxiliary power unit FADEC full authority digital engine ATC air traffic control control BIM blade inspection method FAR Federal Aviation Regulations CAM cockpit area microphone [OBSOLETE term.]

CAR Civil Air Regulations FEM finite element modeling CAS calibrated air speed FHA functional hazard assessment CBIM cockpit blade inspection FMEA failure mode and effects method analysis CDP critical decision point FPM feet per minute CFR Code of Federal Regulations FPS feet per second CG center of gravity FTA fault tree analysis COTS commercial off-the-shelf GPS global positioning system CPS cycles per second GW gross weight CRFS crash resistant fuel system HEC human external cargo CRI certification review item HF high frequency CRT cathode ray tube HIRF high intensity radiated fields CVR cockpit voice recorder HRD high rate of discharge DER designated engineering HUMS health and usage monitoring representative systems DME distance measuring HV height-velocity equipment IAS indicated airspeed EASA European Aviation Safety ICA instructions for continued Agency airworthiness ECAS engine caution advisory ICAO International Civil Aviation systems Organization ECU environmental control unit ICS inter-communication system EDS electronic data system IFR instrument flight rules Page xxv AC 29-2C, Chg 7 2/4/2016 PCDS personnel carrying device IGE in ground effect Systems IIDS integrated instrument display system PCF post crash fire ILS instrument landing system PIO pilot induced oscillation IMC instrument meteorological PSA preliminary safety assessment conditions PSIG pounds per square inch INS inertial navigation system gauge ISA international standard QPL qualified parts list atmosphere RFM rotorcraft flight manual RFMS rotorcraft flight manual ISIS integral spar inspection system supplement ITT inter-turbine temperature KCAS knots calibrated airspeed RGL Regulatory and Guidance KIAS knots indicated air speed Library KTAS knots true airspeed RPM revolutions per minute LDP landing decision point RTCA Radio Technical Commission LPV localizer performance with Of Aeronautics vertical guidance RVR runway visual range LRU line replaceable unit SAE Society of Automotive LWC liquid water content Engineers MC maximum continuous SAR search and rescue MCP maximum continuous power SAS stability augmentation system MEL minimum equipment list SCAS stability and control MG miscellaneous guidance augmentation systems MGB main gearbox SHP shaft horsepower MGT measured gas temperature S/N stress vs. number of cycles MGW maximum gross weight SRM structural repair manual MMEL master minimum equipment SSA system safety assessment list STC supplemental type certificate MSL mean sea level STOL short takeoff and landing MVD median volume diameter TBO time between overhaul NASA National Aeronautics and TC type certificate Space Administration TCDS type certificate data sheet NDI non-destructive inspection TDP takeoff decision point NM nautical mile TIA type inspection authorization NPRM notice of proposed TIR type inspection report rulemaking TOP takeoff power NTSB National Transportation TOT turbine outlet temperature Safety Board TSO technical standard order NVG night vision goggles TVP true vapor pressure NVIS night vision imaging systems VBIM visual blade inspection method OAT outside air temperature VFR visual flight rules OEI one engine inoperative VMC visual meteorological OGE out of ground effect PBA pitch bias actuator conditions Page xxvi

9/30/99 AC 29-2C CHAPTER 1. PART 21 FAR 21 - GENERAL CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS (Amendment 21-50) AC 21.16. § 21.16 SPECIAL CONDITIONS .

a. The Process . Chapter 2, Section 1, Paragraph 8 of the Type Certificate Handbook, Order 8110.4A, provides detailed guidance on the special conditions process. However, much of that material has been outdated with the implementation of the Aircraft Certification Directorate Program. Rotorcraft special conditions are processed through the Rotorcraft Standards Staff, ASW-110. That office will assure coordination with the affected agency and industry elements including the Assistant Chief Counsel. All comments will be considered and the disposition will be documented by the Rotorcraft Standards Staff. ASW-100 will issue the special conditions.

b. Basis for Development .

(1) Special conditions are justified on the basis of the existing Part 29 being inadequate or inappropriate due to novel or unusual design features of the rotorcraft to be certificated.

(2) The phrase “novel or unusual” as used in § 21.16 is a very relative term. As used hereafter in applying § 21.16 to justify the issuance of special conditions, “novel or unusual” will be taken with respect to the state of technology envisaged by the applicable airworthiness standards of this subchapter. It must be recognized that in some areas which will vary from time to time, the state of the regulations may somewhat lag the state of the art in new design because of the rapidity in which the state of the art is advancing in civil aeronautical design and because of the time required to develop the experience base needed by the FAA/AUTHORITY to proceed with general rulemaking.

Applicants for type certification of a new design have the opportunity to mitigate the impact of not knowing the precise airworthiness standards to be applied for “novel or unusual design features” by consulting with the FAA/AUTHORITY early in their certification planning when such features are suspected or known by the applicant to exist. It should also be recognized that, because of the intentional objective nature of the airworthiness standards of this subchapter, many new design features which might be thought of as “novel or unusual” may already be adequately covered by existing regulations, thus obviating the need to issue special conditions.

(3) Before proposing special conditions, the certification staff should very thoroughly analyze the existing regulations and assure they are inadequate or inappropriate in light of a new and novel design feature.

Page FAR 21 - 1 AC 29-2C 9/30/99 AC 21.31 § 21.31 TYPE DESIGN . The regulatory basis for requiring data to define the design is contained in § 21.31. This section is self-explanatory and broad enough in scope to give the certification staff access to sufficient data to determine compliance with Part 29.

AC 21.33 § 21.33 INSPECTION AND TESTS .

a. Applicant Responsibility . Section 21.33 requires the applicant to: (1) Assure the test rotorcraft conforms to the type design. This must be accomplished prior to presentation to the FAA/AUTHORITY for testing.

(2) Conduct all inspections and tests necessary to determine compliance with the airworthiness and noise requirements.

b. FAA/AUTHORITY Responsibility .

(1) The design evaluation engineers should assure that the type design is adequate in their technical area and that the inspections and tests to be conducted are appropriate and sufficient to show compliance with Part 29.

(2) As changes to the rotorcraft are made during the test program, the flight test crew should assure that the appropriate design evaluation engineer concurs with the change and the conformity inspection of the change has been conducted.

AC 21.35 § 21.35 (Amendment 21-59) FLIGHT TESTS .

a. Explanation .

(1) This section outlines the requirements of the applicant for aircraft type certification and should be used in conjunction with FAA Order 8110.4A, Section 5.

Section 21.35 requires, in part, that the applicant conduct sufficient flight tests to show compliance with the flight requirements throughout the proposed flight envelope. The results of the applicant’s flight test should be submitted to the FAA/AUTHORITY in report form for evaluation to determine what verification flight tests the FAA/AUTHORITY may elect to conduct. The report should conclude that in the applicant’s opinion the test aircraft complies with the applicable certification requirements. The FAA/AUTHORITY verification flight test should include, but not be limited to, the critical or marginal results contained in the applicant’s flight test report.

The FAA/AUTHORITY’s role in the certification effort is not envisioned to be one of conducting day-to-day routine flight tests with the applicant, but only to verify his results through limited sampling. In certain tests, such as high altitude testing at a remote mountain site, there is an advantage in conducting flight tests concurrently with the applicant. Additionally, the FAA/AUTHORITY can provide technical flight test Page FAR 21 - 2 9/30/99 AC 29-2C assistance to the applicant in certain cases. This can be done after a cursory review and a letter of authorization is issued to the flight test crew.

(2) Preflight Test Planning . After the applicant’s flight test report is reviewed, it should be determined what FAA/AUTHORITY engineering flight tests are necessary.

These tests are normally specified in the Type Inspection Authorization (TIA). At the same time the FAA/AUTHORITY must know and agree to the applicant’s proposed means of data acquisition, reduction, and expansion of the flight test data. The adequacy of the test instrumentation should be evaluated prior to official type certification tests (reference paragraph AC 21.39).

(3) Order of Testing . The Federal Aviation Regulations are so worded that the results of some flight tests have a definite bearing on the conduct of other tests. For this reason, and to minimize retesting, careful attention should be given to the order of testing. The exact order of testing will be determined only by considering the particular rotorcraft and test program involved. Tests which are particularly important in the early stages of the program are: (i) Airspeed calibration: All tests involving airspeed depend upon the calibration.

(ii) Engine power available determination.

(iii) Engine cooling.

(4) Test Groupings .

(i) Weight and c.g .: In addition to the regulatory relationship of one test to another, efficient testing requires that consideration be given to the accomplishment of as many tests on a single flight as can be accommodated successfully.

(ii) Special Instrumentation . Similarly, consideration should be given to grouping of tests that involve special instrumentation. Examples of these are takeoff and landing tests which usually require group equipment to record horizontal distance, height, and time. Ground calibration of the airspeed indicating system can be accomplished at the same time. It is the applicant’s responsibility to provide the necessary instrumentation.

(5) Functional and Reliability Testing (i) Section 21.35(b)(2) requires that the applicant determine that “there is reasonable assurance that the aircraft, its components, and its equipment are reliable and function properly.” Section 21.35(f)(1) requires a Function and Reliability (F&R) program of 300 hours for turbine engine powered aircraft incorporating engines of a type not previously used in a type certificated aircraft. Section 21.35(f)(2) requires a 150-hour F&R program for all other aircraft. The following reflects general practices that Page FAR 21 - 3 AC 29-2C 9/30/99 have been used during rotorcraft certification programs. FAA/AUTHORITY have supported proposals which gave F&R test time credit for certification testing in lieu of dedicated F&R testing. In establishing such credit, the following should be considered: (A) The point in time in which the rotorcraft reaches substantial conformity with the approved type design.

(B) The extent and complexity of the new design.

(C) For a previously certified rotorcraft, the F&R program requirement should be commensurate with the modification or change in type design and may be zero.

(ii) Historically, for major rotorcraft type certification programs, flight time credit has been limited so as to require an irreducible minimum of 50 hours of dedicated F&R flight time. For rotorcraft programs that involved new engine installations (mature engine design) or drive train/rotor system changes on previously certified aircraft (TC amendments or STC’s), flight time credit has been liberal and often resulted in very little or no dedicated F&R testing.

b. Procedures .

(1) Type Certification Flight Tests .

(i) Prior to initiating official FAA/AUTHORITY flight tests, a conformity inspection of the test aircraft must be accomplished. This is needed to assure that the test aircraft is in the proper configuration or “conforms” to the engineering drawings and documents that have been submitted to FAA/AUTHORITY, evaluated, and approved. It is absolutely essential to know the configuration being tested in any engineering flight evaluation. Conformity inspection prior to TIA flight tests assures that testing will not be wasted because of configuration uncertainties.

(ii) FAA Order 8110.4A, paragraph 67, contains a requirement that the applicant must keep the FAA/AUTHORITY advised of any configuration changes to the aircraft. The manufacturing inspector should keep the FAA/AUTHORITY flight test pilot apprised of any change which may affect safety of the test aircraft or may influence test results.

(iii) Results of the conformity inspection and the engineering flight test program must be documented. This is normally done in the Type Inspection Report (TIR). Results may be documented in any acceptable engineering format. The report should be in sufficient detail to clearly show how compliance with each appropriate section of the rules was determined.

(iv) The flight test pilot must assure that the FAA/AUTHORITY manufacturing inspector and certification engineer are aware of all configuration Page FAR 21 - 4 9/30/99 AC 29-2C changes found necessary as a result of FAA/AUTHORITY tests. The manufacturing inspector is responsible for assuring that all changes are incorporated into production drawings after the design data reflecting the change have been approved by the certification engineer.

(v) Additional flight test responsibilities, procedures, and requirements during the certification flight test process are contained in FAA Order 8110.4A, Section 5, Flight.

(2) Function and Reliability Tests .

(i) A comprehensive and systematic check of all aircraft components must be made to assure that they perform their intended function and are reliable.

(ii) F&R testing should be accomplished on an aircraft which conforms to the type design. Non-conformities must be documented and accepted. F&R testing should follow the type certification testing described in paragraph AC 21.35b(1) above to assure that significant changes resulting from type certification tests are incorporated on the aircraft prior to F&R tests.

(iii) All components of the rotorcraft should be periodically operated in sequences and combinations likely to occur in service. Ground inspections should be made at appropriate intervals to identify potential failure conditions; however, no special maintenance beyond that described in the aircraft maintenance manual should be allowed.

(iv) A complete record of defects and failures should be maintained along with required servicing of aircraft fluid levels. Results of this record should be consistent with inspection and servicing information provided in the aircraft maintenance manual.

(v) A certain portion of the F&R test program may focus on systems, operating conditions, or environments found particularly marginal during type certification tests.

(vi) A substantial portion of the flying should be on a single aircraft. The flying should be carried out to an intensive schedule on an aircraft that is very close to the final certification standard, operated and maintained as though it were in service. A range of representative ambient operating conditions and sites should be considered. It is acceptable for non-F&R flight testing conducted at various sites and in varying ambient conditions to be used to satisfy the F&R requirements for those conditions.

Page FAR 21 - 5 AC 29-2C 9/30/99 AC 21.39. § 21.39 (Amendment 21-59) FLIGHT TEST INSTRUMENT CALIBRATION AND CORRECTION REPORT .

a. Explanation . It is the applicant’s responsibility to provide instrumentation for all parameters needed to show compliance with the airworthiness regulations.

(1) For those data which are necessary to show compliance with the regulations, a permanent record should be established. A permanent record is acceptable in either graphical or photographic form, and in some instances, a manual recording may be satisfactory.

(2) Regardless of the record form, the accuracy of the record must be established by reference to a laboratory standard traceable to the National Bureau of Standards.

(3) If multiplexing is used, the time base must be synchronized to a reference point from which the magnitude of each parameter can unquestionably be determined.

Also, the sampling rate should be sufficiently frequent to assure that the maximums, minimums, and trends of magnitude of the parameter are recorded with respect to time.

b. Procedures . Prior to conducting flight tests, the FAA/AUTHORITY flight test team should review the applicant’s flight test instrumentation calibration and correction report.

(1) Normally the frequency of instrument calibration should not exceed 90 days.

However, the frequency of recalibration varies with the consistency of the instrumentation under consideration. For example, cyclic and collective position is sometimes calibrated immediately before and after a flight where these parameters are used to provide critical flight data. Six months is a typical interval for recording/signal conditioning and nonstrain gage sensors, while one year is typical for strain gauged components. Also, environmental effects such as vibration, humidity, temperature, etc., should be considered when determining whether recalibration is necessary.

(2) The highest and lowest magnitude of the parameter being recorded should be considered when establishing the scale for instrumentation. Ideally, the highest magnitude throughout the flight would fall on the maximum indicating point of the recording.

Page FAR 21 - 6

9/30/99 AC 29-2C CHAPTER 2. PART 29 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT SUBPART A - GENERAL AC 29.1. § 29.1 (Amendment 29-21) APPLICABILITY .

a. Explanation . This section prescribes the rotorcraft categories eligible for certification under this part. There is no minimum weight limit for certification under Part 29; however, Part 27 is applicable to rotorcraft with maximum weights of 6,000 pounds or less so that Part 29, in effect, deals with rotorcraft which have a maximum weight greater than 6,000 pounds. In Part 29, there are two categories of rotorcraft, Category A and Category B.

(1) Category A . Category A provides the most rigid rules, requiring multiengine design with independent engines, fuel systems, and electrical systems. Category A design requires that no single failure can cause loss of more than one engine. Although there is no limit on maximum weight, Category A rotorcraft are certificated at a weight which will assure a minimum climb capability in the event of engine failure and with adequate surface area to assure a safe landing in the event an engine fails early in the takeoff run.

(2) Category B . Category B rotorcraft may be single or multiengine and may not have a maximum weight greater than 20,000 pounds. Category B rotorcraft are not required to have the capability for continued flight with an engine failed.

(i) Without Engine Isolation . For single engine rotorcraft and multiengine rotorcraft without engine isolation, the height-velocity diagram is conducted with sudden failure of all engines and the takeoff distance is measured through the clear area of the diagram to the 50-foot point with all engines operating. The landing distance is determined with all engines inoperative.

(ii) With Engine Isolation . Category B multiengine rotorcraft may be certificated with the Category A design features of Part 29. These rotorcraft meet the design requirements of Category A, but the performance requirements of Category B.

Stay-up ability after an engine failure is not assured. The takeoff is conducted with all engines operating, while the height velocity diagram and landing distances are determined with the most critical engine inoperative.

(3) Dual Certification, Categories A and B . A multiengine rotorcraft may be certificated under both categories provided requirements for both categories are met.

This combination will typically result in conditions (1) and (2)(ii) above with the primary differences being the gross weight allowed and the surface areas required for takeoff.

b. Procedures . None.

Page A - 1 AC 29-2C 9/30/99 AC 29.1A. § 29.1 (Amendment 29-39) APPLICABILITY .

a. Explanation . Amendment 29-39 revised the reference in § 29.1(e) from §§ 29.79 to 29.87, which is a redesignation of the section number for the height-velocity envelope. This section prescribes the rotorcraft categories eligible for certification under this part. There is no minimum weight limit for certification under Part 29; however, Part 27 is applicable to rotorcraft with maximum weights of 6,000 pounds or less so that Part 29, in effect, deals with rotorcraft which have a maximum weight greater than 6,000 pounds. In Part 29, there are two categories of rotorcraft. Category A and Category B.

(1) Category A . Category A provides the most rigid rules, requiring multiengine design with independent engines, fuel systems, and electrical systems. Category A design requires that no single failure can cause loss of more than one engine. Although there is no limit on maximum weight, Category A rotorcraft are certificated at a weight which will assure a minimum climb capability in the event of engine failure and with adequate surface area to assure a safe landing in the event an engine fails anywhere in the flight envelope, including takeoff or landing operations.

(2) Category B . Category B rotorcraft may be single or multiengine and may not have a maximum weight greater than 20,000 pounds. Category B rotorcraft are not required to have the capability for continued flight with one engine inoperative.

(i) Without Engine Isolation . For single engine rotorcraft and multiengine rotorcraft without engine isolation, the height-velocity diagram is conducted with sudden failure of all engines and the takeoff and landing distances are measured with all engines operating.

(ii) With Engine Isolation . Category B multiengine rotorcraft may be certificated with the Category A design features of Part 29. These rotorcraft meet the design requirements of Category A but the performance requirements of Category B.

Stay-up ability after an engine failure is not assured. The takeoff distance is determined with all engines operating. The landing distance, at the option of the applicant, may be determined with the critical engine inoperative or with all engines operating. The height-velocity diagram is determined following failure of the most critical engine.

(3) Dual Certification, Categories A and B . A multiengine rotorcraft may be certificated under both categories provided requirements for both categories are met.

This combination will typically result in conditions (1) and (2)(ii) above with the primary differences being the gross weight allowed and the surface areas required for takeoff.

b. Procedures . The guidance material in paragraph AC 29.1 does not apply to rotorcraft certified with Amendment 29-39 or later.

Page A - 2 9/30/99 AC 29-2C AC 29.2. § 29.2 (Amendment 29-32) SPECIAL RETROACTIVE REQUIREMENTS .

a. Explanation .

(1) Amendment 29-32 requires a combined shoulder harness and safety belt (also called a torso restraint system) at each occupant’s seat for all rotorcraft manufactured after September 16, 1992.

(2) The design features of the restraint system are mainly contained in this section rather than having to refer to other sections within Part 29 except for a general reference to the differing strength standards between earlier static strength only standards and the static and dynamic strength standards of Amendment 29-29.

(3) Combined safety belt and harness strength standards system follows: (i) Those rotorcraft type designs certificated to static strength standards alone prior to Amendment 29-29, such as 4 g’s forward may use belt and harness systems, characterized as 1,500 pounds strength systems, provided they comply with those standards. TSO C22f and earlier restraint systems have such ratings. A combined belt and harness with a 1,500 pounds rating, which comply with the Part 29 standards for the rotorcraft type design, but are not necessarily TSO approved, may be approved as a part of the type design. Such design information for a non-TSO’d item would be included in a note on the aircraft type certificate data sheet (TCDS) or specification sheet by part number as “required equipment.” TSO C114-approved torso restraint systems, characterized as 3,000 pounds strength system, may be used provided the design features comply with this section, but no special information on the TCDS is necessary.

(ii) Those rotorcraft type designs certified to dynamic test requirements of Amendment 29-29 should use torso restraint systems approved under TSO C114 or approved under equivalent standards such as those contained in Part 29.

(4) Load Distribution and Design Requirements . Although not stated in § 29.2, a 60 percent and 40 percent load distribution between the safety belt and harness, respectively, is required in § 29.785(g). The safety belt should withstand 100 percent if the safety belt is capable of being used alone. Also, the safety belt or harness attachments to the seat or structure should include the 1.33 factor described in § 29.785(f)(2) of Amendment 29-24 for those rotorcraft with that certification criteria or should include the 1.15 factor as described in § 29.625 (and predecessor § 7.355(c)(2) CAR Part 7) standards for those rotorcraft with the earlier certification criteria. A factor is used whether test results or analysis methods are used for static substantiation of the seating systems. Refer to paragraph AC 29.785b(1)(i) (§ 29.785).

Page A - 3 AC 29-2C, Chg 1 2/12/03 (5) The companion operating rule change of Amendment 91-220, amended § 91.205 (Amendment 91-223), affecting the aircraft equipment requirements.

Operating rule § 91.107(a) already requires use of the harness whenever the aircraft seat is so equipped.

b. Procedures .

(1) A TSO-approved combined safety belt and harness or torso restraint system may be used provided the installation requirements in § 29.2 are satisfied. A combined belt and harness (not necessarily TSO approved) may be approved as a part of the rotorcraft type design and so noted on the aircraft specification or TCDS.

(2) Structural analysis or static test may be used. For those rotorcraft designs that are subject to the dynamic test standards of § 29.562, the torso restraint system is required to be qualified for the particular use or installation in each rotorcraft type design. A dynamic test may be required for alternate restraint systems as well as the originally approved system. TSO C114 approval does not constitute approval for installation of a restraint system in a rotorcraft design subject to dynamic tests.

(i) Paragraph 27.562 of this AC concerns in part the dynamic test standards of Amendment 29-29.

(ii) AC 23-4 dated June 20, 1986, concerns static test procedures for small airplane seats and restraint systems. (Certain small airplanes manufactured after December 12, 1986, should have harnesses for each seat also.) A test proposal for rotorcraft installations may adopt procedures appropriate to the particular installation.

The 60/40 percent distribution is sufficiently achieved when the blocks in Figure 4 of AC 23-4 are used.

(iii) The static design side load for the harness installation may be proven by test or analysis using the load distribution previously noted. For “older” designs, the side load of § 29.561(b)(3)(iii) is 2.0g, and for later designs (Amendment 29-29 and later), it is 8.0g.

Page A - 4 9/30/99 AC 29-2C CHAPTER 2. PART 29 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT SUBPART B - FLIGHT GENERAL AC 29.21. § 29.21 (Amendment 29-24) PROOF OF COMPLIANCE .

a. Explanation .

(1) This section provides a degree of latitude for the FAA/AUTHORITY test team in selecting the combination of tests or inspections required to demonstrate compliance with the regulations. Compliance must be shown for each combination of gross weight, center of gravity, altitude, temperature, airspeed, rotor RPM, etc.

Engineering tests are designed to investigate the overall capabilities and characteristics of the rotorcraft throughout its operational envelope. Testing will identify operating limitations, normal and emergency procedures, and performance information to be included in the FAA/AUTHORITY-approved portion of the flight manual. The testing must also provide a means of verifying that the rotorcraft’s actual performance, structural design parameters, propulsion components, and systems operations are consistent with all certification requirements.

(2) Section 21.35 requires, in part, that the applicant show compliance with the applicable certification requirements, including flight test, prior to official FAA/AUTHORITY Type Inspection Authorization (TIA) testing. Compliance in most cases requires systematic flight testing by the applicant. After the applicant has submitted sufficient data to the FAA/AUTHORITY showing that compliance has been met, the FAA/AUTHORITY will conduct any inspections, flight, or ground tests required to verify the applicant’s test results. FAA/AUTHORITY compliance may be partially determined from tests conducted by the applicant if the configuration (conformity) of the rotorcraft can be verified. Compliance may be based on the applicant’s engineering data, and a spot check or validation through FAA/AUTHORITY flight tests. The FAA/AUTHORITY testing should obtain validation at critical combinations of proposed flight variables if compliance cannot be inferred using engineering judgment from the combinations investigated.

(3) Performance tests include minimum operating speed (hover), takeoff and landing, climb, glide, height-velocity, and power available. Certain other performance tests, such as Category A, are conducted to meet specific requirements. Detailed performance test procedures and allowable extrapolation or simulation limits are contained in the respective paragraphs in this order.

(i) Hover tests are conducted to determine various combinations of altitude, temperature, and gross weight for both in-ground-effect (IGE) and, if required, Page B - 1 AC 29-2C, Chg 4 9/17/2009 out-of-ground effect (OGE) conditions. From these data the hover ceiling may be calculated.

(ii) Takeoff and landing tests are conducted to determine the total distance to takeoff and land at various combinations of altitude, temperature, and gross weight.

(iii) Climb tests establish the variations of rate-of-climb at the best rate-of-climb or published climb airspeed(s) at various combinations of altitude, temperature, and gross weight.

(iv) Height-velocity tests are conducted to determine the boundaries of the height versus airspeed envelope within which a safe landing can be accomplished following an engine failure.

(v) Power available tests are conducted to verify or reestablish the calculated installed specification engine performance model on which published performance is based.

(4) The purpose of rotorcraft stability and control tests is to verify that the rotorcraft possesses the minimum qualitative and quantitative flying qualities and handling characteristics required by the applicable regulations. In order to assess the handling qualities, standardized test procedures must be utilized and the results analyzed by accepted methods. Section 29.21(a) allows calculation and inference which includes extrapolation and simulation, whereas § 29.21(b) requires demonstration of controllability, stability, and trim. Combinations of §§ 29.21(a) and 29.21(b) may be used to show compliance to the operating envelope limits. Test methods and equipment are described in individual paragraphs of this advisory circular.

b. Procedures .

(1) Efforts should begin early in the certification program to provide advice and assistance to the applicant to insure coverage of all certification requirements. The applicant should develop a comprehensive test plan which includes the required instrumentation.

(2) The tests and findings specified in paragraph a(3) above are required of the applicant to show basic airworthiness and probable compliance with the minimum requirements specified in the applicable regulations. After these basic findings have been submitted and revi 2 ewed, a Type Inspection Authorization, or equivalent, can be issued. The FAA/AUTHORITY will develop a systematic plan to spotcheck and confirm that compliance with the regulations has been shown. The test plan will consider combinations of weight, center of gravity, RPM and cover the range of altitude and temperature for which certification is requested.

Page B - 2 9/30/99 AC 29-2C AC 29.21A. § 29.21(Amendment 29-39) PROOF OF COMPLIANCE.

a. Explanation . Amendment 39 added § 29.83 which changes the requirements for determination of landing distance for Category B rotorcraft. This amendment requires landing distance to be determined with all engines operating within approved limits.

b. Procedures . The guidance material presented in paragraph AC 29.21 continues to apply.

AC 29.25. § 29.25 (Amendment 29-12) WEIGHT LIMITS.

a. Explanation .

(1) This section is definitive and specifies criteria for establishing maximum and minimum certificating weights. These weights may be based on those selected by the applicant, design requirements, or the limits for which compliance with all applicable flight requirements has been shown.

(2) Typical requirements that may establish the maximum and minimum weight limits include: Maximum: Structural limits, performance requirements, stability, and controllability requirements.

Minimum: Autorotative rotor RPM, stability, and controllability requirements.

(3) Jettisonable External Cargo .

(i) Paragraph (c) was added by Amendment 29-12 to provide, in the certification standards, a basis for approving an increase in gross weight (exceed standard limits) that would be an external jettisonable load. The attachment device standards were moved from Part 133 (Amendment 133-5) to Parts 27 and 29.

Section 29.865, “External load attaching means,“ now contains the standards, including design features, for the attaching devices. Cargo hoists and hooks were envisioned.

Prior to these amendments, type design approvals were made under Part 133 and the policy in Review Cases Nos. 37 and 55 of FAA Order 8110.6 whenever the standard limits were exceeded.

(ii) In the preamble of Amendment 29-12 (Proposal 2-99, 41 FR 55454, December 20, 1976) the agency stated, in part, that “...§ 29.25(c) is intended to provide only a total weight standard for approving the rotorcraft structure (and propulsion systems) for operation under Part 133.” As indicated in § 29.865, fatigue substantiation of the external cargo attaching means is not required. The rotorcraft structure, rotors, transmissions, engines, etc., are subject to evaluation under § 29.571 for external cargo approval whenever the “standard” structural limitations are exceeded (Review Case Nos. 37 and 55).

Page B - 3 AC 29-2C, Chg 4 9/17/2009 (iii) Whether or not the standard limitations are exceeded, the flight characteristics evaluations/standards of § 133.41 are appropriate even for engineering approval. This Part 133 standard is also applicable for the individual operator to obtain his operating certificate. The operator may use an FAA/AUTHORITY approved RFM supplement for external load operations to prepare a rotorcraft load combination flight manual required by § 133.47.

b. Procedures .

(1) It may not be possible to demonstrate quantitatively all the flight requirements at the minimum weight because of test instrumentation requirements. The test team must be assured that the rotorcraft complies with the applicable requirements at the lowest permissible flying weight. This evaluation may be done qualitatively, with the test instrumentation removed, and with minimum crewmembers if no critical areas exist or are anticipated. Additionally, reasonable extrapolation may be warranted. However, if critical areas at minimum flying weights are apparent, extrapolation should not be permitted.

(2) Whenever a gross weight increase (§ 29.25(c)) is requested, a TIA evaluation is necessary to evaluate the new limitations and ensure that § 133.41 for typical or representative cargo shapes and weights (density) is satisfactory. All possible combinations of weights and shapes are not evaluated. The representative configurations may be noted in the RFM or RFM supplement for the operator’s information. Sections 133.41 and 133.47 must be satisfied by the individual operator for the particular case at hand. The approved RFM or RFM supplement should provide the necessary limitations and any other information about the representative cargo configurations evaluated. Section 133.41 also permits the operator to obtain approval of additional and unique cargo configurations provided the approved limitations are observed. Paragraph AC 29.1581 concerns the RFM and its contents.

(3) See paragraph AC 29.571, § 29.571, for fatigue substantiation and external cargo considerations.

(4) Refer to AC 133-1A, Rotorcraft External-Load Operations in Accordance with FAR Part 133, October 16, 1979, for further information on airworthiness and flight manual policy.

AC 29.25A. § 29.25 (Amendment 29-51) WEIGHT LIMITS.

a. Explanation .

(1) Amendment 29-51 added a new paragraph (a)(4) that requires that the operating envelope for the controllability demonstrated under § 29.143(c) be included in the limitations section of the Rotorcraft Flight Manual (RFM). The change allows, in addition to the 17-knot controllability requirements, the applicant to provide additional Page B - 4 9/17/2009 AC 29-2C, Chg 4 controllability information within an applicant selected limited azimuth range if the rotorcraft is certified with nine or less passenger seats. This effectively allows increased weights within this limited range. Amendments 29-21 and 29-24 allowed for this relief and subsequent regulatory policy recognized these limitations as they are now required.

In no case should those limits be established at an altitude that is not operationally suitable. In the past, the minimum operationally suitable altitude for takeoff and landing has been established as 3,000 feet density altitude.

(2) The explanation regarding the relief for presentation of hover controllability limits in AC 29.143.a.(2)(ii) (Amendment 29-24) is superseded by this change.

b. Procedures . The policy material pertaining to the procedures outlined in this section remain in effect.

AC 29.27. § 29.27 (Amendment 29-3) CENTER OF GRAVITY LIMITS.

a. Explanation .

(1) This regulation is definitive and requires that the center of gravity limits be defined. Proof of compliance with all applicable flight requirements is required within the range of established CG’s. Along with the longitudinal CG limits, the lateral CG limits should either be established or determined to be not critical.

(2) Ballast is usually carried during the flight test program to investigate the approved gross weight/center of gravity limits. Lead is the most commonly used form of ballast during rotorcraft flight testing although other types of ballast, such as water, may serve just as well. Water may have the added benefit of being jettisonable during critical flight test conditions. Care must be taken regarding the location of ballast. The strength of the supporting structures should be adequate to support such ballast during the flight loads that may be imposed during a particular test and for the ultimate inertia forces of § 29.561(b)(3). Of critical importance is the method of securing the ballast to the desired locations. To avoid any undesired in-flight movements of the ballast, a positive method of constraint is mandatory. The flight test crews should also visually verify the amount, location, and integrity of the ballast. The effects of mass moment of inertia on the flight characteristics due to the ballast locations should also be considered. The mass moment of inertia of the test rotorcraft should, to the extent possible, be the same as that expected in normal, approved loadings, especially during tests involving dynamic inputs.

b. Procedures .

(1) Center of gravity locations and limits are of prime importance to rotorcraft stability and safety of flight. The primary concern is establishment of the longitudinal center of gravity limits. Lateral center of gravity limits with respect to longitudinal center of gravity limits are also important. The design of the rotorcraft is usually such that approximate lateral symmetry exists. This lateral symmetry can be upset by lateral Page B – 5 AC 29-2C, Chg 4 5/1/2014 loadings resulting in the necessity to establish lateral center of gravity limits. There are two characteristics which may be seriously affected by loading outside the established center of gravity limits; these are stability and control. The established center of gravity limits must be such that as fuel is consumed, it is possible for the rotorcraft to remain within the established limits by acceptable loading and/or operating instructions.

(2) Structural limits may restrict the maximum forward longitudinal center of gravity limits. However, in most cases it is the maximum value established wherein adequate low speed control power exists to meet such requirements as § 29.143(c).

Likewise, the maximum aft center of gravity limit may be a “structural limit,” but it usually is determined during flight test after the rotorcraft’s handling qualities tests have been conducted. Additional items which may influence the maximum aft center of gravity limits may be malfunctions of automatic stabilization equipment, excessive rotorcraft attitudes during critical phases of flight, or adequate control power to compensate for an engine failure.

(3) Lateral center of gravity limits have become more critical because of the ever increasing utilization of the rotorcraft for such things as unusual and unsymmetric lateral loads, both internal and external. Maximum allowable lateral center of gravity limits have also influenced the results of the unusable fuel determination.

(4) Summarizing, it is of prime importance that longitudinal and lateral center of gravity limits be determined so that unsafe conditions do not exist within the approved altitude, airspeed, ambient temperature, gross weight, and rotor RPM ranges. All relevant malfunctions must be considered.

AC 29.29. § 29.29 (Amendment 29-15) EMPTY WEIGHT AND CORRESPONDING CENTER OF GRAVITY.

a. Explanation . The empty weight of the rotorcraft consists of the airframe, engines, and all items of operating equipment that have fixed locations and are permanently installed (including both required and optional equipment) in the rotorcraft. It includes fixed ballast, unusable fuel, other unusable fluids, and full operating fluids. “Full operating fluids” such as oils used in an engine, auxiliary power unit, main and auxiliary gearboxes, and hydraulic systems are considered “closed fluid systems” typically filled to a “full mark” indicator level. Fluids necessary for the operation of non-permanently installed equipment (i.e., carry-on equipment) are not considered part of the empty weight.

(1) A ballast is fixed when made a permanent part of the rotorcraft as a means of controlling the certificated empty weight center of gravity (CG).

(2) Installed equipment is any FAA-approved equipment attached to the rotorcraft with hardware and, as a result, becomes an integral part of the rotorcraft. The installation or removal of such equipment must be recorded in the aircraft equipment list. Compliance with paragraph (b) of § 29.29 is accomplished by the use of an Page B – 6 5/1/2014 AC 29-2C, Chg 4 equipment list specifying the installed equipment at the time of weighing and the weight moment arm of the equipment.

b. Procedures .

(1) Determination of the empty weight and corresponding center of gravity is primarily the responsibility of the manufacturer and is normally made on a production rotorcraft rather than a prototype. If the manufacturer has been issued a production certificate and wishes to avoid weighing each production rotorcraft, the manufacturer may make a detailed proposal defining the procedure it would use to establish an empty weight and CG. When the proposal is approved, the manufacturer will weigh the first five to ten production rotorcraft and show that the rotorcraft will be within ±1 percent on empty weight and ±0.2 inches on CG. After this procedure is established, the empty weight and CG may be computed except that at regular intervals a rotorcraft will be weighed to ensure the tolerances are still being maintained (e.g., one in ten rotorcraft).

(2) For prototype and modified rotorcraft, it is only necessary to establish a known basic weight and CG position (by weighing) from which the extremes of weight and CG travel required by the test program may be calculated. See the current version of FAA-H-8083-1 ( Pilots Weight and Balance Handbook ) for a sample weight and balance procedure.

(3) The weight and balance should be recalculated if a modification (or series of modifications) to the rotorcraft results in a significant change to the empty weight.

Additionally, this change in empty weight should be reflected with the weight and balance information contained in the rotorcraft flight manual or rotorcraft flight manual supplement.

c. Ballast Loading and Type .

(1) Ballast loading of the rotorcraft can be accomplished in any manner to achieve a specific CG location. It is acceptable for such ballast to be mounted outside the physical confines of the rotorcraft if the flight test objectives are not affected by this arrangement. In flight test work, loading problems will occasionally be encountered in which it will be difficult to obtain the desired CG limits. Such cases may require loading in engine compartments or other places not designed for load carrying. When this condition is necessary, care should be taken to ensure that local structural stresses are not exceeded or that the rotorcraft flight characteristics are not changed due to increased moments of inertia by attaching the ballast to extreme CG locations that may not be designed for the added weight.

(2) The two types of ballasts that may be used in loading are solid or liquid. The solids are usually high density materials such as lead, while the liquid usually used is water. In critical tests, the ballast may be loaded in a manner so that disposal in flight can be accomplished. In any case, the load should be securely attached in its loaded position so shifting or interference with safety of flight will not result.

Page B – 7 AC 29-2C, Chg 4 5/1/2014 AC 29.31. § 29.31 REMOVABLE BALLAST.

a. Explanation . This regulation provides the option of using removable ballast for operational flights to obtain center of gravity locations that are in compliance with the flight requirement of this Part. Fixed ballast used for flight operations after type certification must be documented in the type design data. Removable ballast is used primarily on small rotorcraft to control the CG with different passenger loadings although this regulation does permit its use on transport rotorcraft. If removable ballast is used, the rotorcraft flight manual must include instructions regarding its use and limitations.

See paragraph AC 29.873 for information on ballast provisions.

b. Procedures . None.

AC 29.33. § 29.33 (Amendment 29-15) MAIN ROTOR SPEED AND PITCH LIMITS.

a. Explanation .

(1) General . This rule requires the establishment of power-on and power-off main rotor speed limits and the requirements for low rotor speed warning.

(2) Power-On . The power-on limits should be sufficient to maintain the rotor speed within these limits during any appropriate maneuver expected to be encountered in normal operations throughout the flight envelope for which certification is requested.

A power-on range of approximately 3 percent has in the past been the minimum range required due to engine governor and engine operating characteristics. With the introduction of advanced engines and electronic engine controls, there may not be a need for a range, but one fixed value may suffice. Transient power-on values may also be acceptable provided they are substantiated.

(3) Power-Off . The power-off rotor speed limits should be sufficient to encompass the rotor speeds encountered during normal autorotative maneuvers except for final landing phase (touchdown) for which rotor RPM may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. The limits should also be sufficient to cover the ranges of airspeed, weight, and altitudes for which certification is requested. It is not the intent of the rule to require the minimum and maximum limit values in conjunction with extremes such as maximum/minimum weights and/or high altitude. The minimum and maximum rotor speed requirements should be thoroughly evaluated at normal operation environment; i.e., at altitudes between approximately sea level and 10,000 feet, temperatures not at extremes, and weights as necessary for other tests and as required to readily establish the limit rotor speeds.

Spot checks of the autorotative requirements should be made at the extremes of the flight envelope and environmental conditions during normal tests at those conditions.

Under conditions where high autorotative rotor speeds may be encountered, it is acceptable for the pilot to adjust the controls to prevent overspeeding of the rotor. At light weight combined with low altitudes and extreme cold temperatures, the normal low pitch setting may not be sufficient to maintain autorotational rotor speed values within Page B – 8 5/1/2014 AC 29-2C, Chg 4 limits. If this occurs, the manufacturer may elect to adjust the low pitch stops as a maintenance procedure at extreme ambient conditions provided the flight and maintenance manuals clearly present the rigging requirements and procedures. There must be sufficient “overlap” of ambient conditions between configurations such that rerigging is not required whenever ambient temperature and surface elevation change slightly. Any down rigging of the low pitch stop must continue to ensure adequate clearance between controls and other rotorcraft structure and should be evaluated during flight test. Both the power-on and power-off limits may also be established by encountering critical flapping limits in some approved flight conditions such as high airspeed or sideward flight.

(4) Additional RPM Ranges . Some applicants have elected to certify their aircraft with additional RPM ranges in an attempt to realize additional performance during certain flight conditions or maneuvers such as Category A OEI continued and rejected takeoffs and balked landings. Such additional RPM ranges have been found acceptable as long as all pertinent FAR requirements are fully substantiated for operation in that range. The substantiation should include drive system endurance and flight test verification of performance and flight characteristics during applicable maneuvers, in the additional RPM range. The FAA/AUTHORITY does not define additional RPM ranges as transient since all applicable requirements must be satisfied for approval of that range.

(5) Low Speed Warning . If it is possible under expected operating conditions for the rotor speed to fall below the minimum approved values, the requirement exists for a low rotor speed warning. This warning is required on all single-engine rotorcraft and on multiengine rotorcraft where there is not an automatic increase in remaining engine(s) power output upon failure of an engine. Although today’s multiengine rotorcraft do not require a low rotor speed warning according to the rule, essentially all have warning systems installed. If the minimum power-on and power-off rotor speed limits are different, the warning signal should be at the higher speed, normally the power-on minimum rotor speed. One rotorcraft has a warning system cutout if the collective is full down, and others have other warnings on the engine speed to indicate engine failure.

All of these related warning systems must be evaluated with emphasis on ensuring adequate rotor speed.

b. Determination and Testing . Refer to paragraph AC 29.1509 (§ 29.1509) for additional information on rotor limits determination and testing.

Page B – 9 AC 29-2C, Chg 4 5/1/2014 SUBPART B - FLIGHT PERFORMANCE AC 29.45. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.

a. Explanation .

(1) Changes to various part 29 sections, which did not include amending § 29.45, added new and revised airworthiness standards for the performance of transport category rotorcraft and renumbered several sections within the performance section of Subpart B. The performance section of this guidance material has been organized for easy use with rotorcraft certificated before or after this amendment. To achieve this, some of the guidance material has been duplicated under different paragraph numbers.

A statement at the beginning of each of these paragraphs indicates where other pertinent information can be found.

(2) Section 29.45 lists the rules and standards under which the performance requirements are to be met. This guidance provides general guidelines that may be used throughout a flight test program. It is impossible to find ideal test conditions and there are many variables that affect the flight test results that must be taken into account. Some of these variables are wind, temperature, altitude, humidity, rotorcraft weight, power, rotor RPM, center of gravity, etc. A thorough knowledge of the testing procedures and data reduction methods is essential and good engineering judgment must be used to determine applicable test conditions. The test results should be analyzed and expanded by an approved methodology. The guidance within this section is considered an approved methodology.

(3) Performance should be based on approved engine power as determined in paragraph b.(4) below and not on any transient limits. Approved transient limits are basically for inadvertent overshoots of approved operational limits. Any sustained operation in these transient limit areas usually require some form of special maintenance. However, for such demonstrations as rejected and continued OEI category A takeoffs and HV determination, low rotor speeds have been authorized based upon additional structural and drive system substantiation (see section 29.33 of this AC).

(4) Where variations in the parameter on which a tolerance is allowed will have an appreciable effect on the test, the results should be corrected to the standard value of the parameter; otherwise, no correction is necessary.

(5) As defined in 14 CFR § 1.1, the 30-second and 2-minute OEI power ratings are based on up to three periods of use during a single flight. The purpose of the three applications is: (i) to initially recover from an engine failure, (ii) to conduct OEI missed approach, and (iii) to conduct the final OEI landing. Rotorcraft performance based on Page B – 10 5/1/2014 AC 29-2C, Chg 4 the use of these time-limited power ratings is only permitted once in each of the above three uses (i.e., 30-second power must not be used more than once during the initial recovery from an engine failure).

(6) All engines operating (AEO) performance must be based upon approved AEO power ratings. OEI power ratings cannot be applied to an AEO condition.

b. Procedures .

(1) Winds For Testing .

(i) Allowable wind conditions will vary with the type of test and will also be different for different types and gross weight rotorcraft. For example, higher winds may be tolerated for takeoff and landing distance tests but not for hover performance.

Likewise higher winds may be tolerated during hover performance testing on large, heavy rotorcraft with high rotor downwash velocities than for smaller rotorcraft with lower rotor downwash velocities. Generally, unless the effects of wind on hover performance tests can be determined and accounted for, hover performance testing should be conducted in winds of 3 knots or less.

(ii) Past experience has shown that a steady wind of 0 to 10 knots will result in acceptable takeoff and landing performance if distances are corrected for the winds measured during these tests. This is not the case for vertical takeoffs and landings. To obtain consistent and repeatable vertical performance data, using wind speeds up to 5 knots is acceptable. In actuality, a rotorcraft may exhibit reduced IGE hover performance in wind speeds from 3 to 15 knots due to partial immersion of the main rotor in its own vortex. Since the height-speed envelope determination is affected by wind just as vertical takeoff and landing performance are, the same allowable winds for testing should be adhered to for HV testing; i.e., 0 to 5 knots. For category A testing, the effects of crosswind and tailwind should also be considered up to the maximum for which category A certification is requested.

(iii) As can be seen from the foregoing, there is no such thing as an exact allowable wind for a particular test or rotorcraft. The flight test team must decide on the allowable wind for each condition based on all available information and their engineering judgment. The following summary of allowable wind conditions is given for general guidance only: (A) Hover performance - 0 to 3 knots.

(B) Conventional takeoff and landing - 0 to 10 (data to be corrected) (C) Vertical takeoff and landing - 0 to 5 knots (D) HV - 0 to 5 knots Page B – 11 AC 29-2C, Chg 4 5/1/2014 (iv) A means should be provided to measure the wind velocity, direction, and ambient air temperature at the rotor height for any particular tests. The wind effects on required runway length for takeoff and landing distances may be shown in the flight manual.

(v) Full wind credit may be given for conventional takeoff and landing field lengths. This credit should not be more than the nominal wind component along the takeoff or landing path opposite to the direction of flight.

(2) Altitude Effects: Extrapolation and Interpolation .

(i) Using FAA/AUTHORITY approved methodology: (A) Hover, takeoff, and landing performance may be extrapolated from test data up to a maximum of ±4,000 feet density altitude from the test altitude.

(B) Experience has shown that IGE handling qualities, height-velocity, and engine operating characteristics may be extrapolated from test data up to a maximum of ±2,000 feet density altitude from the test altitude.

(C) Cruise stability and controllability tests should be evaluated at a minimum of two different altitudes, the lowest practical altitude and approximately the highest cruise altitude requested for approval. This can allow an interpolation of approximately 10,000 feet density altitude.

(ii) As in all testing, extrapolation or interpolation should only be considered if all available information and engineering judgment indicate that regulatory compliance can be met at the untested conditions.

(3) Altitude Limitations .

(i) Explanation . Two altitudes are normally presented in the RFM to define the operating envelope of a rotorcraft: maximum operating altitude and maximum takeoff and landing altitude.

(A) Maximum operating altitude is an operating limitation required by § 29.1527 and delineates the maximum altitude up to which operation is allowed. This altitude normally constitutes the maximum cruise or enroute altitude.

(B) Maximum weight, altitude, and temperature for takeoff and landing constitute a limitation. The maximum takeoff and landing altitude may be coincident with but never above the maximum operating altitude limitation. Takeoff and landing, hover ceiling data, and presentation requirements are presented in §§ 29.51, 29.53, 29.59, 29.63, 29.73, 29.1583, and 29.1587.

Page B – 12 5/1/2014 AC 29-2C, Chg 4 (ii) Procedures .

(A) In establishing the maximum takeoff and landing altitude, the following tests are normally required: ( 1 ) Takeoff (§§ 29.51-29.63) ( 2 ) Climb (§§ 29.64-29.67) ( 3 ) Performance at minimum operating speed (§ 29.49) ( 4 ) Landing (§ 29.75) ( 5 ) HV envelope (§ 29.87) ( 6 ) IGE controllability (§ 29.143(c)) ( 7 ) Cooling (§§ 29.1041-29.1045) ( 8 ) Engine operating characteristics (§ 29.939) Specific guidance on test methodology and data requirements is provided in applicable paragraphs of this guidance section.

(B) As detailed in paragraph b.(2) above, the maximum allowable extrapolation of HV, IGE controllability and engine operating characteristics is ±2,000 feet. Therefore, the maximum takeoff and landing altitude presented in the RFM is not normally more than 2,000 feet above the density altitude experienced at the high altitude test site.

(C) Prior to Amendment 29-21, HV information was an operating limitation.

With the adoption of Amendment 29-21, the HV curve is performance information for category B rotorcraft with nine or less passenger seats but remains a limitation for category A rotorcraft and category B rotorcraft with 10 or more passenger seats.

(D) Prior to Amendment 29-24, IGE controllability was required in 17 knots of wind to the maximum takeoff and landing conditions. With the adoption of Amendment 29-24, if IGE or OGE hover performance is presented for a category B rotorcraft to an altitude in excess of that for which IGE controllability at 17 knots is presented, the maximum safe wind demonstrated for hover operations must be presented in the RFM. The amendment did not change the requirement for category A rotorcraft.

(E) The requirements for data collection and presentation in the RFM vary depending upon the certification basis of the rotorcraft. These requirements are presented by regulation and amendment in Figures AC 29.45-1 and AC 29.45-2.

Page B – 13 AC 29-2C, Chg 4 5/1/2014 (F) The maximum takeoff and landing altitude may be extrapolated no greater than the values given in paragraph b.(2) above and not above the lowest limiting altitude resulting from the requirements listed in paragraph b.(3)(ii)(A) above.

(4) Temperature Effects .

(i) Background.

(A) The regulations prohibit any unsafe design feature throughout the range of environmental conditions for which certification is requested. The regulations also require that the performance and handling qualities be determined over the approved range of atmospheric variables selected by the applicant.

(B) Substantiation of temperature effects on performance and handling characteristics is required throughout the approved temperature range. In the past, approved analyses were frequently accepted for determining the extreme temperature effects on performance and flight characteristics. With the introduction of newer, higher performance rotorcraft, advanced rotor blade designs, higher airspeeds, and blade mach numbers, the previous methods have proven to be insufficient. Therefore, the performance and flight characteristics should be validated at extreme temperatures; however, analysis may be permitted if a suitable methodology is demonstrated.

(C) Various FAA/AUTHORITY cold weather programs have verified that rotorcraft can be affected, sometimes significantly, in both the performance and flying qualities areas. Hot temperature conditions although not shown to be as critical should be given consideration.

(D) Additionally, design deficiencies surfaced when the rotorcraft were exposed to temperature extremes and some of these difficulties were severe enough to require the redesign of equipment and materials. Therefore, to satisfy § 29.1309(a), the applicant needs to substantiate the total rotorcraft at the extreme temperatures for which certification is requested.

(ii) Procedures .

(A) The FAA/AUTHORITY is responsible for verifying the applicant’s predictions of performance and handling characteristics at the temperature extremes for which certification is requested. A limited flight verification, if necessary, could include spot checks of hover and climb performance, IGE controllability, roughness determination, simulated power failure, static stability, height-velocity, V /V NE D evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations.

Page B – 14 5/1/2014 AC 29-2C, Chg 4 (B) Extrapolation or interpolation of test data should only be allowed if the applicant’s predicted or calculated data is verified by actual test. Extrapolations or interpolations should not exceed 10°C below or 20°C above those values tested.

(5) Weight Effects . Test weights should be maintained within +3 percent and -1 percent of the target weight for each data point. Weight may be extrapolated or interpolated only along an established W/ σ line within the allowable altitude extrapolated range.

(6) Engine Power - Turboshaft Engine .

(i) Background .

(A) The purpose of rotorcraft performance flight testing is to obtain accurate quantitative flight test performance data to provide flight manual information.

(B) Flight tests are designed to investigate the overall performance capabilities of the rotorcraft throughout its operating envelope. This testing furnishes information to be included in the flight manual and provides a means of validating the predicted performance of the rotorcraft with a minimum installed specification engine.

(C) The horsepower used to complete the flight manual performance must be based on horsepower values no greater than that available from the minimum uninstalled specification engine after it is corrected for installation losses. A minimum uninstalled specification engine is one that, on a test stand under conditions specified by the engine manufacturer, will produce the certificated horsepower values at specification temperatures and speeds. The specification values may be either a rating or limit . Some engine manufacturers certify an engine to a specified horsepower at a particular engine temperature or speed rating with higher allowable limits . The limit is the maximum value the installed engine is allowed in order to develop the specification horsepower. Prior to installation of each engine in a rotorcraft, the performance is measured by the engine manufacturer. This is done by making a static test run in a test cell and referring the results to standard day, sea level conditions. The performance parameters obtained are presented as uninstalled engine characteristics on a test log sheet. This is commonly referred to as a “final run sheet.” Figure AC 29.45-3 compares a typical engine to one the manufacturer has certified as a minimum uninstalled certified engine.

(D) After engine certification, the engine manufacturer is responsible to ascertain that each engine delivered will produce, as a minimum, the certified horsepower values without exceeding specification operating values; therefore, a “final run sheet” is created for every engine produced. Additionally, if needed, arrangements can usually be made with the engine manufacturer to obtain a torque system calibration for individual engines. This will further optimize the accuracy of the engines used in the flight test program. The engine manufacturer will also provide predicted uninstalled power available for the various power ratings. This information may be derived from an Page B – 15 AC 29-2C, Chg 4 5/1/2014 engine computer “card deck” and from charts and tables in the engine detail installation manual. These data also provide engine performance for the range of altitudes and temperatures approved for the engine and include methods for correcting this performance for installation effects. The parameters contained in a typical “card deck” are plotted for one engine rating in Figure AC 29.45-4.

(E) Several installation losses (i.e., power decrements) may be associated with installing an engine in a rotorcraft. Typical losses are air inlet losses, gear losses, air exhaust losses, and powered accessory losses such as electrical generators.

Additional flight manual performance considerations are the torque indicating system accuracy and torque needle split. The predicted uninstalled power available engine characteristics cannot be assumed to be the actual power available after the engine is installed in the rotorcraft because this procedure would neglect the installation power losses. It is necessary to know the installation losses in order to determine the flight manual performance. Installation losses are reflected reductions in available horsepower resulting from being installed in a rotorcraft. These losses usually consist of those incurred due to engine inlet or exhaust design. The rotorcraft manufacturer usually conducts test to confirm the installed specification. Methods used vary widely between manufacturers, but usually include some combination of ground and flight tests. Figure AC 29.45-5 is a typical example of an installed power available chart for one set of conditions.

(F) This predicted installed power available is, in most cases, lower than obtained on a test stand. This is especially true at lower airspeeds where exhaust reingestion decreases the available horsepower output and changes in airflow routing.

The rotorcraft manufacturer may elect to determine the installation losses for different flight conditions to take any airspeed advantages. This is acceptable if, for example, the hover performance is based on the actual horsepower available from a minimum installed specification engine in a hover . Likewise, it is permissible for the rotorcraft manufacturer to determine his climb performance based on the actual horsepower available from a minimum installed specification engine at the published climb airspeed.

This will allow the manufacturer to take advantage of, for example, increased inlet efficiency.

(ii) Procedures .

(A) To this point the minimum installed specification engine horsepower output has been predicted and calculated for various flight conditions. It is imperative that the predicted values be verified by actual flight test. The flight test involves obtaining engine performance measurements at various power settings, altitudes, and ambient temperatures. The data should be obtained at the actual flight condition for which the performance is to be presented (i.e., hover, climb, or cruise).

(B) Following an initial application of power, engine temperature or RPM can significantly decrease for a period of time as torque is held constant. Said another way, torque will increase if RPM or temperature is held constant. This is a characteristic Page B – 16 5/1/2014 AC 29-2C, Chg 4 typical of turbine engines due largely to expansion of turbine blades and reduced clearances in the engine. Some engines may show a temperature increase at constant power due to engine or temperature sensing system peculiarities. An engine will usually establish a stabilized relationship of power parameters in approximately 2 or 3 minutes. For this reason, the following procedure should be used when obtaining in-flight engine data.

( 1 ) To determine the applicable value (takeoff, 30-second, or 2 1/2-minute power), the engine is first stabilized at a low power setting. After stabilization, rapidly increase the power demand to takeoff, 30-second and 2 1/2-minute power levels as necessary. Record the engine parameters as soon as the specification torque, temperature, or speed is attained. Care must be taken not to exceed a limit.

These readings should be obtained approximately 15 seconds after power is initially applied.

( 2 ) To determine the 30-minute and maximum continuous power values, approximately 2 to 3 minutes of stabilization time is generally used, but up to 5 minutes stabilization time is allowed. The reason for the different procedures is when a pilot requires takeoff or 2 1/2-minute power values he is in a critical flight condition and does not have the luxury of waiting for the engine(s) to produce rated power.

Stabilization time is allowed for the maximum continuous and 30-minute ratings because these values are not associated with flight conditions for which power is needed immediately. An engine may be certified to produce a specification horsepower at a particular temperature or engine speed rating with higher maximum limit value approved. Only the rating values should be used to determine the installation losses.

The limit values of engine temperature and speed are established and certified to allow specification powers to continue to be developed as the engine deteriorates in service.

(C) The in-flight measurements recorded with the engine(s) on the flight test rotorcraft must be corrected downward if the test engine is above minimum specification and corrected upward for a test engine that is below minimum specification. This correction is necessary to verify that a minimum installed specification engine installed on a production rotorcraft is capable of producing the horsepower values used to compute the flight manual performance without exceeding any engine limit. In addition, if the production rotorcraft’s power measurement devices have significant (greater than 3 percent) power error, this error must be accounted for in a conservative manner.

(D) On multiengine rotorcraft, the engine location may result in different installation losses between engines. If this condition exists, multiengine performance should be based on a total of the different minimum installed specification horsepower values. OEI performance must be based on the loss of the engine which has the lowest installation losses. Additionally, the power losses due to such items as accessory bleed air, particle separators, etc., must be accounted for accordingly.

Page B – 17 AC 29-2C, Chg 4 5/1/2014 (E) Power available data should be obtained throughout the test program at various ambient conditions. Some engines have devices which restrict the mechanical N speed to a constant corrected speed at cold temperatures. Others may G limit power to a minimum fuel flow value which would be encountered only at certain ambients. Others may limit by torque limiting devices. Therefore, power available data should be obtained at various ambients to verify that all limiting devices are functioning properly and have not been affected by the installation.

(F) Through use, turbine engine power capabilities decrease with time.

This is called engine deterioration. Deterioration is largely a function of the particular engine design, and the manner and the environment in which the engine is operated.

There is a need, therefore, to provide a method which can be used in service to periodically determine the level of engine deterioration. A power assurance curve is usually provided to allow the flightcrew to know the power producing capabilities of any engine. A power assurance check is a check of the engine(s) which will determine that the engine(s) can produce the power required to achieve flight manual performance.

This check does not have to be done at maximum engine power. Figure AC 29.45-6 is a typical power assurance curve for an installed engine showing minimum acceptable torque which assures that power is available to meet the RFM performance. Some power assurance curves have maximum allowable N limits that must not be exceeded G for a given torque value. An in-flight power assurance check may be used in addition to the pre-takeoff check. The validation of either check must be done by the methodology used to determine the installed minimum specification engine power available. For the in-flight power assurance check there must be full accountability for increased efficiency due to such items as inlet ram recovery, absence of exhaust reingestion, etc. A power assurance check done statically and one conducted in-flight must yield the same torque margin(s). An engine may pass power assurance at low power but still may not be capable of producing the rated horsepower values. This occurs when the curve of measured corrected horsepower and corrected temperature for the engine intersects the minimum uninstalled specification engine curve. If this condition exists, the entire power assurance and power available information may need to be reestablished.

(7) Deteriorated Engine Power - Turboshaft Engine .

(i) Background .

(A) A specific engine model may have been certificated for operation with power which has “normally” deteriorated below specification. This “normal” deterioration refers to a gradual loss in engine performance, possibly caused by compressor erosion, as opposed to a sudden performance loss which may be due to mechanical damage. The application for deteriorated engine power should not be confused with the installed mechanical engine derating which is frequently used to match transmission and engine power capabilities.

(B) The use of deteriorated power is intended to allow continued operations with an engine which is serviceable and structurally sound, although aircraft Page B – 18 5/1/2014 AC 29-2C, Chg 4 performance may be depreciated. The useful life of the engine may, therefore, be extended at a dollar savings to the operator.

(C) Although installed performance is the primary topic in this discussion, considerations must be given to other operational characteristics and systems which may be affected by depreciated engine power. These include: ( 1 ) Engine characteristics (§ 29.939). The reduced compressor discharge pressure, P , would reduce engine surge margin and possibly affect engine C response and engine air-restart capability. These items should be addressed, but flight testing may not be required depending on the individual engine or aircraft installation and fuel scheduling mechanism.

( 2 ) Performance of customer bleed air systems may be degraded slightly. No problem would be anticipated unless certain items within the system depend on a critical engine bleed air pressure for their function.

( 3 ) The maximum attainable gas producers speed, and thus power available under certain ambients, may be affected if engine bleed air pressure is an input to the fuel scheduling mechanism.

( 4 ) Systems for surge protection which schedule on engine bleed air pressure such as bleed valves, flow fences, bleed bands, and variable inlet guide vanes may be influenced. The affect would normally be negligible unless when installed, the installation losses combined with reduced engine bleed air pressure because of deterioration, would cause the bleed device to open and reduce power at any one of the engine ratings.

(ii) Procedures .

(A) The need for flight tests to verify predicted power available with deteriorated engines depends on the scope of testing which occurred during initial certification. If the original rotorcraft certification included flight testing as described in paragraph b.(6) (engine power-turboshaft engines) herein for validation of power available, the need for a demonstration with deteriorated engines, is greatly diminished and perhaps eliminated.

(B) If flight testing to verify deteriorated engine power available is deemed necessary, the procedure used would be the same as that described in paragraph b.(6) (engine power-turboshaft engines), except that the data would be corrected downward to a deteriorated engine runline. Efforts should concentrate on obtaining data in areas of the operational envelope where maximum gas producer speed is likely to be attained, or where bleed valves or other devices which schedule on gas producer discharge pressure are likely to function. On many installations maximum gas producer speed will occur cold and high; bleed valves and other devices which schedule on gas producer Page B – 19 AC 29-2C, Chg 4 5/1/2014 discharge pressure are most likely to function and reduce power on a hot day at low altitude.

(C) The adjustments to the normal power assurance check procedures for deteriorated engines will be influenced by the preferences of the aircraft manufacturer and by any special stipulations of the engine certification region established as a condition for the engine to remain in service when below specification. Possibly, more stringent and more complicated procedures will be introduced for deteriorated power; for example, an in-flight trend monitoring program with the associated bookkeeping duties may be required. Such an in-flight procedure must be evaluated by flight tests as described in paragraph b.(6) above. Normally, however, the manufacturer would be expected to present a modification, or extension of the power assurance procedure already in place for the specification engine, which could eliminate the need for flight test evaluation.

(D) If a complex power assurance procedure is presented with involved data reduction and trending requirements, consideration should be given to restricting the use of deteriorated power to operators where close control over operations is exercised or the operator has demonstrated the ability to operate safely with deteriorated engines.

(8) Engine Failure Testing Considerations (i) For all tests to examine behavior following an engine failure, usually the failure of the engine is simulated in some way. For engines with a hydro-mechanical governing system, it is common practice to close the throttle quickly to idle. For rotorcraft equipped with a FADEC, and particularly those with a 2 minute/30 second OEI rating structure, it is common practice to simulate an OEI condition by using reduced power on all engines by means of a flight test tool.

(ii) In every case, it must be demonstrated that all aspects of rotorcraft and powerplant behavior are identical to those that would occur in the event of an actual engine failure with the remaining engine developing minimum-specification power. Of particular concern are “dead engine” power decay characteristics, “live engine” acceleration characteristics, and rotor RPM control.

(iii) To this end, it is expected that a number of actual engine shut down tests will be conducted to generate sufficient data to validate the fidelity of the flight test tool and methodology, which will then allow its use in developing regulatory performance data. In general, it is best to conduct the tests in a low hover with the rotorcraft stabilized below the HV low point. An engine is then shut down and, following the appropriate pilot intervention time, the collective is raised to cushion the landing.

Page B – 20 5/1/2014 AC 29-2C, Chg 4 AC 29.45A. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.

a. Explanation . Amendment 29-24 adds § 29.45(f) to the regulation. This section establishes the requirement for furnishing power assurance information for turbine powered aircraft. This information is to provide the pilot a means of determining, prior to takeoff, that each engine will produce the power necessary to achieve the performance presented in the RFM.

b. Procedures . All of the guidance material pertaining to AC section 27.45 remains in effect. In addition, the power assurance information included in the RFM should be verified. Although this requirement is normally met with a power assurance curve, other methods of compliance may be proposed.

AC 29.45B. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.

a. Explanation . Although § 29.45 was not changed by Amendment 29-34, that amendment added requirements for certification of 30-second/2-minute OEI power ratings. For rotorcraft approved for the use of 30-second/2-minute OEI, partial power checks currently accomplished with approved power assurance procedures for lower power levels may not be sufficient to guarantee the ability to achieve the 30-second power level.

b. Procedures . MG 9 of this AC includes material on power assurance procedures to ensure that the OEI power level can be achieved. All of the guidance material pertaining to AC sections 29.45 and 29.45A remain in effect.

AC 29.45C. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.

a. Explanation . Although § 29.45 was not changed by Amendment 29-51, that amendment added new performance and handling qualities requirements for transport category rotorcraft. Included within these regulatory changes is OGE handling qualities.

Additionally, hover performance requirements were re-identified from § 29.73 to § 29.49.

b. Procedures . All of the guidance material pertaining to AC sections 29.45, 29.45A, and 29.45B remain in effect. In addition, the following apply: (1) OGE handling qualities may be extrapolated from test data up to a maximum of ±2,000 feet density altitude from the test altitude.

(2) Hover performance guidance that applied to § 29.73 is applicable to § 29.49.

Page B – 21 AC 29-2C, Chg 4 5/1/2014 CERTIFICATION BASIS 14 CFR part CAR 7 Rqms 29-Amdt. 21 29-Amdt. 1 Original Original CAT A HV Cat A & B (>9 Cat A: Cat A: Cat. A: Ref. TEST pax seats): 1. MGW Sea 1. MGW Sea 1. MGW Sea 29.25 CONDITIONS 1. W.A.T. for Level Level Level 29.87 which t.o. and 2. Max. IGE 2. Max. IGE 2. Max. IGE wt.

29.1517 ldg. are wt. Max. wt. Max. Max.

29.1581 approved. alt. capability. alt. capability. alt. Capability.

29.1583 2. Failure of 3. Failure of 3. Failure of 3. Failure of 7.11 critical engine. critical engine. critical critical engine.

7.715 engine.

7.741 AC 29 - CAT A 3. HV is 4. HV is 4. HV is 4. HV is 2C, RFM limitation. limitation. limitation. limitation.

Sections 4. Type of ldg. 5. Type of ldg. 5. Type of ldg. 5. Type of ldg.

29.45 & surface. Surface. surface. surface.

29.79 CAT B Cat B (</=9 Cat B: Cat B: Cat B: TEST pax seats): 1. MGW Sea 1. MGW Sea 1. MGW Sea CONDITIONS 1. MGW Sea Level Level Level Level 2. Max. IGE 2. Max. IGE 2. Max. IGE wt.

2. Max. OGE wt. Max. alt. wt. Max. alt. Max. alt.

wt. Max. alt. Capability Capability Capability Capability 3. Complete 3. Complete 3. Complete 3. Complete power failure, power failure, power failure, or power failure, or failure of or failure of failure of critical or failure of critical engine critical engine engine (w/eng critical engine (w/eng (w/eng isolation.)

(w/eng isolation). isolation.)

isolation).

CAT B 4. HV is perf. 4. HV is 4. HV is 4. HV is RFM Info. limitation info. limitation info. limitation info.

5. Type of ldg. 5. Type of ldg. 5. Type of ldg. 5. Type of ldg.

surface. Surface. surface. surface.

FIGURE AC 29.45-1. HV REQUIREMENTS Page B – 22 5/1/2014 AC 29-2C, Chg 4 CERTIFICATION BASIS 14 CFR part CAR 7 Requirement 29-Amdt. 24 29-Amdt. 3 Original Original s CAT A IGE Cat A 1. Conditions 1. Conditions 1.

CONTROL TEST 1. W.A.T. for selected by the selected by the Conditions Ref. 29.25 CONDITION which applicant. applicant. selected by 29.1583 S t.o. and ldg. 2. Critical CG 2. Critical CG the applicant.

7.121 are Critical Nr Critical Nr 2. Critical 7.743 approved. 3. Wind not 3. Wind not less CG Critical 2. Critical wt. less than 17 than 20 mph. NR Critical CG kts. 3. Wind not Critical Nr less than 20 3. Wind not mph.

less than 17 kts.

AC 29-2C CAT A 4. Max. 4. Max safe 4. Max safe 4. Max.

Sections RFM allowable wind wind above wind above alt. allowable 29.45 and is limitation. max. alt. For for which 17 kt. wind above 29.143 which 17 kt. wind envelope is the altitude Wind envelope established is for which 20 is established perf. info. mph wind is perf. info. envelope is est. is perf.

info.

CAT B Cat B: TEST 1. W.A.T. for CONDITION which t.o. and S ldg. Are approved.

2. Critical wt.

Critical CG Critical Nr 3. Wind speed & quad selected by the applicant.

CAT B 4. Max. safe RFM wind is perf.

info.

FIGURE AC 29.45-2 IGE CONTROLLABILITY REQUIREMENTS Page B – 23 AC 29-2C, Chg 4 5/1/2014 0:: w ENG SIN 12345 :s:: Q.

w

~ 400

(/)

\

.

0:: ::c I- u.

<( MIN CERTIFIED ::c (/) SPECIFICATION ENGINE

~-

w t; ~ 0:: 1000 1100 1200 1300 1400 1500 CORRECTEDTOT°F FIGURE AC 29.45-3 SHAFT HORSEPOWER VS TURBINE OUlLET TEMPERATURE- SEA LEVEL STANDARD DAY Page B – 24 5/1/2014 AC 29-2C, Chg 4 ....

1 4 ~ I- w w LL 12 ' w :::) I- i= _J <( w Cl:'.

H :::) (f) (f) w Cl:'.

CL ,, 0 '---------------'-----'---'---'--'---'------''----'-_,_ 211 0 400 5 00 :1 011 SHAFT HORSEPOWER AVAILABLE FIGU RE AC 29 . 45 -4 UNINS TALL ED TAK EOFF PO W ER AVA I LA BLE Page B – 25 AC 29-2C, Chg 4 5/1/2014 I I I I I I I I of I I I I I 1 8 .,,.1--- +--+---i---r -, 91, I I I I ,o

,. 1< I I I I

S I I I I I I 'ifi 1• I I I I 1s _ L __ 0 1-'t<'r-+--+---i---r I I ' I I Z I I I I I I O f I I I I V5 1 I I I I ~ I I:: I I I __ Tl __ Tl -iL 2_.1 14 - t-- --1- z l _, I 1 I I I I I I 0 I I I o I I I I I I I I I I ~ I X 12 __ L ___ _ -+- +-- ---t ~ I I I I I 1- I I I I I 11:J I I I I I LL I I I I I ' I I I . I ~ I I I -I-- - , ---t- :::l 10 I I I i I I 1- I I I I I I ~ I I I I I I 4'. I I I I I I w I I I I I I 0:: 8 - 1- -- 1 ---t-- -r-- , ---, :::l I I I I I I ~ I I I I I I W I I I I I I

8: I I I I

I I I I I I I I __ L __ J -- r --t---,---r I I I I I I I I I I I I I I I I I I I I I I I I I I _ L __ I __ J ___ f __ - , ---r- I I I I 1 I I I I I I I I I I I I I I I I I I I I I I I __ L 2 - 1- --1---t---J-- -t -- - I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I 0 -------------------------------.._ ..... _....,..___ ...

200 400 SHAFT HORSEPOWER AV Ai LABLE FI GURE AC 29 .4 5-5 INSTA LL ED T AK EOFF POWER AVAILABLE , AN TI -I CE OFF. 400 RPM Page B – 26 5/1/2014 AC 29-2C, Chg 4 L VE FT FT FT FT FT 10000 4000 2000 SEA LE 8000 60()() / / / / / ,/ , V V V / / _,, v 1 / / /v / / / / / / , / 80 RQUE ,, ,, ,,.

I I I I I

..

TO / A / / / NT V V l,L V V V / 70 % v_ V ~ ., ,v 40 ,/ ,/ / / PERCE _/ / RQUE= l,,,L V V I/ , - ✓ V V V V , )/ , - ~ V V V V 0 00 FT - - REQUIRED TO OAT = 17.5°C TO T = 760° C H • = -.

AMPLE: - EX - PS ~ ' °C \

-

E- - '

\

- '\

'

\ \

\ - \ ' \

\

'\ \

-

- \ \ \ '\

\

- i\ i\

' 0% ATOR LOAD 100 AM ' \

\ \

\ \

"" \

O AIRSPEED ' ' \ TI-ICE OFF '\ \ \ \ '\ \.

ENER ER Z RPM MJ G I I\ i\ I\ I '\ 'h '\ \I ,;\ DITI ONS 'i: \ Z, '\ ' \ \ \ \

'\ ' ' \

' ' CON ' '\ \ 10 '\ \ '\ '\ \

'\ ' ,\

'

'

'\ \ \ \ \ \ r-... r-._ " "\ ' OUTSIDE AIR TEMPERATUR o·c 0°C 0°C 0"C o·c o·c o·c o·c 78 76 74 n 1o so ss ss ,, Gi C ::0 m )> (") I',) (0 '.:,. 0, &, "'O 0 ~ m ::0 )> (/) (/) C ~ z (") m (") ::i:: m (") ;,::: (") ::i:: )> ::0 -I Page B – 26.1 AC 29-2C, Chg 4 5/1/2014 AC 29.49. § 29.49 (Amendment 29-39) PERFORMANCE AT MINIMUM OPERATING SPEED.

(For performance at minimum operating speed and for hover performance prior to Amendment 29-39, see § 29.73 and section 29.73 of this AC.)

a. Explanation .

(1) Amendment 29-39 re-designated § 29.73 as § 29.49 to relocate the requirements for rotorcraft hover performance. For the purpose of this manual, the word “hover” applies to a rotorcraft that is airborne at a given altitude over a fixed geographical point regardless of wind. Pure hover is accomplished only in still air.

(2) Under § 29.49, hover performance should be determined at a height consistent with the takeoff procedure for category A rotorcraft and in ground effect (IGE) for category B rotorcraft. Additionally, out of ground effect (OGE) hover performance should be determined for both category A and B rotorcraft. Hover OGE is that condition, where an increase in height above the ground will not require additional power to hover. Hover OGE is the absence of measurable ground effect. It can be less than one rotor diameter at low gross weight increasing significantly at high gross weights. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle 1 ½ rotor diameters above the surface.

(3) The objective of hover performance tests is to determine the power required to hover at different gross weights, ambient temperatures, and pressure altitudes.

Using non-dimensional power coefficients (C ) and thrust coefficients (C ) for p t normalizing and presenting test results, a minimum amount of data are required to cover the rotorcraft’s performance operating envelope.

(4) Hover performance tests must be conducted over a sufficient range of pressure altitudes and weights to cover the approved ranges of those variables for takeoff and landings. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.

(5) The minimum hover height for which data should be obtained and subsequently presented in the flight manual should be the same height consistent with the minimum hover height demonstrated during the takeoff tests. Refer to section 29.51 of this AC for the procedure to determine the minimum allowable hover height.

b. Procedures .

(1) Two methods of acquiring hover performance data are the tethered and free flight techniques. The tethered technique is accomplished by tethering the rotorcraft to the ground using a cable and load cell. The load cell and cable are attached to the ground tie-down and to the rotorcraft cargo hook. The load cell is used to measure the rotorcraft’s pull on the cable. Hover heights are based on skid or wheel height above Page B – 26.2 5/1/2014 AC 29-2C, Chg 4 the ground. During tethered hover tests, the rotorcraft should be at light gross weight.

The rotorcraft will be stabilized at a fixed power setting and rotor speed at the appropriate skid or wheel height. Once the required data are obtained, power should be varied from the minimum to the maximum allowed at various rotor RPM. This technique will produce a large C /C spread. The load cell reading is recorded for each stabilized t p point. The total thrust the rotor produces is the rotorcraft’s gross weight, weight of the cables and load cell plus cable tension. Care must be taken that the cable tension does not exceed the cargo hook limit or load capacity of the tie-down. For some rotorcraft, it may be necessary to ballast the rotorcraft to a heavy weight in order to record high power hover data.

(2) The pilot maintains the rotorcraft in position so that the cable and load cell are perpendicular to the ground. To insure the cable is vertical, two outside observers, one forward of the rotorcraft and one to one side, can be used. Either hand signals or radio can be used to direct the pilot. The observers should be provided with protective equipment. This can also be accomplished by attaching two accelerometers to the load cell which sense movement along the longitudinal and lateral axes. Any displacement of the load cell will be reflected on instrumentation in the cockpit and by reference to this instrumentation, the rotorcraft can be maintained in the correct position. Accurate load cell values may also be obtained by measuring cable angles and, through geometry, determining a corrected load cell value. Increased caution should be utilized as tethered hover heights are decreased because the rotorcraft may become more difficult to control precisely. The tethered hover technique is especially useful for OGE hover performance data because the rotorcraft’s internal weight is low and the cable and load cell can be jettisoned in the event of an engine failure or other emergency.

(3) To obtain consistent data, the wind velocity should be 3 knots or less. Large rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are: (i) Engine and transmission torque.

(ii) Rotor speed.

(iii) Ambient and engine temperatures, such as measured gas temperature (MGT).

(iv) Pressure altitude.

(v) Fuel used (or remaining).

(vi) Load cell reading.

(vii) Generator(s) load.

(viii) Wind speed and direction.

Page B – 27 AC 29-2C, Chg 4 5/1/2014 (ix) Hover height.

As a technique, it is recommended the rotorcraft be loaded to a center of gravity (CG) near the hook to minimize fuselage angle changes with varying powers. All tethered hover data should be verified by a limited spot-check using the free flight technique.

The free flight technique in paragraph b.(4) below will determine if any problems, such as load cell malfunctions, have occurred. The free flight hover data must fall within the allowable scatter of the tethered data.

(4) If there are no provisions or equipment to conduct tethered hover tests, the free flight technique is also a valid method. The disadvantage of this technique as the primary source of data acquisition is that it is very time consuming. In addition a certain element of safety is lost OGE in the event of emergency. The rotorcraft must be reballasted to different weights to allow the maximum C /C spread. When using the t p free flight technique, either as a primary data source or to substantiate the tethered technique, the same considerations for wind, recorded parameters, etc., as used in the tethered technique apply. Free flight hover tests should be conducted at CG extremes to verify any CG effects. If the rotorcraft has any stability augmentation system, which may influence hover performance, it must be accounted for.

(5) Comprehensive hover performance tests are typically conducted at low, intermediate (approximately 7000 feet H ), and high altitude test sites with prepared D landing surfaces, in conjunction with takeoff, landing, controllability, and maneuverability testing. Alternatively, a predicted hover performance model developed for high altitude may be used if verified by limited flight testing. The extrapolation guidelines in section 29.45.b.(2) of this AC are still applicable. These higher altitude hover tests could typically be conducted in conjunction with the limited controllability tests. If the applicant is able to demonstrate to the approving airworthiness authority a method to provide a reliable hover reference, it is acceptable to conduct OGE tests without ground reference.

Page B – 28 9/30/99 AC 29-2C AC 29.51. § 29.51 TAKEOFF DATA - GENERAL.

a. Explanation . Section 29.51 details the conditions under which takeoff performance data can be obtained and presented in the FAA/AUTHORITY approved flight manual. The flight manual must also contain the technique(s) to be used to obtain the published flight manual takeoff performance. Technique should not be confused with exceptional pilot skill and/or alertness as mentioned in § 29.51. Rotorcraft are different from one another and due to this, different pilot techniques are sometimes required to achieve the safest and most optimum takeoff performance. The recommended technique that is published in the flight manual and used to achieve the performance must be determined to be one that the operational pilot can duplicate using the minimum amount of type design cockpit instrumentation and the minimum crew.

b. Background .

(1) Certain special takeoff techniques are necessary when a rotorcraft is unable to takeoff vertically because of altitude, weight, power effects, or operational limitations.

The recommended technique used to take off under such conditions is to accelerate the rotorcraft in-ground-effect (IGE) to a predetermined airspeed prior to climbout. Takeoff tests are performed to determine the best repeatable technique(s) for a particular rotorcraft over the range of weight, altitude, and temperature for which certification is requested.

(2) The primary factor which determines the rotorcraft’s takeoff performance is the amount of excess power available. Excess power available is the difference between the power required to hover at the reference height above the ground and the takeoff power available from a minimum installed specification engine. Utilizing the total power available to execute a takeoff may not be operationally feasible due to such items as HV constraints. In such situations, hover power required plus some power increment may be the maximum that can be used and the resulting performance determined accordingly.

(3) Landing gear height above the ground should not be greater than that demonstrated satisfactorily for HV, rejected takeoff, and that height for which IGE hover performance data is presented in the RFM, or less than that height below which ground contact may occur when accomplishing takeoff procedures. For rotorcraft fitted with wheels, a running takeoff procedure may be accepted. The hover reference height is established as the minimum landing gear height above the takeoff surface, from which a takeoff can be accomplished consistently in zero wind without contacting the runway.

Category B takeoff must be accomplished with power fixed at the power required to hover at the reference height (not greater than the height for which IGE performance data is presented).

c. Procedure . There are different techniques which may be used in order to determine which method is best for a particular rotorcraft. The most commonly accepted method is the hover and level acceleration technique. In this technique, the Page B – 29 AC 29-2C, Chg 4 5/1/2014 rotorcraft is stabilized in a hover at the reference height. From the stabilized hover, the rotorcraft is accelerated to the climbout airspeed using the predetermined takeoff power.

When the desired climbout airspeed is achieved, the rotorcraft is rotated and the climbout is accomplished at the schedule airspeed(s) and constant rotor RPM. Power adjustments may be accomplished to maintain targeted power except where procedure requires high workload outside cockpit (i.e., that portion of takeoff where horizontal acceleration close to the ground has pilot scan outside the cockpit and adjustment of engine torque or temperature would require an undue increase in workload).

AC 29.51A. § 29.51 (Amendment 29-39) TAKEOFF DATA - GENERAL.

a. Explanation . Amendment 39 added takeoff requirements in new §§ 29.55, 29.60, 29.61 and 29.62.

b. Procedures . The guidance material presented in paragraph AC 29.51 continues to apply.

Page B – 30 9/30/99 AC 29-2C AC 29.53. § 29.53 TAKEOFF: CATEGORY A .

a. Explanation .

(1) A Category A takeoff typically begins with an acceleration and/or climb from a hover to a critical decision point. The rule requires that the critical decision point (CDP) be defined for the pilot in terms of an indicated altitude and airspeed combination. However, other parameters to define the CDP have been accepted by the FAA/AUTHORITY on an equivalent safety basis. A regulatory project has been established to change the rule permitting other parameters to be used for CDP definition.

(2) The requirement to define CDP as a combination of both airspeed and height above the takeoff surface is based on a minimum required total energy concept.

A specific minimum combination of kinetic energy (airspeed) and potential energy (height) must be attained at the CDP to be assured that a continued takeoff can be accomplished following the complete failure of one engine. In § 29.53(b), CDP is required to be “…a combination of height and speed selected by the applicant…” Any other method proposed to define CDP must provide the same level of safety as would be obtained using an airspeed-height combination. When using “time,” “height,” or “airspeed” only as alternative methods of identifying the CDP, they must be combined with a precisely defined takeoff path and crew procedure in order to provide the required equivalent level of safety. In addition, it must be demonstrated that the pilot technique used during the takeoff sequence is easily repeatable and consistently produces the required energy (i.e., airspeed and altitude combination) when the CDP “time,” “height,” or “airspeed” is attained. This condition should be verified during the flight test program.

(3) If an engine fails at the CDP or at any point in the takeoff profile prior to attaining CDP, the rotorcraft must be able to land safely within the established rejected takeoff distance. Flight testing to determine the Category A rejected takeoff distance is very similar to height-velocity testing and should be approached with caution. The initial Category A takeoff profiles should be outside of the Category B height-velocity envelope. Previous programs have shown the low speed point immediately after application of power to be particularly critical.

(4) If an engine fails at the CDP or at any subsequent point in the Category A takeoff profile, a continued safe climb-out capability is assured. The continued takeoff for conventional Category A runway profiles is designed to allow acquisition of the takeoff safety speed (V ), at a minimum of 35 feet above the takeoff surface and a TOSS positive rate of climb. During the continued takeoff profile, the pilot is assumed to be flying the rotorcraft via the primary flight controls (cyclic stick, collective, and directional pedals). Manipulation of the throttle controls or beep switches may be permitted as long as such manipulation can be accomplished readily by the pilot flying the rotorcraft without removing his hands from the cyclic and collective flight controls. These manipulations of engine controls should not make major adjustments in power, and Page B - 31 AC 29-2C 9/30/99 should not occur before attaining V . In no case should this be less than 3 seconds TOSS after the critical engine is made inoperative.

(5) Both the rejected takeoff distance and the continued takeoff distance must be determined. Although 29.59(c) suggests a balanced field length requirement, this was not intended. Both rejected and continued takeoff distance should be included in the RFM performance with information stating that the longer distance determines the length of the required takeoff surface. Operations approvals can then determine the required takeoff surface (including stopways and clearways) appropriate for the specific operation.

(6) A typical Category A takeoff profile, assuming an engine failure at the CDP, is shown in figure AC 29.53-1.

b. Procedures .

None.

Page B - 32 9/30/99 AC 29-2C HEIGHT I ABOVE ,I GROUND

I

I (FEEl) VToss I

----- -r-

I

I STOP

I :-:-:;:_

...

.._TAKEOFFCLIIAB S EGMEHT--..; A LL ENG INES OPERATING ~TRANSITION SEGMENT TAKEOFF SEGMENT I RTO SEGMENT ----1 ► 1 ~---DEMONSTRATED SA FE OEI LANDINGS ----1.i HORIZONTAL DISTANCE FIGURE AC 29.53-1 TAKEOFF PERFORMANCE CATEGORY A Page B - 33 AC 29-2C 9/30/99 AC 29.53A. § 29.53 (Amendment 29-39) TAKEOFF: CATEGORY A .

a. Explanation . Amendment 29-39 separated in the text, the Category A takeoff requirement from the definition of a decision point. Category A takeoff performance must be scheduled so that: (1) If an engine failure is recognized at the Takeoff Decision Point (TDP) or at any point in the takeoff profile prior to attaining TDP, the rotorcraft must be able to land safely within the established rejected takeoff distance. Flight testing to determine the Category A rejected takeoff distance is very similar to height-velocity testing and should be approached with caution. The initial Category A takeoff profiles should be outside of the avoid area of the Category B height-velocity envelope. Previous programs have shown the low speed point immediately after application of power to be particularly critical.

(2) If an engine failure is recognized at the TDP or at any subsequent point in the Category A takeoff profile, a continued safe climb-out capability must be assured.

The continued takeoff for conventional Category A runway profiles is designed to allow acquisition of the takeoff safety speed (V ) at a minimum of 35 feet above the takeoff TOSS surface and a positive rate of climb.

(3) Both the rejected takeoff distance and the continued takeoff distance should be determined. A balanced field length is not required by the regulation. Both rejected and continued takeoff distance should be included in the RFM performance section.

Operations approvals can then determine the required takeoff surface (including stopways and clearways) appropriate for the specific operation.

(4) A typical Category A takeoff profile, assuming an engine failure prior to the TDP, is shown in figure AC 29.53A-1.

b. Procedures. None.

Page B - 34 9/30/99 AC 29-2C Vv .:.150fpm 1000 ft (29.67a.2) 1000 ft.

I

::::i

g

-cJ C :::, ~ C) Q) > 200 ft v. ?.100 fpm .0 H-V AVOID (29 87) (29.67 a.1) ,,- ~

~ 1/.//1//.((/(,f /

§-'" .s:::.

-

0) ·;;; "',§$ V,oss .._l;- I 35 ft 35 ft.

15 ft.

- '

15 ft., Oft. I+-- ' -----..

Rejected T.O. Dist (29.62) _____..

Horizontal Distance I+-- T.O. Dist (29.61) FIGURE AC 29.53A-1 TAKEOFF PERFORMANCE CATEGORY A Page B - 35 AC 29-2C 9/30/99 AC 29.55. § 29.55 (Amendment 29-39) TAKEOFF DECISION POINT: CATEGORY A .

a. Explanation .

(1) Amendment 29-39 added a new § 29.55 to redefine the TDP (previously called the CDP) and contained in § 29.53; it further removed the requirement to identify the TDP by height and airspeed, since height alone or other factors may be more appropriate. A Category A takeoff typically begins with an acceleration and/or climb from a hover to TDP. The rule requires that the TDP be defined for the pilot in terms of no more than two parameters such as an indicated height and airspeed combination.

(2) The definition of the TDP is based on a minimum required total energy concept. A specific minimum combination of kinetic energy (airspeed) and potential energy (height) should be attained at the TDP to ensure that a continued takeoff can be accomplished following the complete failure of one engine. In § 29.55(b), TDP is required to be defined by no more than two parameters. When using a single parameter such as time, height, or airspeed as a method of identifying the TDP, the identification must be combined with a precisely defined takeoff path and crew procedure to provide the required equivalent level of safety. In addition, it should be demonstrated that the pilot technique used during the takeoff sequence is easily repeatable and consistently produces the required energy (i.e., airspeed and height combination) when the TDP time, height, or airspeed is attained. This condition should be verified during the flight test program.

b. Procedures . None.

AC 29.59. § 29.59 (Amendment 29-24) TAKEOFF PATH: CATEGORY A .

a. Explanation . The Category A concept limits the rotorcraft takeoff weight such that if an engine failure occurs at or before the CDP, a safe landing can be made or if the engine fails at or after the CDP, the takeoff can be continued. The purpose of these tests is to define the CDP, evaluate the necessary pilot techniques, and determine the required takeoff area for either alternative. The condition of equal distances for either stopping or continuing the takeoff is called a “balanced” field length. The combination of altitude and speed at the CDP which produces a balanced field length is not required for certification. This section deals with the Category A takeoff and rejected takeoff profiles. The profiles necessarily involve consideration of an average pilot skill level as well as a sequence in which it is assumed various configuration adjustments are made to the rotorcraft.

(1) Takeoff . The Category A takeoff path begins with an all-engines-operating acceleration segment to the CDP and continues with a one-engine-inoperative acceleration to takeoff safety speed (V ). (See Conventional Takeoff Profile, TOSS figure AC 29.53-1, paragraph AC 29.53.) CDP is a “go/no-go” condition which is analogous to V speed in transport airplanes. Prior to CDP the pilot is “stop” oriented, Page B - 36 9/30/99 AC 29-2C and when an engine fails in this portion of the takeoff, he will abort because he has not yet achieved sufficient energy to assure continued flight. At the CDP the pilot becomes “go” oriented and when an engine fails at or beyond this point he will continue the takeoff because he no longer has sufficient surface area to abort the takeoff. The takeoff flight path and the CDP must be defined such that a safe landing can be made from any point up to the CDP. This profile may differ significantly from the takeoff flight path developed for Category B weights. The CDP is the last point in the takeoff profile at which a rejected takeoff capability within the scheduled takeoff surface distance is assured. If an engine failure does not occur, the pilot continues the climb and accelerates past the CDP to the recommended climb speed.

(2) Rejected Takeoff . The rejected takeoff profile begins with an all engine acceleration segment to the CDP and ends when the rotorcraft is brought to a complete stop on the designated takeoff surface. The critical engine is made inoperative at the CDP and the landing must be made with the remaining engine(s) operating within approved limits. The rejected takeoff distance is normally measured at a given reference point on the rotorcraft from the start of the takeoff to the same reference point after the rotorcraft has come to a complete stop. This distance should be increased by the rotorcraft length (including main and tail rotor tip paths).

(3) Takeoff Climbout Path .

(i) The “OEI transition segment” is defined as the segment from CDP where the engine becomes inoperative to V . It is assumed that the maximum TOSS approved OEI power is used until the allowable time duration for that power is exhausted. It must be possible for the crew to fly the rotorcraft to V and attain an TOSS altitude of 35 feet and then climb to 100 feet above the takeoff surface by flying the rotorcraft solely by the primary flight controls (including collective). The landing gear may be retracted after attaining a height of 35 feet above the takeoff surface, a speed of V , and a positive rate of climb. Flight manual procedures may recommend TOSS adjustment of auxiliary controls to improve OEI performance. However, compliance with the performance requirements of § 29.67(a)(1) should not be based on use of secondary engine controls such as beepers, etc. Manipulation of the throttle controls or beep switches may be permitted for compliance with the performance requirements of § 29.67(a)(2) as long as such manipulation can be accomplished readily by the pilot flying the rotorcraft without removing his hands from the cyclic and collective flight controls. These manipulations of secondary engine controls should not make major adjustments in the power, and should not occur before attaining V . There should be TOSS a minimum delay of 3 seconds after the critical engine is made inoperative before adjustment of secondary engine controls is allowed during the takeoff path determination. The failure of one engine cannot affect continued safe operation of the remaining engines or require any immediate action by the crew per § 29.903(b). If a 2 ½-minute power rating is used, it should be possible to complete the Category A takeoff profile (assuming an engine failure at CDP), accelerate to V , attain 35 feet TOSS above the surface, and complete landing gear retraction prior to exhausting the 2 ½-minute time limit.

Page B - 37 AC 29-2C 9/30/99 (ii) The takeoff safety speed, V , is a speed at which 100 FPM rate of TOSS climb is assured under conditions defined in § 29.67(a)(1). The takeoff distance is the distance from initial hover to the point at which V and 35 feet in a climbing posture TOSS are attained.

(4) Continued Climbout Path . Continued acceleration and climb capability from 100 feet above the takeoff surface is assured by the 100 FPM V climb requirement TOSS of § 29.67(a)(1) and the 150 FPM requirement of § 29.67(a)(2), normally demonstrated at V . It should be shown that the rotorcraft can be accelerated from V to V in a Y TOSS Y continuous maneuver without losing altitude, including any configurative change (landing gear retraction, etc.).

b. Procedures .

(1) Instrumentation . A photo theodolite, grid camera, or other position measuring equipment is required together with a ground station to measure wind, OAT, humidity (if applicable), and a two-way communication system to coordinate activities with the aircraft. A crash recovery team with support of a fire engine is highly desirable.

Aircraft instrumentation should record with a time scale: engine parameters (speed, temperature, and power), rotor speed, flight parameters (airspeed, altitude, and normal acceleration as a minimum), flight control positions, power lever position, and landing gear loads. Additionally, a method should be devised to allow correlation of the aircraft instrumentation data with the space position data to accurately determine the length of the various takeoff segments.

(2) Establishing the Critical Decision Point (CDP) .

(i) The CDP should be definable with the minimum crew using standard cockpit instrumentation. If a radar altimeter is used, it should be included in the minimum equipment list. If barometric altitude is used to define CDP, the operating conditions at which the altimeter is set should be defined. This is normally done on the ground with the minimum collective pitch. If the wind influences the altimeter reading, the correct relative wind information should be provided. Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the CDP becomes largely a function of the surface area required for takeoff. If takeoff conditions scheduled include considerable surface area (on the order of 2,000 feet), the CDP airspeed may be a high value near V . This will allow a higher takeoff weight and Y demonstrate compliance with the V climb requirement of § 29.67(a)(1). In this case, TOSS the requirements of § 29.67(a)(2) usually become limiting. If required surface area is a small value, CDP will necessarily be some lower airspeed value to allow for an aborted takeoff on the available surface. Weight may need to be reduced at lower values of CDP airspeed (significantly below V ) to allow compliance with the climb requirement of Y § 29.67(a)(1). Compliance with climb requirements can be substantiated initially by testing at a safe altitude above the ground. When OEI climb conditions are verified for Page B - 38 9/30/99 AC 29-2C weight, configuration, pressure altitude, and temperature, the CDP is then evaluated in a rejected takeoff.

(ii) A Category A takeoff procedure for which the CDP is defined as a specific “time,” “height,” or “airspeed” in the takeoff sequence combined with a precise takeoff crew procedure may be approved on the basis of equivalent safety when the following conditions can be satisfied: (A) The flightcrew takeoff procedure must be shown to be consistently repeatable and not require exceptional piloting skill.

(B) It must be documented that the takeoff procedure will produce the required minimum energy level in terms of height and airspeed for all combinations of gross weight, altitude, and ambient temperature for which takeoff data are scheduled.

This may best be accomplished by conducting takeoff procedure abuse tests to show that variations from the established takeoff procedure that could reasonably be expected to occur in service do not result in significant increases in the takeoff distances.

(3) Rejected Takeoff Distance . The rejected takeoff is similar in many respects to the height-velocity (HV) tests described in paragraph AC 29.73. Most of the comments, cautions, and techniques for HV also apply here even though typical flight conditions at CDP are less critical than limiting HV points. As mentioned in paragraph AC 29.79, a minimum 5-knot clearance from any HV limiting condition should be provided throughout the takeoff flight path (see figure AC 29.63-1), and tests should be conducted simulating an unplanned engine cut. The HV diagram appropriate in the Category A test weights may be much less restrictive than that determined for Category B conditions. Normally, a minimum 1-second delay is applied after engine failure before pilot collective control corrections are allowed. However, if pilot cues are strong enough to make engine failure unmistakable, normal pilot reaction time may be utilized following engine failure. As in all engine failure testing, the pilot should not anticipate the failure by changing flight control positions or aircraft attitude. Average pilot techniques should be used. The two primary objectives of rejected takeoff testing are an assured capability to safely return to the takeoff surface when an engine fails at any point prior to CDP and the determination of the rejected takeoff distance that is needed when an engine fails at the CDP. It is important that the surface conditions be defined. For the rejected takeoff distance tests, a minimum of five satisfactory runs should be flown by the FAA/AUTHORITY pilot. The rejected takeoff distances from company and FAA/AUTHORITY runs may be averaged. The rejected takeoff distance tests will be used together with the OEI continued takeoff profiles to establish the required surface area for Category A operations.

(4) Continued Takeoff Distance .

(i) Continued takeoff profiles should be flown to determine the continued takeoff distance. This distance is measured from the point of takeoff initiation to the Page B - 39 AC 29-2C 9/30/99 point in the takeoff profile where the following three conditions have all been attained after a failure of the critical engine at CDP: an airspeed equal to or greater than V , TOSS a positive rate of climb, and a height of at least 35 feet above the takeoff surface. The rotorcraft should not contact the ground at any point after engine failure. If the rotorcraft descends below 35 feet above the takeoff surface while accelerating to V , the TOSS takeoff distance is extended to the point that 35 feet is reattained with a positive rate of climb.

(ii) If the CDP is significantly above 35 feet so that the rotorcraft does not descend below 35 feet during acceleration to V , the takeoff distance then becomes TOSS the distance to the point in the takeoff profile at which both V and a positive TOSS rate-of-climb are attained after failure of the critical engine at CDP. For most applications, the rotorcraft should not be allowed to descend more than one-half the CDP height above the takeoff surface while accelerating to V . In addition, the TOSS rotorcraft should not be allowed to descend below the height above the takeoff surface at which a landing flare would normally be initiated. For example, if a rotorcraft has a CDP of 20 feet but when landing would normally initiate the landing flare at 15 feet, the takeoff profile should not be allowed to descend to 10 feet but should remain above 15 feet in establishing the takeoff distances.

(iii) In establishing the continued takeoff distance, the applicable pilot recognition delay time should be applied following the engine failure at CDP, and the takeoff profile should be established with the pilot using primary flight controls only to control the rotorcraft. The pilot engine failure recognition time delay before adjustment of the collective pitch control should be a minimum of 1 second unless it can be demonstrated that the pilot will have unmistakable engine failure cues sooner than 1 second.

(iv) Engine failure testing should be initially conducted at a safe distance above the ground to assess the continued takeoff profile before conducting the actual profiles for credit. This procedure will serve to validate predicted performance and may prevent an unexpected return to the surface during continued takeoff tests. A minimum of five acceptable runs should be flown by the FAA/AUTHORITY pilot, and these should be averaged with five acceptable runs flown by the manufacturer’s pilot.

(5) Abuse Testing . Takeoff procedure abuse tests should be conducted to show that reasonably expected variations in service from the established takeoff procedures do not result in a significant increase in the established takeoff distances.

Variations should include such considerations as under or over rotation during the takeoff initiation, under or over application of acceleration power, and missed CDP target parameters (e.g., time, height, or airspeed).

(6) Continued Climbout Path . The climb performance requirements of § 29.67(a)(1) should be met at the end of the continued takeoff distance segment.

Beginning at this point, the landing gear may be retracted, and secondary engine controls may be manipulated to adjust power. Any manipulation of secondary engine Page B - 40 9/30/99 AC 29-2C controls should be accomplished readily by the pilot flying the rotorcraft without removing his hands from the cyclic and collective flight controls. The climb should be continued at V until approximately 100 feet above the takeoff surface. It should be TOSS demonstrated that the rotorcraft including any configuration changes can be accelerated from V to V in a continuous maneuver without losing altitude. The airspeed and TOSS Y rotorcraft configuration (landing gear position, rotor RPM engine power, etc.) used to show compliance with the climb requirements of § 29.67(a)(2) should be attained at or prior to reaching 1,000 feet above the takeoff surface.

(7) Power . Power should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping to minimum specification values including consideration of temperature effects on engine power. Turbine engine power does not vary directly with density altitude (H ). At a D given H , turbine engine power available varies with ambient temperature. Turbine D engines typically produce less horsepower as ambient temperature is increased (pressure altitude decreases) at a given density altitude, although some engines produce less horsepower at extremely cold temperatures. In either event, if one test sequence is to be utilized for a given H , it would be appropriate to restrict test power to D the lowest value attainable from a minimum specification engine through the approved ambient temperature range at the density altitude of the test. To attain maximum weights for varying ambient conditions, the applicant may utilize a parametric mapping of power available, pressure altitude, and temperature effects. For this case, engine topping may be adjusted throughout a range appropriate to the test H .

D (8) Aircraft Loading . Both forward and aft CG extremes should be spot checked to determine the critical loading for takeoff distances. Forward center of gravity is usually critical for continued takeoff distance tests while aft CG may be critical for the rejected takeoff because of over-the-nose visibility. A minimum of two weights should be flown at each altitude if the manufacturer elects to schedule field length variation as a function of gross weight. One weight should be the maximum weight for prevailing conditions and the other weight(s) should be low enough to attain a sufficient spread to verify weight accountability.

(9) Extrapolation . Weight cannot be extrapolated above test weight for the same reasons discussed in paragraph AC 29.79. See paragraph AC 29.45 regarding altitude extrapolation of test results.

(10) Ambient Conditions . Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data must be corrected for existing wind conditions during takeoff distance testing. Credit for headwind conditions may be given during flight manual data expansion. Refer to paragraph AC 29.45(b)(1) under “Winds for Testing” for allowable wind credit. Care should be applied in considering headwind credit for vertical operations as previous experience has resulted in difficulty collecting meaningful, repeatable data.

(11) Vertical Takeoffs .

Page B - 41 AC 29-2C 9/30/99 (i) General . Guidelines for rotorcraft certification using vertical takeoff techniques were developed and utilized for civil certification programs many years ago.

As experience has been gained, certain policy decisions have modified these guidelines. The following guidelines incorporate all available policy information as of January 1, 1981. The reader should be familiar with the preceding discussion regarding conventional Category A takeoff profiles because duplicate information is not repeated here.

(ii) Takeoff Profile . A typical vertical takeoff profile for a ground level heliport is shown in figure AC 29.59-1. The maneuver begins with the addition of sufficient power to initiate a climb to the CDP. It must be possible to make a safe landing without exceptional pilot skill if an engine fails at any point up to the CDP. At the CDP, the pilot becomes “go” oriented and continues the takeoff if an engine fails. A typical profile for pinnacle takeoff conditions is shown in figure AC 29.59-2.

Considerations are similar to those of the ground level heliport in figure AC 29.59-1; however, the OEI pinnacle profile allows descent below the takeoff surface, specifies minimum edge clearance criteria, and allows relaxed requirements for final segment climb. Thus far, descent profiles up to 50 feet below the takeoff surface have been allowed; however, there is no reason why greater values could not be determined during engineering flight tests for certification. Use of such a profile, of course, would be dependent on obtaining an operational approval.

(iii) Critical Decision Point (CDP) . For vertical takeoffs, the climb to CDP is nearly vertical, and CDP is typically defined primarily by height. Sufficient testing must be conducted to define a band of CDP conditions (heights) which will be consistent with anticipated variations in pilot technique and the minimum amount of equipment to be installed on the production aircraft. Rejected takeoffs are most critical from high CDPs, and continued OEI takeoffs are most critical from low heights. Tests at the extremes of this band are intended to verify that the anticipated CDP band is safe and repeatable in service for reasonable variations in pilot technique. These extreme points should not be used for distance determination when averaging takeoff performance data.

(iv) Conduct of the Test . Vertical takeoff profiles must be flown from a pad simulating operational conditions because the sight picture may be critical to successful OEI operations, particularly for elevated heliports. At all points on the vertical takeoff flight path up to the CDP, the pilot, with reasonable head movement, shall be able to keep sufficient portions of two heliport boundaries (front and one side) or equivalent markings in view to achieve a safe landing in case of engine failure.

Normally, a minimum 1-second delay is applied after engine failure before pilot collective control corrections are allowed. However, if pilot cues are strong enough to make engine failure unmistakable, normal pilot reaction time may be utilized following engine failure.

Page B - 42 9/30/99 AC 29-2C (A) Establish the rejected takeoff distance as the horizontal distance from the rearmost point of the rotorcraft at the initiation of takeoff to the foremost point after the rotorcraft comes to a stop on the takeoff surface (including rotor tip path), assuming an engine failure in the vertical climb at the CDP; or (B) Establish the continued takeoff distance as the horizontal distance from lift-off to the point at which, following engine failure at CDP, the rotorcraft achieves 35 feet above the takeoff surface and V in a climbing posture. The continued TOSS takeoff profile from elevated heliports must clear the heliport obstructions by at least 15 feet vertically and 35 feet horizontally.

(v) Climb Requirements .

(A) The OEI takeoff profile should include a climb at V to 200 feet TOSS above the takeoff surface prior to accelerating to a higher speed.

(B) For elevated heliports, the climb requirement of § 29.67(a)(2) may be met at 200 feet above the takeoff surface or 1,000 feet above the surrounding terrain, whichever is higher.

(vi) Extrapolation . Basic guidelines for extrapolation are contained in paragraph AC 29.45. If, however, vertical takeoff weights are based upon allowable weights for hovering out-of-ground effect (OGE) with one engine inoperative, all vertical takeoff performance aspects may be extrapolated to the highest altitude requested for takeoff and landing.

(12) Night Operations .

(i) A minimum of three normal takeoffs (and landings) should be conducted to assure that aircraft lighting (internal and external) is adequate to allow normal Category A operations at night.

(ii) Engine failures should be simulated from points along the recommended takeoff profile. Night OEI rejected takeoffs and continued takeoffs from the CDP should be conducted to assure adequate night field of view and realization of Category A field lengths.

(iii) If special airfield markings are used as a reference or to define the CDP, the aircraft external lighting should be evaluated to assure that these airfield markings are adequately visible for night operations.

Page B - 43 AC 29-2C 9/30/99 All Engine Climbout Accelerate to Vv when clear of obstacles Accelerate to Vross CDP--+ Rejected Vertical Takeoff Path Climb Fl GURE AC 29.59-1 CATEGORY A VERTICAL TAKE OFF PROFILE GROUND LEVEL HELIPORT Page B - 44 9/30/99 AC 29-2C All Engine Climbout

~ Accelerate

To Vross

✓

Vertical

t

Climb

t

Minimum & 15 Feet Rejected I Takeoff Path T _______ ...._........ +--Minimum ____.

OEI Flight 35 Feet Path FIGURE AC 29.59-2 CATEGORY A VERTICAL TAKEOFF PROFILE PINNACLE Page B - 45 AC 29-2C 9/30/99 AC 29.59A (AC's 29.60, 29.61, & 29.62) §§ 29.59 (29.60, 29.61 and 29.62) (Amendment 29-39) TAKEOFF PATH, DISTANCE AND REJECTED TAKEOFF; GROUND LEVEL AND ELEVATED HELIPORT: CATEGORY A (For § 29.59 prior to Amendment 39, see paragraph AC 29.59.)

a. Explanation . Amendment 29-39 moved the rejected takeoff requirements from § 29.55 to a new § 29.62 and clearly defined the takeoff path. It also added new §§ 29.60 and 29.61 to introduce the requirements for elevated heliport takeoff path, Category A and to more clearly define the parameters to be used in determining takeoff distance, respectively.

(1) Takeoff Decision Point . The Category A concept limits the rotorcraft takeoff weight such that if an engine failure is recognized at or before the TDP, a safe landing can be made or if an engine failure is recognized at or after the TDP, the takeoff can be continued. The purpose of these tests is to define the TDP, evaluate the necessary pilot techniques, and determine the required takeoff area for either alternative. The condition of equal distances for either stopping or continuing the takeoff is called a “balanced” field length. The combination of altitude and speed at the TDP which produces a balanced field length is not required for certification. This section deals with the Category A takeoff and rejected takeoff profiles. The profiles necessarily involve consideration of an average pilot skill level as well as a sequence in which it is assumed various configuration adjustments are made to the rotorcraft.

(2) Takeoff . The Category A takeoff path begins with an all-engines-operating acceleration segment to the engine failure point and continues with a one-engine-inoperative acceleration through the TDP to the takeoff safety speed (V ). The engine failure point (EFP) and TDP are separated by pilot recognition TOSS time. (See Conventional Takeoff Profile, figure AC 29.53A-1, paragraph AC 29.53A of this advisory circular.) TDP is a “go/no-go condition which is analogous to V1 speed in transport airplanes. Prior to TDP the pilot is “stop” oriented, and when an engine failure is recognized in this portion of the takeoff, the pilot will abort because the rotorcraft has not yet achieved sufficient energy to assure continued flight. At the TDP the pilot becomes “go” oriented and when an engine failure is recognized at or beyond this point, the pilot will continue the takeoff because sufficient surface area no longer remains for an aborted takeoff. The takeoff flight path and the TDP should be defined such that a safe landing can be made from any point up to the TDP. This profile may differ significantly from the takeoff flight path developed for Category B weights. The TDP is the last point in the takeoff profile at which a rejected takeoff capability within the scheduled takeoff surface distance is assured. If an engine failure does not occur, the pilot continues the climb and accelerates past the TDP to the recommended climb speed.

(3) Rejected Takeoff . The rejected takeoff profile begins with an all engine acceleration segment to the EFP and ends when the rotorcraft is brought to a complete stop on the designated takeoff surface. The critical engine is made inoperative prior to Page B - 46 9/30/99 AC 29-2C the TDP, and the landing should be made with the remaining engine(s) operating within approved limits. The rejected takeoff distance is normally measured at a given reference point on the rotorcraft from the start of the takeoff to the same reference point after the rotorcraft has come to a complete stop. This distance should be increased by the rotorcraft length (including main and tail rotor tip paths).

(4) Takeoff Path .

(i) The transition to OEI flight takes place between the engine failure point and the point at which V is achieved. It is assumed that the maximum TOSS approved OEI power is used until the allowable time duration for that power is exhausted. It should be possible for the crew to fly the rotorcraft to V and attain an TOSS altitude of 35 feet and positive rate of climb and then climb to 200 feet above the takeoff surface or the lowest point in the takeoff path by flying the rotorcraft solely by the primary flight controls (including collective). At no time during the takeoff shall the rotorcraft descend below 15 feet above the takeoff surface when the TDP is above 15 feet. The landing gear may be retracted after attaining a speed of V , and a TOSS positive rate of climb. Flight manual procedures may recommend adjustment of auxiliary controls to improve OEI performance, but compliance with the performance requirements of § 29.67(a)(1) may not be based on use of secondary engine controls such as RPM beep switches. During the continued takeoff profile, the pilot is assumed to be flying the rotorcraft via the primary flight controls (cyclic stick, collective, and directional pedals). Manipulation of the throttle controls or beep switches may be permitted as long as such manipulation can be accomplished readily by the pilot flying the rotorcraft without removing his hands from the cyclic and collective flight controls.

These manipulations of engine controls should not make major adjustments in power and should not occur before attaining V . In no case should this be less than TOSS 3 seconds after the critical engine is made inoperative. The failure of one engine cannot affect continued safe operation of the remaining engines or require any immediate action by the crew per § 29.903(b). If a 30-second/2-minute or a 2 ½-minute power rating is used, it should be possible to complete the Category A takeoff profile (assuming recognition of an engine failure at or prior to the TDP), accelerate to V , TOSS attain 35 feet above the surface, stabilize in a climb of at least 100 feet per minute, and complete landing gear retraction prior to exhausting the 2 ½-minute time limit.

(ii) The takeoff safety speed, V , is a speed at which 100 FPM rate of TOSS climb is assured under conditions defined in § 29.67(a)(1). The takeoff distance is the distance from the start of the takeoff to the point at which V , 35 feet above the TOSS takeoff surface, and a positive rate of climb are attained.

(5) Continued Climbout Path . Continued acceleration and climb capability are assured by the 100 FPM V climb requirement of § 29.67(a)(1) and the 150 FPM TOSS requirement of § 29.67(a)(2), normally demonstrated at V . It should be shown that the y rotorcraft can be accelerated from V to V in a continuous maneuver without losing TOSS Y altitude, including any configurative change (landing gear retraction, etc.). The distance Page B - 47 AC 29-2C 9/30/99 required to accelerate from V to V must be considered in determination of the TOSS Y climb and gradients required by § 29.1587(a)(6)(i) and (a)(6)(ii).

b. Procedures .

(1) Instrumentation . A photo theodolite, grid camera, GPS, or other position measuring equipment is normally required together with a ground station to measure wind, OAT, humidity (if applicable), and a two-way communication system to coordinate activities with the aircraft. A crash recovery team with support of a fire engine is highly desirable. Aircraft instrumentation should record with a time scale: engine parameters (speed, temperature, and power), rotor speed, flight parameters (airspeed, altitude, and normal acceleration as a minimum), flight control positions, power lever position, and landing gear loads. Additionally, a method should be devised to allow correlation of the aircraft instrumentation data with the space position data to accurately determine the length of the various takeoff segments.

(2) Establishing the Takeoff Decision Point (TDP) .

(i) The TDP should be definable with the minimum crew using standard cockpit instrumentation. If a radar altimeter is used, it should be included in the minimum equipment list. If barometric altitude is used to define TDP, the operating conditions at which the altimeter is set should be defined. This is normally done on the ground with the minimum collective pitch. If the wind influences the altimeter reading, the correct relative wind information should be provided. Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the TDP becomes largely a function of the surface area required for takeoff. If takeoff conditions scheduled include considerable surface area (on the order of 2,000 feet), the TDP airspeed may be a high value near V . This will allow a higher takeoff weight and Y demonstrate compliance with the V climb requirement of § 29.67(a)(1). In this case, TOSS the requirements of § 29.67(a)(2) usually become limiting. If required surface area is a small value, TDP will necessarily be some lower airspeed value to allow for an aborted takeoff on the available surface. Weight may need to be reduced at lower values of TDP airspeed (significantly below V ) to allow compliance with the climb requirement of Y § 29.67(a)(1). Compliance with climb requirements can be substantiated initially by testing at a safe altitude above the ground. When OEI climb conditions are verified for weight, configuration, pressure altitude, and temperature, the TDP is then evaluated in a rejected takeoff.

(ii) A Category A takeoff procedure should satisfy the following conditions: (A) The flightcrew takeoff procedure should be shown to be consistently repeatable and not require exceptional piloting skill.

(B) It should be documented that the takeoff procedure will produce the required minimum energy level in terms of height and airspeed for all combinations of Page B - 48 9/30/99 AC 29-2C gross weight, altitude, and ambient temperature for which takeoff data are scheduled.

This may best be accomplished by conducting takeoff procedure abuse tests to show that variations from the established takeoff procedure that could reasonably be expected to occur in service do not result in significant increases in the takeoff distances.

(3) Rejected Takeoff Distance . The rejected takeoff is similar in many respects to the height-velocity (HV) tests described in paragraph AC 29.73. Most of the comments, cautions, and techniques for HV also apply here even though typical flight conditions at TDP are less critical than limiting HV points. As mentioned in paragraph AC 29.79, a minimum 5-knot clearance from any HV limiting condition should be provided throughout the takeoff flight path (see figure AC 29.63-1), and tests should be conducted simulating an unplanned engine cut. The HV diagram appropriate to the Category A test weights may be much less restrictive than that determined for Category B conditions. Normally, a minimum 1-second delay (or pilot reaction time, whichever is greater) is applied after engine failure recognition, before pilot collective control corrections are allowed. If the rotorcraft incorporates an engine failure warning device, engine failure recognition should not be less than the time required for the engine to spool down and activate the device. As in all engine failure testing, the pilot should not anticipate the failure by changing flight control positions or aircraft attitude.

Average pilot techniques should be used. The two primary objectives of rejected takeoff testing are an assured capability to safely return to the takeoff surface when an engine failure is recognized at any point prior to TDP and the determination of the rejected takeoff distance required. It is important that the surface conditions be defined. The rejected takeoff distance tests will be used together with the OEI continued takeoff profiles to establish the required surface area for Category A operations.

(4) Takeoff Distance .

(i) Continued takeoff profiles should be flown to determine the continued takeoff distance. This distance is measured from the point of takeoff initiation to the point in the takeoff profile where the following three conditions have all been attained after a failure of the critical engine prior to TDP: an airspeed equal to or greater than V , a positive rate of climb, and a height of at least 35 feet above the takeoff surface.

TOSS If the rotorcraft descends below 35 feet above the takeoff surface while accelerating to V , the takeoff distance is extended to the point that 35 feet is reattained with a TOSS positive rate of climb.

(ii) If the TDP is significantly above 35 feet so that the rotorcraft does not descend below 35 feet during acceleration to V , the takeoff distance then becomes TOSS the distance to the point in the takeoff profile at which both V and a positive rate of TOSS climb are attained after failure of the critical engine prior to the TDP. For all applications, rotorcraft should not be allowed to descend below 15 feet above the takeoff surface while accelerating to V when TDP is above 15 feet. When TDP is TOSS below 15 feet, the aircraft should be able to accelerate in level flight or climb.

Fifteen feet should be considered the absolute minimum clearance allowed with greater Page B - 49 AC 29-2C, Chg 2 4/25/06 clearances required for some rotorcraft dependent on rotorcraft geometry and performance characteristics. In addition, the rotorcraft should not be allowed to descend below the height above the takeoff surface at which a landing flare would normally be initiated. For example, a medium size twin-engined rotorcraft with a TDP of 100 feet or greater, using 20° nose down, would be expected to clear the ground by 25 feet whereas a large multiengined rotorcraft, using similar attitudes and TDP’s, would be expected to clear by 35 feet. For elevated heliports the rotorcraft may descend below the landing surface, but all parts of the rotorcraft must clear the heliport and all other obstacles by not less than 15 feet. These minimum heights would need to be demonstrated with variations in piloting techniques and with pilot recognition and reaction times for engine failures occurring before and after TDP.

(iii) In establishing the continued takeoff distance, the applicable pilot recognition delay time should be applied following the engine failure prior to the TDP, and the takeoff profile should be established with the pilot using primary flight controls only to control the rotorcraft. The pilot engine failure recognition time delay before adjustment of the collective pitch control should be a minimum of 1 second.

(iv) Engine failure testing should be initially conducted at a safe distance above the ground to assess the continued takeoff profile before conducting the actual profiles for credit. This procedure will serve to validate predicted performance and may prevent an unexpected return to the surface during continued takeoff tests. A minimum of five acceptable runs should be flown by the FAA/AUTHORITY pilot, and these should be averaged with five acceptable runs flown by the manufacturer’s pilot.

(5) Abuse Testing . Takeoff procedure abuse tests should be conducted to show that reasonably expected variations in service from the established takeoff procedures do not result in a significant increase in the established takeoff distances.

Variations should include such considerations as under or over rotation during the takeoff initiation, under or over application of acceleration power, and missed TDP target parameters (e.g., time, height, or airspeed).

(6) Continued Climbout Path . The landing gear may be retracted at 35 feet.

The climb should be continued at V until 200 feet above the takeoff surface. The TOSS climb requirements of § 29.67(a)(1) should be met at 200 feet. It should be demonstrated that the rotorcraft, including any configuration changes, can be accelerated from V to V in a continuous maneuver without losing altitude. The TOSS Y airspeed and rotorcraft configuration (landing gear position, rotor RPM engine power, etc.) used to show compliance with the climb requirements of § 29.67(a)(2) should be attained at or prior to reaching 1,000 feet above the takeoff surface.

(7) Power . Power used for demonstrating performance should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping (maximum power available) to minimum specification values including consideration of temperature effects on engine power. If topping results in unrepresentative engine power management, the validity of the Cat A Page B - 50 4/25/06 AC 29-2C, Chg 2 procedure must also be established with representative in-service characteristics. The method used for simulating engine failure must be representative of the power decay characteristics that will occur during a real, sudden engine failure and acceleration of the remaining engine(s). In order to cushion a rejected take-off, it is acceptable for the engine and transmission transient range to be entered in order to droop the rotor provided performance credit is not taken for this additional power above the maximum permitted rating and it can be shown that the engine(s) will remain within these limits in all conditions requested by the applicant. Any excursion beyond established transient limits in this flight phase should be substantiated to the extent that it does not constitute an immediate hazard to the rotorcraft.

(8) Turbine engine power does not vary directly with density altitude (H ). At a D given H , turbine engine power available varies with ambient temperature. Turbine D engines typically produce less horsepower as ambient temperature is increased (pressure altitude decreases) at a given density altitude, although some engines produce less horsepower at extremely cold temperatures. In either event, if one test sequence is to be utilized for a given H , it would be appropriate to restrict test power to D the lowest value attainable from a minimum specification engine through the approved ambient temperature range at the density altitude of the test. To attain maximum weights for varying ambient conditions, the applicant may utilize a parametric mapping of power available, pressure altitude, and temperature effects. For this case, engine topping may be adjusted throughout a range appropriate to the test H .

D (9) Aircraft Loading . Both forward and aft CG extremes should be briefly checked to determine the critical loading for takeoff distances. Forward center of gravity is usually critical for continued takeoff distance tests while aft CG may be critical for the rejected takeoff due to forward and downward field of view. A minimum of two weights should be flown at each altitude if the manufacturer elects to schedule field length variation as a function of gross weight. One weight should be the maximum weight for prevailing conditions and the other weight(s) should be low enough to attain a sufficient spread to verify weight effect.

(10) Extrapolation . Takeoff and landing data may be extrapolated up to 4000 feet along an established W/ σ line, to the maximum gross weight of the rotorcraft.

However, extrapolation will not be considered valid if unacceptable or marginally acceptable landing gear loads are experienced during testing at weights below the W/ σ limit. See paragraph AC 29.77b(5) for further discussion of landing gear loads.

(11) Ambient Conditions . Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data should be corrected for existing wind conditions during takeoff distance testing. Credit for headwind conditions may be given during flight manual data expansion. Refer to paragraph AC 29.1587(a)(3)(iii) under “Wind Accountability” for allowable wind credit.

Care should be applied in considering headwind credit for vertical operations as previous experience has resulted in difficulty collecting meaningful, repeatable data.

Page B - 51 AC 29-2C, Chg 2 4/25/06 (12) Vertical Takeoffs .

(i) General . Guidelines for rotorcraft certification using vertical takeoff techniques were developed and utilized for civil certification programs many years ago.

As experience has been gained, certain policy decisions have modified these guidelines. The reader should be familiar with the preceding discussion regarding conventional Category A takeoff profiles because duplicate information is not repeated here.

(ii) Takeoff Profile . A typical vertical takeoff profile for a ground level heliport is shown in figure AC 29.59A-1. The maneuver begins with the addition of sufficient power to initiate a climb to the TDP. It should be possible to make a safe landing without exceptional pilot skill if an engine fails at any point up to the TDP less engine failure recognition time. At the TDP, the pilot becomes “go” oriented and continues the takeoff if an engine fails. The rotorcraft should not be allowed to descend below 15 feet above the takeoff surface during the continued takeoff. A typical profile for elevated heliports takeoff conditions is shown in figure AC 29.59A-2. Descent profile below the takeoff surface is allowed, after clearing the platform by at least a 15 feet radial margin, provided that the drop down height from the takeoff surface and the distance to reach V with a positive rate of climb is given in the performance chapter TOSS of the RFM.

(iii) Takeoff Decision Point (TDP) . For vertical takeoffs, the climb to the TDP is nearly vertical, and the TDP is typically defined primarily by height. Sufficient testing should be conducted to define a band of TDP conditions (heights) which will be consistent with anticipated variations in pilot technique and the minimum amount of equipment to be installed on the production aircraft. Rejected takeoffs are most critical from high TDP’s, and continued OEI takeoffs are most critical from low heights. Tests at the extremes of this band are intended to verify that the anticipated TDP band is safe and repeatable in service for reasonable variations in pilot technique. These extreme points should not be used for distance determination when averaging takeoff performance data.

(iv) Conduct of the Test . Vertical takeoff profiles should be flown from a pad simulating operational conditions because the sight picture may be critical to successful OEI operations, particularly for elevated heliports. At all points on the vertical takeoff flight path up to the TDP, the pilot, with reasonable head movement, shall be able to keep sufficient portions of two heliport boundaries (front and one side) or equivalent markings in view to achieve a safe landing in case of engine failure.

Normally, a minimum 1-second delay or pilot recognition time interval, whichever is greater, is applied after the EFP before pilot collective control corrections are allowed. If the rotorcraft incorporates an engine failure warning device, engine failure recognition should not be less than the time required for the engine to spool down and activate the device.

Page B - 52 4/25/06 AC 29-2C, Chg 2 (A) Establish the rejected takeoff distance as the horizontal distance from the rearmost point of the rotorcraft at the initiation of takeoff to the foremost point after the rotorcraft comes to a stop on the takeoff surface (including rotor tip path), assuming an engine failure in the vertical climb at the TDP.

(B) Establish the continued takeoff distance as the horizontal distance from lift-off to the point at which, following engine failure prior to the TDP, the rotorcraft achieves; for a ground level heliport, 35 feet above the takeoff surface and V with a TOSS positive rate of climb; for an elevated heliport, the lowest point of the takeoff profile and not less than V with a positive rate of climb. The continued takeoff profile from TOSS elevated heliports should clear the heliport obstructions by at least a 15 feet radial margin.

(C) When used, the back-up technique usually requires the pilot to keep sufficient portions of the helipad in view and involves a rearward movement from the takeoff point to the TDP. In such cases the rearward horizontal distance required should be established as the distance from the rearmost point of the rotorcraft at the initiation of takeoff to the rearmost part of the rotorcraft at TDP. As stated in AC 29.45, crosswinds and tailwinds should be considered if requested by the applicant. Typically, this will require flight-testing to evaluate performance, pilot workload, field-of-view, and visual cueing.

(D) If special helipad markings or other non-standard external references are required to achieve the vertical takeoff performance, these special references should be included in the limitations section of the RFM.

(v) Climb Requirements .

(A) Ground level heliport . The OEI takeoff profile should include a climb at V to 200 feet above the takeoff surface then an acceleration in level flight from TOSS V to V and a climb at V to 1000 feet above the lowest point of the takeoff profile.

TOSS Y Y The climb requirements of § 29.67(a)(1) and (a)(2) may be met at referenced points located respectively at 200 feet and 1000 feet above the takeoff surface. The distance required to accelerate from V to V must be considered in determination of the TOSS Y climb gradient required by § 29.1587 (a)(6)(i) and (a)(6)(ii).

(B) Elevated heliport . The OEI takeoff profile should include a climb at V to 200 feet above the lowest point of the takeoff profile then an acceleration in TOSS level flight from V to V and a climb at V to 1000 feet above the lowest point of the TOSS Y Y takeoff profile. The climb requirements of § 29.67(a)(1) and (a)(2) may be met at referenced points located respectively at 200 feet and 1000 feet above the lowest point of the takeoff profile.

(vi) Extrapolation . Basic guidelines for extrapolation are contained in paragraph AC 29.45. Weight can not be extrapolated above test weight. Altitude extrapolation should be limited to a maximum of ± 4000 feet.

Page B - 53 AC 29-2C, Chg 2 4/25/06 (13) Night Operations .

(i) A minimum of three normal takeoffs (and landings) should be conducted to ensure that aircraft lighting (internal and external) is adequate to allow normal Category A operations at night.

(ii) Engine failures should be simulated from points along the requested takeoff and landing profiles. Night OEI rejected takeoffs and continued takeoffs from the TDP and OEI landings from the LDP should be conducted at the requested WAT limiting conditions to ensure adequate night field of view, suitability of aircraft external lighting, and meets the Category A profiles.

(iii) If special airfield marking or lighting is used as a reference or to define the TDP, the aircraft external lighting should be evaluated to assure that the airfield marking or lighting is adequate for night operations.

Page B - 54 9/30/99 AC 29-2C ALL ENGINE CLIMBOUT ► ACCELERATE TO Vy I/\IHEN CLEAR OF OBSTACLES OEI CLIMBOUT ACCELERATE TO Vross ---- TDP VERTICAL 15' CLIMB FIGURE AC 29.59A-1 CATEGORY A VERTICAL TAKEOFF PROFILE GROUND LEVEL HELIPORT Page B - 55 AC 29-2C 9/30/99 ALL ENGINE CLIMBOUT ACCELERATE TOP-~ TO Vross

VERTICAL i

/

CLIMB & REJECTED TAKEOFF PATH OEI FLIGHT PATH I Vross and poslti11e rate of climb I Takeoff distance FIGURE AC 29.59A-2 CATEGORY A VERTICAL TAKEOFF PROFILE ELEVATED HELIPORT Page B - 56 9/30/99 AC 29-2C AC 29.63. § 29.63 (Amendment. 29-12) TAKEOFF: CATEGORY B .

a. Explanation .

(1) Takeoff distance is the horizontal distance measured from an initial position to a point 50 feet above the takeoff surface with all engines operating within approved limits.

(2) The height-velocity diagram is normally developed and accepted prior to conducting takeoff distance tests. Takeoff distance tests are conducted avoiding the critical areas of the diagram. The amount of power utilized in determining takeoff distance may not be greater than that used in constructing the takeoff corridor and “knee” portions of the height-velocity diagram. Power might also have to be constrained, depending upon the amount of excess power available, so that a “reasonable” nose down pitch attitude is not exceeded during the initial portion of the takeoff run. Acceptable values used during past programs include: (i) Hover power + 10 percent (not to exceed rated engine takeoff power limits) (ii) A percent transmission limiting torque (not to exceed rated engine takeoff power limits), and (iii) Engine (or transmission) limiting power for the particular ambient conditions.

(3) The critical center of gravity should be used for takeoff distance tests.

Critical center of gravity should be established analytically or from previous testing and may be forward or aft depending on the type of rotorcraft. Items that should be considered in determining the critical center of gravity are climb performance and cockpit visibility. At least two gross weights should be flown at each test altitude, if weight accountability is desired, in order to validate the manufacturers prediction of weight effects.

(4) The speed utilized at the 50-foot point in the takeoff profile (V speed) may be largely determined by the ability to obtain reliable, repeatable airspeed indications which can also comply with § 29.1323. Section 29.1323 ties the airspeed system accuracy requirements to the climbout speed. The climbout speed should be that speed attained at 50 feet in complying with § 29.63.

b. Procedures .

(1) Instrumentation . A ground station will measure ambient temperature, humidity (if applicable), and wind. For allowable wind conditions and engine power considerations refer to paragraph AC 29.45. A photo panel or hand recording method Page B - 57 AC 29-2C 9/30/99 may be utilized, as necessary, to record engine and flight parameters. A phototheodolite, takeoff and landing camera, or other approved instrumentation is utilized to measure distance, heights, speed, and time.

(2) Conduct of the Test . If the applicant elects to show weight effects on distance, at least two weights should be flown and, depending on the range of takeoff and landing altitudes to be approved, at least two test altitudes should be flown.

Altitudes should be sufficiently far apart to include a major portion of the approved takeoff and landing altitude range. Takeoff profiles should be started from an initial condition. For takeoffs from a hover, the hover height should be determined by performing fixed collective takeoffs as described in paragraph AC 29.51. “Takeoff” power should be smoothly applied and the aircraft nose lowered as necessary to accelerate without gaining excessive altitude. It must be possible to conduct a consistent takeoff profile clear of the height-velocity diagram with normal pilot effort and skill. A minimum of five good runs should be flown by the FAA/AUTHORITY pilot at each altitude and weight. Runs by the company and FAA/AUTHORITY pilot may be averaged. Effects of missing the V speed by some amount (± 5 knots, for example) or other small changes in profile should be evaluated to determine if gross performance changes result from small piloting errors. Engine failures should be conducted along the takeoff profile to assure safe landing capability. Past programs have shown the low speed point immediately after addition of power to be particularly critical. Night takeoffs should at least be qualitatively evaluated to assure the takeoff procedures are compatible for night operation.

(3) Test Results . Test results are utilized in constructing the flight manual takeoff distance charts required by § 29.1587. The takeoff surface utilized in conducting these takeoff distance and engine failure tests should be included in the flight manual.

The “climbout speed” should also be defined and included in the flight manual. The airspeed utilized at the 50-foot point in the conduct of these tests must be clearly defined to allow compliance with § 29.1323. Test results may be extrapolated in accordance with guidance contained in paragraph AC 29.45.

(4) Test Techniques . For the FAA/AUTHORITY test data runs which will result in rotorcraft flight manual (RFM) performance, only the operational cockpit instrumentation as shown on the minimum equipment list and the piloting procedures from the RFM should be used. A useful technique is to “lead” the targeted V speed by a fixed amount, so that a smooth, consistent, and operationally realistic transition may be made between the acceleration and climbout phases; e.g., begin rotation at 35 knots to achieve 46 knots passing 50 feet. This and other pertinent information defining the takeoff flight path are required flight manual entries per § 29.1587(b).

Page B - 58 9/30/99 AC 29-2C w _J LL 0:: 0...

LL LL z @ :::) ::,:: 0:: ~ w~

> tu

OW AVOID CD LL ~~ ~ w I .,___ 5 KTS MINIMUM MARGIN AVOID VELOCITY (KNOTS) Fl GURE AC 29.63-1 CONVENTIONAL TAKE OFF PROFILE CATEGORY B Page B - 59 AC 29-2C 9/30/99 AC 29.65. § 29.65 (Amendment 29-15) CLIMB: (ALL ENGINES OPERATING) .

a. Explanation .

(1) Section 29.65 requires in part that the steady rate of climb be determined for each Category B rotorcraft with maximum continuous power on each engine for the range of weights, altitudes, and temperatures for which certification is requested. The climb airspeed should be the best rate-of-climb (V ) for standard day sea level Y conditions at maximum weight and at a speed(s) selected by the applicant for other conditions not to exceed V . The applicant can either publish a climb schedule in NE accordance with the above or utilize a constant climb airspeed for all conditions.

Equivalent levels of safety have been found wherein the applicant was allowed to select a climb airspeed that was not the actual V . The selected airspeed must be consistent Y with the speed used to show compliance with such items as cooling, stability, etc. The rate-of-climb resulting from the selected climb airspeed versus that from the actual V Y shall not differ to an extent that a pilot will be encouraged, by appreciable increases in climb performance to fly a climb airspeed different from that published in the Flight Manual.

(2) For Category A rotorcraft, if V at any altitude is less than the maximum NE gross weight sea level standard day condition V , the steady rate-of-climb must be Y determined at the climb speed(s) selected by the applicant not to exceed V . The NE climb performance must be determined from 2,000 feet below the altitude from where V intersects V up to the maximum altitude for which certification is requested. This NE Y should be done utilizing maximum continuous power on each engine with the landing gear retracted.

b. Procedure to Determine V .

Y (1) Sawtooth climbs may be used to determine the best rate-of-climb airspeed V . If such a technique is used, climbs should be flown in pairs on opposite headings Y 90° to the winds at the test altitude. This procedure will minimize any windshear effects.

All testing should be done in smooth air. Windshear is usually an indication of unstable air or a temperature inversion and should be avoided. The climbs are flown on reciprocal headings for approximately 5 minutes through a 1,000-foot band, or a comparable time/altitude band, using maximum continuous power at a constant airspeed. Periodic power adjustments may be necessary. Additional reciprocal heading climbs must also be conducted at different airspeeds sufficient to bracket the lowest point of the power required versus airspeed curve. This technique can be repeated at different altitudes to obtain V throughout the altitude range.

Y (2) Level flight performance (speed power) may also be used to determine the best rate-of-climb airspeed (V ). The testing should be done in smooth air. The Y advantage of this method is that less time is required, and the accuracy is equivalent to the sawtooth climb method. The test can be repeated at various altitudes to determine the V throughout the altitude range desired for the rotorcraft. The test at each altitude Y Page B - 60 9/30/99 AC 29-2C should be conducted at a constant weight over sigma (W/ σ ). The test is normally started at the desired W/ σ with maximum continuous power, or at V , in level flight. A NE series of points should be taken, reducing airspeed 10 to 15 knots between points, with the lowest speed point at approximately 20 to 30 knots. Weight should be computed for each point and the test altitude adjusted to maintain a constant W/ σ . After the data are reduced to standard day conditions, the minimum power required airspeed will be the V speed.

Y (3) Prior to the flight test, the rotorcraft should be ballasted to the desired gross weight and the critical center of gravity. The airspeed should be stabilized prior to data acquisition. Data to be recorded includes time, altitude, airspeed, ambient temperature, engine parameters, torque(s), rotor RPM, fuel reading, aircraft heading, external configuration, etc. Power setting, weight, and climb airspeed should be planned prior to flight. For some turboshaft engines, temperature and/or engine speed limits may be reached prior to a limiting torque. The test team should verify that the resulting power utilized in these tests closely approximates the power producing capabilities of installed minimum specification engine.

c. Procedure to Determine all Engine Operating Climb Performance.

(1) Background . Continuous climbs are conducted at the appropriate climb airspeeds as outlined above in order to obtain the rotorcraft’s climb performance for the flight manual. By-products are a qualitative evaluation of the rotorcraft handling characteristics in a climb and engine data to assist in the determination of installed power available.

(2) Techniques . The techniques used to determine this performance may be the same as those used in the V determination. The climbs are conducted on Y reciprocal headings at the established airspeed(s) through the target altitude range.

The same parameters are recorded. The rotorcraft will usually climb very rapidly during the first few thousand feet; therefore, the data acquisition method must be timely if accurate results are expected. This procedure is usually repeated at weight extremes.

The resulting data must then be corrected for power and weight. Power and weight corrections are satisfactory, provided the test powers and weights closely approximate the target values to make the weight and power corrections accurate. Once this data is finalized and corrected for all the flight test variables, interpolation for intermediate weights can be made with a high degree of reliability. If the rotorcraft has any stability augmentation system, vent systems, etc., which may influence the climb performance, then it must be accounted for. Caution should be taken that anti-ice, air-conditioning, etc., are not on unless the performance is being established specifically for those conditions.

AC 29.65A (AC 29.64) §§ 29.64 and 29.65 (Amendment 29-39) CLIMB (GENERAL AND ALL ENGINES OPERATING) .

a. Explanation .

Page B - 61 AC 29-2C 9/30/99 (1) Amendment 29-39 relocated and clarified the general climb requirements into a new § 29.64 and added requirements to determine Category A climb performance in § 29.65. The guidance material presented in paragraph AC 29.67 does not apply to rotorcraft certified with Amendment 29-39 or later. Sections 29.64 and 29.65 require that the steady rate of climb be determined with maximum continuous power on each engine for the range of weights, altitudes, and temperatures for which certification is requested. The climb airspeed should be the best rate-of-climb (V ) for standard day Y sea level conditions at maximum weight and at a speed(s) selected by the applicant for other conditions not to exceed V . The applicant can either publish a climb schedule in NE accordance with the above or utilize a constant climb airspeed for all conditions.

Equivalent levels of safety have been found wherein the applicant was allowed to select a climb airspeed that was not the actual V . The selected airspeed should be Y consistent with the speed used to show compliance with such items as cooling, stability, etc. The rate-of-climb resulting from the selected climb airspeed versus that from the actual V shall not differ to an extent that a pilot will be encouraged by appreciable Y increases in climb performance to fly a climb airspeed different from that published in the Flight Manual.

(2) If V at any altitude is less than the maximum gross weight sea level NE standard day condition V , the steady rate-of-climb should be determined at the climb Y speed(s) selected by the applicant not to exceed V . The climb performance should NE be determined from 2,000 feet below the altitude from where V intersects V up to the NE Y maximum altitude for which certification is requested. This should be done utilizing maximum continuous power on each engine with the landing gear retracted.

b. Procedure to Determine V .

Y (1) Sawtooth climbs may be used to determine the best rate-of-climb airspeed V . If such a technique is used, climbs should be flown in pairs on opposite headings Y 90° to the winds at the test altitude. This procedure will minimize any windshear effects.

All testing should be done in smooth air. Windshear is usually an indication of unstable air or a temperature inversion and should be avoided. The climbs are flown on reciprocal headings for approximately 5 minutes through a 1,000-foot band, or a comparable time/altitude band, using maximum continuous power at a constant airspeed. Periodic power adjustments may be necessary. Additional reciprocal heading climbs should also be conducted at different airspeeds sufficient to bracket the lowest point of the power required versus airspeed curve. This technique can be repeated at different altitudes to obtain V throughout the altitude range.

Y (2) Level flight performance (speed power) may also be used to determine the best rate-of-climb airspeed (V ). The testing should be done in smooth air. The Y advantage of this method is that less time is required, and the accuracy is equivalent to the sawtooth climb method. The test can be repeated at various altitudes to determine the V throughout the altitude range desired for the rotorcraft. The test at each altitude Y should be conducted at a constant weight over sigma (W/ σ ). The test is normally Page B - 62 9/30/99 AC 29-2C started at the desired W/ σ with maximum continuos power, or at V , in level flight. A NE series of points should be taken, reducing airspeed 10 to 15 knots between points, with the lowest speed point at approximately 20 to 30 knots. Weight should be computed for each point and the test altitude adjusted to maintain a constant W/ σ . After the data are reduced to standard day conditions, the minimum power required airspeed will result in the airspeed for maximum rate of climb. However, aircraft stability may suggest that a higher climb speed may be used for V Y.

(3) Prior to the flight test, the rotorcraft should be ballasted to the desired gross weight and the critical center of gravity. The airspeed should be stabilized prior to data acquisition. Data to be recorded includes time, altitude, airspeed, ambient temperature, engine parameters, torque(s), rotor RPM, fuel reading, aircraft heading, external configuration, etc. Power setting, weight, and climb airspeed should be planned prior to flight. For some turboshaft engines, temperature and/or engine speed limits may be reached prior to a limiting torque. The test team should verify that the resulting power utilized in these tests closely approximates the power producing capabilities of installed minimum specification engine.

c. Procedure to Determine all Engine Operating Climb Performance .

(1) Background . Continuous climbs are conducted at the appropriate climb airspeeds as outlined above in order to obtain the rotorcraft’s climb performance for the flight manual. By-products are a qualitative evaluation of the rotorcraft handling characteristics in a climb and engine data to assist in the determination of installed power available.

(2) Techniques . The techniques used to determine this performance may be the same as those used in the V determination. The climbs are conducted on Y reciprocal headings at the established airspeed(s) through the target altitude range.

The same parameters are recorded. The rotorcraft will usually climb very rapidly during the first few thousand feet; therefore, the data acquisition method should be timely if accurate results are expected. This procedure is usually repeated at weight extremes.

The resulting data should then be corrected for power and weight. Power and weight corrections are satisfactory, provided the test powers and weights closely approximate the target values to make the weight and power corrections accurate. Once this data is finalized and corrected for all the flight test variables, interpolation for intermediate weights can be made with a high degree of reliability. If the rotorcraft has any stability augmentation system, vent systems, etc., which may influence the climb performance, then it should be accounted for. Caution should be taken that anti-ice, air-conditioning, etc., are not on unless the performance is being established specifically for those conditions.

AC 29.67. § 29.67 (Amendment 29-34) CLIMB: ONE ENGINE INOPERATIVE .

a. Explanation .

Page B - 63 AC 29-2C 9/30/99 (1) Section 29.67 requires that Category A rotorcraft must be capable of a steady rate-of-climb without ground effect, of at least 100 feet per minute for all combinations of weight, altitude, temperature, and center of gravity for which takeoffs are to be scheduled. The rate-of-climb is determined with the critical engine inoperative and the remaining engine(s) operating within approved operating limits. The landing gear is extended and the airspeed is the takeoff safety speed (V ) selected by the TOSS applicant.

(2) In addition, the steady rate-of-climb must be at least 150 feet per minute at 1,000 feet above the takeoff surface for which takeoffs are to be scheduled. The rate-of-climb will be determined with the critical engine inoperative and the remaining engine(s) at maximum continuous or the 30-minute minimum specification installed power available values. The landing gear is retracted and the airspeed is that selected by the applicant.

b. Procedures .

(1) One of the acceptable procedures used to obtain the required climb performance is similar to the all engine climb performance determination (paragraph AC 29.65) except that the V and the Category A climb speed may be TOSS selected by the applicant for different weights and ambient conditions. The Category A climb speed could be a single speed, vary as V does, or actually be V . Making a Y Y Category A climbout speed equal to V should be encouraged to simplify cockpit Y procedures. The required results are the allowable weight, altitude, and temperature combinations wherein the rotorcraft is capable of demonstrating 100 feet per minute rate-of-climb at V and 150 feet per minute rate-of-climb at 1,000 feet above the TOSS takeoff surface. Either of these two climb requirements may establish the maximum allowable takeoff weight.

(2) For multiengine Category B rotorcraft with engine isolation, the steady rate of climb or descent must be determined at V , using maximum continuous power and Y 30-minute power if that rating is approved. Appropriate performance data must be included in the Rotorcraft Flight Manual to cover variations in gross weight, altitude, and temperature.

(3) Since climb performance testing is normally conducted separately from Category A and B takeoff performance testing, it is imperative the engine power(s), rotor RPM, and aircraft configuration be the same as those used during the takeoff testing to ensure the climb performance demonstrated will be that attainable immediately after an engine failure during takeoff. The allowable pilot/crew actions during the Category A takeoff and climbout maneuver must be thoroughly evaluated. The pilot’s full attention is required to control the rotorcraft during this phase of flight. Permitting the pilot to readjust (beep) the rotor RPM during this phase of flight should be considered only if such adjustment can be accomplished without a significant increase in pilot workload.

Page B - 64 9/30/99 AC 29-2C (4) A typical sequence for selecting the various speeds to comply with this requirement is as follows: (i) Conduct sawtooth climbs at the various airspeeds (V and below) up Y to the proposed takeoff and landing altitudes. From this a determination can be made regarding the maximum allowable weight that will result in a rate of climb of 150 feet per minute at the selected V for the proposed ambient conditions.

Y (ii) At the same time determine the minimum value of V that will result TOSS in 100 feet per minute rate of climb at the maximum weight determined in (b)(4)(i).

AC 29.67A. § 29.67 (Amendment 29-39) CLIMB: ONE ENGINE INOPERATIVE .

a. Explanation .

(1) Amendment 29-39 expanded the OEI rate of climb requirements. The guidance material presented in paragraph AC 29.67 does not apply to rotorcraft certified with Amendment 29-39 or later. Section 29.67 requires that Category A rotorcraft should be capable of a steady rate-of-climb without ground effect 200 feet above the takeoff surface, of at least 100 feet per minute for all combinations of weight, altitude, temperature, and center of gravity for which takeoffs are to be scheduled. The rate-of-climb is determined with the critical engine inoperative and the remaining engine(s) operating within approved operating limits. The landing gear is extended and the airspeed is the takeoff safety speed (V ) selected by the applicant.

TOSS (2) The steady rate-of-climb should be at least 150 feet per minute at 1,000 feet above the takeoff surface for which takeoffs are to be scheduled. The rate-of-climb will be determined with the critical engine inoperative and the remaining engine(s) at maximum continuous or the 30-minute minimum specification installed power available values. The landing gear is retracted and the airspeed is that selected by the applicant.

(3) Additionally, the steady state rate of climb or descent must be determined with the critical engine inoperative and the remaining engines at OEI maximum continuous power and at 30-minute OEI power if applicable. This performance must be scheduled throughout the ranges of weight, altitude and temperatures for which certification is requested with the landing gear retracted, at an airspeed selected by the applicant.

b. Procedures .

(1) One of the acceptable procedures used to obtain the required climb performance is similar to the all engine climb performance determination (paragraph AC 29.65) except that the V and the Category A climb speed may be TOSS selected by the applicant for different weights and ambient conditions. The Category A climb speed could be a single speed, vary as V does, or actually be V . Making a Y Y Category A climbout speed equal to V should be encouraged to simplify cockpit Y Page B - 65 AC 29-2C 9/30/99 procedures. The required results are the allowable weight, altitude, and temperature combinations wherein the rotorcraft is capable of demonstrating 100 feet per minute rate-of-climb at V at a height of 200 feet above the takeoff surface and 150 feet per TOSS minute rate-of-climb at 1,000 feet above the takeoff surface. Either of these two climb requirements may establish the maximum allowable takeoff weight.

(2) For multiengine Category B rotorcraft with engine isolation, the steady rate of climb or descent should be determined at V , using maximum continuous power, Y maximum continuous OEI power, and 30-minute power if that rating is approved.

Appropriate performance data should be included in the Rotorcraft Flight Manual to cover variations in gross weight, altitude, and temperature.

(3) Since climb performance testing is normally conducted separately from Category A and B takeoff performance testing, it is imperative the engine power(s), rotor RPM, and aircraft configuration be the same as those used during the takeoff testing to ensure the climb performance demonstrated will be that attainable immediately after an engine failure during takeoff. The allowable pilot/crew actions during the Category A takeoff and climbout maneuver should be thoroughly evaluated. The pilot’s full attention is required to control the rotorcraft during this phase of flight. Permitting the pilot to readjust (beep) the rotor RPM during this phase of flight should be considered only if such adjustment can be accomplished without a significant increase in pilot workload.

(4) A typical sequence for selecting the various speeds to comply with this requirement is as follows: (i) Conduct sawtooth climbs at the various airspeeds (V and below) up Y to the proposed takeoff and landing altitudes. From this, a determination can be made regarding the maximum allowable weight that will result in a rate of climb of 150 feet per minute at the selected V for the proposed ambient conditions.

Y (ii) At the same time, determine the minimum value of V that will TOSS result in 100 feet per minute rate of climb at the maximum weight determined in b(4)i.

AC 29.71. § 29.71 (Amendment 29-12) ROTORCRAFT ANGLE OF GLIDE: CATEGORY B.

a. Explanation .

(1) Performance capabilities during stabilized autorotative descent are useful pilot tools to assist in the management of a Category B rotorcraft when all engines fail.

This information is also useful in determining the suitability of available landing areas along a given route segment.

(2) Two speeds are of particular importance, the speed for minimum rate of descent and the speed for best angle of glide. These speeds are required as flight manual entries per § 29.1587. The speed for minimum rate of descent is useful for Page B - 66 9/30/99 AC 29-2C engine failure conditions at higher altitudes and the pilot is required to perform some time-related task, engine restart, float inflation, radio calls, etc. The speed for best angle of glide is a somewhat higher speed that is of particular use when it is necessary to reach a distant landing area. This speed, with appropriate rotor RPM, provides the maximum horizontal distance available from a particular altitude assuming zero wind conditions.

(3) A third speed, recommended autorotation speed, may be provided in addition to minimum rate of descent speed and maximum glide angle speed. The recommended speed for autorotation is usually optimized to assure an effective flare capability and yet be slow enough to allow a controlled, relatively slow touchdown condition. Recommended autorotation speed is ordinarily between the minimum rate of descent and maximum glide angle speeds. The recommended autorotation speed may be provided in the Rotorcraft Flight Manual. The relationship between minimum rate of descent, best glide angle, and recommended autorotation speed is shown in figure AC 29.71-1.

(4) Forward center of gravity is usually critical, however, center of gravity effects should be spot-checked to confirm this for a given design.

b. Procedures .

(1) Tests are conducted at speeds which bracket the anticipated speeds for minimum rate of descent and best glide angle. On a power required plot, the speed for minimum power required approximates the speed for minimum rate of descent. The speed for maximum range glide may be estimated by drawing a tangent from the origin to the power required curve.

(2) Autorotative performance tests may be conducted in conjunction with the climb performance tests. The required data are similar for both tests and it is sometimes convenient and efficient to run alternating climbs and descents through a desired altitude band. Descents should be conducted on reciprocal headings and results averaged in the same manner as climb performance tests.

(3) A reduction in rotor RPM from the normal power-on value may enhance autorotative performance. If the applicant wishes to develop autorotative performance at RPM values significantly below the governing or power-on range, the practicality of reducing and controlling RPM at the lower value and of then increasing RPM as a landing is approached, must be considered. At low weights and low density altitudes, full down collective may automatically produce lower RPM values and this condition is, of course, acceptable provided the approved power-off RPM range is not exceeded.

(4) Care must be taken to make certain that no engine power is delivered to the rotor drive system since a very small amount of power can have a large effect on descent performance.

Page B - 67 AC 29-2C 9/30/99 RATE OF DESCENT RECOMMENDED AUTOROTATIONAL SPEED

~

~

I

BEST GLIDE ANGLE MINIMUM / RATE OF DESCENT /

/

/

/

/

/

/

/

AIRSPEED Fl GURE AC 29. 71-1 AUTO ROTATIONAL CHARACTER! STI CS - TYPICAL Page B - 68 9/30/99 AC 29-2C AC 29.73. § 29.73 (Amendment 29-3) PERFORMANCE AT MINIMUM OPERATING SPEED. HOVER PERFORMANCE FOR ROTORCRAFT .

(For performance at minimum operating speed and for hover performance after Amendment 38, see § 29.49 and paragraph AC 29.49).

a. Explanation .

(1) For the purpose of this manual, the word “hover” applies to a rotorcraft that is airborne at a given altitude over a fixed geographical point regardless of wind. Pure hover is accomplished only in still air.

(2) The regulatory requirement for hover performance, § 29.73, refers to hover in ground effect (IGE). For some applications, such as external load operations, hover performance out-of-ground effect (OGE) is necessary; however, it is not required by this section. Hover OGE is that condition, where an increase in height above the ground will not require additional power to hover. Hover OGE is the absence of measurable ground effect. It can be less than one rotor diameter at low gross weight increasing significantly at high gross weights. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle 1 ½ rotor diameters above the surface.

(3) The objective of hover performance tests is to determine the power required to hover at different gross weights, ambient temperatures, and pressure altitudes.

Using nondimensional power coefficients (C ) and thrust coefficients (C ) for P T normalizing and presenting test results, a minimum amount of data are required to cover the rotorcraft’s operating envelope.

(4) Hover performance tests must be conducted over a sufficient range of pressure altitudes and weights to cover the approved ranges of those variables for takeoff and landings. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.

(5) The minimum hover height for which data should be obtained and subsequently presented in the flight manual should be the same height consistent with the minimum hover height demonstrated during the takeoff tests. Refer to paragraph AC 29.51 for the procedure to determine the minimum allowable hover height.

b. Procedures .

(1) Two methods of acquiring hover performance data are the tethered and free flight techniques. The tethered technique is accomplished by tethering the rotorcraft to the ground using a cable and load cell. The load cell and cable are attached to the ground tie-down and to the rotorcraft cargo hook. The load cell is used to measure the rotorcraft’s pull on the cable. Hover heights are based on skid or wheel height above Page B - 69 AC 29-2C 9/30/99 the ground. During tethered hover tests, the rotorcraft should be at light gross weight.

The rotorcraft will be stabilized at a fixed power setting and rotor speed at the appropriate skid or wheel height. Once the required data are obtained, power should be varied from the minimum to the maximum allowed at various rotor RPM. This technique will produce a large C /C spread. The load cell reading is recorded for each stabilized T P point. The total thrust the rotor produces is the rotorcraft’s gross weight, weight of the cables and load cell plus cable tension. Care must be taken that the cable tension does not exceed the cargo hook limit or load capacity of the tie-down. For some rotorcraft, it may be necessary to ballast the rotorcraft to a heavy weight in order to record high power hover data.

(2) The pilot maintains the rotorcraft in position so that the cable and load cell are perpendicular to the ground. To insure the cable is vertical, two outside observers, one forward of the rotorcraft and one to one side, can be used. Either hand signals or radio can be used to direct the pilot. The observers should be provided with protective equipment. This can also be accomplished by attaching two accelerometers to the load cell which sense movement along the longitudinal and lateral axes. Any displacement of the load cell will be reflected on instrumentation in the cockpit and by reference to this instrumentation, the rotorcraft can be maintained in the correct position. Increased caution should be utilized as tethered hover heights are decreased because the rotorcraft may become more difficult to control precisely. The tethered hover technique is especially useful for OGE hover performance data because the rotorcraft’s internal weight is low and the cable and load cell can be jettisoned in the event of an engine failure or other emergency.

(3) To obtain consistent data, the wind velocity should be less than 3 knots or less as there are no accurate methods of correcting hover data for wind effects. Large rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are: (i) Engine torques.

(ii) Rotor speed.

(iii) Ambient temperatures.

(iv) Pressure altitude.

(v) Fuel used (or remaining).

(vi) Load cell reading.

(vii) Generator(s) load.

As a technique, it is recommended the rotorcraft be loaded to a center of gravity near the hook to minimize fuselage angle changes with varying powers. All tethered hover Page B - 70 9/30/99 AC 29-2C data should be verified by a limited spotcheck using the free flight technique. The free flight technique as contained in paragraph b(4) below will determine if any problems, such as load cell malfunctions have occurred. The free flight hover data must fall within the allowable scatter of the tethered data.

(4) If there are no provisions or equipment to conduct tethered hover tests, the free flight technique is also a valid method. The disadvantage of this technique as the primary source of data acquisition is that it is very time consuming. In addition a certain element of safety is lost OGE in the event of emergency. The rotorcraft must be reballasted to different weights to allow the maximum C /C spread. When using the T P free flight technique, either as a primary data source or to substantiate the tethered technique, the same considerations for wind, recorded parameters, etc., as used in the tethered technique apply. Free flight hover tests should be conducted at CG extremes to verify any CG effects. If the rotorcraft has any stability augmentation system which may influence hover performance, it must be accounted for.

(5) It is extremely difficult to determine when a rotorcraft is hovering OGE at high altitudes above ground level since there is no ground reference. In a true hover, the rotorcraft will drift with the wind. Numerous techniques have been tried to allow OGE hover data acquisition at high altitudes, all of which have resulted in much data scatter. Until a method is proposed and found acceptable to the FAA/AUTHORITY, OGE hover data must be obtained at the various altitude sites where IGE hover data is obtained. Hover performance can usually be extrapolated up to a maximum of 4,000 feet.

AC 29.75. § 29.75 (Amendment 29-17) LANDING .

a. Explanation .

(1) This rule incorporates all of the landing performance requirements for transport category rotorcraft. It consolidates requirements for landing data, Category A landing, Category A flight data, and Category B landing. Parallel takeoff requirements are located in four separate sections of the rule, §§ 29.51 through 29.63. As such, to assure necessary subjects are treated separately, the following discussion will be separated into three parts: (a) a general discussion of basic landing distance requirements, (b) Category A requirements (including vertical landing), and (c) Category B requirements.

(2) All landing performance data are corrected to a smooth, dry, hard, level landing surface condition. As with other flight maneuvers, landings must be accomplished with acceptable flight and ground characteristics using normal pilot skills.

The rule states that Category A and B landing data must be determined at each approved WAT (Weight, Altitude, Temperature) condition. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in paragraph AC 29.45. As in other performance areas, engines must be Page B - 71 AC 29-2C 9/30/99 operated within approved limits. Power considerations are the same as those described under paragraph b(2)(ii)(C).

(3) Unlike fixed wing aircraft, rotorcraft typically require significantly more landing surface area with an engine inoperative than with all engines operating.

Because of this characteristic, the landing distance requirements are met with at least one engine inoperative to assure the most conservative landing distance measurement is achieved.

b. Procedures .

(1) Category A Requirements .

(i) Explanation . The Category A certification concept limits landing weight to a value that will allow the rotorcraft, following an engine failure at the landing decision point (LDP), to land within the available runway or to execute a balked landing, descending no lower than 35 feet above the landing surface. See figure AC 29.75-1.

(A) LDP . The Category A landing profile begins with an assumed engine failure at or prior to the LDP. The LDP is typically defined in terms of airspeed, rate of descent, and altitude above the landing surface. The approach path angle can be defined by LDP airspeed and rate of descent values. Definition of the LDP should include an approach angle because both the landing distance and the missed approach path are significantly influenced by landing approach angle. At any point in the single engine approach path down to and including the LDP, the pilot may elect to land or to execute a balked landing and he is assured both an adequate surface area for OEI landing and adequate climb capability for an OEI balked landing. Said another way, if an engine fails at any point down to and including the LDP, the pilot may safely elect to land or to “go around” by executing a balked landing. The LDP must be defined to permit acceleration to V at an altitude no lower than 35 feet above the landing TOSS surface. The LDP represents a “commit” point for landing. Prior to the LDP in the one engine inoperative approach, the pilot has a choice, he may either land or fly away.

After passing the LDP he no longer has sufficient energy to assure transition to a balked landing condition without contacting the landing surface. If an engine fails after LDP in a normal (all engine) landing the pilot is committed to land. The LDP and landing approach path must be defined such that the critical areas of the height-velocity diagram are avoided. A typical LDP for conventional Category A profiles is 100 feet above the landing surface. LDP should be specified in terms of both actual altitude above the landing surface and indicated barometric altitude. Speed at the LDP should be specified in terms of indicated airspeed.

(B) Landing distance . Approach and landing path requirements are stated in general terms in paragraphs (b)(2) and (4) of § 29.75. The approach path must allow smooth transition for one engine inoperative landing and for balked landing maneuvers and must allow adequate clearance from potentially hazardous HV combinations. Paragraph (b)(4)(ii) implies that a less restrictive HV envelope may exist Page B - 72 9/30/99 AC 29-2C for the Category A approach condition in comparison to that determined under high power conditions in § 29.79. The manufacturer may elect to use this added capability.

The added capability arises from the fact that lower power levels, a lower collective setting, and an established rate of descent accompany typical approach conditions as opposed to the more critical high power conditions of § 29.79. Landing distance is measured from a point 50 feet (25 feet for VTOL) above the landing surface to a stop.

For flight manual purposes, the distance is from the point at which the lowest part of the rotorcraft first reaches 50 feet (25 for VTOL) to the foremost point of the rotorcraft (including rotor tip path) after coming to a stop.

(C) All engine out landing . Section 29.75(b)(5) contains the Category A certification requirement for “last” engine failure and all engine inoperative landing. The rule states that it must be possible to make a safe landing on a prepared surface after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously. See paragraph AC 29.143a(2)(iii)(A) for the Category A sequential engine failure criteria. The conditions for last engine failure are maximum continuous power or 30-minute power if that rating is approved, “wings” level flight, and sudden engine failure with a pilot delay of 1 second or normal pilot recognition time, whichever is greater. Complete power failure has occurred in twin engine Category A rotorcraft.

This requirement ensures that in the event of cockpit mismanagement, fuel exhaustion, improper maintenance, fuel contamination, or unforeseen mechanical failures, a safe autorotation entry can be made and a safe power-off landing can be affected. Two separate aspects of this rule are normally evaluated at different times during the test program. The last engine failure is normally evaluated during cruise or V engine NE failure testing where instrumentation and critical loading have been established for those test conditions. See discussion under paragraph AC 29.143. The all engine out landing is ordinarily conducted in conjunction with an HV or Category A landing distance phase where ground instrumentation and safety equipment are available. The rotorcraft must be capable of conducting the all engine out landing at the takeoff and landing WAT limiting conditions up to the maximum altitude approved for takeoff and landing.

(ii) Procedures .

(A) Instrumentation/Equipment . Instrumentation requirements are basically the same as those for Category A takeoff. A photo theodolite, grid camera, or other position measuring equipment is needed, along with a ground station to measure wind, OAT, and humidity (if applicable). A two-way communication system between the aircraft and the position measuring equipment is essential. Aircraft instrumentation should include engine and flight parameters, control positions, power lever position, landing gear loads, and a method for synchronizing power cuts between the external light normally used for photo theodolite or camera, and onboard instrumentation. A record of rotor RPM at touchdown is necessary to assure it does not exceed transient limits. Rotor RPM at touchdown may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. A crash recovery team with support of a fire engine is highly desirable.

Page B - 73 AC 29-2C 9/30/99 (B) Establishing the LDP .

( 1 ) Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the LDP becomes largely a function of the runway length required for landing. If landing conditions to be scheduled include considerable runway length (on the order of 1,000 feet) the LDP may be defined at a relatively high speed allowing transition to a takeoff safety speed near V which will Y allow the maximum amount of weight for compliance with the balked landing climb requirements of § 29.77(b)/§ 29.67(a)(1). In this case, the requirements of § 29.67(a)(2) usually become limiting. If the runway length is small, LDP will typically be at a lower speed and may be at a higher altitude to allow balked landing transition within the available distance. Landing weight may need to be reduced to allow landing from the lower speed or higher altitude decision point for shorter landing distances. In this case the requirements of § 29.67(a)(1) may be limiting. The climb performance and climb speeds required by § 29.67(a)(1) and (2) should be established prior to Category A landing tests.

( 2 ) The one engine inoperative landing is similar in many respects to the height-velocity tests described in paragraph AC 29.79. Most of the comments, cautions, and techniques for HV also apply here even though typical flight conditions at LDP are less critical than limiting HV points due to a lower power level and an established rate of descent. The approach is made at a predetermined speed and one engine is made inoperative prior to LDP After the LDP, speed is reduced and the rotorcraft is flared to a conventional one engine inoperative landing. Depending on the landing characteristics and landing profile, the flare may be initiated either prior or subsequent to the 50-foot elevation utilized in determining landing distance. Testing should include an engine failure at the LDP with a 1-second pilot delay to assure safe landing capability for this critical case. A minimum of five acceptable runs for distance should be flown by the FAA/AUTHORITY pilot. These may be averaged with an equal number of acceptable runs by the company pilot.

( 3 ) The balked landing portion of the landing profile is addressed under § 29.77, Balked Landing: Category A. For an explanation of that requirement and a discussion of those test procedures refer to paragraph AC 29.71.

(C) Power . Power should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of engine topping to minimum specification values for the range of atmospheric variables to be approved.

This is frequently done by installing an adjustable device in the throttle linkage with a control in the cockpit so that engine topping can be accurately adjusted for varying ambient conditions. With such a device in the control system it becomes vitally important to check topping power prior to each test sequence.

(D) Aircraft Loading . Aft center of gravity is usually most critical for landing distance determination because visibility constraints limit the degree to which the pilot can flare the rotorcraft for landing. If a weight effect is shown, a minimum of Page B - 74 9/30/99 AC 29-2C two weights should be flown at each test altitude. One weight should be the maximum weight for prevailing conditions and the other should provide a sufficient spread to validate weight accountability.

(E) Extrapolation . Weight cannot be extrapolated above test weight. See discussion under Height-Velocity Testing in paragraph AC 29.79. If no marginal areas are apparent and an acceptable analytical method is used, performance data may be extrapolated ±4,000 feet density altitude from test conditions. (See paragraph AC 29.45.)

(F) Ambient Conditions . Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data must be corrected for existing wind conditions during landing distance tests. Credit for headwind conditions may be given during flight manual data expansion. Paragraph AC 29.45 details allowable wind credit.

(G) All engine out landing .

( 1 ) Several procedures can be utilized to demonstrate compliance with the all engine out landing requirement. As discussed in the explanation portion of this paragraph, § 29.75(b) contains two separate requirements. One is the ability to transition safely into autorotation after failure of the last operative engine. This requirement is discussed in paragraph AC 29.143. The second aspect of this rule requires that a landing from autorotation be possible on a prepared surface. The second requirement is discussed below. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above a prepared landing surface. If a complete company test program has documented an all engine out landing to the GW/ σ (gross weight/density ratio) limit for takeoff and landing at each altitude, verification tests may be initiated at those limiting weight conditions. If not, buildup testing should be initiated at light weight. This test is ordinarily conducted at mid center of gravity. Typically, all altitudes may be approved with two weight limit landings: one at sea level and one near maximum takeoff and landing altitude.

( 2 ) Demonstrated compliance with this requirement is intended to show that an autorotative descent rate can be arrested, and forward speed at touchdown can be controlled to assure a reasonable chance of survivability for the all engine failure condition. The touchdown speed (less than 40 KIAS is recommended) should be consistent with the type design limits including landing gear capability, aircraft visibility, and any other factors affecting repeatability of the maneuver. On Category A rotorcraft, rotor inertia is typically much lower than for single engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Also, due to this relatively low inertia level, considerable collective may be needed to prevent rotor overspeed conditions when the rotorcraft is flared for landing. Also, when testing final maximum weight points, the pilot should anticipate a need for considerable collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions. Some designs incorporate features which may lead to rotorcraft damage in Page B - 75 AC 29-2C 9/30/99 testing this requirement (e.g., droop stop breakage or loss of directional control with skids) if landings are conducted to a full stop with the engines cut off.

( 3 ) The intent of this rule is to demonstrate controlled touchdown conditions and freedom from loss of control or apparent hazard to occupants when landing with all engines failed. In these cases compliance can be demonstrated by leaving throttles in the idle position and assuring no power is delivered to the drive train.

Also, computer analysis may be used in conjunction with simulated in-flight checks to give reasonable assurance that an actual safe touchdown can be accomplished.

Another method may be to make a power recovery after flare effectiveness of the rotorcraft has been determined. Other methods may be considered if they lead to reasonable assurance that descent can be arrested and forward speed controlled to allow safe landing with no injury to occupants when landing on a prepared surface with all engines failed. Regardless of the method(s) used to comply with this requirement, careful planning and analyses are very important due to the potentially hazardous aspects of power off simulation and landing of a Category A rotorcraft totally without power. Considerations for weight and altitude extrapolation are the same as those for HV testing (reference paragraph AC 29.79.) The all-engine-inoperative landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests.

(H) Vertical Landings . The reader should be familiar with the preceding discussion of conventional Category A landing profiles because duplicate information is not repeated here. A typical vertical landing profile is shown in figure AC 29.75-2. This profile is equally applicable to both ground level and pinnacle sites. The profile begins at a stabilized single engine approach condition. It must be possible to make a safe OEI landing or go-around at any point prior to the LDP. At the LDP the aircraft becomes committed to landing. A safe landing must be possible in case of an engine failure at any point before or after the LDP. Testing should include a simulated failure at LDP with a 1-second delay or normal pilot response time, whichever is longer, and subsequent landing within the allowable area. The LDP is typically well above the 25-foot point from which landing distance is measured. The landing distance is the distance from the point at which the lowest portion of the rotorcraft reaches 25 feet above the landing surface to the forward-most point after coming to a stop (including main rotor tip path). The LDP becomes very important for landing on small, elevated heliports. The LDP must be clearly defined and flight manual instructions should carefully explain any pilot procedures. An illustration similar to figure AC 29.75-2 with somewhat more detailed information is most useful. Night OEI landings should be conducted to verify suitable visibility for both internal and external vertical landing cues.

c. Category B Requirements .

(1) Explanation . Section 29.75(c) contains the Category B landing requirements. For rotorcraft that do not meet the Category A powerplant installation requirements of this part, landing tests are conducted with all engines inoperative in an autorotative descent condition. Landing distance is measured from the 50-foot point to the point at which the rotorcraft is completely stopped (approximately 3 knots for water Page B - 76 9/30/99 AC 29-2C landings). The autorotative approach speed is selected by the applicant. The landing maneuver is similar to that referred to during normal training flights as a practice autorotation. As in HV tests, care must be taken to assure no power is delivered to the rotor drive system during these tests. A small amount of power can have a significant effect on landing test results. Multiengine rotorcraft incorporating Category A engine isolation features may conduct landing distance tests with only one engine inoperative using the procedures prescribed above for Category A. For these rotorcraft the one engine inoperative condition typically results in much shorter distances due both to a much lower speed at the 50-foot point and the added power available for flaring and cushioning the landing. Instrumentation requirements are the same as those described under Category A above. Appropriate ambient conditions and allowable extrapolation are discussed under paragraph AC 29.45.

(2) Procedures . Prior to conducting these tests the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. For Category B rotorcraft without engine isolation, the flight profile may be entered in the same manner as a straight-in practice autorotation. It is recommended that for safety reasons idle power be used if a “needle split” (no engine power to the rotor) can be achieved. In some cases, a low engine idle adjustment has been set to assure needle split is attained. In other cases a temporary detent between idle and cutoff was used on the throttle. In a third case the engine was actually shut down on sample runs to verify that the engine power being delivered was not materially influencing landing capability or landing distances. The landing flare may be initiated prior to the 50-foot point. The flare is maintained as long as is reasonable to dissipate speed and build RPM. Rotor RPM must stay within allowable limits. Aft center of gravity is ordinarily critical due to visibility and flare-ability. Following the flare, the rotorcraft is allowed to touchdown in a landing attitude. Rotor RPM at touchdown should be recorded and it must be within allowable structural limits. For wheeled rotorcraft, the brakes are applied to an incipient skid for most efficient stopping. For rotorcraft on skids, the collective should be lowered as soon as characteristics allow in order to place a greater weight on the landing skids.

These procedures would be appropriate flight manual entries to show how landing distances can be realized. For flight manual purposes the landing distance should include the horizontal distance from the point at which the lowest part of the rotorcraft first reaches 50 feet above the landing surface to the point at the foremost part of the rotorcraft (including rotor tip path) after coming to a stop. For Category B rotorcraft with engine isolation, the landing procedures are as described for Category A landing.

When conducting Category B landings utilizing Category A “procedures,” § 29.75(b)(2) can be misleading. No transition capability to balked landing is intended for Category B rotorcraft. Section 29.77, Balked Landing, Category A, applies only to Category A rotorcraft and not to Category B rotorcraft which incorporates Category A “design” features. Five acceptable landing runs should be flown by the FAA/AUTHORITY pilot at each test weight. Results may be averaged with an equal number of company runs. If a weight effect on landing distance is to be shown, a minimum of two weight extremes are normally tested.

Page B - 77 AC 29-2C 9/30/99 Vross Vy Vross 1-<1---- LANDING DISTANCE--• FIGURE AC 29.75-1 CATEGORY A CONVENTIONAL LANDING - CLEAR HELIPORT Page B - 78 9/30/99 AC 29-2C FIGURE AC 29.75-2 CATEGORY A VERTICAL LANDING Page B - 79 AC 29-2C, Chg 4 5/1/2014 AC 29.75A. (AC's 29.77, 29.79, 29.81, & 29.83) §§ 29.75, 29.77, 29.79, 29.81, and 29.83 (Amendment 29-39) LANDING.

(For § 29.77 and § 29.79 prior to Amendment 29-39, see paragraphs AC 29.77 and AC 29.79 respectively.)

a. Explanation .

(1) Amendment 29-39 revised and relocated many of the landing requirements of Part 29. Changes were made to the general landing requirements of § 29.75. New requirements were added for designating a landing decision point (LDP) in § 29.77.

The original § 29.79 was redesignated as a new § 29.87. Category A landing requirements were established in a new § 29.79. Requirements were added to determine landing distances in a new § 29.81. Revised Category B landing requirements were relocated from § 29.75(c) into a new § 29.83. The guidance material from paragraph AC 29.75 does not apply to rotorcraft certified with Amendment 29-39 or later.

(2) These rules incorporate all of the landing performance requirements for transport category rotorcraft. They contain the requirements for landing data, Category A landing, and Category B landing. Parallel takeoff requirements are located in eight separate sections of the rule, §§ 29.51 through 29.63. As such, to ensure that necessary subjects are treated separately, the following discussion will be separated into three parts: (a) a general discussion of basic landing distance requirements, (b) Category A requirements (including vertical landing), and (c) Category B requirements.

(3) All landing performance data are corrected to a smooth, dry, hard, level landing surface condition. As with other flight maneuvers, landings should be accomplished with acceptable flight and ground characteristics using normal pilot skills.

The rule states that Category A and B landing data should be determined at each approved WAT (Weight, Altitude, Temperature) condition. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in paragraph AC 29.45. As in other performance areas, engines should be operated within approved limits. Power considerations are the same as those described under paragraph b(1)(ii)(C).

(4) Unlike fixed-wing aircraft, rotorcraft typically require significantly more landing surface area with an engine inoperative than with all engines operating.

Because of this characteristic, the Category A landing distance requirements are met with at least one engine inoperative to ensure the most conservative landing distance measurement is achieved.

b. Procedures - Category A Requirements .

Page B - 80 5/1/2014 AC 29-2C, Chg 4 (1) Explanation . The Category A certification concept limits landing weight to a value that will allow the rotorcraft, following an engine failure at the landing decision point (LDP), to land within the available area or to execute a balked landing descending no lower than 15 feet (or higher depending on rotorcraft geometry and performance characteristics) above the landing surface. For elevated heliports the rotorcraft may descend below the landing surface, but all parts of the rotorcraft must clear the heliport and other obstacles by not less than 15 feet. These minimum heights should be demonstrated with variations in piloting techniques and with pilot recognition and reaction times for engine failures occurring before and after the LDP. See figure AC 29.75A-1. For additional information addressing the OEI landing case at night, refer to AC 29.59A.b.(13).

(i) LDP . The Category A landing profile begins with an assumed engine failure at or prior to the LDP. The LDP is typically defined in terms of airspeed, rate of descent, and altitude above the landing surface. The approach path angle can be defined by LDP airspeed and rate of descent values. Definition of the LDP should include an approach angle because both the landing distance and the missed approach path are significantly influenced by landing approach angle. At any point in the single engine approach path down to and including the LDP, the pilot may elect to land or to execute a balked landing and he is assured both an adequate surface area for OEI landing and adequate climb capability for an OEI balked landing. Said another way, if an engine failure is recognized at any point down to and including the LDP, the pilot may safely elect to land or to “go-around” by executing a balked landing. The LDP should be defined to permit acceleration to V clearing the landing surface by a TOSS minimum of 15 feet. The LDP represents a “commit” point for landing. Prior to the LDP in the one engine inoperative approach, the pilot has a choice, he may either land or fly away. After passing the LDP, he no longer has sufficient energy to assure transition to a balked landing condition without contacting the landing surface. If an engine failure is recognized after LDP in a normal (all engine) landing, the pilot is committed to land.

The LDP and landing approach path should be defined such that critical areas of the height-velocity diagram are avoided. A typical LDP for conventional Category A profiles is 100 feet above the landing surface. LDP should be specified in terms of both actual height above the landing surface and indicated barometric altitude. Speed at the LDP should be specified in terms of indicated airspeed. The applicant may elect to develop an alternate all-engines-operating (AEO) approach procedure which meets the performance after engine failure requirements to execute a go-around before LDP or land after LDP but which could not be executed with OEI following an en route engine failure. If such alternate AEO procedures are provided, the Flight Manual should include the appropriate limitations prohibiting use of the AEO procedures after an en route engine failure. For such alternate AEO approach procedures it should be possible to execute a go-around and use the OEI approach procedure if the landing weight is consistent with such approach (the Flight Manual should indicate this OEI approach procedure and corresponding landing weight).

Page B - 81 AC 29-2C 9/30/99 (ii) Landing distance . Approach and landing path requirements are stated in §§ 29.79(a)(2) and 29.83(a)(2). For Category A rotorcraft, the approach path should allow smooth transition for one-engine inoperative landing and for balked landing maneuvers. For all rotorcraft, the approach and landing paths should allow adequate clearance from potentially hazardous HV combinations. Landing distance is measured from a point 50 feet above the landing surface to a stop. For RFM presentation, the distance is from the aft most portion of the rotorcraft at the point at which the lowest part of the rotorcraft first reaches 50 feet to the foremost point of the rotorcraft (including rotor tip path) after coming to a stop.

(iii) All Engine Out Landing . § 29.79(b) contains the Category A certification requirement for an all engine inoperative landing. The rule states that it should be possible to make a safe landing on a prepared surface after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously.

See paragraph AC 29.143a(2)(iii)(A) for the Category A sequential engine failure criteria. The conditions for last engine failure are maximum continuous power or 30-minute power if that rating is approved, “wings” level flight, and sudden engine failure with a pilot delay of 1 second or normal pilot recognition time, whichever is greater.

Complete power failure has occurred in twin engine Category A rotorcraft. This requirement ensures that in the event of cockpit mismanagement, fuel exhaustion, improper maintenance, fuel contamination, or unforeseen mechanical failures, a safe autorotation entry can be made and a safe power-off landing can be effected. Two separate aspects of this rule are normally evaluated at different times during the test program. The last engine failure is normally evaluated during cruise or VNE engine failure testing where instrumentation and critical loading have been established for those test conditions. See discussion under paragraph AC 29.143. The all engine out landing is ordinarily conducted in conjunction with an HV or Category A landing distance phase where ground instrumentation and safety equipment are available. The rotorcraft should be capable of conducting the all engine out landing at the takeoff and landing WAT limiting conditions up to the maximum altitude approved for takeoff and landing.

(2) Procedures .

(i) Instrumentation/Equipment . Instrumentation requirements are basically the same as those for Category A takeoff. A photo theodolite, grid camera, GPS, or other position measuring equipment is needed, along with a ground station to measure wind, OAT, and humidity (if applicable). A two-way communication system between the aircraft and the position measuring equipment is essential. Aircraft instrumentation should include engine and flight parameters, control positions, power lever position, landing gear loads, and a method for synchronizing aircraft position when the power is cut with onboard instrumentation. A record of rotor RPM at touchdown is necessary to ensure it does not exceed transient limits. Rotor RPM at touchdown may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. A crash recovery team with support of a fire engine is highly desirable.

(ii) Establishing the LDP .

Page B - 82 4/25/06 AC 29-2C, Chg 2 (A) Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the LDP becomes largely a function of the runway length required for landing. If landing conditions to be scheduled include considerable runway length (on the order of 1,000 feet), the LDP may be defined at a relatively high speed allowing transition to a takeoff safety speed near V which will Y allow the maximum amount of weight for compliance with the balked landing climb requirements of § 29.85(b)/§ 29.67(a)(1). In this case, the requirements of § 29.67(a)(2) usually become limiting. If the runway length is small, LDP will typically be at a lower speed and may be at a higher altitude to allow balked landing transition within the available distance. Landing weight may need to be reduced to allow landing from the lower speed or higher altitude decision point for shorter landing distances. In this case the requirements of § 29.67(a)(1) may be limiting. The climb performance and climb speeds required by § 29.67(a)(1) and (2) should be established prior to Category A landing tests.

(B) The one-engine-inoperative landing is similar in many respects to the height-velocity tests described in paragraph AC 29.79. Most of the comments, cautions, and techniques for HV also apply here even though typical flight conditions at LDP are less critical than limiting HV points due to a lower power level and an established rate of descent. The approach is made at a predetermined speed and one engine is made inoperative prior to LDP. After the LDP, speed is reduced and the rotorcraft is flared to a conventional one engine inoperative landing. Depending on the landing characteristics and landing profile, the flare may be initiated either prior or subsequent to the 50 foot elevation utilized in determining landing distance. Testing should include an engine failure such that recognition is at the LDP with a 1-second pilot delay to ensure safe landing capability for this critical case. A sufficient number of acceptable runs should be accomplished to provide confidence in the results. Typically ten acceptable runs are adequate.

(C) The balked landing portion of the landing profile is addressed under § 29.85, Balked Landing: Category A. For an explanation of that requirement and a discussion of those test procedures, refer to paragraph AC 29.77.

(iii) Power . Power used for demonstrating performance should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping to minimum specification values including consideration of temperature effects on engine power. If the management of engine power at topping is beyond normal pilot capability, the validity of the Cat A procedure must also be established with representative in-service characteristics. The method used for simulating engine failure must be representative of the power decay characteristics that will occur during a real, sudden engine failure and acceleration of the remaining engine(s). In order to cushion the OEI landing, it is acceptable for the engine and transmission transient range in order to droop the rotor provided performance credit is not taken for this additional power above the maximum permitted rating and it can be shown that the engine(s) will remain within these limits in all Page B - 83 AC 29-2C, Chg 2 4/25/06 conditions requested by the applicant. Any excursion beyond established transient limits in this flight phase should be substantiated to the extent that it does not constitute an immediate hazard to the rotorcraft.

(iv) Aircraft Loading . Aft center of gravity is usually most critical for landing distance determination because visibility constraints limit the degree to which the pilot can flare the rotorcraft for landing. If a weight effect is shown, a minimum of two weights should be flown at each test altitude. One weight should be the maximum weight for prevailing conditions and the other should provide a sufficient spread to validate weight accountability.

(v) Extrapolation . Landing data may be extrapolated along an established W/ σ line to the maximum gross weight of the rotorcraft. However, extrapolation will not be considered valid if landing gear loads are marginally acceptable at actual landing weights below the W/ σ limit. If no marginal areas are apparent and an acceptable analytical method is used, performance data may be extrapolated up to 4,000 feet density altitude from test conditions. (See paragraph AC 29.45.)

(vi) Ambient Conditions . Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data should be corrected for existing wind conditions during landing distance tests. Credit for headwind conditions may be given during flight manual data expansion.

Paragraph AC 29.1587 details allowable wind credit.

(vii) All Engine Out Landing .

(A) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. As discussed in the explanation portion of this paragraph, §§ 29.79 and 29.83 each require that a landing from autorotation be possible. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface. All-engine-out landing tests should be initiated at light weight with a gradual buildup to the limiting weight conditions. If a complete company test program has documented all-engine-out landings to the GW/ σ limit, the buildup conditions during verification test may be decreased. If not, buildup testing should be initiated at light weight. This test is ordinarily conducted at mid center of gravity. Typically, all altitudes may be approved with two weight limit landings - one at sea level and one near maximum takeoff and landing altitude.

(B) Demonstrated compliance with this requirement is intended to show that an autorotative descent rate can be arrested, and forward speed at touchdown can be controlled to a reasonable value (less than 40 KTAS is recommended) to ensure a reasonable chance of survivability for the all engine failure condition. On multiengine rotorcraft, rotor inertia is typically lower than for single-engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Due to this relatively low inertia level, considerable collective may be needed to prevent rotor overspeed conditions when the Page B - 84 9/30/99 AC 29-2C rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions. Some designs incorporate features which may lead to rotorcraft damage in testing this requirement (e.g., droop stop breakage or loss of directional control with skids) if landings are conducted to a full stop with the engines cut off.

(C) The intent of this rule is to demonstrate controlled touchdown conditions and freedom from loss of control or apparent hazard to occupants when landing with all engines failed. In these cases compliance can be demonstrated by leaving throttles in the idle position and ensuring no power is delivered to the drive train.

Also, computer analysis may be used in conjunction with simulated in-flight checks to give reasonable assurance that an actual safe touchdown can be accomplished.

Another method may be to make a power recovery after flare effectiveness of the rotorcraft has been determined. Other methods may be considered if they lead to reasonable assurance that descent can be arrested and forward speed controlled to allow safe landing with no injury to occupants when landing on a prepared surface with all engines failed. Regardless of the method(s) used to comply with this requirement, careful planning and analyses are very important due to the potentially hazardous aspects of power off simulation and landing of a multiengine rotorcraft totally without power. Considerations for weight and altitude extrapolation are the same as those for HV testing (see paragraph AC 29.79). The all-engine-inoperative landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests.

(D) Prior to conducting these tests, the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. The flight profile may be entered in the same manner as a straight-in practice autorotation. It is recommended that for safety reasons idle power be used if a “needle split” (no engine power to the rotor) can be achieved. In some cases, a low engine idle adjustment has been set to assure needle split is attained. In other cases, a temporary detent between idle and cutoff was used on the throttle. In a third case, the engine was actually shut down on sample runs to verify that the engine power being delivered was not materially influencing landing capability or landing distances. The flare is maintained as long as is reasonable to dissipate speed and build RPM. Rotor RPM should stay with allowable limits. Aft center of gravity is ordinarily critical due to visibility and flare-ability.

Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it should be within allowable structural limits.

(viii) Vertical Landings . The reader should be familiar with the preceding discussion of conventional Category A, landing profiles because duplicate information is not repeated here. A typical vertical landing profile is shown in figure AC 29.75A-2.

This profile is equally applicable to both ground level and elevated heliport sites. The profile begins at a stabilized single engine approach condition. It should be possible to make a safe OEI landing or go-around at any point prior to the LDP unless alternate Page B - 85 AC 29-2C 9/30/99 AEO approach procedures are presented in the Flight Manual according to paragraph AC 29.75b(1)(i)(A). It is possible to have two landing techniques: an “offset” one, which schedules drop down for elevated heliports (but still ensure 15 feet radial deck edge clearance), and a “straight in” approach which utilizes the ground level heliport criteria. These techniques should be stipulated as such in the Flight Manual. At the LDP the aircraft becomes committed to landing. A safe landing should be possible in case of an engine failure at any point before or after the LDP. Testing should include a simulated failure at LDP with a 1-second delay or normal pilot response time, whichever is longer, and subsequent landing within the allowable area. The landing distance is the distance from the point at which the lowest portion of the rotorcraft reaches 50 feet above the landing surface to the forward-most point after coming to a stop (including main rotor tip path). The LDP becomes very important for landing on small, elevated heliports. The LDP should be clearly defined and Flight Manual instructions should carefully explain any pilot procedures. An illustration similar to figure AC 29.75A-2 with somewhat more detailed information is most useful. Night OEI landings should be conducted to verify suitable visibility for both internal and external vertical landing cues. The minimum elevated heliport size demonstrated for the OEI approach procedure and for alternate AEO approach procedures (when provided) should also be provided in the Flight Manual.

c. Category B Requirements .

(1) Explanation . Section 29.83 contains the Category B landing requirements.

Landing distance is measured from the 50-foot point to the point at which the rotorcraft is completely stopped (approximately 3 knots for water landings). The approach speed is selected by the applicant. Appropriate ambient conditions and allowable extrapolation are discussed under paragraph AC 29.45.

(2) Procedures .

(i) Landing Distance . Aft center of gravity is ordinarily critical due to field-of-view and flare ability. For wheeled rotorcraft, the brakes are applied to an incipient skid for most efficient stopping. For rotorcraft on skids, the collective should be lowered as soon as characteristics allow in order to place a greater weight on the landing skids. These procedures would be appropriate flight manual entries to show how landing distances can be realized. For flight manual purposes, the landing distance should include the horizontal distance from the point at which the lowest part of the rotorcraft first reaches 50 feet above the landing surface to the point at the foremost part of the rotorcraft (including rotor tip path) after coming to a stop. Multiengine rotorcraft incorporating Category A engine isolation features may elect to show compliance with § 29.79 and § 29.81. A sufficient number of acceptable runs should be accomplished to provide confidence in the results. Typically ten acceptable runs are adequate. If a weight effect on landing distance is to be shown, a minimum of two weight extremes are normally tested.

(ii) All-Engine-Out Landing .

Page B - 86 9/30/99 AC 29-2C (A) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. Section 29.83(c) requires that a landing from autorotation be possible. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface.

All-engine-out landing tests should be initiated at light weight with a gradual buildup to the limiting weight conditions. If a complete company test program has documented all-engine-out landings to the GW/ σ limit, the buildup conditions during verification test may be decreased. This test is ordinarily conducted at mid center of gravity. Typically, all altitudes may be approved with two weight limit landings - one at sea level and one near maximum takeoff and landing altitude.

(B) Demonstrated compliance with this requirement is intended to show that an autorotative descent rate can be arrested, and forward speed at touchdown can be controlled to a reasonable value (less than 40 KTAS is recommended) to ensure a reasonable chance of survivability for the all engine failure condition. On multiengine rotorcraft, rotor inertia is typically lower than for single-engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Due to low rotor inertia, considerable collective may be needed to prevent rotor overspeed conditions when the rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions.

(C) The intent of this rule is to demonstrate controlled touchdown conditions and freedom from loss of control or apparent hazard to occupants when landing with all engines failed. In these cases compliance can be demonstrated by leaving throttles in the idle position and ensuring no power is delivered to the drive train. Also, computer analysis may be used in conjunction with simulated in-flight checks to give reasonable assurance that an actual safe touchdown can be accomplished. Another method may be to make a power recovery after flare effectiveness of the rotorcraft has been determined. Other methods may be considered if they lead to reasonable assurance that descent can be arrested and forward speed controlled to allow safe landing with no injury to occupants when landing on a prepared surface with all engines failed. Regardless of the method(s) used to comply with this requirement, careful planning and analyses are very important due to the potentially hazardous aspects of power off simulation and landing of a multiengine rotorcraft totally without power. Considerations for weight and altitude extrapolation are the same as those for HV testing (see paragraph AC 29.79). The all-engine-inoperative landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests.

(D) Prior to conducting these tests, the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. The flight profile may be entered in the same manner as a straight-in practice autorotation. It is recommended that for safety reasons idle power be used if a “needle split” (no engine power to the rotor) can be achieved. In some cases, a low engine idle adjustment has been set to Page B - 87 AC 29-2C 9/30/99 assure needle split is attained. In other cases, a temporary detent between idle and cutoff was used on the throttle. In a third case, the engine was actually shut down on sample runs to verify that the engine power being delivered as not materially influencing landing capability or landing distances. The flare is maintained as long as is reasonable to dissipate speed and build RPM. Rotor RPM should stay with allowable limits. Aft center of gravity is ordinarily critical due to visibility and flareability. Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it should be within allowable structural limits.

Page B - 88 9/30/99 AC 29-2C LANDING DISTANCE _____.

FIGURE AC 29.75A-1 CATEGORY A CONVENTIONAL LANDING- CLEAR HELIPORT Page B - 89 AC 29-2C 9/30/99 FIGURE AC 29.75A-2 CATEGORY A VERTICAL LANDING Page B - 90 9/30/99 AC 29-2C AC 29.77. § 29.77 (Amendment 29-24) BALKED LANDING: CATEGORY A (For § 29.77 after Amendment 38, see paragraph AC 29.75A) a. Explanation . This rule has two distinct portions.

(1) Section 29.77(a) states that the rotorcraft must be capable of transitioning smoothly from each approved Category A approach condition to a missed approach with one engine inoperative (OEI). Although not specifically stated in the rule, this requirement must be met for any point prior to the landing decision point (LDP).

(2) Section 29.77(b) requires that the LDP be defined so that it will permit transition to a safe climb condition in the event a balked landing is necessary. (See figure AC 29.75-1.) The safe climb conditions are defined in § 29.67(a)(1) and (2). This suggests establishing a clearly defined balked landing profile similar to the Category A takeoff profile established under § 29.59. The balked landing profile must insure compliance with the climb performance requirements of §§ 29.67(a)(1) and 29.67(a)(2).

b. Procedures .

(1) Instrumentation . Instrumentation requirements are similar to those for Category A takeoff. A ground station with positioning capability is needed along with on-board instrumentation of engine and flight parameters.

(2) Balked Landing Profiles . One engine inoperative balked landing profiles during approach must be conducted at conditions up to and including the LDP. The LDP should be designated so that the balked landing profile may be completed with the rotorcraft descending no lower than 35 feet above the landing surface. The distance from the LDP to the point in the balked landing profile at which a minimum of 35 feet above the landing surface is attained at V in a climbing posture should be recorded.

TOSS This distance should be compared against the landing distance determined under § 29.81 to assure the balked landing maneuver can be completed within the designated landing area. This is especially important for future steep angle, low speed Category A approaches to heliports.

(3) Handling Qualities . Handling qualities features in the balked landing transition should be carefully evaluated. Characteristics such as excessive nose down pitching with power application or excessive engine lag should not be approved.

(4) Climb Performance . In accordance with this rule, the climb requirements of § 29.67(a)(1) and (2) must also be met in the event a balked landing is made. See paragraphs AC 29.65 and AC 29.67.

Page B - 91 AC 29-2C, Chg 4 5/1/2014 AC 29.79. § 29.79 (Amendment 29-21) LIMITING HEIGHT-SPEED ENVELOPE.

(For § 29.79 after Amendment 29-38, see section 29.75A of this AC.)

a. Explanation .

(1) The height-speed envelope is normally referred to as the height-velocity (HV) diagram. It defines an envelope of airspeed and height above the ground from which a safe power-off or one engine inoperative (OEI) landing cannot be made. The diagram normally consists of three portions: (a) the level flight (cruise) portion, (b) the takeoff portion, and (c) the high speed portion (see Figure AC 29.79-1). The high speed portion is omitted on occasions when it can be shown that the rotorcraft can suffer an engine failure at low altitude and high speed (up to V ) and make a successful landing, or climb H out on the remaining engine(s).

(2) Engine power considerations are similar to those in previous takeoff and landing requirements (see sections 29.53, 29.63, and 29.75 of this AC).

(3) The prohibited sections of the HV diagram are separated by the takeoff corridor. This corridor should be wide enough to consistently permit a takeoff flight path clear of the HV diagram using normal pilot skill. The takeoff corridor should always permit a minimum of  5 knots clearance from critical portions of the diagram.

(4) The knee of the curve separates the takeoff portion from the cruise portion and is defined as the highest speed point on the low speed portion of the HV envelope.

Altitudes above this point are considered cruise, or “fly-in,” points and these test points require a minimum time delay of 1-second between throttle chop and control actuation (reference § 29.143(d)). Altitudes below the knee represent takeoff profile points. For test points in the takeoff portion, use takeoff power (or a lower power selected by the applicant as an operating limitation) and normal pilot reaction time.

(5) Since the HV diagram may represent the limiting capabilities of the rotorcraft, each test point should be approached with caution. The manufacturer’s buildup program should be reviewed to determine the amount of conservatism in the HV diagram (if any). It should be remembered that the operational pilot will be operating at or near the HV diagram without the benefit of a buildup program. Buildup testing is necessary, and it is most important to vary only one parameter at a time to prevent surprises. Light weight testing is ordinarily conducted first. High and low hover points are approached from above and below respectively. Portions near the knee are initially evaluated at high speed with subsequent backing down of the speed. In most rotorcraft the effective flare airspeed is critical. At airspeeds slightly below this value, the ability to arrest and control descent rates through use of an aft cyclic flare may be greatly diminished. Extreme care should be exercised when “backing down” to lower speeds.

(6) In addition to the on-board and ground instrumentation, a motion picture camera or other position measuring equipment should cover each run.

Page B - 92 5/1/2014 AC 29-2C, Chg 4 (7) For FAA/AUTHORITY tests, the minimum required crew and minimum instrument panel display should be used. Ground safety equipment should be provided.

(8) This test is the least predictable of all the performance items. Therefore, the expansion and extrapolation of test data are questionable. Weight may not be extrapolated to higher values. In order to extrapolate HV data to higher altitudes, any analytical method must have FAA/AUTHORITY approval. In lieu of pure analytical methods, simulations have been used successfully, especially for multiengine rotorcraft.

In either case, the maximum allowable extrapolation should be limited to 2,000 feet density altitude (H ). HV test weights should be consistent with the takeoff and landing D weight, altitude, temperature (WAT) limit curve which will be placed in the rotorcraft flight manual (RFM). For a given diagram, typical weight reductions that are necessary as altitude is increased can be conservatively estimated by maintaining a constant gross weight divided by density ratio, GW/  (see Figure AC 29.79-2, Part A). If weight is not varied, an enlarged HV diagram is required for safe power-off landing as density altitude is increased (see Figure AC 29.79-2, Part B). Another method of presentation is to show varying weights at a constant density altitude (see Figure AC 29.79-2, Part C.)

(9) The FAA accepts, as a method of extrapolation, a weight penalty of 3% for each 1000 feet above the permitted 2000 feet extrapolation. This weight penalty has been applied to extrapolations along a constant gross weight divided by density ratio, GW/  .

(10) Vertical takeoff and landing (VTOL) testing normally does not require separate HV testing. The takeoff and landing tests take on the combined characteristics of takeoff, landing, and HV tests.

(11) Rotorcraft certificated prior to Amendment 29-21 were required to have the resulting HV diagram as an operating limitation. This limitation restricted opportunities when operating large rotorcraft in various utility applications. Subsequently, Amendment 29-21 allows, under certain conditions, the HV diagram to be placed in the Flight Manual Performance Information Section instead of the Limitations Section.

Specifically, the rotorcraft must be: (1) certificated for a maximum gross weight of 20,000 pounds or less; (2) configured with nine passenger seats or less; and (3) certificated in Category B. Testing must be completed with the aircraft at the maximum gross weight at sea level. For altitudes above sea level, the test aircraft must be at a weight no less than the highest weight the rotorcraft can hover out of ground effect (OGE). Rotorcraft certificated prior to Amendment 29-21 can update their certification basis to take advantage of this provision.

b. Procedures .

(1) Instrumentation .

Page B - 93 AC 29-2C, Chg 4 5/1/2014 (i) Ground Station . The ground station must have equipment and instrumentation to determine wind direction and velocity, outside air temperature, and (if the test rotorcraft has reciprocating engines), humidity. Since the tests must be conducted in winds of 2 knots or less, a smoke generator is highly recommended to show both flightcrew and ground crew personnel the wind direction and velocity at any given time. Additionally, the location of the ground station should be such that it is free of rotor downwash at all times. Motion picture, phototheodolite, and radio equipment will be necessary to properly conduct the test program. The use of telemetry equipment is desirable if the location of the test site and the magnitude of the test program make it practical.

(ii) Airborne Equipment (Test Rotorcraft) . Necessary installed test equipment may include photo panels or recorders for recording engine parameters, control positions, landing gear loads, landing gear deflections, airspeed, altitude, and other variables. An external light attached to the rotorcraft (or any other means of identifying the engine failure point to the ground camera or phototheodolite) is needed to identify the exact time of engine failure and may also be used to synchronize the ground recorder with the airborne recorded data.

(2) Analytical Prediction . The HV diagram can be estimated by analytical means and this is recommended prior to test. HV, however, is the least predictable of all rotorcraft performance and because of this, the expansion and extrapolation of test data must be done with great care. Test weight may not be extrapolated. All test points should be approached conservatively with some speed or altitude margin. If the manufacturer has conducted a comprehensive HV flight test program to validate his analytical predictions, much preliminary testing can be eliminated. In any case, the maximum allowable extrapolation from flight test conditions is 2,000 feet density altitude and an approved analytical or simulation method must be utilized for extrapolation.

(3) Power .

(i) The appropriate power level before engine failure for the low and high hover points is simply the power required to hover at the prevailing hover conditions.

The appropriate power condition prior to failure of the engine for points below the knee is takeoff power or a lower value if approved as an operating limit. For cruise or “fly-in” points above the knee, the appropriate condition is power required for level flight. Rotor speed at execution of the engine failure should be the minimum speed appropriate to the flight condition.

(ii) The applicable power failure conditions are listed in § 29.79(b). Power should be completely cut for category B rotorcraft. For multiengine rotorcraft that comply with the category A engine isolation design requirements, the HV envelope may be determined with OEI and the desired topping power (for the remaining engine(s)) should be set prior to the test. This power value will need adjustment as ambient conditions change. The power can be takeoff power (TOP), 2 1/2-minute power, or Page B - 94 5/1/2014 AC 29-2C, Chg 4 some calculated lower power for simulating hot day or higher density altitude conditions.

Power is verified and recorded by the pilot by “topping” the engine(s) prior to engine failure tests. Care must be taken to assure that this power value is no more than that which would be delivered by a minimum specification engine under the ambient conditions to be approved.

(4) Test Loadings . Weight extrapolation is not permitted for HV. Therefore, the test weight must be closely controlled. Ballast or fuel should be added frequently to maintain the weight within -1 to +5 percent when testing final points. Ordinarily tests are conducted at a mid-center of gravity unless a particular loading is expected to be particularly critical.

(5) Landing Gear Loads .

(i) Instrumented landing gear can be a great help in evaluating test results.

This information can be telemetered to a ground station or otherwise recorded and displayed for direct reference following each landing.

(ii) Any landing which results in permanent deformation of aircraft structure or landing gear beyond allowable maintenance limits is considered an unsatisfactory test point.

(6) Piloting Considerations . In verifying the HV diagram, the minimum required instrument panel display and minimum crew should be used in order not to mislead the operational pilot who has no test equipment available and may have no copilot to assist.

Three distinctly different flight profiles are utilized in developing the diagram.

(i) High Hover . A stabilized OGE hover condition prior to power failure is essential. A minimum 1-second time delay between power failure and initial control actuation is utilized. Following the time delay, the primary concern is to quickly lower collective and to gain sufficient airspeed to allow an effective flare approaching touchdown. While the immediate development of airspeed is necessary, the dive angle must be reasonable and must be representative of that expected in service. While initial aircraft attitude will vary between models and with changing conditions, 10°-20° has been previously applied as a maximum allowable nose down pitch attitude. Use of greater attitudes could result in a diagram which is difficult to achieve and unrealistic for operations in service. Initial testing should start relatively high with gradual lowering of height to the final high hover altitude. A stabilized OGE hover condition prior to power failure is essential. If a stabilized high hover condition cannot be achieved prior to the engine cut, then this point should be tested from a minimum level flight speed. This will result in an open-ended HV diagram. A smoke source or balloon on a long cord is highly desirable since the wind can vary significantly from surface observations to typical high hover altitudes. Vertical speed must be very near zero at the throttle chop.

Any climb or sink rate can have a significant influence on the success of the test point.

Use of a radar altimeter with a cross check to barometric altitude is essential.

Page B - 95 AC 29-2C, Chg 4 5/1/2014 (ii) Low Hover . From the low hover position there is no flare capability and little time for collective reaction. No time delay is applied other than normal pilot reaction. For typical designs the collective may not be lowered after power failure.

Lowering of the collective is not permitted because it is not a pilot action which could be expected if an engine failed without notice during a hovering condition in service. Initial lowering of collective immediately after power failure can result in very high, unconservative low hover altitudes that are unrealistic for operational conditions. If, however, a design is such that a 1-second pilot delay after power failure could be achieved without any appreciable descent, a slight lowering of collective could be allowed.

(iii) Takeoff Corridor . Normal pilot reaction is applied when the engine is made inoperative. At low speeds collective may be lowered quickly to retain RPM and minimize the time between power failure and ground contact. If airspeed is sufficient for an effective flare, the aircraft is flared to reduce airspeed, retain rotor RPM, and control vertical speed prior to touchdown. Considerable surface area may be needed for a sliding or rolling stop.

(iv) Additional Considerations . The “in-between” points utilize similar techniques. The cruise or “fly-in” points are similar to the high hover point although the steep initial pitch attitudes are not needed as altitude is decreased and airspeed is increased along the curve. The low speed points along the takeoff corridor are similar to the low hover point except that the collective may be quickly lowered and some flare capability may be used as the “knee” is approached. The pilot should be proficient in all normal autorotation landings before conducting HV tests in a single-engine rotorcraft.

(7) Ground Support . Motion picture or theodolite coverage and ground safety equipment are necessary. Communication capability among these elements should be provided. Use of a phototheodolite to compare height and speed with cockpit observations is very desirable.

(8) Verifying the HV Diagram .

(i) A sufficient number of test points must be flown to verify the diagram. The key areas are the knee, high altitude hover, low altitude hover, and high speed touchdown. Test points with excessive gear loads, above average skill requirements, winds above permissible levels, rotor droop below approved minimum transient RPM, damage to the rotorcraft, excessive power, incorrect time delay, etc., cannot be accepted.

(ii) After the HV diagram is defined, it should be ascertained that the corridor permits takeoffs within ±5 knots of the recommended takeoff profile.

(9) Flight Manual . The flight manual should list any procedures which may apply to specific points (e.g., high speed points) and test conditions, such as runway surface, wave height for amphibious tests, marginal areas of controllability or landing gear Page B - 96 5/1/2014 AC 29-2C, Chg 4 response, etc. The HV curve should be presented in the RFM using actual altitude above ground level and indicated airspeed.

(10) Night Evaluation . If a rotorcraft is to be certified for night operation, a night evaluation is required. Engine failures should be conducted along the recommended takeoff path. Landings should also be qualitatively evaluated with an engine failed.

Engine failures at critical HV conditions are not required. The intent is to show adequate visibility using aircraft or runway lights without requiring a duplication of the daytime HV test program. See related discussion under AC 29-2C, section 29.63.

(11) Water Landings . For amphibious float equipped rotorcraft, day and night water landings should be conducted under critical loading conditions with an engine failed. Engine failures should be conducted along the recommended takeoff path.

Engine failures at critical HV conditions are not required. The intent is to show similarity to test results over land without requiring a duplication of the HV test program.

Page B - 97 AC 29-2C, Chg 4 5/1/2014 HEIGHT (H) ✓ -KNEE ~-TAKEOFF CORRIDOR HIGH SPEED SPEED (V) FIGURE AC 29.79-1 HEIGHT-VELOCITY (HV) DIAGRAM Page B - 98 5/1/2014 AC 29-2C, Chg 4 Approx.

Constant GW/rr Height Weight Velocity CONSTANT HV DIAGRAM, VARIABLE WEIGHT Increasing Density Altitude Height Velocity L_B __J CONSTANT WEIGHT Increasing Gross Weight Height Velocity L_c_J CON ST ANT DENSITY ALTITUDE FIGURE AC 29.79-2 ALTITUDENVEIGHT ACCOUNTABILITY Page B - 99 AC 29-2C 9/30/99 AC 29.85. § 29.85 (Amendment 29-39) BALKED LANDING: CATEGORY A (For Balked Landing prior to Amendment 39, see § 29.77 and paragraph AC 29.77.)

a. Explanation . Amendment 29-39 revised and relocated the original § 29.77 as a new § 29.85. The guidance material of paragraph AC 29.77 does not apply to rotorcraft certified with Amendment 29-39 or later. This rule has two distinct portions.

(1) Section 29.85(a) states that the rotorcraft must be capable of transitioning smoothly from each approved Category A approach condition to a missed approach with one engine inoperative (OEI). Although not specifically stated in the rule, this requirement must be met for any point prior to the landing decision point (LDP).

(2) Section 29.85(b) requires that the LDP be defined so that it will permit transition to a safe climb condition in the event a balked landing is necessary. (See figure AC 29.75A-1.) The safe climb conditions are defined in § 29.67(a)(1) and (2). A clearly defined balked landing profile similar to the Category A takeoff profile should be established. The balked landing profile must insure compliance with the climb performance requirements of §§ 29.67(a)(1) and 29.67(a)(2).

b. Procedures .

(1) Instrumentation . Instrumentation requirements are similar to those for Category A takeoff. A ground station with positioning capability is needed along with on-board instrumentation of engine and flight parameters.

(2) Balked Landing Profiles . One engine inoperative balked landing profiles during approach must be conducted at conditions down to and including the LDP. The LDP should be designated so that the balked landing profile may be completed with the rotorcraft clearing the landing surface by a minimum of 15 feet. Fifteen feet should be considered the absolute minimum clearance allowed with greater clearances required for some rotorcraft dependent on rotorcraft geometry and performance characteristics.

For elevated or ground level heliports, with significantly lower LDP heights than 100 feet, the minimum clearance is 15 feet vertically and radially. These minimum heights would need to be demonstrated with variations in piloting techniques and with pilot recognition and reaction times for engine failures occurring before/after LDP. The distance from the LDP to the point in the balked landing profile at which a minimum of 35 feet above the landing surface is attained at V in a climbing posture should be TOSS recorded. This distance should be compared against the landing distance determined under § 29.81 to assure the balked landing maneuver can be completed within the designated landing area. This is especially important for future steep angle, low speed Category A approaches to heliports.

(3) Handling Qualities . Handling qualities features in the balked landing transition should be carefully evaluated. Characteristics such as excessive nose down pitching with power application or excessive engine lag should not be approved.

Page B - 100 9/30/99 AC 29-2C (4) Climb Performance . In accordance with this rule, the climb requirements of § 29.67(a)(1) and (2) must also be met in the event a balked landing is made. See paragraphs AC 29.65 and AC 29.67.

AC 29.87. § 29.87 (Amendment 29-39) LIMITING HEIGHT-SPEED ENVELOPE.

(For Limiting Height-Speed Envelope prior to Amendment 39, see § 29.79 and paragraph AC 29.79.)

a. Explanation . Amendment 39 redesignated § 29.79 as § 29.87.

b. Procedures . The guidance material presented in paragraph AC 29.79 continues to apply.

Page B - 101 AC 29-2C, Chg 4 5/1/2014 SUBPART B - FLIGHT FLIGHT CHARACTERISTICS AC 29.141. § 29.141 (Amendment 29-24) FLIGHT CHARACTERISTICS GENERAL.

a. Explanation .

(1) This regulation prescribes the general flight characteristics required for certification of a transport category rotorcraft. Specifically, it states that the rotorcraft must comply with the flight characteristics requirements at all approved operating altitudes, gross weights, center of gravity locations, airspeeds, power, and rotor speed conditions for which certification is requested. The reference to “altitude” in § 29.141(a)(1) refers to “density altitude.” Density altitude is, of course, a function of pressure altitude and ambient temperature, hence the need to account for ambient temperature effects. Additional flight characteristics required for instrument flight are contained in Appendix B of this AC.

(2) Generally the aircraft structural (load level) survey accounts for takeoff power values at speeds up to and including V . At speeds above V , maximum continuous Y Y power is assumed. Stress to rotating components usually increases with airspeed and power. If the takeoff power rating exceeds the maximum continuous power rating, and the structural survey has been conducted under the assumption that takeoff power is not used at speeds above V , the rotorcraft flight manual (RFM) must limit takeoff power Y to speeds of V and below. If takeoff power is structurally substantiated throughout the Y flight envelope, and appropriate portions of the controllability, maneuverability, and trim requirements of §§ 29.141 through 29.161 are met at takeoff power levels, no flight manual entry is needed. Obviously, if transmission limits for maximum continuous (MC) and takeoff power are coincident, no special action is needed.

(3) During the flight characteristics testing, the controls must be rigged in accordance with the approved rigging instructions and tolerances. The control system rigging must be known prior to testing. In addition to the normal rigging procedures, any programmed control surfaces which may be operated by dynamic pressure, electronics, etc., must also be calibrated. During the flight test program, it is frequently necessary to rig a control, such as the swashplate or tail rotor blade angle, to the allowable critical extreme of the tolerance band. For example, it would be necessary to rig the tail rotor to the minimum allowable blade angle if meeting the requirements of § 29.143(c) would be in question. The same consideration must be given to all rotorcraft controls and moveable aerodynamic surfaces where questionable compliance with the regulations may exist. If the rotor-induced vibration characteristics of the rotorcraft are significantly affected and require time-consuming rigging for such things as acceptable ride comfort, then the rotor(s) should be rigged to the allowable extreme tolerance limits to determine compliance, for example, with § 29.251.

Page B - 102 5/1/2014 AC 29-2C, Chg 4 (4) During the FAA/AUTHORITY flight test program, the crew should be especially alert for conditions requiring great attentiveness, high skill levels, or exceptional strength. If any of these features appear marginal, it is advisable to obtain another pilot’s assessment and to carefully document the results of these evaluations.

Section 29.695 requires an alternate system allowing continued safe flight and landing following any single failure of a control system hydraulic boost system. This assessment of ‘safe’ should take into account not only residual post-failure control loads but also workload and pilot fatigue considerations. The following is suggested as an appropriate test sequence, conducted by a range of pilots, for VFR approval: (i) Simulated hydraulic failure at critical flight conditions and Max GW.

(ii) Establish level flight at a cruise speed > V .

Y (iii) Fly for approximately 30 minutes (to assess workload and account for pilot strength variations), demonstrating ability to climb or descend and small bank angle turns. This also allows for the possibility that the hydraulic failure occurs over water or other undesirable landing area.

(iv) Land at a suitable area using a recommended landing technique, appropriate to the emergency.

(5) Because control loads typically increase at higher altitudes, it should be considered if this failure mode should also be investigated during high altitude testing at Max GW/σ. Where approval for any other type of operation is requested (e.g., IFR, category A), an appropriate test sequence must be proposed to the FAA/Authority.

Section 29.141(b) provides the regulatory requirements for these strength and skill requirements, as well as a smooth transition capability between appropriate flight conditions. These requirements must also be met during appropriate engine failure conditions for each category of rotorcraft. Flight characteristics and pilot workload should be evaluated in all expected flight conditions, including actual turbulence.

(6) For night or IFR approval, § 29.141(c) requires additional characteristics for night and IFR flight. The appropriate flight test procedures are included in other portions of this guidance.

b. Procedures. None.

Page B - 103 AC 29-2C, Chg 4 5/1/2014 AC 29.143. § 29.143 (Amendment 29-24) CONTROLLABILITY AND MANEUVERABILITY.

a. Explanation .

(1) This regulation contains the basic controllability requirements for transport rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for control and for maneuverability are summarized in § 29.143(a) which is largely self-explanatory. During the assessment carried out under § 29.143(a)(2)(v) for rotorcraft in glide (i.e., autorotation), in addition to controllability and maneuverability, it should be shown by flight testing that the directional stability characteristics are sufficient to allow the pilot to control the rotorcraft without undue attention to heading at the speeds for minimum rate of descent and best angle of glide.

This should be evaluated at normal trim conditions and in turns up to 30° angle of bank.

The ability to generate a sideslip must be evaluated throughout the autorotation speed envelope. The hover condition is not specifically addressed in § 29.143(a)(2) so that the general requirement may remain applicable to all rotorcraft types, including those without hover capability. For rotorcraft, the hover condition clearly applies under “any maneuver appropriate to the type.” The rotorcraft must still meet the stability requirements of Subpart B and, if applicable, Appendix B.

(2) Paragraphs (b) through (e), § 29.143, include more specific flight conditions and highlight the typical areas of concern during a flight test program.

(i) Section 29.143(b) specifies flight at V with critical weight, center of NE gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at V for NE nose-down pitching of the rotorcraft and for adequate roll control. Nose-down pitching capability is needed for control of gust response and to allow necessary flight path changes in a nose-down direction. Roll control is needed for gust response and for normal maneuvering of the aircraft. In the past, 10 percent control margin has been applied as an appropriate minimum control standard. The required amount of control power, however, has very little to do with any fixed percentage of remaining control travel. There are foreseeable designs for which 5 percent remaining is adequate and others for which 20 percent may not be enough. The key is whether the remaining longitudinal control travel at V generate a clearly positive nose-down pitching moment NE and will the remaining lateral travel allow at least 30° banked turns at reasonable roll rates. Moderate lateral control reversals should be included in this evaluation and since available roll control can diminish with sideslip, reasonable out of trim conditions (directionally) should be investigated. This “control remaining” philosophy must also be applied for other flight conditions specified in this section.

(ii) Section 29.143(c) requires a minimum 17-knot control capability for hover and takeoff in winds from any azimuth. Control capability in wind from zero to at least 17 knots must also be shown for any other appropriate maneuver near the ground such as rolling takeoffs for wheeled rotorcraft. These requirements must be met at all altitudes approved for takeoff and landing. On rotorcraft incorporating a tail rotor, Page B - 104 5/1/2014 AC 29-2C, Chg 4 efficiency of the tail rotor decreases with altitude so that a given sideward flight condition requires more pedal deflection, a higher tail rotor blade angle, and more horsepower. Hence, directional capability in sideward flight (or at critical wind azimuth) is most critical during testing at a high altitude site. Prior to Amendment 29-24, hover controllability, height-velocity, and hover performance were the three regulatory requirements that ordinarily determined the shape of the limiting weight-altitude - temperature (WAT) curve for takeoff and landing. For category A performance rotorcraft operations, of course, the one engine inoperative (OEI) climb performance requirements may also influence the WAT limit curve. Amendment 29-24 allows, under certain conditions, the deletion of any hover controllability condition determined under § 29.143(c) from becoming an operating limitation. Section 29.1587 of Amendment 29 - 24 provides a means wherein category B certificated rotorcraft (in accordance with the requirements of § 29.1, effective with Amendment 29-21) may not be limited by the hover controllability requirements of § 29.143(c). Section 29.1583(g) requirements for category A certificated rotorcraft are unchanged from past regulatory requirements in that if the hover controllability requirements of § 29.143(c) result in the most restrictive envelope it will be published as an operating limitation. Section 29.1587(b) provides a means wherein category B certificated rotorcraft, as defined in § 29.1, may not be restricted in its utilization. It allows such rotorcraft to publish the maximum takeoff and landing capabilities of the rotorcraft, provided something other than the 17 knot hover controllability requirement is not limiting. This may be zero wind IGE hover performance or any other performance the applicant elects to use if the maximum safe wind for operations near the ground is provided. Rotorcraft certificated prior to Amendment 29 - 24 can update their certification basis to take advantage of this provision. If an applicant with a previously type certificated rotorcraft elects to update to this later amendment, caution should be taken to verify that the HV information is done in accordance with Amendment 29-21; that all engine out landing capabilities are satisfactorily accounted for at the new proposed gross weight, altitude, temperature combinations; that takeoff and landing information is provided; and that sufficient information is provided to properly advise the crew of the rotorcraft’s capabilities when utilizing this increased performance capabilities.

(iii) Section 29.143(d) requires adequate controllability when an engine fails.

This requirement specifies conditions under which engine failure testing must be conducted and includes minimum required delay times.

(A) For rotorcraft which meet the engine isolation requirements of category A, demonstration of sudden complete single-engine failure is required at critical conditions throughout the flight envelope including hover, takeoff, climb at V , Y and high speed flight up to V . Entry conditions for the first engine failure are engine NE or transmission limiting maximum continuous power (MCP) (or take-off power where appropriate) including reasonable engine torque splits. For multiengine category A installations (three or more engines) subsequent engine failures should be conducted utilizing the same criteria as that used for first-engine failure. The applicant may limit the flight envelope for subsequent failures. Initial or sequential engine failure tests are ordinarily much less severe than the “last” engine failure test required by § 29.75(b)(5).

Page B -104.1 AC 29-2C, Chg 4 5/1/2014 The conditions for last-engine failure are MCP, or 30-minute power if that rating is approved, level flight, and sudden engine failure with the same pilot delay of 1 second or normal pilot reaction time, whichever is greater.

(B) For category B powerplant installation rotorcraft, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to V (power-on) and conditions of hover, NE takeoff and climb at V . MCP is specified prior to the failure for the cruise condition.

Y Power levels appropriate to the maneuver should be used for other conditions. The corrective action time delay for the cruise failure should be 1 second or normal pilot reaction time (whichever is greater). Cyclic and directional control motions which are part of the pilot task of flight path control are normally not subject to the 1-second restriction; however, the delay is always applied to the collective control for the cruise failure. If the aircraft flying qualities and cyclic trim configuration would encourage routine release of the cyclic control to complete other cockpit tasks during cruise flight, consideration should be given to also holding cyclic fixed for the 1-second delay.

Although the same philosophy could be extended to the directional controls, the likelihood of the pilot’s feet being away from the pedals is much lower, unless the aircraft has a heading hold feature. Rotor speed at execution of the cruise condition power failure should be the minimum power-on value. The term “cruise” also includes cruise climb and cruise descent conditions. Normal pilot reaction times are used elsewhere. Although this requirement specifies MCP, it does not limit engine failure testing to MCP. If a takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions in order to comply with § 29.63(c). Following power failure, rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits.

(iv) Section 29.143(e) addresses the special case in which a V (power-off) NE is established at an airspeed value less than V (power-on). For this case, engine NE failure tests are still required at speeds up to and including V (power-on), and the NE rotorcraft must be capable of being slowed to V (power-off) in a controlled manner NE with normal pilot reactions and skill. There is, however, no controllability requirement for stabilized power-off flight at speeds above 1.1 V (power-off) when V (power-off) NE NE is established per § 29.1505(c).

(v) Application of the controllability requirement for pitch, roll, and yaw at speeds of 1.1 V (power-off) and below is similar to that described above for power-on NE testing at V . Sufficient directional control must exist to allow straight flight in NE autorotation during all approved maneuvers including 30° banked turns up to V NE (power-off) with some small additional allowance for gust control. Adequate controllability margins must exist in all axes throughout the approved autorotative flight envelope. Testing to V at MCP per § 29.143(b), 1.1 V at power for 0.9 V per NE NE H § 29.175(b) or § 29.1505, and to 1.1 V (power-off) in autorotation per § 29.143(e) NE should be sufficient to assure adequate control margin during a descent condition at high speed and low power. The high speed, power-on descent condition should be checked for adequate control margin as a “maneuver appropriate to the type.” There Page B – 104.2 5/1/2014 AC 29-2C, Chg 4 has been one instance where insufficient directional pedal was available to maintain a reasonable trimmed sideslip angle with low power at very high speeds, and a case where there was insufficient forward and lateral cyclic available to reach the power-on V . The insufficient directional pedal margin was due to the offset vertical stabilizers.

NE The lack of cyclic stick margin was because the cyclic stick migrated to the right as power was reduced and the control limits were circular. This provided less total available forward cyclic stick travel when the cyclic was moved right and forward about 45° from the center position. Each of the above rotorcraft was certificated with a rate of descent limitation to preclude operation in the control-limited area.

(vi) An evaluation of the emergency descent capability of the rotorcraft should be made, either analytically or through flight test. Areas of consideration are the rate of descent available, the maximum approved altitude, and the time before a catastrophic failure following the loss of transmission oil pressure or other similar failure. Each rotorcraft should have the capability to descend to sea level and land from the maximum certificated altitude within the time period established as safe following a critical failure.

If the time period does not permit a sea level landing, the maximum height above the terrain must be specified in the limitation section of the rotorcraft flight manual (RFM).

(3) The required controllability and maneuvering capabilities must also be considered following the failure of automatic equipment used in the control system (§ 29.672). Examples include stability augmentation systems (SAS), stability and control augmentation systems (SCAS), automatic flight control systems (AFCS), devices to provide or improve longitudinal static stability such as a pitch bias actuator (PBA), yaw dampers, and fly-by-wire elevator or stabilator surfaces. These systems all use actuators of some type, and they are subject to actuator softover and hardover malfunctions. The flight control system should be evaluated to determine whether an actuator jammed in an extreme position would result in reduced control margins.

Generally, if the flight control system stops are between the actuator and the cockpit control, the control margin will be affected. If the control stops are between the actuator and the rotor head, the control margins may not be affected, but the location of the cockpit control may be shifted. This could produce interference with other items in the cockpit. An example of this would be a lateral actuator jammed hardover causing a leftward shift in the cyclic stick position. Interference between the cyclic stick, the pilot’s leg, and the collective pitch control could reduce the left lateral control available and reduce left sideward flight capability. In the case of fly-by-wire surfaces, both the high speed forward flight controllability and the rearward flight capabilities could be affected.

Flight control systems that incorporate automatic devices should be thoroughly evaluated for critical areas. Every failure condition that is questionable should be flight tested with the appropriate actuator fixed in the critical failure position. These failures may require limitations of the flight envelope. Any procedure or limitation that must be observed to compensate for an actuator hardover or softover malfunction should be included in the RFM.

b. Procedures .

Page B – 104.3 AC 29-2C, Chg 4 5/1/2014 (1) Flight test instrumentation should include ambient parameters, all flight control positions, rotor RPM, main and tail rotor flapping (if appropriate), engine power instruments, and throttle position. Flight controls that are projected to be near their limits of authority should be rigged to the most adverse production tolerance. A very accurate weight and balance computation is needed along with a precise knowledge of the aircraft’s weight and CG variation as fuel is burned.

(2) The critical condition for V controllability testing is ordinarily aft CG, MCP, NE and minimum power-on rotor RPM, although power and RPM variations should be specifically evaluated to verify their effects. The turbine engine is sensitive to ambient temperatures which affect the engine’s ability to produce rated maximum continuous torque. Flight tests conducted at ambient temperatures that cause the turbine temperature to limit MCP would not produce the same results obtained at the same density altitude at colder ambient temperatures where maximum continuous torque would be limiting. Forward CG should be spot checked for any “tuck under” tendency at high speed. The V controllability test is normally accomplished shortly after the NE 1.1 V (or 1.1 V ) point obtained during stability tests required by § 29.175(b).

NE H Controllability must be satisfactory for both conditions. If V varies with altitude or NE temperature, V for existing ambient conditions is utilized for the test. Extremes of the NE altitude or temperature envelope should be analyzed and investigated by flight test.

(3) The critical condition for controllability testing in a hover is ordinarily forward CG at maximum weight with minimum power-on rotor RPM. For rearward flight testing of configurations where the forward CG limit varies with weight, low or high gross weight may be critical. Lateral CG limits should also be investigated. A calibrated pace vehicle is needed to assure stabilized flight conditions. Surface winds should be less than 3 knots throughout the test sequence. Testing can be done in higher stabilized wind conditions (gusting less than 3 knots); however, these conditions are very difficult to find and the method is very time consuming due to the necessity of waiting for stabilized winds. Testing in calm winds is preferred. Hover controllability testing should be accomplished with the lowest portion of the rotorcraft at the published hover height above ground level; however, the test altitude above the ground may be increased to provide reasonable ground clearance. Although the necessary yaw response will vary somewhat from model to model, sufficient control power should be available to permit a clearly recognizable yaw response after full directional control displacement when the rotorcraft is held in the most critical position relative to wind. Testing will be carried out at the power required to achieve stabilized flight conditions. With rotorcraft that are operating in conditions such that the gross weight is limited by the power available, there should always be adequate tail rotor pedal authority to maintain yaw control when using the maximum approved all engines operating (AEO) power, which is takeoff power (TOP) for most designs.

(i) Where the rotorcraft is capable of operating at maximum gross weight with less than maximum approved power, it is appropriate to examine the rotorcraft characteristics with small amounts of additional power applied above the trim power required to allow for typical power variations experienced during normal use of the Page B – 104.4 5/1/2014 AC 29-2C, Chg 4 rotorcraft. For example , maneuvering, turbulence, or rotor governing characteristics may cause the pilot to use power in excess of power required for trim. The maximum power excursions should not exceed maximum approved AEO power, excluding any transient range.

(ii) The rotorcraft should be flown both in ground effect (IGE) and out of ground effect (OGE), with the most adverse wind speed and direction for directional control within the flight envelope proposed, using power variations above trim that might be expected during normal use of the rotorcraft. Consideration should be given to the amount of excess power available, the ease with which power can be controlled via the collective, the effect of tail rotor control inputs on power required, and the characteristics of the rotorcraft if the limits of directional control are approached. There should be no tendency to deviate rapidly or suddenly in yaw. This assessment is normally conducted in conjunction with the critical azimuth testing.

(4) Prior to engine failure testing, it is mandatory that the pilot be fully aware of the engine, drive system, and rotor limits. These limits were established during previous ground and flight tests and they should be specified in the TIA. Particular attention should be given to minimum stabilized and minimum transient rotor RPM limits. These values must be included in the TIA and should be approached gradually with a build-up in time delay unless the company testing has completely validated all pertinent aspects of engine failure testing. On category A installations the maximum power output of each engine must be limited so that when an engine fails and the remaining engine(s) assume the additional load, the remaining engine(s) are not damaged by excessive power extraction and over-temping. This is needed for compliance with § 29.903(b). The propulsion engineer should have assured that this feature was properly addressed in the engine and drive system substantiation; however, it must be assumed that for some period of time the pilot may extract maximum available power from the remaining engine(s) when an engine fails during critical flight maneuvers. Substantiation of this feature should be accomplished primarily by engine and drive system ground tests.

(5) Longitudinal cyclic authority at V with any power setting must permit NE suitable nose-down pitching of the rotorcraft. If the remaining control travel is considered marginal, tests should include applications up to full control deflection to assess the remaining authority. Some knowledge of the aircraft’s response to turbulence is useful in assessing the remaining margin. As a minimum, the rotorcraft must have adequate margin available to overcome a moderate turbulent gust and must not have any divergent characteristic which requires full deflection of the primary recovery control to arrest aircraft motion. If other controls must be utilized to overcome adverse aircraft motion, the results are unacceptable (e.g., if a pitch up tendency resulting from an actual or simulated moderate turbulent gust cannot be satisfactorily overcome by remaining forward cyclic, the use of throttle or collective controls to assist the recovery is not an acceptable procedure; however, the use of lateral cyclic to correct roll in conjunction with forward cyclic to correct pitchup is satisfactory). Obviously during the conduct of these tests, all available techniques should be utilized when the Page B – 105 AC 29-2C, Chg 4 9/17/2009 pilot finds himself “out of control.” However, compliance with this section requires that recovery must be shown by use of only the primary control for each axis of aircraft motion.

(6) Cyclic control authority in autorotation must be sufficient to allow adequate flare capability and landing under the all engine inoperative requirements of § 29.75 (see section 29.75 of this AC).

AC 29.143A. § 29.143 (Amendment 29-51) Controllability and Maneuverability.

a. Explanation . Amendment 29-51 made a minor clarification to assure that in- ground-effect (IGE) controllability is demonstrated at all wind speeds up to 17 knots, for all azimuths. In many rotorcraft, the entry into the regime of translational lift requires the most power, thus potentially causing control difficulties, and frequently occurs at speeds less than 17 knots. The amendment also requires that out-of-ground-effect (OGE) controllability be determined up to a speed of at least 17 knots at a weight selected by the applicant. The amendment clarifies the intent of Amendment 29-21 and Amendment 29-24 with respect to removing hover controllability as a limit.

Section 29.25 is amended to assure that appropriate weight limitations are incorporated into the Rotorcraft Flight Manual (RFM) when the relieving provisions of the previous amendments are adopted by an applicant. The previous amendment and associated AC material indicated that certain Category B rotorcraft were relieved from providing, as a limitation, the conditions of § 29.143(c). In practice, the 17-knot controllability requirement was still treated as a limitation, but, as indicated in the amended § 29.25, additional limits could be included, when demonstrated, that allowed for something other than 17-knot all azimuth controllability. The established weight, altitude, and temperature charts, including any associated wind constraints, could be contained in the performance section of the flight manual when the appropriate reference to those charts were included in the limitations section of the RFM. In addition, the relief of Amendments 29-21 and 29-24 were only intended for those category B rotorcraft with nine or less passenger seats. All the policy material pertaining to this section remains in effect with the following changes: (1) This regulation contains the basic controllability requirements for transport rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for controllability and for maneuverability are summarized in § 29.143(a) which is self-explanatory. The hover condition is not specifically addressed in § 29.143(a)(2) so that the general requirement may remain applicable to all rotorcraft types, including those without hover capability. For rotorcraft, the hover condition clearly applies under "any maneuver appropriate to the type."

(2) Paragraphs (b) through (e) in § 29.143 include more specific flight conditions and highlight the typical areas of concern during a flight test program.

(i) Section 29.143(b) specifies flight at V with critical weight, center of NE gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at V for NE Page B – 106 9/17/2009 AC 29-2C, Chg 4 nose down pitching of the rotorcraft and for adequate roll control. Nose down pitching capability is needed for control of gust response and to allow necessary flight path changes in a nose down direction. Roll control is needed for gust response and for normal maneuvering of the aircraft. In the past, 10 percent control margin has been applied as an appropriate minimum control standard. The required amount of control power, however, has very little to do with any fixed percentage of remaining control travel. There are foreseeable designs for which 5 percent remaining is adequate and others for which 20 percent may not be enough. The key is, can the remaining longitudinal control travel at V generate a clearly positive nose down pitching moment, NE and will the remaining lateral travel allow at least 30° banked turns at reasonable roll rates? Moderate lateral control reversals should be included in this evaluation and since available roll control can diminish with sideslip, reasonable out of trim conditions (directionally) should be investigated. This "control remaining" philosophy must also be applied for other flight conditions specified in this section.

(ii) Section 29.143(c) requires a minimum control capability for hover and takeoff in winds from zero to at least 17 knots from any azimuth. Control capability in wind from zero to at least 17 knots must also be shown for any other appropriate maneuver near the ground such as rolling takeoffs for wheeled rotorcraft. These requirements must be met at all altitudes approved for takeoff and landing. On helicopters incorporating a tail rotor, efficiency of the tail rotor decreases with altitude so that a given sideward flight condition requires more pedal deflection, a higher tail rotor blade angle, and more horsepower. Hence, directional capability in sideward flight (or at critical wind azimuth) is most critical during testing at a high altitude site. Prior to Amendment 29-24, hover controllability, height-velocity, and hover performance were the three regulatory requirements that ordinarily determined the shape of the limiting weight-altitude-temperature (WAT) curve for takeoff and landing. For Category A performance rotorcraft operations, of course, the one-engine-inoperative (OEI) climb performance requirements may also influence the WAT limit curve. Amendment 29-24 allows, under certain conditions, the deletion of any hover controllability condition determined under § 29.143(c) from becoming an operating limitation. Section 29.1587 of Amendment 29-24 provides a means wherein Category B certificated rotorcraft (in accordance with the requirements of § 29.1, effective with Amendment 29-21) may not be limited by the hover controllability requirements of § 29.143(c). Section 29.1583(g) requirements for Category A certificated rotorcraft are unchanged from past regulatory requirements in that if the hover controllability requirements of § 29.143(c) result in the most restrictive envelope it will be published as an operating limitation. Section 29.1587(b) provides a means wherein Category B certificated rotorcraft, as defined in § 29.1, may not be restricted in its utilization. Section 29.1587(b) allows some Category B RFMs to include maximum takeoff and landing performance information, provided that something other than the 17-knot hover controllability requirement is not limiting. This may be zero wind IGE hover performance or any other performance the applicant elects to use, if the maximum safe wind for operations near the ground is provided. Rotorcraft certificated prior to Amendment 29-24 can update their certification basis to take advantage of this provision. If an applicant with a previously type certificated rotorcraft elects to update to this later amendment, caution should be taken to verify that the Page B – 106.1 AC 29-2C, Chg 4 9/17/2009 height-velocity information is done in accordance with Amendment 29-21; that all engine out landing capabilities are satisfactorily accounted for at the new proposed gross weight, altitude, temperature combinations; that takeoff/landing information is provided; and that sufficient information is provided to properly advise the crew of the rotorcraft's capabilities when utilizing this increased performance capabilities.

(iii) Section 29.143(e) requires adequate controllability when an engine fails.

This requirement specifies conditions under which engine failure testing must be conducted and includes minimum required delay times.

(A) For rotorcraft that meet the engine isolation requirements of Category A, demonstration of sudden complete single-engine failure is required at critical conditions throughout the flight envelope including hover, takeoff, climb at V , and high Y speed flight up to V . Entry conditions for the first engine failure are engine or NE transmission limiting maximum continuous power (MCP) (or takeoff power where appropriate) including reasonable engine torque splits. For multiengine Category A installations with three or more engines, the subsequent engine failures should be conducted utilizing the same criteria as that used for first-engine failure. The applicant may limit his flight envelope for subsequent failures. Initial or sequential engine failure tests are ordinarily much less severe than the "last" engine failure test required by § 29.75(b)(5). The conditions for last-engine failure are MCP or 30-minute power if that rating is approved, level flight, and sudden engine failure with the same pilot delay of 1-second or normal pilot reaction time, whichever is greater.

(B) For Category B powerplant installation rotorcraft, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to V (power-on) and conditions of hover, NE takeoff, and climb at V . MCP is specified prior to the failure for the cruise condition.

Y Power levels appropriate to the maneuver should be used for other conditions. The corrective action time delay for the cruise failure should be 1 second or normal pilot reaction time (whichever is greater). Cyclic and directional control motions which are part of the pilot task of flight path control are normally not subject to the 1-second restriction; however, the delay is always applied to the collective control for the cruise failure. If the aircraft flying qualities and cyclic trim configuration encourage routine release of the cyclic control to complete other cockpit tasks during cruise flight, consideration should be given to also holding cyclic fixed for the 1-second delay.

Although the same philosophy could be extended to the directional controls, the likelihood of the pilot having his feet away from the pedals is much lower, unless the aircraft has a heading hold feature. Rotor speed at execution of the cruise condition power failure should be the minimum power-on value. The term "cruise" also includes cruise climb and cruise descent conditions. Normal pilot reaction times are used elsewhere. Although this requirement specifies MCP, it does not limit engine failure testing to MCP. If a takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions in order to comply with § 29.63(c). Following power failure, the rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits.

Page B – 106.2 9/17/2009 AC 29-2C, Chg 4 (iv) Section 29.143(f) addresses the special case in which a V (power-off) NE is established at an airspeed value less than V (power-on). For this case, engine NE failure tests are still required at speeds up to and including V (power-on), and the NE rotorcraft must be capable of being slowed to V (power-off) in a controlled manner NE with normal pilot reactions and skill. There is, however, no controllability requirement for stabilized power-off flight at speeds above 1.1 V (power-off) when V (power-off) NE NE is established per § 29.1505(c).

(v) Application of the controllability requirement for pitch, roll, and yaw at speeds of 1.1 V (power-off) and below is similar to that described above for power-on NE testing at V . Sufficient directional control must exist to allow straight flight in NE autorotation during all approved maneuvers including 30° banked turns up to V NE (power-off) with some small additional allowance for gust control. Adequate controllability margins must exist in all axes throughout the approved autorotative flight envelope. Testing to V at MC power per § 29.143(b) and § 29.175(c), and to 1.1 V NE NE (power-off) in autorotation per § 29.143(f) should be sufficient to assure adequate control margin during a descent condition at high speed and low power. The high speed, power-on descent condition should be checked for adequate control margin as a "maneuver appropriate to the type." There has been one instance where insufficient directional pedal was available to maintain a reasonable trimmed sideslip angle with low power at very high speeds, and a case where there was insufficient forward and lateral cyclic available to reach the power-on V . The insufficient directional pedal margin NE was due to the offset vertical stabilizers. The lack of cyclic stick margin was because the cyclic stick migrated to the right as power was reduced and the control limits were circular. This provided less total available forward cyclic stick travel when the cyclic was moved right and forward about 45° from the center position. Each of the above rotorcraft was certificated with a rate of descent limitation to preclude operation in the control-limited area.

(vi) An evaluation of the emergency descent capability of the rotorcraft should be made, either analytically or through flight test. Areas of consideration are the rate of descent available, the maximum approved altitude, and the time before a catastrophic failure following the loss of transmission oil pressure or other similar failure.

Each rotorcraft should have the capability to descend to sea level and land from the maximum certificated altitude within the time period established as safe following a critical failure. If the time period does not permit a sea level landing, the maximum height above the terrain must be specified in the limitation section of the RFM.

(3) The required controllability and maneuvering capabilities must also be considered following the failure of automatic equipment used in the control system (§ 29.672). Examples include stability augmentation systems (SAS), stability and control augmentation systems (SCAS), automatic flight control systems (AFCS), devices to provide or improve longitudinal static stability such as a pitch bias actuator (PBA), yaw dampers, and fly-by-wire elevator or stabilator surfaces. These systems all use actuators of some type, and they are subject to actuator softover and hardover Page B – 106.3 AC 29-2C, Chg 4 9/17/2009 malfunctions. The flight control system should be evaluated to determine whether an actuator jammed in an extreme position would result in reduced control margins.

Generally, if the flight control system stops are between the actuator and the cockpit control, the control margin will be affected. If the control stops are between the actuator and the rotor head, the control margins may not be affected, but the location of the cockpit control may be shifted. This could produce interference with other items in the cockpit. An example of this would be a lateral actuator jammed hardover causing a leftward shift in the cyclic stick position. Interference between the cyclic stick, the pilot's leg, and the collective pitch control could reduce the left lateral control available and reduce left sideward flight capability. In the case of fly-by-wire surfaces, both the high speed forward flight controllability and the rearward flight capabilities could be affected.

Flight control systems that incorporate automatic devices should be thoroughly evaluated for critical areas. Every failure condition that is questionable should be flight tested with the appropriate actuator fixed in the critical failure position. These failures may require limitations of the flight envelope. Any procedure or limitation that must be observed to compensate for an actuator hardover or softover malfunction should be included in the RFM.

b. Procedures . The policy material pertaining to this section remains in effect with the following changes and additions: (1) Flight test instrumentation should include ambient parameters, all flight control positions, rotor RPM, main and tail rotor flapping (if appropriate), engine power instruments, and throttle position. Flight controls that are projected to be near their limits of authority should be rigged to the most adverse production tolerance. A very accurate weight and balance computation is needed along with a precise knowledge of the aircraft's weight/CG variation as fuel is burned.

(2) The critical condition for V controllability testing is ordinarily aft CG, MC NE power, and minimum power-on rotor RPM, although power and RPM variations should be specifically evaluated to verify their effects. The turbine engine is sensitive to ambient temperatures which affect the engine's ability to produce rated maximum continuous torque. Flight tests conducted at ambient temperatures that cause the turbine temperature to limit MCP would not produce the same results obtained at the same density altitude at colder ambient temperatures where maximum continuous torque would be limiting. Forward CG should be spot checked for any "tuck under" tendency at high speed. The V controllability test is normally accomplished shortly NE after the 1.1 V (or 1.1 V ) point obtained during stability tests required by § 29.175(b).

NE H Controllability must be satisfactory for both conditions. If V varies with altitude or NE temperature, V for existing ambient conditions is utilized for the test. Extremes of the NE altitude/temperature envelope should be analyzed and investigated by flight test.

(3) Controllability (i) The critical condition for controllability testing in a hover is ordinarily forward CG at maximum weight with minimum power-on rotor RPM. For rearward flight Page B – 106.4 9/17/2009 AC 29-2C, Chg 4 testing of configurations where the forward CG limit varies with weight, low or high gross weight may be critical. Lateral CG limits should also be investigated. A calibrated pace vehicle is needed to assure stabilized flight conditions. Surface winds should be less than 3 knots throughout the test sequence. Testing can be done in higher stabilized wind conditions (gusting less than 3 knots); however, these conditions are very difficult to find and the method is very time consuming due to the necessity of waiting for stabilized winds. Testing in calm winds is preferred. IGE hover controllability testing should be accomplished with the lowest portion of the rotorcraft at the published hover height above ground level; however, the test altitude above the ground may be increased to provide reasonable ground clearance. OGE testing should be done with the rotor at a predetermined height above the ground at which it has been determined that there is no ground effect. Although the necessary yaw response will vary somewhat from model to model, sufficient control power should be available to permit a clearly recognizable yaw response after full directional control displacement when the rotorcraft is held in the most critical position relative to wind.

(A) Testing will normally be carried out at the power required to achieve stabilized flight conditions. However, it is also important to show that yaw control remains adequate to allow normal power changes that might be required in normal operational maneuvers typical for the type and use of the rotorcraft. With rotorcraft that are operating in conditions in which the gross weight is limited by the power available, there should always be adequate tail rotor pedal control available to maintain yaw control when using up to Take-off Power. However, this will not be the case if the rotorcraft weight in the low speed flight envelope is limited by yaw control system capability. There may be other conditions where adequate yaw control is not available at high power, for example a rotorcraft which is limited by the CAT A weight (for rotorcraft certificated to § 29.1 (c)).

(B) To cover the case where excess power is available, it is appropriate to examine the rotorcraft characteristics with some small amounts of additional power applied. This will account for typical power variations that will be experienced during normal use of the rotorcraft. For example, maneuvering or turbulence will cause the pilot to use some of the excess power available. The rotorcraft should be flown, both IGE and OGE, with the most adverse wind speed and direction for directional control within the flight envelope proposed. Use power variations above trim that might be expected during normal use of the rotorcraft giving consideration to the amount of excess power available, the ease with which power can be controlled by collective, and the characteristics of the rotorcraft if the limits of directional control are approached.

There should be no tendency to deviate rapidly or suddenly in yaw. This assessment is normally conducted in conjunction with the critical azimuth testing.

(C) It may be appropriate to provide flight manual information on the directional control characteristics, including any relevant maximum power above which it could be expected that directional control might not be maintained.

Page B – 106.5 AC 29-2C, Chg 4 9/17/2009 (ii) Comprehensive controllability tests are typically conducted at low, intermediate (~7000 feet Hd), and high tests sites, with prepared landing surfaces, in conjunction with takeoff, landing, and performance testing.

(iii) Alternatively, a predicted controllability model developed for high altitude may be used if verified by limited flight testing with steady ambient winds. The extrapolation guidelines in AC 29.45 b(2) are still applicable. These high altitude controllability tests could typically be conducted in conjunction with takeoff, landing and performance tests.

(iv) Controllability can usually be extrapolated up to a maximum of 2,000 feet above the highest test site altitude.

Note: Engine operating characteristics must be considered during the limited altitude tests.

(4) Prior to engine failure testing, the pilot should be fully aware of his engine, drive system, and rotor limits. These limits were established during previous ground and flight tests and they should be specified in the TIA. Particular attention should be given to minimum stabilized and minimum transient rotor RPM limits. These values should be included in the TIA and should be approached gradually with a build-up in time delay unless the company testing has completely validated all pertinent aspects of engine failure testing. On Category A installations, the maximum power output of each engine should be limited so that when an engine fails and the remaining engine(s) assume the additional load, the remaining engine(s) are not damaged by excessive power extraction and exceeding a temperature limitation. This is needed for compliance with § 29.903(b). The propulsion engineer should have assured that this feature was properly addressed in the engine and drive system substantiation; however, it must be assumed that for some period of time the pilot may extract maximum available power from the remaining engine(s) when an engine fails during critical flight maneuvers.

Substantiation of this feature should be accomplished primarily by engine and drive system ground tests.

(5) Longitudinal cyclic authority at V with any power setting must permit NE suitable nose down pitching of the rotorcraft. If the remaining control travel is considered marginal, tests should include applications up to full control deflection to assess the remaining authority. Some knowledge of the aircraft's response to turbulence is useful in assessing the remaining margin. As a minimum, the rotorcraft must have adequate margin available to overcome a moderate turbulent gust and must not have any divergent characteristic which requires full deflection of the primary recovery control to arrest aircraft motion. If other controls must be utilized to overcome adverse aircraft motion, the results are unacceptable; e.g., if a pitch up tendency resulting from an actual or simulated moderate turbulent gust cannot be satisfactorily overcome by remaining forward cyclic, the use of throttle or collective controls to assist the recovery is not an acceptable procedure; however, the use of lateral cyclic to correct roll in conjunction with forward cyclic to correct pitch-up is satisfactory. Obviously Page B – 106.6 9/17/2009 AC 29-2C, Chg 4 during the conduct of these tests, all available techniques should be utilized when the pilot finds himself "out of control." However, compliance with this section requires that recovery must be shown by use of only the primary control for each axis of aircraft motion.

(6) Cyclic control authority in autorotation must be sufficient to allow adequate flare capability and landing under the requirements of § 29.143(a)(2)(v) and (vi).

Page B - 107 AC 29-2C, Chg 4 5/1/2014 AC 29.151. § 29.151 (Amendment 29-24) FLIGHT CONTROLS.

a. Explanation . Excessive breakout or preload in the flight controls produces control system force discontinuities, which result in increased workload and even controllability problems for the pilot. Similarly, excessive freeplay results in lost motion, which increases pilot workload and, in an extreme case, could lead to a hazardous pilot- induced oscillation. Although in some designs friction can provide a positive contribution to the function of the flight controls (e.g., masking aerodynamic feedback in reversible systems), friction will eventually have a detrimental effect on the pilot’s ability to properly control the rotorcraft. In the case of an irreversible design equipped with an artificial force feel system in pitch and roll, excessive friction can mask a shallow force gradient making positive stick centering and control force static stability difficult if not impossible to demonstrate. In such an instance, the initial choice of fixes might include implementation of a steeper force gradient or addition of a force preload. Care must be exercised during the initial design phase to ensure that the components and characteristics of the flight control system are well matched.

b. Procedures . Regardless of the flight control system sophistication, it is important that the test pilot understand the system configuration prior to flight evaluation.

Appropriate mechanical characteristics should be documented. For VFR rotorcraft, the mechanical characteristics are typically assessed in flight on a qualitative basis. If a controllability or workload problem is identified, a more detailed investigation would be necessary. Since IFR certification rules include specific trim and force requirements, a more quantitative investigation of mechanical characteristics is normally conducted.

The constantly varying feedback forces of reversible flight control systems generally make such designs unsuitable for IFR application. Irreversible system mechanical characteristics can often be partially documented on the ground with external hydraulic and electrical power supplies connected to the rotorcraft. Characteristics of the flight control system should be qualitatively considered during other flight tests. These characteristics include forces, control harmony, nonlinearities, discontinuities, proper directional senses, breakout, friction, hysteresis, etc.

Page B - 108 5/1/2014 AC 29-2C, Chg 4 AC 29.161. § 29.161 TRIM CONTROL.

a. Explanation .

(1) The pilot has many tasks to perform with each hand during sustained flight conditions. The trim requirement is intended to provide the pilot with a reference cyclic control position for the given flight condition, reduce the physical demands to maintain a given flight condition, and allow the pilot to release the cyclic control for brief periods of time to perform other cockpit duties. A primary flight control which can move when released imposes an additional pilot workload by requiring a continuous hands-on condition. It is not intended to require that control forces be reduced to zero by the trim control during dynamic maneuvers such as takeoff acceleration.

(2) A number of devices may be used to produce the necessary trim characteristics. One popular method of meeting this requirement is through the use of control balance springs in conjunction with a small amount of built-in control system friction. Other methods include use of friction, magnetic brakes, bungees, and irreversible mechanical schemes.

(3) This regulation is not intended to require zero friction or zero breakout force in the control system, nor is it intended to require automatic control recentering. The regulation, in fact, specifically prohibits excessive high friction or high breakout forces which would produce undesirable discontinuities in the primary control force gradient.

b. Procedures .

(1) If comprehensive company flight test data are available, compliance with this requirement can quickly be found by spot checking extreme center of gravity loadings.

Trim tests can ordinarily be done during the course of other flight test activities. To conduct the test, simply release the control at the required flight conditions and determine that the control does not move. The words “any appropriate speed” ordinarily include any speed from hover to V . If the control system trim device might be subject H to temperature or humidity effects, these should be investigated at a minimum of two altitude extremes and during several test phases.

(2) If a pilot controllable variable friction device is incorporated, compliance with this requirement must be shown at the minimum adjustable value. The maximum value of adjustable friction should not completely lock the flight controls.

(3) Continued compliance with this requirement should be assured through a production procedure. If minimum friction or centering springs are used, it is desirable for the manufacturer to include some adjustment capability for production differences.

The explanation and procedures discussed here are applicable for VFR approval under § 29.161. For additional IFR trim requirements, refer to AC 29 Appendix B.

Page B - 109 AC 29-2C, Chg 4 5/1/2014 AC 29.161A. § 29.161 (Amendment 29-24) TRIM CONTROL.

a. Explanation . Amendment 29-24 to the regulation adds the additional requirement that the trim control be capable of trimming collective forces to zero.

b. Procedures . The trim requirement is intended to allow the pilot to release the controls for brief periods to perform other cockpit duties, and to provide the pilot with a reference cyclic position for the given flight condition. The collective should be balanced so that there is no tendency for the collective pitch to change when the collective is released. Any magnetic clutch, friction brake or similar device which modifies the collective characteristics should be capable of being overpowered by the pilot, when fully applied, without requiring excessive force.

AC 29.171. § 29.171 STABILITY: GENERAL.

a. Explanation . This section is intended to require a manageable pilot workload for the minimum crew under foreseeable operating conditions.

b. Procedures .

(1) Compliance with the requirements of this section can often be obtained for the VFR condition without any specific or designated flight testing. If the rotorcraft is marginal in regard to pilot strain and fatigue, the FAA/AUTHORITY pilot should be assured, through special tests if necessary, that the aircraft can be satisfactorily flown throughout the maximum endurance capabilities of the rotorcraft including night and turbulence conditions if those are critical. This test should be conducted with minimum required systems in the aircraft and with minimum flight crew.

(2) Reasonable failure conditions which add to pilot workload, strain, and fatigue should be evaluated (electrical, hydraulic and mechanical failures, etc.). The necessary times associated with flight with a failed system must be appropriate to the flight manual procedures for each failure. A failure condition requiring immediate landing would obviously require shorter evaluation time than a condition allowing continued flight to destination.

(3) IFR approvals necessitate a careful evaluation of paragraphs (1) and (2) above. In IFR operations, weather conditions frequently necessitate continued flight to destination or diversion to alternate airports with critical failures. Immediate landing may not be feasible. The evaluating pilot must assure pilot strain and fatigue are acceptable during typical flight profiles for each type of operation to be approved.

Page B - 110 9/17/2009 AC 29-2C, Chg 4 AC 29.173. § 29.173 (Amendment 29-24) STATIC LONGITUDINAL STABILITY.

a. Explanation .

(1) This rule contains control system design requirements for both stability and control. Paragraph (a) contains the basic control philosophy necessary for all civil aircraft. Forward motion of the cyclic control must produce increasing speeds and aft motion must result in decreasing speeds. For rotorcraft this is accomplished with throttle and collective held constant. This requirement in no way assures aircraft stability. It is simply a control requirement which speaks to direction of control motion.

Rotorcraft with either highly stable or highly unstable static longitudinal stability characteristics can typically comply with the basic requirement for control sense of motion.

(2) The remainder of § 29.173, through reference to § 29.175, contains the basic control position requirements necessary to establish a minimum level of static longitudinal stability. Positive stability is found for conditions of climb, cruise, and autorotation in § 29.175 by requiring a stable stick position gradient through a specified speed range. A defined level of instability is permitted for the hovering condition.

b. Procedures .

(1) The control requirement of this section is so essential to basic flight mechanics that compliance may be found during conventional flight testing for compliance with other portions of the regulations. No special or designated testing should be required.

(2) The procedures necessary to assure compliance with the stability requirements of this section are contained under § 29.175, Demonstration of static longitudinal stability. Refer to paragraph AC 29.175 for an explanation of detailed flight test procedures.

AC 29.173A. § 29.173 (Amendment 29-51) STATIC LONGITUDINAL STABILITY.

a. Explanation .

(1) Amendment 29-51 makes a major change to the requirement by allowing for neutral or negative static longitudinal stability in limited flight domains. Additionally, the requirement for the hover demonstration found in § 29.173(c) has been deleted as this requirement is adequately covered by the controllability requirements. The basic tenants of the rule are unchanged in that the rule contains control system design requirements for both stability and control. Paragraph (a) contains the basic control philosophy necessary for all civil aircraft. Forward motion of the cyclic control must produce increasing speeds and aft motion must result in decreasing speeds. For rotorcraft, this is accomplished with throttle and collective held constant. This requirement in no way assures aircraft stability. It is simply a control requirement that Page B – 111 AC 29-2C, Chg 4 9/17/2009 speaks to direction of control motion. Rotorcraft with either highly stable or highly unstable static longitudinal stability characteristics can typically comply with the basic requirement for control sense of motion. All the policy material pertaining to this section remains in effect with the following changes and additions: (2) Sections 29.173 through 29.175 contain the basic control position requirements necessary to establish a minimum level of static longitudinal stability.

Positive stability is found for conditions of climb, cruise, V , and autorotation in NE § 29.175 by demonstrating a stable stick position gradient through a specified speed range. This is the primary method of demonstrating compliance with the longitudinal static stability requirements.

(3) For aircraft that do not possess positive control position stability for some limited flight conditions or modes of operation, an equivalent level of safety was previously provided that requires a qualitative evaluation of the pilot’s ability to maintain a given airspeed within 5 knots of the desired speed without exceptional piloting skill or alertness. These flight conditions and modes of operation could include various combinations of gross weight, CG, flight regime (climb, cruise, descent), ambient conditions (altitude/temperature), as well as possible variations in the stability augmentation configuration. In the past, the FAA/AUTHORITIES have certified numerous rotorcraft, under equivalent level of safety findings, which have neutral or negative static longitudinal stick position stability in some flight domains. This amendment to § 29.173 is intended to allow for this case without having to resort to an equivalent safety finding. For these previous equivalent safety findings, acceptable qualitative flight characteristics were found on aircraft, which possessed negative longitudinal stick position gradients of up to 2-3% of total control travel in certain flightregimes; however, this value is not intended to be a limit. When this means of compliance is elected by the applicant, in addition to the qualitative pilot evaluation it is still necessary to collect the data associated with the classical static longitudinal stability testing as defined in § 29.175.

b. Procedures . All the policy material pertaining to this section remains in effect with the following changes and additions: (1) The control requirement of paragraph (a) of this section is so essential to basic flight mechanics that compliance may be found during conventional flight testing for compliance with other portions of the regulations. No special or designated testing should be required.

(2) The procedures necessary to assure compliance with the primary stability requirements of this section are contained under § 29.175, Demonstration of Static Longitudinal Stability. Refer to AC 29.175A of this advisory circular for an explanation of detailed flight test procedures.

(3) The procedures necessary to assure compliance with the alternative (i.e., pilot evaluation) method of compliance are provided below.

Page B – 112 9/17/2009 AC 29-2C, Chg 4 (i) For those limited conditions where compliance with the basic control position requirements cannot be shown, the evaluation must focus on the ability of the pilot to maintain airspeed in the flight regime without exceptional piloting skill or alertness under typical flight conditions. “Limited flight conditions” infers that the aircraft should be in reasonable compliance with the stick position stability requirements of § 29.173(b) for most of the flight conditions and configurations tested. Extraordinary means of complying with § 29.173(b) should not be forced on the aircraft design if the airspeed retention task meets the pilot skill and alertness guidelines. The demonstration flight regimes are defined in § 29.175(a) through (d). For those flight regimes, conditions, and configurations where compliance with stick position requirements of § 29.173(b) cannot be shown, the evaluation pilot should assess the ease of maintaining airspeed within the specified +/- 5 knots.

(ii) When assessing the ease of maintaining airspeed the total workload must be considered. Secondary tasks pertinent to the minimum flight crew in each flight regime should be conducted. This may include visual navigation and communication in cruise, traffic avoidance in climb, and landing site selection in autorotation.

(iii) The cues that the aircraft provides are an important contributor to the evaluation, and the nature of these cues should be noted in the compliance report where this alternate qualitative evaluation determines that the aircraft has satisfactory airspeed stability characteristics. The cues that supplant the control position cues may be found to be sufficient if these cues are natural to the speed maintenance task, andprovide adequate guidance to the pilot during the task. One important cue might be the pitch attitude gradient with speed, where a perceptible change in trimmed pitch attitude is required for a perceptible airspeed change. Where pitch attitude is the predominant cue the relationship should be positive (nose down with airspeed increase) and perceptible without exceptional alertness. With this relationship, the evaluation pilot may find that the natural pitch control tasks associated with attitude control result in adequate airspeed retention, and the aircraft would be found to be in compliance. It may be that the power/airspeed relationship of the aircraft can create adequate cues, where a significant rate of descent is created by a nose down pitch attitude change and a subsequent airspeed increase. In this case, the normal cues associated with altitude retention during fixed power cruise flight may prove to be acceptable for airspeed retention if the evaluation pilot finds that, within the context of the overall flight task, airspeed retention is sufficiently accurate. These altitude change cues may not be usable in autorotation or climb, but may be sufficient in cruise, or V tasks.

NE (iv) Other cues may be found for a specific aircraft, such as small but perceptible changes in noise or vibration. It is not intended that the evaluation pilot search for these cues in order to learn how to maintain airspeed in the aircraft under evaluation. These cues should be perceptible to the typical pilot and sufficient to reinforce the airspeed maintenance task.

Page B – 112.1 AC 29-2C, Chg 4 5/1/2014 AC 29.175. § 29.175 (Amendment 29-24) DEMONSTRATION OF STATIC LONGITUDINAL STABILITY.

a. Explanation .

(1) This rule incorporates the specific flight requirements for demonstration of static longitudinal stability. Specific loadings, configurations, power levels, and speed ranges are stated for conditions of climb, cruise, autorotation, and hover.

(2) Some rotorcraft in forward flight experience significant changes in engine power with changes in airspeed even though collective and throttle controls are held fixed and altitude remains relatively constant. For these cases, the guidance in § 29.173, which states that throttle and collective pitch must be held constant, is appropriate for administration of this rule, and the specified power in § 29.175(a), (b), and (c) should be considered as power established at initial trim conditions. This will result in slightly higher or lower torque readings at “off trim” conditions. Collective and throttle controls are held constant when obtaining data during climb, cruise, and autorotation tests.

(3) The effects of rotor RPM on autorotative static stability should be determined, and positive stability demonstrated for the most critical RPM. For Category A rotorcraft this requirement may be satisfied at a nominal RPM value. RPM values can be expected to change as airspeed is varied from the “trimmed” condition.

Manufacturer’s recommended autorotation airspeed is ordinarily used for trim.

(4) Hovering is considered a flight maneuver for which the pilot repeatedly adjusts collective to maintain an approximately constant altitude above the ground. For hover stability tests, collective and throttle adjustments are made as necessary to maintain an approximately constant height above the ground. Also, a limited amount of negative longitudinal control travel is allowed with changes in speed.

b. Procedures .

(1) Instrumentation.

(i) Sensitive control position instrumentation is mandatory. Engine power parameters should be recorded at trim. For testing of minor modifications or when using a “before and after” method, a tape measure or a stick plotting board may be utilized. A stick plotting board consists of a level surface with a clean sheet of paper on it and attached to the cockpit or seat structure. The installation must not interfere when the flight controls are fully displaced. A recording pencil is attached to the cyclic control by an offsetting arm in such a manner that it can be pushed down on the board to record relative cyclic position at key times during test maneuvers. The figure AC 29.175-1 plot is a typical presentation of longitudinal static stability.

Page B – 112.2 5/1/2014 AC 29-2C, Chg 4 (ii) Other necessary parameters include pitch attitude, pressure altitude, ambient temperature, and indicated airspeed (pace vehicle or theodolite speed for hover tests). For hover tests, hover height (radar altitude if available), and surface winds should be documented. Two-way communications with a pace vehicle is highly desirable. Ground safety equipment is desirable.

(2) Ambient Conditions. Smooth air is necessary for stability testing. Allowable wind conditions for hover stability testing are the same as those for hover controllability tests and are described in that section (paragraph AC 29.151). Extrapolation is covered in paragraph AC 29.53.

(3) Loading. Aft center of gravity (CG) is ordinarily critical for longitudinal stability testing, although high speed flight and hover should be checked at full forward CG and maximum weight. At aft CG, light or heavy weight conditions can be critical.

The manufacturer’s flight data should be reviewed to determine critical loading conditions.

(4) Conducting The Test.

(i) The rotorcraft should be established in the desired configuration and flight condition (climb, cruise, autorotation) with the required power and rotor speed at the trim airspeed. The collective stick should be fixed in that position, usually by applying sufficient friction to insure that it is not inadvertently moved. For autorotative tests, a rotor speed should be selected so that the variations in rotor speed as airspeed and altitude change do not exceed the allowable limits. This point is recorded as the trim point. Airspeed is then increased or decreased in about 10-knot increments, stabilizing on each speed and recording the data. At least two points on each side of the trim speed should be taken.

(ii) The cruise test should be conducted by varying airspeed around the desired altitude with throttle and collective fixed. This should be accomplished by first determining V (level flight speed at maximum continuous power) at the test altitude.

H Then reduce power to establish a level trimmed condition at 0.9 V (or 0.9 V if lower).

H NE This point is then recorded as the trim point.

(iii) For climb and autorotation tests, conduct fixed collective tests through an altitude band (usually ± 2,000 feet), first increasing airspeed as data points are collected, then decreasing speed through the same altitude band. It will probably not be possible to obtain the required data on one pass through the altitude band. If repeated passes are required, a trim point should be taken at the beginning of each pass unless very sensitive collective pitch position information is available in the cockpit. Generally, it will be possible to acquire all the high speed points on one pass and the low speed points on the second.

(iv) If extremely precise results are required, an alternate method of testing can be used to acquire the data at a constant altitude. For cruise, data can be Page B – 113 AC 29-2C, Chg 4 5/1/2014 obtained by alternating airspeeds above and below the trim speed to arrive in the vicinity of the test altitude as the point is recorded. This method results in very precise data because collective and throttle are not moved as airspeed is changed at a constant altitude. A typical sequence of speeds that could produce these results would be: 150 (V ), 135 (0.9V ) trim speed, 125, 145, 115, 155, 105, and 165.

H H (v) For rotorcraft with high rates of climb, a series of climbs, each at a different speed, may be required through a given altitude, utilizing sensitive instrumentation to assure collective position is the same for each data point. In autorotation, a similar case arises and a series of descents, each at a different speed, may be required through a given altitude band, using sensitive instrumentation to assure a repeatable collective position.

(vi) Hover tests should be conducted by maintaining an approximately constant altitude above the ground at the hover height established for performance purposes. The test altitude above the ground may be increased to provide reasonable ground clearance during rearward flight. Groundspeed is varied using a pace vehicle, theodolite, or other velocity measuring equipment. A pace vehicle is an aid in maintaining an accurate hover height. The pilot can accurately maintain height by controlling his sight picture of the pace vehicle (level with the roof, antenna, etc.). Hover stability tests are ordinarily conducted in conjunction with hover controllability tests because instrumentation and facilities are essentially the same.

(vii) Normally climb, cruise, and autorotation tests should be conducted at low, medium, and high altitudes. See paragraph AC 29.45 for guidance on interpolation and extrapolation. High speed stability has been critical during cold weather testing. In two recent models, V at cold temperatures has been limited by the stability NE requirements of § 29.176(b). Cold weather testing should be accomplished or a conservative approach for advancing blade tip Mach number should be used to limit cold weather V to tip Mach number values demonstrated during warm weather NE testing.

(viii) Hover stability should be verified at low altitude and, if required, at high altitude. Refer to paragraph AC 29.45b(2) for guidance on expansion and extrapolation of altitude.

Page B – 114 9/30/99 AC 29-2C O CRUISE □ CLIMB b.◊ AUTOROTATION HOVER SHADED POINTS REPRESENT TRIM t---+---+-----,f---+--+----+-+---+---+-----,f---+--+---+-+---+---+------:;j<i!-~·...---,--, 0<: 0 I 1- ti: ~ f>l•>®W• :J 80 l----l----l- .... f---l---l----1---1----1----1- .... f----l-li.e,,i\>· I •~l ---1---1---1-----!- .... -+---I----I c2 ~ ~~ o 70 l----+---+-----,f---+--+----+-+-----+-,,,....+-----,f---+--+---+-+---+---+------,-+--+----1

8: r=,.,l.!J

~ [•)..,1•·..!.

z 60 1----1----1- .... f---l---l----l--'=L!!J"-+----f--+---l----l--l---l-----l----l---l----l----l--f----t i= V5 ~ 50 1----+---+-----,f---+--+----+-+----+---+-----,f---+--+---+-+---+---+------,-+--+----I (.)

::J (.)

(3 40 l----+--+--:---11--<v[••} +--+---t-----,1,;--A----,,ti.-----,ti•;.-4,-4~----'f'--t-+--+---l-+--+----I _J I t~ ~~ '~ •~ •• ~

~ 30 l--~~----l----+-+---l---l----l--+--+---1----1---l-+---+---l----l--+--+---I---I

:::) t: z 20 .... t--+--+-t--+---+--+----+--+----r--+---+--+--+--+---+--+-t--+---1 _J 10 l----+---+-----,1---+--+----+-+----+---+-----,f---+--+---+-+---+---+------,-+--+----I 0 .__...._ ....... ___ .__..,___,___,,_.,__....___._ ___ ...._..,___,_ ....... _.,__...____,_ ____ ..._......___, 0 20 40 60 80 100 120 140 160 CALIBRATED AIRSPEED (KNOTS) Fl GURE AC 29.175-1 STATIC LONGITUDINAL ST ABILITY Page B – 115 AC 29-2C, Chg 4 9/17/2009 AC 29.175A. § 29.175 (Amendment 29-51) DEMONSTRATION OF STATIC LONGITUDINAL STABILITY.

a. Explanation . Amendment 29-51 reduces the speed range for the climb and cruise demonstration points of §§ 29.175(a) and 29.175(b), respectively. A new paragraph (c) was added to require an additional cruise demonstration point in order to compensate for the change in reduced speed range in paragraph (b). Additionally, for autorotation, two typically used trim points are required in place of the current requirement. The requirement for the hover demonstration was eliminated for the reasons given in AC 29.173 (Amendment 29-51). All the policy material pertaining to this section remains in effect with the following changes: (1) This rule incorporates the specific flight requirements for demonstration of static longitudinal stability. Specific loadings, configurations, power levels, and speed ranges are stated for conditions of climb, cruise, V , and autorotation.

NE (2) Some rotorcraft in forward flight experience significant changes in engine power with changes in airspeed even though collective and throttle controls are held fixed and altitude remains relatively constant. For these cases, the guidance in § 29.173, which states that throttle and collective pitch must be held constant, is appropriate for administration of this rule, and the specified powers in § 29.175 should be considered as power established at initial trim conditions. This will result in slightly higher or lower power readings at “off trim” conditions. Collective and throttle controls are held constant when obtaining test data.

(3) The effects of rotor RPM on autorotative static stability should be determined and positive stability demonstrated for the most critical RPM. For Category A rotorcraft, this requirement may be satisfied at a nominal RPM value. RPM values can be expected to change as airspeed is varied from the “trimmed” condition.

The manufacturer’s recommended autorotation airspeed is ordinarily used for trim.

b. Procedures . All the policy material pertaining to this section remains in effect with the following changes: (1) Instrumentation .

(i) Sensitive control position instrumentation is mandatory. Engine power parameters should be recorded at trim. For testing of minor modifications or when using a “before and after” method, a tape measure or a stick plotting board may be utilized. A stick plotting board consists of a level surface with a clean sheet of paper on it and is attached to the cockpit or seat structure. The installation must not interfere when the flight controls are fully displaced. A recording pencil is attached to the cyclic control by an offsetting arm in such a manner that it can be pushed down on the board to record relative cyclic position at key times during test maneuvers. The Figure AC 29.175A-1 plot is a typical presentation of longitudinal static stability.

Page B – 116 9/17/2009 AC 29-2C, Chg 4 (ii) Other necessary parameters include pitch attitude, pressure altitude, ambient temperature, and indicated airspeed.

(2) Ambient Conditions . Smooth air is necessary for stability testing.

(3) Loading . Aft center of gravity (CG) is ordinarily critical for longitudinal stability testing, although high speed flight should be checked at full forward CG and maximum weight. At aft CG, light or heavy weight conditions can be critical. The manufacturer’s flight data should be reviewed to determine critical loading conditions.

(4) Conducting The Test .

(i) The rotorcraft should be established in the desired configuration and flight condition (climb, cruise, V , autorotation) with the required power and rotor speed NE at the trim airspeed. The collective stick should be fixed in that position; usually by applying sufficient friction to insure that it is not inadvertently moved. For autorotative tests, a rotor speed should be selected so that the variations in rotor speed as airspeed and altitude change do not exceed the allowable limits. This point is recorded as the trim point. Airspeed is then increased or decreased in about 5-knot increments, stabilizing on each speed and recording the data. At least two points on each side of the trim speed should be taken.

(ii) The cruise test should be conducted by varying airspeed around the desired altitude with throttle and collective fixed. This should be accomplished by first determining V (level flight speed at maximum continuous power (MCP)) at the test H altitude. Then adjust power to establish a level trimmed condition at V (or 0.8 V if H NE lower). This point is then recorded as the trim point.

(iii) For climb and autorotation tests, conduct fixed collective tests through an altitude band (usually ±2,000 feet). It will probably not be possible to obtain the required data on one pass through the altitude band. If repeated passes are required, a trim point should be taken at the beginning of each pass unless very sensitive collective pitch position information is available in the cockpit.

(iv) If extremely precise results are required, an alternate method of testing can be used to acquire the data at a constant altitude. For cruise and V , data can be NE obtained by alternating airspeeds above and below the trim speed to arrive in the vicinity of the test altitude as the point is recorded. This method results in very precise data because collective and throttle are not moved as airspeed is changed at a constant altitude. A typical sequence of speeds that could produce these results would be: (0.8 V ) trim speed, 135, 145, 130, and 150.

NE (v) For rotorcraft with high rates of climb, a series of climbs, each at a different speed, may be required through a given altitude, utilizing sensitive instrumentation to assure collective position is the same for each data point. In autorotation, a similar case arises and a series of descents, each at a different speed, Page B – 116.1 AC 29-2C, Chg 4 9/17/2009 may be required through a given altitude band, using sensitive instrumentation to assure a repeatable collective position.

(vi) Normally tests should be conducted at low, medium, and high altitudes.

See AC 29.45 for guidance on interpolation and extrapolation. High speed stability has been critical during cold weather testing. Cold weather testing should be accomplished or a conservative approach for advancing blade tip Mach number should be used to limit cold weather V to tip Mach number values demonstrated.

NE Page B – 116.2 9/17/2009 AC 29-2C, Chg 4 ...

Climb (Vy) Cruise (Vne)

•

■ Criuse (Vh) ♦ Autorotation (Vglide) Autorotation (Vmin RoD)

•

•

,.

.4 ,. ■

-

■ I~ .i.

A.&.

.

.. -

~

- ► T

~ - ♦◄

n• t Longitudinal Cyclic Position (% from Full Aft) 0 20 40 60 80 100 120 140 160 Calibrated Airspeed (Knots) FIGURE AC 29.175A-1: STATIC LONGITUDINAL STABILITY Page B – 116.3 AC 29-2C 9/30/99 AC 29.177. § 29.177 (Amendment 29-24) STATIC DIRECTIONAL STABILITY.

a. Explanation . This rule requires that positive static directional stability be demonstrated at the trim airspeeds defined in § 29.175. The trim speed for climb is V Y and for cruise is 0.9V or 0.9V (whichever is less). For autorotation that airspeed H NE defined by the midpoint of the speed range specified in § 29.175(c) may be used.

b. Procedures .

(1) Tests for static directional stability require instrumentation for pedal position and sideslip angle. Lateral cyclic control position instrumentation should be provided for IFR certification tests. To obtain accurate sideslip angle and airspeed information, a “yaw boom” is usually installed for the purpose of mounting a sideslip vane and swiveling airspeed pitot head outside the main rotor downwash region of influence.

Special care should be taken to ensure that the yaw boom installation has been verified to be structurally adequate and free of dynamic instabilities for all combinations of airspeed and rotor speed likely to be experienced during the static directional evaluation. For some installations, the instrumentation yaw boom may influence the flying qualities of the rotorcraft itself. Thus, it is advisable to correlate yaw string displacement or slip indicator ball widths of skid with yaw boom sideslip angle, and then repeat a few critical points with the yaw boom removed.

(2) For some rotor system designs, the main and tail rotor flapping angle may be a critical instrumentation requirement for static directional testing. Both main and tail rotor flapping may increase dramatically at high airspeeds with increasing sideslip angle. Therefore, for rotor systems exhibiting this characteristic, flapping should be monitored carefully during the sideslip maneuver to avoid exceeding limitations. Static directional stability is normally defined in terms of pedal displacement required to maintain a straight flight path sideslip. A single-rotor rotorcraft flying in coordinated flight will exhibit a small inherent sideslip due to tail rotor thrust and fuselage/main rotor sideforces. This condition is normally taken as trim with the inherent sideslip angle noted. Airspeeds should be the trim values described above. A generally accepted technique follows: (i) Stabilize at the trim point, and note indicated airspeed.

(ii) Record trim conditions including inherent sideslip. Maintain fixed collective and throttle for the remainder of the maneuver.

(iii) Smoothly yaw the aircraft with directional control and coordinate with lateral control to establish the desired sideslip angle. A steady heading can best be ensured by maintaining a track over a straight landmark on the ground such as a section line or straight segment of powerline or highway.

(iv) Note airspeed immediately upon completion of the yaw maneuver.

There may be a small change from the trim airspeed. Fly the new airspeed while Page B – 116.4 9/17/2009 AC 29-2C, Chg 4 maintaining a constant heading, and record indicated airspeed, control positions (directional at a minimum), sideslip angle, rotor speed, rate of descent, amount of ball deflection, and bank angle. The pilot should note the physical sideforce feel experienced. A minimum of two sideslip data points on each side of the trim point should be obtained to adequately define the slope of the pedal displacement versus sideslip angle relationship.

(v) Smoothly return the aircraft to the inherent sideslip angle. Static directional stability plots can be expected to differ slightly on either side of the inherent sideslip angle. Positive static directional stability is indicated by increased left pedal displacement for a larger right sideslip and, conversely, increased right pedal for a larger left sideslip angle.

AC 29.177A. § 29.177 (Amendment 29-51) Static Directional Stability.

a. Explanation . Amendment 29-51 makes an extensive change to the current requirement and provides for a clear definition of the sideslip envelope to be evaluated.

Most rotorcraft exhibit satisfactory quantitative and qualitative directional characteristics except for the first 2-3 degrees either side of trim due to inherent airflow blockage of the vertical fin or tail rotor. This amendment takes this blockage into account while requiring that positive directional stability is maintained at larger sideslip angles. The actual demonstration has been increased from a maximum range of ±10° at all speeds, as the previous amendment requires, to ±25° at slow speeds and linearly decreasing to ±10° at V . Alternatively to the previous range specified, the requirement limits the NE maximum sideslip to be demonstrated to at least 0.1g of sideforce or the maximum sideslip attained when full directional control is applied. As in the previous amendment, sufficient cues should alert the pilot when approaching sideslip limits.

b. Procedures . The policy material pertaining to the procedures outlined in this section remain in effect.

Page B – 116.5 AC 29-2C, Chg 4 5/1/2014 [This page intentionally blank.]

Page B – 116.6 5/1/2014 AC 29-2C, Chg 4 AC 29.181. § 29.181 (Amendment 29-24) DYNAMIC STABILITY: CATEGORY A ROTORCRAFT.

a. Explanation . This section requires that Transport Category A rotorcraft, certificated under Amendment 24 of FAR 29, demonstrate positive damping for short-period oscillations (5 seconds or less) at forward speeds from V to V with the Y NE cyclic, collective and directional controls held in the desired test condition or released by the pilot. This requirement would prevent persistent or divergent short-period oscillations and thus alleviate the pilot workload to actively dampen oscillatory motions for all types of operations.

b. Procedures .

(1) Tests for short period dynamic stability are carried out in the same manner as for IFR (reference AC 29 Appendix B) except the oscillation need not be damped as heavily (i.e., to ½ amplitude in not more than one cycle). Similarly pulses and doublets may be used to generate an upset condition that would be expected to be encountered in moderate turbulence for that particular rotorcraft.

(2) Tests should be conducted at the critical gross weight, altitude, center of gravity, rotor RPM, and power conditions during routine climb, cruise, and descent condition for speeds from V to V . This test must be conducted with the minimum Y NE amount of stability augmentation approved for continued safe flight. Consideration should be given to optional equipment that are to be mounted externally.

(3) This requirement is not applicable to transport category rotorcraft certificated as Category B only. The requirements for this situation are unchanged.

Page B - 117 AC 29-2C 9/30/99 SUBPART B - FLIGHT GROUND AND WATER HANDLING CHARACTERISTICS AC 29.231. § 29.231 GENERAL (GROUND AND WATER HANDLING CHARACTERISTICS).

a. Explanation . The rule states: “The rotorcraft must have satisfactory ground and water handling characteristics, including freedom from uncontrollable tendencies in any condition expected in operation.” In addition, §§ 29.235, 29.239, and 29.241, contain specific requirements concerning ground and water handling characteristic evaluations.

b. Procedures .

(1) During the flight test program and the F&R program (§ 21.35(b)(2)), the rotorcraft will be subjected to evaluations at various weight and CG conditions. Any uncontrollable tendencies found during these test programs must be corrected.

(2) Controllable or damped vibrations or oscillations on the ground or in the water are acceptable, provided the design limits of the rotorcraft are not exceeded.

(3) Any significant vibration or oscillation characteristics found during tests should be described in the test report, and the rotorcraft flight manual should contain appropriate descriptions and procedures to describe and either avoid or handle significant characteristics.

(4) For rotorcraft equipped with wheel gear, the evaluation should include takeoff, landing, and taxi at the maximum airspeed and ground speed CG extremes. If a nose or tail wheel lock/swivel control is installed, each position should be evaluated for limiting takeoff, landing, and taxi speeds. Maximum substantiated speed values should be included in the RFM as limitations.

(5) For water operations, the wave height and frequency or “sea state” should be included as a limitation or, if no limit was reached during testing, the demonstrated values should be placed in the Performance Section of the Rotorcraft Flight Manual.

Information or limits on the allowable “sea state” for rotor startup and shutdown should also be included.

AC 29.235. § 29.235 TAXIING CONDITION .

a. Explanation . The rotorcraft is designed for certain landing load factors (§§ 29.471 and 29.473). The rotorcraft must not attain a load factor in excess of the design load factor when taxied over the roughest ground that may reasonably be expected in normal operation at the expected taxi speeds. This rule applies to wheel landing gear equipped rotorcraft.

Page B - 118 9/30/99 AC 29-2C b. Procedures . The structural substantiation data contains the allowable design limits for the rotorcraft. A calibrated accelerometer or load factor “g” meter should be installed, as near as practicable to the rotorcraft CG, to record the maximum vertical load factor attained. Instrumentation of the landing gear and/or related structure may also be an acceptable means of showing compliance.

(1) Calibrated instrumentation should be installed to record the maximum loads or maximum vertical load factor attained during the taxi tests.

(2) The taxi surface should be evaluated for compliance with the rule.

Corrugated surfaces, as well as broken or uneven surfaces, in accordance with the rule, should be used.

(3) Representative typical taxi speeds, up to the maximum selected by the applicant, should be attained over the selected taxi surfaces.

(4) A light and heavy rotorcraft weight condition should be evaluated.

(5) Limitations appropriate for the rotorcraft design should be included in the flight manual. If these tests indicate that it is unlikely that limit load factors will be attained while taxiing, flight manual limitations may not be necessary.

(6) Pertinent taxi information obtained from these test conditions may be included in normal procedures of the flight manual.

AC 29.239. § 29.239 SPRAY CHARACTERISTICS .

a. Explanation . The intent of this requirement is to evaluate by demonstration that water spray does not obscure visibility (day or night) or damage the rotorcraft during normal waterborne operation (for those rotorcraft which have waterborne or amphibious capability).

b. Procedures .

(1) The following maneuvers should be evaluated in ambient conditions up to the proposed sea state or wave height for operation.

Page B - 119 AC 29-2C 9/30/99 Con- Rotor Alti- fig. Condition Weight CG RPM tude Remarks 1 Taxi Max Optional Max SL Speeds up to maximum proposed for water operation.

2 Hover Max Opt Max - Determine critical hover height, if any.

3 Takeoff Max Opt Max SL Unstick at maximum proposed water operation speed.

4 Land Max Opt Max SL Touchdown at maximum proposed for water operation.

5 Shutdown Opt Opt - SL Shut down the rotorcraft.

6 Start Max Opt Max SL Start engines and release rotor brake.

(2) The maximum sea state or wave height evaluated under this rule should be stated and included in the limitations section of the flight manual.

(3) The effect of saltwater contamination and deterioration of turbine engines and other component parts of the rotorcraft should be considered in accordance with § 29.609 and paragraph AC 29.609. Information on saltwater effect and attendant corrective action should be provided in the flight manual, if appropriate, and in the maintenance manual.

AC 29.241. § 29.241 GROUND RESONANCE .

a. Explanation .

(1) The rule states: “The rotorcraft may have no dangerous tendency to oscillate on the ground with the rotor turning.” This rule is a flight requirement that pertains to demonstrating freedom from dangerous oscillations on the ground. CAR Part 7, predecessor to FAR Part 29, originally contained a “strength requirement,” under § 7.203, requiring ground vibration tests. This test would identify critical vibration frequencies and modes of the rotorcraft. CAR Part 7, Amendment 7-4, effective October 1, 1959, removed this ground vibration requirement because the agency concluded that if any major component has a natural frequency which could be excited by some operating parameter, such a condition would be revealed in the course of other ground and flight tests. The Federal Aviation Administration (FAA) apparently was depending on demonstrations under § 7.131/§ 29.241 and the flight load survey data Page B - 120 9/30/99 AC 29-2C (§ 29.571) to satisfy the objective of the vibration test. However, FAR 29, Amendment 29-3, contained new § 29.663 adding reliability and damping action investigation requirements for ground resonance prevention means. A ground vibration survey was not reinstituted by the adoption of § 29.663. Compliance with § 29.663 does require investigation and substantiation as stated. See paragraph AC 29.663.

(2) “Ground resonance” is a mechanical instability of the aircraft while in contact with the ground, often when partially airborne. Stated another way “ground resonance” is a self-excited mechanical instability that involves coupling between the in-plane motion of the rotor blade and the motion of the rotorcraft as a whole on its landing gear (reference “Aerodynamics of the Helicopter,” Gessow & Myers, page 308).

It is caused by the motion of the blade in the plane of rotation (called in-plane vibration) coupled with a rocking or vertical motion of the aircraft as a whole. The tires, landing gear, and rotor restraint pylon structure act as a spring with a vibration frequency which coincides or couples with the natural in-plane frequency of the blade about a real or effective drag hinge in the plane of rotation. When the frequencies of the two motions (rotor and airframe) approach each other and couple, a violent shaking of the aircraft may occur which, if undamped, could result in the destruction of the rotorcraft.

(3) Ground resonance can occur due to flexibility in the rotor pylon restraint system as well as with landing gear flexibilities. This mode of vibration or resonance can happen in-flight (called air resonance) as well as on the ground and should be addressed in the certification program. The evaluation should include variations in stiffness and damping that could occur in service to the rotor pylon restraints (reference “Ground Vibrations of Helicopters,” M.L. Deutsch, JAS, Vol. 13, No. 5, May 1946). See paragraph AC 29.663 for the investigation of the variations.

(4) Ground resonance may be prevented by placing the first order in-plane vibration frequency above the rotor turning speed.

(5) For such configurations which are not susceptible to ground resonance (first order in-plane frequency above rotor turning speed), a simple rotor RPM run-up and run-down with appropriate cyclic control displacement (i.e., excitation of any inherent vibrations) is adequate demonstration that a ground resonance condition does not exist.

Unhinged “rigid” rotors, such as Bell Helicopter 2 blade designs, are this type of rotor system.

(6) For configurations that are susceptible to ground resonance (i.e., first in-plane frequency is below the rotor turning speed), ground resonance is generally prevented by dampers on the blade, acting in the plane of rotation, dampers on the landing gear (sometimes serving as oleo struts), or proper placement of the landing gear frequencies combined with rotor and/or landing gear dampers.

(7) Elastomeric components (in the rotor pylon support system, possibly in the landing gear, and possibly in the rotor head) are significantly affected by ambient temperature prior to warm-up. Their damping characteristics require thorough Page B - 121 AC 29-2C 9/30/99 investigation for the range of rotorcraft operating environment as noted in paragraph AC 29.663.

b. Procedures .

(1) In operation, the resonance characteristics should be checked during takeoff and landing at zero speed and during run-on landings using various power values. Under all conditions, any oscillations which may be introduced should be damped. However, no instability should occur at any operating condition such as during RPM changes from minimum to maximum and idle to maximum. For rotorcraft with wheel gear, uneven taxi surfaces in conjunction with particular taxi speeds, may excite ground resonance and should be evaluated by taxiing on typical surfaces. This evaluation may be conducted in conjunction with tests of § 29.235.

(2) Slow vertical landings for each configuration are made to establish the touchdown collective pitch angle for each rotor speed. For those aircraft equipped with Stability Augmentation Systems (SAS), all ground resonance investigations should be conducted with SAS on and SAS off. This includes the hovering and running takeoffs and landings, taxi tests, and specific ground resonance tests noted herein.

Consideration should be given to conducting tests in various SAS configurations such as roll channel on, pitch channel off, where such configurations are possible and authorized.

(3) For each rotorcraft configuration tested, the aircraft should be positioned on the ground in flat pitch with the rotor stabilized at the minimum practical rotational speed, or optionally, at a speed shown analytically to have significant margin from indicated resonant conditions. Control system inputs should be used to disturb the system for evaluation of subsequent damping.

(4) For each incremental increase in rotor speed and for each rotor speed setting at increments of collective pitch settings, cyclic and collective inputs should be investigated prior to proceeding to the next rotor speed setting. These inputs should cover the appropriate range and combinations of amplitude and frequency.

(5) Cyclic pitch inputs should be made, either by the pilot through the cyclic stick, or through a signal generating device working in conjunction with the cyclic controls. For each frequency of input, amplitude of the inputs should be increased incrementally and ultimately should be large enough to generate responses representative of normal ground and flight operation on the rotor and support system.

The inputs should continue for a time sufficient to execute five complete counterclockwise circles of the cyclic stick (about neutral) at the selected frequency.

(6) At each amplitude of cyclic input, the excitation frequency should be incrementally increased over the range of the blade in-plane frequency in the fixed system. Rotor speed settings should be increased to 1.05 times the maximum power-on rotor speed. Collective pitch settings should be increased in increments of Page B - 122 9/30/99 AC 29-2C not more than 20 percent to maximum collective or alternately to the collective setting required to become partially airborne (when the cyclic is displaced as noted).

(7) Typically, articulated rotor aircraft have natural frequencies on the blade in lag of approximately 0.3 times the power-on main rotor RPM; soft in-plane rotors have natural frequencies approximately 0.7 times the main rotor RPM. Therefore, for example, for a rotorcraft with an in-plane frequency of 0.3/rev, operating at 300 RPM, and with 6 inches of total lateral cyclic stick displacement, the stick should be rotated for 5 revolutions in a 0.6-inch diameter circle at ((1-.03) x 300 RPM) or 3.5 cycles per second to attempt excitation of possible resonant frequencies. At the conclusion of the excitation, the cyclic stick should be returned to the neutral position while continuing the recording of data listed in paragraph b(13).

(8) The complete program should again be repeated with cyclic excitation inputs from the directional and longitudinal controls, if critical for the type of rotorcraft being evaluated.

(9) If onset of ground resonance is encountered, the typical recommended corrective action is to increase the collective pitch and rotor speed and become airborne. However, lowering the collective pitch has been effective for some designs and is considered a satisfactory procedure if resonance can be consistently avoided.

(10) Landings should be made at the maximum touchdown speed proposed with the rotor speed stabilized.

(11) Special Considerations: (i) The influence of variables including environmental effects, corresponding aircraft component characteristic changes, operational parameters, and surface conditions should be investigated over the ranges proposed for certification.

Additionally, the potential of misservicing and possible failure modes should be evaluated. For ground resonance qualification, where practical, variations from the baseline test configuration may be accomplished by either ground run (§ 29.663(b)) requires investigation of probable ranges of damping), analyses, component tests, aircraft shake test, the specification of special operational procedures in the rotorcraft flight manual, or combination thereof. Detailed and rational analyses showing acceptable correlation to the baseline tests, and for which the input parameters were verified by drawings, calculations, component static or dynamic tests, or by aircraft shake tests simulating the conditions/configurations in question, may be used to limit testing to only those variables and operational conditions showing marginal or unacceptable system damping. All operational limitations should be clearly stated in the rotorcraft flight manual. A report of the analytical and/or test results should be permitted per § 29.663.

(ii) Potential instability while airborne, called “air resonance” may occur due to the dynamic coupling of the rotor flexibility and the pylon restraint flexibility. The Page B - 123 AC 29-2C 9/30/99 same considerations apply to air resonance as to ground resonance except that the pylon restraint variables replace the landing gear variables. Air resonance should be addressed in the certification program.

(iii) When operating on the ground, there may be a tendency for the aircraft to exhibit a “ground bounce.” For many configurations, this is a benign, although undesirable phenomenon which may be aggravated by pilot induced oscillations (PIO), particularly if there is little or no friction on the collective.

(12) On rotorcraft with fully articulated rotor heads equipped with landing gear oleos in either skid or wheel configuration, there are tendencies for ground bounce to occur when light on the oleos, either just prior to takeoff or just after landing contact, or during a power assurance check. This bounce may induce ground resonance, particularly if the intensity of the bounce is aggravated by PIO. The corrective action is either to lift off to a hover or to positively lower the collective and remain on the ground.

(13) Instrumentation and Data Acquisition.

(i) Atmospheric Conditions (to be manually noted): Altitude OAT Wind Velocity (ii) Aircraft Configuration (to be manually noted): Gross Weight C.G.

Tire Pressure Landing Gear Oleo Pressure (iii) Instrumentation (for recording during test).

Main Rotor RPM.

Time history of cyclic control fore-and-aft and lateral stick position Time history of collective control stick position Time history of rotor damper motion* Time history of pylon component motion* Time history of landing gear (oleo) motion* Time history of aircraft motions* *As required to obtain modal damping Page B - 124 9/30/99 AC 29-2C SUBPART B - FLIGHT MISCELLANEOUS FLIGHT REQUIREMENTS AC 29.251. § 29.251 VIBRATION .

a. Explanation .

(1) Each part of the rotorcraft must be free from excessive vibration under each appropriate speed and power condition (rule statement).

(2) This flight requirement may be both a qualitative and quantitative flight evaluation. Section 29.571(a) contains the flight load survey requirement that results in accumulation of vibration quantitative data. Section 29.629 generally requires quantitative data to show freedom from flutter for each part of the rotorcraft including control or stabilizing surfaces and rotors. See paragraphs AC 29.571 and 29.629 for these two rules.

(3) Review Case No. 70 (reference FAA Order 8110.6) contains a policy 0statement concerning compliance with this rule. This policy statement is condensed here for convenience: “The rotorcraft must be capable of attaining a 30° bank angle (turn), at V , NE with maximum continuous power (maximum continuous torque) without encountering excessive roughness/vibration. The FAA/AUTHORITY requires the maneuver demonstration to provide the pilot with some maneuver capability at V , and further to NE provide the pilot some margin away from roughness when operating in turbulence.” (This maneuver may result in a descent or a climb.)

(4) Section 29.1505 pertains to V determination. Section 29.1509 pertains to NE rotor speed limits determination. See paragraphs AC 29.1505 and AC 29.1509.

b. Procedures .

(1) During the company flight test program, the rotorcraft is flown to the appropriate rotor and airspeed limits at several weights to prove that the rotorcraft is free from excessive vibration under appropriate speed, power, and weight conditions.

The flight loads survey quantitative data (reference § 29.571) and the applicant’s qualitative and quantitative flight test data must also prove compliance with the requirement prior to issuing an authorization for official FAA/AUTHORITY flight tests.

(2) The flight load survey data obtained under § 29.571(a) will contain measured data concerning proof of freedom from flutter and excessive vibration.

Pertinent critical flight conditions will be reinvestigated during FAA/AUTHORITY flight tests. The specific condition or conditions necessary to demonstrate compliance with § 29.251 varies with the rotorcraft design, and with the minimum and maximum rotor Page B - 125 AC 29-2C 9/30/99 speeds, V and V speeds, and weight and CG position. An illustration of the speed NE D and RPM demonstration is shown in figure AC 29.251-1. Also see subparagraph b(4).

(3) The airspeed and rotor speed limits investigated and established under §§ 29.33, 29.1503, 29.1505, and 29.1509 are also investigated and made a matter of record in the flight loads survey data. During the official FAA/AUTHORITY/TIA flight tests, critical parts of the rotorcraft may have limited instrumentation to reinvestigate and confirm that the critical conditions investigated during the flight load survey are satisfactory and do not result in excessive vibration. Use of instrumentation is optional if the flight loads data (reference paragraph AC 29.571) are conclusive.

(4) FAA policy for certification (Review Case No. 70) requires a “rotor roughness” flight demonstration of a 30° bank angle left and right, at maximum continuous power (MCP) (maximum continuous torque which may be in excess of the maximum continuous temperature limit), at V . To provide the pilot with some margin NE from roughness, the FAA requires maneuver demonstrations of 30°banked turns at V NE without encountering excessive roughness. The maneuver should be conducted with the rotor speed at the minimum RPM and maximum RPM limits. During the flight load survey, this condition should be investigated and data recorded to assure hazardous loads are not encountered for this “unusual” condition. As indicated, the flight condition will be reinvestigated during the FAA/AUTHORITY flight tests. See paragraph b(2) for illustration of this speed and RPM demonstration.

Page B - 126 9/30/99 AC 29-2C o 0° Bank, Autorotati on !J. Level Forward Flight, 1 05% Max RPM Povver-Off 4 Power-On 30° Bank, Power-On Max RPM Power-Off Max RPM Power-On 7 O t;i.6 ROTOR RPM Min RPM Power-On Min RPM Power-Off 95% Min RPM Power-Off 3 O AIRSPEED 1. Autorotation at 1.11 VNE (AA) at minimum placard rotor speed 2. Autorotation at 1.11 NNE (AA) at ma x imum placard rotor speed .

3. Autorotation at NNE (AR) at power-off minimum design limit rotor speed 4. Autorotation at NNE (AR) at power-off ma x imum design limit rotor speed 5. Forward flight 1. 11 VNE at minimum power-on rotor speed.

6. Forward flight 1. 11 VNE at maximum power-on rotor speed 7. Right and left turn at VNE at ma x imum power-on rotor speed with 30° bank angle .

8. Right and left turn at VNE at minimum power-on rotor speed with 30° bank angle Note : VNE (AA) may be less than VNE .

FIGURE AC 29.251-1 DEMONSTRATION POINTS Page B - 127 9/30/99 AC 29-2C CHAPTER 2. PART 29 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT SUBPART C - STRENGTH REQUIREMENTS GENERAL AC 29.301. § 29.301 LOADS .

a. Explanation .

(1) The rule is a general statement concerning limit and ultimate loads and the application of these loads to the rotorcraft.

(2) Ultimate loads are limit loads multiplied by the prescribed factors of safety.

(3) The specified loads must be distributed appropriately or conservatively and significant changes in distribution of the loads, as a result of deflection, must be taken into account.

b. Procedures . The design criteria report and/or design loads report must contain data that comply with the rule.

AC 29.303. § 29.303 FACTOR OF SAFETY .

a. Explanation .

(1) Unless otherwise provided by FAR Part 29, a factor of safety of 1.5 is required and is applied as stated in the rule. This safety margin will assure that the design strength of the rotorcraft is greater than the design loads contained in FAR Part 29.

(2) Other rules, §§ 29.561(b)(3) and 29.787(c), specify use of defined ultimate inertia forces for protection of occupants.

b. Procedures .

(1) The design criteria report and/or design loads report must contain data that include the appropriate factor of safety.

(2) The factor of safety multiplies the limit external and inertia loads. The rule does allow the application of this factor to the resulting “limit internal” stresses if it is more conservative.

Page C - 1 AC 29-2C 9/30/99 AC 29.305. § 29.305 STRENGTH AND DEFORMATION .

a. Explanation .

(1) This general rule defines, in relative terms, allowable deformation for limit and ultimate loads.

(2) If static tests are used to show compliance with this rule, the structure must support ultimate loads for 3 seconds without failure. Alternatively, dynamic tests simulating actual load applications may be used.

(3) Section 29.307 concerns proof of the structure and requires certain specified tests. This rule also allows substantiation by structural analysis. See paragraph AC 29.307.

b. Procedures . Any test results, static or dynamic, must satisfy the limitations or acceptance criteria contained in the rule.

(1) Any test proposals submitted for approval that are used to demonstrate compliance with sections of FAR Part 29 must contain the criteria stated in the rule.

(2) Any test results reports must contain data and information showing the test results comply with the standard.

(3) When dynamic tests are not used to substantiate the ultimate strength of structure subject to significant dynamic response under load, the analytical substantiation should consider flexibility effects and rate of load application (tail boom strength under landing loads is an example of a strength which needs dynamic amplification effects considered).

AC 29.307. § 29.307 (Amendment 29-4) PROOF OF STRUCTURE .

a. Explanation .

(1) The rule requires compliance for each critical loading condition. Certain tests must be conducted as specified. Additional tests for new or unusual design features may be required as noted in § 29.307(b)(6).

(2) “Structural analysis may be used only if the structure conforms to those for which experience has shown the structural analysis method to be reliable.” (3) Fatigue substantiation requirements are explained further in paragraph AC 29.571.

b. Procedures.

Page C - 2 9/30/99 AC 29-2C (1) The design criteria and/or design loads report should contain typical or representative loading conditions from which the critical loading conditions will be selected for analytical substantiation in structural (static and fatigue) reports and dynamics (vibration and stability) reports and fatigue, static, dynamic, or operational test reports.

(2) Whenever tests are used or required, a test proposal or plan must be approved prior to the tests. The test article must have received conformity inspections and must have been accepted by the FAA/AUTHORITY for the test. Test fixtures and instrumentation must also be acceptable to the FAA/AUTHORITY (using DERs as appropriate) prior to the start of the test. The quality control office of the applicant or other qualified personnel may be authorized to conduct inspections of the test fixtures and instrumentation rather than the FAA/AUTHORITY or DER performing this task. The test proposal may be used to define and to authorize the means to accomplish inspection of the test fixtures and instrumentation. Unnecessary drawings, such as test fixture details, or layering of approvals is not intended or envisaged by this policy.

Drawings, sketches, or photographs have been used by the FAA/AUTHORITY to control and to assure correct location, direction, and magnitude of loads and other critical test parameters.

(3) Structural analysis has been accepted for rotorcraft in place of static tests.

Generally the rotorcraft airframe should have frequency placements remote to predominate rotor excitation sources, including rotor harmonics, to avoid undesirable and possibly excessive vibration and potentially high operating stress levels due to this vibration. During the flight load measurement program conducted under § 29.571, critical loaded areas or critical joints may be instrumented with strain gages or other stress strain measuring devices. This actual flight data may be compared to the analytical data to verify accuracy.

(4) Section 29.307(b) specifies certain tests. Test proposals must be approved prior to conducting official FAA/AUTHORITY tests. Other paragraphs in this advisory circular pertain to those tests.

AC 29.307A. § 29.307 (Amendment 29-30) PROOF OF STRUCTURE .

a. Explanation . Amendment 29-30 adds the requirement to account for the environment to which the structure will be exposed in operation. This change is intended to codify recent FAA/AUTHORITY and industry practices for the consideration of environmental effects in showing “proof of structure.” b. Procedures . All of the policy material pertaining to this section remains in effect with the following additions: (1) For either tests or an analysis, environmental effects are now explicitly required. Consideration of loss of strength and stiffness of metals with elevated temperatures and loss of strength and stiffness of composite materials from exposure to Page C - 3 AC 29-2C 9/30/99 heat, moisture, or other operational environments is now required and should be documented in analyses and test reports.

(2) MIL-HDBK-5F, AC 20-107B, or MIL-HDBK-17B, Vol I, Rev. 1E; Vol. II, Rev.

D; Vol. III, Rev. E (or later versions) are acceptable sources of data and procedures to show compliance with environmental effects of metallic and composite materials, respectively.

AC 29.309. § 29.309 DESIGN LIMITATIONS .

a. Explanation .

(1) The rule requires an orderly selection and presentation of the basic structural design limitations of the rotorcraft. The applicant must establish these structural limitations to facilitate design of the rotorcraft.

(2) Refer to the rule for the specific requirements.

b. Procedures .

(1) The design criteria and/or design load report should contain the design limits specified.

(2) These items are structural design limits. Other requirements may result in narrowing the ranges of type design limits or in reducing limits. It is not necessary to revise structural design criteria limits to agree with more conservative operational limits established during the certification program. The operational limits may be subsequently expanded by additional flight tests to agree with design limits.

Page C - 4 9/30/99 AC 29-2 SUBPART C - STRENGTH REQUIREMENTS FLIGHT LOADS AC 29.321. § 29.321 GENERAL - FLIGHT LOADS.

a. Explanation .

(1) The rule specifies the way the loads will be applied to the rotorcraft. It requires load analysis from minimum to maximum design weight. Any practical distribution of disposable loads must be included in the analysis.

(2) Paragraph (a) of the rule states: “The flight load factor must be assumed to act normal to the longitudinal axis of the rotorcraft and to be equal in magnitude and opposite in direction to the rotorcraft inertia load factor at the center of gravity.” b. Procedures .

(1) Derivation of the flight loads is required by and specified in § 29.337 through § 29.351. This rule requires flight load determination from minimum to maximum weight and for disposable loads.

(2) The application of the design loads derived from the flight load factor will be as specified. The flight loads analysis data must comply with the rule.

AC 29.337. § 29.337 (Amendment 29-30) LIMIT MANEUVERING LOAD FACTOR.

a. Explanation . The rotorcraft must be designed and substantiated to load factors as specified to provide a minimum level of structural integrity of the rotorcraft airframe and rotors.

(1) A range of design positive load factors from +3.5 to +2.0 may be used.

(2) A range of design negative load factors from -1.0 to -0.5 may be used.

(3) Load factors inside the range of +3.5 to -1.0 may be used provided the probability of exceeding the design load factors is shown by analysis and flight tests to be extremely remote, and the selected load factors are appropriate to each weight condition between design maximum and minimum weights.

(4) Load factors exceeding these “minimums” may be used.

b. Procedures .

(1) The applicant may elect to substantiate the rotorcraft for a design maneuvering load factor less than +3.5 and more than -1.0. Whenever this option is Page C - 5 AC 29-2C, Chg 4 5/1/2014 used, an analytical study and flight demonstration are required. Maximum available rotor lift with both power on and power off must be considered when substantiating maneuver load factors less than the specified values.

(i) The maximum positive design load factor is +3.5 generally at a weight below maximum gross weight. The maximum thrust capability of the main rotor combined with incremental lift of wings or sponsons, if installed, results in a maximum design positive load factor. An example of a load factor - gross weight curve is shown in figure AC 29.337-1. Note the minimum positive design load factor is +2.0 even though the required analysis and flight demonstration may prove the rotorcraft is not capable of achieving this load factor. This curve also illustrates compliance with § 29.337(b)(2) since the design load factor varies with gross weight.

(ii) The largest negative design load factor is -1.0; however, several current rotorcraft designs are not capable of achieving a negative load factor.

Therefore, -0.5 has been an acceptable structural design negative load factor for certain rotorcraft designs.

(2) Whenever the applicant analytically substantiates the lower load factors allowed by § 29.337(b), the applicant must conduct the flight demonstration required by § 29.337(b)(1). The flight test personnel must determine that the demonstration is conducted in a manner to show that the probability of exceeding the selected design load factors, (those factors less than +3.5 and more than -1.0) is extremely remote.

(3) A numerical value has not been assigned to “extremely remote” in this standard.

Page C - 6 9/30/99 AC 29-2C +3.5------ Represents curve of ~ thrust+ lift Design Load Factor +--+2.0 Nz -1.0--,

Minimum __j

\

Gross Weight Lbs.

Maximum LOAD FACTOR GROSS WEIGHT CURVE FIGURE AC 29.337-1 Page C - 7 AC 29-2C 9/30/99 AC 29.339. § 29.339 RESULTANT LIMIT MANEUVERING LOADS .

a. Explanation . The rule specifies or defines the application of rotor and lift surface loads to the rotorcraft.

(1) The design maneuvering load factors required by § 29.337 will result in or be derived from rotor thrust or lift and from auxiliary surface lift.

(2) The rules §§ 29.321, 29.337, 29.341, and 29.351 all complement one another and result in the derivation of design flight loads that will be imposed to assure structural integrity of the rotorcraft.

(3) The following assumptions and conditions are specified in the rule.

(i) The rule requires application of appropriate loads at each rotor hub and auxiliary lifting surface.

(ii) Power-on and power-off flight with maximum design rotor tip speed ratio and specific conditions that must be considered.

(iii) Rotor tip speed ratio, defined in the rule, has been carried forward from the initial rotorcraft certification rules issued in 1946. The rotor tip speed ratio is a basic parameter used in calculating rotor aerodynamic forces.

b. Procedures .

(1) The rule specifies an acceptable assumption concerning application of the rotorcraft maneuvering loads.

(2) The rotor tip speed ratio is a parameter found in textbooks and other books such as NACA Report No. 716. The equation in the rule contains angle, “a.” Report No. 716 also defines angle, “a,” as the angle of attack of the rotor disk. This definition is more easily understood than the definition contained in the rule.

(3) The rotorcraft design loads are derived as prescribed by §§ 29.321, 29.337, 29.341, and 29.351. These loads are applied to the rotor or rotors and any auxiliary surface as prescribed by this rule.

AC 29.341. §29.341 GUST LOADS .

a. Explanation .

(1) The rotorcraft must be substantiated for the loads derived from 30 feet per second vertical and horizontal gusts from hovering to 1.11 V ; i.e., (V ).

NE D Page C - 8 9/30/99 AC 29-2C (2) Gust loads for any vertical stabilizing surface should be derived for lateral or sideward gusts, as well as the head-on horizontal gusts. See paragraph AC 29.413, § 29.413(a)(2).

(3) Gust loads for any horizontal stabilizing surface should be derived for vertical gusts, upward and downward, as well as for head-on gusts. See paragraph AC 29.413.

b. Procedures .

(1) Either sharp-edged (instantaneous) gusts or sharp-edged gusts modified by an alleviation (attenuation) factor may be used for calculating aerodynamic loads for the rotorcraft and any installed stabilizing surfaces. The following conditions may be used: (i) Vertical gusts may be considered normal to the flight path of the rotorcraft except during hover or low speed flight (20 knots or less) when the gusts may be assumed normal to the longitudinal axis of the rotorcraft.

(ii) For a vertical stabilizing surface, the horizontal gusts are normal to the flight path of the rotorcraft except during hover or low speed flight when the gusts may be assumed normal to the longitudinal axis of the rotorcraft.

(iii) A primary effect of encountering the gust is to change the lift of the rotors and rotorcraft surfaces. Of primary concern is the gust load or lift created by the main rotor or rotors. The lift increment of the horizontal stabilizing surface and fuselage are generally negligible when compared to the rotor and may be neglected for the rotorcraft gust load determination if proven negligible by analysis.

(iv) The rotorcraft shall be assumed in stabilized level flight prior to meeting the gust.

(v) The gust velocity may be assumed uniform across the rotorcraft.

(vi) Gust loads on the stabilizing surfaces are required as stated in paragraph AC 29.413.

(2) The rotorcraft design maneuvering load factors may generally exceed the design gust load factors calculated in compliance with this rule. This may be attributed to the small incremental change in lift due to the 30 FPS gust. Nonetheless, design gust loads for the rotorcraft shall be calculated as specified in the rule to assure the rotorcraft maneuvering load factors do, in each case, exceed the design gust load factor.

(3) For further information about rotorcraft gust response characteristics, see Paper No. 9 presented at the AHS/NASA - Ames Specialist’s Meeting on Rotorcraft Dynamics, February 13-15, 1974. The paper, entitled, “Helicopter Gust Response Page C - 9 AC 29-2C 9/30/99 Characteristics Including Unsteady Aerodynamics Stall Effects,” was written by P.J.

Arcidiacono, R.R. Berquist, and W.T. Alexander, Jr. References listed in the paper may be helpful also.

AC 29.351. § 29.351 YAWING CONDITIONS .

a. Explanation . The rule requires proof of a rotorcraft “structural” yaw or sideslip design envelope. This sideslip envelope must cover minimum forward speed or hover to V or V , whichever is less. The rotorcraft must be structurally safe for the thrust H NE capability of the directional control system.

(1) The rotorcraft structure must be designed to withstand the loads for the specified yaw conditions. The standard does not require a structural flight demonstration. It is a structural design standard.

(2) Maximum displacement of the directional control, except as limited by pilot effort (130 pounds; § 29.397(a)), is required for the conditions cited in the rule. A control system rate limiter or a yaw damper may be used. The total displacement is therefore a function of time as well as the maximum effort applied (130 pounds).

(i) At low airspeeds, 90° yaw (sideward flight) should be the design limit.

(ii) At high airspeeds, stabilized yaw angle (stabilized sideslip) must be substantiated as stated in the rule.

(iii) At high airspeeds, the maximum tail rotor thrust will be combined with the vertical (directional) stabilizer surface load, if a stabilizer is used, as specified by § 29.351(b)(1).

(iv) At high airspeeds, while the rotorcraft is in the sideslip condition, the directional control is then returned to the neutral position, attendant with the flight condition. The tail rotor thrust will be added to the restoring force of the vertical stabilizer.

(v) Both right and left yaw conditions should be proven.

(3) The tail rotor attachment structure must comply with § 29.403.

(4) The vertical stabilizing surface must also comply with § 29.413.

b. Procedures .

(1) Many of the current single main rotor rotorcraft designs have vertical (directional) stabilizing surfaces. These surfaces may be solely vertical stabilizing fins as on the Bell Model 206, or a swept vertical extension of the tail boom as on the Hiller Page C - 10 9/30/99 AC 29-2C Model FH1100. The Hiller FH1100 tail surface houses the tail rotor drive shaft and the tail rotor output gearbox.

(i) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust.

(ii) For vertical stabilizers that house the tail rotor output gearbox, such as the Hiller Model FH1100, the tail surface air loads will add to or subtract from the tail rotor thrust according to the flight condition under consideration.

NOTE: For one example: At stabilized yaw to the right (left pedal depressed to limit) (§ 29.351(b)(2)), the tail rotor thrust moment should equal the restoring moment of the tail boom, vertical stabilizer and main rotor torque. As stated by § 29.351(b)(3), the tail rotor thrust moment then is added to the vertical stabilizer restoring moment. The addition of tail rotor thrust (§ 29.351(b)(3)) and vertical stabilizer load is generally one of the critical design conditions for the fuselage/tail boom.

(iii) For vertical stabilizers or fins that have an offset incidence angle with respect to the rotorcraft axis, the vertical fin moment is added, or subtracted as applicable, to the tail rotor thrust moment. The condition stated in § 29.351(b)(1) may result in adding the fin load to the tail rotor thrust.

(iv) Low airspeed maneuvers, such as sideward, rearward, and hover turns over a spot, typically impose insignificant aerodynamic loads on the fuselage and/or tail boom. The aerodynamic loads at V or V , whichever is required, are H NE generally the significant aerodynamic design loads.

(v) A rational assessment of the various yaw conditions may be used to reduce the load deviation and analysis to the critical rotorcraft design conditions.

(vi) The rotorcraft structure shall be analyzed or tested for loads derived from the critical design conditions.

(vii) A simple structural design envelope may be derived from these design data. If the right or left yaw limits are not very different, common, conservative design limits may be used. A sample yaw/forward speed diagram, as derived from design analysis of the characteristics of a hypothetical rotorcraft, is presented in figure AC 29.351-1. A table of values would also suffice. This figure reflects characteristics which include a 90° yaw when the directional control inputs are applied at low airspeeds (up to 30 knots presumably the maximum sideward flight speed of which this aircraft is capable) and 10° yaw when they are applied at V , with a straight H line variation from 30 knot forward speed to V .

H Page C - 11 AC 29-2C 9/30/99 ✓ -go• go• ,-----------.:::_ / 10° MAX.

YAW DEGREES AIRSPEED 30 KNOT LESSER OF SIDEWARD VH OR VNE FLIGHT FIGURE AC 29.351-1 SAMPLEY AW/FORWARD SPEED DIAGRAM Page C - 12 9/30/99 AC 29-2C (viii) During flight test evaluations, yaw angles have been measured using a yaw angle probe (swiveling vane type) on a nose boom. Both a visual readout for the pilot and a record, such as an oscillograph trace, have been used. This test may be conducted in the flight test program or in the flight load survey program. This record should confirm the yaw angle used in design as conservative with respect to operational and actual flight characteristics. This test is not a requirement however.

AC 29.351A. § 29.351 (Amendment 29-30) YAWING CONDITIONS .

a. Explanation . Amendment 29-30 adds maximum sideslip angles to the existing § 29.351 for structural design purposes. The standard should apply to power-on conditions; not power-off, since V is a part of the standard. For airspeeds up to 0.6 H V , sideslip angles larger than 90° (or sideward flight) need not be considered. For NE airspeeds at V or V (whichever is less), sideslip angles larger than 15° need not be NE H considered.

b. Procedures .

(1) All of the policy material pertaining to this section remains in effect with the addition of the maximum sideslip limits of 90° and 15° specified above. The rotorcraft does not need to be capable of attaining these conditions. A revised yaw/forward speed diagram is presented in figure AC 29.351A-1.

(2) FAR § 29.351(b)(1) incorrectly references § 29.395(a) for maximum pilot forces. The correct reference should be § 29.397(a).

Page C - 13 AC 29-2C, Chg 1 4/25/06 go• YAW DEGREES (LON GI TUDI N AL AXIS) +-15° MAX.

AIRSPEED LESSER OF VH OR VNE FIGURE AC 29.351A-1 SAMPLE YAW/FORWARD SPEED DIAGRAM Page C - 14 4/25/06 AC 29-2C, Chg 2 AC 29.351B. § 29.351 (Amendment 29-40) YAWING CONDITIONS .

a. Definitions .

(1) Suddenly . For the purpose of this section, ‘suddenly’ is defined as an interval not to exceed 0.2 seconds for complete control input. A rational analysis may be used to substantiate an alternative value.

(2) Zero Yaw . Normal, 1-g, level flight condition with either zero bank angle or zero sideslip.

b. Explanation . The rule requires a rotorcraft “structural” yaw or sideslip design envelope. This sideslip envelope must cover minimum forward speed, or hover, to V or V , whichever is less. The rotorcraft must be structurally H NE safe for the thrust capability of the directional control system.

(1) The rotorcraft structure must be designed to withstand the loads for the specified yawing conditions. The standard does not require a structural flight demonstration. It is a structural design standard.

(2) This standard applies only to power-on conditions. Autorotations need not be considered.

(3) This standard requires the maximum allowable rotor RPM consistent with the flight conditions, including special operational rotor settings.

(4) For the purposes of this section, the analysis may be performed at international standard atmosphere (ISA) sea level conditions.

(5) The rotorcraft structure must be designed to withstand the loads for the specified sideslip conditions. This includes, but is not limited to: (i) Main cabin, tailboom, and vertical control surfaces.

(ii) Tail rotor structures, including the fitting attachments to the frame.

(iii) Windows, doors, and other transparencies.

(iv) Landing gear and retracting mechanism.

(v) Fairings and cowlings.

(6) Maximum displacement of the directional control, except as limited by pilot effort (§ 29.397(a)), is required for the conditions cited in the rule. Control system limiting devices may be used, however the probability of failure or malfunction of these system(s) should be considered (see Figure AC 29.351B-2). This evaluation may Page C - 15 AC 29-2C, Chg 2 4/25/06 include Flight Manual Limitations, if failure of the system is reliably indicated to the crew.

(7) Both right and left yaw conditions should be evaluated.

(8) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust (superpositioning).

(9) Loads associated with sideslip angles exceeding the values of Figure AC 29.351B-1 do not need to be considered. The corresponding points of the maneuver may be deleted.

c. Procedure . The design loads should be evaluated within the limits of Figure AC 29.351B-1 or the maximum capability of the rotorcraft, whichever is less; at speeds from zero to V or V , whichever is less, for the following phases of the H NE maneuver: (1) With the rotorcraft at an initial trim condition (1 g level flight and zero yaw), the cockpit directional control is suddenly displaced to the maximum deflection limited by the control stops or by the maximum pilot force specified in § 29.397(a). This is intended to generate a high tail rotor thrust.

(2) While maintaining maximum cockpit directional control deflection, within the limitation specified in c(1) of this AC paragraph, allow the rotorcraft to yaw to the maximum transient sideslip angle or to the value defined in Figure AC 29.351B-1, whichever is less. This is intended to generate high aerodynamic loads.

(3) Allow the rotorcraft to stabilize at the maximum steady-state sideslip angle.

In the event that the maximum steady state angle is greater than the value defined in Figure AC 29.351B-1, the rotorcraft should be trimmed to the value of the angle using less than maximum cockpit directional control deflection.

(4) With the rotorcraft yawed to the static equilibrium sideslip angle specified in c(3) of this AC paragraph, the cockpit control is suddenly returned to its initial trim position. This is intended to combine a high tail rotor thrust and high aerodynamic restoring forces.

Page C - 16 4/25/06 AC 29-2C, Chg 2 SIDESLIP ANGLE ( deg. ) 0.6 V V or V 0 ne h ne whichever is less AIRSPEED FIGURE AC 29.351B-1 YAW/FORWARD SPEED DIAGRAM Page C - 17 AC 29-2C, Chg 2 4/25/06 • For static strength substantiation, each part of the structure should be able to withstand without failure, the loads generated by the maneuver described in the rule multiplied by a factor of safety depending on the probability of being in this failure state. The factor of safety is defined in the figure below: FS 1 .5 -~ 1 .0

I

10-9 Qj - Probability of being in failure condition j Q = (T )(P ) j j j where: T = Average flight time spent with a failed control limiting system j (in hours) j P = Probability of occurrence of failure of the control limiting system j j (per hour) -3 Note : If P is greater than 10 per flight hour then a 1.5 factor of safety should be j applied to all limit load conditions specified in this standard.

FIGURE AC 29.351B-2 Safety Factors for Probability of Failure AC 29.361. § 29.361 (Amendment 29-26) ENGINE TORQUE .

a. Explanation .

(1) The rotorcraft should be designed for limit engine torque values, as prescribed by the rule, to account for maximum engine torque, including certain transients and torsional oscillations. The rule recognized that reciprocating (piston) engines generate higher torque oscillations than turbine engines.

(i) A factor of 1.25 applies to maximum continuous power for turbine engines. Section 29.923 refers to torque output and § 29.927(b) refers to other torque output conditions for use in an “endurance test.” Page C - 18 9/30/99 AC 29-2C (ii) Torque factors are also specified for reciprocating engines having two or more cylinders in § 29.361(a)(2) or § 29.361(b) of Amendment 29-26. The appropriate torque factor applies to takeoff power torque as well as maximum continuous power and other power conditions.

(2) Amendment 29-26 introduced additional turbine engine installation considerations for the following: (i) Engine torque loads associated with emergency operation of governor-controlled turboshaft engines.

(ii) Torque reaction loads from sudden turbine engine stoppage which is applied to the engine and the engine suspension and restraint system.

(3) Paragraph AC 29.549 concerns § 29.549(c) and (e) that contains design standards for engine mounts and adjacent structure for flight and landing and also flight with 2 ½-minute OEI power rating. Amendment 29-26 added OEI power to the standard.

(4) Section 29.547(e)(1)(ii) concerns the application of limit engine torque to design of the main rotor structure.

b. Procedures .

(1) The engine torque associated with the maximum continuous power condition should be multiplied by the appropriate torque factor to obtain the engine torque value used for structural substantiation purposes of the rotorcraft.

(2) The torque values associated with the minimum power-on RPM limit should be used. Maximum power-on speed limit will result in a lower torque value when calculating torque from design horsepower values. However, due to piston engine power output characteristics, an engine may produce a higher torque at higher engine speeds contrary to the previous statement. The torque factor should account for this characteristic.

(3) For turbine engines limit torque values are determined for the four cases cited. Two cases are related to “endurance” test standards.

(4) For sudden stoppage of turbine engines the engine manufacturer can reasonably provide engine rotating inertia and deceleration time expected in the event of sudden engine stoppage which generates these critical loads in the engine mounting and restraint system. These manufacturer’s data should be acceptable for use in complying with this part of the design standard.

Page C - 19 AC 29-2C 9/30/99 SUBPART C - STRENGTH REQUIREMENTS CONTROL SURFACE AND SYSTEM LOADS AC 29.391. § 29.391 CONTROL SURFACE AND SYSTEM LOADS - GENERAL .

a. Explanation . This general rule concerns requirements for design loads of tail rotors, control or stabilizing surfaces, and their control system.

b. Procedures . The design criteria and/or the design loads report must contain the loads dictated by the referenced rules. See paragraphs AC 29.395, AC 29.397, AC 29.399, AC 29.401, AC 29.403, AC 29.411, and AC 29.413.

AC 29.391A. § 29.391 (Amendment 29-30) CONTROL SURFACE AND SYSTEM LOADS - GENERAL .

a. Explanation . Amendment 29-30 adds an explicit reference to § 29.427, Unsymmetrical Loads (paragraph AC 29.427), to clarify that substantiation for unsymmetrical loads is a general control surface requirement. A reference to § 29.399, Dual Control System (paragraph AC 29.399), is also added for clarification. In addition, §§ 29.401, 29.403, 29.413 were removed by this amendment since these references and requirements were adequately addressed in other standards.

b. Procedures . The referenced AC paragraphs become AC 29.395, AC 29.395A, AC 29.397, AC 29.399, AC 29.411, and AC 29.427.

AC 29.395. § 29.395 CONTROL SYSTEM .

a. Explanation . Control system design loads and the application of these loads are contained in this rule.

(1) Paragraph (a) of the rule specifies the way or means of reacting the design loads specified in §§ 29.397 and 29.399 (for dual control systems). The design loads must be imposed on any locks and stops and irreversible mechanisms in the control system. Both rotor blade horns and control surface horns must react without failure, the specified loads while the controls are in critical positions.

(2) Paragraph (b) of the rule specifies application of limit pilot forces or of the maximum loads that can be obtained in normal operation, including any single power boost system failure, whichever is greater. However, minimum limit pilot force 0.60 of the loads specified in §§ 29.397 and 29.399, may be used, as specified, in parts of the primary control system that are not stiff enough to react to the loads specified in the first part of Paragraph (b) of the rule. Note the objective for a rugged control system.

Page C - 20 9/30/99 AC 29-2C (3) Control system design feature and test requirements are found in §§ 29.671 through 29.695. Bearing factors and fitting factors are specified in §§ 29.623 and 29.625, respectively.

b. Procedures .

(1) The design criteria and/or a design loads report that includes the primary control system design loads should be submitted for FAA/AUTHORITY approval.

(2) The rotorcraft control system may be tested to ultimate design loads or may be analyzed for the ultimate design loads. See paragraph AC 29.307.

(i) It is advisable that the applicant prepare a proposal describing the procedures and techniques to be used in the static testing of the control system which reflects compliance with the condition specified. It is further advisable that the FAA/AUTHORITY concur that the tests proposed achieve that objective. Omission of these steps may result in the need for retesting. The test results should be documented.

(ii) If tests are not conducted, a structural analysis of the control system is required. Appropriate factors from §§ 29.685(e), 29.623, and 29.625 must be used as specified. A structural analysis report should be used to document compliance with § 29.685(d)(1) and (4), and § 29.685(f).

(3) If a part of the control system is not stiff or rigid enough to react the design loads specified in § 29.397, that part of the system may be substantiated for lower loads as prescribed.

(i) The limit design loads are those loads specified in § 29.397; (ii) The limit design loads are the maximum that can be obtained in normal operation, including any single power boost system failure, except for objectives stated for a rugged system; and (iii) In lieu of a rational analysis, the limit design loads may be 0.60 of the loads specified in § 29.397.

(iv) For example, if a control surface servo tab or a small elevator is a part of the rotorcraft design, the control system for this part must be stiff enough to react the control surface loads without failure and to provide enough surface deflection to control the rotorcraft. These limit loads may be 60 pounds fore and aft and 40 pounds laterally on the cyclic control stick in lieu of a rational analysis and may be the maximum loads that can be obtained in normal operation.

(v) If a hydraulic power actuation or boost system is part of the rotorcraft design, the design limit load for the affected parts of the control system will be the Page C - 21 AC 29-2C 9/30/99 maximum output force of the boost at normal operating pressure added to the limit design loads resulting from the loads specified in § 29.397. If a single failure in the power portion of the hydraulic system results in actuator forces that exceed the maximum output force at normal operating pressure, the highest output loads must be used as noted in subparagraph (3)(ii). This hydraulic system failure standard is specified in § 29.695(a)(1) as well.

(4) Controls proof and operation test is required by §§ 29.307(b), 29.681, and 29.683. This test is conducted using the design limit loads approved under § 29.395(b).

See paragraphs AC 29.681 and AC 29.683.

AC 29.395A. § 29.395 (Amendment 29-30) CONTROL SYSTEM .

a. Explanation . Amendment 29-30 clarifies that the loads in § 29.395(b) apply to power “control” systems not just power “boost” systems; and the limit pilot forces prescribed in § 29.397 are required to be applied in conjunction with the forces from normally energized power devices. The amendment may increase required loads for systems if operational loads may be exceeded through jamming, ground gusts, control inertia, or friction. If so, the system is required to withstand 100 percent of limit pilot forces specified in § 29.397, rather than 60 percent of the limit pilot forces as specified previously.

b. Procedures . The procedures of paragraph AC 29.395 continue to apply except that the increased loads in new paragraph § 29.355(b)(4) of 100 percent of limit pilot forces are specified for systems where operational loads may be exceeded by jamming, ground gusts, control inertia, or friction.

AC 29.397. § 29.397 (Amendment 29-12) LIMIT PILOT FORCES AND TORQUES .

a. Explanation . Design forces are contained in the rule.

(1) Primary controls, pilot and copilot, must be designed for the limit pilot forces specified in paragraph (a) of the rule.

(2) For other operating controls, such as flap, tab, stabilizer, rotor brake, and landing gear, design limit forces are specified in paragraph (b) of the rule.

b. Procedures .

(1) Design loads specified in the rule must be used in required structural tests and in any structural strength analysis of the control systems submitted in compliance with other rules.

(2) Operation tests of the control systems noted in other rules require application of these forces also.

Page C - 22 9/30/99 AC 29-2C AC 29.399. § 29.399 DUAL CONTROL SYSTEM .

a. Explanation . Design limit loads are specified for dual control systems. Pilot effort forces applied in opposition and in the same direction are required for dual control systems.

b. Procedures .

(1) Design loads specified in the rule must be used in required structural tests and in any structural strength analysis submitted for compliance with the other rules.

(2) Operation tests of the control systems, noted in other rules, require application of these forces also.

AC 29.401. § 29.401 (Amendment 29-4) AUXILIARY ROTOR ASSEMBLIES .

a. Explanation .

(1) For rotorcraft equipped with auxiliary rotors, normally called tail rotors, an endurance test is required by § 29.923 and structural strength substantiation is required.

Section 29.401(b) specifically refers to structural strength substantiation for centrifugal loads resulting from maximum design rotor RPM. Due to the pitch feathering requirements, auxiliary rotors typically have detachable blades.

(2) The rotor blade structure must have sufficient strength to withstand not only aerodynamic loads generated on the blade surface, but also inertial loads arising from centrifugal, coriolis, gyroscopic, and vibratory effects produced by this blade movement.

Sufficient stiffness and rigidity must be designed into the blades to prevent excessive deformation and to assure that the blades will maintain the desired aerodynamic characteristics. As a design objective, the structural strength requirements should be met with the minimum material. Excess blade weight imposes extra centrifugal loads that may increase the operating stress levels. Blade weight and strength should be optimized. Even though a structural strength analysis for the blade design loads is required, a flight load survey and fatigue analysis are also required by § 29.571.

(3) Section 29.1509 defines the design rotor speed as that providing a 5 percent margin beyond the rotor operating speed limits.

b. Procedures .

(1) The endurance tests prescribed by §§ 29.923 and 29.927 require achieving certain speeds, power, and control displacement for the auxiliary (tail) rotor as well as the main rotor. The parts must be serviceable at the conclusion of the tests.

Page C - 23 AC 29-2C 9/30/99 (2) Structural substantiation of the auxiliary (tail) rotor is required to assure integrity for the minimum and maximum design rotor speeds and the maximum design rotor thrust in the positive and negative direction. Thrust capability of the rotor should offset the main rotor torque at maximum power as required by § 29.927(b).

(i) The maximum and minimum operating rotor speed, power-off, is 95 percent of the maximum design speed and is 105 percent of the minimum design speed, respectively.

(ii) The rotor operating speed limits shown during the official FAA/AUTHORITY flight tests must include the noted 5 percent margin with respect to the design speeds.

(iii) The auxiliary rotor generally has a positive and negative pitch limit that assures adequate directional control throughout the operating range of the rotorcraft.

The power-off rotor speed limits are generally broader than the power-on rotor speed limits because of the required autorotational rotor speed characteristics. Thus, the auxiliary rotor design conditions concern the maximum and minimum design rotor speeds in conjunction with the maximum positive or negative pitch thrust as appropriate.

Thrust capability and precone angle of the rotor, if any, will significantly influence the rotor design loads. The variations in rotor design features and an example of substantiation would be too lengthy to include here. However, ANC-9, “Aircraft Propeller Handbook,” contains principles that may be applied to tail rotor designs. Tail rotors may be considered a special propeller design.

(iv) Bearings are generally used in the tail rotor installation to allow flapping and feathering motion of the blades. The bearings manufacturer’s ratings of these bearings must not be exceeded. Bearings generally used in main and tail rotors are classified as ABEC Class 3, 5, or 7. Class 7 is the highest quality presently available. Satisfactory completion of the endurance tests of §§ 29.923 and 29.927 is a means of proving that use of a particular bearing is satisfactory.

(v) The analysis must include appropriate special factors, casting factors, bearing factors, and fitting factors prescribed by §§ 29.619, 29.621, 29.623, and 29.625, respectively. The fitting factor of 1.15 must be applied in the analysis of the tail rotor installation.

AC 29.401A. § 29.401 (Amendment 29-31) AUXILIARY ROTOR ASSEMBLIES .

a. Explanation . Amendment 29-31 removed this section since the requirements are adequately addressed in §§ 29.337, 29.339, and 29.341.

b. Procedures . The policy material pertaining to this section is retained as supplemental information.

Page C - 24 9/30/99 AC 29-2C AC 29.403. § 29.403 AUXILIARY ROTOR ATTACHMENT STRUCTURE .

a. Explanation .

(1) The auxiliary rotor attachment structure(s), which is considered to include gear boxes, must be designed to withstand design limit loads that occur in flight and on landing. These design loads that generally consist of the following must be established for the particular flight and landing condition under consideration.

(i) Inertia loads generated by linear and angular accelerations of the auxiliary rotors and their gear boxes, combined with (ii) Thrust and torque loads developed by the auxiliary rotors.

The linear and angular acceleration loads imposed by the weight of the tail rotor and gearbox are generally derived from airframe loads data. Thrust and torque output of the tail rotor are derived during external aerodynamic and landing loads development for pertinent flight and landing conditions.

(2) General rules related to proof of structure loads and factor of safety are §§ 29.307, 29.301, 29.303, and 29.305.

b. Procedures .

(1) The angular and linear acceleration loads combined with appropriate tail rotor thrust and torque for the critical conditions shall be imposed on the tail rotor gearbox mount lugs, the airframe mounting structure, and the attaching hardware.

(2) The yaw and maximum power climb conditions are generally critical.

Landing and maneuvering conditions with and without power may also impose high inertia and rotor thrust and torque loads on the attachment structure.

(3) The derivation of the loads and conditions are too extensive to include here.

Additional information can be found in the U.S. Army Material Command Report AMCP 706-201, “Engineering Design Handbook: Helicopter Engineering, Part One, Preliminary Design.” AC 29.403A. § 29.403 (Amendment 29-31) AUXILIARY ROTOR ATTACHMENT STRUCTURE .

a. Explanation . Amendment 29-31 removed this section since the requirements are adequately addressed in §§ 29.337, 29.339, and 29.341.

b. Procedures . The policy material pertaining to this section is retained as supplemental information.

Page C - 25 AC 29-2C 9/30/99 AC 29.411. § 29.411 GROUND CLEARANCE: TAIL ROTOR GUARD .

a. Explanation .

(1) The rule requires specific protection to prevent the tail rotor from contacting the landing surface during a normal landing if it is possible that the tail rotor will contact the surface. The rule states that it must be impossible for the tail rotor to contact the surface during a normal landing.

(2) If a guard is required, the guard and its supporting structure must withstand suitable design loads.

(3) Section 29.501(c)(1) contains skid landing gear drag requirements that may be applied to the guard design loads.

b. Procedures .

(1) The applicant may submit sketches or drawings showing probable clearance with typical level landing surfaces during normal landings. Typical attitudes such as nose high autorotation, or autorotation with power-on landing, or other possible tail low attitudes should be investigated. If the drawings or sketches reveal that it is not likely the tail rotor will contact the landing surface, this minimum clearance with the landing surface may be confirmed during official FAA/AUTHORITY flight tests, such as HV and landing tests. The clearance may be confirmed by having a frangible device of suitable length (i.e., a balsa wood dowel) extending beyond the guard and attached to the tail rotor guard or other appropriate fuselage part. If the device is not damaged, broken, or no contact is made with the surface, compliance has been demonstrated.

(2) If it is possible for the tail rotor guard to contact the landing surface suitable design loads must be established for the guard. ANC-2a dated March 1948, “ANC Bulletin Ground Loads,” paragraph 6.4, entitled “Tail Bumper Criteria,” is an acceptable means of deriving the rotorcraft kinetic energy that shall be absorbed by the guard. This method is noted here for convenience.

(i) The tail rotor guard shall be able to absorb the kinetic energy of the rotorcraft in its most unfavorable CG position in the tail down landing attitude. The kinetic energy that the tail rotor guard shall be capable of absorbing must be determined as follows: Page C - 26 9/30/99 AC 29-2C 2 2 WV K S Y KE = ________ X ____________ 2 2 2g (K + 1 ) Y B where-- V = vertical speed ft/sec, derived from § 29.725(a) S K = pitching radius of gyration - ft. from pitching axis Y 1 = distance from most critical CG location to the guard B or bumper contact point - ft.

W = gross weight less rotor lift from § 29.473(a) - lbs.

G = 32.2 ft./sec (ii) Other, more recent, analytical techniques (most utilizing computer programs) may, of course, be used rather than the ANC-2a means after proper substantiation for applicability and validity.

(iii) The tail rotor guard shall not fail when the limit and ultimate load, which is derived from a combination of the limit kinetic energy and the guard resulting limit deflection required to dissipate the energy, is imposed on the guard and the rotorcraft tail (see § 29.305).

(3) Substantiation of the guard, skid, or bumper for the design loads derived may be accomplished by test or analysis as stated in § 29.307(a).

(4) Several rotorcraft tail rotor guards are installed solely for the protection of ground personnel from the rotating tail rotor. For guards installed for this purpose, the applicant should use prudent and reasonable design loads and features. Such guards should not present a hazard to the rotorcraft because of its design features.

AC 29.413. § 29.413 STABILIZING AND CONTROL SURFACES .

a. Explanation . Minimum design loads are specified for stabilizing as well as control surfaces.

(1) Paragraph (a) of the rule requires application of minimum empirical design loads, application of critical maneuvering loads, and application of critical maneuvering loads combined with vertical or horizontal gust loads (30 feet per second per § 29.341).

(2) Paragraph (b) requires load distributions that closely simulate actual pressure distributions. Both spanwise and chordwise distributions are intended.

(3) These surfaces are used for stability and control thereby hopefully extending the CG range and increasing the airspeed of modern designs.

Page C - 27 AC 29-2C 9/30/99 (4) To “closely simulate actual pressure condition” on the surfaces, unsymmetrical loads are also required on horizontal surfaces. An arbitrary distribution, if conservative, may be used.

(5) It is noted § 29.571 requires fatigue substantiation of the flight structure which will include control and stabilizing surfaces.

(6) If the surface is controllable, a proof and operation test of the surface control system is required by §§ 29.681 and 29.683.

b. Procedures . Modern rotorcraft designs have generally employed a fixed or a wholly movable, not split or divided, stabilizing or control surface.

(1) Design Loads .

(i) Limit loads of 15 pounds per square foot will apply up to approximately 90-knot design airspeed. Above a 90-knot design airspeed (V ), the D coefficient (C = 0.55) imposes higher limit loads on the surface.

N (ii) In addition, combined maneuvering and gust loads may impose the highest limit loads on the control surfaces of rotorcraft. This is attributed to the increase in speed (horizontal gust) and to the change in angle of attack and change in airspeed (vertical gust). Imposing the horizontal gust (30 feet per second or 17.8 knots) on the surface in combination with 130-knot design speed results in a 30 percent increase in the design load. The gust conditions cause a significant increase in design loads due to a change in angle of attack, with a change in resultant airspeed, or due to the increase in airspeed.

(iii) The applicant may choose to derive the limit loads using maximum aerodynamic coefficients for the surface under consideration at the maximum design airspeed combined with a 17.8-knot gust. This would be acceptable provided these design loads exceed the minimum loads derived from a C = 0.55 at design airspeed or N exceed 15 pounds per square foot load on the surface.

(2) The load distribution on the surface should closely simulate actual pressure distributions.

(i) The spanwise load may be rectangular or other acceptable conservative distributions may be used. The method developed by O. Schrenk in NACA TM 948, 1940, is an acceptable method for approximation of spanwise distribution.

NOTE: The method is valid for aspect ratios of 5 to 12 and for rectangular planforms such as used on rotorcraft, other planforms may be acceptable as prescribed in the TM.

Page C - 28 9/30/99 AC 29-2C (ii) The chordwise distribution appropriate for the aerodynamic shape should be used.

(iii) The flight load survey conducted under § 29.571 may be used to confirm design parameters and possible load distribution data. On controllable surfaces, the pitching moment (control loads) is measured for fatigue substantiation of the control system. The control stabilizing surfaces are subject to loads measurement and possible fatigue tests for fatigue substantiation also.

(3) Proof of the structure for the required loads is specified in §§ 29.301, 29.303, 29.305, and 29.307. Tests or analysis may be used as prescribed. If analysis is used, fitting factors and other appropriate factors prescribed by the rules of §§ 29.625, 29.621, and 29.623 will be required in the analysis.

AC 29.413A. § 29.413 (Amendment 29-31) STABILIZING AND CONTROL SURFACES .

a. Explanation . Amendment 29-31 removed this section since the requirements are adequately addressed in §§ 29.337, 29.339, and 29.341.

b. Procedures . The policy material pertaining to this section is retained as supplemental information especially as reference material for paragraph AC 29.341 (§ 29.341).

AC 29.427. § 29.427 (Amendment 29-31) UNSYMMETRICAL LOADS .

a. Explanation . Amendment 29-30 added the standard and Amendment 29-31 amended it. Minimum unsymmetrical design loads are specified for horizontal tail surfaces and also vertical tail surfaces whenever they support the horizontal tail surfaces.

(1) Loads are derived by rational analysis, or for earlier certification bases, the prescribed empirical loads of § 29.413 may be used. Section 29.413 was removed by Amendment 29-31 since the requirements are adequately addressed in §§ 29.337, 29.339, and 29.341.

(2) Rational loads, appropriate for the aerodynamic surfaces, should be distributed according to the standard.

(3) When vertical tail surfaces support the horizontal tail surfaces, the vertical tail surfaces and supporting surfaces are required to support the critical combination of vertical and horizontal surface loads distributed as shown.

b. Procedures . Two basic loading conditions are required by § 29.427 for each of the two basic empennage configurations shown.

Page C - 29 AC 29-2C 9/30/99 (1) Horizontal surfaces supported by the tail boom or fuselage . Structural substantiation should be provided for all six combinations shown in figure AC 29.427-1.

All of these empirical loading distributions should be used unless rational analysis shows one or more of each set of conditions to be non-critical or equal or more realistic distributions are substantiated. Rectangular spanwise air load distribution should be used unless more rational distribution is substantiated. If end plates are used, the air loads should be distributed accordingly.

(i) First unsymmetrical loading condition: (A) 100 percent of the flight load is applied to one side of the plane of symmetry; and 0 percent of the flight load is applied on the other side of the plane of symmetry.

(B) For surfaces with end plates or other similar devices, the load distribution will be changed accordingly.

(ii) Second unsymmetrical loading condition: 50 percent of the flight load on one side of the plane of symmetry acting up; and 50 percent of the flight load on the other side of the plane of symmetry acting down.

(2) Horizontal surfaces supported by a vertical surface . Structural substantiation should be provided for all six combinations shown in figure AC 29.427-2.

All of these empirical loading distributions should be used unless rational analysis shows one or more of each set of conditions to be non-critical or equal or more realistic distributions are substantiated. Rectangular spanwise air load distribution should be used unless more rational distribution is substantiated. If end plates are used, the air loads should be distributed accordingly.

(i) First unsymmetrical loading condition: 100 percent of the flight load on one side of the plane of symmetry; and 0 percent of the flight load on the other side of the plane of symmetry.

(ii) Second unsymmetrical loading condition: 50 percent of the flight load on one side of the plane of symmetry acting up; and 50 percent of the flight load on the other side of the plane of symmetry acting down.

Page C - 30 9/30/99 AC 29-2C

, _ JJtttr

!!Ill!

- ! l l I

Fi gur e AC 29.4 27 -1 (Vi ew Lo oking Fo rward) Page C - 31 AC 29-2C 9/30/99

t

Figure AC 29 .427-2 (Vi ew Lo ok ing F or wa rd) Page C - 32 9/30/99 AC 29-2C SUBPART C - STRENGTH REQUIREMENTS GROUND LOADS AC 29.471. § 29.471 GROUND LOADS - GENERAL .

a. Explanation . This regulation specifies that limit ground loads must be considered which are: (1) External loads caused by landing (ground) conditions and by ground taxiing loads as specified in § 29.235.

(2) Loads considering the rotorcraft structure as a rigid body.

(3) Loads in equilibrium with linear and angular inertia loads.

(4) The critical center of gravity “must be selected so that the maximum design loads are obtained in each landing gear element.” b. Procedures .

(1) The standards to be considered are specified in §§ 29.473 through 29.511.

These associated standards cover landing gear arrangements, landing conditions, and ground handling conditions.

(2) Drop tests are required for determination of landing load factors. See paragraph AC 29.723.

(3) The application of the design loads derived from the landing load factors will be as specified for each element affected by landing or ground handling loads.

(4) During the applicant’s flight test program, the ground, landing, and taxiing load factors may be monitored to assure the design load factors used are adequate.

See paragraph AC 29.235 for § 29.235 guidance.

AC 29.473. § 29.473 (Amendment 29-3) GROUND LOADING CONDITIONS AND ASSUMPTIONS .

a. Explanation . The rotorcraft is to be designed for the maximum weight. A rotor lift of two-thirds of the design maximum weight may be used. The minimum limit landing load factor is determined by the drop tests of § 29.725. Provisions are made for supplementary energy absorption devices that have triggering mechanisms.

b. Procedures . Loads for the landing conditions are derived considering mass (equal to the maximum weight) and rotor lift (equal to two-thirds of the maximum weight) acting through the center of gravity throughout the landing impact. Unbalanced external Page C - 33 AC 29-2C 9/30/99 loads resulting from asymmetric loading conditions are reacted as specified in the individual subparagraphs.

NOTE: If supplementary energy absorption devices are used, neither they nor their triggering devices may fail under the loads established by the limit drop tests or the reserve energy absorption drop tests.

AC 29.475. § 29.475 TIRES AND SHOCK ABSORBERS .

a. Explanation . This section specifies the tire and shock absorber position to be used in ground load derivations.

b. Procedures . Ground loads are to be derived with the tires in static (1g) position and the shock absorbers “in their most critical position.” The determination of the “most critical position” for the shock absorbers generally requires a load versus deflection test or analysis of the shock absorber system and a determination of the effect of both load and deflections on the shock absorber, attachment structure, and substructure designed by ground loads.

AC 29.477. § 29.477 LANDING GEAR ARRANGEMENT .

a. Explanation . This section specifies the individual standards to be used for ground load conditions for rotorcraft having two wheels aft and one or more wheels forward of the center of gravity.

NOTE: § 29.497 gives ground loading conditions for landing gear with tail wheels, and § 29.501 gives ground loading conditions for landing gear with skids.

b. Procedures . The ground loading conditions of §§ 29.235, 29.479 through 29.485, and 29.493 will be used for rotorcraft having two wheels aft and one or more wheels forward of the center of gravity. This includes forward wheels on separate axles.

AC 29.479. § 29.479 LEVEL LANDING CONDITIONS .

a. Explanation . This section provides explicit level landing load criteria for landing gear with two wheels aft and one or more wheels forward of the center of gravity.

(1) Level landings- - (i) Each wheel contacting the ground simultaneously; and (ii) Aft wheels contacting the ground with forward wheels just clear of the ground.

(2) Application of loads- - Page C - 34 9/30/99 AC 29-2C (i) Maximum design vertical loads applied alone; (ii) The maximum design vertical loads applied with a drag load of at least 25 percent of the vertical load (applied at the ground contact area); and (iii) The vertical load at the instant of peak drag load in conjunction with the peak drag load. A ground speed and load application is specified.

(3) A 40 percent/60 percent load distribution between wheels for configurations having two forward wheels including quadricycle. This distribution between wheels on a common axis is to be applied for the conditions of vertical loads only, and for vertical loads combined with drag loads of 25 percent of the vertical loads. Section 29.511 concerns a 60 percent to 40 percent ground load distribution between multiple-wheel units. See paragraph AC 29.511 for dual wheels on a common axle or axis.

(4) Aircraft pitching moments are to be reacted by the forward landing gear or by the angular inertia forces when the forward landing gear is clear of the ground as specified.

b. Procedures .

(1) The specified loading conditions will be used in load derivations.

(2) The critical center of gravity condition will be used for each gear and gear support structure.

(i) The aft center of gravity condition with the forward gear clear will normally be critical for the aft gear and gear supports.

(ii) The forward center of gravity condition with each gear contacting the ground simultaneously will normally design forward gear elements critical for vertical loads.

(iii) The forward center of gravity condition with the forward gear clear may result in high load factors, angular plus linear, that will greatly affect security of items of significant mass.

(3) The vertical load, at the instant of peak drag load combined with the peak drag component, can be determined from drop tests utilizing wheel spin-up or it can be analytically determined. If analysis is used, it must successfully correlate with the results of a previous well-instrumented test program.

Page C - 35 AC 29-2C 9/30/99 AC 29.481. § 29.481 TAIL-DOWN LANDING CONDITIONS .

a. Explanation . This section provides the criteria for tail-down landing conditions, i.e., “the maximum nose-up attitude allowing ground clearance” with ground loads acting “perpendicular to the ground.” b. Procedures .

(1) The tail-down landing condition will be used to check (by analysis or test) for criticality of landing gear or support structure. This attitude generally creates the highest forward loads on the landing gear in combination with vertical loads.

(2) The tail-down landing condition may be the critical condition for both landing load factor and for energy absorption by the main gear. Section 29.725 requires that “each landing gear must be tested in the attitude simulating the landing condition that is most critical.” Where questions exist as to the critical attitude, both level landing and tail-down landing attitudes should be used in drop tests required by § 29.725.

AC 29.483. § 29.483 ONE-WHEEL LANDING CONDITIONS .

a. Explanation . This section gives the condition to be used for one-wheel landing conditions. Only the vertical load condition of § 29.479(b)(1) is required.

b. Procedures . The one-wheel landing condition is generally critical for the landing gear-to-fuselage attachments and the landing gear elements between the attachments. Unbalanced external loads are reacted by rotorcraft inertia. Large items of mass located radially from the center of gravity (aircraft centerline may be used) should also be structurally substantiated for the combined rolling (angular) and linear accelerations of this loading condition.

AC 29.485. § 29.485 LATERAL DRIFT LANDING CONDITIONS .

a. Explanation .

(1) This section provides the loading conditions which impose side (and vertical) loads on the landing gear. A level landing attitude is specified. Two main conditions required are- - (i) Only the aft wheels in contact with the ground; and (ii) All wheels contacting the ground simultaneously.

(2) Loads. The vertical loads to be applied with the side loads are specified as “one-half of the maximum ground reactions of § 29.479(b)(1).” These vertical loads are the level landing loads considering both contact and noncontact with the ground by the forward wheels.

Page C - 36 9/30/99 AC 29-2C (i) One side load condition is specified as “0.8 times the vertical reaction acting inward on one side and 0.6 times the vertical reaction acting outward on the other side” when only the aft wheels contact the ground.

(ii) The other side load condition (for all wheels contacting the ground) specifies the 80 percent inward/60 percent outward distribution for the aft wheels and 0.8 times (80 percent) the vertical reaction for the forward wheels.

b. Procedures . The loading conditions, as specified, are applied to the landing gear and attaching structure. The loads are applied at the ground contact point, except for full swiveling gear which has the load applied at the center of the axle. In other words, full swiveling gear is considered to have swiveled to a static position under the side load before the design vertical and side loads are achieved. The landing gear backup structure, as well as the landing gear itself, will be substantiated for these side load conditions.

AC 29.493. § 29.493 BRAKED ROLL CONDITIONS .

a. Explanation . This section provides two loading conditions for ground braking operations. Specific vertical loads in conjunction with drag loads (due to braking) are to be considered. The limit vertical load factor is 1.33 for condition of all wheels in contact with the ground, and 1.0 for condition of aft wheels only in contact with the ground and nose wheel clear. The drag load on wheels with brakes is 0.8 times the vertical load or the drag load value based on limiting brake torque, whichever is less.

b. Procedures . The braking loads are calculated from the specified criteria with the shock absorbers in their static (normal) positions and with the drag loads applied at the ground contact point. Structural substantiation of the affected structure may be accomplished by test or analysis. If tests are used, the wheel and tire assembly is commonly replaced with a test fixture so the limit loads and static deflections specified can be more accurately controlled. The test specimen should be complete enough to assure that the landing gear structure and the attach and backup structure are adequately substantiated.

AC 29.497. § 29.497 GROUND LOADING CONDITIONS: LANDING GEAR WITH TAIL WHEELS .

a. Explanation . This section provides the loading conditions for landing gear designs with tail wheels.

(1) Level landings are to consider the following: (i) All wheels (main and tail) contacting the ground simultaneously, as well as only forward main wheels contacting the ground.

Page C - 37 AC 29-2C 9/30/99 (ii) Maximum design vertical loads applied alone.

(iii) The maximum design vertical loads combined with a drag load of at least 25 percent of the vertical loads for both conditions.

(2) Nose-up landings with only the rear wheel or wheels initially contacting the ground must be considered unless shown to be extremely remote.

(3) Level landings on one forward wheel only are to be considered. Drag loads are not required.

(4) Side load conditions are imposed on the main wheels and tail wheels for level landing attitudes. Criteria for full swiveling and locked tail wheels are included in this standard.

(5) Braked roll conditions are specified for the level landing attitudes.

(6) Rear wheel turning loads are also specified for swiveling and locked tail wheels.

(7) Taxiway condition loads for the landing gear and rotorcraft are those that “occur when the rotorcraft is taxied over the roughest ground that may reasonably be expected in normal operation.” The aircraft design load factors should not be exceeded during the evaluation. Section 29.235 contains an identical standard that applies to all types of wheel landing gear.

b. Procedures .

(1) The specified loading conditions are to be used in load derivations.

(2) The critical center of gravity condition is used for each gear and gear support structure.

(i) The forward center of gravity condition with the tail gear clear will normally be critical for the forward gear and gear supports.

(ii) The aft center of gravity condition with the tail gear clear should be checked for criticality of security of large mass items located forward of the center of gravity. Vertical and angular accelerations are additive under this landing condition.

(iii) The aft center of gravity condition with each gear contacting the ground simultaneously will generally design tail gear elements critical for vertical loads.

The other conditions are generally less severe but must be proven.

(3) For nose-up landing procedures use § 29.481. The reference to “extremely remote” in § 29.497(d)(2) predates current §§ 25.1309, 29.1309, and AC 25.1309.1.

Page C - 38 9/30/99 AC 29-2C This phrase has been used to require consideration of nose-up landings unless features of design are present which prevent nose-up landings or where such landings are unlikely during the life of the rotorcraft. See paragraph AC 29.481.

(4) Use § 29.483 for one-wheel landing procedures, paragraph AC 29.483.

(5) Use § 29.485 procedures for side load conditions, paragraph AC 29.485.

(6) Use § 29.493 procedures for braked roll conditions, paragraph AC 29.493.

(7) For rear wheel turning loads, swiveling of tail landing gears is allowed as in basic side load conditions. The side load is applied at the axle, or if the wheel is locked, the load is applied at ground contact. Rear wheels are loaded with the critical vertical static load in conjunction with an equal side load to substantiate the tail gear.

(8) Since the rotorcraft is to be designed for load factors that will not be exceeded during taxi tests or other conditions, an instrumented taxi test program will be necessary. Use § 29.235, paragraph AC 29.235.

AC 29.501. § 29.501 (Amendment 29-3) GROUND LOADING CONDITIONS: LANDING GEAR WITH SKIDS .

a. Explanation . This section provides the ground loading conditions for landing gear with skids. The loading conditions are similar to those for wheeled gear except for the following criteria which are unique to skid gears: (1) Structural yielding (plastic deformation) of elastic spring members under limit loads is allowed.

(2) Design ultimate loads for elastic spring members need not exceed the loads obtained in a drop test with a drop height of 1.5 times the limit drop height. The rotorcraft and the landing gear attachments are subject to the prescribed design ultimate loads.

(3) The gear must be in its most critically deflected position (similar to § 29.475).

(4) Ground reactions are rationally distributed along the bottom of the skid unless otherwise specified. Paragraph (f) concerns specific “concentrated” and arbitrary load conditions.

(5) Drag loads are 50 percent of vertical reactions rather than the 25 percent for wheeled gear.

(6) Side loads are 25 percent of the total vertical reaction rather than the 60-80 percent for wheeled gear.

Page C - 39 AC 29-2C 9/30/99 (7) Side loads are applied to one skid only (inward acting and outward acting) with resulting unbalanced moment resisted by angular acceleration.

(8) A ground reaction load of 1.33 times the maximum weight is to be applied at 45° from the horizontal axis: (i) Distributed among or between the skids; (ii) Concentrated at the forward end of the straight portion of the skid tube; and (iii) Applied only to the forward end of the skid tube and its attachment to the rotorcraft.

(9) A concentrated vertical load equal to one-half of the design limit vertical load is to be applied at a point midway between the skid tube attachments.

b. Procedures .

(1) The specified loading conditions are to be used in load derivations.

(2) The critical center of gravity conditions are to be used for each gear and gear support structure. Asymmetry of the skid tubes, cross tubes, and gear attachments are to be considered in determining the critical center of gravity condition.

(3) The rotorcraft and landing gear attachment must be substantiated for ultimate landing loads by either test or analysis utilizing an ultimate load factor of 1.5 in accordance with § 29.303. The elastic spring members may be analyzed or static tested for ultimate loads (and deflections) using either a factor of safety of 1.5 or one associated with an “ultimate” drop height of 1.5 times the limit drop height.

Substantiation by “ultimate” drop tests may be used provided all combinations of critical parameters are included in the total substantiation effort. This method will require a series of tests using several test specimens, or a limited number of drop tests plus further substantiations by static tests or analyses for additional critical conditions not covered by the drop test(s).

AC 29.501A. § 29.501 (Amendment 29-30) GROUND LOADING CONDITIONS: LANDING GEAR WITH SKIDS .

a. Explanation . Amendment 29-30 relaxes previous requirements in two cases by: (1) Allowing the total sideload of § 29.501(d)(3) to be distributed “equally between skids” rather than being “applied along the length of one skid only;” and, Page C - 40 9/30/99 AC 29-2C (2) Allowing the concentrated load of § 29.501(f)(2)(ii) to be distributed over 33.3 percent of the skid (between skid tube attachments) rather than being “concentrated at a point midway between the skid tube attachments.” b. Procedures . The previous procedures (through Amendment 29-19) continue to apply to Amendment 29-30 except for the use of the new load distributions.

AC 29.505. § 29.505 SKI LANDING CONDITIONS .

a. Explanation . This is an optional requirement for ski operations. The regulation specifies vertical loads, side loads, and torque loads (M ) to be applied to ski Z installations. The four loading conditions to be applied at the pedestal bearings are: (1) Simultaneous application of Pn, up load, and Pn/4, horizontal load.

(2) Up load of 1.33 P.

(3) Side load of 0.35 Pn.

(4) Torque load of 1.33 P (in foot-pounds), about vertical axis through the centerline of the pedestal bearings.

NOTE: Where P is the maximum static weight on each ski and n is the limit load factor obtained from drop tests. The load factor obtained from wheel or skid landing gear drop tests may be used.

b. Procedures . Structural substantiation may be accomplished by static test or analysis using the specified loads. Skis generally have a limit load rating. The design loads derived for this standard must not exceed the rating. TSO-c28 concerns, in part, standards for aircraft skis.

AC 29.511. § 29.511 (Amendment 29-3) GROUND LOAD: UNSYMMETRICAL LOADS ON MULTIPLE-WHEEL UNITS .

a. Explanation . Two loading conditions are provided to account for unsymmetrical loads on multiple-wheel units due to landing and normal operations over crowned runways and taxiways and to account for deflated tires. They are: (1) Sixty percent of total ground reaction applied to one wheel of a dual wheel unit and 40 percent to the other.

(2) Sixty percent of the “specified load for the gear unit” is applied to the wheel with an inflated tire when the other tire is deflated (the 60 percent load may not be less than the 1g static load).

Page C - 41 AC 29-2C 9/30/99 NOTE: The 60:40 distribution also applies to nose wheel units as noted in § 29.479(b)(4).

b. Procedures . Structural substantiation may be accomplished by static test or analysis using the specified load. As provided by the standard, the total load on the gear units may neglect the transverse shift of the load centroid due to unsymmetrical load distribution; i.e., the external load for each gear may be calculated considering the same load centroid as with symmetrical wheel loads, and then the external load for each gear is divided in accordance with the distributions of § 29.511(a) and (b) between the wheels.

Page C - 42 9/30/99 AC 29-2C SUBPART C - STRENGTH REQUIREMENTS WATER LOADS AC 29.519. § 29.519 (Amendment 29-30) HULL TYPE ROTORCRAFT: WATER-BASED, AMPHIBIAN .

a. Explanation .

(1) This regulation provides design criteria for amphibian rotorcraft with hull provisions.

(2) The most severe wave heights for which approval is desired are to be considered. A minimum of sea state 4 condition wave heights should be considered (reference paragraph AC 29.801 for a description of sea state 4 conditions).

(3) A rotor lift of two-thirds of the rotorcraft weight may be applied during landing impact.

(4) Vertical landing conditions are specified as: (i) Zero forward speed.

(ii) Likely pitch and roll attitudes.

(iii) Vertical descent velocity ≥ 6.5 FPS.

(5) Forward speed landing conditions are specified as: (i) Forward velocities of zero to 30 knots (a 30-knot limit may be reduced if it can be demonstrated that the maximum forward velocity selected would not be exceeded in a normal one-engine-out landing).

(ii) Likely pitch, roll, and yaw attitudes.

(iii) Vertical descent velocity ≥ 6.5 FPS.

(6) Auxiliary float immersion conditions are specified to be applied unless it can be shown that full immersion is unlikely. If full immersion is unlikely, the highest float buoyancy load is specified that considers loading of the float immersed to create restoring moments which compensate for upsetting moments caused by side wind, asymmetrical rotorcraft loading, water wave action, and rotorcraft inertia.

b. Procedures.

Page C - 43 AC 29-2C 9/30/99 (1) Tests should be conducted to establish procedures for water entry. These tests should include determination of optimum pitch attitude and forward velocity for landing in a calm sea as well as entry procedures for the highest sea state to be demonstrated (e.g., the recommended part of the wave on which to land and direction of landing relative to crest/trough direction).

(2) The landing structural design consideration should be based on water impact with a rotor lift of not more than two-thirds of the maximum design weight acting through the center of gravity under the following conditions: (i) Vertical Landing Conditions.

(A) Zero forward velocity.

(B) The optimum pitch attitude as determined in paragraph AC 29.519b(1) with consideration for pitch attitude variations that would reasonably be expected to occur in service.

(C) Vertical descent velocity of 6.5 FPS or greater.

(D) Likely roll attitudes.

(ii) Forward Speed Landing Conditions.

(A) Forward velocities of zero to 30 knots (or a reduced maximum forward velocity if it can be demonstrated that a lower maximum velocity would not be exceeded in a normal one-engine-out landing).

(B) The optimum pitch attitude as determined in paragraph AC 29.519b(1) with consideration for pitch attitude variations that would reasonably be expected to occur in service.

(C) Vertical descent velocity of 6.5 FPS or greater.

(D) Likely roll and yaw attitudes.

(3) Landing load factors may be determined by- - (i) Landing gear drop tests for limited amphibian; (ii) Water drop tests for amphibian; or (iii) Analysis based on tests.

(4) Water load distribution should be determined by tests or analysis based on tests.

Page C - 44 9/30/99 AC 29-2C (5) Auxiliary float loads should be determined by full immersion or restoring moments required to react upsetting moments caused by side wind, asymmetrical rotorcraft loading, water wave action, and rotorcraft inertia. Auxiliary float loads may be determined by analysis. Load distributions should be determined by tests or analysis based on tests.

AC 29.521. § 29.521 (Amendment 29-3) FLOAT LANDING CONDITIONS .

a. Explanation . This is an optional requirement for float operations, and it applies only when float operations are requested. The regulation specifies vertical loads, aft loads, and side loads to be applied to the float installations. The two loading conditions to be applied are: (1) Up-load Condition .

(i) A vertical load appropriate to a landing load factor determined under § 29.473(b).

(ii) The resultant water reaction passes vertically through the aircraft CG.

(iii) An aft load equal to 25 percent of the vertical load.

(2) Side-load Condition .

(i) A vertical load equal to 75 percent of the vertical load for the up-load condition.

(ii) Vertical load equally divided among the floats.

(iii) A side load at each float equal to 25 percent of the vertical load at each float.

b. Procedures .

(1) The vertical load factor is determined by drop tests in accordance with §§ 29.473(b) and 29.725. The floats may be drop tested, or they may be assumed to have the same load factor as wheeled gear which have been drop tested.

(2) Structural substantiation may be accomplished by either static tests or analysis using the specified loads. The load distribution on the floats may be realistically based on hydrostatic pressure distributions or conservative pressure distributions.

Page C - 45 AC 29-2C 9/30/99 SUBPART C - STRENGTH REQUIREMENTS MAIN COMPONENT REQUIREMENTS AC 29.547. § 29.547 (Amendment 29-4) MAIN ROTOR STRUCTURE .

a. Explanation . This regulation requires the main rotor structure to be designed to the static load requirements of §§ 29.337 through 29.351 (vertical maneuvering loads, vertical and horizontal gust loads, and yawing maneuver loads). In addition, the main rotor blades, hubs, and flapping hinges are specified to be designed for impact forces of each blade against its stop during ground operation and for specified limit torque at any rotational speed including zero. The torque forces (from the drive system) are distributed to the rotor blades as specified.

b. Procedures .

(1) Substantiation in compliance with this standard is accomplished by application of the flight loads of §§ 29.337 through 29.351 and the torque loads of § 29.361 to the rotor structure by stress analyses and/or static tests. The use of wind tunnel data as well as flight loads survey data may be used to generate and/or check the external load magnitudes and distributions.

(2) Where new materials are used in the main rotor structure, such as composites containing plastics, the effects of temperature and humidity are to be considered in accordance with § 29.603, and the effects of uncertainties in manufacturing processes or inspection methods are to be considered in accordance with § 29.619.

(3) The design impact forces of each blade must be imposed against the blade stop or stops. Impact loads from 2 to 3 g’s have been commonly used to provide rotor structure protection against blades impacting against lower (droop) stops. Different values may be used for flapping and lag stops as determined by a rational basis.

Appropriate monitoring of the blades, hubs, flapping hinges, and stops during laboratory tests, ground endurance tests, and flight tests should ensure that the stops are sufficient for ground operation loads (taxiing, backing, etc.), training, and offshore platform landings. Taxiing should consider typical obstacles such as pavement edges, ropes, air lines, and so forth. The design torque loads are derived as prescribed.

AC 29.547A. § 29.547 (Amendment 29-40) MAIN ROTOR AND TAIL ROTOR STRUCTURE .

a. Explanation . Amendment 29-40 revised § 29.547 to add requirements to perform a design assessment. Section 29.547 (a) and (b) set forth a definition of a rotor and its associated components and requires a design assessment to be performed.

The intent of these paragraphs is to identify the critical components and/or clarify their Page C - 46 9/30/99 AC 29-2C design integrity to show that the basic airworthiness requirements which are applicable to the rotors will be met.

A design assessment of the rotors should be carried out in order to substantiate that they are of a safe design and that compensating provisions are made available to prevent failures classified as hazardous and catastrophic in the sense specified in paragraph b below. In carrying out the design assessment, the results of the certification ground and flight testing (including any failures or degradation) should be taken into consideration. Previous service experience with similar designs should also be taken into account (see also § 29.601(a)).

b. Definitions . For the purposes of this assessment, failure conditions may be classified according to the severity of their effects as follows: (1) Minor . Failure conditions which would not significantly reduce rotorcraft safety, and which involve crew actions that are well within the crew capabilities. Minor failure conditions may include, for example, a slight reduction in safety margins or functional capabilities, a slight increase in crew workload, such as routine flight plan changes, or some inconvenience to occupants.

(2) Major . Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be, for example, a significant reduction in safety margins or functional capabilities, a significant increase in crew work load or in conditions impairing crew efficiency, or discomfort to occupants, possibly including injuries.

(3) Hazardous . Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be - - (i) A large reduction in safety margins or functional capabilities.

(ii) Physical distress or higher workload such that the flight crew cannot be relied upon to perform their tasks accurately or completely.

(iii) Serious or fatal injury to a relatively small number of the occupants.

(iv) Loss of ability to continue safe flight to a suitable landing site.

(4) Catastrophic . Failure conditions which would prevent a safe landing.

(5) Minimize . Reduce to the least possible amount by means that can be shown to be both technically feasible and economically justifiable.

(6) Health Monitoring . Equipment, techniques, and/or procedures by which selected incipient failure or degradation can be determined.

Page C - 47 AC 29-2C 9/30/99 c. Procedures .

(1) Failure Analysis . The first stage of the design assessment should be the failure analysis, by which all the hazardous and catastrophic failure modes are identified. The failure analysis may consist of a structured, inductive bottom-up analysis, which is used to evaluate the effects of failures on the system and on the aircraft for each possible item or component failure. When properly formatted, it will aid in identifying latent failures and the possible causes of each failure mode. The failure analysis should take into consideration all reasonably conceivable failure modes in accordance with the following: (i) Each item/component function(s).

(ii) Item/component failure modes and their causes.

(iii) The most critical operational phase/mode associated with the failure mode.

(iv) The effects of the failure mode on the item/component under analysis, the secondary effects on the rotors and on the rotor drive system, on other systems, and on the rotorcraft. Combined effects of failures should be analyzed where a primary failure is likely to result in a secondary failure.

(v) The safety device or health monitoring means by which occurring or incipient failure modes are detected, or their effects mitigated. The analysis should consider the safety system failure.

(vi) The compensating provision(s) made available to circumvent or mitigate the effects of the failure mode (see also paragraph c(2) below) (vii) The failure condition severity classification according to the definitions given in paragraph b above.

When deemed necessary for particular system failures of interest, the above analysis may be supplemented by a structured, deductive top-down analysis, which is used to determine which failure modes contribute to the system failure of interest.

Dormant failure modes should be analyzed in conjunction with at least one other failure mode for the specific component or an interfacing component. This latter failure mode should be selected to represent a failure combination with potential worst case consequences.

When significant doubt exists as to the effects of a failure, these effects may be required to be verified by tests.

Page C - 48 4/25/06 AC 29-2C, Chg 2 (2) Evaluation of Hazardous and Catastrophic Failures : The second stage of the design assessment is to summarize the hazardous and catastrophic failures and appropriately substantiate the compensating provisions which are made available to minimize the likelihood of their occurrence. Those failure conditions that are more severe should have a lower likelihood of occurrence associated with them than those that are less severe. The applicant should obtain early concurrence of the cognizant certificating authority with the compensating provisions for each hazardous or catastrophic failure.

Compensating provisions may be selected from one or more of those listed below, but not necessarily limited to this list.

(i) Design features; i.e., safety factors, part derating criteria, redundancies, etc.

(ii) A high level of integrity: All parts with catastrophic failure modes and critical characteristics are to be identified as Critical Parts and be subject to a Critical Parts Plan (see AC 29.602). Where a high level of integrity is used as a compensating provision, parts with a hazardous failure mode which would prevent continued safe flight may be included in a Critical Parts Plan or subjected to other enhancements to the normal control procedures for parts.

(iii) Fatigue tolerance evaluation.

(iv) Flight limitations.

(v) Emergency procedures.

(vi) An inspection or check that would detect the failure mode or evidence of conditions that could cause the failure mode.

(vii) A preventive maintenance action to minimize the likelihood of occurrence of the failure mode including replacement actions and verification of serviceability of items which may be subject to a dormant failure mode.

(viii) Special assembly procedures or functional tests for the avoidance of assembly errors which could be safety critical.

(ix) Safety devices or health monitoring means beyond those identified in (vi) and (vii) above.

Page C - 49 AC 29-2C 9/30/99 AC 29.549. § 29.549 (Amendment 29-26) FUSELAGE AND ROTOR PYLON .

a. Explanation . This regulation requires that the fuselage and rotor pylon (including the tail fin, if any) be designed to withstand the flight loads of §§ 29.337 through 29.351, the ground loads of §§ 29.235, 29.471 through 29.497, skid loads of § 29.501, ski loads of § 29.505, water loads of § 29.521, and rotor loads of § 29.547(d) and (e). The ski and water loads pertain to optional features.

(1) Consideration is also required of -- (i) Auxiliary rotor thrust; (ii) The torque reaction of each rotor drive system; and (iii) Balancing air and inertia loads.

(2) Each engine mount and adjacent fuselage must be substantiated as prescribed. In addition, if 2 ½-minute power is used, “each engine mount and adjacent structure must be designed to withstand the loads resulting from a limit torque equal to 1.25 times the mean torque for 2 ½-minute power combined with 1g flight loads.” Amendment 29-26 extended paragraph (e) of the standard to 2 ½-minute “OEI power.” b. Procedures . Compliance with this standard is accomplished by application of the specified aircraft loads including engine torque to the fuselage and rotor pylon structure by stress analyses and/or static tests. Drive system torque factors to be used are noted in § 29.547 for the main rotor structure as well as in § 29.549(e).

AC 29.551. § 29.551 AUXILIARY LIFTING SURFACES .

a. Explanation . This regulation specifies that auxiliary lifting surfaces be designed to withstand critical flight and ground loads derived for conditions specified and any “other critical condition expected in normal operation.” Stub wings would comply with this standard.

b. Procedures . The surface design loads are derived from the conditions specified. Conservative aerodynamic data, including load distributions, may be used in place of data derived from wind tunnel or instrumented flight testing of the exact aerodynamic shapes involved. Special attention should be placed on concentrated load effects from fuel tanks or other large mass items that may be located in lifting surfaces.

These types of load concentrations are to be considered in conjunction with inertia and aerodynamic loads.

Page C - 50 5/1/2014 AC 29-2C, Chg 4 SUBPART C – STRENGTH REQUIREMENTS EMERGENCY LANDING CONDITIONS AC 29.561. § 29.561 EMERGENCY LANDING CONDITIONS - GENERAL.

a. Explanation .

(1) Occupant protection . The occupants should be protected as prescribed from serious injury during an emergency, minor crash landing on water or land for the conditions prescribed in the standard. The standard states that each occupant should be given every reasonable chance of escaping serious injury in a minor crash landing.

In addition, the occupants must be protected from items of mass inside the cabin as well as outside the cabin. For example, a cabin fire extinguisher must be restrained for the load factors prescribed in this section. A transmission or engine must be restrained to the load factors in § 29.561(b)(3) if located adjacent to, above, or behind the occupants.

(2) Load factor determination . Section 29.561(b)(3) specifies certain ultimate inertial load factors but allows a lesser downward vertical load factor by virtue of a 5 FPS ultimate rate of descent at maximum design weight.

(3) Retractable landing gear . For rotorcraft equipped with retractable landing gear only the retracted configuration must be considered.

(4) Fuel tank protection .

(i) Underfloor fuel tanks are specifically addressed in § 29.561(d). The fuselage structure must be designed to resist crash impact loads prescribed in § 29.561(b)(3) and to also protect the fuel tank from rupture as prescribed. The landing gear must be retracted if the rotorcraft is equipped with retractable gears.

(ii) Section 29.963(b), a general rule tank design standard, also refers to § 29.561. This standard specifies that each tank and its installation must be designed or protected to retain fuel without leakage under the emergency landing conditions in § 29.561. Section 29.963 of this AC relates to this standard.

(5) External load considerations . The load factors of § 29.561 and the criteria of § 29.562 are not directly applicable to external load systems. This is because in emergency crash scenarios that involve external loads, the external load is neither typically subjected to the same minor crash loads (§ 29.561) as is the rotorcraft hull and its internal occupants nor are all of the occupant protection criteria (§ 29.562) needed or practicable to apply. Appropriate safety for external load carriage systems is provided by the criteria of § 29.865. Safety standards for external load attaching means are provided in § 29.865.

Page C - 51 AC 29-2C, Chg 4 5/1/2014 b. Procedures .

(1) The design criteria report or another similar report of the rotorcraft structural limits should contain the (ultimate) minor crash condition load factors.

(2) Section 29.785 (section 29.875 of this AC) concerns application of this design standard to seats (berths, litters), belts, and harnesses.

(3) The ultimate design landing and maneuvering load factors may exceed the minor crash condition load factors. The highest load factor derived must be used.

(i) For example, for light weight conditions, the ultimate maneuvering load factor may be 5.25g as specified in § 29.337.

(ii) The ultimate vertical landing load factors derived from §§ 29.471 through 29.521, whichever are appropriate for the design, may exceed the 4.0g down load factor in this section. The rotorcraft landing case design limit contact velocity must be at least 6.5 FPS (see §§ 29.473 and 29.725).

(4) As specified in § 29.561(b)(3)(iv), the downward load factor is 4.0, or a lower design load factor may be used at maximum design weight.

(i) The lower load factor relates to a rotorcraft impacting a flat, hard landing surface at 5 FPS (ultimate) vertical rate of descent. The load factor derived for each unique design is a function of the rotorcraft impact and crushing characteristics.

(ii) The 4.0g down load factor case is related to either a fixed or retractable gear rotorcraft. This condition is not dependent on impact characteristics of the rotorcraft.

(iii) As noted in paragraph b.(3) above, the design landing load factors may exceed each of the two previous cases and would then become the prominent design (vertical load) parameter for seats, transmissions, fire extinguishers, and so forth.

(5) Items of mass such as fire extinguishers, radio equipment, life rafts, engines, and transmissions must be restrained for the appropriate load factors.

(6) Cargo or baggage compartments separated from the passenger compartment must be designed for load factors specified in § 29.787. The conditions in § 29.561 are excepted from that standard.

(7) Each fuel tank and its installation are subject to the loads stated in this standard whether “under floor” or located elsewhere. (See § 29.963(b) also.) Under - floor fuel tanks are specifically addressed in § 29.561(d); however, an acceptable means of compliance with CAR 7.261 which is identical to and preceded § 29.561(d) is quoted here for information.

Page C - 52 5/1/2014 AC 29-2C, Chg 4 Notes: Fuselage keels whose design and structural strength are such as to resist crash impacts associated with the emergency landing conditions of § 7.260 (§ 29.561) without extreme distortion which might tend to rupture the fuel tank may be considered to comply with the requirements of this section (7.261).

Puncture resistant “bladder” fuel cells that are adequately designed and also protected from the stated impact loads imposed on the fuselage may also satisfy the standards.

(8) For rotorcraft with retractable landing gear, alternative landing gear positions and the resulting effects on potential fuel release should be evaluated.

AC 29.561A. § 29.561 (Amendment 29-29) EMERGENCY LANDING CONDITIONS - GENERAL.

a. Explanation . Amendment 29-29 adds or increases the design static load factors of § 29.561 in three different areas. The addition of these load factors eliminates the 5 FPS descent velocity criteria of unamended § 29.561(b)(3).

(1) The design static load factors for the cabin in § 29.561(b)(3) are increased in concert with the dynamic test requirements of new § 29.562.

(2) Design static load factors are added in § 29.561(c) for external items of mass located above or behind the crew and passenger compartment.

(3) The static load factors, which were formerly only referenced in § 29.561(d), are now included explicitly in § 29.561(d) for substantiation of internal fuel tanks which are below the passenger floor.

b. Procedures . The procedures in section 29.561 of this AC continue to apply except the new load factors of § 29.561 should be used. Penetration of any items of mass into the cabin or occupied areas should be prevented. In addition, each fuel tank and its installation are subject to specific load factors that are based on the fuel tank location.

(1) The crash impact load factors for the airframe structure surrounding the underfloor fuel tanks are specified in § 29.561(d). The fuselage structure must be designed to resist the specified crash impact loads and to help protect the fuel tank from rupture. If equipped with retractable landing gear, the effects of the landing gear on fuel system rupture should be considered in both the retracted and unretracted configurations.

(2) Section 29.952(b) (see section 29.952 of this AC) specifies the design load factors for crash resistant fuel systems in an otherwise survivable impact. This section relates to § 29.561(d) as follows. The § 29.952 load factors are for the fuel tanks, other Page C - 53 AC 29-2C, Chg 4 5/1/2014 significant mass items in the fuel system, and their attachment to the rotorcraft airframe for both occupant survivability and retention of fuel in a survivable impact; whereas, the § 29.561(d) load factors only apply to the rotorcraft airframe surrounding the underfloor fuel tanks and their installation for the same reasons. These two sets of load factors are not additive. They are applied separately (as design ultimate load factors) to the portions of the rotorcraft to which they are specified to apply. The application of the § 29.561(d) load factors is described as follows. The loads generated by § 29.561(d) are intended to be applied to the airframe structure surrounding the fuel cell to ensure that the entire airframe structure provides the appropriate level of crash resistance (i.e., stiffness, crushability, crushing rate, energy absorption capability, etc.) and to ensure that the airframe structure’s failure modes (e.g., buckling, creation of sharp edges, structural spears, etc.) are such that fuel cell rupture (and the resultant post crash fire potential) is mitigated to the maximum practicable extent in a otherwise survivable emergency landing. Each fuel cell (and major fuel cell component) creates an applied load on the airframe in an emergency landing condition. These loads are determined by multiplying the worst case mass of the fuel cell (i.e., a full fuel cell) by the load factors of § 29.561(d). These loads are then applied (utilizing the appropriate design load paths) to the airframe structure surrounding the fuel cell to help design the structure for optimal crash resistance. Added stiffness effects for both a full and less than full fuel cell should be considered in the design process. A significantly less than full fuel cell will typically not have any significant stiffness effects, since in a less than full condition the fuel cell cannot typically transfer load hydraulically.

(3) The minor crash ultimate load factors for doors and others emergency exits are specified in §§ 29.783(d) and 29.809(e). The related inertial forces are not applicable to cargo or to service doors (not suitable for use as an exit in an emergency).

If any item of mass installed in the cabin can possibly interact with the fuselage and cause higher deformation, then § 29.561 (b)(3) loads factors should be applied to the design of doors and emergency exits.

AC 29.561B. § 29.561 (Amendment 29-38) EMERGENCY LANDING CONDITIONS - GENERAL.

a. Explanation. Amendment 29-38 adds a new rearward emergency load factor of 1.5g to both §§ 29.561(b)(3)(v) and 29.561(c)(5). The addition of the 1.5g rearward load factor in § 29.561(b)(3)(v) is to provide an aft ultimate load condition for substantiation of the restraints required for retention of both occupants and significant items of mass inside the cabin that could otherwise come loose and cause injuries in an emergency landing. The addition of the 1.5g rearward load factor to § 29.561(c)(5) is to provide an aft ultimate load condition for substantiation of the support structure for retention of significant items of mass above and forward of the occupied volume(s) of the rotorcraft that could otherwise come loose and injure an occupant in an emergency landing. Amendment 29-38 also increases the forward, sideward, and downward emergency load factors of § 29.561(c)(2), (c)(3), and (c)(4), respectively, for retention of items of mass above and behind the occupied volume(s) that could otherwise come loose and injure an occupant in an emergency landing.

Page C - 54 5/1/2014 AC 29-2C, Chg 4 b. Procedures . The procedures in sections 29.561 and 29.561A of this AC continue to apply except the newly specified load factors must be used. A list of the significant items of mass to be considered should be compiled by the applicant and approved by the certifying authority.

Note: For doors and emergency exit design, when applicable, the rearward load factor to consider is in § 29.561(b)(3)(v).

Page C – 54.1 AC 29-2C, Chg 4 5/1/2014 [This page intentionally blank.]

Page C – 54.2 5/1/2014 AC 29-2C, Chg 4 AC 29.562 § 29.562 (Amendment 29-29) EMERGENCY LANDING DYNAMIC CONDITIONS.

a. Explanation . Amendment 29 -29 adds new requirements for the dynamic testing of all seats in rotorcraft. This paragraph is rewritten to incorporate the guidance previously documented in AC 20-137 dated 3/30/92.

b. Background . Improved occupant restraint in civil rotorcraft is addressed in Amendments 29-29 to the airworthiness standards, which add two dynamic crash impact design conditions for seat and occupant restraint systems. This amendment also prescribes a shoulder harness for each occupant and adopts human impact injury criteria as a measure for occupant protection for the dynamic crash impact conditions.

In addition, these amendments significantly improve occupant protection for normal category rotorcraft in a survivable emergency landing. This advisory material addresses the dynamic test conditions and the related pass-fail injury criteria for the dynamic test conditions. This material pertains to single as well as multiple seats and tandem arrangements of the seats in rotorcraft.

(1) Dynamic test methods . This guidance focuses on the use of dynamic test methods for evaluating the performance of rotorcraft seats, restraint systems, and certain related interior systems for demonstrating structural strength and the ability of those systems to protect an occupant from possible injuries in an emergency landing environment represented by the standard. These test methods differ from static test methods, which are limited to demonstrating only the structural strength of the seat or restraint system under ultimate load for at least 3 seconds. This guidance contains sources for appropriate test procedures and provides some insight into the logic of these procedures. It also defines, in part, test facility and equipment characteristics necessary for conducting these tests.

(2) Standardized test methods . Dynamic tests are often conducted at a specially equipped facility, one other than that owned by the designer or manufacturer of the test article. To obtain consistent test results, it is necessary to specify the critical test procedures in detail in the test plan, and then carefully follow these procedures when conducting the tests. This guidance defines certain critical procedures for accomplishing the tests of the seat and restraint systems and assessing the data obtained in the tests. Many of these procedures are accepted as standards by government and commercial test facilities and have been modified in this guidance only as necessary for the specific testing of rotorcraft systems.

(3) Relationship of dynamic tests to design standards . This guidance describes test procedures useful in assessing the performance of a rotorcraft seat, Page C - 55 AC 29-2C, Chg 1 2/12/03 restraint system, and interior system. However, it is impractical to conduct sufficient tests for assessing the performance of the system throughout its entire range of possible uses in unique interior arrangements. The seat, restraint system, and related interior system should be designed for the range of occupants and environments for which it is expected to perform, not just for the dynamic test conditions described in this guidance.

(i) Occupant size . The dynamic tests are conducted with a specific, th acceptable, standard anthropomorphic test dummy (ATD) representing a 50 percentile male occupant. Energy absorbing systems, restraint system loads and anchorage locations, seat adjustments, seat pitch (for multiple seat rows), head strike envelopes, etc., are typical factors directly influenced by occupant size.

(ii) Seat position and location . The tests should be sufficient to represent the range of performance expected of a seat and restraint system. A seat, especially an adjustable flight crew seat, should be qualified for those positions approved for take-off and landing. As with static test procedures the seat is also tested to the most critical condition for the dynamic tests. For an adjustable flight crew seat, as an example, the full-up position and longitudinal impact case are expected to be the critical condition.

But these dynamic tests and occupant injury assessment provide a systems approach to qualification. It is therefore necessary to test adjustable seats at the design position for the ATD. Two tests would be required to demonstrate compliance with the strength standards and with the occupant injury criteria. Alternatively adjusting the flight crew seat to its highest position with the interior features, such as an instrument panel shield, raised to maintain the proper perspective or relation to the ATD, is considered an acceptable test procedure for demonstrating compliance with the structural and occupant injury requirements for the seat and its location in a particular cockpit arrangement.

(iii) Test conditions . Only two minimum impact tests are described in the dynamic test procedures discussed in this guidance. These procedures address the tests needed to demonstrate compliance for a typical seat and restraint system installation. Additional tests may be necessary to demonstrate compliance for other types or variations of seat and restraint system installations. For example, while only one lateral load direction is specified in the tests, the system should perform properly when similarly loaded from either side.

(iv) Floor deformation . The test procedures require evaluating the effect of certain sidewall or floor deformation. The seat and its attachments and the restraint system should also perform properly if no floor deformation is present.

(v) Head impact . Should such contact occur, head impact with a seat back or the interior of the rotorcraft is evaluated by using a Head Injury Criterion (HIC), which can be measured directly in the tests discussed in this guidance or in supplementary tests of the interior. The design of the interior should protect the head Page C - 56 2/12/03 AC 29-2C, Chg 1 (vi) Emergency egress . Standards for emergency evacuation of the rotorcraft are contained in FAR Part 29. The objective is to allow each occupant to leave the seat and rapidly evacuate the rotorcraft using an exit on either side of the rotorcraft.

c. Dynamic Test Methods and Facilities .

(1) General . A minimum of two dynamic tests are used to assess the performance of the rotorcraft seat, restraint system, and related interior system. The seat, the restraint, and the nearby interior all function together as a system to protect the occupant during emergency landing. The specific test conditions are shown in Figure AC 29.562-1. Explanations of the test conditions are as follows: (i) Test 1 . The test determines the protection provided when the impact environment is such that the resulting predominant impact load component (vertical) is directed along the spinal column of the occupant in combination with a horizontal (longitudinal) component. Protection against spinal injury is important and it may be necessary to provide energy absorbing (load limiting) or attenuation capability in the seat system in order to comply with the human injury criteria specified in § 29.562 (c)(7).

(ii) Test 2 . The test determines the protection provided in an impact where the predominant impact load component is in the longitudinal direction in combination with a lateral component. Evaluation of head injury protection is important in this test if the head could strike some interior portion of the rotorcraft or a forward seat. Chest or spinal column injury, which might result from the upper torso restraint, is also evaluated in this test.

(iii) Tests 1 and 2 . These test conditions are also significant for the structural strength of the system. Both tests should be used to assess submarining (where the seat belt slips above the ATD pelvis) and rollout of the upper torso restraint system particularly with single, diagonal torso restraint belts. Since external crash forces frequently cause significant structural deformation, simulated floor deformation is specified for the tests to prove the seat design can accommodate the relative deformation between the seat and the floor or sidewall and still function without imposing excessive loads on the seat, the attachment fittings, or floor tracks.

Page C - 57 AC 29-2C, Chg 1 2/12/03 Illustration shows Test 1 Test 2 a forward facing seat ATD inertial ~~~ load shown by arrow:

f,i-

•

•

Min . fps 30 42 V1, Max. sec. . 031 tr , . 071 Min. G 18.4 Defont floor: degrees roll 10 10 degrees pitch 10 10 t r time ....

0-tr-7"'7":l~ll'"r-~~--- • deceleration ' tr = rise the

v1 = tnpact velocity

G FIGURE AC 29 .562-1 . Seat Restraint System Dynamic Tests Normal and Transport Category Rotorcraft Page C - 58 2/12/03 AC 29-2C, Chg 1 (2) Test facilities . A test proposal is prepared for certification authority approval and should reflect the capability of the facility. It should be noted that a number of test facilities could be used to accomplish dynamic testing. Test facilities can be grouped into categories based on the method they use to generate the impact pulse (i.e., accelerators, decelerators, or impact with rebound) and whether the facility is a horizontal (sled) design or a vertical (drop tower) arrangement. As in all certification compliance tests, a test proposal, which may refer to certain specific or generic test equipment, is approved prior to testing. The test may be conducted anywhere, within certain availability or mutually convenient constraints, as long as the test is conducted in accordance with the approved test plan and properly witnessed.

(i) Facility Characteristics or Features . Each of the facilities has characteristics that may have advantages or disadvantages with regard to the dynamic tests discussed in this guidance. One concern is the rapid sequence of acceleration and deceleration that must take place in the tests. In a landing impact, the acceleration phase (flight) is gradual and usually well separated in time from the deceleration (crash impact) phase. In a test, the deceleration usually closely follows the acceleration.

When assessing the use of a facility for the specific test procedures outlined in the recommendations, it is necessary to assess the possible consequences of this rapid sequence of acceleration and deceleration on the test articles and ATD. The standard accommodates the different facilities that are or may be available for the applicant’s use. That is, the standards dictate the peak acceleration with a tolerance as stated in this AC. The “decay” in deceleration with respect to time is not dictated, thereby allowing for the different test facility equipment characteristics.

(A) Deceleration sled facilities . In an aircraft crash, the impact takes place as a deceleration, so loads are applied more naturally in test facilities that create the test impact pulse as a deceleration. Since it is simpler to design test facilities to extract energy in a controlled manner than to impart energy in a controlled manner, several different deceleration sled facilities can be found. The deceleration sled facility at the FAA’s Civil Aeromedical Institute (CAMI) was referred to in developing the test procedures discussed in this and similar AC’s related to airplanes.

( 1 ) The Acceleration Phase. Sufficient velocity for the test impact pulse acquired in this phase can distort the test results if the acceleration is so high that the test articles or ATD are moved from their intended pre-test position. This inability to control the initial or onset conditions of the test would directly affect the test results.

This can be avoided by using a lower acceleration for a relatively long duration and by providing a coast phase (in which the acceleration or deceleration is nearly zero) prior to the impact. This allows any dynamic oscillation in the test articles or the ATD that might be caused by the acceleration to decay. To guard against errors in data caused by pre-impact accelerations, data from the electronic test measurements (accelerations, loads) should be reviewed for the time period just before the test impact pulse to make sure all measurements are at the baseline (zero) level. Photometry film taken of the Page C - 59 AC 29-2C, Chg 1 2/12/03 test should also be reviewed to make certain that the ATD’s used in the test and the test articles were all in their proper position prior to the test impact pulse.

( 2 ) Orientation of test article. The horizontal test facility readily accommodates forward-facing seats in both tests discussed in this guidance, but problems can exist in positioning the test ATD in Test 1 if the seat is a rear or side- facing seat. In these cases, the ATD’s tend to fall out of the seat due to the force of gravity and must be restrained in place using breakaway tape, cords, or strings. Since each installation will present its own problems, there is no simple, generally applicable, guidance. Attention should be given to positioning the ATD against the seat back and to proper positioning of the ATD’s arms and legs. It will probably be necessary to build special supports for a breakaway restraint so that the restraint will not interfere with the function of the seat and occupant restraint system during the test. Photos of the test from “side of track cameras” should be reviewed to make sure that the breakaway restraint did break (or become slack) in a manner that did not unduly influence the motion of the ATD or the test articles during the test.

(B) Acceleration sled facilities. Acceleration sled facilities, usually based on the Hydraulically Controlled Gas Energized (HYGE) accelerator device, provide the impact test pulse as a controlled acceleration at the beginning of the test. The test item and the ATD are installed facing in the opposite direction from the velocity vector, opposite from the direction used on a deceleration facility, to account for the change in direction of the impact. There should be no problem with the ATD or the test items being out of position due to pre-impact sled acceleration, since there is no sled movement prior to the impact test pulse. Because of this characteristic the applicant may prefer this type of a facility.

( 1 ) Test pulse. After the impact test pulse, when the sled is moving at the maximum test velocity, stop the sled safely. Most of the facilities of this design have limited track length available for deceleration, so that the deceleration levels can be relatively high and deceleration may begin immediately after the impact test pulse.

Since the maximum response of the system usually follows (in time) the impact test pulse, any sled deceleration, which takes place during that response will affect the response and change the test results. The magnitude of change depends on the system being tested, so that no general “correction factor” can be specified. The effect can be minimized if the sled is allowed to coast, without significant deceleration, until the response is complete.

( 2 ) Test results. If the seat or restraint system experiences a structural failure during the test pulse, the post impact deceleration can increase the damage and perhaps result in failures of unrelated components. This will complicate the determination of the initial failure mode and make product improvement more difficult.

One other consideration is that the photometry film coverage of the response to impact test pulse must be accomplished when the sled is moving at near maximum velocity.

Onboard cameras or a series of trackside cameras are usually used to provide film coverage of the test. Since onboard cameras frequently use a wide-angle lens placed Page C - 60 2/12/03 AC 29-2C, Chg 1 close to the test items, it is necessary to account for the effects of distortion and parallax when analyzing the film. The acceleration sled facility faces the same problems in accommodating rearward-facing or side-facing seats in Test 1 as the deceleration sled facility, and the corrective action is the same for both facilities.

(C) Impact-with-rebound sled facilities . One other type of horizontal test facility used is the “impact-with-rebound” sled facility. On this facility, the impact takes place as the moving sled contacts a braking system, which stores the energy of the impact, and then returns the stored energy back to the sled, causing it to rebound in the opposite direction. This facility has an advantage over acceleration or deceleration facilities in that only one-half of the required velocity for the impact would need to be generated by the facility (assuming 100 percent efficiency). Thus the track length can be shortened, and the method of generating velocity is simplified. The disadvantages of this facility combine the problems mentioned for both acceleration facilities and deceleration facilities. Since one of the reasons for this type of facility is to allow short track length to be used, it may be difficult to obtain sufficiently low acceleration just before or after the impact pulse to resolve data error problems caused by significant pre-impact and post-impact accelerations.

(D) Drop towers . Vertical test facilities can include both drop towers (decelerators) and vertical accelerators. Vertical accelerators, which can produce a longer duration/displacement impact pulse, may not be available. However, drop towers are one of the easiest facilities to build and operate and are frequently used.

( 1 ) Acceleration phase. In these facilities, the pull of earth’s gravity is used to accelerate the sled or guided test fixture and test article to specified impact velocity to avoid the use of a complex mechanical accelerating system. Reproducing the required impact pulse may require extensive development tests for the facility.

Unfortunately, these facilities are more difficult to use for conducting Test 2, particularly for typical forward-facing seats.

( 2 ) Test article. In preparing for (longitudinal) Test 2, the seat should be installed at an angle according to the standards such that the ATD’s tend to fall from the seat due to gravity. The restraint system being tested cannot hold the ATD against the seat unless tightened excessively and will not usually locate the head, arms, or legs in their proper position relative to the seat. Design and fabrication of an auxiliary “break-away” ATD positioning restraint system just for this test are usually a complex task. The auxiliary restraint should not only position the ATD against the seat (including maintaining proper seat cushion deflection) during the pre-release condition of 1 g, it should also maintain the ATD in that proper position during the free fall to impact velocity when the system is exposed to zero g, and then it should release the ATD in a manner that does not interfere with the ATD response to impact. The usual sequence of 1 g/0 g impact, without the possibility of a useful “coast” phase, as done in horizontal facilities, causes shifts in initial conditions for the test impact pulse that can affect the response to the impact. The significance of this undesired movement will depend on Page C - 61 AC 29-2C, Chg 1 2/12/03 the dynamic characteristics of the system under test, and these characteristics are seldom known with sufficient accuracy to achieve the response initially.

( 3 ) Other facets. In addition, the earth’s gravity will oppose the final rebound of the ATD into the seat back, so that an adequate test of seat back strength and support for the ATD cannot be obtained. The problems in Test 1, or with rear-facing seats in Test 2, are not as difficult because the seat will support the ATD prior to the free fall. However, the zero g condition free fall that exists prior to impact will allow the ATD to “float” in the seat restraint system, perhaps changing position and certainly changing the initial impact conditions if movement occurs. Again, the development of a satisfactory auxiliary breakaway restraint system to assure correct pre-impact condition is difficult.

(ii) Test Fixtures. A test fixture is usually required to position the test article on the sled or drop carriage of the test facility and to represent the aircraft’s structure floor, sidewall, bulkhead, etc. It holds the attachment fittings or floor tracks for the seat, provides the floor and sidewall deformation needed for the test, and provides a floor or footrest for the ATD, and it positions the pertinent interior items, such as instrument panels, sidewalls, bulkheads, a second row of seats, if required, for successful performance of the tests, and otherwise simulates the rotorcraft for the test.

The text fixture is usually fabricated of heavy structural steel and does not necessarily simulate lightweight aircraft design or construction. The details of the fixture will depend upon the requirements of the test articles, but provisions for the specified floor and sidewall deformation are needed.

(A) Purpose of floor or sidewall deformation. The purpose of using pitch and roll deformation for the tests is to demonstrate that the seat/restraint system will remain attached to the airframe and perform properly for the tests, although the structure and seat may be more severely deformed by the forces associated with a particular crash. Typical design deficiencies addressed by the test conditions include, but are not limited to, the following: ( 1 ) Concentrated loads may be imposed on floor fittings (studs) or tracks by seat leg attachment fittings which fit tightly or are clamped to a track or fitting, and which do not have some form of relief (especially lateral roll relief) incorporated in the design. These joint fittings can concentrate the forces on one lip of the floor and sidewall track or stud and may break the joint (track or the fitting).

( 2 ) Similarly, loads can be concentrated on one edge of a floor track or stud fitting having an “I,” “bulb head” or “mushroom” cross section and may prematurely break the flange or the fitting.

( 3 ) Detents, pins, or collars which lock the seat leg fitting to the floor track can become disengaged, or the mechanism which is used to disengage the detents, pins, or “dogs” can be actuated and release the seat as the seat or airframe deforms.

Page C - 62 2/12/03 AC 29-2C, Chg 1 ( 4 ) Seat assemblies that provide an energy absorbing system between a seat “bucket or pan” and a seat frame attached to the floor may not perform properly for a pre-loaded seat frame attached to the floor or sidewall. Deformation of the seat frame may cause the energy absorbers to receive unanticipated loads or cause excessive friction in the guides between the seat bucket and seat frame to lock the energy absorber in place.

( 5 ) Restraint system anchorages attached to the airframe structure may be significantly displaced relative to the seat if the seat deforms during the test, and that displacement may inhibit proper performance of the seat/restraint system. This is especially critical for the necessary vertical stroking or displacement.

(B) Floor Deformation. The pitch and roll displacement is intended to evaluate the track or stud and leg fitting joint (axis) tolerance to angular misalignment and not necessarily axis translational displacement.

( 1 ) For the typical aircraft seat. For a multiple or single person seat, with four seat legs mounted in the airframe on two parallel tracks, the floor deformation test fixture may consist of two parallel beams, a “pitch beam” which pivots about a lateral (y) axis, and a “roll beam” which pivots about a longitudinal (x) axis. The beams can be made of any fairly rigid structural form, box, I-beam, channel, or other appropriate cross section. The pitch beam should be capable of rotating in the x-z plane up to +/- 10° relative to the longitudinal (x) axis. The roll beam should be capable of a +/- 10° roll about the axis of the seat attachment/fitting joint (centerline of floor track or fittings).

(See Figure 29.562-2 for a schematic of an installation.) A means should be provided to fasten the beams in the deformed positions.

( 2 ) Seat and floor interface. The beams should have provision for installing floor tracks or other attachment fittings on their upper surface in a manner that does not alter the above-floor strength of the track or fitting. The track or other attachment fittings should be representative (in above-floor configuration shape and strength) of those used in the rotorcraft. Structural elements below the surface of the floor that are not considered part of the floor track or fitting may be omitted in the installation. The seat having four legs should then be installed on the beams so that the rear seat leg attachment point is near the pitch beam axis of rotation, and the seat positioning pins or locks are fastened in the same manner as specified in the test proposal and as would be used in the rotorcraft, including the adjustment of “anti-rattle” mechanisms, if employed.

( 3 ) Test set-up. The remainder of the test preparations would then be completed (ATD installation and positioning, instrumentation installation, adjustment and calibration, camera checks, etc.). The “floor deformation” would be induced as the final action before the test is accomplished. The roll beam should first be rotated 10° and locked in place, and then the pitch beam should be rotated 10° and locked in place.

The direction of rotation would be selected to produce the most critical loading condition on the seat and floor track or fitting. If the seat is fairly flexible, it may be possible to Page C - 63 AC 29-2C, Chg 1 2/12/03 rotate the beams by manual effort, perhaps using removable pry bars to gain mechanical advantage. However, rotation of the beams used for testing a stiff seat frame is likely to require greater effort than can be accomplished manually, and the use of removable hydraulic jacks or other devices may be necessary. If this condition is expected, provision should be made for appropriate loading points when designing the fixture. This condition is most likely to be encountered when rotating the pitch beam.

The test facility personnel should adhere to appropriate safety provisions during the deformation process. The test fixture may be designed to adjust to fit a wide range of seat designs, including leg spacing, that may be encountered.

Page C - 64 2/12/03 AC 29-2C, Chg 1 ORIGINAL POSITION NOTE: FOR CLARITY, LOAD TRANS DUCERS ARE NOT SHOWN POSITION ., AFTER FLOOR DEFORMATION PITCH AXIS UNDER REAR LEG ATTACHMENT

o/J

MINIMUM ~EAT TRACK ~-. 100 ROLL TO BE TlSTED •

LEFT OR RIGHT- ~

DETAIL (A) EXAMPLE OF AIRCRAFT TRACK BEAM FIGURE AC 29.562-2 – Schematic Floor Deformation Fixture; Seat Legs Attached at Floor Level Page C - 65 AC 29-2C, Chg 1 2/12/03 (C) Alternative configurations. The preceding discussion described the fixture and floor deformation procedure that would be used for a typical seat that has four seat legs and four attachments to the fuselage floor. These test procedures may be adapted to seats having other designs. Special test fixtures may be necessary for different configurations. The following methods, while not covering all possible seat designs, provide guidance for the more common configurations of seats: ( 1 ) Rotorcraft seats with three legs may have one central leg in front or back of the seat and one leg on each side of the seat. The central leg should be held in its undeformed position as pitch deformation is applied to one side leg and roll to the other.

( 2 ) Seats that are “integral” with the structure without floor or sidewall attachment devices and with continuous attachments such as rows of rivets or screw, etc., are excluded from the deformation, misalignment, or preload prior to test impact.

Similarly bulkhead-mounted seats, solely mounted to a bulkhead, are excluded from the deformation requirement. The test fixture could represent the seat and structure or a rigid bulkhead or an actual bulkhead panel. If a rigid bulkhead installation is used, the test fixture should transfer loads to the seat restraint system through components equivalent to the seat attachment fittings and surrounding bulkhead panel, which exist in the actual installation. Similar guidelines apply to integral seats.

( 3 ) Seats that are attached to both the floor and a bulkhead would be tested on a fixture that positions the bulkhead surface in a plane through the axis of rotation of the pitch beam. The bulkhead surface should be located perpendicular to the plane of the floor (the rotorcraft floor surface, if one were present) in the undeformed condition or in a manner appropriate to the intended installation. Either a rigid bulkhead simulation or replica or an actual bulkhead panel may be used. If a rigid bulkhead simulation is used, the test fixture should transfer loads to the seat restraint system through components equivalent to the attachment fitting and surrounding bulkhead panel that would exist in the actual installation. The seats would be attached to the bulkhead and the floor in a manner representative of the rotorcraft installation, and the floor, as represented in the test, would then be deformed as described in paragraph b.(2)(ii)(B).

( 4 ) Seats mounted between fuselage sidewalls or to the sidewall and floor of an airplane should be tested in a manner simulating rotorcraft fuselage cross-section deformation (e.g., from circular or rectangular to flattened circular or rectangular or ellipsoidal shape) during a severe impact. The 10° roll would simulate the change in fuselage shape. Brackets should be fabricated to attach the seat to the sidewall test fixture at the same level above the fixture “floor” that would represent the installation above the rotorcraft floor. The sidewall bracket or rail should be located on the “roll” beam. It is envisaged that the sidewall rolls outward 10° about an axis at the floor and sidewall juncture. Then, as the beams are rotated to produce the critical loading condition, the combined angular and translational deformation would simulate the Page C - 66 2/12/03 AC 29-2C, Chg 1 deformation at the sidewall attachment during a landing impact. (See Figure 29.562-3 for a schematic of an installation.)

( 5 ) Seats that are cantilevered from one sidewall without connection to another structure would not be subject to floor deformation. However, sidewall deformation is likely, and should be considered by warping the entire sidewall attachment plane, or the attachment points of the seat, 10° to represent the most likely fuselage sidewall deformation. This is intended to evaluate a critical condition for seat attachment or seat and occupant restraint system performance. Either a rigid sidewall simulation or an actual sidewall panel may be used. If a rigid sidewall simulation is used, the test fixture should transfer loads to the seat through components equivalent to the attachment fitting as well as the surrounding sidewall to replicate the actual installation.

( 6 ) Side-facing seats, occupiable for takeoff and landing, are subject to the specified dynamic test conditions. Compliance with the structural requirements should be demonstrated for side-facing seats using the same conditions for the test and pass/fail criteria as for fore- and aft-facing seats. The seat should be loaded in the most critical case structurally. Means of restraining the ATDs may need to be adapted to ensure adequate retention during the test. The application of floor distortion will need to be assessed on an individual basis, depending on the design and the method of attaching the seat.

( 7 ) A seat assembly for multiple occupants may have more than two pairs of legs. If the assembly uses a uniform cross section, deformation of only the outer leg assemblies is sufficient. The inner leg pairs may be maintained in the normal, undeformed position for the dynamic tests.

(D) Multiple Row Test Fixtures. In tests of passenger seats normally installed in rows in a rotorcraft, head impact conditions should be evaluated by tests using at least two rows of seats. This allows direct measurements of the head injury data if secondary head impact occurs and demonstrates the effect of the interaction loads between rows; e.g., due to occupant contact with the front row. (That is, ATD leg contact does not overload the front row.) These conditions are usually critical only on Test 2. The single seat row fixture used for the test should be used to position the front (first) seat row and provide appropriate floor deformation to that row. The test is critical for the first row strength. An additional simple fixture may position the second seat row in the proper location and need not provide floor deformation. The second row should be fully occupied unless it is not as critical a condition for the first seat row.

Representative seat cushions and torso restraint systems should be used on both seat rows. The allowable seat pitch (longitudinal spacing) can be determined by analysis of previous test data or limited by type design data and information for the most critical condition for head or leg impact against relatively stiff structure in the first seat row.

Operational limitations that specify the allowable seat pitch of the seats in rotorcraft may be considered also. No impact surface such as seats, bulkheads, etc., may be needed Page C - 67 AC 29-2C, Chg 1 2/12/03 for the ATD in the first seat row unless such a surface is within the expected head strike envelope whenever the seats are installed in rotorcraft.

Page C - 68 2/12/03 AC 29-2C, Chg 1 ORIGINAL POSITION NOTE : "' POSITION FOR CLARITY, AFTER FLOOR LOAD TRANSDUCERS DEFORMATION ARE NOT SHOWN REPRESENTATION SEAT TO SIDEWALL AT'fACHMENT IGID TEST STRUCTURE PITCH AXIS UNDER REAR LEG ATTACHMENT FIGURE AC 29.562-3 – Schematic Test Fixture; Side Wall Mounted Seat Page C - 69 AC 29-2C, Chg 1 2/12/03 (E) Other fixture applications. Test fixtures should provide a flat footrest for an ATD used in tests of passenger and attendant seats. Flightcrew seats associated with special foot rests or foot-operated controls may use simulated footrests. The surface of the footrest should be covered with carpet (or other appropriate material) and be at a position representative of the undeformed floor or control. Test fixtures may also be necessary to provide guides or anchors for torso restraint systems or for holding instrument panels or bulkheads if necessary for the proposed tests. If these provisions are necessary, the installation should represent the configuration of the installation and be of adequate structural strength to withstand the expected test loads.

(iii) Instrumentation. Electronic and photographic instrumentation systems are essential to properly record the information for the tests discussed in this AC.

Electronic instrumentation is used to measure accelerations and forces required for verifying the test environment and for measuring most of the pass/fail criteria and the floor (seat) attach loads. Photographic instrumentation is used for recording the overall qualitative results of the tests, for confirming that the lap safety belt remained on the ATD’s pelvis (no submarining), and that the upper torso restraint straps remained on the ATD’s shoulder, and for recording the relative deformation of the seats as it may influence rapid evacuation of the rotorcraft by the occupants. Paragraph d.(10), of this guidance contains allowable seat deformation information related to an aisle, passageway, access to exits, and so forth.

(A) Electronic instrumentation. Electronic instrumentation should be accomplished in accordance with the Society of Automotive Engineers Recommended Practice SAE J211, Instrumentation for Impact Tests. In this practice, a data channel is considered to include all of the instrumentation components from the transducer through the final data measurement, including connecting cables and any analytical procedures that could alter the magnitude or frequency content of the data. Each dynamic data channel is assigned a nominal channel “class” equivalent to the high frequency limit for that channel, based on a constant output/input ratio vs. frequency response plot which begins at 0.1 Hz (+1/2 to –1/2 db) and extends to the high frequency limit (+1/2 to – 1 db). Frequency response characteristics beyond this high frequency limit are also specified. When digitizing data, the sample rate should be at least five times the –3 db cutoff frequency of the pre-sample analog filters. Since most facilities set all pre-sample analog filters for Channel Class 1,000 and since the –3 db cutoff frequency for Channel Class 1,000 is 1,650 Hz, the minimum digital sampling rate would be about 8,000 samples per second. For the dynamic tests discussed in this guidance, the dynamic data channels should comply with the following channel class characteristics: ( 1 ) Sled or drop tower vehicle acceleration should be measured in accordance with the requirements of Channel Class 60, unless the acceleration is also integrated to obtain velocity or displacement, in which case, it should be measured in accordance with the Channel Class 180 requirements.

( 2 ) Belt restraint system loads should be measured in accordance with the requirements of Channel Class 60.

Page C - 70 2/12/03 AC 29-2C, Chg 1 ( 3 ) ATD head accelerations used for calculating the HIC should be measured in accordance with the requirements of Channel Class 1,000.

( 4 ) ATD femur forces may be measured if desired in accordance with Channel Class 600.

( 5 ) ATD pelvic/lumbar spinal column force should be measured in accordance with the requirements of Channel Class 600.

( 6 ) The full-scale calibration range for each channel should provide sufficient dynamic range for the data being measured.

( 7 ) Digital conversion of analog data should provide sample resolution of not less than 1 percent of full-scale input.

(B) Photographic instrumentation. Photographic instrumentation is used for documenting the response of the ATD and the test items to the dynamic test environment. Both high speed motion picture and still systems are used.

( 1 ) High-speed motion picture cameras that provide data used to calculate displacement or velocity should operate at a nominal speed of 1,000 pictures per second. Photo instrumentation methods should not be used for measurement of acceleration. The locations of the cameras and of targets or targeted measuring points within the field of view should be measured and documented. Targets should be at least 1/100 of the field width covered by the camera and should be of contrasting colors or should contrast with their background. The center of the target should be easily discernible. Rectilinearity of the image should be documented. If the image is not rectilinear, appropriate correction factors should be used in the data analysis process.

A description of photographic calibration boards or scales within the camera field of view, the camera lens focal length, and the make and model of each camera and lens should be documented for each test. Appropriate digital or serial timing should be provided on the image media. A description of the timing signal, the offset of timing signal to the image, and the means of correlating the time of the image with the time of electronic data should be provided. A rigorous, verified analytical procedure should be used for data analysis.

( 2 ) Cameras operating at a nominal rate of 200 pictures per second or greater can be used to document the response of ATD and test items if measurements are not required. For example, actions such as movement of the pelvic restraint system webbing (lap safety belt) off of the ATD pelvis or movement of upper torso restraint webbing off of the ATD’s shoulder can be observed by documentation cameras placed to obtain a “best view” of the anticipated event. These cameras should be provided with appropriate timing and a means of correlating the image with the time of electronic data.

Page C - 71 AC 29-2C, Chg 1 2/12/03 ( 3 ) Still image cameras can be used to document the pretest installation and the posttest response of the ATD’s and the test items. At least four pictures should be obtained from different positions around the test items in pretest and posttest conditions. Where an upper torso restraint system is installed, posttest pictures should be obtained before moving the ATD. For the posttest pictures, the ATD’s upper torso may be rotated to the approximate upright seated position so that the condition of the restraint system may be better documented, but no other change to the posttest response of the test item or ATD’s should be made. The pictures should document that the seat remained attached at all point of attachment to the test fixture. Still pictures can also be used to document posttest yielding of the seat for the purpose of showing that it would not impede the rapid evacuation of the airplane occupants. The ATD’s should be removed from the seat in preparation for still pictures used for that purpose.

Targets or an appropriate target grid should be included in such pictures, and the views should be selected so that potential interference with the evacuation process can be determined. For tests where the ATD’s head impacts a fixture or another seat back, pictures should be taken to document the head contact areas.

(iv) ATD . The tests discussed in this guidance were developed using modified forms of the ATD specified by the United States Code of Federal Regulations, th Title 49, Part 572 Anthropomorphic Test Dummies, Subpart B – 50 Percentile Male.

These “Part 572B” ATD’s were developed for automobile impact testing and have been shown to be reliable test devices capable of providing reproducible results in repeated testing. However, since ATD development is a continuing process, the standards allow use of “equivalent” dummies. See paragraph c.(2)(iv)(D) of this guidance. Dummy types should not be mixed when the tests discussed in this guidance are performed.

(A) Modification for measuring pelvic/lumbar column load . Since ATD’s have been developed for use in automobile testing to evaluate injury protection in forward, rearward, and sideward impacts, the ATD’s must be modified to measure the spinal load to comply with the § 29.562(c)(7). This load is influenced by a vertical direction component and by upper torso restraints which may produce a downward force component on the shoulders. To measure the load, a load (force) transducer is inserted into the ATD pelvis just below the lumbar column. This modification is shown in Figure 29.562-4. A commercially available “femur” load cell with end plates removed has been adapted to the modified ATD to measure the compression load between the pelvis and the lumbar spine column of the ATD. A “femur” load cell is commonly available to most test facilities and (according to specifications) is insensitive to bending and twisting moments. This feature prevents load transmission through the load cell as it measures the ATD lumbar/pelvis compression forces. To maintain the correct seated height of the ATD, the load cell is fixed in a rigid cup inserted into a hole bored in the top surface of the ATD pelvis, the top flange of which is bolted to the pelvis. If necessary, ballast should be added to the pelvis to maintain the specified weight of the assembly.

Alternative approaches to measuring the axial force transmitted to the lumbar spinal column by the pelvis are acceptable if the method— Page C - 72 2/12/03 AC 29-2C, Chg 1 ( 1 ) Accurately measures the axial force but is insensitive to moments and forces other than that being measured; ( 2 ) Maintains the intended alignment of the spinal column and the pelvis, the correct seated height, and the correct weight distribution of the ATD; and ( 3 ) Does not alter the other performance characteristics of the ATD.

FIGURE AC 29.562-4 – Installation of Pelvic—Lumbar Spine Load Cell In Part 572B Anthropomorphic Dummy.

(B) Figure 29.562-4 shows an acceptable installation of a femur load cell (d) at the base of the ATD lumbar spine (a). The load cell is in line with the centerline of the lumbar spine and set below the top surface of the pelvis casting to maintain the seated height of the ATD. A rigid adapter cup (e) is fabricated to hold the load cell, and a hole is bored in the ATD pelvis to accept the cup. Provide clearance between the walls of the adapter cup and the load cell and the wires leading from the cell to avoid possible interference loads. The bottom of the load cell is bolted to the adapter cup.

Adapter plates having similar hole patterns in their periphery are fabricated for the lower surface of the lumbar spine (b) and the upper surface of the load cell (c). These plates are fastened to the lumbar spine and load cell with screws through holes matching threaded holes in those components and are then joined together by bolts through the peripheral holes. The flange on the adapter cup has a bolt hole pattern matching that on the pelvis. The cup is fastened to the pelvis using screws to the threaded holes in the pelvis. Spacers (f) may be placed under the flange of the cup to obtain the specified ATD seating height. Additional weight should be placed in the cavity below the adapter Page C - 73 AC 29-2C, Chg 1 2/12/03 cup to compensate for any weight lost because of this modification. The instrument cavity plug (g) is cut to provide clearance for the adapter cup and added weight.

(C) Other ATD modifications . Flailing of the ATD arms often causes the “clavicle” used in the Part 572B ATD to break. To reduce the frequency of this failure, the clavicle may be replaced by a component having the same shape but made of higher strength material. This may increase the ATD weight slightly, but it would be acceptable for the tests discussed in this guidance. Another useful modification is the use of “submarining indicators” on the ATD pelvis. These electronic transducers are located on the anterior surface of the ilium of the ATD pelvis without altering its contour and indicate the position of the lap safety belt as it applies loads to the pelvis. Thus they can provide a direct record that the lap safety belt remains on the pelvis during the test and eliminates the need for careful review of high-speed camera images to make that determination.

(D) Equivalent ATD . The continuing development of ATD for dynamic testing of seat restraint/crash-injury-protection systems is guided by goals of improved biofidelity (human-like response to the impact environment) and reproducibility of test results. The following criteria can be used to assess whether or not an ATD is equivalent to the present Part 572B ATD: ( 1 ) Fabrication in accordance with design and production specifications established and published by a regulatory agency responsible for crash injury protection systems; ( 2 ) Capability of providing data for the measurements discussed in this guidance or of being readily altered to provide the data; ( 3 ) Evaluation by comparison with the Part 572B ATD and shown to generate similar response to the impact environment discussed in this guidance; and ( 4 ) Any deviations from the Part 572B ATD configuration or performance are representative of the occupant of a civil aircraft in the impact environment discussed in this guidance.

(E) Temperature and humidity . Since extremes of temperature and humidity can change the performance of ATD, the tests discussed in this guidance should be conducted at a temperature from 66° F to 78° F, and at a relative humidity from 10 percent to 70 percent. The ATD should have been maintained under these conditions for at least 4 hours prior to the test.

(3) Test Preparation . Preparations for the tests should include selection of the test articles to be used in the tests, determination of the “most critical” conditions for the tests, and installation of the test articles, instrumentation, and ATD on the test fixture.

Preparations pertaining to the normal operation of the test facility, such as safety Page C - 74 2/12/03 AC 29-2C, Chg 1 provisions and the actual procedure for accomplishment of the tests, are particular to the test facility. These may be included in a test proposal or plan.

(i) Selection of test articles . Many seat designs compose a “family or type” of seats which have the same basic structural design but differ in detail. For example, a basic seat frame configuration can allow for several different seat leg locations to permit installation in different airplanes. If these differences are of such a nature that their effect can be determined by rational analysis, then the analysis can determine the most highly stressed (“most critical”) configuration. The most highly stressed configuration would normally be selected for the dynamic tests so that the other configurations could be accepted by analysis and comparison with that configuration. The HIC depends on head impact (secondary impact after rotorcraft ground impact) and is more dependent on seat pitch for multiple row seats and on location for others than on seat structural stress for a given “family” of seats, so that the selection of the most highly stressed seat structure and the most critical seat pitch or location will permit these factors to be evaluated in one dual row test under the conditions of Test 2. Critical pelvic/lumbar spinal column forces are usually found under the vertical impact conditions of Test 1 but are influenced by the upper torso restraint in Test 2. Certain factors should be considered when employing that assumption. For example: (A) If the test item incorporates some energy absorbing or load limiting design concept necessary to meet the test criteria or other requirement, a less severe loading condition may adversely affect the performance of that design concept as related to the pass-fail criteria. In such a case, it should be shown by rational analysis or additional testing that the design concept would continue to perform as intended even under the lower loads.

(B) If different configuration of the same basic design incorporated load-carrying elements, especially joints or fasteners, which differed in detail design, the performance of each detail design should be demonstrated in a dynamic test.

Experience has shown that small details in the design often cause problems in meeting the test performance criteria.

(C) If structural strength is not the critical condition for achieving the performance criteria of the dynamic test, the true critical condition should be evaluated in a dynamic test. For example, if in one of the design configurations the restraint system attachment points are located so that the lap safety belt was more likely to slip above the ATD pelvis during the impact, then that configuration should also be dynamically tested even though the structural loading might be less. In all cases, the test item should be representative of the final production item in all structural elements and should include seat cushions, armrests and armcaps, functioning position adjustment mechanism, and correctly adjusted seat back breakover (if present), food trays or any other service or accoutrements required by the seat manufacturer or customer, and any other items of mass carried or positioned by the seat structure (e.g., weights simulating luggage carried or restrained by luggage restraint bars, fire Page C - 75 AC 29-2C, Chg 1 2/12/03 extinguishers, survival equipment, etc.). If these items of mass are placed in a position that could limit the function of an energy absorbing design concept in the test item, they should be of representative shape and stiffness as well as weight. That is, seat stroking should perform properly when used in rotorcraft interiors.

(ii) Consideration of test criteria . The test proposal or plan should be planned to achieve “most critical” conditions for the criteria that make up each test.

(A) For multiple occupant seat assemblies, a rational structural analysis should be used to determine the number and seat location for the ATD and the direction for seat yaw in Test 2 to provide the most critical seat structural stress. This will usually result in unequally loaded seat legs. The seat deformation procedure should be selected to increase the load on the highest loaded seat leg and to stress the floor track or fitting in the most severe manner. The seat position in Test 2 depends on the upper torso restraint design. See c.(3)(ii)C below.

(B) If multiple row testing is used to gather data for HIC in passenger seats, the seat pitch distance between seat rows should be selected within the allowable range, so that the head would be most likely to contact hard structure in the forward seat row. The effect of the 10° yaw in Test 2 and of any seat back breakover should be considered. Results from previous tests or rational analysis can be used to estimate the head strike path. Upper torso restraints may prevent head strike; however, leg kick loads into the front seat row require use of two rows. This kick load is a seat structural test not an ATD consideration.

(C) If nonsymmetrical upper torso restraints (such as single diagonal shoulder belts) are used in a system, they should be installed on the test fixture in a position representative of that in the aircraft and that would most likely allow the ATD to move out of the restraint. For example, in a forward facing crew seat equipped with a single diagonal shoulder belt, the seat should be yawed in Test 2 in a direction such that the belt passes over the trailing shoulder. This is a part of the pass/fail criteria evaluation.

(D) If a seat has sitting height adjustment, it should be tested in the th highest position that could be used by a 50 percentile male occupant in the aircraft installation. See b.(3)(ii) of this guidance.

(E) Floor deformation need not be considered in assessing the consequence of any seat deformation as related to the possible impairment of rapid evacuation of the airplane. After the test, the pitch and roll floor beams can be returned to their neutral position and the necessary measurements of the seat deformation made to determine the effect, if any, on rapid evacuation.

(F) In some cases, it may not be possible to measure data for HIC during the test of the seat and torso restraint system. The design of the surrounding interior, such as the instrument panel, may not be known to the designer of the seat and torso Page C - 76 2/12/03 AC 29-2C, Chg 1 restraint system, or the system may be used in several applications with different interior configurations. In such cases, it will be necessary to document the head strike path and the velocity along the path. This will require careful placement of photo instrumentation cameras and location of targets on the ATD representing the ATD head center of mass so that the necessary data can be obtained. These data can be used by the interior designer to ensure that head impact with the interior will not take place or that if possible head impact occurs, it will remain within the limits of the HIC. In the event the head impacts the specific interior, the interior under evaluation should be subjected to an individual special test to measure the head impact or HIC. The test is done using a rigid 6.5-inch diameter spherical head form weighing 15 pounds, (which includes necessary mass to represent the neck and a portion of the torso). The center of the head form is guided along the previously determined head strike path so that the form contacts the interior components at the velocity previously determined during the seat and torso restraint system dynamic test. Accelerometers located at the center of the head form would provide the data necessary for the HIC computation. If the interior component to be impacted by the ATD has significant inertial response to the impact environment, it will be necessary to evaluate those features or systems, such as breakover seatbacks or instrument panels designed to move forward, relative to the seat, in a dynamic test program which includes the full ATD occupant/seat/restraint system. See b.(3)(ii) of this guidance for ATD and panel location for adjustable crew seats.

(iii) Use of ATD . ATD used in the tests discussed in this guidance should be maintained to perform in accordance with the requirements described in their specification. Periodic teardown and inspection of the ATD should be accomplished to identify and correct any worn or damaged components, and appropriate ATD calibration tests (as described in their specification) should be accomplished if major components are replaced. Each ATD should be clothed in form-fitting cotton stretch garments with short sleeves, mid-calf length pants, and shoes (size 11E) weighing about 2.5 pounds.

The head and face of the ATD can be coated with chalk dust if it is desired to mark head contact areas on seats or other structure. The friction in limb joints should be set so that the joints barely restrain the weight of the limb when extended horizontally. The ATD should be placed in the seat in a uniform manner for reproducible test results. For the tests discussed in this guidance, the following procedures are adequate: (A) The ATD should be placed in the center of the seat in as nearly a symmetrical position as possible.

(B) The ATD’s back should be against the seat back without clearance.

This condition can be achieved if the ATD’s legs are lifted as it is lowered into the seat.

Then, the ATD is pushed back into the seat back as it is lowered the last few inches into the seat pan. Once all lifting devices have been removed from the ATD, the ATD should be “rocked” slightly to settle it in the seat.

(C) The ATD knees should be separated about 4 inches.

Page C - 77 AC 29-2C, Chg 1 2/12/03 (D) The ATD hands should be placed on the top of the legs, just behind the knees. If tests on crew seats are conducted in a mockup with aircraft controls, the ATD hands should be lightly tied to the controls. If only the seat and occupant restraint system are tested, the ATD hands should be tied together with a lack cord that provides about 24 inches of separation before the cord becomes tight. This will prevent excessive arm flail during the ATD rebound phase.

(E) To the extent that they influence the injury criteria, all seat th adjustments and controls should be in the design position intended for the 50 percentile male occupant. If seat and occupant restraint systems being tested are to be used in applications where requirements (placards) dictate particular positions for landing and takeoff, those positions should be used in the tests.

(F) The feet should be in the appropriate position for the type of seat tested (flat on the floor for a passenger seat or on control pedals or on a 45° footrest for flightcrew systems). The feet should be placed so that the centerlines of the lower legs are approximately parallel, unless the need for placing the feet on aircraft controls dictates otherwise.

(iv) Installation of instrumentation . Professional practice should be followed when installing instrumentation. Care should be taken when installing the transducers to prevent deformation of the transducer body from causing errors in data.

Lead-wires should be routed to avoid entanglement with the ATD or test item, and sufficient slack should be provided to allow motion of the ATD or test item without breaking the lead wires or disconnecting the transducer. Calibration procedures should consider the effect of long transducer lead-wires. Head accelerometer (transducer) should be installed in the ATD in accordance with the ATD specification and the instructions of the transducer manufacturer. The load cell between the pelvis and the lumbar spinal column should be installed as shown in Figure 29.562-4 of this guidance or in a manner that would provide equivalent data.

(A) An upper torso restraint is required by §29.785(b). The tension load should be measured in a segment of webbing between the ATD’s shoulders and the first contact of the webbing with hard structure (the anchorage point or a webbing guide). Restraint webbing should not be cut to insert a load cell in series with the webbing, since that would change the characteristics of the restraint system.

Commercially available load cells can be placed over the webbing without cutting. They should be placed on free webbing and should not contact hard structure, seat upholstery, or the ATD during the test. They should not be used on double-reeved webbing, multiple-layered webbing, locally-stitched webbing, or folded webbing unless it can be demonstrated that these conditions do not cause errors in the data. These load cells should be calibrated using a length of webbing of the type used in the restraint system. If the placement of the load cell on the webbing causes the restraint system to sag, the weight of the load cell can be supported by light string or tape that will break away during the test.

Page C - 78 2/12/03 AC 29-2C, Chg 1 (B) Loads in restraint systems attaching directly to the test fixture can be measured by three-axis load cells fixed to the test fixture at the appropriate location.

These commercially available load cells measure the forces in three orthogonal directions simultaneously, so that the direction as well as the magnitude of the force can be determined. If desired, similar load cells can be used to measure forces at other boundaries between the test fixture and the test item, such as the forces transmitted by the legs of the seat into the floor track. It is possible to use independent, single axis load cells arranged to provide similar data, but care should be taken to use load cells that can withstand significant cross-axis loading or bending without causing errors in the test data, or use careful (often complex) installation to protect the load cells from cross-axis loading or bending. Since load cells are sensitive to the inertial forces of their own internal mass and to the mass of fixtures located between them and the test article, as well as to forces applied by the test article, it may be necessary to compensate the test data for that inaccuracy if the error is significant. Data for such compensation will usually be obtained from an additional dynamic test replicating the load cell installation but will not include the test item.

(v) Restraint system adjustment . The ATD should be sitting in the normal upright position. Care should be taken not to tighten the restraint system beyond the level reasonably expected in use and do not lock any emergency locking device (inertia reel) prior to the impact. Automatic locking retractors should be allowed to perform the webbing retraction and automatic locking function without assistance. Care should be taken that emergency locking retractors sensitive to acceleration do not lock prior to the impact test because of pre-impact acceleration applied by the test facility that is not present in a landing impact. If “comfort zone” retractors are used, they should be adjusted in accordance with instructions given to the user of the system. If manual adjustment of the restraint system is required, it should be sufficient to remove slack in the webbing, but it should not be adjusted so that it is unduly tight. Since the force required to adjust the length of the webbing can be as high as 11 pounds, a preload of 12-15 pounds is commonly recommended. This load is too small to be accurately measured by transducers selected to measure the high loads encountered in the impact test, so it should be measured manually as the restraint is being adjusted. Special gauges are commercially available to assist in this measurement. The preload should be checked and adjusted, if necessary, just prior to the floor deformation phase of the test.

(vi) Repetition of tests . It may be necessary to repeat the tests discussed in this AC if accurate data are not collected in critical data channels or if some other error occurs (e.g., cameras fail to operate, impact pulse inadequate, etc.). Preparation for a repeated test should follow the same steps as for the initial test. The seat should be removed from the fixture, and its attachment fittings or floor track examined and replaced, if necessary, to correct any damage. The ATD should be carefully examined and repaired or adjusted, if necessary. It is usually preferable to use a new seat and restraint system for all repeated tests to preclude system failures due to undetected damage. A new seat and restraint system should be used if there is any detectable variation from the intended design configuration.

Page C - 79 AC 29-2C, Chg 1 2/12/03 d. Data Analysis And Compliance With The Criteria (1) General . All data obtained in the dynamic tests should be reviewed for errors. Baseline drift, “ringing,” and other common electronic instrumentation problems should be detected and corrected before the tests. Loss of data during the test is readily observed in a plot of the data vs. time and is typically indicated by sharp discontinuities in the data, often exceeding the amplitude limits of the data collection system. If these occur early in the test in essential data channels, the data should be rejected and the test repeated. If they occur late in the test, after the maximum data in each channel has been recorded, the validity of the data should be carefully evaluated, but the maximum values of the data may still be acceptable for the tests described in this guidance. The HIC does not represent a maximum data value, but represents an integration of data over a varying time base. The head acceleration measurements used for that computation should not be accepted if errors or loss of data are apparent in the data at any time from the beginning of the test until the ATD and all test articles are at rest after the test.

(2) Impact pulse shape . Data for evaluating the impact pulse shape are obtained from an accelerometer that measures the acceleration in the direction parallel to the line of inertial response shown in Figure 29.562-1 of this guidance. The impact pulse intended for the tests discussed in this guidance has a symmetrical (isosceles) triangular shape. Since this ideal pulse is considered a minimum test condition, it is possible to evaluate the actual test pulse by comparing it with the ideal triangular pulse.

The ideal pulse can be drawn to scale on the data plot of the test sled or carriage acceleration vs. time. The test pulse is acceptable if the plotted data are equal to or greater than the ideal impact pulse. This method can lead to a practical necessity of exceeding the ideal pulse by a significant degree, unless the test facility has precise control in generating the test pulse. A graphic technique may be used to evaluate test impact pulse shapes that are not precise isosceles triangles. A graphic technique is contained in paragraph f. (1) of this guidance.

(3) Head Injury Criterion (HIC) . Data for determining the HIC need to be collected during the tests discussed in this guidance only if the ATD’s head is exposed to secondary impact. The HIC is a method for defining an acceptable limit; i.e., the maximum values of the HIC should not exceed 1,000 for head impact against broad interior surfaces in a crash. The HIC is reported as the maximum value, and the time interval during which the maximum value occurs is also given. Most facilities will make this computation if requested. The HIC is calculated by computer-based data analysis systems because manual attempts to use this method with real data are likely to be tedious. The HIC is calculated according to the following equation: Page C - 80 2/12/03 AC 29-2C, Chg 1 2.5 MAX Where: t and t are any two points in the time range during the head impact. The 1 2 range should not exceed 0.050 seconds, and a(t) is the resultant head acceleration at the center of gravity (expressed in g’s) during the head form impact.

(i) Data collection . The HIC is commonly based on data obtained from three mutually perpendicular accelerometers installed in the head of the ATD in accordance with the ATD specification. Data from these accelerometers are obtained using a data system conforming to Channel Class 1,000 as described in SAE Recommended Practice J211. For the tests discussed in this guidance (both ATD and head form), only the data taken during secondary head impact with the aircraft interior need be considered. Head impact is often indicated in the data by a rapid change in the magnitude of the acceleration. Alternately, a film of the test may show head impact which can be correlated with the acceleration data by using the time base common to both electronic and photographic instrumentation, or simple contact switches on the impacted surface can be used to define the initial contact time.

(ii) HIC methodology . The following discussion outlines the basic method for computing the HIC. The magnitude of the resultant acceleration vector obtained from the three accelerometers is plotted against time. Then, beginning at the time of initial head contact (t ), the average value of the resultant acceleration is found for each increasing increment of time (t – t ), then integrating the curve between the range of t 2 1 1 and t and then dividing the integral value by the time (t – t ). This calculation should 2 2 1 use all data points provided by the minimum 8,000 samples per second digital sampling rate for the integration. However, the maximizing time intervals need be no more precise than 0.001 seconds. The average values are then raised to the 2.5 power and multiplied by the corresponding increment of time (t – t ). This procedure is then 2 1 repeated, increasing t by 0.001 seconds for each repetition. The maximum value of the set of computations obtained from this procedure is the HIC. The procedure may be simplified by noting that the maximum value will only occur in intervals where the resultant magnitude of acceleration at t is equal to the resultant magnitude of acceleration at t and when the average resultant acceleration in that interval is equal to 5/3 times the acceleration at t or t .

1 2 (iii) Limitations . HIC does not consider injuries that can occur from contact with surfaces having small contact areas or sharp edges, especially if those surfaces are relatively rigid. These injuries can occur at low impact velocities, and are often described as “cosmetic” injuries; however, they can involve irreversible nerve damage and permanent disfigurement. While there is no generally accepted test procedure to provide quantitative assessment of these injuries, a judgmental evaluation of soft tissue injuries can be made by assessing tears or cuts in a synthetic skin placed Page C - 81 AC 29-2C, Chg 1 2/12/03 over the ATD’s head or a head form during the test. Synthetic skins are discussed in the Society of Automotive Engineers Information Report SAE J202, Synthetic Skins for Automotive Testing.

(4) Impact velocity . Impact velocity can be obtained by measurement of a time interval and a corresponding sled displacement occurring just before or after (for acceleration facilities) the test impact, and then dividing the displacement by the time interval. When making such a computation, the possible errors of the time and displacement measurements should be used to calculate a possible velocity measurement error, and the test impact velocity should exceed the velocity shown in Figure 29.562-1 by at least the velocity measurement error. If the sled is changing acceleration during the immediate pre-impact interval, or if the facility produces significant rebound of the sled, the effective impact velocity can be determined by integrating the plot of sled acceleration vs. time. If this method is used, the sled acceleration should be measured in accordance with Channel Class 180 requirements.

(5) Upper torso restraint system load . The maximum load in the upper torso restraint system webbing can be obtained directly from a plot or listing of webbing load transducer output. If a three-axis load transducer, fixed to the test fixture, is used to obtain these data, the data from each axis should be combined to provide the resultant vector magnitude. If necessary, corrections should be made for the internal mass of the transducer and the fixture weight it supports. This correction will usually be necessary only when the inertial mass or fixture weight is high or when the correction becomes critical to demonstrate that the measurements fall below the specified limits.

(6) Compressive load between the pelvis and lumbar column . The maximum compressive load between the pelvis and the lumbar column of the dummy can be obtained directly from a plot or listing of the output of the load transducer at that location. Since most load cells will indicate tension as well as compression, care should be taken that the polarity of the data has been correctly identified.

(7) Retention of upper torso restraint straps . Retention of the upper torso restraint webbing straps on the ATD’s shoulders can be verified by observation of photometry or documentary camera coverage. The webbing should remain on the sloping portion of the ATD’s shoulder until the ATD rebounds after the test impact and the upper torso restraint straps are no longer carrying any load. The webbing straps should not bear on the neck or side of the head and should not slip to the upper rounded portion of the upper arm during that time period.

(8) Retention of lap safety belt . Retention of the lap safety belt on the occupant’s (ATD) pelvis can be verified by observation of photometry or documentary camera coverage. The lap safety belt should remain on the ATD’s pelvis, bearing on or below each prominence representing the anterior superior iliac spines, until the ATD rebounds after the test impact and the lap safety belt becomes slack. If the lap safety belt does not become slack throughout the test, the belt should maintain the proper position throughout the test. Movement of the lap safety belt above the prominence is Page C - 82 2/12/03 AC 29-2C, Chg 1 usually indicated by an abrupt displacement of the belt into the ATD’s soft abdominal insert which can be seen by careful observation of photo data from a camera located to provide a close view of the belt as it passes over the ATD’s pelvis. This movement of the belt is sometimes indicated in measurements of lap safety belt load (if such measurements are made) by a transient decrease or plateau in the belt force, as the belt slips over the prominence, followed by a gradual increase in belt force as the abdominal insert is loaded by the belt. Retention of the lap safety belt can also be verified by “submarining indicators” located on the ATD’s pelvis. These transducers are essentially a series of small, uncalibrated load cells placed in or above the rim of the ATD’s pelvis without changing its essential geometry. They indicate the position of the lap safety belt by producing an electrical signal when they are under load from the belt.

(9) Femur load . Measuring femur loads is not required by the rotorcraft standards. If a seat is installed in an aircraft in a manner that will expose the system to loads from an occupant seated behind the seat system as well as the occupant seated in the seat system, the tests discussed in this guidance should be conducted in a manner to demonstrate that the system will perform properly under the combined loading. For example, Test 2 should be conducted with at least two rows of seats in place, as the seats in the first row carry the loads from the occupants in the first row, as well as the leg kick loads from the second row (also noted in c.(3)(ii)(A) of this guidance).

(10) Seat attachment . Documentation that the seat and restraint system has remained attached at all points of attachment should be provided by still photographs that show the intact system components in the load path between the attachment points and the occupant.

(11) Seat deformation . Occupant seats evaluated in the tests discussed in this guidance can deform permanently, either due to the action of discrete (impact) energy absorber systems included in the design or due to residual plastic deformation of their structural components. If this deformation is excessive, it could impede emergency evacuation. Each seat design may differ in this regard and should be evaluated according to its unique deformation characteristics. Permanent seat deformations are measured on the critically loaded seat subsequent to conduct of the tests required in §29.562. The seat deformation is measured subsequent to completion of the dynamic tests and, where applicable, release of the applied pre-test floor deformation.

(i) Seats . The following post-test deformations and limitations regarding emergency egress and access to exits may be used for showing compliance with § 29.785(j): (A) Forward or Rearward Directions . The forward or rearward deformations should not exceed a maximum of 4.0 inches (100 mm). In addition, the clearance between undeformed seat rows, measured as shown in Figure AC 29.562-5 (Dimension A), should be a minimum of 9.0 inches, except where seat rows lead to Type III or IV exits, where it should be a minimum of 11.0 inches. For seats with Page C - 83 AC 29-2C, Chg 1 2/12/03 deformations exceeding 4.0 inches, the undeformed clearances between seats should be increased accordingly. In addition, at seat rows leading to Type III or IV exits, a minimum of 20 inches clearance, measured above the arm rests, must be maintained between adjacent seat rows. This measurement may be made with the seat backs returned, using no more than original seat back breakover forces, to their pretest upright or structurally deformed position. At other seat rows, the most forward surface of the seat back should not deform to a distance greater than one half of the original distance to the forwardmost hard structure on the seat (see Figure 29.562-6).

(B) Downward Direction . There is no limitation on downward deformation, provided it can be demonstrated that the feet or legs of occupants seated aft would not be entrapped. Additionally, the seat bottom rotational deformation from the horizontal, measured at the centerline of each seat pan, should not exceed 20° forward (pitch down) or 35° aft (pitch up). This measurement should be made between the fore and aft extremities of the seat pan structure, considering the final position of the seat pan structure. In no case should rotation of the seat pan cause entrapment of the occupant.

(C) Sideward Direction .

( 1 ) The deformed seat should not encroach more than 1.5 inches (40 mm) into the required space for longitudinal aisle at heights up to 25 inches (635 mm) above the floor. Determine which parts of the seat are at what heights prior to testing.

( 2 ) The deformed seat should not encroach more than 2.0 inches (50 mm) into the longitudinal aisle space at heights 25 inches (635 mm) or more above the floor.

(D) Additional Considerations. In addition, none of the above deformations permit the seat to: ( 1 ) Affect the operation of any emergency exit or encroach into an emergency exit opening for a distance from the exit not less than the width of the narrowest passenger seat installed except as stated in § 29.813(c)(2).

Page C - 84 2/12/03 AC 29-2C, Chg 1 Measurement to be taken over full width of seat bottom cushion FIGURE AC 29.562 - 5 Pre - test condition Post - test condition Dimension “C” must be at least 50% of Dimension “B” FIGURE AC 29.562 - 6 Page C - 85 AC 29-2C, Chg 1 2/12/03 ( 2 ) Encroach into any required passageway to large exits, § 29.813(a) and (b).

( 3 ) Encroach more than 1.5 inches into any cross aisle or evacuation (flight attendant) assist space for certain exits.

(ii) Stowable Seats . Stowable seats, if used, should stow post-test and remain stowed without projecting into any required passageways. In addition, they should not project more than 1.5 inches into any flight attendant assist space or cross-aisle.

(A) Seats that are Stowed Manually . A post-test stowage force no greater than 10 pounds (22kg) above the original stowage force may be used to stow the seat.

(B) Seats that Stow Automatically . For a seat that may interfere with the opening of any exit, it shall automatically retract to a position that does not interfere with the exit opening as prescribed in § 29.807. For determining encroachment into passageways, cross-aisles, and assist spaces, a posttest stowage force no greater than 10 pounds (22kg), applied at a single point, may be used to assist automatic retraction.

e. Test Documentation .

(1) General . The tests discussed in this guidance should be documented in reports describing the test procedures and results. The test proposal, a description of the required tests, approved by the FAA should be referenced in the test report and contain the following: (i) Facility data .

(A) The name and address of the test facility performing the tests.

(B) The name and telephone number of the individual at the test facility responsible for conducting the tests.

(C) A brief description and/or photograph of each test fixture.

(D) The date of the last instrumentation system calibration and the name and telephone number of the person responsible for instrumentation system calibration.

(E) A statement confirming that the data collection was done in accordance with the recommendations in this guidance or a detailed description of the actual calibration procedure used and technical analysis showing equivalence to the recommendations of this AC (Paragraph c(2)(iii)(A)).

Page C - 86 2/12/03 AC 29-2C, Chg 1 (F) Manufacturer, governing specification, serial number, and test weight of ATD used in the tests, and a description of any modifications or repairs performed on the ATD which could cause them to deviate from the specification.

(G) A description of the photographic-instrumentation system used in the tests (Paragraph c(2)(iii)(B)).

(ii) Seat/Restraint system data .

(A) Manufacturers name and identifying model numbers of the seat/ restraint system used in the tests, with a brief description of the system, including identification and a functional description of all major components and photographs or drawings as applicable.

(B) For unsymmetrical systems, an analysis supporting the selection of most critical conditions used in the tests.

(2) Test Proposal or Plan and Description . The description of the test should be documented in enough detail so that the tests could be reproduced by following the guidance given in the report. The procedures outlined in this guidance can be referenced in the report but should be supplemented, as necessary, to describe the unique conditions of the individual seat design.

(i) Pertinent dimensions and other details of the installation not included in the drawings of the test items should be provided. This can include footrests, restraint system webbing guides and restraint anchorages, “interior surface” simulations, bulkhead or sidewall attachments for seats or restraints, etc.

(ii) The floor deformation procedure, guided by goals of most critical loading for the test articles, should be documented.

(iii) Placement and characteristics of electronic and photographic instrumentation chosen for the test, beyond that information provided by the facility, should be documented. This can include special targets, grids or marking used for interpretation of photo documentation, and transducers and data channel characteristics for lap belt loads, floor reaction forces, or other measurements beyond those discussed in this guidance.

(iv) Any unusual or unique activity or event pertinent to conducting the test should be documented. This could include use of special “breakaway” restraints or support for the ATD’s, test items or transducers, operational conditions or activities such as delayed or aborted test procedures, and failures of test fixtures, instrumentation system components or ATD.

(3) Test results report . The documentation should include copies of all test results, analysis, and conclusions. As a minimum, the following should be documented: Page C - 87 AC 29-2C, Chg 1 2/12/03 (i) Impact pulse shape (Paragraph d(2)).

(ii) HIC results for all ATD exposed to secondary head impact with interior components of the rotorcraft (Paragraph d(3)), or head strike paths and velocities if secondary head impact is likely for future use in unique interiors (Paragraph c.(3)(iii)).

(iii) Impact velocity (Paragraph d(4)).

(iv) Upper torso restraint system load if applicable (Paragraph d(5)).

(v) Compressive load between the pelvis and the lumbar column (Paragraph d(6)6).

(vi) Retention of upper torso restraint straps if applicable (Paragraph d(7)).

(vii) Retention of lap safety belt (Paragraph d(8)).

(viii) Femur thigh loads, optional measurement.

(ix) Seat attachment (Paragraph d(10)).

(x) Seat deformation (Paragraph d(11)).

(xi) Seat attachment reaction time histories (Paragraph f).

(4) Dynamic Impact Test – Pass/Fail Criteria : The dynamic impact tests should demonstrate that: (i) The seat structure remains intact that is attached to the tracks or fittings, etc.

(ii) The occupant retention system is capable of carrying the dynamic loads.

(iii) The seat permanent deformations are within defined limits and will not significantly impede an occupant from releasing the torso restraints, standing and exiting the seat.

(iv) If the ATD’s head is exposed to impact during the test, a HIC of 1,000 is not exceeded. Data may be obtained for use with other unique installations.

(v) Where upper torso restraint straps are used, tension loads in individual straps do not exceed 7.78 kN (1,750 lbs.). If dual straps are used for restraining the upper torso, the total strap tension load does not exceed 8.90 kN (2,000 lbs.).

Page C - 88 2/12/03 AC 29-2C, Chg 1 (vi) The maximum compressive load measured between the pelvis and the lumbar column of the (ATD) does not exceed 6.67 kN (1,500 lbs.).

(vii) Each upper torso restraint strap remains on the ATD shoulder during impact.

(viii) The pelvic restraint remains on the ATD pelvis during impact.

f. Procedures for Evaluating Impact Pulse Shapes .

(1) Acceptable Evaluation Method . Data for evaluating the impact pulse shape are obtained from an accelerometer which measures the acceleration on the test fixture or sled at the seat location or equivalent location in the direction parallel to the line of inertial response shown in Figure 29.562-1 of this guidance. The impact pulses intended for the tests discussed in this guidance have an isosceles triangle shape.

These ideal pulses are considered minimum test conditions. Since the actual acquired test pulses will normally differ from the ideal, it may be necessary to evaluate the acquired test pulses to insure the minimum requirements are satisfied.

(2) An acceptable method to evaluate the pulse shape should use the following steps: (i) Extend the calibration baseline (zero G) (ii) Locate the maximum peak deceleration (G ) indicated on the plot.

p (iii) Construct reference lines parallel to the baseline at levels of 0.1 G , p 0.9 G , and 1.0 G .

p p (iv) Construct an onset line through the intersection points of the 0.1 G p and 0.9 G reference lines with the increasing (onset) portion of the data plot. The data p plot should not return to zero G between the two points selected.

(v) Locate the intersection points of the onset line with the baseline and with the 1.0 G reference line. The interval between these two points, measured along p the time axis of the data plot, is considered the rise time (t ) of the test impact pulse.

r (vi) The rise time of the test impact pulse should not exceed the value of (t ) given in Figure AC 29.562-1 for each test.

r (vii) The area under the data plot curve within the rise time of the test impact pulse, V , should represent at least one half of the impact velocity given in ra Figure AC 29.562-1 for each test. If the value of peak acceleration measured in the test exceeds the level given in Figure AC 29.562-1 by no more than 10 percent, the pelvis to lumbar spinal column force and the upper torso restraint force measured in the test may Page C - 89 AC 29-2C, Chg 1 2/12/03 be adjusted by multiplying the measured values by the ratio of the peak acceleration given in Figure AC 29.562-1, divided by the measured peak acceleration, if necessary.

(viii) The magnitude of G should equal or exceed the minimum G given in p Figure AC 29.562-1 for each test.

(ix) The area under the data plot curve from the intersection point of the onset line and the zero G baseline and a time not more than twice the appropriate rise time specified in Figure AC 29.562-1, plus 30 percent of the rise time later, should represent at least the impact during the test.

Page C - 90 9/30/99 AC 29-2C Sled Impact Pulse u, I

-,-

1Gp I I 0 50 100 150 200 250 300 350 Time (Milliseconds) FIGURE AC 29.562-7. Impulse Shape Sled Impact Pulse 0.0 0 50 100 150 200 250 300 350 Time (Milliseconds) FIGURE AC 29.562-8. Change in Velocity (V, 0) Page C - 91 AC 29-2C, Chg 4 5/1/2014 AC 29.563. § 29.563 (Amendment 29-12) STRUCTURAL DITCHING PROVISIONS.

a. Explanation . Amendment 29-12 included certification requirements for ditching approvals. The rotorcraft must be able to sustain an emergency landing in water as prescribed by § 29.801(e).

b. Procedures . Refer to paragraph AC 29.801, § 29.801, for procedures.

AC 29.563A. § 29.563 (Amendment 29-30) STRUCTURAL DITCHING PROVISIONS.

a. Explanation . Amendment 29-30 added specific structural conditions to be considered to support the overall ditching requirements of § 29.801. These conditions are to be applied to rotorcraft for which over-water operations and associated ditching approvals are requested.

(1) The forward speed landing conditions are specified as: (i) The rotorcraft should contact the most critical wave for reasonable, probable water conditions in the likely pitch, roll, and yaw attitudes.

(ii) The forward velocity relative to wave surface should be in a range of 0 to 30 knots with a vertical descent rate of not less than 5 FPS relative to the mean water surface.

NOTE: A forward velocity of less than 30 knots may be used for multiengine rotorcraft if it can be demonstrated that the forward velocity selected would not be exceeded in a normal one-engine-out touchdown.

(iii) Rotor lift of not more than two-thirds of the design maximum weight may be used to act through the CG throughout the landing impact.

(2) For floats fixed or deployed before water contact, the auxiliary or emergency float conditions are specified in § 29.563(b)(1). Loads for a fully immersed float should be applied (unless it is shown that full immersion is unlikely). If full immersion is unlikely, loads resulting from restoring moments are specified for sidewind and unsymmetrical rotorcraft landing.

(3) Floats deployed after water contact are normally considered fully immersed during and after full inflation. An exception would be when the inflation interval is long enough that full immersion of the inflated floats does not occur; e.g., deceleration of the rotorcraft during water impact and natural buoyancy of the hull prevent full immersion loads on the fully inflated floats.

b. Procedures .

Page C - 92 9/30/2008 AC 29-2C, Chg 3 (1) The rotorcraft support structure, structure-float attachments, and floats should be substantiated for rational limit and ultimate ditching loads.

(2) The most severe wave heights for which approval is desired are to be considered. A minimum of Sea State 4 condition wave heights should be considered (reference paragraph AC 29.801 (§ 29.801) for a description of Sea State 4 conditions).

(3) The landing structural design consideration should be based on water impact with a rotor lift of not more than two-thirds of the maximum design weight acting through the center of gravity under the following conditions: (i) Forward velocities of 0 to 30 knots (or a reduced maximum forward velocity if it can be demonstrated that a lower maximum velocity would not be exceeded in a normal one-engine-out landing).

(ii) The rotorcraft pitch attitude that would reasonably be expected to occur in service. Autorotation flight tests or one-engine-inoperative flight tests, as applicable, should be used to confirm the attitude selected. This information should be included in the Type Inspection Report.

(iii) Likely roll and yaw attitudes.

(iv) Vertical descent velocity of 5 FPS or greater.

(4) Landing load factors and water load distribution may be determined by water drop tests or analysis based on tests.

(5) Auxiliary or emergency float loads should be determined by full immersion or the use of restoring moments required to react upsetting moments caused by sidewind, asymmetrical rotorcraft landing, water wave action, rotorcraft inertia, and probable structure damage and punctures considered under § 29.801. Auxiliary or emergency float loads may be determined by tests or analysis based on tests.

(6) Floats deployed after initial water contact are required to be substantiated by tests or analysis for the specified immersion loads (same as for (5) above and for the specified combined vertical and drag loads).

[The remainder of this page intentionally left blank.]

Page C - 93 AC 29-2C, Chg 3 9/30/2008 SUBPART C - STRENGTH REQUIREMENTS FATIGUE EVALUATION AC 29.571. § 29.571 FATIGUE EVALUATION OF FLIGHT STRUCTURE .

a. Explanation . An evaluation is required to assure structural reliability of the rotorcraft in flight. This evaluation may take the form of either tests or analysis. During the certification process, fatigue testing is more effective than analysis alone in identifying and preventing cracking that may occur during service. Analysis used for substantiation should be validated by tests. AC 27 MG 11 contains background information and acceptable means of compliance with the requirements pertaining to the safe life methodology. A safe life may be assigned or the structure may be determined to be fail safe or a combination of these may be used. AC 29 MG 11 contains background information and acceptable means of compliance with the requirements pertaining to fatigue and flaw tolerance.

b. Procedures .

(1) The fatigue evaluation requires consideration of the following factors: (i) Identification of the structure/components to be considered.

(ii) The stress during operating conditions.

(iii) The operating spectrum or frequency of occurrence.

(iv) Fatigue strength, and/or fatigue crack propagation characteristics, residual strength of the cracked structure.

(2) Since the design limits, e.g., rotor RPM (maximum and minimum), airspeed, and blade angles (thrust, weight, etc.) affect the fatigue life of the rotor system, it is necessary that flight conditions be conducted at limits that are appropriate for the particular rotorcraft and at the correct combination of these limits. It will be the responsibility of flight test personnel to determine that the flight strain program includes conditions of flight at the various combinations of rotor RPM, airspeed, thrust, etc., that will be representative of the limits used in service. The flight test personnel should assure that the severity of the maneuvers to be investigated is such that actual service use will not be more severe. Flight test verification may be achieved through: (i) Flying a representative set of maneuvers with the applicant’s pilot in the test aircraft at noncritical combinations of weight, CG, and speed. (An FAA/AUTHORITY letter for specific test authorization would ordinarily be required.)

(ii) Flying a representative set of maneuvers with the applicant’s pilot in a similar (certified) model to assess and agree upon the required maneuvers, control C – 94 9/30/2008 AC 29-2C, Chg 3 deflections, and aircraft rates. The required maneuvers or conditions will be specified in the flight strain program plan.

(iii) Flying a chase aircraft which has a flight envelope appropriate to allow visual confirmation of the proposed and programmed flight maneuvers.

(iv) Observation of telemetered flight data to assure desired control deflections, rates, and aircraft attitudes.

(v) Some combinations of items b(2)(i) through b(2)(iv) above.

(3) Assessing the operation spectrum and the flight loads or strain measurement program will involve airframe, propulsion, and flight test personnel.

(4) Variation in the operating or loading spectrum among models, and variations in the spectrum for a particular model rotorcraft, should be evaluated.

Figure AC 27 MG 11-7 contains typical flight load measurement program conditions to be investigated. An example of a twin turbine spectrum is presented in Figure AC 27 MG 11-9. The tables should be used only as a guide and should be modified as necessary for each particular rotorcraft design.

(5) The difference in loading spectrum for different models that may be anticipated is illustrated by comparing the percentage of time assigned to level flight conditions, specifically 0.8 V to 1.0 V for three different rotorcraft designs where V is H H H the maximum airspeed at maximum continuous power in level flight. The first column applies to a single-piston-engine powered small rotorcraft used in utility operations. The second was obtained from data for a single-turbine-engine powered seven-place small business and utility rotorcraft. The third was obtained from data for a twin-engine-powered 13 passenger transport rotorcraft. It should be noted that the level flight percentage of occurrences shown in the table below for the turbine utility business and turbine transport rotorcraft are examples of a particular design. The high percentage of time shown in this flight regime could be unconservative for some designs, especially if the stresses under these design conditions produce an infinite fatigue life for the particular component. The fatigue spectrum percentage of occurrences may be modified according to the intended operation usage of the rotorcraft. However, a conservative application should be considered. This variation illustrates the "tailoring" of the loading spectrum for the type of rotorcraft and the anticipated usage.

[The remainder of this page intentionally left blank.]

C – 95 AC 29-2C, Chg 1 2/12/2003 FIGURE AC 29.571-1 Comparison Percent of Time in Level Flight Turbine Piston Utility Twin Turbine Utility Business Transport 0.8 V 25% 0.8 V 16% 0.8 V 15% NE H H 1.0 V 15% 0.9 V 21% 0.9 V 20% H H H 1.0 V 3% 1.0 V 24% 1.0 V 38% NE H H Total 43% 61% 73% This variation illustrates the “tailoring” of the loading spectrum for the type of rotorcraft and the anticipated usage.

(6) External cargo operations are a unique and demanding operation. A “logging” operator may use 50 maximum power applications per flight hour to move logs from a cutting site to a hauling site. Power is used to accelerate, decelerate, or hover prior to load release. Lifting loads over an obstruction or natural barrier is another example of very frequent high power applications for takeoff and for hovering over the release area. Similar types of operations require flight loads data to assess the effects on fatigue critical components.

(7) Frequently the applicant may request approval of a gross weight for an external cargo configuration that exceeds the standard configuration gross weight. The external cargo V is typically significantly lower than the standard configuration V NE NE possibly due to adverse effects on flight loads at the increased weight.

(8) The impact of the external cargo operation on standard configuration limits should be assessed to determine whether or not the component service lives will be affected. The assessment may be done by calculating an “external cargo configuration” service life for each critical component. The lowest service life obtained from standard configuration flight loads data and loading spectrum, or from external cargo configuration flight loads data and loading spectrum is generally the approved service life. This procedure avoids prorating the operating time between the two types of operations. This procedure is necessary since the regulatory maintenance and operating rules do not require recording time in service for the different types of operations.

(9) The applicant should plan to conduct a flight loads survey program for both a standard configuration and an external cargo configuration, if appropriate. This procedure will avoid delays associated with reinstallation and calibration of equipment.

Page C - 96 5/1/2014 AC 29-2C, Chg 4 AC 29.571A. §29.571 (Amendment 29-28) FATIGUE EVALUATION OF STRUCTURE.

a. Explanation . Amendment 29-28 adds a requirement to substantiate tolerance to flaws during the fatigue evaluation of structure. A flaw tolerant safe-life evaluation or a fail-safe (residual strength after flaw growth) evaluation is required by § 29.571(b) unless “the applicant establishes that these fatigue flaw tolerant methods for a particular structure cannot be achieved within the limitations of geometry, inspectability, and good design practices.” b. Procedures .

(1) AC 29 MG 11 provides acceptable general procedures for complying with Amendment 29-28.

(2) Specific rotorcraft drive system gear fatigue evaluation procedures, which supplement AC 29 MG 11, follow: (i) Fatigue test evidence is necessary for the fatigue evaluation of gears.

The test evidence should be provided by rotating tests of complete gearbox specimens operating under power. The tests provide the basis for analysis leading to the establishment of safe-life.

(ii) The tests are conducted specifically for the purpose of gear tooth evaluation, and components subjected to the tests do not have to be considered serviceable on completion of the test. Excessive wear on bearings and shafts and marking (including spalling) of bearings and gear teeth are acceptable provided no fatigue damage is evident on the gear teeth. However, fatigue damage other than tooth fatigue should be considered for test validity and the integrity of the affected part confirmed as necessary.

(iii) The test conditions (torque versus number of cycles) should permit the setting of mean strength curve(s) to be associated with each primary gear in the drive train. The minimum test condition should encompass those power levels for which repeated application in service is expected under normal conditions. The S-n curve(s), for the material and type of gear, should be reduced by a factor of safety to take into account material and manufacturing variability. The factored curve will then be used in conjunction with the flight power spectrum to determine a life (limited or unlimited) for the gears in the primary drive system.

(iv) Special procedures, which do not affect fatigue evaluation of the gear teeth, may be allowed to facilitate completion of the test provided they have been justified and they do not affect life determination. These include periodic interruption for inspection, replacement of non-critical parts and the use of special lubricants, special cooling systems, and methods to prevent unrepresentative deflections at the test torque levels.

Page C - 97 AC 29-2C, Chg 4 12/2/2011 (v) From evidence in relation to the strength of steel gears of conventional design, it is accepted that adequate fatigue strength can be demonstrated by the use of the above safety factor of 1.4 for a single test, 1.35 for two tests, 1.32 for three tests, and 1.3 for four or more tests. Where several tests are to be conducted, specimens should be selected from different manufacturing batches if practicable.

(vi) Demonstration of infinite life for gear teeth will normally require tests of a minimum of 10 cycles duration at factored power levels. Use of shorter duration tests should be justified.

AC 29.571B. § 29.571 (Amendment 29-55) FATIGUE TOLERANCE EVALUATION OF METALLIC STRUCTURE.

a. Purpose . This advisory material provides an acceptable means of compliance with the provisions of § 29.571 Amendment 29-55 of the FAA regulations dealing with the fatigue tolerance evaluation of transport category rotorcraft metallic structure. This guidance applies to conventional metallic materials. (Corresponding guidance for composite structure can be found in AC 29–2C MG 8, supplemented by AC 20-107B.

Note: once § 29.573 is effective, AC 29.573 will be the current guidance for composite structures.) The fatigue evaluation procedures outlined in this advisory material are for guidance purposes only and are neither mandatory nor regulatory in nature. Although a uniform approach to fatigue tolerance evaluation is desirable, it is recognized that in such a complex area, new design features and methods of fabrication, new approaches to fatigue tolerance evaluation, and new configurations may require variations and deviations from the procedures described herein. It should be noted that § 29.571 requires that the methodology used by the applicant be approved by the FAA to assure compliance with the regulatory requirements.

b. Special Considerations . The unique performance capabilities of rotorcraft and their typical operational environment make fatigue tolerance evaluations both complex and critically important. Due to the many rotating elements inherent in their design, rotorcraft structures are potentially subject to damaging cyclic stresses in practically every regime of flight. The complexity of the fatigue loading is compounded by the fact that rotorcraft are highly maneuverable and are utilized for many widely varying roles.

Corrosion and other environmental damages are not uncommon in rotorcraft operations; neither are inadvertent damages from maintenance that is typically frequent and intensive. For these reasons, special attention should be focused on the fatigue tolerance evaluation of rotorcraft structure.

c. Background .

(1) Fatigue of rotorcraft dynamic components was first addressed in the 1950’s by means of a Safe-Life methodology. The application of this methodology, as Page C - 98 12/2/2011 AC 29-2C, Chg 4 described in AC 27-1B MG 11, has proven to be successful in providing an adequate level of reliability for transport category rotorcraft. However, it was recognized in the 1980’s that higher levels of reliability might be realized by taking into account the fatigue strength-reducing effects of damage that experience has shown can occur in manufacture or in operational service. The introduction of composites led the manufacturers and regulatory authorities to develop a robust Safe-Life methodology by taking into account the specific static and fatigue strength-reducing effects of aging, temperature, moisture absorption, impact damage, and recognition of an accepted industry standard. Furthermore, where clearly visible damages resulted from impact or other sources, inspection programs were developed to maintain safety. In parallel, crack growth methodology has been successfully used for solving short-term airworthiness problems in metallic structures of rotorcraft, and as the certification basis for civil and military transport aircraft applications. These advances in design, analytical methods, and industry practices made it feasible to address certain types of damage, which could result in fatigue failure. Consistent with this, the regulatory requirements of § 29.571 were substantially revised by Amendment 28. While many years have passed since its introduction, Amendment 28 has had little exposure to use for certification of completely new rotorcraft designs. However, the general understanding of rotorcraft fatigue tolerance evaluation has developed considerably in the interim and an additional amendment was determined to be appropriate. The latest Amendment 29-55 of part 29 and the associated revisions to advisory material were introduced to improve the currency and understanding of the rule and clarify the differing approaches and methods available for accomplishing fatigue tolerance evaluation of rotorcraft metallic structure.

(2) This guidance provides material with respect to the fatigue tolerance requirements for metallic structure and is supplemented by AC 27-1B MG 11 for evaluations using the Safe-Life methodology and other general fatigue considerations.

d. Introduction .

(1) Definitions. The following definitions are applicable when used within the context of this guidance material.

(i) As-manufactured structure is a structure that passes the applicable quality control process and has been found to conform to an approved design within the allowable tolerances.

(ii) Barely Detectable Flaw (BDF) is the worst-case flaw that is expected to remain on the structure for its operational life.

(iii) Catastrophic failure is an event that could prevent continued safe flight and landing.

Page C - 99 AC 29-2C, Chg 4 12/2/2011 (iv) Clearly Detectable Flaw (CDF) is the worst-case detectable flaw that would not be expected to remain in place for a significant period of time without corrective action.

(v) Damage is a detrimental change to the condition of the structure or assembly. In the context of this guidance material it is used as a generic term to describe all types of flaws including those caused by environmental effects and accidental damage arising in manufacture, maintenance or operation.

(vi) Damage Tolerance is the attribute of the structure that permits it to retain its required residual strength without detrimental structural deformation for a period of un-repaired use after the structure has sustained a given level of fatigue, corrosion, accidental, or discrete source damage.

(vii) Discrete flaw is a flaw that is not inherent in the design and is caused by an external action, such as corrosion, scratches, gouges, nicks, fretting, wear, impact, and potentially cracks initiated by fatigue.

(viii) Fatigue is a degradation process of a structure subject to repeated loads that may involve four phases (e.g., nucleation of many micro-cracks, coalescence of some micro-cracks to one major macro-crack, stable crack growth, unstable crack growth) and immediate failure. The boundaries between these phases are, in practice, not always easily defined. Crack initiation methods (e.g., using the S-N curve and the Miner’s Rule) are generally used to address the first two phases. Linear Fracture Mechanics methods (e.g., using da/dn - ΔK and fracture toughness data) are generally used for the latter two phases.

(ix) Fatigue Loads are repeated loads, which induce a repeated variation of stress versus time in a structure.

(x) Fatigue Tolerance is the ability of a structure, either in an as-manufactured or damaged condition, to tolerate specified operational loading for a given period of use without initiating cracks, and assuming they initiate, tolerate their growth, without failure, under specified residual strength loads.

(xi) Flaw is an imperfection, defect, or blemish and may be either discrete or intrinsic.

(xii) Inspection interval is the maximum period of usage allowed for a structure between inspections. At the end of this period, the structure is inspected and if there is no damage detected, the structure may be returned to service for another inspection interval.

(xiii) Intrinsic flaw is a flaw that is inherent in the design and manufacture of the part, situated within it or peculiar to it, such as inclusions, cracks, forging laps, or porosity.

Page C - 100 12/2/2011 AC 29-2C, Chg 4 (xiv) Limit Loads are the maximum loads to be expected in service, as defined in § 29.301(a).

(xv) Multiple Load Path is identified with a redundant structure of multiple and distinct elements, in which the applied loads would be safely redistributed to other load carrying members after complete failure of one of the elements. These may be Active, where two or more elements are loaded during operation to a similar load spectrum, or Passive, where one or more of elements of the structure are relatively unloaded until failure of the other element(s).

(xvi) Principal Structural Elements (PSE) are structural elements that contribute significantly to the carrying of flight or ground loads and the fatigue failure of which could result in catastrophic failure of the rotorcraft.

(xvii) Residual Strength is the level of strength retained by a structure with damage present.

(xviii) Retirement (Replacement) Time of a component is that number of events such as flight hours or landings at which the part must be removed from service regardless of its condition.

(xix) Safe-Life is the number of events, such as flight hours or landings, for a structural component during which there is a low probability that the strength will degrade below its design ultimate value due to fatigue damage initiating cracks.

(2) General. The objective of fatigue tolerance evaluation is to prevent catastrophic failure of the structure by mitigation of the effects of damage in combination with fatigue throughout the life of the rotorcraft.

(i) Fatigue tolerant design as substantiated by fatigue tolerance evaluation methods such as those outlined in this guidance is required for all PSE’s, unless it entails such complications that an effective structure that is tolerant to damage cannot be achieved within the limitations of geometry, inspectability, or good design practice.

In such cases, the particular type of damage at issue must be identified and alternative measures should be taken to minimize both the risk of acquiring that damage and its consequences.

(ii) To perform an evaluation first requires an understanding of the potential threats (resulting in damage) that may modify the fatigue behavior of the component.

The principal concerns of this guidance are consideration of all damage sources and of the fatigue loads and rotorcraft usage. Further mitigation of the sources of damage may be achieved by adoption of a critical parts plan to help ensure that the condition of the part remains as envisaged by the designer throughout its life cycle (see § 29.602).

Page C - 101 AC 29-2C, Chg 4 12/2/2011 (iii) The need for the use of complex inspection techniques or equipment or highly trained personnel (resources that may not be available to the small operator or in remote areas of operation) should be considered when establishing the methodology.

When inspections cannot be relied upon for detection of small cracks or other damage, then retirement times must be established that account for the probable types and locations of the damage, including consideration of cracks.

(iv) A retirement time should be provided for all components, including those subject to inspection, whose fatigue behavior is not reliably established to a point well beyond the life of the rotorcraft. This is intended to prevent the continued use of components beyond the point that ultimate load capability may no longer be assumed to exist in the rotorcraft due to the onset of fatigue cracking. This is particularly important for single load path components or a structure prone to widespread fatigue damage.

(v) Experience with the application of methods of fatigue tolerance evaluation indicates that a relevant test background should exist in order to achieve the design objective. It is general practice within industry to conduct tests to obtain design information and for certification purposes. Damage location, fatigue characteristics, and crack growth data based on test results and service history of similar parts, if available, should be considered when establishing inspections and retirement times.

The FAA should agree upon the extent of supporting evidence necessary for each phase of the evaluation process outlined below.

(3) Essential Considerations. In order to satisfy the requirements of § 29.571, consideration should be given to the following issues in order to demonstrate compliance.

(i) Selection of PSE. All structure, structural elements, and assemblies, the failure or undetected failure of which could result in catastrophic failure of the rotorcraft, should be identified as PSE [see paragraph f.(2)]. To do this, a failure mode and effects analysis or similar method may be used. Specific areas of interest within the PSE that may require particular attention include the following: (A) irregularly shaped parts, or those containing numerous or super - imposed fillets, holes, threads, or lugs; (B) parts of unique design for which no past service experience is available; (C) new materials or processes for which there is no previous experience; (D) bolted or pinned connections; (E) parts subject to fretting; Page C - 102 12/2/2011 AC 29-2C, Chg 4 (F) complex casting; and (G) welded sections.

(ii) In-flight measurement to determine the loads or stresses (steady and oscillatory) for the PSEs in all critical conditions throughout the range of limitations in § 29.309 (including altitude effects), except that maneuvering load factors need not exceed the maximum values expected in operations. See paragraph f.(3).

(iii) Loading spectra as severe as those expected in operation including external load operations, if applicable, and other high frequency power cycle operations. See paragraphs f.(3) and f.(4).

(iv) A threat assessment of probable damage, including a determination of the probable locations, types, and sizes should be performed. In particular, the assessment should include an evaluation of the details of the specific work processes used on each component, operational environment, and maintenance practices to determine the potential for damage. See paragraph f.(5).

(v) Inspectability of the rotorcraft, inspection methods, and detectable flaw sizes should be compatible with the chosen fatigue tolerance methods and validated by trials conducted under realistic conditions. See paragraph f.(6).

(vi) For each PSE, one or more fatigue tolerance methodologies should be selected to ensure each specific damage resulting from the threat assessment is addressed and to satisfy the requirement for inspections and retirement times as discussed in paragraph e. of this guidance. The fatigue tolerance characteristics (including variability) of the structure and materials therein should be evaluated as necessary to support the evaluation. Generally, this will include understanding the fatigue strength, fatigue crack propagation characteristics of the materials used, and the structure and the residual strength of the damaged structure. See paragraphs e., f.(7) and f.(8).

(vii) Fatigue Tolerance Results of the evaluation should be used to provide data in the Limitations Section of the Instructions for Continued Airworthiness. See paragraph f.(9).

e. Fatigue Tolerance Evaluation . A fatigue tolerance evaluation, by analysis and tests, of the PSE is required to establish inspections and retirement times, or approved equivalent means, to avoid catastrophic failure due to fatigue cracking during the operational life of the rotorcraft. The evaluation should consider the impact of the probable threats identified on the fatigue performance and residual strength of all critical areas of each PSE. A number of different fatigue evaluation methods have evolved over the years. Seven of these methods are recognized and discussed in detail in this guidance. The seven methods are summarized as a table in Figure AC 29.571B-1.

Also noted in the table is the safety management strategy the specific method supports, Page C - 103 AC 29-2C, Chg 4 12/2/2011 the analysis category in which they belong, and whether the specific method can be used to address the types of damage identified in the threat assessment.

(1) Each approach results in information that can be used to support establishment of retirement times or inspection requirements. Four methods are used to support safety-by-retirement strategies and they result in retirement times. The other three methods are used to support safety-by-inspection strategies and the result is in-service inspection requirements.

(2) In some cases, application of one method may be sufficient to achieve acceptable fatigue tolerance. In other cases more than one method may be needed.

For example, use of Safe-Life Retirement in combination with Crack Growth Inspections could be an effective way to manage fatigue due to all possible sources.

(3) All the methods listed, with the exception of Safe-Life Retirement, were developed to explicitly address some level of damage. All the methods can theoretically be implemented analytically or by test. However, some of the methods are more practically implemented analytically and some are best implemented by test.

Page C - 104 12/2/2011 AC 29-2C, Chg 4 Method PARAGRAPH Strategy Analysis Threat Category Assessment Results Safe - Life Retirement e.(6)(i)(A) Retire Crack Not Included Initiation Safe-Life Retirement e.(6)(i)(B) Retire Crack Not Including with BDF(s) Initiation Cracks Safe-Life Retirement e.(6)(i)(C) Retire Crack Not Including with CDF(s) Initiation Cracks Safe-Life Inspection e.(6)(i)(D) Inspect Crack Included for CDF(s) Initiation Safe-Life Inspection e.(6)(i)(E) Inspect Crack Included if for a failed element Initiation Considered for all Elements Crack Growth e.(6)(ii)(A) Retire Crack Included if Retirement Growth Crack Bounds Damage Crack Growth e.(6)(ii)(B) Inspect Crack Included Inspection Growth Figure AC 29.571B-1: Seven Fatigue Evaluation Methods discussed in this guidance (4) From an analytical standpoint, these methods fall into one of two categories, crack initiation or crack growth. Each of the seven methods is briefly described below in paragraphs e.(6)(i) and e.(6)(ii), depending on the category.

(5) In-service experience may be used to support establishing fatigue tolerance characteristics when it is shown on a similar structure.

(6) Fatigue Evaluation Methods.

Page C - 105 AC 29-2C, Chg 4 12/2/2011 (i) Crack Initiation Methods. The methods described in this section are categorized as crack initiation methods since they involve quantifying the time it takes for a crack to initiate at a critical area in an as-manufactured part or at a critical area that has sustained some level of damage. Analytically these methods depend on fatigue data (e.g., stress versus number of cycles (S-N) curves) and cumulative fatigue damage algorithms (e.g., Miner’s Rule) to establish a high margin retirement time.

Testing that supports these methods employs specimens that are as-manufactured or ones that have been preconditioned with damage as identified in the threat assessment.

(A) Safe-Life Retirement. Safe-Life Retirement is a crack initiation method that accounts for damage induced by fatigue loading but does not account for flaws and defects due to manufacturing and in-service conditions. Application of this method results in a replacement time based on the time to initiate a crack in an as- manufactured part. Analysis or tests may be used to determine the crack initiation life.

The rationale behind this method is based on part replacement before the probability of initiating a crack becomes significant. This method needs to be supplemented by other methods to account for damage. For compliance details, see paragraph f.(7)(i).

(B) Safe-Life Retirement with a Barely Detectable Flaw (BDF). Safe-Life Retirement with a BDF is a crack initiation methodology that explicitly addresses the effect of damage that is considered barely detectable and is therefore likely to go unnoticed for the life of the part. Application of this method results in a replacement time based on the time to initiate a crack from a BDF. Analysis or tests may be used to determine the crack initiation life. The rationale behind this method is based on part replacement before the probability of initiating a crack is significant. Damage in excess of the BDF must be addressed using other methods. For compliance details, see paragraph f.(7)(ii).

(C) Safe-Life Retirement with a Clearly Detectable Flaw (CDF). Safe-Life Retirement with a CDF is a crack initiation methodology that explicitly addresses the effect of damage that is considered clearly detectable but conservatively recognizes that it would remain in place without corrective action prior to the retirement time of the part. Application of this method results in a retirement time based on the time to initiate a crack from a CDF. Analysis or tests may be used to determine the crack initiation life.

The rationale behind this method is based on part replacement before the probability of initiating a crack is significant. Use of this method by itself could achieve acceptable fatigue tolerance and may preclude the need for any mandated directed inspections.

See paragraph f.(7)(iii) for compliance details.

(D) Safe-Life Inspection for a CDF. Safe-Life Inspection for a CDF is a crack initiation method that explicitly addresses the effect of damage that is considered clearly detectable and would therefore not be expected to remain in place without corrective action for any significant period of time. Application of this method results in a directed inspection task with an interval based on the time to initiate a crack from a clearly detectable flaw. Analysis or tests may be used to determine the crack initiation Page C - 106 12/2/2011 AC 29-2C, Chg 4 life. The rationale behind this method is based on visual detection and disposition of the flaw before the probability of initiating a crack is significant. Damage that is not detectable must be addressed by other methods and the cumulative effects of fatigue prior to and following the advent of the damage should be considered. For compliance details, see paragraph f.(8)(i).

(E) Safe-Life Inspection for a failed element. Safe-Life Inspection for a failed element is a crack initiation method. It results in an inspection for a completely failed load path with an interval based on the crack initiation life of the adjacent structure accounting for internal load redistribution due to failure of the load path that is to be inspected. This method can only be applied if the structure is initially designed for limit load capability with the failed element. The rationale behind this method is based on visual detection and disposition of the failed load path before the probability of initiating a crack in the adjacent structure becomes significant. Therefore it may not be appropriate if the damage that has led to the failure of the first load path could similarly affect the remaining path. For compliance details, see paragraph f.(8)(iii).

(ii) Crack Growth Methods. The methods described in this paragraph are categorized as crack growth methods since they involve quantifying the time it takes a crack at a critical area to grow from some initial size to some final size. Analytically these methods depend on crack growth rate properties (e.g., da/dN vs. ΔK vs. R) and fracture properties (e.g., K ). Using these properties, Fracture Mechanics based tools IC are used to predict crack growth and final fracture. Testing that supports these methods employs specimens that contain cracks and involves close monitoring to document actual crack growth and final fracture.

(A) Crack Growth Retirement is a crack growth method that explicitly addresses the largest damage that could occur during manufacture or operation of the rotorcraft. This damage is modeled as a bounding equivalent crack (BEC) established based on the results of the threat assessment. Application of this method results in a retirement time based on the time for the initial crack to grow large enough to reduce the residual strength to design limit level. Since typical BECs are relatively small and thus difficult to induce in test specimens, this method is typically implemented analytically. The rationale behind this method is based on part retirement before the largest probable damage, modeled as a crack, would reduce the residual strength below design limit. Use of this method by itself could achieve acceptable fatigue tolerance and preclude the need for any mandated inspections provided all threats are accounted for by the BECs. For compliance details, see paragraph f.(7)(iv) (B) Crack Growth Inspection is a crack growth method that explicitly addresses damage that could occur during manufacture or operation of the rotorcraft.

An in-service inspection method is selected that defines a detectable crack size, which could be as large as a completely failed load path. An inspection interval is established based on the time for the detectable crack to grow to critical size or for the residual strength of the adjacent structure to drop to design limit due to continuing crack growth in it. This method is applicable to single or multiple load path structure and inspection Page C - 107 AC 29-2C, Chg 4 12/2/2011 for a completely failed load path or less. This method may be addressed by analysis supported by test depending on the difficulty of introducing into the specimen the inspectable crack or failed load path. The rationale behind this approach is based on detection and disposition of a crack or failed load path before residual strength is reduced below the design limit load. For compliance details, see paragraph f.(8)(ii).

f. Means of Compliance .

(1) General. The results of the fatigue tolerance evaluation required by § 29.571 are used to establish operational procedures that are meant to minimize the risk of catastrophic failures during the operational life of the rotorcraft. It is required that the evaluation performed considers the effect of damage that could result from potential threats present during manufacture and operation. An assessment of probable threats is required to identify the damage that must be considered in the fatigue tolerance evaluation.

(i) The fatigue tolerance evaluation should establish both retirement times and inspection intervals, or approved equivalent means, to prevent any catastrophic failures. Retirement times should be set to ensure that baseline ultimate strength capability is not compromised for as-manufactured structures and structures where the damage is likely to be undetected during the operational life. Intervals for inspections for detectable damage must be established so that strength capability will never fall below maximum design limit level. The intent is that if damage does occur, the structure will retain the capability to withstand reasonable loads without catastrophic failure or excessive structural deformation until the damage is detected and the structure is replaced or repaired. If inspections cannot be established within the limitations of geometry, inspectability, or good design practice, then supplemental procedures, when available, should be established that would minimize the risk of damage being present or leading to a catastrophic failure.

(ii) The following considerations will assist the successful design of a fatigue tolerant structure.

(A) Use multiple-element and multiple load path construction with provisions for crack stoppers that can limit (arrest) the growth of cracks while maintaining adequate residual strength.

(B) Select materials and stress levels that preclude crack growth or crack initiation from flaws or that provide a controlled slow rate of crack propagation combined with high residual strength after initiation of cracks. Test data should substantiate material properties.

(C) Design for detection of damage (i.e., cracks and flaws) and retirement or repair.

Page C - 108 12/2/2011 AC 29-2C, Chg 4 (D) Provide provisions that limit the occurrence of damage and the probability of concurrent multiple damage, particularly after long service.

(iii) Section 29.571 requires that the applicant's proposed compliance methodology must be submitted and be approved by to the Administrator. Therefore, the applicant should coordinate the involvement of the FAA from an early stage. The proposed means of compliance should include the following items.

(A) A list of PSEs to be evaluated.

(B) The results of threat analyses for each PSE including type, location, and size of the damage that will be considered in order to establish retirement times, inspections, or other procedures.

(C) Inspection criteria that includes an estimate of detectability or inspectability, along with any supplemental procedure to minimize the risk of damage.

(D) The analysis methods and supporting test data that will establish retirement times, inspections, or other procedures.

(2) Identification of PSE. The fatigue tolerance evaluation should first consider all airframe structure and structural elements, and assemblies in order to identify the PSE. The structural elements and assemblies identified as PSE should be formally submitted to the FAA with justification for the PSE based on good design practice, service history with similar structure, drawing reviews, static analysis issues, or other appropriate means.

(i) A Failure Mode and Effects Analysis or similar method may be used to identify structures whose failure due to fatigue can lead to catastrophic failure of the rotorcraft. The need to design a PSE for fatigue tolerance when they are supplied by third parties (e.g., actuators) should be clearly identified in the rotorcraft manufacturer’s specification for the part. The list of PSE will likely include structural elements and assemblies that will be subjected to significant fatigue loading expected during the operational life of the rotorcraft. This may include the following rotorcraft parts: (A) Rotors: blades, hubs, hinges, attachment fittings, vibration dampening devices; (B) Rotor drive systems (parts connecting rotors to engines): gears, shafts, gear housings, couplings; (C) Rotor control systems: actuators, pitch control system, swashplate, servo flaps; (D) Fuselage (airframe): rotor system support structure, landing gear attachment; Page C - 109 AC 29-2C, Chg 4 12/2/2011 (E) Fixed and movable control surfaces: stabilizer; (F) Engine, transmission or equipment mountings: APU, auxiliary gearbox; (G) Landing gear; (H) Folding systems: main blade, tail beam.

(ii) Analyses and fatigue tests on complete structures or representative sub- element structures can determine the locations within PSE that need to be identified for fatigue tolerance evaluation. The following should be considered: (A) Strain gauge data on undamaged structure that can identify high stress points.

(B) Analysis that shows high stress or small margin of safety values.

(C) Locations where permanent deformation occurred in static tests.

(D) Locations where failure has occurred in as-manufactured structure fatigue tests.

(E) Locations where the potential for fatigue damage has been identified by analysis.

(F) Locations where the maximum allowed stress occurs when an adjacent element fails.

(G) Locations in structure needed to maintain adequate residual strength that has high stress concentration values.

(H) Locations where detection would be difficult.

(I) Locations where service experience with similar components indicates potential for fatigue or other damage (e.g., fretting, corrosion, wear).

(3) Flight Loads Measurement Program. The simulation of expected spectrum loads for each PSE should be based on flight recorded strain gauge data collected as part of a structured flight test program. The PSE spectrum loads include the steady state, transient, and vibratory loads that are expected in operation. AC 27-1B MG 11, provides further detail for development and use of flight measured loads as the basis for spectrum loads used in the fatigue tolerant evaluations.

(4) Rotorcraft Usage Spectrum.

Page C - 110 12/2/2011 AC 29-2C, Chg 4 (i) The usage and loading spectrum should be developed so that it is unlikely that the actual usage and loads will cause fatigue damage or crack growth rates beyond those associated with the defined spectrum used in the fatigue tolerance evaluation.

The usage spectrum allocating percentage of time or frequencies of occurrence to flight conditions or maneuvers should be based on the expected usage of the rotorcraft.

Considerations should include flight history, recorded flight data, design limitations established in static strength requirements, and recommended operating conditions and limitations specified in the rotorcraft flight manual.

(ii) The fatigue load spectrum developed for fatigue testing and analysis purposes should be representative of the anticipated service usage. Low amplitude load levels that can be shown not to contribute fatigue damage may be omitted (truncated). Simplification of the spectrum loads may also include summing (binding) of percent times or cycles with common steady and vibratory load values.

(iii) The steady state, transient, and vibratory flight load assigned to each regime in the spectrum and utilized in the fatigue tolerance evaluations for each condition should take into account combinations of altitude, center of gravity (CG), gross weight (GW), airspeed, etc., considered to be representative of expected GW/CG mission configurations.

(iv) The usage spectrum should be presented to the FAA for their concurrence. It should include normal operation over the range of rotorcraft configurations including a percent time under ‘external load’ conditions. This spectrum should represent a “composite worst-case” compilation that includes all of the critical conditions that the rotorcraft is expected to experience during performance of the design missions.

(v) AC 27-1B MG 11, provides further detail for the development of the usage spectrums used in the fatigue tolerance evaluations.

(5) Threat Assessment.

(i) A determination should be made of all potential threats that could occur during the manufacturing and service life that may cause damage to each PSE. A threat assessment should be performed for each PSE. To acquire sufficient knowledge of the component and of its global environment, the following items must be identified: (A) manufacturing process.

(B) quality control process.

(C) prescribed storage, transport, handling, assembly and maintenance aspects of the component, and of the surrounding components.

Page C - 111 AC 29-2C, Chg 4 12/2/2011 (D) operational environment.

(E) potential for corrosion including that from contamination by corrosive fluids.

(F) potential for impact damages from debris, dropped tools, hail, tramping underfoot during maintenance, etc.

(G) potential for wear.

(ii) To determine types, locations, and sizes of the probable damages, considering the time and circumstances of their occurrence, the following should be considered: (A) Intrinsic flaws and other damage that could exist in an as-manufactured structure based on the evaluation of the details and potential sensitivities involved in the specific manufacturing work processes used.

(B) Damage that could be expected to occur during prescribed activities associated with storage, transport, handling, assembly, maintenance, overhaul, repair and operation of the component and of the surrounding components including impacts, scratches, fretting, corrosion, contamination, wear, and loss of bolt torque.

(C) Previous experience and data collected on similar events and on similar components; materials, and processes should be considered in identifying risks and causes of damages and their effects in inducing flaws or cracks.

(D) Metallurgical evaluations, manufacturing records and overhaul and repair reports, field service reports, incident and accident investigations, and engineering judgment may be used as supporting data.

(E) When data are not available, the threat should be experimentally simulated and the effect established through tests and analysis. With agreement of the FAA, an upper cut-off value may be established for each class of damage.

(F) Credit may be given to manufacturing, transport, handling, installation, and maintenance instructions finalized to minimize or avoid damages. Examples of these processes or instructions could be: "frozen manufacturing processes," Flight Critical Parts programs, material selection to mitigate intrinsic flaws like inclusions and defects, procedures to reduce deviations from nominal structures, etc.

(G) Credit may be given to protection of structures, such as the use of protective coatings, shielding and plating against corrosion, fretting, and impacts.

Page C - 112 12/2/2011 AC 29-2C, Chg 4 (H) Critical areas will be assumed as a typical location of the damage, unless proper justification is provided to limit the applicability to specific areas or sections of the part.

(iii) Classification of Damage.

(A) The results of the threat assessment are used to classify the damage used in the fatigue tolerance evaluation. The process employed to classify the damage will depend on the fatigue tolerance evaluation method to be used. Depending on the method, a BDF, a CDF, a BEC, or an initial inspectable crack must be established.

(B) For each damage type identified, the sizes to be considered should be representative of the maximum sizes that might not be detected by the inspection techniques established for the component. Sizes exceeding those that are likely to occur do not need to be considered. Standard sizes of damage or standard level of aggression may be derived from previous experience. Each applicant will be required to present justification for damage and crack sizes to be used in the fatigue tolerance evaluations. Within the operational life, defect sizes that have been found in service should be correlated with the sizes used in the design certification.

(C) Barely Detectable Flaw (BDF). For retirement time analysis, flaw sizes that are “barely detectable” may be used to conservatively represent the worst case of undetectable flaws. Alternatively, when the detectable size is larger than the one identified by the threat assessment, a smaller size, but one not less than the flaw size likely to occur, can be used. Sometimes an “allowable” detectable size is established as acceptable for a specific manufacturing process, such as castings, to remain in place for the life of the structure. When it is impossible to simulate that maximum allowable size in the test specimen, the sizes available in the specimen may be used, provided the subsequent analysis of the test result conservatively accounts for the shortfall in the damage size.

(D) Clearly Detectable Flaw (CDF). For inspection intervals, flaw sizes that are “clearly detectable” may be used. The largest discrete size of a CDF to be considered may be limited to the maximum size of the CDF that is likely to remain in place for a significant period of time and not be detected during routine inspections for general conditions and normal observations by knowledgeable personnel. The damage size used may be limited to the maximum probable size identified in the threat assessment. For multiple load path structure, the number of failed load paths to be considered should be established.

(E) Bounding Equivalent Crack (BEC). A Bounding Equivalent Crack must be defined to determine a retirement time using the Crack Growth Retirement method.

The size of the BEC should bound the life reducing effect of damage that could occur as a result of manufacturing, maintenance, or the service environment. The size may be established by analytical back calculations from coupon or service fatigue life data accounting for material variability effects in the data. In any case, there should be no Page C - 113 AC 29-2C, Chg 4 12/2/2011 probable damage from any source that would lead to failure of the part in less time that it would take the BEC to reach critical size. Each applicant must justify the BEC sizes used in the analysis; however, there has been some limited experience that indicates that the following BEC sizes could be appropriate.

( 1 ) 0.015 inch or 0.380 mm radius semicircular surface crack for precision-machined mechanical parts ( 2 ) 0.050 inch or 1.270 mm radius quarter-circular corner crack in fastener holes for typical aluminum airframe structure (F) Initial Inspectable Crack. The size and shape of the initial inspectable crack (a ) must be established when the Crack Growth Inspection approach is used.

DET The inspection interval is based on the time for the initial inspectable crack to grow to a size (a ) that would result in catastrophic failure of the rotorcraft if limit loads were CRIT applied. The initial inspectable crack is a function of the inspection method that is used to detect it. Regardless of the inspection method, the probability of detecting this size crack should be high and it should be substantiated.

(6) Inspectability and Inspection Methods. This section provides guidance on selecting and substantiating damage detection methodology for use with the methods of paragraphs f.(8) (Inspection Intervals) and f.(10) (Approved Equivalent Means). The methods of paragraph f.(8) can result in a mandated inspection program that must be included in the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness in accordance with § 29.1529 of the regulatory requirements. Qualified personnel must conduct these inspections at the specified interval using the approved method or methods. Additionally, § 29.571 allows that substantiation may be accomplished by “Approved Equivalent Means,” which is discussed in paragraph f.(10).

These Approved Equivalent Means may include actions that detect damage or flaws indirectly, and are substantiated using the methods of paragraph f.(8). These actions should be shown to be reliable and systematically conducted by knowledgeable personnel. The following are considerations for establishing inspections, inspection methods, or indirect damage detection.

(i) Inspectability. The ease of conducting an inspection should be a design goal for principal structural elements. Design features such as open construction, access panels or ports, or other easy access to fatigue critical areas for needed inspections should be considered. A design that requires disassembly in order to conduct a required inspection, other than during a scheduled maintenance disassembly, should be avoided.

(ii) The specific inspection methods that are used to accomplish fatigue substantiation should be: Page C - 114 12/2/2011 AC 29-2C, Chg 4 (A) Compatible with the threats identified in the threat assessment, paragraph f.(5), and provide a high probability of detection in the threat assessment and their development, under the operational loads and environment.

(B) Consistent with the capabilities, facilities, and resources of the potential operators of the helicopter. The need to conduct complex or difficult field-level inspections should be avoided, especially when the projected usage of the helicopter may include extended periods of operation in remote areas.

(C) Developed and substantiated for each specific application by means of a full-scale test program, or by experience with similar methods in similar applications.

(D) Included in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness in accordance with § 29.1529 as required by § 29.571(g).

(iii) Detectable Damage Size Assessment.

(A) In the case where the substantiation is predicated on the detection of a specific flaw or crack size, an assessment should be conducted to assure that the selected inspection method would be highly reliable in detecting that size of damage in service. This assessment may be based on the known capability of currently available inspection methods and equipment, provided that this capability is verified by a full- scale test program or by experience with the method in service for similar structure and damage.

(B) If the current capability of a specific inspection method is in question, or if the capability of a specific method needs to be extended to a smaller damage size, then a systematic assessment and substantiation of the method for the intended purpose is appropriate. This assessment could include the determination of the Probability of Detection (POD) as a function of damage size and should consider the capabilities of the potential operators of the helicopter and the environment in which the inspections will be conducted.

(iv) Indirect Detection of Damage. Several damage detection procedures are available that could be used as “Approved Equivalent Means” to support substantiation of a structure [reference paragraph f.(10)]. These procedures, if systematically required and conducted by knowledgeable personnel, can be used in conjunction with the methods presented in paragraph f.(8) to achieve the substantiation. Examples of this type of substantiation are: (A) In-flight damage detectable by vibration, noise, or observing a blade - out-of-track tip path plane. Consideration should be given to the background levels of noise and vibration, as well as whether the indication is of a different character (more detectable) rather than just a change in level (less detectable).

Page C - 115 AC 29-2C, Chg 4 12/2/2011 (B) Damage that is obvious in a preflight check or routine visual examination. This could include obvious flaws or cracking, but also could include structure that is found to be loose, broken, or soft when deflected by hand. Other obvious damage detection could include fluid leaks, missing fasteners, structure bent or out of alignment, or jamming of mechanical parts.

(C) Damage that is indicated following flight completion. Spectrographic oil analysis would be an example.

(D) Damage detection by automated means. This includes crack detection by foil, fiber, or wire break, load monitoring (to detect a change in internal load distribution), acoustic emission monitoring, or other on-board sensors that meet the goals of damage detectability and reliability.

(7) Retirement Times. Each of the four methods below provides a means to establish a retirement time for each PSE. The determination of the fatigue tolerance characteristics should include an assessment using the conventional Safe-Life methodology. In addition, this serves as a baseline for comparison to retirement times determined with flaws and defects included, and should be used as the structure’s retirement time if it is the lowest calculated time.

(i) The conventional Safe-Life methodology accounts for damage induced by fatigue loading but does not account for flaws and defects due to manufacturing and in-service conditions. If the retirement time is established using this method, then the damage identified in paragraph f.(5) (as required by § 29.571(d)(iii)) must be addressed by inspections or other equivalent means. Information to guide a fatigue evaluation based on a conventional Safe-Life approach is provided in detail in AC 27-1B MG 11.

The method consists of: (A) Establishing mean fatigue curves (e.g., stress-life or strain-life) based on crack initiation in constant-amplitude or spectrum testing of as-manufactured structure; (B) Establishing working fatigue curves with strength and life margins; and (C) Conducting a cumulative damage working life calculation using known flight loads and estimated usage.

(ii) A Safe-Life retirement time substantiation with BDF provides a safe period of operation of a structure with probable flaws that may remain in place without detection for that period. Barely detectable flaws are intended to conservatively represent a worst-case of undetectable flaws. The substantiation is accomplished by testing and analysis employing conventional Safe-Life methodology except that an intrinsic and discrete critical flaw in critical locations on the structure is considered. It should be noted that this method, since it is a Safe-Life (crack initiation) method, is not appropriate for use when the flaw being considered is already a crack.

Page C - 116 12/2/2011 AC 29-2C, Chg 4 (A) The types, sizes, and locations of flaws to be considered are determined by the threat assessment (paragraph f.(5)). These flaws may be represented by “equivalent flaws” if it is demonstrated that they have the same or a more severe strength-reducing effect than the corresponding representative flaws.

(B) The mean fatigue strength of the structure with flaws may be determined by one of the following three methods: ( 1 ) Testing a full-scale structure with flaws: (i) Representative flaws as determined by the threat assessment, or equivalent flaws if substantiated, are imposed at the critical locations on the structure where flaws are likely to occur.

( ii ) S-N or spectrum safe-life fatigue testing is conducted; see paragraph e of AC 27-1B MG 11.

( iii ) A mean S-N curve with flaws is derived directly from this data.

( 2 ) As-manufactured structure strength modified by the effect of flaws.

(i) A mean strength for as-manufactured structure (without flaws) can be determined using full-scale S-N or spectrum safe-life fatigue testing.

( ii ) The effect of flaws may be determined by analysis, by similarity to components where the effect of the flaws has previously been determined, or by a specimen test program incorporating the pertinent features of the full-scale component.

Consideration should be given to the material form, geometric features, surface finish, and steady and vibratory load levels, in combination with flaws representative of those identified in the threat assessment.

( iii ) The effect of the flaws is combined with the fatigue result determined on the as-manufactured structure without flaws.

( 3 ) Analytical mean strength modified by the effect of flaws: (i) A mean strength for as-manufactured structure (without flaws) can be determined analytically, provided that correlation with a similar design can be accomplished, or if additional conservatism is included in the working curve reductions employed in paragraph f.(7)(ii)(C).

( ii ) The effect of flaws may be determined by analysis, by similarity to components where the effect of the flaws has previously been determined, or by a specimen test program incorporating the pertinent features of the full-scale component.

Consideration should be given to the material form, geometric features, surface finish, Page C - 117 AC 29-2C, Chg 4 12/2/2011 and steady and vibratory load levels in combination with flaws representative of those identified in the threat assessment.

( iii ) The effect of the flaws is combined with the fatigue result analytically determined for the as-manufactured structure without flaws.

(C) Working Curve Determination. Reduction factors should be applied to the mean curve determined above to derive a working fatigue curve. As outlined in AC 27-1B MG 11, working curve reduction factors should include consideration of the number of specimens tested, variability (scatter), previous test data on the same materials or similar structures, as well as service experience. Different reduction factors from those used for conventional Safe-Life methodology may be employed if appropriately justified.

(D) Retirement Time Determination. The working fatigue curve, flight loads (paragraph f.(3)), and usage spectrum (paragraph f.(4)) are used with a cumulative damage analysis such as shown in AC 27-1B MG 11, to calculate a safe retirement time.

(iii) Safe-Life Retirements with Clearly Detectable Flaws.

(A) A retirement time may also be based on flaws larger than the BDF case, up to the clearly detectable size described in paragraph f.(5), if the applicant chooses. This could be the case, for example, if it was desired to allow a specific manufacturing-related flaw of detectable size to remain in place for the life of the structure without further inspection.

(B) The substantiation for this case can be the same as described in paragraph f.(7)(ii), except that the larger flaws selected for the replacement time substantiation are used instead of the BDFs.

(iv) Crack Growth Retirement.

(A) General.

( 1 ) This approach depends on retirement rather than inspection to ensure the continued airworthiness of a PSE. The retirement time is established based on consideration of crack growth characteristics. Fatigue with damage is addressed by timely retirement and there are no explicit inspection requirements that are derived from this approach.

( 2 ) This approach requires demonstration either by analysis, testing, or both, that the BEC (a ), the most severe crack consistent with manufacturing, BEC maintenance, and service environment, will not grow or will not grow to critical size (a ) under the service loading and environment before the structure is retired. The CRIT critical crack size (a ) is established by limit load. The crack should be assumed at CRIT Page C - 118 12/2/2011 AC 29-2C, Chg 4 the critical location, as defined by the largest stress intensity factor range under the expected service loading range including the ground–air–ground cycle. It is recommended that full scale fatigue testing be undertaken to provide an understanding of the fatigue behavior of the component in support of the chosen methodology. In particular it ensures hot spots are identified, which experience has shown analysis often fails to identify.

( 3 ) A threat assessment (see paragraph f.(5)) should be performed to support establishing the BEC size to be used. It is intended that the BEC conservatively bounds the most severe defect resulting from manufacturing, maintenance, or the service environment. That is, there should be no probable defect, from any source, that would lead to failure of the part in less time than it would take the BEC to reach critical size. It should be noted that the resulting crack is a mathematical expedient that may not represent a true physical crack. If the BEC is defined by analytical back calculations from coupon or service fatigue life data, it will be highly dependent on the predictive tool used (i.e., growth algorithm, material data, etc.).

Therefore, the same predictive tool must be used to perform the fatigue tolerance evaluation. When the BEC is based upon test or service data, it must account for material variability in initiation and growth.

( 4 ) To determine the retirement, the BEC should be assumed at the critical location and the crack growth characteristics should be determined for the expected load and environment spectrum. There are three different scenarios that could result from a crack growth assessment and be used for establishing a retirement time. These scenarios are illustrated in Figure AC 29.571B-2, Figure AC 29.571B-3, and Figure AC 29.571B-4.

(B) No Growth. The no crack growth scenario is illustrated in Figure AC 29.571B-2. Here the BEC does not grow when using top-of-scatter crack growth rate data. In this case the retirement time should not exceed the design service life (L ).

DES Page C - 119 AC 29-2C, Chg 4 12/2/2011 E IZ S CK CRA

a

R

L

DES

TIME

Figure AC 29.571B-2: No Growth (C) Slow Growth of Undetectable Crack. Figure AC 29.571B-3 illustrates the scenario where the BEC grows relatively slowly but becomes critical prior to becoming detectable (a ). In this case, the retirement time should be set equal to the DET total crack growth life (L ) divided by a factor N.

T E IZ S CK a DE T CRA L T a CRIT a R TIME Figure AC 29.571B-3: Slow Growth of Undetectable Crack Page C - 120 12/2/2011 AC 29-2C, Chg 4 (D) Slow Growth of Detectable Crack. Figure AC 29.571B-4 illustrates the scenario where the BEC grows to a detectable size (at L ) before becoming critical (at L +L ). In this case, the retirement time should be set equal to the total crack growth 1 2 life (L +L ) divided by a factor N.

1 2 L L 1 2 E a IZ CRIT S CK CRA a DET a R TIME Figure AC 29.571B-4: Slow Growth of Detectable Crack (E) Life Factors for Crack Growth Retirement.

(1) In determining the factor of N to be used for determining the retirement time, consideration should be given to the crack growth data used (e.g., top of scatter data versus average data, number of specimens used to generate data, etc.).

( 2 ) The minimum suggested N value should be N=2 in the case where the conservative top-of-scatter crack growth data are used in the crack growth analysis, or N=4 when the average crack growth data are used in the crack growth analysis, or N=4 when the crack growth life is obtained from the crack growth test of one specimen (for two or more full scale specimens, N=3 of the shortest crack growth life can be used).

( 3 ) It should also be noted that with this approach, the validity of the crack growth threshold, Δ K , is especially important since there is no element of th inspection to ensure continued airworthiness. Consistent with this, additional attention may be required for validating the crack growth threshold value(s) used in the analyses.

Consideration should be given to the influence of the test procedure used to develop values, microstructure, heat treatment, crack size, loading conditions, environment, grain size and orientation, etc. In general, a coupon-testing program may be necessary Page C - 121 AC 29-2C, Chg 4 12/2/2011 to develop a consistent ΔK database and the use of published data may require th additional conservatism.

(8) Inspection Intervals. Each of the following three methods provides a means to establish inspection intervals for detectable damage or detectable damage growth.

The time of the first inspection should coincide with the repetitive interval established unless the applicant can substantiate an alternate time.

(i) Safe-Life Inspection for a CDF provides a safe interval of operation between repetitive inspections for the presence of probable detectable flaws. The substantiation is accomplished by testing and analysis employing conventional Safe-Life methodology except that intrinsic and discrete critical flaws are considered.

The size of flaws considered should be “clearly detectable”, which is intended to be a conservative representation of detectable flaws that could remain in place for the entire interval in spite of routine inspections for general condition. It should be noted that this method, since it is a Safe-Life (crack initiation) method, is not appropriate for use when the flaw being considered is already a crack.

(A) The method described in paragraph f.(7)(iii), Safe-Life Retirements with Clearly Detectable Flaws, may be employed for this case, except that the calculated retirement time is used as a repetitive inspection interval.

(B) The repetitive inspection consists of examination of the structure for the presence of the flaw using the substantiated inspection method. If no flaw is found, the structure may be returned to service for another inspection interval period, up to the established retirement time. If the flaw is found, the structure is retired; or, if a repair procedure for the specific flaw type has been substantiated, the structure is repaired and returned to service for another inspection interval period, up to the established retirement time for the structure.

(C) Substantiation of repairs should include careful consideration as to whether undetectable cracks may now exist and whether the original certification approach is still applicable.

(ii) Crack Growth Inspection. This approach depends on detection of cracks before they become critical to ensure the continued airworthiness of a PSE. While any inspections that are capable of detecting cracks with high reliability may be used with this approach, the criteria stated in paragraph f.(6), Inspectability and Inspection Methods, should be considered in making the selection. The inspection method chosen will define the initial inspectable crack that will be used to perform the fatigue tolerance evaluation. Once the initial inspectable crack is defined, crack growth, and residual strength assessments must be performed to determine the time for the initial inspectable crack (a ) to grow to a size (a ) that would result in a catastrophic DET CRIT failure of the rotorcraft if limit loads were applied. This assessment could be theoretically done analytically or by test; however, in most cases it is performed analytically using fracture mechanics methods. The resulting life for a to grow to DET Page C - 122 12/2/2011 AC 29-2C, Chg 4 a is used to set the inspection interval. This general process applies to both single CRIT and multiple load path structure regardless of the level of inspection (e.g., for complete load path failure or less than load path failure in a multiple load path structure). The details of defining the interval once the crack growth life has been determined are discussed later.

(A) Single Load Path Structure . The time for a detectable crack (a ) to DET grow to critical size (a ) in a structure is denoted as L in Figure AC 29.571B-4. If CRIT 2 this were a single load path structure, the inspection interval would be established as L divided by N. (See paragraph f.(8)(ii)(C) for guidance on values of N.) This interval is valid until the part is retired.

(B) Multiple Element Structure .

( 1 ) Depending on inspectability considerations and residual life characteristics of the structure following a load path failure, it may be beneficial to take advantage of the redundancy of a multiple load path structure. On the other hand, the safety of a multiple load path structure can be managed without taking advantage of its redundancy. In this case, each load path would be considered independently and inspection intervals established for each load path consistent with paragraph f.(8)(ii)(A).

This may be necessary for similarly stressed load paths when damage according to the threat assessment could occur in each element at the same time.

( 2 ) When considering redundancy in a multiple load path structure, two scenarios might be possible; one where the required inspection is for a completely failed load path and one where the inspection is for less than a load path failure. In either case, the remaining life of the secondary load path after primary load path failure is used to determine the inspection interval. Consistent with this, the resulting intervals are only valid until the cumulative fatigue damage or crack growth in the intact structure is taken into account. This issue is illustrated in a crack growth context in Figure AC 29.571B-5. Crack growth in the secondary load path from an initial crack as detailed in paragraph f.(8)(ii)(B)( 3 )(i) will proceed along curve A-B as long as the primary load path remains intact and load redistribution is negligible. However, at the time of primary load path failure, loading on the secondary load path will increase due to load redistribution and crack growth will be accelerated (e.g., subsequent growth from point 1, 2, or 3 depending on if the failure occurs at time t , t or t ). Note that the 1 2 3 residual life, L , in the secondary load path is inversely proportional to the time at which r primary load path failure occurs. This should be considered whenever L is used in r establishing repeat inspection intervals.

Page C - 123 AC 29-2C, Chg 4 12/2/2011 Primary Load Path Failure

L L L Primary Load Path

r1 r2 r3 Crack Growth CRACK SIZE .... -· ....... --- ........ ---·-· _____ .... ---· _ ..... -----·-·-· .... -•-· ------- _________ .,.

·-------------- --------·-------------------------- B Secondary Load Path Crack Growth - Intact Structure A ----- )t_

--- ----------

t t t 1 2 3 CRACK GROWTH LIFE Figure AC 29.571B-5: Decreasing Residual Life in Secondary Load Path for Multiple Element Crack Growth with Inspections.

( 3 ) Inspect for Load Path Failure. If a failed load path is easily detectable and the residual life and strength of the remaining structure is sufficient, this approach may be optimum. Analysis or tests as described in the following paragraphs can determine the inspection interval.

(i) Evaluation by analysis. Figure AC 29.571B-6 illustrates an example of multiple load path structure for which a completely failed load path is easily detectable. The inspection interval is based on the life of the secondary load path (L ) r after primary load path failure at time N . Consistent with this, damage accumulated in F the secondary load path prior to primary load path failure must be accounted for in the analysis. In order to do this within the context of a crack growth analysis, it is necessary to assume some initial crack, of size a , exists in the secondary load path at time zero.

i This initial crack size should be representative of a normal manufacturing quality unless the threat assessment indicates that larger damage could exist. Crack growth accumulated prior to a load path failure is accounted for by calculating the amount of growth, (Δa ), between time zero and N . Load redistribution that may occur prior to N i F F should be considered. The residual life, (L ), then becomes the time for a crack of size r a + Δa to grow to critical size, assuming a complete load path failure has occurred (i.e., i i “failed” condition loads used). It should be noted that the assumed time of load path failure would also represent an upper limit of validity for any repeat inspection period based on L . It is therefore recommended that N be assumed equal to the retirement r F Page C - 124 12/2/2011 AC 29-2C, Chg 4 time for the structure being inspected or the rotorcraft design life if the structure has no declared retirement time. Based on the above, ( A ) Inspection Interval = L /N [For N refer to paragraph r f.(8)(ii)(C)].

( B ) Limit of validity = N (i.e., repetitive inspection time would F not be valid for operation beyond N ).

F Primacy Load Failure ..

~

Secondacy Path Failure Limit CRACK GROWIB Figure AC 29.571B-6: Multiple Load Path Structure Analytical Evaluation to Support Inspection for a Failed Load Path.

( ii ) Evaluation by Test. Figure AC 29.571B-7 illustrates some key points if an inspection for a complete load path failure is to be developed based on testing. The inspection interval is based on the test demonstrated residual life (L ) r subsequent to load path failure. Because the residual life decreases with the time accumulated prior to a load path failure, there will be a limit of validity to the L and it will r be dependent on the time at which a load path failure is simulated, (N ).

D ( A ) The test article should consist of as-manufactured production parts. Representative “well” condition loading should be applied for some Page C - 125 AC 29-2C, Chg 4 12/2/2011 predetermined period of time, (N ). It is recommended that the “well” condition loading D be of sufficient duration so that N /L is not less than the retirement time minus one D SF inspection interval for the structure being inspected or the rotorcraft design life if the structure has no declared retirement time. At the end of this period, the load path that is to be inspected for complete failure should be disabled (e.g., saw cutting, attachment(s) removal, member removal) to simulate its failure. The test should then be restarted with a representative “failed” condition loading. (Note that the external loads may be the same as for the “well” condition if the member failure simulation results in the correct “failed” condition internal load redistribution.) The test should continue until the desired residual life has been achieved or to the time at which the secondary load path can no longer support limit loads without failure, whichever is less, (N ).

( B ) In developing the test spectrum, consideration should be given to proper use of representative loads, truncation of non-damaging loads, inclusion of ground-air-ground cycles, clipping of high magnitude loads, and load sequence.

( C ) Based on the above, ( a ) Demonstrated residual life = L = N -N .

r 0 D ( b ) Repetitive inspection time = L /N [For N refer to r paragraph f.(8)(ii)(C)].

( c ) Limit of validity = N /L .

D SF ( d ) L = 2, Life safety factor.

SF Page C - 126 12/2/2011 AC 29-2C, Chg 4 Primary Load Path Intentionally Disabled ''Failed" Condition Loading ~

u

Secondary Load

<

~ Path Not Capable U "Well" Condition --------------------------.

of Limit Load ~ Loading IN TEST Figure AC 29.571B-7: Multiple Load Path Structure Evaluation by Test to Support Inspection for a Failed Load Path.

( 4 ) Inspect for Less Than a Load Path Failure. Inspection for less than a load path failure may require special non-destructive Inspection (NDI) procedures but will result in longer inspection intervals. Figure AC 29.571B-8 illustrates how inspection intervals could be established on the basis of crack growth and residual strength evaluation.

(i) In this case, the inspection interval is based on the life of the secondary load path (L ) subsequent to primary load path failure at N plus the time (L ) r F P for a detectable crack (a ) in the primary load path to grow to critical size under DET in-service loads. The determination of L is the same as discussed in paragraph r f.(8)(ii)(B)( 3 )(i).

( ii ) Based on the above, ( A ) Repetitive Inspection = (L + L )/N [For N refer to paragraph P r f.(8)(ii)(C)].

( B ) Limit of validity = N .

F Page C - 127 AC 29-2C, Chg 4 12/2/2011 L L r P Failure of Secondary Load Path at Limit Load Failure of Primary Load Path Under Crack Growth in Inservice Primary Load Path Loads Crack Growth in CRACK SIZE Secondary Load Path after Primary a Load Path Failure DET Crack Growth in Secondary Load Path with Primary Load Path Intact D a a i i N F CRACK GROWTH LIFE Figure AC 29.571B-8: Multiple Load Path Structure Analytical Evaluation to Support Inspection for Less than a Failed Load Path.

(C) Safety Factors .

( 1 ) In determining the factor of N to be used for determining the inspection time, consideration should be given to the crack growth data used (e.g., top of scatter data versus average data, number of specimens used to generate data, etc.)

and the capability of the inspection procedure.

( 2 ) The minimum suggested N value should be N=2 in the case where the conservative top-of-scatter crack growth data are used in the crack growth analysis, or N=4 when the average crack growth data are used in the crack growth analysis, or when the crack growth life is obtained from the crack growth test of one specimen (for two or more full scale specimens, N=3 of the shortest crack growth life can be used).

(iii) Safe-Life Inspection for a Failed Element.

(A) A Safe-Life Inspection substantiation for a Failed Load Path provides a safe interval of operation between repetitive inspections for the failed load path. The substantiation is accomplished by testing and analysis employing conventional Safe-Life methodology except that the configuration of the structure substantiated is with the Page C - 128 12/2/2011 AC 29-2C, Chg 4 critical load path inoperative and appropriate flaws imposed on the remainder of the structure, as determined by the threat assessment.

(B) The method described in paragraph f.(8)(i) can be employed for this case with the following differences: ( 1 ) The principal “flaw” considered is failure or loss of the most critical load path. The load path failure can be the result of fatigue cracking, static failure, or a fractured or missing fastener, as determined by the threat assessment, paragraph f.(5).

( 2 ) The remainder of the structure may be representative of normal manufacturing quality unless the threat assessment indicates that larger damage should exist.

( 3 ) The mean strength for the substantiation should be based on the number of cycles from the first load path failure to the first initiation of cracking at any other point in the remaining structure. Any applied load changes or load distribution changes that occur as a consequence of the load path failure should also be included (bending due to increased deflection, for example).

( 4 ) When the remaining structure may have some pre-existing fatigue damage at the time the first load path fails (due to both load paths being highly loaded, for example), this should be factored into the analysis.

( 5 ) The remaining structure after first load path failure must be shown to have limit load capability, considered as the ultimate loading, except in some cases where no retirement life is provided and fatigue damage is expected (see paragraph f.(10).

( 6 ) The inspection conducted is for the failed or missing load path.

(9) Retirement Time and Inspection Interval Schedules.

(i) Based on the evaluations required by § 29.571, inspections, retirement times, combinations thereof, or other procedures have been established as necessary to avoid catastrophic failure. These inspections, retirement times, or approved equivalent means must be included in the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness (ICA) as required by § 29.1529 and Appendix A29.4 of the regulatory requirements. These inspections, retirement times, or a combination of both are normally stated in hours time-in-service, but may be stated in other terms, such as engine starts, landings, external lifts, etc.

(ii) The design service life should be specified in the fatigue evaluation methodology that must be approved by the FAA. In any case, routine inspections for wear, fretting, corrosion, cracking, and service damage are appropriate. These routine inspections should be noted in the ICAs (maintenance manual) but are not required to Page C - 129 AC 29-2C, Chg 4 12/2/2011 be contained within the ALS of the ICAs unless they are structural inspection intervals or related structural inspection procedures approved under § 29.571.

(10) Approved Equivalent Means. The requirement includes the possibility that in place of setting retirement times or inspections for damage, some other means may be used. All proposals for ‘equivalent means’ must be submitted to the FAA for approval. Potentially equivalent means to inspection include, but are not limited to: (i) Indirect detection of damage used to establish a period of safe operation for a structure with the damage present. In this case, the detection is based on the effect of the damage, which may be recognized through: (A) A warning in flight or during maintenance from a specific feature, sensor, or health monitor, including: oil analysis, chip detector, crack detection wire or foil, health monitoring, fluid leaks or pressure change in a sealed chamber; or by (B) Pilot sensitivity to a change in the rotorcraft’s behavior (such as poor blade tracking, noise generation, vibration generation) provided it is well defined and does not require exceptional piloting skills to recognize these behaviors.

(ii) In all cases, an adequate level of residual strength is demonstrated for the period of operation concerned. Generally, limit load will be considered the minimum residual strength requirement. However, load levels less than the critical limit load conditions may be acceptable for consideration of obvious damage sustained in flight and for the completion of that flight only, provided it allows for continued safe flight and landing.

(iii) Two instances are considered here where it may not be necessary to provide a retirement time in the ALS of the ICAs. However, this does not preclude the investigation of fatigue behavior throughout the life of the rotorcraft or of the part if longer.

(A) When fatigue cracking occurs, or is expected to occur, for a specific PSE while in service, then the first approach allows the PSE to operate until the damage is found. Therefore, the inspection must find the damage prior to loss of ultimate load capability. This approach may not be appropriate for a single load path structure. For such a process to be safe, the behavior of the part and associated parts that influence its fatigue behavior must be substantiated for as long as they remain in service. All potential failure modes throughout the life of the rotorcraft must be identified and shown to be consistent, repeatable and addressed by the inspection program. In order to meet the intent of the new fatigue tolerance requirements, a high probability of ultimate load capability is required throughout the lifetime of the component. Therefore, for cracks or other damage that are allowed or highly likely to exist, ultimate load capability should be substantiated for that damage and any growth that may occur during the subsequent inspection period.

Page C - 130 12/2/2011 AC 29-2C, Chg 4 (B) It may be acceptable that a PSE does not have a specific retirement time when the fatigue tolerance of the part, including any damage not controlled by an acceptable inspection program, has been demonstrated to be in excess of the rotorcraft design life to such an extent that no safety benefit arises from imposing that requirement.

(11) Supplemental Procedures.

(i) The requirement states that if inspections, for any of the damage types identified during the threat assessment, cannot be established within the limitations of geometry, inspectability or good design practice, then supplemental procedures must be established that will minimize the risk of each of these types of damage being present or leading to catastrophic failure. When assessing good design practice, measures such as improved protection against impact, scratches, and corrosion should already have been considered. If the part cannot be redesigned to reduce the acquisition and influence of damage, then supplemental procedures should be introduced.

(ii) Supplemental procedures that should be considered include, but are not limited to: (A) Specifying shorter than usual calendar inspection intervals to reduce the probability of occurrence and the extent of the damage.

(B) Improving control of maintenance processes associated with the component and damage type, such as by providing specifically designed tooling and requiring additional quality checks after each operation is performed.

(C) Introducing an overhaul program.

(D) Restricting the allowable repair limits for the part.

(E) Modifying the PSE design based on service experience if this shows the original design assumptions to be overly conservative with respect to demonstrating impracticality at certification.

(F) Specifying a conservative inspection interval, if the calculated interval cannot be established and there are no other alternatives.

Page C - 131 AC 29-2C, Chg 4 12/1/2011 AC 29.573. § 29.573 (Amendment 29- 54) DAMAGE TOLERANCE AND FATIGUE EVALUATION OF COMPOSITE ROTORCRAFT STRUCTURES a. Purpose . This advisory material provides an acceptable means of compliance with the provisions of § 29.573, Amendment 29-54, Title 14 of the Code of Federal Regulations (CFR) dealing with the damage tolerance and fatigue evaluation of transport category composite rotorcraft structures. Paragraph f.(6) specifically addresses the advisory guidance applying to damage tolerance and fatigue evaluation as required by § 29.573, Amendment 29-54. Some information contained in AC 29-2C, MG 8 (Amendment 29-42) is repeated and updated, as appropriate, to preserve the “building block” approach for analyses of composite rotorcraft structure for compliance to § 29.573, Amendment 29-54 . (Supplemental guidance can be found in AC 20-107, “Composite Aircraft Structure.”) These procedures address the substantiation requirements for composite material system constituents, composite material systems, and composite structures common to rotorcraft. A uniform approach to composite structural substantiation is desirable, but it is recognized that in a continually developing technical area, which has diverse industrial roots both in aerospace and in other industries, variations and deviations from the procedures described here may be necessary. Deviations from this advisory material should be coordinated in advance with the Rotorcraft Directorate.

b. Special Considerations . Since rotorcraft structure is configured uniquely and is inherently subjected to severe cyclic stresses, special consideration is required for the substantiation of all rotorcraft structure, including composites. This special consideration is necessary to ensure that the level of safety intended by the current regulations are attained during the type certification process for all structure with special emphasis on composite structure because of its unique structural characteristics, manufacturing quality and operational considerations, and failure mechanisms.

c. Background .

(1) Historically, rotorcraft have required unique, conservative structural substantiation because of unique configuration effects, unique loading considerations, severe fatigue spectrum effects, and the specialized comprehensive fatigue testing required by these effects. Rotorcraft structural static strength substantiation for both metal and composite structure is essentially identical to that for fixed wing structure once basic loads have been determined. However, rotorcraft structural fatigue substantiation is significantly different from fixed wing fatigue substantiation. Since AC 20-107, as developed, applies to both fixed wing aircraft and rotorcraft, it, of necessity, was finalized in a broad generic form. Accordingly, a need to supplement AC 20-107 for rotorcraft was recognized during type certification programs. One significant difference in traditional rotorcraft fatigue substantiation programs and fixed wing fatigue programs is the use of multiple component fatigue tests for rotorcraft programs rather than just one full-scale test. Also, constant amplitude, accelerated load tests are typically used rather than spectrum tests because of the high frequency Page C - 132 12/1/2011 AC 29-2C, Chg 4 loads common to rotorcraft operations. These rotorcraft fatigue tests have traditionally involved the generation of stress versus life or cycle (S-N) curves for each critical part (most of which are subjected to the cyclic loading of the main or tail rotor system) using a monotonic (sinusoidal) fatigue spectrum based on maximum and minimum service stress values. Unless configuration differences or flight usage data dictate otherwise, the monotonic fatigue spectrum’s period is typically based on six ground-air-ground (GAG) cycles for each flight hour of operation. The S-N curves for the substantiation of each detailed part are typically generated by plotting a curved line through three data points (see AC 29-2C, AC 29 MG 11, “Fatigue Tolerance Evaluation of Transport Category Rotorcraft Metallic Structure”). The three data points selected are a short specimen life (low-cycle fatigue), an intermediate specimen life and a long specimen life (high-cycle fatigue). Each raw data point is generated by monotonically fatigue testing at least two full-scale parts to failure or run out for each data point on the S-N curve.

The raw data point values are then reduced by an acceptable statistical method to a single value for plotting to ensure proper reliability of the associated S-N curve.

Order 8110.9, “Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopter and Other Power Transmission Systems” and AC 27-1B, AC 27 MG 11, “Fatigue Evaluation of Rotorcraft Structure”, contain comprehensive discussions of the S-N curve generation process. The rotorcraft S-N curve process contrasts sharply with the fixed wing process of using a single full-scale fatigue article (usually an entire wing or airframe, which constitutes a single full-scale assembly data point), generic material or full-scale assembly S-N data (e.g., Metallic Materials Properties Development and Standardization (MMPDS), formerly the MIL-HDBK-5 for metals; Composites Materials Handbook-17 (CMH-17), formerly the MIL-HDBK-17 for composites; or AC 23-13, “Fatigue, Fail-Safe, and Damage Tolerance Evaluation of Metallic Structure for Normal, Utility, Acrobatic, and Commuter Airplanes”, which replaced AFS-120-73-2 for full-scale assemblies), a non-monotonic spectrum, and relatively large scatter factors to verify or determine the design fatigue life of the full-scale airplane.

(2) Additionally, rotorcraft have employed and mass-produced composite designs in primary structure (typically main and tail rotor blades) since the early 1950’s.

This was 10 or more years before composites were type certificated for primary fixed-wing structure in either military or civil aircraft applications (with some notable limited production exceptions, such as the Windecker fixed wing aircraft). In any case, the early 1950 period was well before a clear, detailed understanding of composite structural behavior (especially in the areas of macroscopic and microscopic failure mechanisms and modes) was relatively common and readily available in a usable format for the average engineer working in this field. It also predated the initial issuance of AC 20-107. Currently, much composite design information is proprietary, either to government, industry or both, and many data gathering methods have not been completely standardized. Consequently, a significant variation from laboratory to laboratory in material property value determination methods and results can exist. The early rotor blade designs (as well as current designs) are by nature relatively low strain, tension structure designs. Also, by nature, these designs are not damage or flaw critical. Thus, by circumstance as much as design, early composite rotor blade and other composite rotorcraft designs incorporated an acceptable fatigue tolerance level of Page C - 133 AC 29-2C, Chg 4 12/1/2011 safety. In the 1980’s, more test data, analytical knowledge, and analytical methodology became available to more completely substantiate a composite design. Current 14 CFR parts 27 and 29 contain many sections to be considered in substantiating composite rotorcraft structure. This advisory material provides the current or updated information from AC 29-2C, MG 8, Amendment 29-42 to supplement the general guidance of AC 20-107 and provides compliance guidance for the requirements of § 29.573 Amendment 29-54 for rotorcraft composite structure.

d. Definitions . The following basic definitions are provided as a convenient reading reference. CMH-17, and other sources, contain more complete glossaries of definitions.

(1) A-Basis Allowable . The “A” mechanical property value is the value above which at least 99 percent of the population of values is expected to fall, with a confidence level of 95 percent.

(2) Accidental Damage. Discrete damage, which may occur in service use or in manufacturing due to impacts or collisions, such as dents, scratches, gouges, abrasions, disbonds, splintering, and delaminations.

(3) Active Multiple Load Path . Structure providing two or more load paths that are all loaded during operation to a similar load spectrum.

(4) Allowables . Both A-basis and B-basis values statistically derived and used for a particular composite design.

(5) As-Manufactured . Product or component that has passed the applicable quality control process and has been found to conform to the approved design within the allowable tolerances.

(6) Autoclave . A closed apparatus usually equipped with variable conditions of vacuum, pressure, and temperature. It is used for bonding, compressing or curing materials.

(7) B-Basis Allowable . The “B” mechanical property value is the value above which at least 90 percent of the population of values is expected to fall, with a confidence level of 95 percent.

(8) Balanced Laminate . A composite laminate in which all laminae at angles other than 0° occur only in ± pairs (not necessarily adjacent).

(9) Bond . The adhesion of one surface to another, with or without the use of an adhesive as a bonding agent.

(10) Catastrophic Failure. An event that could prevent continued safe flight and landing.

Page C - 134 12/1/2011 AC 29-2C, Chg 4 (11) Cocure . The process of curing several different materials in a single step.

Examples include the curing of various compatible resin system pre-pregs, using the same cure cycle, to produce hybrid composite structure or the curing of compatible composite materials and structural adhesives, using the same cure cycle, to produce sandwich structure or skins with integrally molded fittings.

(12) Component . A major section of the airframe structure (e.g., wing, fin, body, horizontal stabilizer), which can be tested as a complete unit to qualify the structure.

(13) Coupon. A small test specimen (e.g., usually a flat laminate) for evaluation of basic lamina or laminate properties or properties of generic structural features (e.g., bonded or mechanically fastened joints).

(14) Cure . To change the properties of a thermosetting resin irreversibly by chemical reaction (i.e., condensation, ring closure, or addition). Cure may be accomplished by addition of curing (crosslinking) agents, with or without a catalyst, and with or without heat.

(15) Damage. A generic term for structural anomalies caused by manufacturing (processing, fabrication, assembly or handling) or service usage. Trimming, fastener installation, or foreign object impact are potential sources of damage, along with fatigue and environmental effects.

(16) Damage Tolerance . The attribute of the structure that permits it to retain its required residual strength for a period of use after the structure has sustained a given level of fatigue, corrosion, accidental or discrete source damage.

(17) Damage Tolerant Fail-Safe . The capability of structure remaining after a partial failure to withstand design limit loads without catastrophic failure within an inspection period.

(18) Damage Tolerant Safe Life . Capability of structure with damage present to survive expected repeated loads of variable magnitude without detectable damage growth and to maintain ultimate load capability throughout service life of the rotorcraft.

(19) Delamination . The separation of the layers of material in a laminate.

(20) Design Limit Loads . The maximum loads to be expected in service, as defined by § 29.301(a).

(21) Detail . A non-generic structural element of a more complex structural member (e.g., specific design configured joints, splices, stringers, stringer runouts, or major access holes).

Page C - 135 AC 29-2C, Chg 4 12/1/2011 (22) Disbond . A lack of proper adhesion in a bonded joint. This may be isolated or may cover a majority of the bond area. It may occur at any time in the cure or subsequent life of the bond area and may arise from a wide variety of causes.

(23) Element . A generic part of a more complex structural member (e.g., skin, stringers, shear panels, sandwich panels, joints, or splices).

(24) Environment . External, non-accidental conditions (excluding mechanical loading), separately or in combination, that can be expected in service and which may affect the structure (e.g., temperature, moisture, UV radiation, and fuel).

(25) Fatigue or Environmental Damage . Structural damage related to fatigue or environmental effects such as delaminations, disbonds, splintering, or cracking.

(25) Fiber . A single homogeneous strand of material, essentially one-dimensional in the macro-behavior sense, used as a principal constituent in advanced composites because of its high axial strength and modulus.

(26) Fiber Volume . The volume of fiber present in the composite. This is usually expressed as a percentage volume fraction or weight fraction of the composite.

(28) Fill . The 90° yarns in a fabric, also called the woof or weft.

(29) Glass Transition . The reversible change in an amorphous polymer or in amorphous regions of a partially crystalline polymer from (or to) a viscous or rubbery condition to (or from) a hard and relatively brittle one.

(30) Glass Transition Temperature . The approximate midpoint of the temperature range over which the glass transition takes place.

(31) Hybrid . Any mixture of fiber types (e.g., graphite and glass).

(32) Impregnate . An application of resin onto fibers or fabrics by several processes: hot melt, solution coat, or hand lay-up.

(33) Intrinsic or discrete manufacturing defects. Intrinsic or discrete imperfections or flaws related to manufacturing operations, processing or assembly, such as voids, gaps, porosity, inclusions, fiber dislocation, disbonds, and delaminations.

(34) Lamina . A single ply or layer in a laminate in which all fibers have the same fiber orientation.

(35) Laminate . A product made by bonding together two or more layers or laminae of material or materials.

Page C - 136 12/1/2011 AC 29-2C, Chg 4 (36) Low Strain Level . As used herein, is defined as a principal, elastic axial gross strain level that for a given composite structure provides for no flaw growth and thus provides damage tolerance of the maximum defects allowed during the certification process using the approved design fatigue spectrum.

(37) Material System . The combination of single constituents chosen (e.g., fiber and resin).

(38) Material System Constituent . A single constituent (ingredient) chosen for a material system (e.g., a fiber, a resin).

(39) Matrix . The essentially homogeneous material in which the fibers or filaments of a composite are embedded in resins, which are mainly thermoset polymers in aircraft structure.

(40) Maximum Structural Temperature . The temperature of a part, panel or structural element due to service parameters such as incident heat fluxes, temperature, and air flow at the time of occurrence of any critical load case, (i.e., each critical load case has an associated maximum structural temperature). This term is synonymous with the term “maximum panel temperature.” (41) Multiple Load Path . Structure providing two or more separate and distinct paths of structure that will carry limit load after complete failure of one of the members.

(42) Passive Multiple Load Path . Structure providing load paths with one or more of the members (or areas of a member) relatively unloaded until failure of the other member or members.

(43) Point Design . An element or detail of a specific design, which is not considered generically applicable to other structure for the purpose of substantiation (e.g., lugs and major joints). Such a design element or detail can be qualified by test or by a combination of test and analysis.

(44) Porosity . A condition of trapped pockets of air, gas, or void within a solid material, usually expressed as a percentage of the total nonsolid volume to the total volume (solid + nonsolid) of a unit quantity of material.

(45) Pre-Preg, Preimpregnated . A combination of mat, fabric, nonwoven material, tape, or roving already impregnated with resin, usually partially cured, and ready for manufacturing use in a final product that will involve complete curing. Pre - preg is usually drapable, tacky, and can be easily handled.

(46) Principal Structural Element (PSE ). A structural element that contributes significantly to the carrying of flight or ground loads and whose failure can lead to catastrophic failure of the rotorcraft.

Page C - 137 AC 29-2C, Chg 4 12/1/2011 (47) Residual Strength . The strength retained for some period of unrepaired use after a failure or partial failure due to fatigue, accidental, or discrete source of damage.

(48) Resin . An organic material with indefinite and usually high molecular weight and no sharp melting point.

(49) Resin Content . The amount of matrix present in a composite by either percent weight or percent volume.

(50) Secondary Bonding . The joining together, by the process of adhesive bonding, of two or more already-cured composite parts, during which the only chemical or thermal reaction occurring is the curing of the adhesive itself. The joining together of one already-cured composite part to an uncured composite part, through the curing of the resin of the uncured part, is also considered for the purposes of this advisory circular to be a secondary bonding operation. (See COCURE).

(51) Shelf Life . The lengths of time a material, substance, product, or reagent can be stored under specified environmental conditions and continue to meet all applicable specification requirements and remain suitable for its intended function.

(52) Strain Level . As used herein, is defined as the principal axial gross strain of a part or component due to the principal load or combinations of loads applied by a critical load case considered in the structural analysis (e.g., tension, bending, bending-tension). Strain level is generally measured in thousandths of an inch per unit inch of part or microinches/inch (e.g., .003 in/in equals 3000 microinches/inch).

(53) Subcomponent . A major three-dimensional structure, which can provide complete structural representation of a section of the full structure (e.g., stub box, section of a spar, wing panel, wing rib, body panel, or frames).

(54) Symmetrical Laminate . A composite laminate in which the ply orientation is symmetrical about the laminate midplane.

(55) Tape . Hot melt impregnated fibers forming unidirectional pre-preg.

(56) Thermoplastic . A plastic that repeatedly can be softened by heating and hardened by cooling through a temperature range characteristic of the plastic, and when in the softened stage, can be shaped by flow into articles by molding or extrusion.

(57) Thermoset (Or Chemset) . A plastic that once set or molded cannot be re-set or remolded because it undergoes a chemical change; (i.e., it is substantially infusible and insoluble after having been cured by heat or other means).

(58) Warp . Yarns extended along the length of the fabric (in the 0° direction) and being crossed by the fill yarns (90° fibers).

Page C - 138 12/1/2011 AC 29-2C, Chg 4 (59) Work Life . The period during which a compound, after mixing with a catalyst, solvent, or other compounding constituents, remains suitable for its intended use.

e. Related Regulatory and Guidance Material .

Document Title FAA Order 8110.9 Handbook on Vib ration Substantiation and Fatigue Evaluation of Helicopter and other Power Transmission Systems AC 27 - 1B, MG 11 “Fatigue Evaluation of Rotorcraft Structure” AC 20 - 107 “Composite Aircraft Structure” AC 21 - 26 “Quality Control for the Manufacture of Compos ite Materials” CMH - 17 “Composite Materials Handbook” AC 29 - 2C, MG 11 “Fatigue Tolerance Evaluation of Transport Category Rotorcraft Metallic Structure” DOT/FAA/CT - 86/39 Whitehead, R.S., Kan, H.P., Cordero, R., and Seather, R., “Certification Testing Me thodology for Composite Structures”, October 1986.

f. Procedures for Substantiation of Rotorcraft Composite Structure . The composite structures evaluation has been divided into eight basic regulatory areas to provide focus on relevant regulatory requirements. These eight areas are: fabrication requirements; basic constituent, pre-preg and laminate material acceptance requirements, and material property determination requirements; protection of structure; lightning protection; static strength evaluation; damage tolerance and fatigue evaluation; dynamic loading and response evaluation; and special repair and continued airworthiness requirements. Original as well as alternate or substitute material system constituents (e.g., fibers, resins), material systems (combinations of constituents and adhesives), and composite designs (e.g., laminates, cocured assemblies, bonded assemblies) should be qualified in accordance with the methodology presented in the following paragraphs. Each regulatory area will be addressed in turn. It is important to remember that proper certification of a composite structure is an incremental, building block process, which involves phased FAA/AUTHORITY involvement and incremental approval in each of the various areas outlined herein. It is recommended that a FAA/AUTHORITY certification team approach be used for composite structural substantiation. The team should consist of FAA/AUTHORITY and cognizant members of the applicant’s organization. Personnel who are composites specialists (or are otherwise knowledgeable in the subject) should be primary team member candidates.

Once selected, it is recommended that team meetings be held periodically (possibly in conjunction with type boards) during certification to ensure the building block certification process is accomplished as intended. The team should assure that permanent documentation in the form of reports or other FAA/AUTHORITY acceptable documents are included in the certification data package. The documentation includes but is not limited to the structural substantiation reports (both analysis and test), Page C - 139 AC 29-2C, Chg 4 12/1/2011 manufacturing processes and quality control, and Instructions for Continued Airworthiness (maintenance, overhaul, and repair manuals). The Airworthiness Limitations Section of the Instructions for Continued Airworthiness is approved by FAA engineering. Engineering practices for many of the areas identified below are available in CMH-17.

(1) The first area is the fabrication requirements of § 29.605: (i) The quality control system should be developed considering the critical engineering, manufacturing, and quality requirements and a guidance standard such as AC 21-26, “Quality Control for the Manufacture of Composite Materials." This ensures that all special engineering, or manufacturing quality instructions for composites are presented, evaluated, documented, and approved, using drawings, process and manufacturing specifications, standards, or other equivalent means. This should be one of the early phases of a composite structure certification program, since this represents a major building block for sequential substantiation work. Some important concepts of AC 21-26 are included below.

(ii) Specific allowable defect limits (e.g., fiber waviness, warp defects, fill defects, porosity, hole edge effects, edge defects, resin content, large area disbonds, and delaminations) for a particular material system component, laminate design, detailed part, or assembly should be jointly established by engineering, manufacturing, and quality, and the associated inspection programs created, validated, and approved for defect detection. Each critical engineering design should consider the variability of the manufacturing process to determine the worse case effects (maximum waviness, disbonds, delaminations, and other critical defects) allowed by the reliability limitations of the approved inspection program.

(iii) If bonds or bond lines such as those typical of rotorcraft rotor blade structure are used, special inspection methods, special fabrication methods, or other approved verification methods (e.g., engineering proof tests - see paragraph f.(6)) should be provided to detect and limit disbonds or understrength bonds.

(iv) Structurally critical composite construction fabrication process and procurement specifications, for fabricating reproducible and reliable structure, must be provided and FAA approved early during the certification process and should, as a minimum, cover the following: (A) Vendor and Qualified Parts List (QPL) Control . Applicants should be able to demonstrate to FAA certification team members (both the manufacturing inspection district office (MIDO) and FAA engineering) at any time, that their quality control systems ensure on a continuous basis, that only qualified suppliers provide the basic material constituents or material systems (e.g., pre-pregs) that meet approved material specifications. Recommended guidelines for qualification of alternate material systems and suppliers are contained in CMH-17. These methods can also be used periodically for qualification status renewals of existing material systems and suppliers.

Page C - 140 12/1/2011 AC 29-2C, Chg 4 (B) Receiving Inspection and In-Process Inspection . Applicants should be able to demonstrate to FAA certification team members (both MIDO and engineering), at any time, that their receiving and in-process quality control systems provide products, which continuously meet approved material and process specifications. Quality systems should be designed with appropriate checks and balances, so that the necessary statistical reliability and confidence levels for the items being inspected (that are specified by engineering) are continuously maintained. This will require periodic standard inspections and engineering characterization tests on basic constituent and material system samples, which should be conducted, as a minimum, on a batch-to-batch basis. The periodic testing necessary to maintain the quality standard should be conducted by the applicants on conformed samples and should be FAA witnessed.

(C) Material System Component Storage and Handling . Applicants should be able to demonstrate to FAA certification team members (both MIDO and engineering), at any time, that their composite material system (or constituent) storage and handling procedures and specifications provide products, which continuously meet approved material and process specifications. Quality systems should be designed with appropriate checks and balances, so that the necessary statistical reliability and confidence levels for the items being inspected (that are specified by engineering) are continuously maintained. This should require, as a minimum, periodic inspections to ensure that proper records are kept on critical parameters (e.g., room temperature “bench” exposure, shelf life) and that periodic basic constituent and material system characterization tests are conducted, on a batch-to-batch basis. The periodic testing necessary to maintain the quality standard should be conducted by the applicants on conformed samples and should be FAA witnessed.

(D) Statistical Validation Level . It is necessary to maintain the minimum required statistical validation level of the quality control system, which should be specified for each critical item or constituent by the approved quality and engineering specifications. The statistical validation level should be defined and approved early in certification. Also, approval and proper usage should be continuously maintained during the entire procurement and manufacturing cycles.

(v) Alternate fabrication and process specifications should be approved and must comply with § 29.605. Any alternate specifications should provide at least the same level of quality and safety as the original specification. Any changes should be presented for FAA approval well in advance of the effective date of the production change.

(2) The second area is the basic raw constituent, pre-preg and laminate material acceptance requirements, and material property determination requirements of §§ 29.603 and 29.613. These criteria require application of the critical environmental limits such as temperature, humidity, and exposure to aircraft fluids (such as fuel, oils, and hydraulic fluids), to determine their effect on the performance of each composite Page C - 141 AC 29-2C, Chg 4 12/1/2011 material system. Temperature and humidity effects are commonly considered by coupon and component tests utilizing preconditioned test specimens for each material system selected. Material “A” and “B” basis allowable strength values and other basic material properties (based on CMH-17 or equivalent procedures) are typically determined by small scale tests, such as coupon tests, for use in certification work. In the case of composites, determination of these basic constituent and material system properties will almost invariably involve the submittal, acceptance, and use of company standards. This is currently necessary because the FAA (new managers of CMH-17) has not completed development of “B” basis allowables for inclusion in CMH-17. Also, test methods vary somewhat from manufacturer to manufacturer; therefore, individual company results will exhibit some scatter in final material property values. Any company standard that is used should meet or exceed related CMH-17 requirements.

Material structural acceptance criteria and property determination should, as a minimum, include the following: (i) Property characterization requirements of all material systems (e.g., pre-pregs, adhesives) and constituents (e.g., fibers, resins) should be identified, documented, and approved. These requirements, once approved, should be placed in all appropriate procedures and specifications such as those in paragraph f.(1).

(ii) Moisture conditioning of test coupons, parts, subassemblies, or assemblies should be accomplished in accordance with CMH-17, other similar approved methods or per FAA approved programs.

(iii) The maximum and minimum temperatures expected in service (as derived from test measurements, thermal analyses on panels and other parts, experience, or a combination) should be determined and accounted for in static and fatigue strength (including damage tolerance) substantiation programs considering associated humidity-induced effects.

(iv) The wet glass transition temperature, T , is an important characteristic g parameter of amorphous polymers, such as epoxies. It is the temperature below which the polymer behaves like a “glassy” solid and above which it behaves like a “rubbery” solid (i.e., it is the temperature at which there is a very rapid change in physical properties). The change from a hard polymeric material to a rubbery material takes place over a narrow temperature range. A composite material will experience a drastic reduction in matrix-controlled mechanical material properties when loaded in this temperature range. Since the resin is the critical structural element in a composite matrix and the T is critical to structural integrity, a T determination is necessary. The g g T margin methodology of CMH-17 should be implemented (i.e., the T should be 50° F g g higher than the maximum structural temperature (see definition)). For any type of resin or adhesive, an acceptable temperature margin using CMH-17 techniques (e.g., consideration of limited high temperature excursions) or equivalent methodologies based on tests or experience, or both, should be established and approved early in the certification process.

Page C - 142 12/1/2011 AC 29-2C, Chg 4 (v) Local design values should be established by analysis and characterization tests and approved for specific structural configurations (point designs), which include the effects of stress risers (e.g., holes, notches) and structural discontinuities (e.g., joints, splices). Proper determination of these values for full-scale design and test should be considered one of the most critical building blocks in substantiating and evaluating a composite structure. These transitional load transfer areas typically produce the highest stresses (and strains) and serve as the initiation sites for many of the failures (including those due to the relatively low interlaminar strength of composites) that occur in service in a full-scale part or assembly. Small scale tests (such as coupon, element, and subcomponent tests), or equivalent approved testing programs, and analytical techniques should be carefully designed, prepared, and approved to evaluate potential “hot spots” and provide accurate simulations and representations of full-scale article stresses and strains in the critical transition areas. Proper certification work in this area will ensure initial safety and continued airworthiness in full-scale production articles.

(vi) The design strain level for each major component and material system should be established so that specified impact damage considerations are defined and properly limited. The effects of the strain levels may be established for each composite material using small-scale characterization tests and the results should be used to establish or verify the maximum allowable design strain level for each full-scale article.

The maximum allowable design strain values selected should also take into account the reliability and confidence levels established for the relevant portions of the quality control system. This methodology is necessary because the amount and size of flaws in the production article may restrict the allowable level of design strain. In a no-flaw-growth design, the maximum specified impact damage and manufacturing flaw size at the most critical location on the part will be a major factor in determining the maximum allowable elastic strain. This design approach is currently selected for nearly all civil and most military applications; since, under normal conditions, only visual inspections are required in the field (unless unusual external damage circumstances such as a hail storm occur) to maintain the initial level of airworthiness (safety).

However, many military applications, because of their demanding missions, employ scheduled field non-destructive inspection (NDI) maintenance, (such as comparative ultrasonics) to ensure that flaw growth either does not occur, is controlled by approved structural repair, or by replacement of affected parts. To date, civil applicants have not requested a flaw growth, phased NDI approach. Therefore, selection of the full-scale article’s design strain limit based on small-scale tests for a no flaw growth design is extremely important.

(vii) Composite and adhesive properties should be determined so that detrimental structural creep does not occur under the sustained loads and environments expected in service. Small-scale characterization tests (such as coupon, element, and subcomponent tests) and analysis, which verify and establish the full-scale design criteria and parameters necessary to ensure that detrimental structural creep in full-scale structure does not occur in service, should be conducted early in certification and should be FAA-approved.

Page C - 143 AC 29-2C, Chg 4 12/1/2011 (viii) Material allowable strength values for full-scale design and testing should be developed using the coupon procedures presented in CMH-17 or equivalent.

The intent is to represent the material variability including the effects that can occur in multiple batches of material and process runs. At least three batches of material samples should be used in material allowable strength testing. Company standards should be prepared, evaluated and FAA-approved early in certification (as part of the building block process), that reflect the material property determination considerations recommended in CMH-17 on an equal to or better than basis.

(3) The third area is the protection of structure as required by § 29.609.

Protection against thermal, humidity, and other environmental effects (e.g., weathering, abrasion, fretting, hail, ultraviolet radiation, chemical effects, accidental damage) should be provided, or the structural substantiation should consider the results of those effects for which total protection is impractical. Determination and approval of worst-case or most conservative operating limits, and damage scenarios should be accomplished.

Appropriate flammability and fire-resistance requirements should also be considered in selecting and protecting composite structure. Usually, a threat analysis is conducted early in the certification process that identifies the various threats and threat levels for which protection must be provided. This data is then used to construct and submit for approval the methods-of-compliance necessary to provide proper structural protection.

(4) The fourth area is the lightning protection requirements of § 29.610.

Protection should be provided and substantiated in accordance with analysis and with tests such as those of AC 20-53, “Protection of Aircraft Fuel Systems Against Fuel Vapor Ignition Caused by Lightning” and FAA Report DOT/FAA/CT-86/8. For composite structure projects involving rotorcraft certificated to earlier certification bases (which do not automatically include the lightning protection requirements of § 29.610), these requirements should be imposed as special conditions. The design should be reviewed early in the certification process to ensure proper protection is present. The substantiation test program should also be established, reviewed and approved early to ensure proper substantiation.

(5) The fifth area is the static strength evaluation requirements of §§ 29.305 and 29.307 for composite structure. Structural static strength substantiation of a composite design should consider all critical load cases and associated failure modes, including effects of environment, material and process variability, and defects or service damage that are not detectable or allowed by the quality control, manufacturing acceptance criteria, or maintenance documents of the end product. The static strength demonstration should include a program of component ultimate load tests, unless experience exists to demonstrate the adequacy of the analysis, supported by subcomponent tests or component tests to accepted lower load levels. The necessary experience to validate an analysis should include previous component ultimate load tests with similar designs, material systems, and load cases.

Page C - 144 12/1/2011 AC 29-2C, Chg 4 (i) The effects of repeated loading and environmental exposure, both of which may result in material property degradation, should be addressed in the static strength evaluation. This can be shown by analysis supported by test evidence, by tests at the coupon, element or subcomponent levels, or alternatively by existing data. Earlier discussions in this AC address the effects of environment on material properties (see paragraph f.(2)) and protection of structure (see paragraph f.(3)). Static strength tests should be conducted for substantiation of new structure. For the critical loading conditions, two approaches to account for prior repeated loading or environmental exposure for structural substantiation exist.

• In the first approach, the large-scale static test should be conducted on structure with prior repeated loading and conditioned to simulate the environmental exposure and then tested in that environment.

• The second approach relies upon coupon, element, and sub-component test data to assess the possible degradation of static strength after application of repeated loading and environmental exposure. The degradation characterized by these tests should then be accounted for in the static strength demonstration test (e.g., load enhancement), or in the analysis of these results (e.g., showing a positive margin of safety with allowables that include the degrading effects of environment and repeated load).

In practice, the two approaches may be combined to get the desired result (e.g., a large-scale static test may be performed at a temperature with a load enhancement factor to account for moisture absorbed over the aircraft structure’s life).

(ii) The strength of the composite structure should be statistically established, incrementally, through a program of analysis and tests at the coupon, element, subcomponent, or component levels. As part of the evaluation, building block tests and analyses at the coupon, element, or subcomponent levels can be used to address the issues of variability, environment, structural discontinuity (e.g., joints, cut-outs or other stress risers), damage, manufacturing defects, and design or process-specific details.

Figure AC 29.573-1 provides a conceptual schematic of tests included in the building block approach. The material stress-strain curve should be clearly established, at least through the ultimate design load, for each composite design. As shown in Figure AC 29.573-1, the large quantity of tests needed to provide a statistical basis comes from the lowest levels (coupons and elements) and the performance of structural details are validated in a lesser number of sub-component and component tests. The static strength substantiation program should also consider all critical loading conditions for all critical structure including residual strength and stiffness requirements after a predetermined length of service (e.g., end of life (EOL)), which takes into account damage and other degradation due to the service period.

Page C - 145 AC 29-2C, Chg 4 12/1/2011 ..

••

COMP ON EN TS

I

I

•

I •

;-

TAIL ROTOR BLADE

'

•

c

·- - - '

SUB-CO MP ONENTS

I I

I I •

'

u, "' a:

·~· :: - : : -: · : .:, ii , :,;: - :· ::~)·tt-:; · _ 7

s

I \

- - - - - ...

\

I

DET AI LS I

I \

I \

I

I \

In

I \

\

I

,,, ____________ -\

/ I ELEME NTS I \

~ 0

!ti

w : SK IN :IE I

~

i----------------l

u, (.)

: I COUPONS I \

ii: w

z

w C)

: - - ~ \

/ ttfi:dif \

I --'

, ~ UJID rrm fZZ3 ea \

I \

~--------------------·

Figure AC 29.573-1: Schematic Diagram of Building Block Tests.

Page C - 146 12/1/2011 AC 29-2C, Chg 4 (iii) Allowables should be evaluated and used as specified in § 29.613.

These allowables may be generated at the lamina, laminate, or specific design feature level (e.g., filled hole, lap joint, stringer run-out), provided they accurately reflect the actual value and variability of the structural strength for the critical failure modes being considered, at each point design where margins need to be established.

(iv) The static test articles should be fabricated and assembled in accordance with production specifications and processes so that they are representative of production structure including defects consistent with the limits established by manufacturing acceptance criteria.

(v) The material and processing variability of the composite structure should be considered in the static strength substantiation. This can be achieved by establishing sufficient process and quality controls to manufacture structure and reliably substantiate the required strength in tests and analyses, which support a building block approach. If sufficient process and quality controls cannot be achieved, it may be necessary to account for greater variability with special factors (§ 29.619) applied to the design. Such factors should be accounted for in the component static tests or analysis.

(vi) It should be shown that impact damage (or other minor discrete source damage) that can be realistically expected from manufacturing and service, but not more than the established threshold of detectability for the selected inspection procedure, will not reduce the structural strength below ultimate load capability. This static strength capability can be shown by analysis supported by test evidence, or by a combination of tests at the coupon, element, subcomponent, and component levels.

Later discussions in this AC address the issues associated with damage in excess of that considered in f.(5) and drops in residual strength below ultimate load capability (see paragraph f.(6) below).

(6) The sixth area is the damage tolerance and fatigue evaluation requirements of § 29.573.

(i) Background. The static strength determination required by §§ 29.305 and 29.307 establishes the ultimate load capability for composite structures that are manufactured, operated, and maintained with established procedures and conditions.

The damage tolerance and fatigue evaluation required by § 29.573 mandates procedures that allow the composite structure to retain the intended ultimate load capability when subjected to expected fatigue loads and conditions during its operational life. The requirements established for the damage tolerance and fatigue evaluation include component replacement times, inspection intervals, or other procedures as necessary to avoid catastrophic failure. These evaluations assume that the baseline ultimate strength capability might be compromised by damage caused by fatigue, environmental effects, intrinsic or discrete flaws, or accidental damage. This damage includes flaws or defects, which may occur in manufacturing or maintenance and which are used to set the ultimate strength capability and establish the Page C - 147 AC 29-2C, Chg 4 12/1/2011 manufacturing acceptance criteria. The damage tolerance assessment establishes standards that allow the static strength capability to degrade below the ultimate strength capability assuming such damage occurs within the operational life of the structure.

However, when this damage occurs, the remaining structure must withstand expected loads without failure or excessive structural deformations until the damage is detected and the component is either repaired to restore ultimate strength capability or retired.

(ii) General. The nature and extent of the required analysis or tests on complete structures and portions of the primary structure can be based on applicable previous fatigue or damage tolerant designs, construction, tests, and service experience on similar structures. In the absence of experience with similar designs, FAA/AUTHORITY approved structural development tests of components, subcomponents, and elements should be performed. The following considerations are unique to the use of composite material systems and should be observed for the method of substantiation selected by the applicant. Rotorcraft structure provides a broad range of composite applications that are quite different in terms of functionality, geometry and inspectability. These include the rotors, the drive shafts, the fuselage, control system components (e.g., push-pull rods), and the control surfaces. When selecting the approach, attention should be given to the composite application under evaluation, the type of potential damage and degradation of the structural design details, the materials used and margin over flight loads. Whatever the approach selected, the following considerations will apply for tests and analysis: (A) The test articles should be fabricated and assembled in accordance with production specifications and processes so that the test articles are representative of production structure.

(B) The test articles should include material imperfections whose extent is not less than the limits established under the inspection and acceptance criteria used during the manufacturing process and consistent with the inspection techniques used in service (e.g., visual, ultrasonic, X-ray). The initial extent of these imperfections should be discussed and agreed with the FAA, taking into account experience in manufacturing and routine in-service inspections. Typical defects to be considered include but are not limited to the following: ( 1 ) Disbonds and weak bonds (considered as disbonds).

( 2 ) Delaminations, fiber waviness, porosity, voids.

( 3 ) Scratches, gouges, and penetrations.

( 4 ) Impact damage.

All of the damages identified in the preceding paragraph (B) above should be derived from the threat assessment described in the following paragraph (C).

Page C - 148 12/1/2011 AC 29-2C, Chg 4 (C) For each PSE, a threat assessment must be made of the probable locations, types, and sizes of damage considering fatigue, environmental effects, intrinsic and discrete flaws, and impact or other accidental damage. This determination must be submitted with accompanying rationale to the FAA/AUTHORITY for approval.

This rationale may include experience with similar materials, designs, processes (manufacturing, maintenance, and overhaul), structural details, or structure, and may also include service failure evaluations, manufacturing records, overhauls and repair reports, field service reports, incident and accident investigations, service impact surveys, inspectability surveys, and engineering judgment. Consideration should also be given to factors that: - Reduce scatter and deviations from nominal structures, such as frozen processes, Flight Critical Parts programs, and materials and manufacturing processes to mitigate intrinsic flaws (inclusions and defects).

- Preclude a type of damage by use of a specific design feature (material selection, surface treatment, protective coating, or shielding), a specific stress level (for fatigue damage), or a specific manufacturing inspection process (if it can be shown to be highly reliable, well-controlled, documented, and systematically required).

The assessment should include: • A systematic evaluation of all the location, types, and sizes of damage and their estimated probability of occurrence.

• A selection or elimination of this damage based on the above estimate.

• A verification that the inspection method selected is capable of detecting the damage at the size and location determined.

The types of damage to consider include: ( 1 ) Intrinsic Flaws (imperfections), which are probable to exist in an as- manufactured structure based on the evaluation of the details and potential sensitivities of the specific manufacturing work processes used. The types of flaws to be considered include voids, disbonds, inclusions, foreign objects, resin-rich and resin- starved areas, and improper ply orientation or ply ending. The sizes of the intrinsic flaws considered should be based on the limits established under the manufacturing inspection and acceptance criteria and are expected to remain in service for the life of the structure.

( 2 ) Impact Damage, which may occur during manufacturing and in service based on an evaluation of the threats by means of an impact survey and/or service experiences. This type of damage can include dents, penetrations, gouges, abrasions, and scratches. A threat assessment is needed to identify impact damage Page C - 149 AC 29-2C, Chg 4 12/1/2011 severity and detectability for design and maintenance. A threat assessment usually includes damage data collected from service plus an impact survey. An impact survey consists of impact tests performed with configured structure, which is subjected to boundary conditions characteristic of the structure. Many different impact scenarios and locations are typically considered in the survey, which has a goal of identifying the most critical impacts (i.e., those causing the most serious damage but are least detectable). When simulating accidental impact damage, blunt or sharp impactors should be selected to represent the maximum criticality versus detectability, according to the load conditions (e.g., tension, compression or shear). Until sufficient service experience exists to make good engineering judgments on energy and impactor variables, impact surveys should consider a wide range of conceivable impacts, including runway or ground debris, hail, tool drops, and vehicle collisions. Service data collected over time, can better define impact surveys and design criteria for subsequent products, as well as establish more rational inspection intervals and maintenance practice. Refer to paragraph f.(6)(ii)(H) for various combinations of detectability and energy levels to be considered in the damage tolerance and fatigue evaluation.

( 3 ) Discrete Source Damage. The structure should be able to withstand limit static loads (considered as ultimate loads) and fatigue loads, which are reasonably expected during a completion of a flight on which damage resulting from obvious discrete source occurs (e.g., hail damage, bird strike, uncontained engine failure, and uncontained high energy rotating machinery failure). The extent of damage should be based on a rational assessment of service mission and potential damage relating to each discrete source.

(D) The use of composite secondary bonding in manufacturing or maintenance requires strict process and quality controls to achieve the reliability needed to use such technology in critical structures (see AC 21-26). Assuming good process and quality controls, service history has shown that additional damage tolerant design considerations are also needed to ensure the safety of structure with secondary bonds (i.e., random, but an unacceptable number of weak bonds discovered in service).

Unless the ultimate strength of each critical bonded joint can be reliably substantiated in production by NDI techniques (or other equivalent, approved techniques), then the limit load capability should be ensured by any or a combination of the following: ( 1 ) Consider isolated disbonds and weak bonds (represented by zero bond strength) in structural elements that use secondary bonding for primary load transfer. The associated disbond size should be up to the limitations provided by redundant design features (i.e., mechanical fasteners or a separate bonding detail).

The structure containing such damage should be shown to carry limit load by tests, analyses, or some combination of both. For purposes of test or analysis demonstration, each disbond should be considered separately as a random occurrence (i.e., it is not necessary to demonstrate residual strength with all structural elements disbonded simultaneously).

Page C - 150 12/1/2011 AC 29-2C, Chg 4 ( 2 ) Each critical bonded joint on each production article should be proof-tested to the critical limit load.

( 3 ) Critical bonded joints that have high static margins of safety (e.g., some rotor blades) may be accepted based on satisfactory service history of like or similar components.

(E) The fatigue load spectrum developed for fatigue testing and analysis purposes should be representative of the anticipated service usage. Low amplitude load levels that can be shown not to contribute to fatigue damage may be omitted (truncated). Reducing maximum load levels (clipping) is generally not accepted.

(F) Environmental effects (temperature and humidity representative of the expected service usage) on the static and fatigue behavior and damage growth should be considered. Unless tested in the environment, appropriate environmental knock down factors for the static and the fatigue test articles should be derived and applied in the evaluation.

(G) Variability in fatigue behavior should be covered by appropriate load or life scatter factors and these factors should take into account the number of specimens tested.

(H) The following Figure AC 29.573-2 illustrates the extent of the impact damage that needs to be considered in the damage tolerance and fatigue evaluation.

Page C - 151 AC 29-2C, Chg 4 12/1/2011 Detectability Obvious Damage ~ Detectable damage [

~

Barely Detectable Impac~ Damage

+

Undetectable damage

+

+

I Zone 1

Energy

I

Level of energy selected for compliance with §§ 29.305, 29.307 (Static Strength) and fl:5)(vi) of this AC.

Maximum level of energy selected for the Damage Tolerance evaluation. This level should not be exceeded in service.

Figure AC 29.573-2: Characterization of Impact Damage.

( 1 ) Both the energy level associated with the static strength demonstration and the maximum energy level associated with the damage tolerance evaluation (depicted in Figure AC 29.573-2) are dependent on the part of the structure under evaluation and a threat assessment.

( 2 ) Obvious impact damage is used to define the threshold from which damage is readily detectable and appropriate actions may be taken before the next flight.

( 3 ) Barely Detectable Impact Damage (BDID) is the state of damage at the threshold of detectability for the approved inspection procedure. Barely Visible Impact Damage (BVID) is that threshold of visually detectable damage associated with a detailed visual inspection procedure.

( 4 ) Detectable Damage is the state of damage that can be reliably detected at scheduled inspection intervals. Visible Impact Damage (VID) is that threshold associated with the type of damage that should be detectable during a detailed visual inspection.

( 5 ) Three Zones are depicted by this figure: Page C - 152 12/1/2011 AC 29-2C, Chg 4 Zone 1: Since the damage is not detectable, Ultimate Load capability is required. The provisions of paragraph f.(5) provide a means of compliance.

Zone 2: Since the damage can be detected at a scheduled inspection, Limit Load (considered as Ultimate load) capability is the minimum requirement for this damage.

Zone 3: Since the damage is not detectable with the proposed in- service inspection procedures, ultimate load capability is required, unless an alternate procedure can show an equivalent level of safety. For example, residual strength lower than ultimate may be used in association with improved inspection procedures or with a probabilistic approach showing that the occurrence of energy levels is low enough so that an acceptable level of safety can be achieved.

Of the three zones, only Zone 3 may have a residual strength requirement that can vary with alternate procedures or the probability of damage occurrence or both. In either case, any compromise for residual strength requirements less than the ultimate load requirement should only be considered when pursuing one of the options under the damage tolerant fail-safe means of compliance, as described in the following section, f.(6)(iii)(B). One example of the use of alternate procedures is for the rare damage threat from a high energy, blunt impact (e.g., service vehicle collision). Depending on the selected maintenance inspection scheme, such damage may fall under Zone 3.

When considering such damage in the design of a part, it may be shown to be damage tolerant fail safe, even though the damage is not detectable, based on a very low probability of occurrence. As a result, the design may have sufficiently high residual strength (e.g., below Ultimate, but well above limit load capability to ensure safety without detection for long periods of time). If it is further determined that such impact events usually occur with the knowledge of maintenance or aircraft service personnel, then the alternate procedures may be added to the Instructions for Continued Airworthiness. For example, advanced inspection methods, which can detect damage from high-energy blunt impacts, may be used as alternate procedures to minimize the risk of catastrophic failure for such Zone 3 damage.

(iii) Means of compliance. For each PSE, inspections, replacement times, or other procedures must be established as necessary to avoid catastrophic failure.

Compliance with the requirements of § 29.573(d) and (e) should be shown by one, or a combination of, the methods described subsequently. Generally, replacement times are established using Damage Tolerance Safe Life Evaluations and Inspection Intervals are established using Fail Safe Evaluations. From current state-of-the-art rotorcraft applications, it is widely accepted that composite materials have good flaw and damage tolerance capabilities and therefore the supplemental procedures may only be rarely necessary. Damage tolerance evaluations are best suited for composite structures, particularly those with structural redundancy and inherent resistance to damage growth.

Damage resulting from anomalous or accidental events must be considered in the damage tolerant evaluations. The damage tolerant evaluation for replacement times and inspection intervals is to be used unless it is established that neither can be Page C - 153 AC 29-2C, Chg 4 12/1/2011 achieved within the limitations of geometry, inspectability, or good design practice. In that case, supplemental procedures must be established and submitted to the FAA for approval. In any case, the FAA must approve the methodology used for compliance to § 29.573. The substantiation method(s) should be chosen so that the structure is protected against catastrophic failure from each of the threats identified in paragraph f.(6)(ii)(C) of this AC by a specific procedure (inspection, replacement time, or other procedure). For example, a manufacturing-related void of a specific allowable size could be substantiated by means of a replacement time method with no scheduled inspection. An accidental impact in the same area could be substantiated by an inspection method with no specific replacement time. The result could be one structure with several different inspection requirements (location, method, and interval) and a fixed replacement time as well. This combination of procedures assures that each threat is covered. The fatigue substantiation should include sufficient coupon, element, sub-element, or component tests to establish the fatigue scatter, curve shapes, and the environmental effects. The substantiation should include full-scale, component, or sub - component fatigue testing but also may be accomplished by analysis supported by test evidence. When spectrum testing is used, the lowest load levels can be eliminated from the spectrum if they can be shown to be non-damaging. The substantiation should include a static strength evaluation to show that the required residual strength and adequate stiffness, accounting for the effects of environment, are retained for the life of the structure or the appropriate inspection interval. Damage as determined in paragraph f.(6)(ii) of this AC for the specific structure being substantiated should be imposed at each critical area of the structure.

(A) Damage Tolerant Safe-Life Evaluation . This is a “No-Growth” method in which the structure, with damage present, is able to withstand repeated loads of variable magnitude without detectable damage growth for the life of the rotorcraft or within a specified replacement time. This evaluation may be used to substantiate any type of damage that will remain in-service for the life of the part. No specific inspection requirements are generated from the test program in this method. However, compliance with routine inspections for cracking, delaminations, and service damage and other limitations prescribed in accordance with § 29.1529 are always required.

Compliance using full-scale, component, or sub-component fatigue testing can be accomplished by either of the following methods: ( 1 ) S-N Method . This method is based on determining the point where initiation of growth occurs for the damage present at critical locations in the structure.

AC 27-1B, AC 27 MG 11, provides guidance that may be appropriate for this method.

The method utilizes one or more full-scale, component, or sub-component test specimens subjected to constant-amplitude or spectrum loading applied in a distribution on the structure that is representative of critical flight conditions. Any indication of growth of the imposed damage and defects, or structurally significant cracking, disbonding, splintering, or delaminating of the composite, defines the fatigue initiation characteristic of the structure in terms of applied load and cycles. Working S-N curves are established from the mean curve using strength or cycle reductions or both to account for fatigue scatter and environmental effects. Flight loads are compared to this Page C - 154 12/1/2011 AC 29-2C, Chg 4 working curve, and if any intercepts occur, a cumulative damage calculation is conducted to establish the component retirement time. Compliance with the ultimate load requirements should be demonstrated at the completion of the fatigue test.

( 2 ) Life-Test Method . This method uses spectrum fatigue testing to verify the absence of damage growth over a large number of cycles that are equivalent to a lifetime of expected usage. The method uses one or more full-scale, component, or sub-component test specimens subjected to spectrum fatigue loading applied in a representative distribution of flight loads, including Ground-Air-Ground (GAG) loads.

Fatigue test loads should be increased by factors for environment and fatigue strength scatter. The load may also be increased using an S-N curve approach to reduce the duration of the test. Any significant growth of the imposed damage, or structurally significant cracking, disbonding, splintering, or delamination of the composite during the test constitutes failure to achieve the desired lifetime. However, the equivalent life demonstrated at the time of inception of damage growth or cracking can be used as a retirement time for the component. Compliance with the ultimate load requirements should be demonstrated at the completion of the fatigue test.

(B) Damage Tolerant Fail-Safe (Residual Strength with Detectable Damage) Evaluation . This method establishes inspection intervals to ensure that the structure remaining after a partial failure is able to withstand design limit loads without failure or excessive structural deformations within a specified inspection interval. If the damage is detected in an inspection, the structure should be either replaced or repaired to restore ultimate load capability. Evaluation of Zone 3 damage should have sufficiently high residual strength and, if necessary, supplemental procedures should be established to minimize the risk of catastrophic failure. Full-scale, component, or sub - component testing should be accomplished using one or more specimens subjected to constant amplitude or spectrum loading applied in a manner representative of flight load conditions. The test loads should be increased by factors that account for environment and fatigue strength scatter. The results of the testing can be used to manage the structure in one or a combination of the three methods described subsequently.

( 1 ) No Growth Evaluation . This approach is appropriate for inspectable in-service damage, which does not grow in service (see Figure AC 29.573-3). Damage growth should be substantiated using either method described in f.(6)(iii)(B)( 2 ) or f.(6)(iii)(B)( 3 ). Structural details, elements, sub-components, and components of critical structural areas, or full-scale structures, should be tested under repeated loads for validating a no-growth approach to the damage tolerance requirements. The number of cycles applied to validate a no-growth concept should be statistically significant, and may be determined by load or life considerations or both. Residual strength testing or evaluations should be performed after repeated load cycling demonstrating that the residual strength of the structure is equal to or greater than limit load considered as ultimate. Moreover, it should be shown that stiffness properties have not changed beyond acceptable levels. Inspection intervals should be established, considering the residual strength capability associated with the assumed damage. The intent of this is to assure that structure is not exposed to an excessive period of time with static margins Page C - 155 AC 29-2C, Chg 4 12/1/2011 less than ultimate, providing a lower safety level than in the typical slow growth situation, as illustrated by the Figure AC 29.573-3. Once the damage is detected, the component is either repaired to restore ultimate load capability or replaced.

Damage init iation

t

Res id u al s tren g th o r o ccu rren ce ~ , Slo w-g ro wt h ap p ro ach * ~ Ult i mat e lo ad s No -g ro wt h ap p ro ach ** Li mit lo ad s * Rep air t o Res t o re U lt i mat e St ren g t h * * No g ro wt h wit h o u t rep air is n o t accep t ab le Ti me

--

-.....

Figure AC 29.573-3: Residual Strength vs. Time.

The lower the residual strength of a structure after an accidental damage event, the shorter the inspection interval should be. Considerations of both inspectability and impact surveys (including probability of occurrence) for specific structure may be used to isolate the most critical threats to consider in setting a maintenance inspection interval. Knowledge of the residual strength for a given critical damage is also needed for such an evaluation. If it is known that the design is capable of handling large and clearly detectable damage, while maintaining a residual strength well above limit load, a less rigorous engineering approach may be applied in establishing the inspection interval.

Slow Growth Evaluation. (2) This method is applicable when the damage grows in t h e test and the growth rat e is shown to be slow, stable, and predictable, as illustrated in Figure AC 29.573-4. An inspection program should be developed consisting of the frequency, extent, and methods of inspection for inclusion in the maintenance plan. Inspection intervals should be established so that the damage will have a very high probability of detection between the time it becomes initially inspectable and the time at which the extent of the damage reduces the residual static strength to limit load (considered as ultimate), including the effects of environment. For any damage size that reduces the load capability below ultimate, the component is either repaired to restore ultimate load capability or replaced. Should functional impairment (such as unacceptable loss of stiffness) occur before the damage becomes Page C - 156 12/1/2011 AC 29-2C, Chg 4 otherwise critical, this should be accounted for in the development of the inspection program.

Ultimate Slow Growth

L_

Static Strength I ______ _ Arrested Growth Limit Critical Size

-----------

_ _J Arrested Growth Damage Size

_J r-·

No Growth Slow Growth i Threshold of Detectability _ ..... ,.......... ,.1 ........ , ... ,....... , ... , ... • .. , ... ,....... , ... ,...... .

time Slow Growth Inspection Interval Basis Figure AC 29.573-4: Illustration of Residual Strength and Damage Size Relationships for Fail-Safe Substantiation.

(3) Arrested Growth Evaluation . This method is applicable when the damage grows, but the growth is mechanically arrested or terminated before becoming critical (residual static strength reduced to limit load), as illustrated in Figure AC 29.573-4. Arrested Growth may occur due to design features such as a geometry change, reinforcement, thickness change, or a structural joint. This approach is appropriate for inspectable arrested growth damage. Structural details, elements, and sub-components of critical structural areas, or full-scale structures, should be tested under repeated loads for validating an arrested growth approach to the flaw tolerance requirements. The number of cycles applied to validate an arrested growth concept should be statistically significant, and may be determined by load or life considerations, or both. Residual strength testing or evaluation should be performed after repeated load cycling and a demonstration that the residual strength of the structure is equal to or greater than limit load considered as ultimate. Moreover, it should be shown that stiffness properties have not changed beyond acceptable levels. Inspection intervals should be established, considering the residual strength capability associated with the arrested growth damage. The intent of this is to assure that structure is not exposed to an excessive period of time with static margins less than ultimate, providing a lower safety level than in the typical slow growth situation, as illustrated by Figure AC 29.573-3. For any damage size that reduces the load capability below ultimate, the component is either repaired to restore ultimate load capability or replaced. The lower the residual strength of a structure after an arrested growth event, the shorter the Page C - 157 AC 29-2C, Chg 4 12/1/2011 inspection interval should be. Considerations of both inspectability and impact surveys (including probability of occurrence) for specific structure may be used to isolate the most critical threats to consider in setting a maintenance inspection interval.

Knowledge of the residual strength for a given critical damage is also needed for such an evaluation. If it is known that the design is capable of handling large and clearly detectable damage, while maintaining a residual strength well above limit load, a less rigorous engineering approach may be applied in establishing the inspection interval.

(C) Combination of Damage Tolerant Safe Life and Fail Safe Evaluations.

Generally, it may be appropriate to establish both a replacement time and an inspection program for a given structure as calculated by the Damage Tolerant Safe Life and Fail Safe Evaluations.

(D) Other Procedures . Other procedures are allowed according to § 29.573(d). Such alternative procedures must still provide the same degree of damage tolerance to the same identified threats as the replacement time or inspection interval methods. One possible alternate approach is the use of indirect damage detection methods instead of the specific mandated inspection procedures that are determined in the Fail Safe Evaluations of f.(6)(iii)(B). These indirect detection methods should be documented and shown to have the same degree of reliability, repeatability, and margin provided by a conventional inspection approach. These methods could include: (1) establishing measurable vibration or blade out-of-track conditions and limits, (2) defining indirect inspections, which would detect damage, and (3) in-flight detecting of damage by means of monitoring and warning devices.

(E) Supplemental Procedures. If the damage tolerant evaluations as described previously cannot be achieved within the limitations of geometry, inspectability, or good design practice, a fatigue evaluation using supplemental procedures may be proposed to the FAA/AUTHORITY per § 29.573(e). The applicant must establish that the damage tolerance criteria are impracticable and cannot be satisfied for the specific PSE, locations, and threats considered. In addition, the types of damage considered in the evaluations must be identified. Finally, supplemental procedures must be established to minimize the risk of catastrophic failure with the damages considered.

(iv) Additional considerations for damage tolerance and fatigue evaluations.

(A) Experience with the application of methods of fatigue and damage tolerance evaluations indicates that a relevant test background should exist in order to achieve the design objective. It is the general practice within industry to conduct damage tolerance tests for design information and guidance purposes. It is crucial that the critical structure be identified and tested to the proper flight and ground loads.

(B) Identification of the structure to be considered in each evaluation (a failure mode and effects analysis or similar method should be used).

Page C - 158 12/1/2011 AC 29-2C, Chg 4 ( 1 ) Identification of Principal Structural Elements. Principal structural elements are those that contribute significantly to carrying flight and ground loads and whose failure could result in catastrophic failure of the rotorcraft. Typical examples of such elements are: (i) Rotor blades and attachment fittings.

( ii ) Rotor heads, including hubs, hinges, and some main rotor dampers.

( iii ) Control system components subject to repeated loading, including control rods, servo structure, and swashplates.

( iv ) Rotor supporting structure (lift path from airframe to rotorhead).

( v ) Fuselage, including stabilizers and auxiliary lifting surfaces, airframe provisions for engine and transmission mountings.

( vi ) Main fixed or retractable landing gear and fuselage attachment structure.

( 2 ) Identification of Locations Within Principal Structural Elements to be Evaluated . The locations of damage to structure for damage tolerance evaluation can be determined by analysis or by fatigue test on complete structures or subcomponents. However, tests will be necessary when the basis for analytical prediction is not reliable, such as for complex components. If less than the complete structure is tested, care should be taken to ensure that the internal loads and boundary conditions are valid. The following should be considered: (i) strain gauge data on undamaged structure to establish points of high stress concentration as well as the magnitude of the concentration; ( ii ) locations where analysis shows high stress or low margins of safety; ( iii ) locations where permanent deformation occurred in static tests; ( iv ) locations of potential fatigue damage identified by fatigue analysis; ( v ) locations where the stresses in adjacent elements will be at a maximum with an element in the location failed; ( vi ) partial fracture locations in an element where high stress concentrations are present in the residual structure; Page C - 159 AC 29-2C, Chg 4 12/1/2011 ( vii ) locations where detection would be difficult; and ( viii ) design details that service experience of similarly designed components indicates are prone to fatigue or other damage.

( 3 ) In addition, the areas of probable damage from sources such as a severe corrosive or fretting environment, a wear or galling environment, or a high maintenance environment should be determined from a review of the design and past service experience.

(C) The stresses and strains (steady and oscillatory) associated with all representative steady and maneuvering operating conditions expected in service.

(D) The frequency of occurrences of various flight conditions and the corresponding spectrum of loadings and stresses.

(E) The fatigue strength, fatigue crack propagation characteristics of the materials used and of the structure, and the residual strength of the damaged structure.

(F) Inspectability, inspection methods, and detectable flaw sizes.

(G) Variability of the measured stresses of paragraph f.(6)(iv)(C), the actual flight condition occurrences of paragraph f.(6)(iv)(D), and the fatigue strength material properties of paragraph f.(6)(iv)(E).

(v) Flight strain measurement program.

(A) General. Subsequent to design analysis, in which aircraft loads and associated stresses are derived, the stress level or loads are to be verified by a carefully controlled flight strain measurement program. (This guidance is similar to that of AC 27-1B, MG 11.)

(B) Instrumentation.

( 1 ) The instrumentation system used in the flight strain measurement program should accurately measure and record the critical strains under test conditions associated with normal operation and specific maneuvers. The location and distribution of the strain gauges should be based on a rational evaluation of the critical stress areas.

This may be accomplished by appropriate analytical means supplemented, when necessary, by strain sensitive coatings or photoelastic methods. The distribution and number of strain gauges should cover the load spectrum adequately for each part essential to the safe operation of the rotorcraft as identified in § 29.573(d)(1). Other devices such as accelerometers may be used as appropriate.

( 2 ) The corresponding flight parameters (airspeed, rotor RPM, center - of-gravity accelerations, etc.) should also be recorded simultaneously by appropriate Page C - 160 12/1/2011 AC 29-2C, Chg 4 methods. This is necessary to correlate the loads and stresses with the maneuver or operating conditions at which they occurred.

( 3 ) The instrumentation system should be adequately calibrated and checked periodically throughout the flight strain measurement program to ensure consistent and accurate results.

(C) Parts to be Strain-Gauged. Fatigue critical portions of the rotor systems, control systems, landing gear, fuselage, and supporting structure for rotors, transmissions, and engine are to be strain-gauged. For rotorcraft of unusual or unique design, special consideration might be necessary to ensure that all the essential parts are evaluated.

(D) Flight Regimes and Conditions to be investigated.

( 1 ) Typical flight and ground conditions to be investigated in the flight strain measurement program are given in paragraphs c. and d. of AC 27-1B, MG 11.

( 2 ) The determination of flight conditions to be investigated in the flight strain measurement program should be based on the anticipated use of the rotorcraft and, if available, on past service records for similar designs. In any event, the flight conditions considered appropriate for the design and application should be representative of the actual operation in accordance with the rotorcraft flight manual. In the case of multiengine rotorcraft, the flight conditions concerning partial engine-out operation should be considered in addition to complete power-off operation. The flight conditions to be investigated should be submitted in connection with the flight evaluation program.

( 3 ) The severity of the maneuvers investigated during the flight strain survey should be at least as severe as the maneuvers likely to occur in service.

( 4 ) All flight conditions considered appropriate for the particular design are to be investigated over the complete rotor speed, airspeed, center of gravity, altitude, and weight ranges to determine the most critical stress levels associated with each flight condition. The temperature effects on loads as affected by elastomeric components are to be investigated. To account for data scatter and to determine the stress levels present, a sufficient amount of data points should be obtained at each flight condition. Consideration can be given to the use of scatter factors in determining the sufficiency of data points. In some instances, the critical weight, center of gravity, and altitude ranges for the various maneuvers can be based on past experience with similar design. This procedure is acceptable where adequate flight tests are performed to substantiate such selections. The combinations of flight parameters that produce the most critical stress levels should be used in the evaluation.

(vi) Frequency of loading.

Page C - 161 AC 29-2C, Chg 4 12/1/2011 (A) Types of Operation.

( 1 ) The probable types of operation (transport, utility, etc.) for the rotorcraft should be established. The type of operation can have a major influence on the loading environment. In the past, rotorcrafts have been substantiated for the most critical general types of operation with some consideration of special, occasional types of operation. To assure that the most critical types of operation are considered, each major rotorcraft structural component should be substantiated for the most critical types of operation as established by the manufacturer. The types of operation shown below should be considered and, if applicable, used in the substantiation: (i) Long flights to remote sites (low ground-air-ground cycles but high cruising speeds).

( ii ) Typical, general types of operation.

( iii ) Short flights as used in logging operations.

( 2 ) One means is to substantiate for the most severe type of operation; however, this method is not always economically feasible.

( 3 ) A second means is to quantify the influence of mission type on fatigue damage by adding to or replacing hour limitations by flight cycle limitations (if properly defined and easily identifiable by the crew, for example: one landing, one load transportation). A special type of flight hour limitation replacement using factorization of flight hours for multiple types of operations may be feasible if continuing manufacturers’ technical support is provided and documented (i.e., the manufacturer either provides the factorization analyses or checks them on a continuing basis for each type of rotorcraft operation).

( 4 ) Where one or more operations are not among the general uses intended for the rotorcraft, the rotorcraft flight manual should state in the limitations section that the intended use of the rotorcraft does not include certain missions or repeated maneuvers (e.g., logging with its high number of takeoffs and landings per hour). A note to this effect should also appear in the rotorcraft airworthiness limitations section of the maintenance manual prepared in accordance with §§ 29.573 and 29.1529.

( 5 ) Should subsequent usage of the rotorcraft encompass a mission outside the original structural substantiation, the effects of this new mission environment on the frequency of loading and structural substantiation should be addressed and where practicable, in the interest of safety, a reassessment made. If this reassessment indicates the necessity for revised retirement times, those new times may be limited to specific rotorcraft model involved in the added mission provided: Page C - 162 12/1/2011 AC 29-2C, Chg 4 (i) changes are adopted through the airworthiness directives process and proper part re-identification is established; or ( ii ) a Rotorcraft Flight Manual (RFM) supplement outlining the limitations is approved; or ( iii ) an airworthiness limitations section (ALS) supplement is approved; or ( iv ) an appropriate combination of part re-identification, RFM supplement, or airworthiness limitation section supplement is approved.

(B) Loading Spectrum. The spectrum allocating percentage of time or frequencies of occurrence to flight conditions or maneuvers is to be based on the expected usage of the rotorcraft. This spectrum is to be established so that it is unlikely that actual usage will subject the structure to damage beyond that associated with the spectrum. Considerations to be included in developing this spectrum should include prior knowledge based on flight history recorder data, design limitations established in compliance with § 29.309, and recommended operating conditions and limitations specified in the rotorcraft flight manual or instructions for continued airworthiness (ICA).

The distribution of times at various forward flight speeds should reflect not only the relation of these speeds to V but also the recommended operating conditions in the NE rotorcraft flight manual or ICA that govern V or cruise speed. It is desirable to conduct c the flight strain-gauge program by simulating the usage as determined previously, with continuous recording of stresses and loads, thus obtaining directly the stress or load spectra for structural elements.

(7) The seventh major area is the dynamic loading and response requirements of §§ 29.241, 29.251, and 29.629 for vibration and resonance frequency determination and separation for aeroelastic stability and stability margin determination for dynamically critical flight structure. Critical parts, locations, excitation modes, and separations should be identified and substantiated. This substantiation should consist of analysis supported by tests, including tests that account for repeated loading effects and environment exposure effects on critical properties, such as stiffness, mass, and damping. This must be accomplished to assure that the initial stiffness, residual stiffness, proper critical frequency design, and structural damping are provided as necessary to prevent vibration, resonance, and flutter problems.

(i) All vibration and resonance critical composite structures must be identified and properly evaluated.

(ii) All flutter-critical composite structures must be identified and properly evaluated. This structure must be shown by analysis to be flutter free to 1.1 V (or any NE other critical operating limit, such as V , for a VSTOL aircraft) with the extent of damage D for which residual strength and stiffness are demonstrated.

Page C - 163 AC 29-2C, Chg 4 12/1/2011 (iii) Where appropriate, crash impact dynamics considerations should be taken into account to ensure proper crash resistance and a proper level of occupant safety for an otherwise survivable impact.

(8) The eighth area is the special repair and continued airworthiness requirements of §§ 29.611, 29.1529, and 14 CFR part 29 Appendix A, for composite structures. When repair and continued airworthiness procedures are provided in service documents (including approved sections of the maintenance manual or instructions for continued airworthiness), the resulting repairs and maintenance provisions should be shown to provide structure, which continually meets the guidance of paragraphs (1) through (7) of this AC. All certification-based repair and continued airworthiness standards, limits, and inspections must be clearly stated, and their provisions and limitations clearly documented to ensure continued airworthiness. No composite structural repair should be attempted that is beyond the scope of the applicable approved Structural Repair Manual (SRM) without an engineering design approval by a qualified FAA/AUTHORITY designated representative.

Page C - 164 9/30/99 AC 29-2C CHAPTER 2. PART 29 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT SUBPART D - DESIGN AND CONSTRUCTION DESIGN AND CONSTRUCTION - GENERAL AC 29.601. § 29.601 DESIGN .

a. Explanation .

(1) This rule requires that no design features or details be used that experience has shown to be hazardous or unreliable.

(2) Further, the rule requires that the suitability of each questionable design detail and part must be established by tests.

b. Procedures .

(1) This rule is met partially by a review of service history of earlier model rotorcraft, or for a new model, review of service experience of models with similar design features. Specifically, this rule covers “features or details” such as the following: (i) Seat track-to-seat interface fittings should have adequate locking devices to prevent both premature structural failure and premature unlatching.

(ii) Seat belt and harness should be of a type and construction that service experience has shown to be easy to don, unlatch, and remove. They should also be of a type that is reliable, does not interfere with egress, and does not sustain unnecessary wear and tear under normal operations.

(iii) Metallic parts less than a certain thickness gauge and composite materials less than a certain number of plies should not be used. The minimum thickness and number of plies should be based to a large degree on service experience (normal wear and tear) with similar designs.

(2) The effects of service wear on the loading of critical components should be considered. Flight testing, ground testing, and analyses may be used in these considerations.

(3) Tests are required for details and parts which the applicant chooses to use after questions have arisen concerning their suitability.

Page D - 1 AC 29-2C, Chg 2 4/25/06 AC 29.602 § 29.602 CRITICAL PARTS .

a. Explanation .

(1) Critical parts requirements apply to structural components, rotor drive systems, rotors, and mechanical control systems.

(2) The objective of identifying critical parts is to ensure that critical parts are controlled during design, manufacture, and throughout their service life so that the risk of failure in service is minimized by ensuring that the critical parts maintain the critical characteristics on which certification is based.

(3) Definitions with respect to § 29.602: (i) The use of the word “could” in paragraph 29.602(a) of the rule means that this failure assessment should consider the effect of flight regime (i.e., forward flight, hover, etc.). The operational environment need not be considered.

(ii) With respect to this rule, the term “catastrophic” means the inability to conduct an autorotation to a safe landing, without exceptional piloting skills, assuming a suitable landing surface is available.

(iii) The use of the word “and” in paragraph 29.602(a) of the rule means the part must have both a catastrophic failure mode together with one or more critical characteristics.

(iv) With respect to this rule, the term “part” means one piece, or two or more pieces permanently joined together.

(v) With respect to this rule, the term “critical characteristic” means any dimension, tolerance, finish, material, or any manufacturing or inspection process, or other feature which cannot tolerate variation from type design requirements and, if nonconforming, would cause failure of the critical part.

(4) Many rotorcraft manufacturers already have procedures in place within their companies for handling “critical parts.” These plans may be required by their dealings with other customers, frequently military (e.g., US DoD, UK MoD, Italian MoD).

Although these plans may have slightly different definitions of “critical parts” which have sometimes been called “Flight Safety Parts,” “Critical Parts,” “Vital Parts,” or “Identifiable Parts,” they have in the past been accepted as meeting the intent of this requirement and providing the expected level of safety. It is acceptable for these plans to use alternative names and terminology provided they meet the intent of this requirement.

b. Procedures . The rotorcraft manufacturer should establish a Critical Parts Plan that identifies and controls the critical characteristics. The policies and procedures which constitute that plan should be such as to ensure that- - Page D - 2 4/25/06 AC 29-2C, Chg 2 (1) All critical parts of the rotorcraft are identified by means of an appropriate failure assessment and a Critical Parts List is established.

(2) Documentation draws the attention of the personnel involved in the design, manufacture, maintenance, inspection, and overhaul of a critical part to the special nature of the part and details the relevant special instructions. For example all drawings, work sheets, inspection documents, etc., could be prominently annotated with the words "Critical Part" or equivalent and the Instructions for Continued Airworthiness and Overhaul Manuals (if applicable) should clearly identify critical parts and include the needed maintenance and overhaul instructions. The documentation should: (i) Contain comprehensive instructions for the maintenance, inspection and overhaul of critical parts and emphasize the importance of these special procedures; (ii) Indicate to operators and overhaulers that unauthorized repairs or modifications to critical parts may have hazardous consequences; (iii) Emphasize the need for careful handling and protection against damage or corrosion during maintenance, overhaul, storage, and transportation and accurate recording and control of service life (if applicable); (iv) Require notification of the manufacturer of any unusual wear or deterioration of critical parts and the return of affected parts for investigation when appropriate; (3) Procedures should be established for identifying and controlling critical characteristics.

(4) To the extent needed for control of critical characteristics, procedures and processes for manufacturing critical parts (including test articles) are defined (for example material source, forging procedures, machining operations and sequence, inspection techniques, and acceptance and rejection criteria). Procedures for changing these manufacturing procedures should also be established.

(5) Any changes to the manufacturing procedures, to the design of a critical part, to the approved operating environment, or to the design loading spectrum are evaluated to establish the effects, if any, on the fatigue evaluation of the part.

(6) Materials review procedures for critical parts (i.e., procedures for determining the disposition of parts having manufacturing errors or material flaws) are in accordance with paragraphs (4) and (5) above.

(7) Critical parts are identified as required, and relevant records relating to the identification are maintained such that it is possible to establish the manufacturing history of the individual parts or batches of parts.

Page D - 3 AC 29-2C, Chg 4 5/1/2014 (8) The critical characteristics of critical parts produced in whole or in part by suppliers are maintained.

AC 29.603. § 29.603 (Amendment 29-17) MATERIALS.

a. Explanation . The rule requires that the suitability and durability of materials, the failure of which could adversely affect safety, must be determined by three-fold considerations: (1) Considerations based on experience or tests.

(2) Meeting approved specifications.

(3) Taking into account environmental conditions such as temperature and humidity.

b. Procedures .

(1) Experience may be used to show a material’s resistance to wear and deterioration from environmental effects (environmental effects include both natural environmental effects such as exposure to sunlight, water, salt spray, etc., and installation environmental effects such as exposure to fuel, hydraulic fluids, deicing fluids, etc.). Installation environmental effects should consider both direct exposure contact and expected migration of potentially deleterious fluids and compounds.

Testing for environmental effects may use either coupon testing, full-scale testing, or a combination. A combination of testing and experience may also be used.

(i) MIL-HDBK’s-5, -17, and -23 include consideration of some environmental effects and contain reference to additional methods of testing for environmental effects.

(ii) The use of AC 20-107A, Composite Aircraft Structure, is recommended for environmental and damage tolerance considerations of advanced composite materials. (Also see sections 29.573 and AC 29 MG 8 of this AC.)

(iii) The effects of excessive wear and delamination of elastomeric and self-lubricated bearings used in critical load carrying applications in relation to redistribution of loading should be considered.

(2) Where possible, materials that meet widely accepted specifications such as AISI, SAE, MIL, or AMS and alloys which have favorable experience or tests should be used. Where company-developed materials are used, approved specifications are required to ensure the developed properties are duplicated in each lot of material.

Documented specification usage is necessary to maintain quality assurance of materials.

Page D -4 9/30/99 AC 29-2C (3) Section 29.613 concerns strength properties and design values. (See paragraph AC 29.613.)

AC 29.605. § 29.605 ( Amendment 29 -17) FABRICATION METHODS .

a. Explanation . The basic requirement of this rule is that the methods of fabrication must produce sound structure and produce it consistently.

(1) A process specification is required for fabrication processes requiring close control .

(2) A test program is explicitly required for each new aircraft fabrication method.

b. Procedures .

(1) The approved specifications required by this rule may either be established government/industry specifications such as MIL, AISI, ASTM, or SAE, or the specifications may be company-developed proprietary specifications. Sufficient data should be provided to the FAA/AUTHORITY aircraft engineering offices to show that the desired features are provided by the process specification. In addition, sufficient process controls, inspections, and tests should be coordinated with FAA/AUTHORITY manufacturing inspection personnel to ensure that continued quality of the process is provided.

(2) In addition to the examples given by the rule; i.e., gluing, spot welding, and heat treating process, specifications should also be prepared for types of welding other than spot welding, for platings of metals, for protective finishes (other than decorative), for sealing, and for unique fabrication methods such as those used for composite materials.

(3) The required test programs should consider static strength effects, fatigue strength effects, and environmental effects as appropriate to the processes.

(4) During the fabrication of advanced composite materials, the effects of fabrication anomalies (i.e., disbonds, voids, porosity) should be considered. Special nondestruct testing inspection techniques and procedures should be developed to cover fabrication with allowable anomalies and permitted repair procedures. (See also paragraph AC 29 MG 8.)

AC 29.607. § 29.607 (Amendment 29-5) FASTENERS .

a. Explanation . Section 29.607 of Amendment 29-5 requires dual locking removable fasteners in critical locations. A nonfriction locking device is specifically required in any bolt subject to rotation, as stated in the rules.

Page D - 5 AC 29-2C 9/30/99 b. Procedures . Advisory Circular 20-71, Dual Locking Devices or Fasteners, December 8, 1970, contains information, procedures, and means of complying with § 29.607 of Amendment 29-5.

AC 29.609. § 29.609 PROTECTION OF STRUCTURE .

a. Explanation . The structure should be suitably protected as specified in the rule to maintain its design strength. Ventilation and drainage provisions must be provided as specified in the rule. Overboard drains should be furnished for corrosive or waste liquids. Drains for flammable fluids are specified in other rules such as §§ 29.999 and 29.1187.

b. Procedures .

(1) The structure may be preserved, painted, or treated with chemical films to protect it from strength deterioration. An approved process specification should be used for these types of treatments.

(2) Parts may be plated or chemically treated, such as anodized, for protection.

An evaluation and substantiation may be required to assure the structure or parts are not adversely affected during, or as a result of, the plating or treatment process.

(§ 29.605 concerns approval of process specifications and fabrication methods.)

(3) Plating or material surface hardness or composition changes may require fatigue substantiation to assure the fatigue strength is not altered or is otherwise properly assessed. An approved process specification should be used for these types of treatments.

(4) To prevent water accumulation, drain holes should be placed at possible dams such as bulkheads, and at low points in the fuselage and in the stabilizing surfaces.

(5) Control tubes and tubes used as primary mount structures (i.e., transmission support structure and engine mount structure) should be designed to prevent entry and collection of corrosive fluids or vapor, including water.

(i) A closed insert in each tube end may be used.

(ii) A sealant applied around the tube ends and around each rivet head may be used.

(6) Overboard drains should discharge clear of the entire rotorcraft. Dyed water discharged in flight, may be used to assure fluids are properly drained.

Page D - 6 5/1/2014 AC 29-2C, Chg 4 (7) Welded tubes should be flushed and sealed after welding in accordance with an approved process specification.

(8) Refer to AC 43-4, “Corrosion Control for Aircraft,“ for further procedures.

Page D - 7 AC 29-2C, Chg 4 5/1/2014 AC 29.610. § 29.610 (Amendment 29-53) LIGHTNING AND STATIC ELECTRICITY PROTECTION.

a. Background . During the initial development and promulgation of airworthiness standards, rotorcraft operated primarily in a VFR and non-icing environment. A prudent pilot avoided thunderstorms where the possibility of encountering severe weather and a lightning strike was much greater. Now, many rotorcraft are authorized to fly under IFR and into known icing conditions. Because many rotorcraft now use the same advanced technologies in structures and systems as fixed-wing aircraft, a specific rule on lightning protection of rotorcraft was adopted in Amendment 29-24. Amendment 29-40 revised the title of § 29.610 to include protection from static electricity, and a new § 29.610(d) was added specifying requirements for electrical bonding and protection against lightning and static electricity. Amendment 29-53 removed the system lightning protection requirements for electrical and electronic equipment in § 29.610(d)(4) because those requirements were implemented in the new § 29.1316.

b. Explanation .

(1) The regulation requires protection of rotorcraft from the catastrophic effects of lightning. This means that lightning must not prevent the continued safe flight and landing of the rotorcraft.

(2) Rotorcraft structural components, propulsion system, gearboxes, and mechanical and hydraulic control systems should be designed to ensure lightning will not prevent continued safe flight and landing of the rotorcraft (i.e., damage due to lightning currents that flow through any of the components must not result in catastrophic failure).

c. Procedures .

(1) Certification Plan . A formal written certification plan is an acceptable means to ensure demonstration of regulatory compliance. This plan is also useful to identify and define an acceptable resolution to the critical issues early in the certification process. These are the usual steps to follow when utilizing a certification plan: (i) Prepare a certification plan that describes the analytical procedures and qualification tests to be utilized to demonstrate lightning and static electricity protection effectiveness. Test proposals should describe the rotorcraft and system to be utilized, test drawing(s) as required, the method of installation that simulates the production installation, the lightning zone(s) applicable, the lightning simulation method(s), test voltage or current waveforms to be used, diagnostic methods, and the appropriate schedules and location(s) of proposed test(s).

(ii) Obtain FAA/AUTHORITY concurrence that the certification plan is adequate.

Page D - 8 5/1/2014 AC 29-2C, Chg 4 (iii) Obtain FAA/AUTHORITY detail part conformity of the test articles and installation conformity of applicable portions of the test setup. Obtain FAA/AUTHORITY approval of the test proposal. A comprehensive test proposal may be used.

(iv) Schedule FAA/AUTHORITY witnessing of the test or tests proposed.

(v) Submit a test report describing all results and obtain FAA/AUTHORITY approval of each report prepared.

(2) Lightning Environment and Zones . AC 20-155 refers to SAE documents ARP5412 (or EUROCAE ED-84) and ARP5414 (or EUROCAE ED-91), which provide acceptable definitions for the rotorcraft lightning environment and for rotorcraft lightning attachment zones.

(3) Testing . Tests may be required to check the adequacy of the lightning and electrostatic charge protection. Refer to SAE ARP5416 (or EUROCAE ED-105) for acceptable test methods to show that the lightning protection is effective and to SAE ARP5672 for acceptable test methods to show rotorcraft electrostatic charge control.

(4) Aircraft Lightning Protection Design Features . The Aircraft Lightning Protection Handbook (DOT/FAA/CT-89/22) provides information on aircraft lightning protection design. The following are examples where lightning protection should be considered.

(i) Rotors and Control Systems.

(A) It should be established that an adequate bonding path exists between the rotors and the airframe, such that a lightning strike to a rotor will not result in damage to or seizure of gearbox or swashplate bearings, control jacks, etc.

(B) Each hinge and bearing of rotor blades and control surfaces should either: ( 1 ) be capable of withstanding lightning without damage or seizure leading to loss of function, or ( 2 ) be provided with at least one bonding conductor, as flexible and short as possible and installed so that there is no danger of the conductor jamming the hinge or bearing following a lightning strike.

(ii) External Non-Metallic Parts.

Page D - 9 AC 29-2C, Chg 4 5/1/2014 (A) Where non-metallic parts are fitted externally to the rotorcraft (e.g., rotors, radomes, composite skin panels) and may be subjected to lightning, they should be protected. The protection should consider disruption of the materials because of: (1) rapid expansion of gases within them (e.g., water vapor), (2) rapid build-up of pressure in voids or in the enclosure provided by the parts resulting in mechanical disruption of the parts themselves or of the structure enclosed by them, and (3) fire caused by the ignition of the materials themselves or of the materials contained within the enclosures.

(B) Materials used for external non-metallic parts should have low water absorption characteristics, should not occlude gases, and should be of high dielectric strength in order to encourage surface flashover rather than puncture.

(C) Rotors and other external parts of nonmetallic construction should be provided with effective lightning conductors that are capable of safely carrying lightning current, unless it can be shown that damage due to lightning will not endanger the rotorcraft or its occupants. Bonding straps and leads are not required for small gaps between metallic structure and diverters in non-conducting panels in order to comply with the lightning protection criteria. However, an electrical bonding path may be required to achieve static electricity protection.

(5) Protection Against the Effects Static Electricity .

(i) General. Rotorcraft structure, rotor systems, and equipment should be electrically bonded together to minimize the accumulation and discharge of electrostatic charge, which could result in electrical shock, ignition of flammable vapors, or interference with essential equipment such as radio communications and navigational aids.

(ii) Intermittent Contact. The design should prevent random intermittent contact between metallic or metallized parts such that unwanted radio interference or degradation of the components due to sparking will not occur.

(iii) Rotors and other external parts of nonmetallic construction should be provided with effective electrical conductors that are capable of safely conducting electrostatic charge.

(iv) High Pressure Refueling and Fuel Transfer. Where provision is made for high pressure refueling or high rates of fuel transfer, it should be established, by test, or by consultation with the appropriate fuel manufacturers, that dangerously high voltages Page D - 10 5/1/2014 AC 29-2C, Chg 4 will not be induced within the fuel system. If compliance with this requirement involves any restriction on the types of fuel to be used or on the use of additives, an appropriate operating limitation should be established under § 29.1501(a). The critical refueling rates are related to the rotorcraft refueling installations, and the designer should seek the advice of fuel suppliers on this problem.

(A) With standard refueling equipment and standard aircraft turbine fuels, voltages high enough to cause sparking may be induced between the surface of the fuel and metal parts of the tank at refueling rates above approximately 250 gal/min. These induced voltages may be increased by the presence of additives and contaminants (e.g., anti-corrosion inhibitors, lubricating oil, free water) and by splashing or spraying of the fuel in the tank.

(B) The static charge can be reduced as follows: ( 1 ) by means taken in the refueling equipment such as increasing the diameter of refueling lines and designing filters to give the minimum of electrostatic charging, or ( 2 ) by changing the electrical properties of the fuel by the use of anti- static additives and thus reducing the accumulation of static charge in the tank to a negligible amount.

(6) Fuel Systems . Requirements for lightning protection for fuel systems are in § 29.954. AC 20-53B provides guidance on compliance with § 29.954. Section 29.954 of this AC addresses the protection required for systems.

Page D - 11 AC 29-2C, Chg 4 5/1/2014 AC 29.611. § 29.611 INSPECTION PROVISIONS.

a. Explanation . The rotorcraft must have access panels, or openings, that will allow for proper maintenance and/or adjustment of the rotorcraft systems.

(1) The rule states: There must be means to allow close examination of each part that requires recurring inspection, adjustment for proper alignment and functioning, or lubrication.

(2) “Structural” or load-carrying access panels may be used to comply with the rule. Structural panels should have stencils or permanent labels (§ 29.1541(a)(2)) stating the panels must be installed prior to ground or flight operation.

(3) Holes or “nonstructural” access panels should be used whenever possible.

b. Procedures .

(1) The determination of compliance can be accomplished in conjunction with the following activities: (i) Reviewing type design drawings.

(ii) Conformity inspections accomplished during certification testing.

(iii) Be evaluated during the control system proof and operation tests (§§ 29.681 and 29.683).

(iv) During type inspection tests and functioning and reliability testing.

(2) Equipment requiring frequent inspections (at less than 25-hour intervals), lubrication, or adjustments should be accessible through “nonstructural” doors. Areas or items requiring daily attention should be accessible through “nonstructural” doors since properly rated maintenance personnel are required to “open and close,” or reinstall structural panels and special design features, such as multiple pins and latches, are generally necessary for structural doors.

(3) If the rotorcraft is subject to an FAA Maintenance Review Board Approval Program, further review of the rotorcraft inspection provisions will be obtained.

Page D - 12 5/1/2014 AC 29-2C, Chg 4 AC 29.613. § 29.613 (Amendment 29-17) MATERIAL STRENGTH PROPERTIES AND DESIGN VALUES.

a. Explanation . The rule requires the use of materials that have a known minimum strength value. The structure must not be understrength and must be designed to minimize fatigue failure.

(1) Material design values in certain specified documents may be used. The FAA/AUTHORITY may approve other material design values thus allowing the applicant greater flexibility in selection of materials by proving their strength properties and design values as stated in § 29.613(d).

(2) Other materials that may be new or are not included in the specified documents may be tested and design values established as provided by § 29.613(a) and (d).

(3) Section 29.613(d) requires the selection of materials that will retain design values and properties in the type of service environment and for the length of service time intended for the structure.

(4) Section 29.613(c) is an objective rule concerning minimizing fatigue failures.

Paragraph AC 29.571, § 29.571, concerns quantitative fatigue substantiation requirements.

b. Procedures .

(1) The properties and design values in the documents noted in the rule may be used.

(2) MIL-HDBK-5, Metallic Materials and Elements for Flight Vehicle Structure, Chapter 9, contains procedures for establishing design values of additional materials.

Uniform means of presenting the data is also contained in this chapter.

(3) Design values and properties must include effects of the service environment and service time. An example is exposure at elevated temperatures on the ultimate tensile strength of 7079-T6 aluminum alloys as found in figure 3.7.4.1.1(c) of MIL-HDBK-5C.

(4) The probability of disastrous fatigue failures must be minimized. This may be accomplished by using design features usually identified as fail-safe features, such as the following. (See paragraph AC 29.571 for the fatigue requirement information.)

(i) Selection of materials and stress levels that provide a controlled slow rate of crack propagation combined with high residual strength after initiation of cracks (lightly loaded structures).

Page D - 13 AC 29-2C, Chg 4 5/1/2014 (ii) Use of multipath construction and the provision of crack stoppers to limit the growth of cracks.

(iii) Use of composite (multielement) duplicate structures so that a fatigue crack or failure occurring in one element of the composite (multielement) member will be confined to that element and the remaining structure will still possess adequate load-carrying ability.

(iv) Use of backup structure wherein one member carries all the load, with a second member available and capable of assuming the extra load if the primary member fails.

(v) Design to permit detection of cracks including the use of crack detection systems, in all critical structural elements before the cracks can become dangerous or result in appreciable strength loss, and to permit replacement or repair.

(5) Acceptable standards for pressurized containers or cylinders, such as cylinders of nitrogen, used to inflate emergency floats may be found in 49 CFR 178 Subpart C, §§ 178.36 through 178.68. Specifically, § 178.44 concerns standards for steel cylinders used in aircraft that are subjected to at least 900 PSI service pressure.

This standard includes strength, test, material property, inspection, quality, design features, identification and inspection report requirements. As an example, § 178.44-14, entitled “Hydrostatic Test,” requires that each cylinder must be (proof) tested to at least 5/3 times the service pressure. Section 178.44-16, entitled “Burst Test,” also states that one cylinder taken at random out of each lot of cylinders shall be hydrostatically tested to destruction.

(6) Other design criteria may be developed and approved under the provisions of FAR Part 29 as a unique part of the aircraft type design.

AC 29.613A. § 29.613 (Amendment 29-30) MATERIAL STRENGTH PROPERTIES AND DESIGN VALUES.

a. Explanation . Amendment 29-30 added explicit probability standards criteria to § 29.613(b). This amendment also provided for testing or proving the strength of selected individual items rather than conducting coupon tests to develop generic material strength properties that would be used for design purposes.

b. Procedures . The basic procedures of paragraph AC 29.613 still apply, except: (1) Probability criteria common with MIL-HDBK-5D are explicitly allowed to determine strengths for metallic materials whose data are not available in MIL-HDBK-5D. These specific probability criteria should be used in conjunction with MIL-HDBK-17B whenever determining material strength properties for non-metallics.

(Also, reference paragraph AC 29 MG 8.)

Page D - 14 5/1/2014 AC 29-2C, Chg 4 (2) New § 29.613(e) provides for the premium selection of materials. The premium selection of materials method uses a specimen from each individual item (part) to determine its properties before its use is allowed. This is a highly specialized and possibly costly method which applies only to parts that have areas available from which specimens can be obtained without destroying the part. The rotorcraft type design data of those parts made from premium selection should have the necessary information, such as minimum allowable strength, on the part drawing.

Page D - 15 AC 29-2C, Chg 4 5/1/2014 [This page intentionally blank.]

Page D - 16 9/30/99 AC 29-2C AC 29.619. § 29.619 SPECIAL FACTORS .

a. Explanation .

(1) This is a general rule to complement other rules. Special factors are employed for reasons cited in the rule to ensure an airworthy aircraft structure. The 1.5 ultimate load factor in § 29.303 is multiplied by a special factor as specified in the rule.

(2) Specific factors are prescribed for castings and fittings in §§ 29.621 and 29.625 respectively. Factors may be prescribed for bearings with free clearance as stated in § 29.623. In addition, any other factor may be prescribed “to ensure that the probability of the part being understrength because of the uncertainties specified in § 29.619(a) “is extremely remote.” b. Procedures .

(1) One example of fitting factor use follows: 1,000 pounds limit design load x 1.15 fitting factor x 1.5 ultimate load factor equals 1,725 pounds ultimate design load.

(2) Other specific factors may be similarly applied. Refer to §§ 29.623, 29.625, 29.685, and 29.785.

(3) Other factors may be imposed as cited in the rule. Advisory Circular 20-107, paragraphs 5 and 6, are examples of requiring tests of component and subcomponent structure to account for variability of strength and stiffness of composite structures. Factors appropriate for the particular design are obtained and used in substantiation of the composite structure.

(4) The rule complements §§ 29.603 and 29.613. Regardless of the rule invoked, the variability of the material and/or assembly properties must be accounted for.

(5) Ground resonance can occur due to flexibility in the rotor pylon restraint system as well as with landing gear flexibilities. This evaluation should include variations in stiffness and damping of the rotor pylon restraints that may occur in service (reference “Ground Vibrations of Helicopters,” M.L. Deutsch, JAS, Vol. 13, No. 5, May 1946).

AC 29.621. § 29.621 CASTING FACTORS .

a. Explanation . Casting design, test, and inspection criteria are included in this rule for critical and noncritical structural castings. Hydraulic or other fluid containers are Page D - 17 AC 29-2C 9/30/99 not subjected to “structural loads” but are subject to pressure testing as a part of hydraulic or other flight systems. Critical and noncritical castings are defined in the rule.

(1) Factors, tests, and inspections are specified for structural castings.

Additional factors, tests, and inspections may be applied, as prescribed by §§ 29.603, 29.605, or 29.613, for foundry quality control.

(2) For castings that have surfaces subject to bearing structural design loads, the casting factor need not exceed 1.25 with respect to bearing stresses and need not be used with respect to the bearing surfaces if the bearing factor of § 29.623 exceeds the applicable casting factor.

(3) Critical castings must have a casting factor not less than 1.25 and must receive 100 percent inspection as specified including radiographic inspection. Static test requirements are also specified in addition to the inspection requirements.

(4) Noncritical structural castings may have a casting factor as small as 1.0 with attendant increased inspection and quality control requirements. Use of larger casting factors reduces the inspection and quality control requirements.

(5) Structural static and fatigue substantiation, by test or analysis, are still required in addition to any casting static tests required by this rule.

b. Procedures .

(1) The rotorcraft castings should be classified as critical, or noncritical, or nonstructural, or fluid container as soon as possible in the certification program. The applicant should then be prepared to propose the tests required for certification.

(2) The casting factors and associated inspection requirements dictated by § 29.621(c) and (d) are shown below: Page D - 18 9/30/99 AC 29-2C INSPECTION REQUIREMENTS CRITICAL CASTINGS NONCRITICAL CASTINGS <(2)> <(3)> CASTING FACTOR FAA REQUIRE - OTHER FAA REQUIRE - OTHER RANGE MENT 29.621(c) CLASSIFICATION MENT 29.621(d) CLASSIFICATION <(1)> 2.01 OR <(7)> <(4)> GREATER 1.50 TO <(7)> <(5)> 2.00 1.250 TO <(7)> <(8)> <(6)> 1.499 1.00 TO NOT NOT <(7)> <(8)> 1.249 ALLOWED ALLOWED <(9)> <(1)> Ultimate load = Casting factor x 1.5 x limit load. CAUTION: For casting factor range of 1.25 to 1.5 see yield test requirements of NOTE <(8)>. The mechanical properties to be used for analysis shall be based on the tabulated values of MIL-HDBK-5 or other approved sources, reference § 29.613.

<(2)> Critical castings are those castings whose failure would preclude continued safe flight and landing or result in injury to any occupant, reference § 29.621(c).

<(3)> Noncritical castings are castings other than those defined by NOTE <(2)>.

<(4)> Each casting shall receive 100 percent visual inspection.

<(5)> Each casting shall receive 100 percent visual and reduced magnetic particle or penetrant inspection or approved equivalent methods.

<(6)> Each casting shall receive 100 percent visual and reduced Page D - 19 AC 29-2C 9/30/99 radiographic and magnetic particle or penetrant inspection, or approved equivalent methods.

<(7)> Each casting shall receive 100 percent inspection by visual, radiographic and magnetic particle or penetrant inspections or approved equivalent methods.

<(8)> Three sample castings shall be static tested and shown to meet: No failure at 1.25 x 1.5 x limit load, and No yielding at 1.15 x limit load.

<(9)> Castings shall be procured to a specification that guarantees the mechanical properties of the material in the casting and provides demonstration of these properties by test of coupons cut from the castings on a sampling basis.

This chart may be included in the casting test proposal report. It is recommended that the applicant include in the test proposal report additional information such as shown in paragraph AC 29.621b(3).

(3) The casting test report may include the following sections or items in a Part I of the report. The report may also have a Part II that contains the test results as shown in the following example report. The following sections are a recommended format content of the report. Appropriate changes should be made as desired to accommodate the applicant’s system.

EXAMPLE OF REPORT INTRODUCTION This report presents the proposal for the static test of the castings used on the Model XYZ. The castings will be tested in compliance with Federal Aviation Regulations, Part 29, § 29.621. The purpose of this test is to substantiate the structural strength of the castings used on the Model XYZ. Part II of this report, which will be published after static tests have been completed, will present test results.

All test specimens will be selected as radiographic standards of acceptance for the particular castings (see Test Specimen). Additional information on selecting the specific castings may be included in the test specimen section of this report.

Load sheets giving direction and magnitude of loads for each of the castings are presented in numerical order by part number at the end of this report. The test loads and design criteria for the castings are discussed in detail in the test loads section of this report.

Page D - 20 9/30/99 AC 29-2C The test loads will be applied and reacted using mating aircraft parts or special fixtures which simulate the mating parts. The methods and apparatus to be used for the static tests of the castings are discussed in the apparatus and method section of this report.

Testing will be conducted in…(location).

TEST SPECIMEN The castings which will be tested are listed in numerical order in figure AC 29.621-2.

Those castings which, after structural analysis, show less than a 1.5 casting factor will be tested. All directions are given with reference to a forward facing position in the rotorcraft.

On the basis of a radiographic examination, the three castings which are of the poorest acceptable quality in the first production lot of castings will be selected as test specimens. The poorest of the three castings will be selected as the initial test casting and its radiograph or ASTM standard will be used as the standard for accepting future castings of the particular part unless later standards are approved. Three castings must be tested for each critical condition for each part.

Conformity Inspection Each machined casting will be subjected to an FAA/AUTHORITY conformity inspection prior to testing to determine compliance with the type design drawings. A conformity report for each casting may be incorporated in Part II, Test Results, of this report.

The test specimen will be permanently marked or defaced after testing to preclude its use on a rotorcraft.

See figure AC 29.621-2 for an example of a convenient means of listing castings.

TEST LOAD The test load(s) to be applied to each casting represents the critical loading condition(s) for that casting. The critical conditions on each of the castings were determined by the design criteria and substantiating data approved by the FAA/AUTHORITY.

The design criteria for all of the castings to be static tested may fall into one of two categories. The load factors and structural acceptability requirements for each category are discussed below. Casting factors that are included on the load sheets of each part do not apply in the discussion below. (See paragraph AC 29.621b(2) for casting factors.)

Page D - 21 AC 29-2C 9/30/99 Castings Designed to Limit Load Conditions A structural analysis of each test casting showing the critical design limit load conditions is given in the data, (reference report number here). The load factors for the static test of the castings are as follows: 1.15 x design limit load = design yield load 1.50 x design limit load = design ultimate load Castings Designed Only to Crash Landing Conditions The castings in this category were designed using a crash landing load factor for the design ultimate load. The design yield load criteria of 1.15 x limit load need not apply to these castings. The test loads for these castings may be given in terms of design ultimate load on the individual casting load sheets shown in Part I of this report.

Test Procedures Depending on the results of the initial static test of each casting, the following procedure will be used.

a. If in the initial test of critical castings the casting is found to have a casting factor of 1.5 (1.5 x design ultimate load), the casting will be considered acceptable and no further tests will be conducted.

b. If in the initial test(s) the critical casting is found to have a casting factor less than 1.5 but equal to or greater than 1.25, two additional castings will be tested for each critical load condition. Each must also show a minimum casting factor of 1.25.

c. If in the initial test, or in one of two additional tests, a casting shows a casting factor less than 1.25 times design ultimate or yields prior to reaching 1.15 times design limit load, the casting will be redesigned and retested. The yield criteria are also applicable to the first two procedures with the exception of critical castings designed to crash landing conditions.

TEST APPARATUS AND METHOD The Model XYZ casting static tests will be conducted using fixtures designed to simulate the installation of the castings in the aircraft. Where practical, mating aircraft parts will be used to apply and react test loads. When practical, the static tests will be conducted with mating castings assembled when the critical loads for the mating castings are compatible; otherwise, fixtures simulating the mating parts may be designed and fabricated for the tests. Assembly hardware used to mount test castings will be the same as hardware used on the rotorcraft. All bolt torques and other assembly notes will conform to the type design assembly instructions.

Page D - 22 9/30/99 AC 29-2C The tests will be conducted using calibrated load measuring devices such as hydraulic cylinders and pressure gages, load cells, strain gage bridges, or dead weights.

Deflections of the casting may be measured using graduated dial indicators or scales in all tests. The deflection indicators will be based or mounted on the casting and will measure casting deflection only, when possible, otherwise the indicators will be based on the fixture and measure deflection of the casting relative to the fixture. Deflection readings will be made at 20 percent increments of limit load through 100 percent of limit load and at 115 percent of limit load. These increments may be changed if necessary.

Permanent deformation readings will be made after relieving 115 percent and 150 percent of limit load.

See figure AC 29.621-1 as an example of a load sheet.

Page D - 23 AC 29-2C 9/30/99 p

+5,000 # TENS. i

ULTIMATE LOADS -1,000 # COMP

- -

r-1--,

--~

-- -- --

/ ~

-

--

-

Tension Comp

-

.....

-

.

Tension Reaction Compression Reaction FIGURE AC 29.621-1 EXAMPLE OF CASTING LOAD SHEET RETRACT ACTUATOR SUPPORT - LANDING GEAR Page D - 24 9/30/99 AC 29-2C Include spherical bearing with clamped-up bolt and a link in the test setup to confirm the stability. Loads are based on a jam condition with actuator operating at 1,700 PSI pressure maximum.

A 1.25 casting factor is included in these loads.

These loads were derived from data in approved structural loads and analysis report.

END OF SAMPLE REPORT (4) The format of the previous guidance material may be changed to accommodate the applicant’s method of data presentation.

(5) Nonstructural castings may be tested and included in the test report.

(6) Cast fluid containers, including hydraulic fluid containers, may be tested as prescribed in other rules of FAR Part 29 and a test proposal and test results report may be included in the casting test report or an appropriate report may be referenced for convenience. We recommend use of one report to contain test data or reference to test data for all castings used on the rotorcraft.

Page D - 25 AC 29-2C 9/30/99 FIGURE AC 29.621-2 EXAMPLE CASTINGS TO BE STATIC TESTED FOR MODEL XYZ REF. LOAD MACHINE OR SHEET CASTING NO. ASSY. NO. NAME AND LOCATION MATERIAL FIG. NO.

Base Assembly, Pilot’s Collective Column Page D - 26 9/30/99 AC 29-2C AC 29.623. § 29.623 BEARING FACTORS.

a. Explanation .

(1) The rule requires use of a minimum bearing factor in free fit joints to account for effects of typical relative motion. A minimum value is not specified in the rule. The factor, appropriate for the application, is applied to the ultimate bearing strength of the softest material used as a bearing. A definition of free fit (clearance fit) is noted in Subparagraph b(7) below.

(2) Specific bearing factors are specified by § 29.685(e) for control system joints subject to angular rotation. These factors are applied to the ultimate bearing strength of the softest material used as a bearing in the control system. Control systems ball, roller, or needle bearings are covered by § 29.685(f).

(3) MIL-HDBK-5C, paragraph 8.3, refers to design standards for plain or journal bearings or bushings. These standards are found in Air Force Systems Command Design Handbook 2-1, Airframe, Chapters 2 and 6.

b. Procedures .

(1) Control system joint bearings are discussed under paragraph AC 29.685, § 29.685 of this document but the bearing factors are noted here for convenience.

Section 29.685(e) requires a 2.0 bearing factor for cable systems and a 3.33 bearing factor for push-pull systems other than ball and roller bearing systems. The manufacturer’s static, non-Brinell rating of ball and roller bearings may not be exceeded as stated in § 29.685(f).

(2) A landing gear pivot, grease lubricated, plain bearing is one example of a free fit subject to pounding or vibration. A bearing factor of 2.0 may be used or another factor may be proven for grease lubricated plain bearing or bushing to account for the anticipated higher loads caused by pounding or vibration. See subparagraph AC 29.623b(6) for ball or roller bearings.

(3) A typical engine mount bolt installation with a plain bearing having a free or loose fit (not interference fit), is another example of a sleeve bearing application subject to a design bearing factor. As an example, a bearing factor of 1.85 may be applied to the design loads on the softest material reacting the bearing loads. A different factor will be acceptable if proven. For example, the design limit load may be calculated for a .312-inch-diameter bolt in a 2-inch-long bearing. The bearing projected area is .312 x 2 = .624-inch-square. The design limit load is 3,000 pounds. The design limit bearing stress is 3,000 pounds/.624-inch-square x 1.85 = 8,894 PSI. If a free or loose fit is not used; i.e., tighter than free fit, a bearing factor is not required.

(4) Military standard part specification, MS 21240, “Bearing, Sleeve Plain, TFE Lined” and MS 21241, “Flanged Bearing, Sleeve Plain, TFE Lined contain allowable Page D - 27 AC 29-2C, Chg 4 5/1/2014 load ratings, static, and dynamic that apply to the particular use of the bearing. An appropriate bearing factor should be applied to the static rating. Military Specification MIL-B-8943A, Amendment 3, “Bearing, Sleeve, Plain, and Flanged, TFE Lined” (temperature range -65° F to 250° F) shows that MS 21240 and MS 21241 sleeve bearings have been superseded by MS 1934/1 and MS 81934/2 sleeve bearings, respectively. Military Specification MIL-B-81934, Amendment 2, “Bearings, Sleeve, Plain and Flanged, Self-Lubricating,” uses TFE liners. These bearings are intended for use in a temperature range from -65° F to +325° F. Whenever a sleeve bearing is used an appropriate bearing factor should be applied to the static rating that is contained in the specification or standard. Other sleeve bearings are contained in standards NAS 72 through NAS 77, NAS 537, and NAS 538. The installation design information is only contained in standards NAS 72 through NAS 74. These types of plain sleeve bearings are designed for clamping to the shaft or bolt with relative motion occurring on the bearing outside diameter. An appropriate bearing factor is required for the application.

(5) The minimum fitting factor 1.15, specified by § 29.625, must be applied as specified to account for load distribution at the fitting. This fitting factor need not apply to plain or journal “bearings” whose “bearing factor” exceeds 1.15.

(6) For airframe and landing gear structural joints, the manufacturer’s static, non-Brinell rating of ball and roller bearings may not be exceeded. ABEC Class 1 bearings or better quality bearings may be used in airframe structural joints and landing gear; ABEC Class 3, 5, or 7 bearings should be used in rotor pivot joints. The non-Brinell rating includes consideration of the bearing factor and no other bearing factor is required.

(7) A free fit was described in American Standards Association (ASA) Standard B4a-1925. The “free fit” clearances and tolerances of this old standard are now called Class RC6, Medium Running Fit, in ASA Standard B4.1, 1955. As an illustration using these standards, a 1-inch diameter shaft and a plain sleeve bearing would have a clearance ranging from .0014 to .0040 inch.

Page D - 28 5/1/2014 AC 29-2C, Chg 4 AC 29.625. § 29.625 FITTING FACTORS.

a. Explanation . A 1.15 factor is specified to assure that the calculated load and stress distribution within any fitting is conservative. Application of the factor is excluded or excepted as stated in the rule.

b. Procedures .

(1) The factor may be applied to the calculated load or stress for the fitting.

(2) The structural substantiating data for the rotorcraft, including the rotor system, must include the prescribed fitting factor. The rotor system includes the flight control system rotor head and hubs and rotor blade attachments, rotor head and hubs, and boosted control system elements. Other typical areas that may be considered are tail rotor gearbox attachment, tailboom to fuselage fittings, transmission pylon attachments, and landing gear attachment to the rotorcraft.

(3) The fitting factor is not required in the following applications: (i) Joints such as continuous joints in metal plating, welded joints and scarf joints in wood.

(ii) Elements proven by limit or ultimate load tests such as nonboosted control system parts.

(iii) Elements for which a larger load factor is used such as a casting factor, a 1.33 retention factor when required for seats and safety belts, a fatigue factor, bearing factor or special factor greater than 1.15, crash load factors that are the only design case, and crash load factors that exceed limit load factors x 1.5 x 1.15.

(iv) Elements for which the failure mode does not affect safety of flight or occupant safety.

AC 29.625A § 29.625 (Amendment 29-42) FITTING FACTORS.

a. Explanation. Amendment 29-42 added § 29.625(d) that requires a 1.33 factor applied to the emergency landing loads of § 29.561(b)(3) for the substantiation of attachments of each seat, berth, and litter to the structure and each safety belt or harness to the seat, litter, or structure.

b. Procedures. All of the advisory material pertaining to this section remains in effect with the following additions.

(1) A fitting factor of 1.33 must be applied to the emergency landing loads of § 29.561(b)(3) when evaluating the attachments of the seat, berth, and litter to the Page D – 28.1 AC 29-2C, Chg 4 5/1/2014 structure, and each safety belt and harness attachment to the seat, berth, litter, or structure.

(2) The 1.33 factor is required whether analysis or test is used.

Page D – 28.2 5/1/2014 AC 29-2C, Chg 4 AC 29.629. § 29.629 FLUTTER.

a. Explanation .

(1) The rotorcraft must be free from flutter.

(2) Section 29.251 vibration is an associated flight requirement concerning flight demonstrations. See paragraph AC 29.251 for this standard.

(3) Section 29.571(a)(3) concerns in-flight measurement of loads or stresses.

b. Procedures .

(1) Freedom from flutter may be shown by analysis or appropriately instrumented flight flutter tests.

(2) The flight loads survey proposal submitted for compliance with § 29.571 may also contain tests to fulfill compliance with § 29.629. The flight loads survey program encompasses the envelope of design airspeed and rotor RPM, and sufficient aerodynamic excitation is generally present to excite any latent flutter modes.

(3) Flight loads survey data or flight flutter test data submitted should be reviewed to assure that excessive oscillatory loads of rotors or surfaces will not be encountered.

AC 29.629A. § 29.629 (Amendment 29-40) FLUTTER AND DIVERGENCE.

a. Explanation . Amendment 29-40 adds the requirement that each aerodynamic surface of the rotorcraft must be free from divergence in addition to the requirement of freedom from flutter. The aeroelastic stability evaluations required by this regulation include flutter and divergence. Compliance with this regulatory requirement should be shown by analysis and/or flight test, supported by any other means found necessary by the Administrator. The aeroelastic evaluation of the rotorcraft should include an investigation of the significant elastic, inertia and aerodynamic forces on all aerodynamic surfaces (including rotor blades) and their supporting structure. The forces associated with the rotations and displacements of the plane of the rotors should be considered.

b. Procedures .

(1) It should be shown by analysis that the rotorcraft is free from flutter and divergence (unstable structural distortion due to aerodynamic loading) under any condition of operation including: (i) Airspeeds up to 1.11 V (power on and power off).

NE Page D - 29 AC 29-2C, Chg 4 5/1/2014 (ii) Main rotor speeds from 0.95 x the minimum permitted speed up to 1.05 x the maximum permitted speed (power on and power off).

(iii) The critical combinations of weight, CG position, load factor and altitude.

(2) Adequate tolerances should be established on those physical quantities which could affect flutter, divergence, or structural distortion to a degree sufficient to cause a significant deterioration in the characteristics of the rotorcraft, such that likely variations in these quantities will not result in flutter or divergence within this envelope.

(3) All physical properties which could contribute to a reduction in the predicted flutter or divergence margins are to be investigated, including stiffness, damping, mass balance, and aerodynamic coefficients. Parametric variations should be sufficient to cover any possible variation due to manufacturing and maintenance tolerances and environmental factors, and to provide conservatism where estimated values are used.

Linear approximations to non-linear variations may be used.

(4) Where approval for flight in icing conditions is being sought, the effects of ice accretion on unprotected surfaces, including that which might occur as a result of a single system malfunction, should be considered.

(5) Rotorcraft should be demonstrated by suitably instrumented flight tests to be free from flutter and divergence at all combinations of forward speed and rotor RPM (power off and power on), up to 1.11 V and 1.05 times the maximum permitted RPM NE (except that combinations of speed in excess of V and rotor speed in excess of the NE maximum permitted are not required to be tested). Flight tests to demonstrate compliance with flutter and divergence requirements may normally be addressed simultaneously with testing in compliance with §§ 29.251, 29.571, 29.1505 and similar regulations. Special flight tests for flutter and divergence would not normally be required.

(6) Stabilizing surfaces may be addressed by analysis alone if flutter and divergence margins can be shown to provide adequate conservatism. Flight testing at 1.05 times the maximum power-on rotor speed may also be waived if it is considered impractical, and can be adequately addressed by analysis.

Page D - 30 5/1/2014 AC 29-2C, Chg 4 AC 29.631. § 29.631 (Amendment 29-40) BIRD STRIKE.

a. Explanation . Amendment 40 adds requirements for continued safe flight and landing after a bird strike. Compliance with § 29.631 must be shown for a 2.2 lb (1.0 kg) bird at a relative velocity equal to the lesser of V or V at altitudes up to 8,000 feet.

NE H For Category A certification, the rotorcraft must be capable of continued safe flight and landing after the described bird strike. For Category B certification, the rotorcraft must be capable of a safe landing after the bird strike.

b. Procedures . For compliance with FAR 29.631, it should be demonstrated by test or analysis supported by test evidence that, (1) The windshields will withstand the bird strike, without penetration, and, (2) The rotorcraft is capable of continued safe flight and landing following impact with a 2.2-lb (1.0 kg) bird at V or V (whichever is the lesser) at altitudes up to NE H 8,000 feet. Areas of impact that are of particular interest include flight control surfaces (which includes main and tail rotors) and exposed flight control system components.

Page D - 31 AC 29-2C 9/30/99 SUBPART D - DESIGN AND CONSTRUCTION ROTORS AC 29.653. § 29.653 (Amendment 29-3) PRESSURE VENTING AND DRAINAGE OF ROTOR BLADES.

a. Explanation . The rule requires each rotor blade to be provided with venting and drainage means (i.e., holes, etc.) or the blade must be sealed and designed to withstand internal pressure.

b. Procedures . Although the rule provides for venting and drainage features, recently certificated blades have been designed to be sealed and to sustain the “maximum pressure differentials expected in service.” For modern blade designs, the internal pressure buildup due to environmental effects and centrifugal acceleration effects (near the tip) can be readily sustained with moisture sealing accomplished. The use of sealed blades is highly advantageous and recommended because of the possibility for severe corrosion damage resulting from trapped moisture and because of the difficulty in finding internal corrosion damage by use of field level inspections.

AC 29.659. § 29.659 (Amendment 29-3) MASS BALANCE.

a. Explanation . The rule requires that mass balancing of rotors and blades be provided, as necessary, to prevent excessive vibration and flutter. Further, the rule requires structural substantiation of the mass balance installation.

b. Procedures .

(1) The weight, geometry, and location of rotor and blade mass balance devices are determined as the requirements of §§ 29.571 and 29.629 are met.

(2) The structural substantiation should show static strength to meet the maneuver and gust loads of §§ 29.337, 29.339, and 29.341. In addition, the main rotor loads of § 29.547(c) should be substantiated. The fatigue strength of the mass balance devices (including structural supports) should meet the requirements of § 29.571.

(3) In addition to the appropriate strength requirements, some recent designs have included features which trap the balance weight inside a limited area even if the primary attachment means (adhesive, bolts, etc.) fail. This type of design feature is recommended because of the severe loading environment to which balance devices are subjected.

AC 29.661. § 29.661 (through Amendment 29-3) ROTOR BLADE CLEARANCE.

a. Explanation .

Page D - 32 AC 29-2C 9/30/99 SUBPART D - DESIGN AND CONSTRUCTION ROTORS AC 29.653. § 29.653 (Amendment 29-3) PRESSURE VENTING AND DRAINAGE OF ROTOR BLADES .

a. Explanation . The rule requires each rotor blade to be provided with venting and drainage means (i.e., holes, etc.) or the blade must be sealed and designed to withstand internal pressure.

b. Procedures . Although the rule provides for venting and drainage features, recently certificated blades have been designed to be sealed and to sustain the “maximum pressure differentials expected in service.” For modern blade designs, the internal pressure buildup due to environmental effects and centrifugal acceleration effects (near the tip) can be readily sustained with moisture sealing accomplished. The use of sealed blades is highly advantageous and recommended because of the possibility for severe corrosion damage resulting from trapped moisture and because of the difficulty in finding internal corrosion damage by use of field level inspections.

AC 29.659. § 29.659 (Amendment 29-3) MASS BALANCE .

a. Explanation . The rule requires that mass balancing of rotors and blades be provided, as necessary, to prevent excessive vibration and flutter. Further, the rule requires structural substantiation of the mass balance installation.

b. Procedures .

(1) The weight, geometry, and location of rotor and blade mass balance devices are determined as the requirements of §§ 29.571 and 29.629 are met.

(2) The structural substantiation should show static strength to meet the maneuver and gust loads of §§ 29.337, 29.339, and 29.341. In addition, the main rotor loads of § 29.547(c) should be substantiated. The fatigue strength of the mass balance devices (including structural supports) should meet the requirements of § 29.571.

(3) In addition to the appropriate strength requirements, some recent designs have included features which trap the balance weight inside a limited area even if the primary attachment means (adhesive, bolts, etc.) fail. This type of design feature is recommended because of the severe loading environment to which balance devices are subjected.

AC 29.661 § 29.661 (through Amendment 29-3) ROTOR BLADE CLEARANCE .

a. Explanation .

Page D - 32 9/30/99 AC 29-2C (1) This paragraph discusses the regulatory requirement contained in § 29.661.

That requirement is that there must be enough clearance between the rotor blades (main and tail rotor blades) and other parts of the structure to prevent the blades from striking any part of the structure during any operating condition.

(2) In the past, some rotorcraft that have been shown to comply with § 29.661 during the certification process have experienced subsequent accidents involving in-flight contact between the main rotor and airframe (rotor/airframe contact).

Completion of developmental and TIA flight testing without a rotor/airframe contact incident has proven not to be adequate demonstration of compliance with § 29.661 in all cases.

(3) Historically, in-flight rotor/airframe contact accidents have occurred as a result of mast bumping, rotor stall, or excessive rotor flapping due to control manipulation. For some rotorcraft, a more thorough examination may be required to ensure adequate clearances.

b. Procedures . Testing should be conducted by the applicant, prior to FAA/AUTHORITY participation, to ensure that the rotorcraft is in compliance with § 29.661 in all areas of the envelope during all operational maneuvers expected throughout the life of the aircraft. The tests should be performed concurrently with performance, flight characteristics, and flight loads testing. Tests should include: (1) A blade flapping survey to determine flapping angles/margins, blade bending, and blade clearance from the entire airframe. Data may be gathered from instrumented flapping hinges, instrumented blades, high-speed video from airframe mounted cameras, a chase aircraft, or other acceptable means.

(2) Determine that margin exists between the minimum rotor RPM encountered during testing for compliance with § 29.143(d) and the RPM (power off) at which analysis shows that the rotor will experience a significant stall. A significant stall condition may be defined by the rotor reaching an RPM from which normal operating RPM is unrecoverable due to drag on the main rotor blades or, a stall that results in excessive main rotor flapping. The rotor RPM decay rate under the critical conditions of weight, density altitude, minimum approved power-on rotor RPM must provide a margin between the minimum rotor speed achieved during demonstration of compliance with §§ 29.87 and 29.143(d) and the analytically derived rotor stall RPM for the same conditions. For example, the minimum rotor RPM resulting during H-V tests must allow for a margin above the rotor stall value to allow for variations that may occur during operational flying.

(3) During parts of the certification flight test program, frangible devices (wood dowels) or other means of measuring clearance, may be requested to confirm that the clearances shown in the drawings and verified during company flight tests are adequate in all operating conditions. Balsa wood dowels or styrofoam pads may be clamped to Page D - 33 AC 29-2C 9/30/99 the aft part of the fuselage and cabin roof within the rotor arc. Such devices may be especially helpful in determining clearance during autorotation and controllability testing under FAR 29.143. If such measuring devices are used, the type inspection report should contain a record of clearance found during the tests. During TIA flight testing, it is not necessary to precisely determine the clearance but only necessary to determine "enough clearance" as stated in the rule.

AC 29.663. § 29.663 GROUND RESONANCE PREVENTION MEANS .

a. Explanation .

(1) This section, adopted in Amendment 29-3, and amended by Amendment 29-30, requires reliability and damping action investigation for the ground resonance prevention means which typically includes the shock struts. Section 29.1529 requires associated maintenance information in the maintenance manual. The probable range of variations in service, not just the allowable range, should be established and investigated as prescribed. This probable range includes operation on the ground, or other appropriate landing surface applicable to the rotorcraft design. Quantitative test data are generally obtained in compliance with this rule although analysis or tests may be employed. The preamble to Amendment 29-3 contains additional information.

(2) Note that the maintenance information is not contained in the approved mandatory section of the maintenance manual.

(3) Paragraph AC 29.241 concerns demonstrating freedom from ground resonance during certain applicant and TIA verification evaluations or tests of the rotorcraft. Section 29.241 complements the requirements of § 29.663. As noted in paragraph AC 29.241, a specific requirement for a ground vibration survey was removed from CAR Part 7. However, § 29.663 was adopted by Amendment 29-3 to investigate possible sources of ground resonance and to assure that the reliability of the ground resonance prevention means; i.e., dampers, shock struts, etc., would preclude the occurrence of ground resonance. The total rotorcraft system, including landing gear, struts, tires, etc., is evaluated under this standard.

(4) Viscous dampers in the rotor head have been used for many years to prevent ground resonance. Modern rotorcraft designs may also use elastomeric dampers and may use elastomeric bearings in the rotor head and rotor pylon attachment to the airframe. The standard applies to viscous and elastomeric dampers.

The “probable” range in damping shall be investigated. The standard also requires investigation of the probable range of variations of these dampers, whether viscous or elastomeric, and elastomeric bearings to preclude ground resonance.

(5) Ground resonance can occur due to flexibility in the rotor pylon restraint system as well as with landing gear flexibilities and/or shock struts. See paragraph AC 29.663b(2) for an explanation. An analysis may be done to show the effect of the rotor pylon mount stiffness on ground resonance stability. If the analysis Page D - 34 9/30/99 AC 29-2C shows that rotor pylon mount stiffness could affect ground resonance, the evaluation should include variations in stiffness and damping of the rotor pylon restraints that may occur in service (reference “Ground Vibrations of Helicopters,” M.L.. Deutsch, JAS, Vol. 13, No. 5, May 1946).

b. Procedures .

(1) The reliability of the means for preventing ground resonance may be substantiated as stated in the standard. An analysis report or a test proposal and subsequent test report may be used to show compliance. The probable range of variations, in service, of the damping action are an important part of the assessment.

The test may be conducted in conjunction with the testing required by § 29.241. See paragraph AC 29.241.

(i) Analysis and tests may be used.

(ii) Reliable service history of identical or closely similar systems may be used. The materials and fluids used, clearance or fits, seals, and physical installation are important items to be evaluated and considered for “closely similar” systems.

(iii) Testing of the complete rotorcraft may be used to prove that malfunction of a single means of the damping system will not cause ground resonance.

One method of demonstrating acceptable compliance is by removing all seals, if practicable, from one damper. Another method is to remove all or most of the fluid, in conjunction with considering the allowable ranges of damping of the other parts of the rotorcraft damping system and operating the rotorcraft throughout the rotor speed range from start to maximum rotor speed. Investigation of elastomeric dampers may require innovative test procedures and preliminary discussions of these prior to preparation of a test proposal. The rotorcraft cyclic control should be displaced as noted in paragraph AC 29.241 to assure that the possible rotorcraft resonance frequencies are excited. If vibrations are damped in all tests, the damping system is satisfactory. Each critical rotor damper and landing gear damper, which includes shock struts and tires, should simulate a malfunction to comply with the standard. The testing discussed, however, could become very extensive if one were to attempt to test all combinations of all maintenance adjustments of all components which contribute to the prevention of ground resonance, while at the same time rendering each of the pertinent components ineffective in turn and then repeating all of the maintenance tolerance testing each time.

Fortunately, rational analytical methods are available which will permit the evaluation of such combinations so that only the combinations with the least amount of margin used are physically tested.

(2) The pylon damper variation can affect ground resonance. The variations in stiffness and/or damping of pylon mounts should be evaluated except the pylon mounts on contemporary conventional rotorcraft may have little influence on “classical” ground resonance stability. The dynamics of the rotorcraft on its landing gear is generally established by the airframe properties and the landing gear properties under the Page D - 35 AC 29-2C 9/30/99 influence of the rotor system, with the “pylon” having little effect. For air or flight resonance, the rotor generally couples with the rigid body modes of the fuselage. For a specific design, a relatively simple analysis may be used to show the effect of the pylon mount system stiffness on air and ground resonance stability, and if not important, variations in the system may be omitted from the test program.

(3) The probable ranges of damping shall be established and investigated as prescribed and noted in paragraph (b) of § 29.663. An approved test proposal and test results report should be used for complying with § 29.663(b). For example, if a conventional wheel landing gear is used on the rotorcraft, the probable ranges of tire pressure or the lowest probable tire pressure should be stated in the test proposal and effects of the tire pressure investigated during the test. In addition, the effects of strut pressures should be investigated also. See paragraph AC 29.241, § 29.241, concerning tests and instrumentation of the test associated with complying with § 29.241. The instrumentation noted in paragraph AC 29.241 also applies to § 29.663(b).

(4) If the wheel landing gear is equipped with wheel brakes, the evaluation should include brakes “on” and “off.” The nose or tail wheel should be locked and unlocked if it swivels to evaluate any possible adverse effects of this feature.

AC 29.663A. § 29.663 (Amendment 29-30) GROUND RESONANCE PREVENTION MEANS .

a. Explanation . Amendment 29-30 clarifies that analysis as well as tests may be used to show freedom from ground resonance after malfunction or failure of a single means of ground resonance prevention. This amendment primarily clarifies that the probable range of damping should be established as well as investigated.

b. Procedures . The procedures of paragraph AC 29.663 continue to apply with the addition of the need to document the establishment of probable range of damping of ground resonance prevention means. Acceptable tire and oleo minimum and maximum pressures as well as other identified factors should be documented in maintenance instructions if necessary to assure the desired characteristics.

Page D - 36 9/30/99 AC 29-2C SUBPART D - DESIGN AND CONSTRUCTION CONTROL SYSTEMS AC 29.671. § 29.671 (Amendment 29-24) CONTROL SYSTEMS - GENERAL .

a. Explanation .

(1) The rule requires that controls operate easily and smoothly and provide positive response of the rotorcraft to control input.

(2) In addition, the rule requires that incorrect assembly be prevented by special design features or special markings.

(3) After Amendment 29-24, November 6, 1984, the rule requires that the flight control system be designed such that the full range of flight control authority can be verified by the pilot before flight. This check would normally have to be completed prior to turning the rotor since control extremes typically cannot be reached with the rotor operating on the ground.

b. Procedures .

(1) Easy, smooth operations of controls are substantiated by the operations tests of § 29.683 and the FAA/AUTHORITY flight testing under TIA procedures.

Positive response of the rotorcraft to control inputs is also evaluated during company flight testing and FAA/AUTHORITY TIA flight testing to the requirements of §§ 29.141 through 29.175.

(2) To meet the requirement that incorrect assembly be prevented, the preferred method is providing design features which make incorrect assembly impossible. Typical design features which can be used are different lug thicknesses, different member lengths, or significantly different configurations for each system component. In the event that incorrect assembly is physically possible (because of other considerations), the rule may be met by the use of permanent, obvious, and simple markings. Permanent (durable) decals or stencils may be used.

(3) Design features of the control systems are checked when reviewing the type design drawings. During the proof and operation tests of §§ 29.681 and 29.683, the controls should be thoroughly reviewed for possible incorrect assembly and for any required markings supplied for compliance with this standard.

AC 29.672. § 29.672 (Amendment 29-24) STABILITY AUGMENTATION, AUTOMATIC, AND POWER-OPERATED SYSTEMS .

a. Explanation .

Page D - 37 AC 29-2C, Chg 2 4/25/06 (1) This rule requires that the pilot be made aware of stability augmentation, automatic, or power-operated system failures which could lead to an unsafe condition.

It should be understood that this requirement applies to stability augmentation and supplementary controls and not the primary flight control system, which is dealt with under § 29.695 and associated advisory material. Examples of clearly distinguishable warnings include, but are not limited to, an obvious aircraft attitude change following the failure or an audio warning tone. A visual indication itself may not be adequate since detection of a visual warning would normally require special pilot attention. The use of devices such as stick pushers or shakers is not acceptable as a warning means since the automatic flight control systems (AFCS) may provide a hands-off capability or normal helicopter vibrations could mask a control shaker. However, this rule is not intended to eliminate the use of such devices for other purposes. Examples of automatic control systems other than a stability augmentation system would be a pitch axis actuator used for the purpose of demonstrating compliance with longitudinal static stability requirements or a fly-by-wire elevator. The design of such systems must not interfere with completion of the control checks described in § 29.671(c). Further, for control systems where a series actuator malfunction could degrade control authority, a means should be provided to the pilot to determine actuator alignment (see § 29.1329(b)).

(2) The corrective flight control input following a system failure should be in the logical direction. For example, a malfunction resulting in a nosedown pitch of the aircraft should require a corrective cyclic control input in the aft direction. The system deactivating means does not have to be located on the primary flight control grips; however, it should be easily accessible to the pilot. Consideration should be given to the consequences of inadvertent de-selection of the automatic stabilization system, especially if the deactivation control is mounted on a primary control grip. Malfunctions and subsequent recoveries must be shown throughout the operating envelope of the aircraft. In a case where control authority is decreased following a malfunction, a practical flight envelope must be defined wherein compliance with controllability and maneuverability requirements can be demonstrated. This practical flight envelope must be presented in the flight manual. Compliance with trim and stability characteristics is not required following a malfunction; however, a pilot workload assessment should be made to show that these characteristics are not impaired below that needed for continued safe flight and landing.

b. Procedures . A discussion of malfunction test procedures is presented in paragraph AC 29 Appendix B b(6). Controllability and maneuverability test procedures are addressed in paragraph AC 29.143.

AC 29.673. § 29.673 (Amendment 29-24) PRIMARY FLIGHT CONTROLS .

a. Explanation . This section basically defines primary flight controls as “those used by the pilot for immediate control of pitch, roll, yaw, and vertical motion of the Page D - 38 9/30/99 AC 29-2C rotorcraft.” This section clarifies the application of § 29.1555 which requires markings for controls other than “primary flight controls or control(s) whose function is obvious.” b. Procedures . The primary flight controls (e.g., cyclic stick, collective, and tail rotor pitch control pedals) are excluded from the marking requirements of § 29.1555.

AC 29.674. § 29.674 (Amendment 29-30) INTERCONNECTED CONTROLS .

a. Explanation . A new § 29.674 is added by Amendment 29-30 which requires that the rotorcraft be capable of safe flight and landing after a malfunction, failure, or jam of any auxiliary interconnected control.

b. Procedures .

(1) Section 29.674 requires that the rotorcraft be shown to be capable of safe flight and landing after a malfunction, failure, or jam of an auxiliary control interconnected with a primary control. The section does not apply to interconnected primary controls; e.g., cyclic and collective controls.

(2) Examples of auxiliary controls covered by this section may include certain autopilot or stability augmentation or trim system components. Section 29.1309 methods may be used in determining failure effects of autopilot and stability augmentation system components.

(3) If an engine control could jam and result in a collective control jam, these controls should be designed to relieve that connection.

AC 29.675. § 29.675 (Amendment 29-17) STOPS .

a. Explanation .

(1) Stops are required to prevent unrestrained movements of pilot/autopilot inputs from causing interferences or overloads.

(2) The rule requires that the stop must not appreciably affect the control system range of travel due to wear, slackness, or take-up adjustments.

(3) Each stop is required to withstand loads corresponding to design conditions.

(4) In addition, each main rotor blade, if appropriate for the design, must have stops to limit its travel about its hinge points. For rotors with hingeless design, stops may be provided as appropriate to limit blade travel. Loads which may result from the blade hitting the stops (during starting or stopping the rotor, or during any large but allowable pilot control inputs such as autorotation cyclic traverse or when subjected to ground gusts, etc.) shall not overload the stops nor any rotor component.

Page D - 39 AC 29-2C 9/30/99 b. Procedures .

(1) Stops are generally provided in the cockpit area and near any controllable surface end of the control system (i.e., main rotor hub, tail rotor hub, and stabilizer activators). For systems with control coupling or series actuators, stops have been located further away from the cockpit to permit increased control output during malfunction (hardover) or extreme control position cases.

(2) Location of stops in close proximity to each end of a control system will allow the stops to function most efficiently without undue deflections between the stops and the adjacent surface or the adjacent cockpit control lever or pedals. The location of stops close to the control lever or surface will help meet the requirement that the stop and its function not be appreciably affected by wear, slackness, or take-up adjustments.

Consideration should be given to limiting the total amount of take-up adjustments of both the stop and the control systems to preclude a hazardous adjustment of the control surface range of travel.

(3) Each stop is to be substantiated for critical design conditions from either pilot effort, aerodynamic loads, hydraulic loads, or other critical loads, as applicable.

The stops can be substantiated for limit loads by the tests of § 29.681. (Deliberate misrigging of the controls on the test aircraft may be necessary to assure that the maximum limit load which the stop will be subjected to in service is applied to the stop during these tests.)

(4) The stops to limit the main rotor blade about its hinge points should be positioned to prevent the blades from striking any part of the structure, particularly during startup and shutdown operations. These stops should also limit the flapping of the static main rotor blades of the rotorcraft when they are subjected to ground gusts or rotor wash from nearby taxiing rotorcraft. Provisions should be made to prevent overloading the stops or the blade under conditions of ground gusts, rotor wash effects, or during autorotation landing flares. The need for provisions to prevent possible overloads due to ground gusts and close taxiing by adjacent rotorcraft and by autorotation landings can be determined using the instrumented flight load survey aircraft by hover-taxiing another rotorcraft near the instrumented aircraft and by conducting autorotation landing flares with the instrumented aircraft. Substantiation for the final main rotor flapping stop design can be demonstrated by similar tests.

(5) If features of design are added to the main rotor stop assembly which activate certain portions of the stop assembly only on the ground to meet the requirement that the blade not hit the droop stop during any operation other than starting and stopping the rotor, such features of design must be substantiated to reliably operate by both ground tests and flight tests, as appropriate. Wear and rigging tolerances should be considered in these demonstration tests.

Page D - 40 9/30/2008 AC 29-2C, Chg 3 AC 29.679. § 29.679 CONTROL SYSTEM LOCKS .

a. Explanation .

(1) Whenever a control system lock or locks are used, the standard requires design features to prevent flight or limit operation before flight begins with the lock engaged. Locks are not required by the standard.

(2) After flight begins, design features shall be used when needed to prevent possible lock engagement while the rotorcraft is in flight or ground operation.

(3) The standard applies to external control locks as well as internal locks.

b. Procedures .

(1) Locks that release or disengage automatically, as stated, may be used.

Attention should be directed to reviewing possible means of lock engagement while in flight. Fault analysis of the system should be used to ensure possible failures are determined. Design features may be used or needed to preclude this event.

(2) Manually applied and released locks may be used. Design features of the locks must prevent engagement in flight also.

(3) Any “unmistakable” warning to prevent takeoff with a lock engaged should be easily discernable during day and night operations. It should be possible to apply the lock only in such a manner that the required warning is provided. Color, location, shape (identification), and accessibility of the device or its control and legibility of any device placards or markings are important considerations in the evaluation.

(4) During a “compliance inspection,” and during TIA evaluations, the locks shall be evaluated to the standards. When a lock is not automatically disengaged, the operation of the rotorcraft should be limited. Unmistakable warning may be achieved as follows.

(i) Prevent sufficient power for takeoff.

(ii) The pilot shall be unable to move the collective control from the lowest pitch limit.

(iii) One or more aural devices that cannot be disengaged (turned off) until all locks are removed.

(5) The rotorcraft Instructions for Continued Airworthiness should include appropriate maintenance checks and procedures to be completed following modification (for example, via STC or field approval), maintenance, alignment, or adjustment that affects the flight control system locks.

Page D - 41 AC 29-2C, Chg 3 9/30/2008 AC 29.681. § 29.681 LIMIT LOAD STATIC TESTS .

a. Explanation .

(1) The rule requires static tests of the control system in showing compliance with limit load requirements.

(2) The tests are specified to include each fitting, pulley, and bracket of the control system being tested and to include the “most severe loading.” (3) Also, the rule requires that compliance with bearing factors (reference § 29.623) be shown by individual tests or by analyses for control system joints subject to motion.

b. Procedures .

(1) Compliance with the requirements of this rule is obtained by static tests conducted on either a static test airframe or on a prototype flying ship. In either case, conformity of the control system and related airframe is necessary to validate the tests.

(2) The rotor blades or aerodynamic surfaces may be used to react pilot effort loads through the control system or they may be replaced with fixtures. If fixtures are used, they should be evaluated for geometric and stiffness effects to assure test validity.

(3) The loads to be applied during the limit load static tests are specified in §§ 29.395, 29.397, and 29.399. The loads are applicable to collective, cyclic, yaw, and rotor blade control systems as well as any other flight control systems provided by the design.

(4) Section 29.585(e) specifies bearing factors for control system joints subject to angular motion. These factors are 3.33 for push-pull systems and 2.0 for cable systems for joints with plain bearings. For joints with ball or roller bearings, use the manufacturer’s ratings.

AC 29.683. § 29.683 OPERATION TESTS .

a. Explanation . The rule requires that the control system be free from jamming, excessive friction, and excessive deflection. An operational test is required in which specified loads are applied at the pilot controls and carried through an operating control system.

b. Procedures .

(1) Compliance with the requirements of this rule is obtained by use of a test setup similar to that used for the limit load tests of § 29.681, except the load reactions at Page D - 42 9/30/2008 AC 29-2C, Chg 3 the blades (or surfaces) must allow for movement of the blades (or surfaces) as the system is operated through its operating range.

(2) Fixtures are normally affixed to the surfaces (or replace the surfaces) to allow pulley arrangements which provide for movement under load. These fixtures should be evaluated to assure that system loads up to limit will be applied during the full range of operations of each system.

(3) Each flight control system should be operated through its entire range under a light load and under limit load. As the controls are being operated, the system should be checked for jamming, excessive friction, and excessive deflection. Excessive deflection includes deflection sufficient to contact other systems or structure. Also, if under these limit load conditions the components deflect, the deflection would be considered excessive if there is permanent deformation of any component or supporting structure. Also any deflection that results in an uncorrected condition when the load is released, e.g., if a bellcrank is forced off-center or over-center during load and does not return to the normal position after load release is excessive deflection. Floor panels, wall panels, and other access panels may have to be removed to permit visual checks of the entire control system. However, care should be taken when removing panels so that airframe structure is not weakened enough to deflect from its normal position when test loads are applied to the control system.

AC 29.685. § 29.685 (Amendment 29-12) CONTROL SYSTEM DETAILS .

a. Explanation . The rule requires that the control system be designed to prevent chafing, jamming, and interference from cargo, passengers, loose objects, or the freezing of moisture. Specifically, means are required in the cockpit to prevent the entry of foreign objects into places where they would jam the system, and means are required to prevent the slapping of cables or tubes against other parts. Specific design considerations to prevent binding and overloads within the control system are required such as- - (1) Assure pulley-cable combinations as specified in MIL-HDBK-5 are used unless inapplicable.

(2) Assure close fitting pulley guards are provided.

(3) Assure pulley-cable alignment sufficient to prevent excessive pulley flange loads is provided.

(4) Assure fairlead-cable alignment is within 3°.

(5) Assure no clevis pins are retained only by cotter pins.

(6) Assure turnbuckles do not bind other structures throughout the range of travel.

Page D - 43 AC 29-2C, Chg 3 9/30/2008 (7) Assure means for inspection of control system components are provided.

(8) Assure control system joints subject to angular motion incorporate special bearing factors, 3.33 for push-pull systems and 2.0 for cable systems.

(9) Assure that manufacturer’s ratings for ball or roller bearing ratings are not exceeded.

b. Procedures .

(1) The geometry of the control system components and installations is the primary control to prevent chafing, jamming, and interference. The control system from cockpit to surface should be checked for clearances both unloaded and loaded. The control system should be checked under load during both the limit load static tests (reference § 29.681) and the operational tests of § 29.683. Location of guides or fairleads and pulleys may be used in cable systems to prevent chafing and interference with other structure. Generally, tubes should clear adjacent structure by location and design geometrical considerations. If supplemental means are provided to assure the tubes do not chafe or interfere, the means should be evaluated for possible jamming.

(2) Rubber (or other elastomeric) boots connected to both the cockpit control arm or shaft and to the floor are acceptable means to prevent the entry of foreign objects into underfloor areas where they may cause jamming of controls. Control systems should, in general, be routed around cargo compartments. If routing of the control system components is in or near cargo areas, the control system components should be protected by bulkheads, panels, or other enclosures which have sufficient strength and stiffness to prevent possible interference with the control system components when subjected to cargo loading and handling deflections.

(3) Control system details should be reviewed for possible moisture collection.

Areas should drain free. Exposed or open control areas should drain free, and areas of possible freezing moisture collection should not accumulate ice that would cause a jam of the controls. Simulated or actual ice collection on the controls may be used to prove questionable features. The areas to be considered for moisture collection include both external and internal areas where moisture may accumulate by direct impingement of water, entrapment of water particles, or condensation of moisture.

(4) The latest revisions of MIL-HDBK-5 do not explicitly give approved pulley-cable combinations, but appropriate MIL specifications are given in Chapter 8.3 for use in determining pulley-cable combinations and ratings.

(5) Provide ratings, factors, and alignment as specified.

(6) Provide inspection means as specified.

(7) Provide close fitting pulley guards as specified.

Page D - 44 9/30/99 AC 29-2C AC 29.687. § 29.687 SPRING DEVICES .

a. Explanation .

(1) This standard for control systems assures that springs and spring devices used to prevent flutter, control oscillations, or vibrations are either - - (i) Reliable; or (ii) The failure is not critical to the rotorcraft.

(2) Tests simulating service conditions are required in either instance.

b. Procedures .

(1) Springs and spring devices used in the control system, including balance springs, should be identified early in the certification program.

(2) If a spring cannot be shown by observation or analysis to be noncritical, then ground or flight tests may be required.

(3) Springs that are critical to safe operation may be subject to fatigue substantiation to prove they are reliable for the operating conditions imposed in service.

(4) Springs used in conjunction with hydraulic actuator spool valves may be subject to the standards of § 29.695.

AC 29.691. § 29.691 AUTOROTATION CONTROL MECHANISM .

a. Explanation .

(1) Rotorcraft designs generally have a main rotor blade collective pitch control system that does not have detents or other devices to limit pitch control in the control mid-range. Autogyro and other rotorcraft designs may include detents or other finite position control for collective pitch control. This rule requires that the control design allow rapid entry into autorotation after a power failure.

(2) Section 29.33 contains standards concerning establishment and control of the main rotor speed limits. The standard requires flight tests and demonstrations. The standard also concerns rotorcraft design features that are related to control of the main rotor speed limits. Paragraph AC 29.33, § 29.33, pertains to this standard.

(3) Other design requirements for control systems are contained in § 29.685.

b. Procedures .

Page D - 45 AC 29-2C 9/30/99 (1) If high and low main rotor pitch stops are employed in the collective control and if the control may be rapidly moved from one limit to the other, compliance is shown.

(2) If detents or intermediate stops are employed, the pilot must be able to easily and readily override, disconnect, remove, or bypass the device to allow rapid autorotation entry prior to exceeding transient low speed rotor limits. An early assessment for design deficiencies may be accomplished by the flight test personnel with the evaluation completed in the Type Inspection Authorization (TIA) test program.

(3) It is acknowledged that modern rotorcraft designs may have an autorotation V that is lower than power-on V or normal cruise speed. For rotorcraft designs with NE NE this characteristic, the speed must be reduced after entry into autorotation. The rule also applies to rotorcraft designs with this characteristic and no relief from the rule is required since many phases of operation occur at speeds less than power-on V . For NE example, a critical phase of flight occurs during takeoff. Rapid entry into autorotation is essential during this phase also.

(4) The features of the autorotation control mechanism and ability to control the rotor speed within the design limits for any rotorcraft will be evaluated as an integral part of the TIA test program.

AC 29.695. § 29.695 POWER BOOST AND POWER-OPERATED CONTROL SYSTEM .

a. Reference Regulations . The following sections of Part 29 are either incorporated in the provisions of § 29.695 or are otherwise applicable to power boost and power-operated control systems: (1) Section 29.307Proof of structure.

(2) Section 29.571Fatigue evaluation of fight structure.

(3) Section 29.681Limit load static tests.

(4) Section 29.685Control system details.

(5) Section 29.861Fire protection of structure, controls, and other parts.

(6) Section 29.863Flammable fluid fire protection.

(7) Section 29.1301 Function and installation.

(8) Section 29.1309 Equipment, systems, and installations.

b. Explanation .

Page D - 46 9/30/1999 AC 29-2C (1) The rule requires an alternate system if a power boost or power-operated control system is used.

(2) The alternate system must, in the event of any single failure in the power portion of the system, or in the event of failure of all engines: (i) Be immediately available.

(ii) Allow continued safe flight and landing.

(3) The alternate system may be: (i) A duplicate power portion of the system; or (ii) A manually operated mechanical system.

(4) The power portion of the system includes: (i) The power source (such as hydraulic pumps); and (ii) Items such as valves, lines, and actuator.

(5) The failure of mechanical parts (such as piston rods and links) must be considered unless their failure is extremely improbable.

(6) The jamming of power cylinders must be considered unless their jamming is considered extremely improbable.

c. Procedures . It is assumed in the following discussion that the power boost or power-operated control system being utilized is a typical aircraft hydraulic system.

(1) The rule requires, without regard to the probability of failure, an alternate system for the power portion of the system. The power portion of the system, by example in the rule, includes hydraulic pumps, valves, lines, and actuators. It has also been interpreted to include seals, servo valves, and fittings.

(2) If a duplicate power portion of the system is used to meet the requirements of the rule, the requirements may be met by providing a dual independent hydraulic system, including the reservoirs, hydraulic pumps, regulators, connecting tubing, hoses, servo valves, servo-valve cylinder, and power actuator housings. There must be no commonality in fluid-carrying components. A break in one system should not result in fluid loss in the remaining system.

(3) Dual actuators should be designed to assure that any single failure in the duplicated portion of the system, such as a cracked housing, broken interconnecting Page D - 47 AC 29-2C, Chg 3 9/30/2008 input, or broken interconnecting output link, does not result in loss of total hydraulic system function.

(4) A manually operated mechanical system may be used as the alternate system to a single hydraulic system if, after the loss of the single hydraulic system, the pilot can control the rotorcraft without exceptional piloting skill and strength in any normal maneuver for a period of time as long as that required to effect a safe landing.

The control forces should not exceed those specified in § 29.397 and flight characteristics should meet the requirements of §§ 29.141 (b) and (b)(3).

(5) The substantiation of the various system components should include consideration for operation in the normal and alternate system modes.

(6) The “extremely improbable” criteria noted in § 29.695(c) for failure of mechanical parts may be satisfied by performing component fatigue testing and establishing a service life through this technique.

(7) Fatigue substantiation of the control actuator is required under § 29.571 and should consider both the stresses imposed by flight loads and the stresses imposed by hydraulic pump pressure pulses. Flight loads factored in a suitably conservative manner may be an acceptable means to take into account both effects.

(8) The possibility of jamming of the power cylinder may be shown as “extremely remote” through a failure analysis that considers every possible system component failure such as, but not limited to, ruptured lines, pump failure, regulator failure, ruptured seals, clogged filters, jammed servo valves, broken interconnecting servo valve inputs, broken interconnecting output links, etc.

(9) Three acceptable means to meet the requirements of § 29.695(a)(2) could be as follows: (i) Provide two transmission-driven hydraulic pumps, provided the pumps are driven by the transmission during all flight conditions including autorotation.

(ii) Use two electrically driven hydraulic pumps if electrical power is available to drive the pumps with all engines failed. If this approach is used, the battery must be capable of running both pumps plus all other required equipment necessary for continued safe flight.

(iii) Use a single transmission driven pump and an electrically driven pump.

[The remainder of this page intentionally left blank.]

Page D - 48 9/30/99 AC 29-2C SUBPART D - DESIGN AND CONSTRUCTION LANDING GEAR AC 29.723. § 29.723 SHOCK ABSORPTION TESTS .

a. Explanation .

(1) Limit and “reserve energy” drop tests are required as prescribed in §§ 29.725 and 29.727, respectively. These tests must be conducted on the complete rotorcraft or on units consisting of wheel, tire, and shock absorber in their proper relation. For rotorcraft with skid landing gear, the tests may be conducted on the complete rotorcraft or on a simulated fuselage with the complete skid landing gear system.

(2) The rotorcraft must be designed to limit load factors that equal or exceed the limit load factor substantiated by these drop tests. In practical application, the rotorcraft may be designed to a limit load factor, such as 2.8 g. Thus, it is necessary that the limit landing load factor derived from the landing gear drop tests be equal to or less than 2.8 g. If not, the rotorcraft must be redesigned for the higher load factor derived from the drop tests. It must be shown in accordance with § 29.723 that the limit load factors selected for design under § 29.473 will not be exceeded in landings with the limit descent velocity corresponding to the drop height specified in that section. In addition, reserve energy absorption capacity of the landing gear must be shown for a descent velocity of 1.22 times the limit descent velocity selected under § 29.473 by increasing the drop height to 1.5 times the “limit” drop height. The test requirements or procedures outlined in FAR 29 for obtaining the landing load factors are empirical; however, these procedures are based on and supported by satisfactory experience.

(3) As stated in § 29.725(c), each landing gear unit should be tested in the attitude simulating the landing condition that is most critical from the standpoint of the energy to be absorbed by it.

(i) For wheel landing gear designs, the level landing or tail down landing and level landing with drag are generally the most critical attitude. A test of more than one attitude is generally required to comply with the standard. The landing attitudes or conditions prescribed are level (vertical loads), inclined (loads at 14.5° aft from the vertical axis), level with wheel spin-up and tail down. These attitudes are specified in §§ 29.479(b)(1), (2), and (3) and 29.481.

(ii) For skid landing gear designs, the level landing and level landing with drag are generally the most critical attitudes. These attitudes are specified in § 29.501(b) and (c).

(4) Drop tests are required. If analytical methods and/or means are proposed by the applicant, the data presented for approval must be equal to or conservative with Page D - 49 AC 29-2C 9/30/99 respect to that data obtained from physical drop tests. Section 21.21(b)(1) of FAR 21 concerns “equivalency” determinations. Presenting an acceptable means of “equivalency” here would circumvent the necessary scrutiny of an analytical method or means and is also beyond the scope of this document.

b. Procedures . The test plan or proposal must be approved prior to official FAA/AUTHORITY tests unless satisfactory resolution of outstanding proposal or conformity inspection items can be accomplished after the test.

(1) The following headings would be a typical table of contents for the test proposal, and a generalized explanation of the contents that may be included under each of these headings for a wheel landing gear follows.

(i) Purpose . The regulations to which compliance is being shown by the drop tests should be identified (usually §§ 29.723, 29.725, and 29.727). Also, the rotorcraft landing gear including the wheels and tires to be dropped, should be positively identified in the report by the manufacturer’s or applicant’s previously FAA/AUTHORITY approved drawing, technical standard orders (TSO’s), or other identifying FAA/AUTHORITY approved data as applicable.

(ii) Description of test setup . This section should present a description of the test fuselage or jig, method of attaching landing gear to jig, and type of accelerometer to be used to measure load factors. Proof of calibration of accelerometer should be available. The accelerometer should be mounted at the aircraft CG if a free drop of the aircraft is used, or as close as practical to the centerline of the main shock absorbing component of each landing gear (oleo strut, etc.) if each gear is tested separately. The description of the test jig, including platforms on which the gears are to be dropped, should be defined by sketches in addition to the required mathematical calculations. This data should show that the landing gear will be at the proper attitude, relative to the platform, on impact for the particular landing condition. Drawings or other approved data from which the geometry is taken should be referenced in the proposal.

The tire and oleo pressures at the time of the test should be specified. The method of measuring the deflection of the tire plus the vertical travel of the axle under impact should be described. This measurement may be accomplished by telescoping tubes attached to the point on the jig that would measure the total (tire and oleo) vertical deflection of the landing gear. Other vertical and horizontal deflections should be measured as required to determine if the landing gear has experienced permanent deformation after each drop test. The effect of surface roughness should be considered. Smooth surfaces tend to give maximum deflections where rough surfaces tend to restrict deflection and to result in maximum values of N . Preliminary company Z drop tests (at less than limit drop height) may be used to determine the critical surface roughness, or engineering evaluations may be used (without tests) when the gear configurations are such that the critical surface condition can be analytically determined (or when the load factor is shown to be negligibly affected by surface roughness).

NACA Report 1154, dated 1953, contains information that surface coefficients of friction may vary from 0.4 to 0.7. Skid landing gear standards, § 29.501(c), indicate an Page D - 50 9/30/99 AC 29-2C acceptable coefficient of friction is 0.5. A wheel landing gear design standard, § 29.479(b), indicates an acceptable coefficient of friction is 0.25. In the case of a small rotorcraft, the entire aircraft may be dropped. This may be accomplished by establishing pivot points at the main gear axles for the tail (or a point forward of the nose gear) drops, and a pivot point at the tail (or nose gear) axle for the main gear drops. It is the responsibility of the applicant to distribute the aircraft inertia items, including added weight to get the proper effective drop weight (W ) at the landing gear, e so that no local failures of the aircraft occur as a result of the limit or reserve energy drop tests.

(iii) Test data . Computations for the required drop height (h) and the effective drop weight (W ) should be shown for each design level landing and tail down e landing condition in compliance with §§ 29.479 and 29.481. The computations should be in accordance with § 29.725(a) for h and § 29.725(b) for W for the limit drop tests.

e W and h are computed in accordance with § 29.725 for the limit drop test and with e § 29.727 for reserve energy drop test. The computation of the static weight on the gear being dropped (W , W , or W ) and used in the computation of W should be shown.

M N T e This static weight is defined as W , W , or W for the main gears, tail gear, or nose M T N gear, respectively, in § 29.725(d). It should be shown that the critical CG and proposed certificated maximum landing weight have been used in the computation of W , W , or M T W . The computation of the slope of the platforms required for the inclined reaction N conditions should be presented also.

NOTE: Effective drop weight (W ) is used only for free drops. It provides a technique E for accounting for rotor lift without applying lift during the test. If rotor lift is applied during the drop tests, actual weights (W , W , or W ) will be used, not effective weights, M T N W .

E (iv) Test results . The results of the test are based on the values of W , h, E d, W, and L used and obtained for each drop test and the value of Nj obtained from the accelerometer. These results should be summarized, and the method of computing the aircraft limit inertia load factor should be shown for each drop in accordance with § 29.725(d). A print or copy of the film or other recording trace from the accelerometer, if not a direct readout type of accelerometer, should be included in the test results.

Each critical condition should have several preliminary drops as many times as required to obtain reasonable correlation.

(2) Skid landing gear may be tested using similar procedures except a level landing attitude drop test is all that is required by § 29.501. The design load conditions specified in § 29.501(c) through 29.501(f) are derived from this level drop test condition.

(i) Section 29.501, paragraphs (a)(2) and (3), contain special considerations for skid landing gear.

(ii) Section 29.501(a)(2) specifies that structural yielding of elastic spring members under limit load is acceptable. This yielding or deformation is a means of Page D - 51 AC 29-2C 9/30/99 absorbing the landing impact. For skid landing gear that use oleo or other types of shock absorbers, the standard does not allow structural yielding under limit load.

During the limit load and reserve energy (ultimate for skid landing gear with elastic spring numbers) drops, the yielding energy absorbing members will probably deform or yield. After a limit drop test, the gear may be used for a reserve energy drop at the discretion of the applicant but a gear that has been subjected to a reserve energy drop should not be used unless it can be shown that no yielding has occurred in that gear.

(3) Wheel landing gear is tested in attitudes prescribed in paragraph a(3)(i) above. Each unit, nose or main gear, is generally tested separately.

(4) Skid landing gear is tested in attitudes prescribed in subparagraph a(3)(ii) above. Due to the construction of skid landing gear, the complete skid landing gear is tested as a unit. Thus, the level landing with drag condition is probably the critical attitude for the forward cross-tube and its attachments. The level landing condition is probably the critical attitude for the aft cross-tube and its attachments.

(5) An FAA/AUTHORITY or FAA/AUTHORITY designated or delegated person need only witness the drop tests for “record” or “compliance.” Preliminary or developmental drops do not require an FAA/AUTHORITY witness.

AC 29.725. § 29.725 (Amendment 29-3) LIMIT DROP TEST .

a. Explanation . Limit drop tests, in the critical aircraft attitude or critical attitude of each gear, are required for the landing gear. The drop height must be at least 8 inches, which equates to 393 feet per minute (free fall) vertical descent speed. Rotor lift may be simulated and an effective mass may be used in the drop test as prescribed.

b. Procedures . See paragraph AC 29.723, § 29.723.

AC 29.727. § 29.727 RESERVE ENERGY ABSORPTION DROP TEST .

a. Explanation .

(1) In addition to the limit drop tests, a reserve energy drop test is required.

The landing gear must not collapse in this test to the extent that the fuselage impacts the ground. Fracture (to separation) of landing gear parts is considered collapse of the landing gear. This test is not an ultimate load drop test for the landing gear, except as specified in § 29.501(a)(3) for certain skid landing gear designs using elastic spring members.

(2) All other types of landing gear must be substantiated for design ultimate loads in addition to this reserve energy drop test.

Page D - 52 9/30/99 AC 29-2C (3) Shock absorbing devices, such as oleos, must not “bottom” during the reserve energy drop test. “Bottoming” occurs when displacement of the device no longer occurs with increasing load.

(4) Requirements for proof of the landing gear and airframe structure are found in §§ 29.305, 29.307, and 29.473.

b. Procedures . See paragraph AC 29.723, § 29.723.

AC 29.727A. § 29.727 (Amendment 29-30) RESERVE ENERGY ABSORPTION DROP TEST .

a. Explanation . Amendment 29-30 defines the word “collapse” as used in § 29.727(c). Collapse of the landing gear during reserve energy absorption drop tests occurs when: (1) A member of the landing gear will not support the rotorcraft in the proper attitude; or, (2) A member deforms sufficiently to allow the rotorcraft structure other than the landing gear and external accessories to impact the landing surface.

b. Procedures . The procedures of paragraph AC 29.727 continue to apply with the following supplemental guidance.

(1) The proper attitude for the rotorcraft after the reserve energy absorption drop test is an attitude which allows for permanent deformation of landing gear elements but provides for adequate egress from the rotorcraft.

(2) External accessories that may not impact the landing surface during drop testing include devices such as externally mounted fuel tanks or accessories likely to cause post-landing fires. Cameras, loudspeakers, and search lights may be damaged during deformations resulting from reserve energy drop tests if electrical connections are sufficiently protected to preclude electrical fires and the devices are not likely to penetrate fuel tanks. The expendable accessories, if installed, should also be designed to not have “hard points” that would unacceptably damage the rotorcraft structure under landing impacts by penetration into the occupied areas or fuel tanks. These expendable accessories should be designed with frangible fittings, frangible devices, or comparable design features. Also, these devices should be designed to not significantly alter the energy absorbing ability or design features of the landing gear.

AC 29.729. § 29.729 (Amendment 29-24) RETRACTING MECHANISM, LANDING GEAR .

a. Explanation.

Page D - 53 AC 29-2C 9/30/99 (1) Structural substantiation is required for the gear, retracting mechanism, doors, gear supporting structure for landing loads, maneuvering, gusts, and yawing flight condition loads.

(2) An emergency means to extend the gear after failure of the retraction/extension system is required for all except solely manual mechanical systems.

(3) This regulation requires an indication to the pilot when the gear is secured in the extreme positions. This rule does not apply to rotorcraft with fixed gear. The rule also applies to amphibious rotorcraft with retractable gear.

(4) A landing gear down-lock is required. An optional up-lock may be used if it meets reliability requirements.

(5) A (ground) operation test must be conducted to ensure proper functioning of the system.

(6) Location and operation of the control lever or device must comply with § 29.777. This section includes identification of controls to prevent confusion and inadvertent operation. Sections 25.779 and 25.781 of FAR Part 25 contain large airplane design requirements for motion, effect, and shape of cockpit controls and their knobs and should be consulted for further guidance.

b. Procedures .

(1) The design load factors and resulting loads should be derived from the design data. The landing gear, while retracted, operating, and extended, and its supporting structure should be substantiated for the critical aerodynamic and inertia loads. Yawed conditions should be considered. The specific conditions are noted in paragraphs (a)(1), (2), and (3) of § 29.729.

(2) Wheel well doors, if installed, should be designed for the aerodynamic loads, including loads from yawing conditions (angles proven under § 29.351) for airspeeds up to the design maximum landing gear extended speed. Aerodynamic effects on both open and closed doors must be considered in the door and door support substantiations. The applicant may choose to substantiate the rotorcraft for a “landing gear operating” and “extended” speed V and V , respectively, that is equal to the LO LE rotorcraft V . This option will alleviate an airspeed “structural limitation” because of the NE landing gear design substantiation. Any airspeed “structural limitation” should be listed in the structural limitations part of the TIA.

(3) The required “down-lock” should be checked during the operation test. The design drawing should be reviewed for compliance prior to conducting an operation test.

The “down-lock” system should be evaluated for § 29.1309 function and reliability requirements.

Page D - 54 9/30/99 AC 29-2C (4) If an optional “up-lock” is installed (including hydraulic locking), the landing gear should be extended during the operation test after simulation of critical failure mode of the retraction system (reference § 29.1309).

(5) An “operation” test plan or proposal submitted for compliance with § 29.729(d) should include the items noted in the two previous subparagraphs and should include a functional check of the position indicator system. Those tests must be satisfactorily completed before issuing the TIA.

(6) During the official FAA/AUTHORITY flight tests, compliance with the emergency operation, position indicator, and control aspect of § 29.729(c), (e), and (f), respectively, will be verified or accomplished. In addition, the F and R test program plan (§ 21.35) will specify certain tests or evaluations for the retraction system.

(7) Position Indicator Evaluation.

(i) When evaluating the position indicator system, emphasis should be placed on the switches and their installations, and on the cockpit presentation. Each gear must have its own set of switches to indicate when it is secured in its extreme “up” position and its extreme “down” position. The switches must be located to give a valid indication of the arrival of the gear at its extreme position.

(ii) The reliability and environmental qualifications of the switches to be used should be carefully considered. An example of a condition that has potential for trouble is operation on wet areas. Trouble starts when water is picked up by the tires and deposited on the switches. During winter months the water can freeze, and the resulting ice may prevent the switch from functioning properly.

(iii) An acceptable cockpit presentation consists of two lights for each gear. One light is colored “green” and indicates when its gear is secured in the extreme “down” position. The other light is colored “amber” or “red” and indicates when its gear is in transit. When the gear is in either extreme position, the in transit light is “out.” For this presentation, the indication to the pilot that the gear is in the extreme “up” position is an all-gear, lights-out condition.

(iv) Some manufacturers have also included a warning system to alert the crew if the landing gear has not been extended prior to landing. If a warning system is presented, §§ 29.1301 and 29.1309 should be used to evaluate its functional characteristics and the impact of its failure modes.

AC 29.731. § 29.731 WHEELS .

a. Explanation . This standard requires use of approved wheels, either approved under TSO-C26 or a later revision or approved under the type certificate for the aircraft.

Wheels must satisfy both a design static (1g) load and design limit landing or taxiing Page D - 55 AC 29-2C 9/30/99 load determined under the applicable ground load requirements. Standards for a tire installed on a wheel are contained in § 29.733.

b. Procedures .

(1) The structural design loads data shall contain both a static load and a landing and taxiing load for each wheel. These loads are determined by virtue of compliance with the standards of § 29.731(b) and (c). The ratings of the wheel shall not be exceeded. TSO-C26c contains minimum performance standards for TSO approval of aircraft wheels and wheel-brake assemblies. Ratings are assigned in accordance with this performance standard.

(2) If a wheel selected for an aircraft design has TSO approval, the wheel manufacturer will supply the rating to the aircraft manufacturer. Each wheel shall be marked as prescribed which includes a listing of the TSO number. Even though a wheel is TSO approved, the application on the aircraft (loads imposed on the wheel) requires proof that the rating is not exceeded.

(3) If a wheel selected for an aircraft design is not approved under a TSO, the necessary data, both detail design and assembly drawings and qualification tests and test report data, will be required to comply with the standards contained in Part 29.

Design control and inspections will be accomplished as a part of the aircraft type design. Structural substantiation and any appropriate qualification tests shall be accomplished. See §§ 29.471 through 29.497 and § 29.511 for the ground load conditions.

(4) The Tire and Rim Association, Inc., generally issues a yearbook listing tire and rim sizes and ratings. The dimensions and contours for aircraft wheel rims are contained in Section 9 of this yearbook.

AC 29.733. § 29.733 (Amendment 29-12) TIRES .

a. Explanation .

(1) This standard specifies both design and performance criteria for tires. The tire must fit the wheel rim. The maximum static ground reaction for the condition specified must not exceed the maximum static load rating of each tire. In addition, any tire of retractable gear systems must have adequate clearance from surrounding structure and systems as specified.

(2) Main, nose, and tail wheel tires must comply.

(3) Rotorcraft design maximum weight shall be used. Static and “dynamic” conditions are specified for rotorcraft tires.

(4) Tire performance standards are contained in TSO-C62c.

Page D - 56 9/30/99 AC 29-2C b. Procedures .

(1) The aircraft structural design loads should contain a maximum static load imposed on the tires. The load is derived for a static ground reaction assuming the design (maximum) weight and the critical center of gravity for each tire of the landing gear. The wheel loads are determined under § 29.731(b). Reduced weight but forward CG conditions may result in the highest static load on a nose wheel tire. Thus, combinations of weight and CG locations require investigation for the maximum tire load of each main, nose, and tail wheel tire. Nose wheel tires are subject to a specific dynamic condition.

(2) The maximum possible size of the tires considering appropriate temperatures, aging, and pressure should be obtained to check wheel well and cover clearances. Tire dimensions (for clearances) may be found in the yearbook noted in paragraph AC 29.733b(4). If the tire clearance is questionable, objects may be taped to the tire to simulate tire growth or oversize dimensions expected and the wheel retracted and rotated by hand to check for possible interferences. Minimum clearance, such as one-half inch, may be adequate as a design objective. The design drawings should be reviewed for information of correct systems installations and landing gear rigging within the wheel wells and wheel covers, if installed. If necessary to control tire sizes, specific manufacturer’s tires should be used as “required equipment” and the tire manufacturer and the part number should be specified in the design data and on the type certificate data sheet as “required equipment.” (3) As specified in § 29.729(d), an operation test of any retractable landing gear should be performed. During this operation test, the tire clearances should be determined and recorded for the maximum tire size expected in service. Only the least or minimal clearance found, if adequate, should be recorded.

(4) The Tire and Rim Association, Inc., generally issues a yearbook listing tire and wheel rim sizes and ratings. This information is advisory as stated in the yearbook.

Section 9 concerns aircraft tires and rims. Table AP-5 in Section 9 of the yearbook concerns tires used on rotorcraft. The tire may be selected initially from the yearbook, but qualification data for the specific tires used shall be furnished with the type design data in compliance with the standards. Section 9 also contains tire size and tire growth dimensions.

(5) Minimum performance standards for aircraft tires, excluding tail wheel tires are found in TSO-C62c, Aircraft Tires. Tires meeting the TSO are marked as prescribed in the standards. The load rating (reference § 29.733) is marked on the tire.

TSO tires are not required but should be used whenever possible. The manufacturer’s information, such as load rating, should be included in the aircraft type design structural substantiation data.

Page D - 57 AC 29-2C 9/30/99 AC 29.735. § 29.735 (Amendment 29-24) BRAKES .

a. Explanation .

(1) Brakes are required for wheel landing gear aircraft. Minimum performance standards are contained in this section. During the course of the FAA/AUTHORITY flight test program and of any F&R program conducted under § 21.35, the brakes shall be used and evaluated.

(2) Design criteria are contained in this standard.

(i) The braking device must be controllable by the pilot. It is optional for the second pilot station except as may be specified under the provisions of § 29.771.

(ii) The braking device must be usable during power-off landings.

(3) Performance criteria are also contained in this standard.

(i) The brakes must be adequate to counteract any normal unbalanced torque when starting or stopping the rotor or rotors.

(ii) The brakes must be adequate to hold the rotorcraft parked on a 10° slope on dry, smooth pavement.

(4) In §§ 29.493(b)(2) and 29.497(g)(2)(ii), limiting brake torque is one ground load standard for design of the landing gear.

(5) Although not specifically noted in a standard, the position of the brake on the wheel is important. The brake should be positioned to avoid ground contact whenever the tire is deflated.

(6) TSO-C26c contains minimum performance standards for aircraft landing wheels and wheel-brake assemblies. For rotorcraft, a wheel-brake assembly design rating is established by the manufacturer. The TSO standard for rotorcraft brakes specifies a 20° slope standard (rather than a 10° slope) for an over-pressure hydraulic brake test.

(7) The brake application device at the pilot station is subject to other structure strength standards in this Part, such as the limit pilot forces or torque specified in § 29.397.

b. Procedures .

(1) Wheel-brake assemblies approved under TSO-C26 or a later revision will have various (rotorcraft) ratings as specified in the standard. One rating of TSO standard for a rotorcraft wheel-brake assembly is the kinetic energy capacity in Page D - 58 9/30/99 AC 29-2C foot-pounds at the design landing rate of absorption. The design takeoff and landing weight and rotorcraft speed in knots for brake application are a part of the equation.

The brake manufacturer should furnish this rating and the two noted parameters for the selected design or designs. The ratings of selected brakes should be included in a structural design data report such as a design criteria report. The use or application of each brake design on the particular rotorcraft design should not exceed capacity of the brake or the ratings established under the TSO. If appropriate, the part number and manufacturer of each brake may be listed in the structural data reports as well as listed in the type design drawings.

(2) The limiting brake torque obtained from the brake manufacturer should be used in complying with §§ 29.493(b)(2) and 29.497(g)(2)(ii).

(3) Compliance with the brake standards should be confirmed, demonstrated, and recorded as a part of the flight test type inspection report. This applies to TSO brakes and to brakes approved as a part of the aircraft type design.

(4) If found necessary under the provisions of § 29.771, the second pilot station should have brake control devices. The brake control devices should be listed with the other required equipment that defines the equipment necessary for a second pilot station.

(5) A brake assembly may be evaluated and approved under Part 29 as a part of the aircraft type design. TSO-approved brakes are not specifically required but are recommended. For non-TSO-approved brakes, all detail and assembly drawings, required test proposals, and test results reports may be submitted and processed as a unique part of the particular aircraft type design.

(6) During an inspection of the landing gear, such as an engineering compliance inspection, the brake location should be checked to ensure the brake does not contact the ground when the tire is deflated. Type design drawings should control the proper location of the brake on the landing gear.

AC 29.737. § 29.737 SKIS .

a. Explanation . This standard is, in part, derived from small airplane standards.

Aircraft skis approved under TSO-C28 may be used on rotorcraft. TSO-C28 for aircraft skis refers to Sections 4 and 5 of National Aircraft Standards Specification 808, dated December 15, 1951, for strength and performance standards. The standard also addresses flight/aerodynamic loads.

(1) A maximum limit load rating is assigned to each ski approved under TSO-C28.

(2) This limit load rating must not be exceeded by the maximum limit ground load determined under the standards of § 29.505, Ski landing conditions.

Page D - 59 AC 29-2C 9/30/99 (3) The ski installation is also subject to the maximum aerodynamic and inertia loads and to the ground rotation or torque load per § 29.505(c).

(4) Ski mounting or installation parts used in the particular application are subject to substantiation as any landing gear member is subject to substantiation.

(5) Ski installations are also subject to flight and ground operation evaluations.

(6) Pads or “bear paws” on skid or wheel landing gears for use in snow or soft soils are unique to rotorcraft. These shall be approved also. For new type certificate applications after November 27, 1989, § 29.571, Amendment 29-28 requires fatigue substantiation of the landing gear. The effect of pads, etc., shall be evaluated in compliance with the standard.

b. Procedures .

(1) The limit load rating for the ski selected shall be obtained from the ski manufacturer. This information shall be included in the design criteria and/or structural substantiation reports. The type design drawings will include the appropriate part number for the TSO-approved product and the necessary installation information.

(2) The design limit loads derived in compliance with § 29.505 shall not exceed the ski limit load rating. The skis shall be substantiated for the torque load in § 29.505(c) since the TSO standard does not contain a similar requirement.

(3) Skis that are not TSO approved may be approved as a part of the aircraft type design by complying with the strength and performance standards contained in TSO-C28 (NAS 808).

(4) The aerodynamic loads shall be based on a limit load design speed of 1.11 V . The maximum V used in design may be reduced only for a “ski NE NE configuration” airspeed limitation.

(5) Pads or “bear paws” installed on skid or wheel landing gear to facilitate operations in snow conditions or marsh lands may be approved as a part of or as an alteration to the aircraft type design. Rational flight and landing design loads applicable to the particular pad design must be developed and strength substantiating data submitted proving compliance with the strength and performance standards contained in Part 29. In addition, skid landing gear may be subject to excessive vibratory loads while in flight whenever the weight and mass distribution is altered by adding “bear paws.” The effect of additional weight should be investigated over the flight operating regimes, including the approved range of rotor speeds. Resonant vibratory conditions should be avoided or highly damped, thus avoiding a potential change in service life. In compliance with § 29.571, Amendment 29-28, stress measurement, etc., may be necessary, if the standard is applicable.

Page D - 60 9/30/99 AC 29-2C SUBPART D - DESIGN AND CONSTRUCTION FLOATS AND HULLS .

AC 29.751. § 29.751 (Amendment 29-3) MAIN FLOAT BUOYANCY .

a. Explanation .

(1) This section specifies standards for single and multiple float buoyancy in fresh water. The standard does not apply to ditching/emergency flotation devices, but to amphibian rotorcraft devices.

(2) It is a design and a performance standard. Rigid or inflatable floats may be used. Enough water tight compartments (per Amendment 29-3) rather than a specific number are required to minimize the probability of capsizing when one compartment is flooded or deflated.

b. Procedures .

(1) Excess buoyancy . A minimum of 50 or 60 percent in excess of the maximum certificated weight of the rotorcraft is required for single or multiple floats, respectively. The weight of fresh water (density 62.42 pounds per cu. ft.) displaced by fully submerged float or floats (total volume of each float at operating pressure is used) should be a minimum of 50 or 60 percent greater than the maximum certificated weight of the rotorcraft.

(2) Capsizing .

(i) Each float should have enough sealed, separate and approximately equal volume compartments to minimize the probability of capsizing when the critical compartment is flooded or deflated. Five or more compartments in each float are usually necessary to meet the standard. Ten compartments per float have been employed in certain designs.

(ii) An analysis or test or combination thereof may be used, if necessary, to prove a positive margin of stability with the most “critical” compartment in one float flooded or deflated.

(iii) The location of the floats, and the most critical compartment, the rotorcraft weight, mass moment of inertia, and center of gravity location are also important considerations for capsize stability.

AC 29.753. § 29.753 MAIN FLOAT DESIGN .

a. Explanation.

Page D - 61 AC 29-2C 9/30/99 (1) Strength or design load standards are encompassed in the standard for inflatable bag and rigid floats. Bag pressure loads are included. The standard applies to an amphibious rotorcraft.

(2) The float landing loads are derived from the drop test of the float landing gear, or the load may be derived from tests of the wheel (or skid) landing gear (reference § 29.521). Bag type floats are not subject to the side loads according to the standard. Rigid floats, whether single or dual, are subject to the side load in each direction.

(3) Inflatable bag type floats should also be designed for the maximum pressure differential developed for the maximum operating altitude difference requested. That is, the resulting pressure difference between an operational altitude and a take-off site elevation should be established, and proven and may become an operating limitation.

(4) Landing loads suffice for the aerodynamic loads for typical rotorcraft float designs. Nonetheless, design and/or support of the forward part of bag type floats should be evaluated for maximum design speeds to prevent collapse or significant distortion of the bag while in flight.

(5) Resistance to puncture and abrasion at attach/wear points is not in the standard but is an important design consideration. “Girt” or attachment design loads shall be sufficient to withstand the loads imposed by the standards.

(6) The water or sea conditions (wave heights) evaluated in §§ 29.231 and 29.239 tests are not limitations but should be noted in the procedures section of the flight manual.

(7) The standard does not apply to ditching/emergency floatation devices.

b. Procedures .

(1) Landing load factor .

(i) A drop test of the float landing gear may be conducted to obtain the limit landing load factor (reference § 29.725). Level landing attitude should be used for the float assembly.

(ii) The limit load factor for wheel or skid landing gear may be used (reference § 29.521) for the floats.

(iii) The float design ultimate load factor is 1.5 multiplied by the limit load factor.

Page D - 62 9/30/99 AC 29-2C (2) Flight aerodynamic loads--bag type floats .

(i) Evaluate collapse or significant distortion of bag type floats for speeds up to V (1.11 V ) with the minimum operating bag pressure.

D NE (ii) External tubes to support the bag may be employed.

NOTE: Design landing loads may exceed the flight loads.

(3) Altitude differential loads .

(i) Bag type floats should not rupture due to the change in absolute pressure from take-off to the operating altitude. The applicant should select and prove the maximum operating altitude differential desired. A 5,000 to 8,000 feet operating differential may be a sufficient limitation. That is the rotorcraft with bags properly inflated could not operate more than 5,000 to 8,000 feet above the take-off site elevation.

(ii) A proof and ultimate pressure test should be conducted for the design.

If operating or inflation pressure is 2.62 PSI (including a tolerance) and 5,000 feet (pressure) differential is desired (use sea level to 5,000 feet pressures), the proof or limit pressure should be 2.62 + 2.47 = 5.09 PSI. The pressure relief valves may be set at this value also. The change in size during inflation should be recorded. Significant changes may adversely affect flight characteristics and should be evaluated. The ultimate or burst free pressure should be proof pressure (5.09 PSI) multiplied by 1.5 or 7.635 PSI. A video or photographic record may be used as a reference of the change in size or shape for this test.

(iii) Each compartment should be equipped with a pressure relief valve to further protect the bag from excessive internal pressure.

(iv) At least one float should be subjected to a burst pressure test.

Record the gauge pressure of burst.

(4) Landing loads .

(i) Rigid float vertical and a combined vertical and aft load conditions. A vertical or up-load only and a vertical combined with an aft load component for a resulting vector angle of 14.03° from the vertical axis of the rotorcraft shall be used.

Reference § 29.521(a). The resulting design load is the same load in both cases.

(ii) Rigid float side and vertical load condition . For each rigid float, whether single or dual, a vertical load combined with a side load resulting in a vector angle of 26.6° from the vertical axis of the rotorcraft shall be used. The side load is applied to each float individually. Both inward and outward acting side load conditions shall be substantiated separately for the design of dual floats.

Page D - 63 AC 29-2C 9/30/99 (iii) Load distribution on rigid floats shall be appropriate for the critical conditions. ANC-3 or § 25.533 and FAR Part 25, Appendix B may be useful.

(iv) Bag type float . The loads and the distribution of the loads are rather simple according to the standards. Only vertical loads and vertical with aft (drag) component are specified in the standard. These shall be distributed along the length over 75 percent of the projected area of the bag. Side loads are not required.

(5) Operating limitations .

(i) V with floats installed is typically lower than the V for wheel or skid NE NE landing gear rotorcraft configurations.

(ii) Bag inflation pressure shall be placarded or stenciled near inflation fittings.

(iii) The operating attitude differential proven for bag floats shall be an operating limitation. In addition, the flight manual should caution pilots about the effect of a significant decrease in altitude from the take off level which causes or reduces pressure in the bag. Placards may be employed as well.

(iv) Flight test results may dictate a further reduction in V or changes in NE other operating limitations.

AC 29.755. § 29.755 (Amendment 29-30) HULL BUOYANCY .

a. Explanation .

(1) This section contains performance standards for an integral fuselage hull and auxiliary (such as outrigger) floats. Water-based, amphibian and limited amphibian rotorcraft were encompassed in the standard.

(2) Amendment 29-3 added but Amendment 29-30 removed paragraph (b) which concerned Limited Amphibian Rotorcraft. Rotorcraft of that type used a “boat type hull” which is not desirable now and are certificated to the standards of § 29.801, Ditching, and § 29.563, Structural ditching provisions. (Limited amphibian rotorcraft were converted to the ditching configuration.)

(3) The worst combination of wave height and surface winds selected by the applicant shall be used in compliance with the standard.

b. Procedures .

(1) Capsizing .

Page D - 64 9/30/99 AC 29-2C (i) The hull and auxiliary floats shall have enough sealed compartments to allow failure of the critical, single, compartment in either the hull or auxiliary float and minimize the probability of capsizing.

(ii) Location of the most critical compartment (whether hull, sponson, or auxiliary), rotorcraft weight, mass moment of inertia, and CG location are also important considerations to prove stability or not capsizing.

(iii) The lightweight rotorcraft configuration and wind and wave condition should be considered, as well as the heavy weight configuration.

(iv) The sea state (worst combination of wave height and surface winds) is selected by the applicant. The condition proven is included in the procedures or information section of the flight manual. (It is not an operating limitation.)

(2) Buoyancy .

(i) Excess buoyancy is necessary to comply with the standard but the amount is dependent on several factors, such as number, size, and location of the sealed, watertight, compartments.

(ii) Wheel tires may be used for buoyancy if appropriate to the design.

(iii) Fuel tanks, if properly located and protected from potential rupture and if the aircraft has a system to rapidly empty the tanks, may be used also for buoyancy.

(iv) Buoyancy may be determined using the displacement of fresh water, with 62.42 pounds per cubic ft. density.

(3) Tests .

(i) If necessary, scale models may be used to prove the stability of the rotorcraft design for the sea state and wind conditions selected by the applicant.

(ii) The rotorcraft is subject to water tests per § 29.231. Compliance with part of this standard may be demonstrated or proven for the sea state or wave height, and wind conditions selected in conjunction with the TIA flight test program. This information is not an operating limitation.

(iii) Proposals should be submitted for evaluation.

AC 29.757. § 29.757 (Amendment 29-3) HULL AND AUXILIARY FLOAT STRENGTH .

a. Explanation . The standard is an objective or performance strength standard.

The water loads in § 29.519 shall be imposed for the hull and auxiliary floats in a Page D - 65 AC 29-2C 9/30/99 conservative manner. The hull and float are “rigid” conventional amphibian or water-based aircraft structures.

b. Procedures .

(1) The water loads and conditions specified in § 29.519 shall be used. The pressures or load distributions should be appropriate to the design. ANC-3 and §§ 25.523 through 25.535 and Appendix B to FAR Part 25 may be of use.

(2) The water loads and applications of the loads are objective standards. A proposal and early discussions in the life of a project should be used to agree on an appropriate avenue or means of compliance. Tests or analysis supported by tests may be appropriate.

Page D - 66 9/30/99 AC 29-2C SUBPART D - DESIGN AND CONSTRUCTION PERSONNEL AND CARGO ACCOMMODATIONS AC 29.771. § 29.771 (Amendment 29-24) PILOT COMPARTMENT .

a. Explanation .

(1) Volumes have been written on human factors and their contribution to pilot workload and fatigue. This document cannot begin to address the myriad of considerations involved in pilot compartment design. The intent of the rule is simply to ensure that reasonable human factor engineering practices have been followed.

Equipment should be logically grouped within the pilot’s reach and view and be easy to operate. Seats should provide a reasonable level of comfort for the normal anthropometric range of pilots for a typical mission duration. Environmental considerations such as radiation from the sun through overhead windows should be addressed. Heating, cooling, and ventilation systems should be adequate for the range of expected operating conditions.

(2) Each pilot compartment and its equipment should allow the minimum flightcrew to perform their duties without unreasonable concentration or fatigue. If there is a provision/requirement for a second pilot, his station should be equipped with primary flight controls and have easy access to powerplant controls. Duplicate wheel brakes are recommended. Duplication of miscellaneous controls such as idle detent switches, RPM beep functions, nosewheel locks, and parking brakes has not been required. The need for duplicate instruments for the second pilot tends to be a function of cockpit size and panel configuration.

(3) Webster defines appurtenances as “accessory objects or apparatus.” Items such as blowers, fans, and gyros should not have noise or vibration characteristics which could contribute to pilot fatigue or distraction. Instrument panel vibration is specifically addressed in § 29.1321.

(4) Although the rule prohibits in-flight rain or snow leaks that distract the crew or harm the structure, leaks occurring on the ground should also meet these requirements. In extreme cases where an offensive leak could not be stopped, the moisture has been rerouted to a noncritical area. In the context of this rule, “structure” is interpreted to include any part of the pilot compartment to include systems and equipment.

b. Procedures .

(1) Initial evaluation of the pilot compartment should be conducted on the ground. However, the cockpit assessment should be an ongoing effort throughout the flight test program. If a second pilot position is provided/required, the adequacy of Page D - 67 0 7 /0 2 /18 AC 29 - 2C, Chg 8 controls and instruments should be evaluated under all normally expected operating conditions. If a second pilot position is not provided/required, any passenger position in the pilot compartment should be evaluated to ensure that a passenger, properly br iefed by the flightcrew, can sit comfortably without inadvertent interference with normal control operations. All equipment should be operated during at least one flight of typical mission profile and duration.

(2) Although many pilot compartment rain or snow leaks can be located on the ground by dousing the aircraft with a hose, in - flight leaks often occur in varying intensity and in different locations. Flight in rain should therefore be included during flight test.

AC 29.773. § 29.773 (Amendment 29 - 56 ) PILOT COMPARTMENT VIEW.

a. Nonprecipitation Conditions.

(1) Explanation. Amendment 29 - 56 changed 29.773(a)(2) by adding the option to perform a ground test in lieu of a flight test. Paragraph (a)(1)(iii) below contains guidance to reflect this change.

(i) The procedure s paragraph following this explanation discusses one means of demonstrating an adequate field of view.

(ii) Since glare and reflection often differ with the sun’s inclination, consideration should be given to evaluating the cockpit at midday and in early mornin g or late afternoon. Windshields with embedded wire heating elements should be evaluated for distortion with the system both “ON” and “OFF.” (iii) If night approval is requested then compliance with 29.773 may be shown by either a ground or flight test. Internal lighting can be evaluated with a ground test. External lighting modifications may require a flight test. If a flight test is conducted, then the lighting system should be evaluated in likely combinations and unde r expected flight condition s.

(A) Internal lighting evaluations may be accomplished by a ground test with the aircraft in a dark hangar or with external light blocked to create a dark cockpit.

Locations and intensity of instrument panel and control panel light reflections should be assessed to ensure they do not interfere with the pilots’ ability to see outside the aircraft. Figure ac29.773 - 1 contains guidance on the areas in pilots’ field of vision where reflections should be eliminated. Although a certain amount of equipment reflec tion (avionics control heads, etc.) In the windshield may be unavoidable, the pilot’s normal field of view should be unobstructed. If reflections are mitigated by modifying the glare shield, then the pilots’ field of view should be revaluated. If a reflect ion’s effect on the pilots’ ability to see outside the aircraft is uncertain, a night flight test may be required .

(B) External lighting evaluations should be don e during night flight tests. Evaluating external lights on the ground may result in unacceptable light entering the cockpit and causing reflections or glare due to reflection off the ground or surrounding structures. Landing and taxi lights should be exercised throughout their adjustment range (if applicable) to check for reflections, particularly in chin windows.

Anticollision and strobe lights should be evaluated to ensure that frequency interaction Page D-68.1 05/04/18 AC 29 - 2C , Chg 8 and reflections off the rotor do not result in distractions to the pilot. The effect of cabin lighting on the pilot compartment view should be assessed, particularly on EMS - configured aircraft where the in - flight use of cabin lights may be mandatory (2) Procedures. The following procedures are one acceptable means of evaluating pilot compartment field of view considering only those objects in the pilot compart ment, the windshield, and its support structure in nonprecipitating conditions.

The applicant’s design is not required to meet these guidelines, and each design should be evaluated on its own merits. The area of visibility established in the following para graphs will provide an acceptable level of visibility for a minimum crew of one (pilot). In the event that a minimum crew of two (pilot and copilot) is required, the Page D-68.2 9/30/99 AC 29-2C second pilot should have an area of visibility equivalent to that provided for the pilot but on the opposite side. In this event, the pilot’s area of visibility to the left as shown in figure AC 29.733-1 needs only to comply to 60° left, and the area of visibility for the second pilot needs only to comply to 60° right.

(i) A single point established in accordance with the provisions of this paragraph constitutes the referenced eye position (i.e., a point midway between the two eyes) from which the central axis may be located. The referenced eye position is a reference datum point based on the eye location that permits the specified vision envelope required by figure AC 29.733-1, allows for posture slouch, and is the datum point from which the aircrew station geometry is constructed. The referenced eye position should be located by means of ship’s coordinates that contain station reference number, water line, and butt line for both pilot and copilot, if applicable, and complies with: (A) The pilot’s seat in a normal operating position from which all controls can be utilized to their full travel, by an average subject, and which should provide for vertical adjustment of the seat of not less than 2.5 inches above and 2.5 inches below this initial vertical position.

(B) The seat back in its most upright position.

(C) The seat cushion depression being that caused by a subject weighing 170 to 200 pounds.

(D) The longitudinal axis of the rotorcraft to be that of “cruise attitude” (0.9V or 0.9 V whichever is lower).

H NE (E) The point established not beyond 1 inch to the right or left of the longitudinal centerline of the pilot’s seat.

(F) All measurements made from the single point established in accordance with this paragraph.

(ii) A dual lens camera, as photo recorder, should be used in measuring the angles specified in the paragraphs listed below. Other methods, including the use of a goniometer, are acceptable if they produce equivalent areas to those obtained with a dual lens camera. When not using a dual lens camera, compensation should be made for one-half the distance which exists between the eyes, or 1¼ inches. With the referenced eye position located as indicated in paragraph AC 29.733a(2)(i), and utilizing binocular vision and azimuthal movement of the head and eyes about a radius, the center of which is 3 and 5/16 inches behind the referenced position (this point to be known as the central axis), the pilot should have the following minimum areas of vision measured from the appropriate eye position. (See figure AC 29.733-1.)

(A) 20° forward and above the horizon between 0° and 100° left.

Page D - 69 AC 29-2C 9/30/99 (B) 20° forward and below the horizon between 10° and 100° left.

(C) 20° forward and below the horizon at 10° left increasing to a point 30° forward and below the horizon at 10° right.

(D) 50° forward and below the horizon between 10° right and 135° right.

(E) 20° forward and above the horizon at 0° increasing to a point 40° above the horizon at 80° right and 100° right and then decreasing to a point 20° forward and above the horizon at 135° right.

(iii) Any vertical obstruction which falls within the minimum area of visibility outlined in paragraph AC 29.733a(2)(ii) should be governed by the following: (A) No vertical obstruction between 20° right and 20° left.

(B) Between 20° right and 135° right, vertical obstruction should not exceed 2.5 inches in width.

(C) Between 20° left and 100° left no vertical obstruction greater than 2.5 inches in width.

(iv) Any horizontal obstruction which falls within the minimum area of visibility outlined in paragraph AC 29.733a(2)(ii) should be governed by the following: (A) The area 15° forward and above the horizon between 135° right and 40° left decreasing to a point 10° above the horizon at 100° left, and 15° forward and below the horizon between 135° right and 100° left should be free from horizontal obstructions.

(B) The area above and below the horizon which is between the minimum area of vision specified in paragraph AC 29.733a(2)(ii) and paragraph AC 29.733a(2)(iv)(A) is limited to one horizontal obstruction above the horizon, and one below the horizon. These horizontal obstructions should not be greater than 4 inches in width. An overhead window which will provide twice as much additional visibility as was lost due to the obstruction, should be located immediately above any obstruction which is above the horizon. This requirement is in addition to any area of visibility specified by paragraph AC 29.733a(2)(ii) which may be included in the overhead window area.

(C) If the instrument panel obstructs any required area between 10° left and 10° right below 20° forward and below the horizon, a window which affords triple equivalent additional visibility should be located immediately below and between the angles of 20° left and 20° right above 65° below the horizon.

Page D - 70 9/30/99 AC 29-2C (v) For steep rejected takeoffs and steep approaches such as used for oil rigs or confined heliports, the visibility should be such that the pilot can see the touchdown pad and sufficient additional area to the side and forward to provide both an accurate approach to the touchdown point as well as a satisfactory degree of depth perception. A 5-inch head movement, by the pilot, forward and/or sideward of the normal position is acceptable in determining compliance.

b. Precipitation Conditions .

(1) Explanation .

(i) Heavy rainfall is defined by the National Weather Service as one resulting in accumulation in excess of 0.03 inches in 6 minutes. On past designs, the windshield wipers required by § 29.1307 have been adequate to ensure satisfactory view at low to medium airspeeds. Airflow over the windshield and/or wipers has normally been sufficient to keep the windshield clear at higher airspeeds. Obscuration of side windows by rainfall should be addressed, particularly for confined area approaches.

(ii) If icing certification is requested, a means must be provided to ensure that a sufficiently large viewing area is kept clear of ice to permit safe operation. As a minimum, a clear area on the windshield should be available, although some configurations could require clear view in other areas, in order to provide an adequate level of safety in certain operations.

(iii) An openable “clear view” window must be provided for the first pilot.

The rule requires that the window be openable in heavy rain at forward speeds to V H and in the worst icing conditions requested for certification. The rule further requires a field of view through this opening which is adequate for safe operation. Although the rule implies that a safe field of view must be provided for airspeeds up to V , it has not H been interpreted as such. In most designs, the only practical location for an openable window is in a side panel or door. Aircraft sideslip limits normally restrict useful view from this window opening at high airspeeds. The intent is to provide the pilot with an adequate view for safe approach and landing in the event that normal windshield clearing systems malfunction.

(2) Procedures . Compliance with the requirements of this rule should be checked by flying the aircraft in the applicable environmental conditions. Although wipers can be partially checked on the ground with a hose, their effectiveness at higher airspeeds should also be verified. Likewise, additional or alternate rain removal systems should be exercised throughout the required airspeed range. The need for windshield wash systems should be assessed, particularly if the aircraft will be used in an offshore salt spray environment. Systems provided to ensure clear view in icing conditions should be evaluated during icing flight tests. The location and effectiveness of the openable window should be evaluated following failure of the rain removal and anti-ice system (if applicable). The view through the window opening should permit safe Page D - 71 AC 29-2C 9/30/99 operation from hover up to a reasonable approach airspeed. Care should be exercised during flight test to stay within airframe sideslip limits.

Page D - 72 9/30/99 AC 29-2C o· 20° 20° 40° so· 40° 120° 100° WIDTH) IN WIDTH SEAT) ONE CONSIDERED IN so· TO INCHES OBSTRUCTION INCHES LIMITED OBSTRUCTIONS RIGHT-HAND 2.

THAN COMMAND so· TO IN AREA VISIBILITY OCCUPY VERTICAL THIS HORIZONTAL PILOT GREATER TO LIMITED o• 135° 4 (NOT (NOTE: COCKPIT - f0" __..

10" ONS 29.773-1 o• VERT.

AC NO 10" BSTRUCTI

i

FIGURE 40° WIDTH IN so· INCHES OBSTRUCTIONS 2.5 TO so· VERTICAL LIMITED Page D - 73 AC 29-2C 9/30/99 AC 29.775. § 29.775 WINDSHIELDS AND WINDOWS .

a. Explanation. Nonsplintering safety glass is specified in windshields and windows containing glass to protect crew and passengers if window fracturing occurs.

In any case, windshields and windows are to be made of transparent materials which will not break into dangerous fragments.

b. Procedures .

(1) Use nonsplintering safety glass in windshield or window applications which contain glass rather than plastic acrylics, polycarbonates, epoxies, etc. The glass selected should meet a specification such as MIL-G-25871, and if new vendors are selected by an airframe manufacturer, test data should be obtained from the vendor to demonstrate the safety glass provided meets an acceptable specification and provides adequate nonsplintering capability.

(2) Windshields and windows should be designed so that either - - (i) They are made of material which will not cause a serious reduction in the field of view by becoming suddenly opaque; or (ii) Any one panel becoming opaque will not cause a serious reduction in the field of view (reference § 29.773).

(3) In the event of any reasonably probable failure, a transparency heating system must be incapable of raising the temperature of any windshield or window to a point where there would be a danger of fire or structural failure (reference § 29.1309).

AC 29.775A. § 29.775 (Amendment 29-31) WINDSHIELDS AND WINDOWS .

a. Explanation . Amendment 29-31 changed § 29.775 to allow the use of material other than nonsplintering safety glass; i.e., plastics are allowed. Additionally, whatever material is used should not break into dangerous fragments upon impact.

b. Procedures . The procedures contained in paragraph AC 29.775 apply equally to glass or plastics.

AC 29.775B. § 29.775 (Amendment 29-40) WINDSHIELDS AND WINDOWS .

a. Explanation . Amendment 29-40 added § 29.631 which requires the rotorcraft be designed to ensure capability of continued safe flight and landing (Category A) or safe landing (Category B) after impact with a 2.2 lb (1.0 kg) bird when the velocity of the rotorcraft (relative to the bird along the flight path of the rotorcraft) is equal to V or V NE H (whichever is lesser) at altitudes up to 8,000 feet.

Page D - 74 4/25/06 AC 29-2C, Chg 2 b. Procedures . In addition to the procedures outlined above, compliance with § 29.631, should be demonstrated by tests or analysis supported by test evidence that the windshield will withstand, without penetration, the impact with a 2.2 lb (1.0 kg. bird) at V or V (whichever is lesser) at altitudes up to 8,000 feet. See NE H paragraph AC 29.631 (§ 29.631) for additional information.

AC 29.777. § 29.777 COCKPIT CONTROLS .

a. Explanation . This section defines the general cockpit control requirements.

Cockpit control location and arrangement, with respect to the pilot’s seat, must be designed to accommodate pilots from 5’2” to 6’0” in height. Pilots within this range should be able to reach and operate all required controls and have sufficient clearance with the structure, panels, etc.

b. Procedures .

(1) The applicant should have a cockpit design report which documents the anthropometric suitability of the cockpit. Subsequent cockpit evaluations of control movement and location should be conducted with adjustable seats and/or controls positioned in a flight position for the subject pilot. Essential controls should be evaluated with the shoulder harness locked in the retracted position. Evaluation pilots should be aware of their individual anthropometric measurements and temper their assessments based on this information. Ideally, a new design should include evaluations by a range of different sized subject pilots. Control considerations for a second pilot position are the same as for the pilot station. Paragraph AC 29.771 discusses current philosophy concerning duplication of controls.

(2) As background, the following are examples of cockpit control issues which should be avoided: (i) Collective control blocking the lateral movement of a pilot’s leg, which in turn restricts the left lateral cyclic displacement.

(ii) Seat or seat cushion impeding the aft cyclic movement.

(iii) Inadequate space for large feet equipped with large flight boots.

(iv) Control/seat relationship which requires unusual pilot contortions at extreme control displacements.

(v) Control/seat relationship or control system geometry which will not permit adequate mechanical advantage with unboosted controls or in a boost OFF situation.

(vi) Addition of control panels or equipment to instrument panels or consoles which restrict full control throw.

Page D - 75 AC 29-2C 9/30/99 (vii) Brake pedal geometry which results in inadvertent brake application upon displacement of the directional controls.

(viii) Controls for accessories or equipment which require a two-handed operation.

(ix) Emergency external cargo release controls which cannot be activated without releasing the primary flight controls.

(x) Essential controls which cannot be actuated during emergency conditions with the shoulder harness locked.

(xi) Throttle controls which can be inadvertently moved through idle to the cutoff position.

(xii) Switches, buttons, or other controls which can be inadvertently activated during routine cockpit activity including cockpit entry.

(xiii) Failure to account for operation with the pilot wearing bulky winter clothing.

(xiv) Aft cyclic movement limited by the pilot’s body with a fore and aft adjustable seat in the full forward position.

AC 29.779. § 29.779 (Amendment 29-24) MOTION AND EFFECT OF COCKPIT CONTROLS .

a. Explanation . The section standardizes motion and effect of cockpit controls.

While this paragraph specifically addresses primary flight controls, engine power controls, and landing gear controls, it applies to all cockpit controls not addressed in other paragraphs.

b. Procedures .

(1) The cyclic should be mechanized such that movement of the control results in a corresponding sense of aircraft motion in the same axis. While a certain amount of coupling may be present following a pure control input in a given axis, that coupling should not be objectionable to the pilot. Collective pitch control should be mechanized such that an upward movement of the collective results in a corresponding relative motion of the aircraft in the vertical plane. Again, coupling should not be objectionable.

Care should be taken to ensure that the primary pilot perception of collective motion is in the vertical plane. The objective is to clearly differentiate collective motion from that associated with an airplane throttle. The rule is self-explanatory on the subject of engine power controls. A distinction is made between normal landing gear controls an Page D - 76 9/30/99 AC 29-2C emergency controls. Emergency controls may operate in a sense which might not correspond to the direction of resultant gear motion.

(2) The recommended operating convention and “switchology” for miscellaneous controls is: (i) Up/forward = on/increase.

(ii) Down/aft = off/decrease.

(iii) Variable rotary controls should move clockwise from the OFF position, through an increasing range, to the full ON position. For some variable intensity controls such as instrument lighting, the desired minimum setting may not be completely off. Pushbuttons not giving an obvious indication of mechanical position should be configured such that the flightcrew has a clear indication of switch actuation under both day and night (if applicable) conditions. Failure of the indication should be shown to be free of hazards.

(3) Slew or “beep” switches associated with flight control system applications warrant special attention. The recommended conventions for control-mounted single, or multifunction, two or four-way “beep” switches are: (i) Cyclic.

Flight Control System Switch Direction /Autopilot Configuration Aircraft Response Forward/up basic trim nose down airspeed/groundspeed increased airspeed forward mode selected speed reference vertical speed mode increased rate of selected (without descent/decreased rate of airspeed mode engaged) climb hover mode selected increased ground-speed or forward acceleration reference Left basic trim left wing down heading mode selected slew heading reference left hover mode selected increased ground-speed or acceleration reference to left Page D - 77 AC 29-2C 9/30/99 (ii) Collective (assumes switch is mounted on top of grip).

Flight Control System Switch Direction /Autopilot Configuration Aircraft Response Forward control position hold down collective vertical speed mode increased rate of selected descent/decreased rate of climb hover mode selected decreased hover height reference Left control position hold increase left pedal hover mode selected slew heading reference left (iii) Opinions are divided concerning the preferred convention for forward and rearward motion of slew switches mounted atop the collective grip. Part of the reason appears to stem from the fact that such a switch is never used in a purely control position trim capacity. The switch has normally remained nonfunctional until a vertical autopilot mode is selected. At that point, the switch is viewed by one pilot/engineer contingent as either an autopilot reference slew function or a power increase/decrease switch, which should follow the “forward equals increase” convention. The other group views the switch as a form of control position trim and finds the “forward equals down collective” convention to be more consistent with the sensing used for the cyclic beep switches. An obvious solution is to mount collective/vertical axis switches in a vertical orientation on the grip. Barring that alternative, viable arguments can be made for either philosophy. The recommended convention was selected following a survey of manufacturers and test pilots.

AC 29.783. § 29.783 (Amendment 29-20) DOORS .

a. Explanation . This regulation requires at least one door for all closed cabin rotorcraft. Standards for all doors and airstair doors are included. To assure that the doors provide normal entry and egress without causing or contributing to hazardous conditions, even after a minor crash, the following requirements are imposed: (1) Passenger doors may not be located with respect to any rotor to endanger persons using the doors as instructed.

Page D - 78 9/30/2008 AC 29-2C, Chg 3 (2) Means are required for locking crew and external passenger doors to prevent their opening in flight due- - (i) To inadvertent operation; or (ii) To mechanical failure.

(3) External doors are required to be openable from the inside or outside by simple and obvious means.

(4) Reasonable provisions to prevent jamming of external doors are required as specified and to assure that an “airstair door” is useable.

(5) The following visual indications of external doors being closed and locked are required: (i) Direct visual inspection means by crewmembers of the locking mechanism of all external doors.

(ii) Visual means to signal to crewmembers “when normally used external doors are closed and fully locked.” (6) For certain outward opening doors, an auxiliary safety latching device is required “to prevent the door from opening when the primary latching mechanism fails.” Suitable operating procedures to prevent this device from being used during takeoff and landing are required if the door cannot be opened from outside the rotorcraft (reference § 29.783(c)) with the device in place.

(7) If the door is a sliding door and intended to be opened and closed in flight, the sliding mechanism should positively attach the door to the airframe (e.g., sliding hinge) to minimize the likelihood of the door departing the aircraft in flight. Appropriate flight limitations should also be established to minimize any hazard while operating the door.

b. Procedures .

(1) Passenger doors should be located as far as possible from the auxiliary rotors. The doors may be hinged and door open stops may be provided to separate entering and egressing passengers from the auxiliary rotor blades. If necessary for the design, “appropriate instructions” should be provided for all passenger doors concerning entering and leaving the rotorcraft and safe use of each door relative to all rotors.

These instructions should be obvious to a passenger using the door, contain large enough letters to be readily legible, and use letters or background colors associated with danger (i.e. orange or red).

(2) Means to prevent the opening of doors in flight.

Page D - 79 AC 29-2C, Chg 3 9/30/2008 (i) Means to prevent the opening of doors in flight due to inadvertent operation may be provided by recessing door handles to prevent their inadvertent operation by the normal movement of passengers about the cabin. If recessing the door handle is impractical, a cover may be provided which will prevent inadvertent operation of the handle, but the cover should be of such design that it does not obscure the door handle or its operating instructions. It must not unduly interfere with deliberate operation of the door handle by passenger or crew. Transparent or nonsolid covers, easily displaced by deliberate actions, have been used to prevent inadvertent door handle operation. Some rotorcraft designs meet this requirement by requiring that passengers wear their seat belts at all times during flight. This design requires that the “fasten seat belt” sign be on at all times the rotorcraft is in flight (for practical purposes, the “fasten seat belt” light is generally designed to be on when power is applied to the rotorcraft).

(ii) Means to prevent inadvertent door opening in flight due to “mechanical failure” is most efficiently provided by multiple door latches and multiple load path door locking mechanisms so that the door will remain locked after a single failure. Care should be taken in the design of multiple load path latches and mechanisms to assure independence of all failures and to consider the effort of deflections after failures (if a failure allows deflections into the airstream sufficient to increase aerodynamic loads, the increase in loads should be accounted for; if a failure allows significant movement of latching components, the deflections should be accurately accounted for to assure that disengagement of nonfailed latches does not occur).

(3) The means to open normally used external doors is required to be simple (such as a rotating handle) and to be accessible from the inside or the outside. To prevent the inadvertent use of emergency exits (separate from normal entry doors) for routine entry and exit with the resulting “wear and tear,” the normally used doors for entry and exits should be equipped with operating handles and instructions distinctly different from those of the emergency exits. Obviously, the above does not apply to normally used exits which are also the primary (or only) emergency exits.

(4) Reasonable provisions to prevent jamming of external doors include the following: (i) Design features of doors which are insensitive to large fuselage deflections for door operation.

(ii) Provision of clearance between door and door frame latching devices sufficient to allow some relative deflection between the door and door frame and still allow door operation. The relative deflections may be determined by static test or by an analysis approved by the FAA/AUTHORITY.

(iii) Sliding doors are frequently used in transport rotorcraft for versatility and utility reasons. If sliding doors are used, one of the following features of design may be required to assure that the requirements of § 29.783(d) are met: Page D - 80 9/30/2008 AC 29-2C, Chg 3 (A) The sliding door(s) must be provided with jettison features which allow release of the door(s) from the tracks (to preclude jamming). The emergency release is generally separate and distinct from the normal door handle.

(B) Separate emergency exits of appropriate size and number may be installed in the sliding door(s).

(C) Separate emergency exits of appropriate size and number may be installed in addition to the sliding door(s).

(iv) Whether or not the sliding door is qualified as an emergency exit , it must meet the remaining door design standards.

(5) Direct visual inspection means by crewmembers of the locking mechanism of external doors may provide for visual observation of the door frame and the latching components for engagement or for visual observation of “flag” areas of the locking mechanism. If “flag” areas are used (such as tabs or shoulders which protrude into the crewmember’s line of sight when the latches are engaged (locked)), care should be taken to assure that the tab is permanently affixed (or an integral part) to the locking mechanism; and it should not give erroneous readings to the crewmembers under any foreseeable operation or failure of the latching mechanism. “Visual means to signal” to crewmembers “when normally used external doors are closed and fully locked” may be provided by annunciator panel lights or equivalent means. The visual indicating system may consist of an indicator for each individual door, or a system connecting all doors in series. If the latter system is used, it need not necessarily show which door is not fully locked. It is not necessary that more than one crewmember be able to ascertain by a visual signal that all external doors normally used by the crew in supplying the rotorcraft, or in loading and unloading passengers and cargo, are fully closed and locked. The visual signal should be located so that it may easily be seen by the appropriate crewmember from his station.

(6) For § 29.783(f), the auxiliary safety latching device to “prevent the door from opening when the primary latching mechanism fails” can be provided by the same multiple load path features which meet the § 29.783(c) requirement for prevention of door opening in flight after a “mechanical failure.” If a completely separate “auxiliary safety latching device” is used, it should allow the door to be opened from the inside, or outside, when in place. If the device must be removed to allow use of the door, “suitable operating procedures” (i.e., placards and RFM instructions) will be required for removal of the device during takeoff and landing.

(7) Additional standards for “airstair doors” were added by Amendment 29-20.

(i) An analysis or test may be used to prove compliance with deformation standards in § 29.783(g)(1).

Page D - 81 AC 29-2C, Chg 3 9/30/2008 (ii) A sketch, drawing, or demonstration may be used to prove the door is useable for the conditions described in § 29.783(g)(2).

AC 29.783A. § 29.783 (Amendment 29-31) DOORS .

a. Explanation .

(1) Amendment 29-30 extends the requirements of § 29.783 to: - include each external door, not just passenger doors; and, - require provision of door location and/or door operation procedures to protect persons from danger from propellers, engine intakes, and engine exhausts. (Protection from rotors are already included in the standard.)

(2) Amendment 29-31 adds a new paragraph (h) to § 29.783 which requires for doors used for ditching egress to have a means to secure the “ditching exits” in an open position and remain securely open in the appropriate Sea State used for compliance with § 29.801, paragraph AC 29.801.

b. Procedures . The procedures of paragraph AC 29.783 continue to apply to § 29.783 (and Amendment 29-31) with the following additions: (1) Occupants of the rotorcraft and servicing personnel are now required to be protected from injury when using any external door to enter or egress the rotorcraft and when loading cargo or servicing the rotorcraft. Consideration should be given to door location and/or operating procedures to include protection from propellers (if equipped) and engine inlets and exhausts, as well as from rotors.

(2) These new standards clarify that engine exhausts, engine inlets, and propellers, as well as rotors, are potentially hazardous and should be located or designed to protect rotorcraft occupants and ground personnel or use door latching and operating procedures to protect those persons. Operating procedures for the door, including readily visible markings, should be provided to minimize injury to personnel when practical component locations or component design features, alone, do not assure possible freedom from injury.

(3) For § 29.783(h), a means such as a cable, chain, pin, or mechanical linkage should be provided to secure doors used as ditching exits in the open position. The means should be shown to be effective under rotorcraft attitudes and dynamic conditions common to ditching. The sea states for ditching approval in accordance with § 29.801 are found in paragraph AC 29.801. Demonstrations under actual ditching conditions are not mandated for substantiation purposes, but the substantiation methodology should be reliable, i.e., an analytical or test method demonstrated to be reliable and used in previous structural substantiation programs.

Page D - 82 9/30/99 AC 29-2C AC 29.785. § 29.785 SEATS, SAFETY BELTS, AND HARNESSES .

a. Explanation .

(1) This section requires that seats, belts, harnesses, and adjacent parts of the rotorcraft be substantiated for the structural loads resulting from the inertia forces of § 29.561 as well as normal flight and ground inertia forces on a 170-pound occupant.

The inertia forces of § 29.561 are ultimate loads and must be multiplied by a factor of 1.33 in determining the “strength of attachment” of each seat to structure and each belt or harness to structure. The seat, belt, etc., are required to sustain applied loads and to protect the occupant from serious injury. The pilot seats must also sustain the effects of the pilot forces of § 29.397.

(2) In addition, the “occupant must be protected from head injury” by the seat belt and one of the following: (i) A harness to prevent the head from contacting an injurious object.

(ii) Elimination of injurious object within striking distance of the head.

(iii) A cushioned rest as specified.

(3) Handholds are required to steady occupants using the aisle in moderately rough air.

(4) Projecting objects which would injure occupants “in normal flight must be padded.” b. Procedures .

(1) Each seat with its belts and harnesses are to be substantiated for the flight, ground, and emergency landing loads of § 29.561 by structural test or stress analysis.

Section 29.785(b) states that “each seat must be approved.” Certification approval can be gained by Technical Standard Order (TSO) approval or by accomplishing sufficient structural substantiation to gain FAA/AUTHORITY approval of the seat and its belt(s) as part of the Type Design of the rotorcraft. TSO No. C-39 concerns standards for aircraft seats, including rotorcraft seats. If TSO No. C-39 is used as an approval basis for a specific rotorcraft seat, the seat should be checked to assure it has been substantiated for the vertical (up and down) and side loads imposed by installation in the aircraft. For example, TSO No. C-39 (and NAS 809) specifies an ultimate down load of 4.0g which is in agreement with the 4.0g emergency landing load factor of § 29.561, but it may be less than the design maneuver load factor (which can be as high as 3.5g limit or 5.25g ultimate).

Page D - 83 AC 29-2C 9/30/99 (i) The 1.33 factor is specified for substantiation of attachments of each seat to the structure and each safety belt or harness to the seat or structure for § 29.561 loads, whether analysis or test is used.

(ii) If static testing of seats, belts, and harnesses is used, the body block of NAS 809 may be used. The corners of the NAS 809 body block may be radiused and padded if it is found that the small radii cause premature, unrealistic crippling of thin wall tubing or other structure used in the seat.

(iii) The substantiation of the pilot seats is required to include pilot forces of § 29.397 in conjunction with normal flight and ground loads. For example, the pilot foot force (195 pounds ultimate) must be reacted by the seat.

(2) The following criteria have been found satisfactory for preventing occupant head injuries: (i) If a harness is used, it should support the shoulders without applying hazardous loads to the side or front of the neck. It should be easily donned and a single point release with the seat belt is preferred. If separate release is provided, it must be simple, compatible with the seat belt release, and near the seat belt release. The harness should be tested in conjunction with the seat belt using a “body block” similar to that of NAS 809 if possible. If the harness is tested separately from the belt, it should be tested to 50 percent of the forward crash loads for the entire occupant weight of 170 pounds, unless that percentage distribution is found to be unrealistic by a rational analysis.

(ii) Elimination of injurious objects within striking distance of the head and other vital parts can be accomplished by removal of objects with sharp edges or rigid surfaces from within striking distance of vital parts of the occupant. Dimensions and weights for typical occupants are available in U.S. Army USAAULABS Reports 70-22 (August 1969) and 66-39 (June 1966) and NACA Report TN 2991 (August 1953).

Because of the range of occupant head striking distance, a combination of “elimination of injurious objects” and “cushioned rests” may be required for some interior configurations.

(iii) An acceptable cushioned rest can be provided by use of a 1-inch thickness of foamed polyvinyl chloride (PVC), or equivalent energy absorbing material.

The density of material should be in the 5 to 10 pounds per cubic foot density range.

PVC foam has the property of absorbing energy efficiently with negligible rebond effects. PVC foam recovers slowly to the original configuration after deformation. If PVC foam is used, however, care must be taken in its application relative to its flammability characteristics (reference § 29.853).

(3) Handholds for the occupants are generally provided by seat backs adjacent to the aisle. If the seat backs fold, the amount of support provided by the seat backs before they fold must be evaluated in a furnished interior or mock up. To provide Page D - 84 9/30/99 AC 29-2C adequate support, the seat back may use an easily disengaged latch or adequate friction in the hinge mechanism to obtain adequate support. Handholds along the aisle are, of course, not needed for rotorcraft with no aisles or where seat belts must be fastened during flight.

(4) Projecting objects which could injure occupants in normal flight should be padded. The amount of padding required depends on the location, size, and minimum radius of the projecting object. In general, this requirement will mean that sharp edges must be padded with one-half inch of PVC foam or equivalent (5 to 10 lbs. density), while objects with radii in excess of 1 inch may meet the requirements of § 29.785(e) with a lesser amount of energy absorbing padding, if it can be contacted only by persons “moving about in the rotorcraft in normal flight.” AC 29.785A. § 29.785 (Amendment 29-29) SEATS, BERTHS, BELTS, SAFETY BELTS, AND HARNESSES .

a. Explanation . Amendment 29-29 makes the following changes to § 29.785: (1) The title of § 29.785 now includes berths (which would include litters).

(2) Section 29.785(a) has been revised to include reference to the new § 29.562, “Emergency Landing Dynamic Conditions.” (3) Section 29.785(b) has been revised to include a reference to the new § 29.562(c)(5) head injury criteria and to describe a torso restraint system that is contained in TSO-C114.

(4) Section 29.785(f) has been revised to change the percentage of load distribution for safety belt and harness combination to 60-40.

(5) A new § 29.785(i) has been added which provides a list of “seating device system” components.

(6) A new § 29.785(j) provides for deformations of the seat energy absorption device system installed to meet the requirements of § 29.562 but requires that the system “remain intact and not interfere with rapid evacuation of the rotorcraft.” Further “structural” performance standards are contained in §§ 29.562(c)(1) and (2).

(7) A new § 29.785(k) provides static strength and restraint requirements for litters and berths. Litters may be oriented laterally as well as longitudinally in the rotorcraft. Dynamic tests of litters are not required. For longitudinally oriented litters, features should be provided to protect the occupant from the increased loads in § 29.561(b) of Amendment 29-29.

Page D - 85 AC 29-2C, Chg 1 2/12/03 b. Procedures . The procedures of paragraph AC 29.785 still apply to static substantiation of the seats, berths, safety belts, and harness. In addition: (1) Compliance with § 29.562 (except litters are not included) and § 29.561(b) is required.

(2) Section 29.562 includes a specific pass fail criteria, which includes head injury criteria.

(3) Shoulder harnesses need only be substantiated for 40 percent of total occupant load rather than the former 60 percent adopted by Amendment 29-24.

(4) AC 29.562 provides guidance for evaluating the functioning of a seating energy absorption device system under dynamic test conditions. Stroking is associated with the vertical-horizontal impact case and is recognized in the static strength substantiation.

(5) Berths or litters installed within 15° or less of the rotorcraft longitudinal axis (oriented longitudinally) shall use a combination of restraint devices, such as are required to be equipped with a padded end-board, cloth diaphragm, or equivalent means to withstand and distribute the occupant loads resulting from § 29.561(b) requirements. Other berths or litters may be equipped with straps or safety belts to withstand the forward reaction of § 29.561(b) as well as other loads, including flight loads.

(i) Berths/litters may be substantiated by static load tests, analysis, or a combination thereof and need not be substantiated to the 1.33 fitting factor of seat installations.

(ii) The berth/litter occupant’s head, neck, and spine should be protected from (landing) impact forward loads by appropriate design means; e.g., • non-longitudinal orientation of the berth/litter; or • “feet forward” orientation; or • distribution of an appropriate percentage of forward loads on the shoulders (not solely to the head and spine).

(iii) Recommendations for litter occupants: • If the occupant’s head is oriented forward, a shoulder harness should be provided, in conjunction with body and leg straps that prevents the occupant’s head from falling off the litter. A padded end board, diaphragm, etc., may be used, provided head and spinal loads are alleviated or prevented.

Page D - 86 5/1/2014 AC 29-2C Chg 4  If the occupant’s feet are oriented forward, the padded end board may also be used in combination with the body and leg straps or other such restraints.

 Multiple or combinations of devices should be used to distribute the occupant loads as well as protect the occupant from possible neck and spine compression.

AC 29.785B § 29.785 (Amendment 29-42) SEATS, BERTHS, LITTERS, SAFETY BELTS, AND HARNESSES.

a. Explanation. Amendment 29-42 revised the title of the rule to now include litters to distinguish from berths. Additionally, § 29.785(k) was revised to require that the 1.33 fitting factor of § 29.625(d) must be applied to the emergency landing loads of § 29.561(b)(3) for the substantiation of attachments of each berth and litter to the structure.

b. Procedures. All of the advisory material pertaining to this section remains in effect with the following additions: (1) A fitting factor of 1.33 must be applied to the emergency landing loads of § 29.561(b) when evaluating the berth or litter attachment and the occupant restraint system to the structure.

(2) The 1.33 factor is required whether analysis or test is used.

Page D - 87 AC 29-2C, Chg 4 5/1/2014 AC 29.787. § 29.787 (Amendment 29-12) CARGO AND BAGGAGE COMPARTMENTS.

a. Explanation .

(1) This section requires that cargo and baggage compartments be designed for normal flight and ground loads and for a 4g ultimate forward load condition.

Maximum placarded weights and critical distributions are to be considered.

(2) Means to prevent cargo shifting and contact between any cargo lamp bulb and cargo is to be provided.

b. Procedures . Structure tests or analyses may be used for substantiation for the design loads.

(1) Nets or straps may be used to prevent cargo shifting. The nets or straps are required to be substantiated for the structural loads. They need a means for adjustment to assure proper restraint for different sizes and shapes of cargo.

(2) Cargo lamp bulbs need to be guarded, recessed, or placed in upper inside corners to prevent contact with cargo.

AC 29.787A. § 29.787 (Amendment 29-31) CARGO AND BAGGAGE COMPARTMENTS.

a. Explanation . Amendment 29-31 adds two subparagraphs to § 29.787 (c) which clarify that cargo and baggage compartments should be designed to protect occupants from injury by the compartment contents during emergency landings. This may be done by location or by retention provisions . The new paragraphs also add a requirement that the compartment contents should not cause injury when subjected to the loads of § 29.561.

b. Procedures . The procedures of paragraph AC 29.787 are still applicable. In addition to the forward load, the cargo and baggage compartments should be designed to withstand loads in other directions as specified in § 29.561. Also, the compartment may be shown to provide protection of occupants by location; i.e., cargo and baggage compartments may be shown to be located in a position where loose contents will not endanger occupants in an emergency landing. If the compartment is located above or behind the occupied area, § 29.561(c) still applies. If a compartment is in the occupied area, § 29.561(b) may apply.

Page D - 88 5/1/2014 AC 29-2C, Chg 4 AC 29.801. § 29.801 (Amendment 29-12) DITCHING.

a. Explanation .

(1) Ditching certification is accomplished only if requested by the applicant.

(2) Ditching may be defined as an emergency landing on the water, deliberately executed, with the intent of abandoning the rotorcraft as soon as practical. The rotorcraft is assumed to be intact prior to water entry with all controls and essential systems, except engines, functioning properly.

(3) The regulation requires demonstration of the flotation and trim requirements under “reasonably probable water conditions.” The FAA/AUTHORITY has determined that a sea state 4 is representative of reasonably probable water conditions to be encountered. Therefore, demonstration of compliance with the ditching requirements for at least sea state 4 water conditions is considered to satisfy the reasonably probable requirement.

(4) A sea state 4 is defined as a moderate sea with significant wave heights of 4 to 8 feet with a height-to-length ratio of: (i) 1:12.5 for category A rotorcraft.

(ii) 1:10 for category B rotorcraft with category A engine isolation.

(iii) 1:8 for category B rotorcraft.

The source of the sea state definition is the World Meteorological Organization (WMO) Table (see Figure AC 29.801-1).

(5) Ditching certification encompasses four primary areas of concern: rotorcraft water entry, rotorcraft flotation and trim, occupant egress, and occupant survival.

(6) The rule requires that after ditching in reasonably probable water conditions, the flotation time and trim of the rotorcraft will allow the occupants to leave the rotorcraft and enter liferafts. This means that the rotorcraft should remain sufficiently upright and in adequate trim to permit safe and orderly evacuation of all personnel.

(7) For a rotorcraft to be certified for ditching, emergency exits must be provided with visible markings and must meet the requirements of §§ 29.807(d) and 29.811(a).

(8) The safety and ditching equipment requirements are addressed in §§ 29.1411, 29.1415, and 29.1561 and specified in the operating rules (parts 91, 121, 127, and 135). As used in § 29.1415, the term ditching equipment would more properly be described as occupant water survival equipment. Ditching equipment is required for extended overwater operations (more than 50 nautical miles from the nearest shoreline Page D - 89 AC 29-2C, Chg 4 5/1/2014 and more than 50 nautical miles from an offshore heliport structure). However, ditching certification should be accomplished with the maximum required quantity of ditching equipment regardless of possible operational use.

(9) Current practices allow wide latitude in the design of cabin interiors and consequently, the stowage provisions for safety and ditching equipment. Rotorcraft manufacturers may deliver aircraft with unfinished (green) interiors that are to be completed by the purchaser or modifier. These various “configurations” present problems for certifying the rotorcraft for ditching.

(i) In the past, “segmented” certification has been permitted to accommodate this practice. That is, the rotorcraft manufacturer shows compliance with the flotation time, trim, and emergency exit requirements while the purchaser or modifier shows compliance with the equipment provisions and egress requirements with the completed interior. This procedure requires close cooperation and coordination between the manufacturer, purchaser or modifier, and the FAA/AUTHORITY.

(ii) The rotorcraft manufacturer may elect to establish a “token” interior for ditching certification. This interior may subsequently be modified by a supplemental type certificate or a field approval. Compliance with the ditching requirements should be reviewed after any interior configuration and limitations changes where applicable.

(iii) The rotorcraft flight manual (RFM) and supplements (RFMS) deserve special attention if a “segmented” certification procedure is pursued.

b. Procedures . The following guidance criteria has been derived from past FAA/AUTHORITY certification policy and experience. Demonstration of compliance to other criteria may produce acceptable results if adequately justified by rational analysis.

Model tests of the appropriate ditching configuration may be conducted to demonstrate satisfactory water entry and flotation and trim characteristics where satisfactory correlation between model testing and flight testing has been established. Model tests and other data from rotorcraft of similar configurations may be used to satisfy the ditching requirements where appropriate.

(1) Water entry .

(i) Tests should be conducted to establish procedures and techniques to be used for water entry. These tests should include determination of optimum pitch attitude and forward velocity for ditching in a calm sea as well as entry procedures for the highest sea state to be demonstrated (e.g., the recommended part of the wave on which to land). Procedures for all engines operating, one engine inoperative, and all engines inoperative conditions should be established. However, only the procedures for the most critical condition (usually all engines inoperative) need to be verified by water entry tests.

Page D - 90 5/1/2014 AC 29-2C, Chg 4 (ii) The ditching structural design consideration should be based on water impact with a rotor lift of not more than two-thirds of the maximum design weight acting through the center of gravity under the following conditions: (A) For entry into a calm sea- - ( 1 ) The optimum pitch attitude as determined in 337(b)(1)(i) with consideration for pitch attitude variations that would reasonably be expected to occur in service; ( 2 ) Forward speeds from zero up to the speed defining the knee of the height-velocity (HV) diagram; ( 3 ) Vertical descent velocity of 5 feet per second; and ( 4 ) Yaw attitudes up to 15°.

(B) For entry into the maximum demonstrated sea state- - ( 1 ) The optimum pitch attitude and entry procedure as established in (b)(1)(i); ( 2 ) The forward speed defined by the knee of the HV diagram reduced by the wind speed associated with each applicable sea state; ( 3 ) Vertical descent velocity of 5 feet per second; and ( 4 ) Yaw attitudes up to 15°.

(C) The float system attachment hardware should be shown to be structurally adequate to withstand water loads during water entry when both deflated and stowed and fully inflated (unless in-flight inflation is prohibited). Water entry conditions should correspond to those established in paragraphs b.(1)(ii)(A) and (B).

The appropriate vertical loads and drag loads determined from water entry conditions (or as limited by flight manual procedures) should be addressed. The effects of the vertical loads and the drag loads may be considered separately for the analysis.

(D) Probable damage due to water impact to the airframe or hull should be considered during the water entry evaluations (i.e., failure of windows, doors, skins, panels, etc.).

(E) The ditching maximum demonstrated sea state for water entry is the same or greater than the maximum demonstrated sea state for flotation and trim.

(2) Flotation Systems .

Page D - 91 AC 29-2C, Chg 4 5/1/2014 (i) Normally inflated . Fixed flotation systems intended for emergency ditching use only and not for amphibian or limited amphibian duty should be evaluated for: (A) Structural integrity when subjected to: ( 1 ) Air loads throughout the approved flight envelope with floats installed; ( 2 ) Water loads during water entry; and ( 3 ) Water loads after water entry at speeds likely to be experienced after water impact.

(B) Rotorcraft handling qualities throughout the approved flight envelope with floats installed.

(ii) Normally deflated . Emergency flotation systems which are normally stowed in a deflated condition and inflated either in flight or after water contact during an emergency ditching should be evaluated for: (A) Inflation . The float activation means may be either fully automatic or manual with a means to verify primary actuation system integrity prior to each flight. If manually inflated, the float activation switch should be on one of the primary flight controls and should be safeguarded against spontaneous or inadvertent actuation for all flight conditions.

( 1 ) The inflation system design should minimize the probability of the floats not inflating properly or inflating asymmetrically. This may be accomplished by use of a single inflation agent container or multiple container system interconnected together. Redundant inflation activation systems will also normally be required. If the primary actuation system is electrical, a mechanical backup actuation system will usually provide the necessary reliability. A secondary electrical actuation system may also be acceptable if adequate electrical system independence and reliability can be documented.

( 2 ) The inflation system should be safeguarded against spontaneous or inadvertent actuation for all flight conditions. It should be demonstrated that float inflation at any flight condition within the approved operating envelope will not result in a hazardous condition unless the safeguarding system is shown to be extremely reliable.

One safeguarding method that has been successfully used on previous certification programs is to provide a separate float system arming circuit which must be activated before inflation can be initiated.

( 3 ) The maximum airspeeds for intentional in-flight actuation of the float system and for flight with the floats inflated should be established as limitations in the RFM unless in-flight actuation is prohibited by the RFM.

Page D - 92 5/1/2014 AC 29-2C, Chg 4 ( 4 ) The inflation time from actuation to neutral buoyancy should be short enough to prevent the rotorcraft from becoming more than partially submerged assuming actuation upon water contact.

( 5 ) A means should be provided for checking the pressure of the gas storage cylinders prior to takeoff. A table of acceptable gas cylinder pressure variation with ambient temperature and altitude (if applicable) should be provided.

( 6 ) A means should be provided to minimize the possibility of overinflation of the float bags under any reasonably probable actuation conditions.

( 7 ) The ability of the floats to inflate without puncture when subjected to actual water pressures should be substantiated. A full-scale rotorcraft immersion demonstration in a calm body of water is one acceptable method of substantiation.

Other methods of substantiation may be acceptable depending upon the particular design of the flotation system.

(B) Structural Integrity . The flotation bags should be evaluated for loads resulting from: ( 1 ) Airloads during inflation and fully inflated for the most critical flight conditions and water loads with fully inflated floats during water impact for the water entry conditions established under paragraph b.(1)(ii) for rotorcraft desiring float deployment before water entry; or ( 2 ) Water loads during inflation after water entry.

(C) Handling Qualities . Rotorcraft handling qualities should be verified to comply with the applicable regulations throughout the approved operating envelopes for: ( 1 ) the deflated and stowed condition; ( 2 ) the fully inflated condition; and ( 3 ) the in-flight inflation condition.

(3) Flotation and Trim . The flotation and trim characteristics should be investigated for a range of a sea states from zero to the maximum selected by the applicant and should be satisfactory in waves having height/length ratios of 1:12.5 for Category A rotorcraft, 1:10 for Category B rotorcraft with Category A engine isolation, and 1:8 for Category B rotorcraft. Model tests in a wave basin on a number of different rotorcraft types have indicated that an improvement in sea keeping, response of the rotorcraft to waves, performance of approximately one sea state can consistently be achieved by fitting float scoops. If the basic flotation system (without scoops) has Page D - 93 AC 29-2C, Chg 4 5/1/2014 demonstrated compliance with the minimum flotation and trim requirements, credit for float scoops to achieve stability in more severe water conditions may be allowed.

However, the effect of scoops on improved sea keeping must be demonstrated during model testing.

(i) Flotation and trim characteristics should be demonstrated to be satisfactory to at least sea state 4 conditions.

(ii) Flotation tests should be investigated at the most critical rotorcraft loading condition.

(iii) Flotation time and trim requirements should be evaluated with a simulated, ruptured deflation of the most critical float compartment. Flotation characteristics should be satisfactory in this degraded mode to at least sea state 2 conditions.

(iv) A sea anchor or similar device should not be used when demonstrating compliance with the flotation and trim requirements but may be used to assist in the deployment of liferafts. If the basic flotation system has demonstrated compliance with the minimum flotation and trim requirements, credit for a sea anchor or similar device to achieve stability in more severe water conditions (sea state, etc.) may be allowed if the device can be automatically, remotely, or easily deployed by the minimum flightcrew.

(v) Probable rotorcraft door and window open or closed configurations and probable damage to the airframe or hull (i.e., failure of doors, windows, skin, etc.)

should be considered when demonstrating compliance with the flotation and trim requirements.

(4) Float System Reliability . Reliability should be considered in the basic design to assure approximately equal inflation of the floats to preclude excessive yaw, roll, or pitch in flight or in the water.

(i) Maintenance procedures should not degrade the flotation system (e.g., introducing contaminants which could affect normal operation, etc.).

(ii) The flotation system design should preclude inadvertent damage due to normal personnel traffic flow and excessive wear and tear. Protection covers should be evaluated for function and reliability.

(iii) Float design should provide a means to minimize the likelihood of damage or tear propagation between compartments. Single compartment float designs should be avoided.

(iv) Where practical, design of the flotation system should consider the likely effects of an uncontrolled water entry and locate system components away from the major effects of structural deformity.

Page D - 94 5/1/2014 AC 29-2C, Chg 4 (v) Visual identification of the helicopter following a ditching (and possible capsize) is made easier by the choice of material for the construction of the floats that has high visual conspicuity properties.

(5) Occupant Egress and Survival . The ability of the occupants to deploy liferafts, egress the rotorcraft, and board the liferafts should be evaluated. For configurations which are considered to have critical occupant egress capabilities due to liferaft locations and ditching emergency exit locations and floats proximity, an actual demonstration of egress may be required. When a demonstration is required, it may be conducted on a full-scale rotorcraft actually immersed in a calm body of water or using any other rig or ground test facility shown to be representative. The demonstration should show that floats do not impede a satisfactory evacuation. Service experience has shown that it is possible for occupants to have escaped from the cabin but have not been able to board a liferaft and have had difficulties finding handholds to stay afloat and together. Where practical, handholds or lifelines should be provided. The normal attitude of the rotorcraft and the possibility of a capsize should be considered when locating the handholds or lifelines.

(6) Rotorcraft Flight Manual . The RFM is an important element in the approval cycle of the rotorcraft for ditching. The material related to ditching may be presented in the form of a supplement or a revision to the basic manual. This material should include: (i) The information pertinent to the limitations applicable to the ditching approval should include the range of sea state conditions that has been demonstrated for water entry and flotation stability. If the ditching approval is obtained in a segmented fashion (i.e., one applicant performing the aircraft equipment installation and operations portion and another designing and substantiating the liferaft or lifevest and ditching safety equipment installations and deployment facilities), the RFM limitations should state “Not Approved for Ditching” until all segments are completed. The requirements for a complete ditching approval not yet completed should be identified in the “Limitations” section.

(ii) Procedures and limitations for flotation device inflation.

(iii) Recommended rotorcraft water entry attitude, speed, and wave position.

(iv) Procedures for use of emergency ditching equipment.

(v) Ditching egress and raft entry procedures.

(vi) Information stating the flotation system has been certificated for Ditching (as opposed to Emergency Flotation) to facilitate compliance with operational requirements.

Page D - 95 AC 29-2C, Chg 4 5/1/2014 SEA STATE CODE (WORLD METEOROLOGICAL ORGANIZATION) Sea State Significant Wave Height Wind Speed Code Description of Sea Meters Feet Knots 0 Calm (Glassy) 0 0 0-3 1 Calm (Rippled) 0 to 0.1 0 to 1/3 4-6 2 Smooth (Wavelets) 0.1 to 0.5 1/3 to 1 2/3 7-10 3 Slight 0.5 to 1.25 1 2/3 to 4 11-16 4 Moderate 1.25 to 2.5 4 to 8 17-21 5 Rough 2.5 to 4 8 to 13 22-29 6 Very Rough 4 to 6 13 to 20 28-47 7 High 6 to 9 20 to 30 48-55 8 Very High 9 to 14 30 to 45 56-63 9 Phenomenal Over 14 Over 45 64-118 Notes: (1) The Significant Wave Height is defined as the average value of the height (vertical distance between trough and crest) of the largest one-third of the waves present.

(2) Maximum Wave Height is usually taken to be 1.6 x Significant Wave Height; e.g., Significant Wave Height or 6 meters gives Maximum Wave Height of 9.6 meters.

(3) Winds speeds were obtained from Appendix R of the “American Practical Navigator” by Nathaniel Bowditch, LL.D.; Published by the U.S. Naval Oceanographic Office, 1966.

FIGURE AC 29.801-1 Page D - 96 9/30/99 AC 29-2C AC 29.803. § 29.803 (Amendment 29-3) EMERGENCY EVACUATION .

a. Explanation . The regulation specifies that “means for rapid evacuation in a crash landing” be provided considering the landing gear extended or retracted, and “considering the possibility of fire.” Any external exits, whether normal entrance doors or service doors, can be considered as emergency exits if the requirements of §§ 29.805 through 29.815 are met. “Limited amphibian rotorcraft” emergency exits are required to be designed for probable maximum local water pressure (or shown to have nonhazardous failure characteristics) and to have a specified number of exits above the water level. Limited amphibian rotorcraft are approved under the provisions of §§ 29.519 and 29.755(b). Sections 29.801 and 29.807(d) refer to similar standards that pertain to “rotorcraft ditching configurations.” b. Procedures . Exits, arrangement, markings, access, and aisle widths as specified in § 29.805 through 29.815 are to be provided. Recent rotorcraft designs have been approved under the “ditching” standards of § 29.801. Previous “limited amphibian rotorcraft” were designed to the same standards.

AC 29.803A. § 29.803 (Amendment 29-30) EMERGENCY EVACUATION .

a. Explanation .

(1) Amendment 29-30 removed § 29.803(c) which concerned limited amphibians, now obsolete with adoption of § 29.801 ditching standards, and added § 29.803(d) for evacuation criteria of certain rotorcraft designs. Part 29, Appendix D evacuation procedures was adopted concurrently. In addition, newly adopted § 29.803(e) allows use of analysis and tests for compliance with the standard.

(2) This amendment adds explicit demonstration requirements even for certain smaller but “dense” interior arrangements as stated in § 29.803(d)(2).

(3) The 90-second duration for an evacuation demonstration through all exits on one side of the rotorcraft is a primary addition to the standard.

b. Procedures . All of the policy material pertaining to this section remains in effect with the following additions: (1) For rotorcraft with a seating capacity of more than 44 passengers, conduct an emergency evacuation in accordance with the provisions of Part 29, Appendix D.

(2) For certain smaller rotorcraft with a van or limousine-type “dense” interior as defined in § 29.803(d)(2), conduct an emergency evacuation in accordance with the provisions of Part 29, Appendix D. The rotorcraft should meet all three requirements before a demonstration is specifically required.

Page D - 97 AC 29-2C 9/30/99 (3) Part 29, Appendix D contains procedures. Safety equipment for alleviating “ground” injuries is contained in paragraph (c) of the Appendix.

(4) A combination of analysis and tests may be used in lieu of test only. A combination of tests and analysis is particularly intended to evaluate emergency evacuations from rotorcraft from 10 to 44 passengers with van or limousine-type interiors. Test other than full-scale evacuation tests may be used in conjunction with analyses to evaluate specific design features such as folding seat backs which affect only one or two passengers. That is, sections of an interior may be used to evaluate a feature and its effects on prompt evacuation of the rotorcraft.

AC 29.805. § 29.805 (Amendment 29-3) FLIGHTCREW EMERGENCY EXITS .

a. Explanation . Flightcrew emergency exits are required when passenger exits are not convenient. The placement of litters, cargo, or bulkheads may prevent passenger exits from being convenient to the flightcrew. Flightcrew exits, if required, are to be of sufficient size and located on both sides of the rotorcraft (or one top hatch) to “allow rapid evacuation of the flightcrew.” A test or tests are required.

b. Procedures . Flightcrew emergency exits, if required, may consist of one overhead hatch or two side exits (one on either side). The size is not explicitly defined except that it be “of sufficient size . . . to allow rapid evacuation of the flightcrew.” The ability for “rapid evacuation” should be demonstrated by test. For side exits located immediately adjacent to the crew seat and exceeding Type IV exits (§ 29.807) in size, the test demonstration can be accomplished by normal use and evaluation of the exits by the FAA/AUTHORITY crew during Type Inspection Authorization (TIA) testing. For any overhead exit or side of fuselage exits not meeting Type IV dimensions, a special demonstration test should be accomplished. This demonstration should show that 2.5 percentile to 97.5 percentile men could egress rapidly through the crew exit(s), i.e., men 5 feet 4 inches to 6 feet 5 inchesto 6 feet in height and up to 225 pounds in weight, based on the Civil Aeromedical Institute’s (CAMI) 1998 Aeromedical Certification Statistical Handbook . If an overhead hatch type exit is utilized, on conventional rotorcraft designs, hazards associated with the proximity of rotor blades should be considered.

AC 29.805A. § 29.805 (Amendment 29-30) FLIGHTCREW EMERGENCY EXITS .

a. Explanation . Amendment 29-30 adds a new paragraph § 29.805(c) which requires that water or flotation devices not obstruct the flight crew emergency exits after a ditching. Test, demonstration, or analysis is required for substantiation.

b. Procedures .

Page D - 98 9/30/99 AC 29-2C (1) The tests, demonstrations, or analysis required by § 29.805(c) for flight crew exits is analogous to those of § 29.807(d)(3) except the crew exit threshold may be slightly below the water line but should not obstruct use of the exit.

(2) Tests in water (tanks or large bodies of water) or demonstrations in the laboratory may be used for compliance if the deflections of flotation devices relative to the exits are accurately or conservatively achieved.

(3) Obstructions should be identified, should be minor, and should not interfere with exit removal or opening, or with crew egress.

AC 29.807. § 29.807 (Amendment 29-12) PASSENGER EMERGENCY EXITS .

a. Explanation . The normal passenger exits (type and number in each side of fuselage) are specified as follows: (1) For overland operations .

Passenger Emergency exits (rectangular with corner radii of width/3) Seating for each side of the fuselage Capacity Floor level Step-up -29” Max.

Type I Type II Type III Type IV 24” X 48” 20” X 44” 20” X 36” 19” X 26” 1 through 10 1 11 through 19 1 or 2 20 through 39 1 1 40 through 59 1 1 60 through 79 1 1 or 2 (2) For overwater operations (related to ditching an optional standard).

Passenger Emergency exits (rectangular with corner radii of width/3) Seating for each side of the fuselage Capacity Threshold Above Waterline Type III Type IV 20” X 36” 19” X 26” w/step-up - 29” MAX 1 through 9 1 10 through 35 1* Each Additional or Partial Unit of 35 1* *The passenger seat-to-exit ratio may be increased by using larger exits if proven by analyses or tests.

Page D - 99 AC 29-2C 9/30/99 (3) For crash rollover conditions . Sufficient top, bottom, or ends of fuselage exits are to be provided for evacuation unless the probability of the rotorcraft coming to rest on its side in a crash landing is extremely remote.

(4) Ramp exits to replace Type I or II exits are permitted.

(5) Each emergency exit must be functionally tested.

b. Procedures .

(1) The number and size of overland and overwater operation exits will be as specified. The use of oversize exits is allowed if the threshold is flat and of the specified width.

(2) The top, bottom, or end fuselage exits should be provided unless features of design are provided which prevent the rotorcraft from coming to rest on its side in a crash landing, and unless sufficient fail-safe and fatigue tests and analyses are conducted of the landing gear and support structure to show it is unlikely that the rotorcraft will come to rest on its side as a result of a single structural failure. An analysis is generally necessary to prove compliance with § 29.807(c).

(3) Ramp exits may be used in place of one Type I or one Type II exit if the required Type I or Type II exit is impractical, and if the § 29.813 exit access requirements are met by the ramp exits.

(4) Each emergency exit is to be opened from the inside and the outside as a functional test. Interior panels and seats should be installed for the exit functional tests to check for interferences and other effects. Section 29.813 pertains to access to the exits.

AC 29.807A. § 29.807 (Amendment 29-30) EMERGENCY EXITS .

a. Explanation . Amendment 29-30 added § 29.807(d)(3) which requires proof that all ditching configuration exits will be free of interference from emergency flotation devices, whether stowed or deployed (inflated). The threshold for each of these “ditching” exits should be above the water line in calm water.

b. Procedures .

(1) Test, demonstration, compliance inspection, or analysis is required to show freedom from interference from stowed and deployed emergency flotation devices. In the event an analysis is insufficient or a given design is questionable, a demonstration may be required. Such a demonstration would consist of an accurate, full-size replica (or true representation) of the rotorcraft and the flotation devices while stowed and after their deployment.

Page D - 100 9/30/99 AC 29-2C (2) The type inspection authorization may be used to perform compliance evaluation utilizing a full-scale rotorcraft in calm water. Designs may be accepted “by compliance inspection” if location of exit and flotation devices relative to each other ensures that interference is impossible. In this case, a demonstration is unnecessary.

AC 29.809. § 29.809 (Amendment 29-3) EMERGENCY EXIT ARRANGEMENT .

a. Explanation . Emergency exits are to be provided which result in an unobstructed opening to the outside. The following emergency exit requirements are the same as passenger door requirements of § 29.783 and noted for convenience.

(1) Openable from inside or outside.

(2) Simple and obvious means for opening.

(3) Means for locking.

(4) Means to prevent opening in flight inadvertently or as a result of mechanical failure.

(5) Means to minimize jamming in a minor crash landing.

NOTE: In addition the following emergency exit requirements are: (1) the means of opening may not require exceptional effort; and (2) a slide (for floor level exits) or rope must be provided as prescribed for exits whose thresholds are more than 6 feet from the ground (unless located over the wing). Sections 29.1411(c) and 29.1561 contain other standards for the descent devices.

b. Procedures . Subparagraphs 1 through 5 of the above explanation are covered in the procedure for § 29.783, paragraph AC 29.783.

(1) The effort required to open the exit can be evaluated when the tests of § 29.807(f) are conducted. If the effort required to open the exit is in the range of 40 to 50 pounds, it is recommended that a person of slight stature, such as a female in the 90 to 110 pound weight range, be used for the exit opening demonstration/test. In any case, the average load required to operate the exit release mechanism and open the exit should not exceed 50 pounds, and the maximum individual load of a test series should not exceed 55 pounds.

(2) If an approved escape slide, or its equivalent, is provided for exits more than 6 feet from the ground with the landing gear extended, it should be located near the door and conspicuously marked. Automatic inflation and deployment under emergency conditions are the preferred means of operation but are not required by § 29.809. If automatic inflation and deployment features are provided, design features should prevent inadvertent deployment if the exit is a door used for normal entry and/or service. If manual deployment methods are used, they must be simple and easily Page D - 101 AC 29-2C 9/30/99 carried out by a person of slight build and strength. The slide should rapidly inflate upon deployment. See § 29.809(f) for standards concerning an escape rope.

AC 29.809A. § 29.809 (Amendment 29-30) EMERGENCY EXIT ARRANGEMENT .

a. Explanation .

(1) Amendment 29-29 added the phrase, “under the ultimate forces in § 29.783(d),” to clarify that the following inertial load factors previously stated in § 29.809 were not altered by Amendment 29-29 and that the previous design conditions still apply to § 29.809(e) exits as well as the doors: (i) Upward - 1.5g (ii) Forward - 4.0g (iii) Sideward - 2.0g (iv) Downward - 4.0g (2) Amendment 29-30 further revised the requirements of § 29.809 by: (i) Amending requirements of § 29.809(f) to include landing gear malfunction or failure in determining the distance from the exit to the ground. (A means is required to assist occupants in descending to the ground when that distance is more than 6 feet); (ii) Adding specific requirements for automatic slides, automatic slide deployment (not optional), and slide qualification in a new § 29.809(g); (iii) Allowing relaxation in § 29.809(h) such that a rope or other assist means may be used rather than a slide for rotorcraft having 30 or fewer passenger seats provided an evacuation demonstration is successfully accomplished; and, (iv) Moving but not changing the egress rope requirements formerly in § 29.809(f) to a new § 29.809(i).

b. Procedures .

(1) The procedures of paragraph AC 29.809 continue to apply except compliance should consider landing gear collapse, breaking, or not extending as well as slide deployment and proper inflation in 25 knot winds.

(2) Automatic deployment of slides is now a requirement, not an option.

(3) Procedures for slide qualification tests are explicitly provided in § 29.809(g)(5).

Page D - 102 9/30/99 AC 29-2C AC 29.811. § 29.811 (Amendment 29-24) EMERGENCY EXIT MARKING .

a. Explanation .

(1) This regulation covers both the marking and exit interior illumination by emergency lighting prior to Amendment 29-24.

(2) With adoption of Amendment 29-24, the interior emergency lighting standards were moved to § 29.812, and exterior emergency lighting standards were added. However, the standards for emergency lighting in § 29.812 apply to transport Category A rotorcraft. Transport Category B rotorcraft shall have the “emergency” lighting required in § 29.811(d). General interior lighting standards are no longer specified in § 29.811.

(3) Locating and marking signs are specified for each emergency exit with the following features: (i) Locating signs and marking signs are to- - (A) Be recognizable from a distance equal to the width of the cabin; (B) Have 1-inch white letters on a 2-inch red background (colors may be reversed); and (C) Be self- or electrically illuminated to a minimum brightness of 160 microlamberts.

(ii) Locating signs visible to occupants approaching along the main aisle are required for each exit.

(A) The sign is required next to or above the aisle for floor level exits.

(B) Bulkheads or dividers obscuring exits must have exit locating signs except as stated.

(4) Exit operating or release handle instructions are to be- - (i) Readable from a distance of 30 inches; and (ii) Supplemented with a red arrow and sign (for Type I or Type II exits with a handle having rotary motion) with the following features provided: (A) A red arrow with a ¾-inch shaft, a head of twice the shaft width, and a 70° arc at 75 percent of handle length.

(B) The word “open” in red letters 1 inch high near the head of the arrow.

Page D - 103 AC 29-2C 9/30/99 (5) Emergency lighting .

(i) Prior to Amendment 29-24, an independent source of light, as prescribed, shall be installed in transport Category A or B rotorcraft to: (A) Illuminate marking and locating signs; (B) Provide general lighting of 0.05 foot-candles at 40-inch intervals at armrest height along the main aisle; and (C) Operate manually and automatically in a crash landing and when the normal electrical power is interrupted.

(ii) Amendment 29-24 requires for transport Category B rotorcraft either self- or electrically illuminated exit marking and locating signs. General lighting standards are not specified. See § 29.812 for transport Category A standards.

(6) External exit markings are required which include a 2-wide band around the exit, identification, and instructions for opening. The external markings are to have a reflectance difference of 30 percent from the fuselage surface finish.

(7) Emergency exits signs may read simply “EXIT.” (8) Excess exits should meet all of the “EXIT” standards or should not be identified as an exit.

b. Procedures .

(1) Emergency exit locating signs may be located to the side of the aisle for small fuselage heights, rather than over the aisle where they may present a hazard to the occupant’s head and possibly impede egress. For small passenger cabins one self-illuminated sign stating “EXIT” may be used as both the locating and marking sign for an individual exit on one side of the cabin (operating instructions will, of course, still be required). If one “EXIT” sign is used to both locate and mark the exit, it should be attached to the fuselage above the exit and not to the exit itself. If it is attached to the exit itself and the exit is discarded from the cabin after opening, the locating function of the exit sign is lost when the exit is removed. That is, there is no sign to locate the exit for passengers other than for the one who discarded the exit. The exit locating sign is a necessity to direct all occupants.

(2) Operating instructions should be provided as specified. They should be kept short but clear; e.g., “rotate handle,” “push,” “pull,” etc.

(3) Lighting should be provided as specified to illuminate the cabin for egress paths and to supplement lighting of the exit operating instructions signs.

Page D - 104 9/30/99 AC 29-2C (4) The reflectance of external exit markings can be checked by appropriate electro-optical instrumentation or by use of photometer card sets. AC 20-47, Exterior Colored Band Around Exits on Transport Airplanes, provides information for complying with identical standards contained in § 25.811. These are also acceptable for § 29.811.

The Munsell Color Company, 2441 North Calvert Street, Baltimore, Maryland 21218, provides a set of cards which includes shades of most commonly used colors.

AC 29.811A. § 29.811 (Amendment 29-31) EMERGENCY EXIT MARKING .

a. Explanation .

(1) Amendment 29-30 changes § 29.811(f)(1) to allow marking or outlining the handles, release devices, levers, etc., of passenger emergency exits which are “normally used doors,” rather than outline the entire door of smaller transport rotorcraft.

If an exit, other than a normally used door, such as a hatch, window, etc., is approved, that exit would be marked around the perimeter as described.

(2) Amendment 29-31 added two requirements to § 29.811(a): (i) A clarification that emergency exit markings should be conspicuously marked for egress in darkness as well as in daylight.

(ii) A requirement for visibility of emergency exit markings when the “rotorcraft is capsized (in water) and the cabin is submerged. This standard applies to rotorcraft configurations complying with § 29.801.

b. Procedures . The procedures of paragraph AC 29.811 are still applicable plus: (1) The release device, handle, etc., of the normally used door(s) may be separate from the normally used handle of the door (such as a release system lever for sliding door rollers). To preclude jamming a sliding door, which is also an exit, in an emergency landing impact, the door should be released from the track. An emergency release handle for releasing door rollers may be used to allow the exit door to be “pushed off” the track. For smaller rotorcraft, such a release lever should comply with the necessary operating procedures and exit markings but should use a distinct, separate 2-inch wide band around the release lever per § 29.811(f)(1). That is, a distinct “band” is necessary to comply rather than a solid block of color around the release lever. Large rotorcraft should have exits marked with a distinct 2-inch band around the exit perimeter as stated in subparagraphs § 29.811(f)(1). Refer to paragraph AC 29.811 for color contrast.

(2) The interior compliance checklist should report that emergency exit markings have been evaluated by “interior compliance inspections” conducted in darkness as well as daylight, and visibility of interior emergency exit markings should be checked under submerged cabin conditions or alternate/equivalent means for those rotorcraft configurations equipped for over-water flights that are approved under § 29.801.

Page D - 105 AC 29-2C 9/30/99 AC 29.812. § 29.812 (Amendment 29-24) EMERGENCY LIGHTING .

a. Explanation . Section 29.812 was added by Amendment 24. This change unified the requirements for an emergency lighting system into a single paragraph and required these systems only for Category A rotorcraft. The purpose of this change was to afford passengers flying in Transport Category A rotorcraft the same level of safety in an emergency evacuation at night as passengers flying in transport category airplanes.

b. Procedures . This paragraph is quite similar to the emergency lighting system required for Part 25 airplanes. The exception is there are no requirements in this paragraph for floor proximity emergency escape path markings. The following items should be considered in the design of emergency lighting systems: (1) There is a requirement for two controls of the system. One of these controls is located in the cabin, where it can be operated by a flight crew member or a passenger. The other control is located in the cockpit. These switches must have an “ON,” “OFF,” and “ARMED” position. These switches should operate independently of each other, and any other systems in the rotorcraft. The emergency lights must become lighted or remain lighted if the switch is either turned on, or the switch is in the armed position and there is an interruption of the rotorcraft electrical power supply.

Inertia switches should not be used to satisfy this requirement.

(2) Sharing of light bulbs with the normal cabin lighting is acceptable provided there is sufficient isolation of the emergency lighting system from the normal cabin lighting circuits. No single failure of the shared portion should render the emergency lighting system inoperative.

(3) The luminosity tests of the emergency lighting system should be accomplished with the emergency exits open.

AC 29.813. § 29.813 (Amendment 29-12) EMERGENCY EXIT ACCESS .

a. Explanation . Paragraph (a) of § 29.813 prescribes design details for passageways, both between passenger compartments and for access to Type I and II emergency exits, should they be provided. Such passageways are not made mandatory by § 29.813 although most larger rotorcraft have used them. Some utility or “wide-body” rotorcraft may have open areas between the crew area (pilots) and passenger area (cabin). These configurations may have lateral seating arrangements providing access to emergency exits of Type I or II size, even though they may not be required by § 29.807(b). These designs may not have a main aisle.

(1) Paragraph (c) of this standard concerns access to Type III and Type IV exits. Although “passageways” with explicit requirements are not required for Type III and Type IV exits, “ access from each aisle to each Type III and Type IV exit” is required.

Page D - 106 9/30/99 AC 29-2C (2) For exits whose thresholds are more than 6 feet above the ground, additional space adjacent to the exit is required to allow room for a crewmember to assist passengers with the descent device such as an escape slide or rope noted in § 29.809(f).

(3) In addition to requiring passageways and crewmember space adjacent to exits over 6 feet above the ground, this standard does not allow obstructions in the projected opening of Type III or Type IV emergency exits for one seat width from the exit, except as noted. For passenger seating configurations of 19 or less, minor obstructions into the projection of the exit are allowed only if “compensating factors to maintain the effectiveness of the exit” are provided.

b. Procedures .

(1) The provision for unobstructed passageways, at least 20 inches wide as specified, is straightforward for medium or large cabins with a main aisle and a typical rectangular floor plan. Care should be taken to assure that seats (with lateral or fore-and-aft movement) or galleys (with doors or drawers) are not installed so that they can encroach upon the required passageway. Design features such as stops in seat tracks, seat back mechanisms, stops in galley door (or drawer) mechanisms may be required to assure that unobstructed passageways are provided.

(2) The requirement (added by Amendment 29-12) that “access from each aisle to each Type III and Type IV exit” be provided may add design features to the interior of many typical compact interiors of medium-size rotorcraft. Rotorcraft with emergency exits located in either hinged or sliding doors and having passenger area encroachment or protrusions by compartments for fuel cells, gear boxes, etc., may require special design features to assure that passengers seated to one side or one area of the cabin have “access” to all Type III or Type IV exits on the same or other side of the rotorcraft.

The cabin must not be separated into compartments or partitioned. For example, fold down seat back mechanisms may be required for compact cabin configurations having only lateral aisles rather than longitudinal aisles and having Type III or Type IV exits located on each side of the cabin at the end of the lateral seat row or rows.

(3) The space adjacent to an exit that requires a crewmember to assist passengers with descent devices must be large enough to prevent the crewmember from becoming an obstruction in access to the exit. Twenty inches of access must be maintained.

(4) Minor obstructions are allowed in the projected opening of Type III or Type IV exits (for 19 or less passenger seat configurations) if “compensating factors to maintain the effectiveness of the exit” are provided. Compensating factors may include such design features as larger than required exit opening, additional exits beyond the minimum number required, or steps or other assist features which facilitate egress through the exit with the obstruction. Test or analysis may be required to prove the effectiveness of the compensating feature.

Page D - 107 AC 29-2C 9/30/99 AC 29.815. § 29.815 (Amendment 29-12) MAIN AISLE WIDTH a. Explanation. Main aisle widths are specified in the following table: Minimum main passenger aisle width Passenger seating capacity Less than 25 inches 25 inches and more from floor from floor Inches Inches 10 or less---------------------------------- 12* 15 11 through 19---------------------------- 12 20 20 or more-------------------------------- 15 20 *A narrower width not less than 9 inches may be approved when substantiated by tests found necessary by the Administrator.

b. Procedures .

(1) Provide the specified aisle minimum width where a longitudinal main aisle is provided in the type design.

(2) Historically, certain rotorcraft with short, wide cabins were initially designed without a longitudinal main aisle for military and cargo use, but were later fitted and approved for civil passenger configuration. These craft generally have 19 or less passenger seats and have either (1) outboard facing passenger seats, (2) a limited number of lateral rows with fold down seats/seat backs, or (3) a combination of lateral and longitudinal rows with and without main aisles to facilitate entrance and egress.

Page D - 108 9/30/99 AC 29-2C AC 29.831. § 29.831 VENTILATION .

a. Explanation .

(1) This rule specifies minimum ventilation requirements for each passenger and crew compartment. The minimum requirement for fresh air in the crew compartment is that amount that will allow the crew to accomplish their duties without undue discomfort or fatigue which shall be at least 10 ft /m per crewmember. The passenger and crew compartments are also required to be free from harmful or hazardous concentrations of gases or vapors. Specifically for carbon monoxide, the concentration may not exceed 1 part in 20,000 parts of air during forward flight. Failure conditions must also be considered when applying this rule.

(2) This rule becomes more significant when engine bleed air is used for conditioning of the passenger and crew compartments’ air. Certain data are necessary in order to properly analyze the bleed air provided under normal and malfunction conditions. The airframe manufacturer can normally look to the engine manufacturer for a specification of the maximum amount of air that can be extracted and the temperature of the extracted air. The engine manufacturer also normally provides a failure analysis that identifies ways the bleed air can be contaminated and the associated oil flow rates under each failure condition. The oil manufacturers are in a position to provide information regarding breakdown of the oil under different temperature conditions and the impact of that breakdown on the quality of the air being provided to the passenger and crew compartments.

b. Procedures .

(1) The passenger and crew compartments should be initially analyzed to ensure that at least 10 ft /m per crewmember of ventilation air is being provided. The emphasis has been placed on forward flight and, “air scoops” have been one way of showing compliance with this rule. Most installations also include blowers; however, they are normally provided primarily for defogging the windshields, and a secondary benefit is some circulation during ground or hover operation. In addition, the flight test crew should be asked to do a qualitative evaluation to ensure the amount of ventilation air actually provided meets the requirement for the crew to be able to accomplish their duties without undue discomfort or fatigue. In addition, the ventilation devices provided should not excessively increase the noise level in the cockpit. Compliance with the first requirement of § 29.831(a) can therefore be shown by an analysis showing the existence of at least 10 ft /m per crewmember, and a report from the flight test crew indicating that the amount actually provided is satisfactory.

(2) The passenger and crew compartment should be monitored under normal operating conditions for the presence of carbon monoxide. A carbon monoxide test kit is normally used for this evaluation. Air is monitored around outlets and different combinations of windows closed/open, heat off/on, air-conditioner off/on, etc., are checked to ensure all conditions are evaluated.

Page D - 109 AC 29-2C 9/30/99 (3) When engine bleed air is used to condition the passenger and crew compartments’ air, it should be initially substantiated that under normal operation, the amount of air being extracted does not exceed the limit established by the engine manufacturer. To accomplish this, determine the flight condition that will give the maximum bleed air flow through the flow limiter (venturi). The flow calculations should use this maximum flow condition and should also be made using the maximum tolerance diameter of the venturi throat.

(4) The engine bleed air should also be evaluated under malfunction conditions to determine a worst-case air contamination condition. (A typical worst-case malfunction is for an oil seal to fail in the engine that allows the engine oil supply to be introduced into the airflow.) With information regarding the contaminant, flow rate calculations can be made to predict the contamination levels that will be reached in the passenger and crew compartments and also the associated time duration of passenger and crew exposure. The severity of the exposure to the contaminated air is related to the temperature of the oil when it is introduced into the airflow. For example, synthetic base oils manufactured to MIL-L-7808 or MIL-L-23699 begin to break down into toxic components when the temperature exceeds 300° C (572° F). The oil manufacturers have evaluated this problem and should be in a position to provide data regarding the amount and type of toxic components to be expected, and the effect of introducing those components into the passenger and crew compartments. Therefore, from information supplied by the engine manufacturer, the worst-case air contamination condition can be calculated, and this can be compared with results of the oil manufacturers’ tests to determine if the concentrations are harmful or hazardous.

AC 29.833. § 29.833 HEATERS .

a. Explanation . This standard provides that each combustion heater must be approved. The standard contains no provisions regarding functioning of the system, environmental considerations, or malfunctions, therefore, the provisions of §§ 29.1301 and 29.1309 should be used to evaluate those aspects of an installation. The provisions of § 29.831, ventilation, should also be considered, as well as § 29.859, concerning combustion heater fire protection.

b. Procedures .

(1) Technical Standard Order, TSO-C20, was issued June 15, 1949, and amended on April 16, 1951, and concerns Combustion Heaters. If a heater chosen for installation has been qualified to the provisions of TSO-C20, it is considered to be approved. If a unit is not qualified to TSO-C20, a qualification program for the heater itself should be established with FAA/AUTHORITY certification engineers participating in the program as early as possible. The program should be based on the provisions of the TSO.

(2) The TSO refers to the SAE Aeronautical Standard, AS 143B, which specifies certain additional devices, design features, air supply considerations, performance tests, safety controls, environmental considerations, and so forth.

Page D - 110 9/30/99 AC 29-2C Consideration of all of the provisions of the aeronautical standard should result in an approved unit; however, it will not necessarily result in a satisfactory installation. For environmental considerations, it should be possible to specify an environmental spectrum more suitable to rotorcraft by referencing the latest version of Document No. RTCA/DO-160, Environmental Conditions and Test Procedures for Airborne Equipment, rather than AS 143B. Other specifications may also be satisfactory.

(3) The installation evaluation should consider functioning of the system based on the provisions of § 29.1301. Section 29.1309(a) is the regulatory basis for consideration of environmental conditions, and the expected environmental conditions resulting from the installation should be compared to those specified in the TSO. If the two are not compatible, additional environmental considerations are appropriate. The provisions of § 29.1309(b) should be used to evaluate the possible malfunctions of the installed system, and this evaluation should be documented in a fault analysis. The provisions of § 29.831 should be considered since certain standards of ventilation air quality under normal and malfunction conditions are specified. Additionally the provisions of § 29.859 should also be considered.

Page D - 111 AC 29-2C 9/30/99 SUBPART D - DESIGN AND CONSTRUCTION FIRE PROTECTION AC 29.851. § 29.851 FIRE EXTINGUISHERS .

a. Explanation .

(1) The standard concerns objective performance criteria for both handheld fire extinguishers in the crew and passenger compartments and built-in fire extinguisher systems if the system is required.

(2) Section 29.853(e) and (f) dictate the quantity and general location of the handheld fire extinguishers.

(3) Section 29.855(d) contains standards for cargo/baggage compartments.

(4) Sections 29.1541 and 29.1561 concern durable and conspicuous markings and placards for location and operation or use of the equipment.

(5) The rotorcraft flight manual should contain appropriate information as well.

(6) Advisory Circular 20-42C, Handheld Fire Extinguishers for use in Aircraft, provides an acceptable means of compliance with the standard.

b. Procedures.

(1) Advisory Circular 20-42C provides valuable information to select the type and size of the handheld extinguishers.

(2) The type design data shall contain appropriate information. One location should be used (recommended) for the crew compartment. Several locations may be selected to allow for evaluation and approval of several extinguishers and their locations in the passenger compartment.

(3) During a compliance inspection of a complete interior, the installation of required and optional extinguishers shall be checked for compliance.

(4) Whenever an extinguisher is installed, even though not required by § 29.853(f), it shall also comply with the standards.

Page D - 112 9/30/99 AC 29-2C AC 29.853. § 29.853 (Amendment 29-23) COMPARTMENT INTERIORS .

a. Explanation .

(1) Interior materials and components, windows, linings, etc., must meet certain flammability standards as set forth in Amendment 29-17. The rule refers to Part 25, Appendix F (Amendment 25-32), for procedures. Flight Standards Service Release No. 453 contained acceptable flammability standards for specified interior materials prior to adoption of Amendment 29-17.

(2) Smoking may be permitted with use of self-contained removable ashtrays as specified in § 29.853(c).

(3) Fire resistant waste containers may be used as specified.

(4) Hand fire extinguishers are required for flight crewmembers and passengers as specified. Section 29.851 and AC 20-42C, Hand Fire Extinguishers for use in Aircraft, dated March 7, 1984, contain standards for the extinguishers.

Section 29.1561(b) concerns identification and operating information signs for the safety equipment, and § 29.1411 concerns accessibility of the equipment.

(5) Amendment 29-23 adopted new flammability requirements for passenger seat and seat back cushions. Section 29.853(b) was added to require tests of “fire blocking” features of the cushions including upholstery materials. The rule refers to Part II, Appendix F, FAR Part 25 or an equivalent for the test procedures and test specimen requirements. Appendix F to Part 25, effective November 26, 1984, is the correct reference.

b. Procedures .

(1) With adoption of Amendment 29-17, materials subject to the flammability standards were significantly expanded. Acrylic windows and signs and transparencies were, for example, included. The rules list the materials and components subject to the flammability standards and refer to Appendix F of FAR Part 25 for the test procedures.

Specific burn chambers are also required for the tests. See paragraph b(5) below for flame application time and reference to Appendix F.

(2) A placard prohibiting smoking at all times may be used if ashtrays are not provided. If ashtrays are provided, the installation must have an inner fire resistant liner to close off the ashtray cavity or receptacle when the ashtray is removed. An illuminated sign or signs must be used if prescribed. Each crewmember must be able to control illumination of the sign.

(3) Fire resistant waste containers must have self-closing lids, such as a spring-loaded lid. If a removable container is installed in the receptacle, it must meet the same fire resistant standards as the receptacle. The receptacle must not have any openings outside the galley or an opening into the rotorcraft structure. An opening may Page D - 113 AC 29-2C 9/30/99 allow accumulation of trash and may allow flames and smoke to go throughout the rotorcraft in case of fire.

(4) A fire extinguisher must be adjacent to crew seats and must be readily accessible to the crew (§§ 29.1411 and 29.853(e)). The extinguisher should be accessible to the crewmember while he is seated. Fire extinguishers are also required in the passenger compartment for seven or more passengers. If one passenger is allowed in the left forward crew seat and six passengers are allowed in the passenger compartment, an extinguisher is not required for the passenger compartment. The extinguisher specified in § 29.853(e) should be located, whenever possible, so that it is visible and convenient to the passengers. If the passenger compartment extinguisher or extinguishers (§ 29.853(f)) are not visible to the passengers when seated, locating signs will be required. See § 29.1561(b).

(5) FAR Part 25, Appendix F, Part 1, established in Amendment 25-32 (effective May 1, 1972) contains flammability test procedures that must be used when complying with § 29.853 of Amendment 29-17. Appendix F refers to sections of FAR Part 25 that do not coincide with sections of the FAR Part 29. To preclude confusion the following statements should be used to develop company test procedures that will provide for compliance with § 29.853.

(i) Section 29.853(a)(1) materials are tested (vertically) to procedures in Appendix F, paragraph (d), and the flame must be applied for 60 seconds and then may be removed.

(ii) Section 29.853(a)(2) materials are tested (vertically) to procedures in Appendix F, paragraph (d), and the flame must be applied for 12 seconds and then may be removed.

(iii) Section 29.853(a)(3) and (4) materials are tested (horizontally) to procedures in Appendix F, paragraph (e), and the flame must be applied for 15 seconds and then may be removed.

(iv) Appendix F, paragraph (h) contains criteria for burn length measurement.

(v) Appendix F, paragraph (f) contains a procedure that does not apply to FAR Part 29, certification rules through Amendment 29-19.

(vi) Appendix F, paragraphs (a), (b), and (c) contain appropriate test procedures. It is noted § 29.853(a)(4) materials are equivalent to the materials specified in § 25.853(b-3).

(vii) Electrical wire and cable materials are tested in accordance with FAR 25, Appendix F, paragraph (g). (Refer to §§ 29.1351(d)(3), 29.831, and 29.863, and possibly special conditions for some rotorcraft.)

Page D - 114 9/30/99 AC 29-2C (6) AC 23-2, Flammability Tests, dated August 20, 1984, pertains to small airplanes and their materials. This AC includes information from Flight Standards Service Release No. 453 and may be useful in preparing a test proposal for flash-resistant, flame-resistant, fire-resistant, and fireproof materials. FAR Part 1 contains a further definition of these four terms.

(7) The “fire blocking layer” features of the seat cushions must be tested as prescribed in Appendix F, Part II, Part 25. Specific test equipment and devices are prescribed. AC No. 25.853-1, Flammability Requirements for Aircraft Seat Cushions, dated September 17, 1986, provides guidance material for demonstrating compliance with the seat cushion flammability standards.

AC 29.855. § 29.855 (Amendment 29-24) CARGO AND BAGGAGE COMPARTMENTS .

a. Explanation . This section contains standards for accessible and inaccessible compartments. The rotorcraft should be able to contain a fire until it is detected and extinguished or until a safe landing and evacuation are accomplished. The cabin may be used as a cargo compartment for rotorcraft used for carriage of cargo only.

Protective breathing equipment is required (§ 29.1439) for an appropriate crewmember or crewmembers when a compartment is accessible in flight. The rule does not provide for classification of cargo compartments. Reference is made to § 29.853 for flammability standards of certain materials.

(1) The compartment must be constructed of, or lined with, materials that are at least fire resistant. Accessible and inaccessible compartments must comply.

(2) Inaccessible compartments must be sealed and designed to completely contain a compartment fire or to allow detection as stated in § 29.855(c) and (d).

(3) Inaccessible compartments must have a detector unless the compartment can contain a fire as stated. Accessible compartments must have a detector or be designed to ensure detection by a crewmember while at his station as stated in § 29.855(d). Flight evaluations assure that an inaccessible compartment is sealed and will contain smoke, gases, etc., as stated.

(4) The cabin area may be used for carriage of cargo only as stated in § 29.855(e). Crew emergency exit must be accessible; sources of heat protected, and air flow must be stopped.

(5) Section 29.853 of Amendment 29-17 provides flammability standards for cargo compartment liners, covers, cargo, baggage tiedown equipment, etc., as stated in that section. This section pertains to compartments used by passengers or crew.

Section 29.855(a) requires a fire resistant liner and is the overriding requirement.

b. Procedures . It is intended to provide for adequate protection of the crew and passengers in the event of an in-flight fire. For Category B rotorcraft, one objective as Page D - 115 AC 29-2C 9/30/99 stated in § 29.861 is that the rotorcraft should be protected for at least 5 minutes (after recognition) in the event of a fire. The correct time interval to consider for Category A or B rotorcraft may be derived from the policy stated in paragraph AC 29.861, § 29.861.

(1) An aluminum inner skin, fire resistant liner, or closure of the compartment, whether the compartment is accessible or inaccessible is required by the rule. In the event of a compartment fire, the inner skin or liner will protect the load-carrying structure from direct flame impingement until the fire is detected and appropriate action is taken.

Flight Standards Service Release No. 453 provides the standards for fire resistant materials.

(2) Inaccessible compartments, in addition to having the inner skin or liner, must be sealed to prevent entry of air and thereby contain a fire in the compartment.

Flight tests are generally necessary to assure the compartment, primarily doors, do not leak in flight. Sensitive pressure measuring equipment (range of 10 inches of H O) may be used to prove the compartment is sealed by finding no appreciable change in compartment pressure during ground and flight conditions. The appropriate tests should also be conducted to determine that no accumulation of harmful quantities of smoke, flame , extinguishing agents, or other noxious gases occur in any crew or passenger compartment. For compartments having a volume not in excess of 500 cubic feet, an airflow of not more than 1,500 cubic feet per hour is considered acceptable. For larger compartments lesser airflow may be applicable to assure fires are contained.

(3) Inaccessible compartments may have a detector as prescribed. A smoke detector is preferable in place of a fire detector. The instrument panel will have an illuminated red indicator, such as baggage/cargo, as a warning signal for the flightcrew.

Although no specific standards for the detectors are contained in FAR Part 29, the following standards are recommended. The detection system should be designed to provide a visual indication to the flightcrew within one minute after start of a fire or within 5 minutes after smoke initiation appropriate to the detector used (30 seconds is allowed under TSO C 1b, for smoke detector actuation). There should be a means to allow the crew to check in flight the functioning of each fire or smoke detector circuit. For large compartments, the effectiveness of the detection system should be proven and the detection system should be capable of detecting a fire at a temperature significantly below the temperature at which the structural integrity of the rotorcraft would be substantially decreased.

(4) Accessible compartments must have a detector or detectors unless a crewmember can detect a fire while at his station. Flight evaluations are necessary to assure accessible compartments may be isolated from crew and passenger compartments as stated. The rule envisaged separate compartments for passengers or crew and cargo/baggage.

(5) Insulation blankets, cargo covers, cargo and baggage tie-down equipment, including containers, bins and pallets used in accessible and inaccessible Page D - 116 5/1/2014 AC 29-2C, Chg 4 compartments should meet the flammability standards specified in § 29.853 for the same counterparts noted therein.

AC 29.855A. § 29.855 (Amendment 29-30) CARGO AND BAGGAGE COMPARTMENTS.

a. Explanation . Amendment 29-30 relaxes previous requirements by allowing small, accessible cargo and baggage compartments to be lined with passenger compartment materials rather than fire resistant materials. Materials may meet the § 29.853(a)(1), (a)(2), and (a)(3) requirements for cargo or baggage compartments if: (1) The presence of a compartment fire would be easily discovered by a crew member while at the crew member’s station.

(2) Each part of the compartment is easily accessible in flight.

(3) The compartment has a volume of 200 cubic feet or less.

b. Procedures . The previous procedures continue to apply to Amendment 29-30 except for allowing the use of passenger compartment materials for accessible compartments.

AC 29.859. § 29.859 (Amendment 29-2) COMBUSTION HEATER FIRE PROTECTION.

a. Explanation . This regulation ensures that onboard combustion heating systems (of all type designs) are safe during normal and survivable emergency operations. Thus as a minimum, each combustion heater design must meet the requirements of § 29.859.

b. Definitions .

(1) Backfire . An improperly timed detonation (or explosion) of a fuel mixture which results in higher than normal temperatures and pressures.

(2) Reverse flame propagation . An event that occurs when the flame from a controlled combustion process such as a heater, goes in an abnormal path (i.e., either a reverse or different path than the intended path) as a result of a change in internal pressure or internal pressure gradient (e.g., a backfire) from a detonation or a similar event.

(3) Safe distance . A maximum flow length dimension determined from the thermodynamics of a worst-case flow reversal (backfire) and the local heater system geometry.

(4) Heater zone (or region) . A geometric zone defined by the heater type, heater size, the location of heater system components, and the maximum safe distance Page D - 117 AC 29-2C 9/30/99 determined under (3) above. The heater system components may affect the heater zone’s size if they are closely located to the heat source. For example a heater fuel tank would not be part of the heater zone if it were located far away from the zone boundary; however, if it were adjacent or close to the boundary, it would be included in the heater zone.

(5) Fireproof . Fireproof is defined in § 1.1, “General Definitions.” (6) Severe Fire . The following thermodynamic definitions are based on AC 20-135, “Powerplant Installation and Propulsion System Component Fire Protection Test Methods, Standards and Criteria” and on the definitions in § 1.1 for fire resistant and fireproof materials. These definitions are provided for analytical purposes. A severe fire, when used with respect to fireproof materials, is one which reaches a steady state temperature of 2,000 ±150° F for at least 15 minutes. A severe fire, when used with respect to fire resistant materials, is one which reaches a steady state temperature of 2,000 ±150° F for at least 5 minutes.

(7) Hazardous accumulation of water or ice . An accumulation of water or ice that causes a device to not perform its intended function in either normal operation or a survivable emergency situation.

c. Procedures . When suitable data is available, the heating system design should be thoroughly reviewed to determine which system components and arrangements must comply with each subsection of § 29.859. The method of compliance relative to each subsection of § 29.859 should then be determined. Acceptable, but not the only, methods of compliance are discussed on a section-by-section basis as follows.

(1) For compliance with § 29.859(a), combustion heater designs, their installations and their heater zones must be identified and thoroughly evaluated. The most direct method of compliance for the heater, itself, is to procure units that already have internal design features that meet the relevant requirements of this section; otherwise, design features should be provided and evaluated during certification that meet these same requirements. Several combustion heaters are approved under TSO-C20. TSO-C20 provides the procurement sources and the detailed approval standards for these combustion heaters. Each heater, its installation and its heater zone should be reviewed against the criteria of §§ 29.1181 through 29.1191 and §§ 29.1195 through 29.1203 (reference paragraphs AC 29.1181 through AC 29.1191 and AC 29.1195 through AC 29.1203) to ensure compliance. Next, the fire detector installation drawings and specifications should be reviewed for each heater region. The review should consider all reasonable hazards and failure modes of the heater and the detection system, itself. If not previously TSO approved, the detectors themselves should be evaluated and approved during the overall system certification effort. Then, the drainage and venting system for each heater installation should be reviewed to ensure that areas of fuel or fuel vapor collection are properly drained or vented. The capacity of each drain or vent should be determined and, unless impracticable, the flow capacity should be a minimum of 3-to-1 over the worst-case leakage anticipated Page D - 118 9/30/99 AC 29-2C clogging should be considered. Finally, the drainage and ventilation systems should be reviewed to ensure that discharges do not create external hazards by entering or contacting external ignition sources such as engine inlets and hot exhausts. If an accurate determination cannot be made by a design review, ground and/or flight test work with dyed, inert fluids or vapors should be conducted to accurately display discharge patterns.

(2) For compliance with § 29.859(f), the ventilating air duct design should be reviewed to determine what ducts are routed through heater zones. Once this has been determined, each duct section running through the heater zone should be made fireproof by either using a fireproof shroud around the existing duct or by using fireproof material for the duct wall. A primary purpose of these certification measures is to eliminate any system leakage that would allow carbon monoxide (a poisonous gas) to enter occupied areas, incapacitate the crew or passengers, and cause a crash.

Regardless of the method-of-compliance chosen, periodic checks should be performed during certification using carbon monoxide detection equipment to certify the leak-free integrity of the system. Several such checks should be done during flight test, especially after rigorous maneuvers, to ensure no leakage. It is also recommended that periodic checks using a carbon monoxide detector be conducted in conjunction with phased visual inspections (typically at a less frequent interval than each visual inspection) to ensure continued airworthiness. Carbon monoxide tests are reliable and quickly accomplished without any system disassembly. Continued airworthiness considerations are very important since carbon monoxide is a colorless, odorless, tasteless, poisonous gas that incapacitates an occupant without warning. Carbon monoxide’s ability to incapacitate increases with altitude, and has long been suspected as a probable cause for many aircraft accidents. It is the subject of General Aviation Airworthiness Alert No. 137, dated December 1983.

(3) For compliance with § 29.859(c), any design using combustion air ducts should be reviewed to ensure that the ducts are either made from fireproof material or shrouded with a fireproof shroud over a safe distance (see definition). The safe distance should be determined analytically, by test, or a combination, if the analytical results are not conclusive. The design should be reviewed to ensure that combustion air ducts are not connected to the ventilating airstream, except when an informal equivalent safety finding can be made that shows backfires or reverse burning cannot induce flames or fumes into the ventilating airstream under any failure condition or malfunction of the heater or its associated components. Such a finding should require analysis, testing, or a combination for a proper determination. A hazard FMEA should be conducted to ensure that no flames or fumes can be induced under any failure mode.

(4) For compliance with § 29.859(d), the design and installation of all standard heater control components, control tubing and safety controls should be reviewed to determine the probable points of water or ice accumulation (e.g., sumps, rough surfaces, joints, etc.). If a design review cannot accurately determine these accumulation points, then bench tests and flight tests should be conducted for proper determination. Once these points are identified, the ability of the effected part (or parts) Page D - 119 AC 29-2C 9/30/99 to perform its intended function when water or ice has fully accumulated must be determined for both normal and survivable emergency operations. If the part (or parts) either has not lost its ability to function; has lost part of its ability to function; or has lost all of its ability to function; and the entire system’s function is not impaired, then nothing further should be required. However, if the overall system’s function is hazardously impaired or lost, as a result of water or ice accumulation on a part (or parts), then rectifying design improvements should be made prior to final approval. These improvements should either alter the part’s environment (e.g., relocation, enclosure, insulation, etc.) or eliminate the hazardous accumulation of water or ice (e.g., provide drainage, better sealing, better location, different surface finish, etc.).

(5) For compliance with § 29.859(e), combustion heaters, if used, must have separate, independent safety controls from their standard controls (e.g., air temperature, air flow, fuel flow, etc.) which are remotely located in case of a heater fire, are operable by the crew and automatically shut off the ignition and fuel supply when a hazardous condition exists, (as defined by § 29.859(g)). These separate safety controls must comply with § 29.859(g)(1), must keep the heater off until restarted by the crew or ground maintenance, and must warn the crew when an essential heater is automatically shut down. The safety control system design should be thoroughly reviewed and tested to ensure that it complies and that no hazardous failure modes exist. An FMEA should be conducted to ensure proper compliance.

(6) For compliance with § 29.859(f), each combustion and ventilating air intake’s location should be identified, reviewed, and tested to ensure that no flammable fluids or vapors can enter the heater system, ignite and create a fire. If a combustion or ventilating air intake’s location is critical or questionable, it should be relocated, shielded, drained, or other equivalent means provided to eliminate the potential fire hazard. If engineering analysis and evaluation are not adequate to make an acceptable safety finding, testing using dyed, inert, leaked fluids or vapors should be conducted.

(7) For compliance with § 29.859(g), each heater exhaust system design should be reviewed, tested, or a combination to ensure proper compliance with § 29.1121 and § 29.1123 (reference paragraphs AC 29.1121 and AC 29.1123, respectively). Each exhaust shroud should be sealed to ensure that leaked flammable fluids or vapors do not contact the hot exhaust and cause a fire. The seal design should be reviewed to ensure that the sealing material is fireproof, is chemically compatible with the relevant fuels and vapors, is durable and is functionally adequate. If the design review is not conclusive for compliance purposes, then the seal system should be bench tested under pressure while undergoing critical service loads and motions to ensure no leakage occurs. Phased seal inspections should be considered to ensure continued airworthiness. An analysis should be conducted to determine the structural effects on the exhaust system of the worse case restricted backfire (typically a shock wave analysis can be used to determine the peak internal pressure and the resultant load on the exhaust system.) If structural failure would occur, based on the analysis, either the backfire restriction should be reduced or the exhaust design should be structurally improved to eliminate the failure.

Page D - 120 9/30/99 AC 29-2C (8) For compliance with § 29.859(h), each heater’s fuel system design must be reviewed to ensure compliance with the powerplant fuel system requirements of Part 29 that are necessary for safe operation to be achieved. An equivalent safety finding should be made if an application is received that requests partial compliance or non-compliance with the powerplant fuel system requirements of Part 29. The finding should ensure that the safety intent of § 29.859(j) is achieved. Analysis, engineering evaluation, testing, or a combination should be used to substantiate the heater fuel system design. Heater fuel system components that, by leakage or other failures, can induce flammable fluids or vapors into the ventilating air stream should be shrouded by drainable, fireproof shrouds.

(9) For compliance with § 29.859(i), the drain system design should be reviewed to identify parts that may be subjected to high temperature and parts that may be subjected to hazardous ice accumulation in service. The high temperature parts should be evaluated using the methods of compliance for heater exhausts (reference paragraph c(7), above and paragraph AC 29.1123). Drains that would be stopped up from ice accumulation should be protected by relocation, size, shields, heating, or a combination to ensure hazardous fluids and vapors are properly drained away.

AC 29.861. § 29.861 FIRE PROTECTION OF STRUCTURE, CONTROLS, AND OTHER PARTS .

a. Explanation .

(1) As stated in the rule, a Category B rotorcraft must be controllable until landed and a Category A rotorcraft must be controllable and continue its flight after a powerplant fire. For Category B rotorcraft designs with Category A powerplant isolation or Category A rotorcraft, a powerplant fire in one engine compartment must not adversely affect the remaining engine or engines. (Refer to § 29.903(b)). A policy statement on powerplant fire protection provisions was contained in the following note that appeared after Civil Air Regulation (CAR), Part 7, § 7.480, Designated Fire Zones.

NOTE: For Category B rotorcraft, the powerplant fire protection provisions are intended to ensure that the main and auxiliary rotors and controls remain operable, that the essential rotorcraft structure remains intact, and that the passengers and crew are otherwise protected for a period of at least 5 minutes after the start of an engine fire to permit a controlled autorotative landing.

(2) To achieve the objectives of the rule, each part of the rotorcraft, as stated in the rule, must be isolated from a powerplant fire by a firewall (§ 29.1191), or for (i) Category A, must be fireproof and must also comply with § 29.903(b).

(ii) Category B, must be protected so that they can perform their essential functions for at least 5 minutes under any foreseeable powerplant fire condition.

Page D - 121 AC 29-2C 9/30/99 Review Case No. 26 pertains to CAR, Part 6, §§ 6.384 and 6.483. These rules were replaced by §§ 27.861 and 27.1191 respectively. Even though these rules pertain to normal category rotorcraft requirements, the objective statements contained in the review case pertain to the interpretation of the time interval specified by CAR, Part 7, § 7.384(b) and the note under CAR, Part 7, § 7.480 for Category B rotorcraft. These rules have been replaced by § 29.861(b) and § 29.1191, respectively. In the review case, the FAA stated, in part, that the firewall must be fireproof, support appropriate flight and landing condition loads, and prevent flame penetration when subjected to a flame of 2000° F for 15 minutes. Essential structure and controls must be protected for the duration of time appropriate to the rotorcraft operation and be able to carry loads and resist any failure that could cause hazardous loss of control when subjected to the temperature resulting from any foreseeable powerplant fire. Insufficient protection to provide enough time for a controlled landing would represent an unsafe feature or characteristic for the rotorcraft design.

(3) In addition, paragraph AC 29.1193 (§ 29.1193(c)) pertains to allowable opening in engine cowls and to fireproof skins in specified cases.

b. Procedures .

(1) If each part described in the rule is isolated completely by firewalls, compliance is obtained for Category A or B.

(2) If each part described by the rule is made of fireproof material such as steel, compliance is obtained for Category A or B.

(3) For some Category A rotorcraft, § 29.903(b) also imposes additional considerations where structure, controls, and other parts are common to the engine installation. For example, an interconnected engine mount must be fireproof and also perform its function and not affect the remaining engine in case of a powerplant fire. An evaluation should involve propulsion and airframe disciplines.

(4) For Category B certification, if each part described by the rule does not comply as stated in (1) or (2), it must be proven that it will perform its function under the prescribed conditions. Compliance for Category B may be demonstrated by the following criteria: (i) The parts must have a positive margin of safety for the appropriate flight and landing condition, including appropriate engine power conditions, under any foreseeable powerplant fire condition. The time interval under consideration here is the time necessary to complete an emergency descent (as described in the flight manual) and landing from the maximum operating altitude for which certification is requested. In no case is the total time interval to be less than 5 minutes.

(ii) The factors affecting the time interval should include the maximum height above the terrain, the maximum operating altitude, the flight manual Page D - 122 9/30/99 AC 29-2C recommendations for rate of descent, and a reasonable time for recognizing a powerplant fire.

(iii) The factors affecting the change in physical characteristics (strength primarily) of the parts are the temperature of the part, time interval at the elevated temperature, size, heat absorption or rejection.

(iv) The factors affecting the temperature of the part are location and distance from fire and flames, and temperature of the flames (2,000° F ±50° F should be used unless proven otherwise).

(v) The rule requires substantiations for any foreseeable powerplant fire condition. Each rotorcraft design is unique and an evaluation of each design is necessary to establish the fire and flight conditions under consideration.

(vi) A very brief and simple example of compliance noted here may be helpful. This example pertains to a single engine Category B rotorcraft with the engine mounted on top at the fuselage center line. The engine is supported by all steel tubular mounts. The fuselage panel serves as a work deck as well as a firewall. A 15-minute duration is appropriate for this design. A representative panel of the firewall (deck) skin may be subjected to the autorotation flight loads and the landing load. A flame from an appropriate size burner, measuring 2,000° ±50° F at the skin surface, should impinge on the loaded panel for 15 minutes. The panel may deform but must remain intact and sustain the appropriate load. The flame must not penetrate the panel skin.

(vii) Other rotorcraft designs may have engines located on top of the fuselage and under the main rotor. If cowls or firewalls do not isolate the rotors and essential controls, it must be determined by a rational analysis or by temperature measurement that the rotor and essential controls will perform their functions. Air flow through the rotor and factors noted in paragraphs b(4)(ii), (iii), and (iv) are important to an analysis. Compliance with § 29.1193(e)(3), fireproof skins will involve airframe and propulsion disciplines for rotorcraft designs that do not have cowls.

Page D - 123 AC 29-2C 9/30/99 AC 29.861A. § 29.861 (Amendment 29-30) FIRE PROTECTION OF STRUCTURE, CONTROLS, AND OTHER PARTS .

a. Explanation .

(1) Amendment 29-30 revised the standard for Category B rotorcraft to allow use of parts made from standard fireproof materials of known acceptable dimensions in areas affected by powerplant fires without further proof of qualification. Previously the standard imposed a performance criterion for Category B applications regardless of the materials and part dimensions used.

(2) Fireproof and fire resistant are defined in FAR Part 1, § 1.1.

b. Procedures .

(1) A part with acceptable geometry that is made of steel, or another fireproof material, may be used to comply with the standard.

(2) A material system, panel, or assembly would be equivalent to steel provided it successfully completes the flammability tests described in paragraph AC 29.861b4(vi) for the appropriate time period that includes fire recognition.

AC 29.863 § 29.863 (Amendment 29-17) FLAMMABLE FLUID FIRE PROTECTION .

a. Background.

(1) The development of current § 29.863 can be traced through CAR 7.483, § 29.863 (1968), NPRM 68-18 (1968), and NPRM 75-26 (Airworthiness Review Notice November 7, 1975) and subsequent Amendment 29-17.

(2) Investigation of two accidents disclosed evidence of in-flight fires caused by leakage of flammable fluids to ignition sources. The revisions to § 29.863 adopted by Amendment 29-17 require significantly more attention to overall fire protection and prevention.

b. Explanation .

(1) Prior to Amendment 29-17, this rule only required either a means to prevent ignition of flammable fluids or vapors or a means to control any resulting fire. Isolation of flammable fluids and vapors from ignition sources by shrouding or sealing was the normal method of compliance. The revised rule further requires the assumption that these means fail or are ineffective and a fire does actually occur. Means to minimize the consequence of these fires should be provided. Specifically identified considerations should include the flammability of any combustible or absorbing materials, electrical faults, malfunction of protective devices, etc.

Page D - 124 5/1/2014 AC 29-2C, Chg 4 (2) The rule does not go so far as to require the entire rotorcraft to be a “designated fire zone.” Zonal analysis of areas containing flammable fluids may be used to show compliance with the requirements of this section. The general philosophy to be adopted for demonstrating compliance with § 29.863 is illustrated in AC 29.863-1.

c. Methods of Compliance .

(1) To minimize the probability of ignition of fluids and vapors after single failure of a component or systems, the following methods may be used. In considering compliance, the actual extent of protective measures required may be related to the situation considering the quantity and flammability characteristics of the fluid, the fire damage tolerance of the area, and the means available to the crew to minimize hazards from a fire.

(i) Shroud and drain flammable fluid systems (including steel fluid lines), fittings, etc. and/or provide fuel and vapor seals with respect to ignition sources (electrical wiring and equipment, hot bleed air lines, etc.). Drains should be designed and positioned to enable systems to be drained until any remaining flammable fluid residue is negligible. The arrangement of drains should be such that the discharge of flammable fluid from the outlet would not constitute a fire hazard, nor could flammable fluid or vapor enter personnel compartments or other portions of the rotorcraft where a hazard of ignition may exist. If flammable fluid drains are routed through personnel compartments, means for protection from damage should be provided to prevent possible entry of flammable fluids or vapors into these compartments.

(ii) Provide other effective separation, ventilation, or overheat shutdown devices, etc., to preclude ignition. Systems using flammable fluids should be separated from potential sources of ignition, including equipment or parts with hot surface temperatures above the ignition temperature of the fluid, such that the risk of fires as a result of leakage or bursting of the fluid system is minimized.

(iii) Ensure that potential ignition sources, such as bleed air lines and electrical equipment in the areas subject to flammable fluids and vapors is either hermetically sealed, shrouded, insulated or ventilated as necessary to minimize the possibility of ignition, or has been tested and shown to be free of ignition capability.

(iv) Place a restricting orifice in fluid pressure lines routed to instruments and transducers.

(v) Ensure flammable fluid carrying lines and drain lines are not located so as to be subject to abrasion during normal operations. Cargo compartments should be evaluated for potential line damage due to cargo movement.

(2) To minimize the hazards if ignition occurs: (i) Provide fireproof designs, fire wall isolation, or equivalent means for critical structure, equipment and personnel areas, e.g.: Page D - 125 AC 29-2C 9/30/99 (A) Flammable fluid lines and reservoirs of flammable fluids should be adequately protected against the anticipated type and duration of fire should ignition occur. Drain lines and their fittings, the failure of which would not result in, or add to, a fire hazard need not be Fire-resistant.

(B) Where there is a risk of leaking flammable fluids re-entering the rotorcraft through joints in the cowling or other rotorcraft surfaces to areas where a hazard of ignition may exist, the ventilation of such compartments should, where practical, be arranged to provide an air pressure within the compartment higher than that of the pressure of the ambient air.

(C) Absorbent materials in areas where leakage or spillage of flammable fluids (i.e., liquids, vapors, gases) could occur as a result of normal operation, failures of the equipment, or leakage from joints or unions should be covered or treated to prevent the absorption of hazardous quantities of fluid. Whenever insulation made of absorbent materials is used on pipes, tanks or equipment containing Flammable fluids, suitable precautions should be taken to prevent the wetting of the insulation by Flammable fluids.

(D) All electrical equipment including cables and their accessories should, as far as is practicable, be constructed of material which do not support combustion and which meet the relevant requirements of FAR/JAR25 Appendix F Part 1. Other materials should be applied and/or protected so that the risk arising from a fire is not increased by their use.

(E) It should be shown by analysis and/or tests that there are adequate means to prevent hazardous quantities of smoke, flame, extinguishing agents or other noxious gases produced as a result of a fire from entering any crew or passenger compartment.

(F) All components of the overheat or fire detector system (if applicable) should be at least Fire resistant.

(G) If located in an area where flammable fluids are present, critical structural components, controls, and essential indicating systems required for safe flight must be able to withstand the conditions resulting from a flammable fluid fire in the area so that a safe landing may be made. Under these conditions the structural members and the control devices should be able to carry the loads appropriate to the expected maneuvers including any vibrations normally experienced in flight. The quantity of flammable fluid likely to be present assuming all fluid drains function correctly and the maximum temperature characteristics of the particular flammable fluid may be taken into account in the analysis of the effect of a fire on critical structure. When making the determination that these components can withstand the flammable fluid fire conditions, the time required to detect such a fire should be taken into account.

Page D – 126 9/30/99 AC 29-2C (ii) In considering compliance, the actual protective measures required may be related to the situation considering the quantity and flammability characteristics of the fluid, the fire damage tolerance of the area, and the means available to the crew to minimize hazards from the fire. Provisions for fire detection, extinguishment, shutoff valves, fire suppression systems, etc., may be considered as alternate means to limit the duration of a fire in lieu of the protective measures listed in (2)(i) above. However, the consequences of spurious or erroneous operation of these systems should also be considered when evaluating the requirement for their provision. If action by the crew is necessary, quick-acting means (not necessarily fire detectors) should be provided to alert the crew in the event of a fire. Details of any action required by the crew must be in the Rotorcraft Flight Manual in accordance with § 29.1585(a).

(3) Compliance with § 29.863(d) requires as a minimum, type design data defining each area where flammable fluids or vapors might escape.

Page D - 127 AC 29-2C 9/30/99 Completed Design Review Are Systems No Containing Flammable Fluids Present?

Yes Could Flammable Fluids be Yes No Present After Leakage or Draining From Other Areas?

Could Ignition Sources Be Present, No Before or After FAR Section 29.863 Not System/Component Applicable Failure?

No Hazard Yes Minimize the Probability of Ignition, Reference Paragraph Stop AC 29.863c(1) of AC 29-2B Minimize the Hazard if Ignition does Occur, Reference Paragraph AC 29.863c(2) of AC 29-2B Stop, FAR Section 29.863 Satisfied FIGURE AC 29.863-1 Page D - 128 5/1/2014 AC 29-2C, Chg 4 SUBPART D - DESIGN AND CONSTRUCTION EXTERNAL LOADS.

AC 29.865. § 29.865 (Amendment 29-12) EXTERNAL LOAD ATTACHING MEANS.

a. Background . The external load attaching means standards for transport and normal category rotorcraft were originally contained in Subpart D, “Airworthiness Requirements of FAR Part 133, Rotorcraft External-Load Operations.” Amendment 29-12, in 1977, added a new § 29.865, which moved these standards from Part 133 to Part 29. An identical transfer occurred in 1977 for Part 27. Transport Category A and B rotorcraft were initially used under Part 133 operations and, after Amendment 133-6, restricted category rotorcraft were also included under Part 133 operations. The use of restricted category first came about when an operator, exempt from Part 133, transferred harbor pilots to and from ships by a hoist and sling. The exemption was granted to study the feasibility of passenger transfer outside of the cabin. Subsequently, Amendment 133-9, adopted in January 1987, established a new Class D rotorcraft load combination for transporting passengers external to the rotorcraft. Amendment 133-9 also provided for the limitations and conditions for external passenger transportation and the necessary, associated safety requirements.

Part 29 rules have not yet been changed to reflect the Class D requirements.

b. Explanation . While the regulation only addresses external load attaching means, this advisory material also includes guidance for certification of external load carrying devices for rotorcraft to be used in conjunction with Part 133, “Rotorcraft External Load Operations.” Subpart D of Part 133 contains supplemental airworthiness requirements. Part 1 defines four classes of rotorcraft load combinations which are operationally approvable under the Part 133 operating rules and, thus, are eligible for certification under § 29.865. Parts 1 and 133 (through Amendment 133-9) contain a new rotorcraft load combination, Class D, that addresses personnel carried externally.

The four classes of rotorcraft load combinations are summarized in FIGURE AC 29.865-1 and are discussed in detail in paragraph c. For further information, AC 133-1A, “ Rotorcraft External-Load Operations in Accordance with Federal Aviation Regulations Part 133 ,” October 16, 1979, may be reviewed. Also, paragraph AC 29.25 (reference § 29.25) concerns, in part, jettisonable external cargo.

c. Procedures .

(1) The applicant should clearly identify the Parts 1 and 133 rotorcraft load combination classes (A, B, C, or D) that are being applied for. The loads and operating envelopes for each class should be determined and used to formulate the flight manual supplement and basic loads report. The applicant should show by analysis, test, or both, that the rotorcraft structure, the external load attachment means, and (for Class D operations) the personnel carrying device meets the requirements of §§ 29.865(a), 133.41, 133.43, and 133.45(e)(3) for the proposed operating envelope.

Page D - 129 AC 29-2C 9/30/99 (2) For rotorcraft load combination classes A, B, and C, § 29.865 requires use of 2.5 g vertical limit load factor (N ) at the maximum substantiable cargo load (which ZW is typical for cargo hauling configurations). This 2.5 g limit load factor is based on an engineering evaluation and a rationalization of § 29.337 for high gross weight applications. However, for lower gross weight configurations (which are more typical of a Class D application; i.e., personnel transport or evacuation), a higher limit load factor is recommended to ensure that limit load is never exceeded in service. For example, a Class D external load carrying device which is certified to a limit vertical load factor of 2.5 g and is installed in a minimum gross weight configuration rotorcraft capable of generating a vertical limit load factor of 3.2 g’s could experience ((3.2/2.5 X 1.5) X 100) - 85 percent of ultimate load under emergency conditions with new external hardware. However, if factors such as wear and corrosion have effected the structural integrity of the external hardware ultimate load could be exceeded in emergency service. In any case, FAA/AUTHORITY policy is to not exceed limit load in service. The higher load factor for Class D cases should be the analytically derived maximum vertical limit load factor for the restricted operating envelope being applied for; or, as a conservative option, a vertical limit load factor of 3.5 g’s (reference § 29.337). Unless a more rational proposal is received, for Class D cases where maximum operating gross weight for external load is between design maximum weight and design minimum weight, linear interpolation can be used between N and ZW MIN N versus gross weight for design limit load factor determination.

ZW MAX (3) For applications that employ winches (or hoists) to raise or lower an external load from a hover (or another phase of flight), limit load must be properly determined based on the characteristics of the winch system and its installation such as mechanical advantage, static strength of the winch, static strength of its installation and the payload for any operating scenario being applied for. One acceptable method of determining limit load is by the following procedure: (i) Determine the basic loads that fail and unspool the winch or its installation, respectively (Note: This determination should be based primarily on static strength; however, any dynamic load magnification factors that are significant should be accounted for).

(ii) Select the lower of the two values from (i) as the ultimate load of the winch system installation.

(iii) Divide the selected ultimate load by 1.5 to determine the limit load of the system.

(iv) Compare the system’s derived limit load to the applied for one “g” payload multiplied by the maximum downward vertical load factor (N ) from ZW MAX paragraph (2) to determine the critical payload’s limit value.

(v) If the critical limit payload is equal to or less than the system’s derived limit load the installation is structurally approvable as presented.

Page D – 130 9/30/99 AC 29-2C (vi) If the critical limit payload exceeds the system’s derived limit load then one of the following options should be considered: (A) Disapproval.

(B) Application for exemption.

(C) Reduction of the applied for critical limit payload to less than or equal to the system’s derived limit load.

(D) Redesign of the winch system (and installation) to increase its derived limit load to equal to or greater than the critical payload.

(E) A combination of options (C) and (D).

(F) Approvable operating restrictions to reduce N and, the ZW MAX corresponding critical limit payload to less than or equal to the system’s derived limit load.

(4) In all approved cases, appropriate winch system placards and flight manual restrictions should be provided. Also, for Class D load combinations, the winch or hoist should have a demonstrated, acceptable level of reliability (for the phases of flight in which it is operable and in which the Class D load is carried externally). The winch should be disabled (or utilize an overriding mechanical safety device such as a flagged removable shear pin) to prevent inadvertent load unspooling or release during the phases of flight that the load is carried externally and operation is not intended. The maximum allowable winch cable angle should be determined and approved. This is primarily a structural requirement but should also be reviewed from an interference and flight handling criteria standpoint.

(5) It is recommended that winch or hoist systems be demonstrated as follows: (i) At least 1/3 of the demonstration cycles should include the maximum aft angular displacement of the load from the drum applied for under § 29.865(a).

(ii) The load versus speed combinations of the winch should be demonstrated by showing repeatability of the no load-speed combination, the 50 percent load-speed combination, the 75 percent load-speed combination and the system limit load-speed combination.

(iii) A minimum of six consecutive, complete operation cycles should be conducted at the system’s critical limit load speed combination.

(iv) In addition, the demonstration should cover all normal and emergency modes of intended operation and should include operation of all control devices, limit switches braking devices, and overload sensors in the system.

Page D - 131 AC 29-2C 9/30/99 (v) Quick disconnect devices, and cable cutters should be demonstrated at 25 percent, 50 percent, 75 percent, and 100 percent of system limit load. Any electrical load release devices for Class D loads should be treated as a novel design feature and should be coordinated with the Rotorcraft Directorate.

(vi) Any devices or methods used to increase the mechanical advantage of the winch should also be demonstrated.

(vii) During each demonstration cycle, the winch should be operated from each station from which it can be controlled.

(viii) Operating manuals, flight manuals, and associated placards should be used and proofed during the demonstration.

(6) For all applications, it is good practice to obtain the gross weight range limits, the corresponding limit load factors (N ), and substantiate the system, ZW accordingly, for the critical loads. This procedure determines the critical basic loads and associated operating envelope for the rotorcraft load combination categories requested.

(7) For a request involving more than one class of rotorcraft load combinations, structural substantiation is required only for the critical case if accurately determinable from analysis.

(8) Appropriate placards, markings, and flight manual restrictions should be provided as determined by load capacities and operational restrictions. Each placard, marking, and flight manual supplement should be checked during TIA flight testing.

(9) For load Classes A, B, C, and D, the basic vertical limit load factor (N ) ZW from (c)(2) is converted to ultimate by multiplying the maximum applied load (i.e., the sum of the carrying device load and cargo or personnel loads) by 1.5 (for restricted category approvals, see guidance in paragraph AC 29 MG 5.) This load is used to substantiate all existing structure affected and all added structure associated with the external load carrying device and its attachments. Casting and/or fitting factors are to be applied where appropriate. For load Class D, the weight of each occupant carried externally should be assumed, for analysis purposes, to be that of the 95 percentile (202 pound) man (reference MIL-STD-1472, “Human Engineering Design Criteria for Military Systems, Equipment and Facilities).

(10) For load Classes B, C, and D, the maximum limit external load for which certification is requested, even though it may otherwise be much less than the maximum system capacity; e.g., cargo hook capacity, etc., should not exceed the rated capacity of the quick release device used in the applicant’s proposed design or, for Class D only, the rated capacity of the personnel carrying device. The quick release and personnel carrying devices should be strength tested (with FAA/AUTHORITY witness) or otherwise structurally substantiated to determine their allowable limit load capacity, if it has not been previously approved or was not produced to a recognized and approvable industry or military standard.

Page D - 132 9/30/99 AC 29-2C (11) For load Classes B, C, and D, in substantiating analyses and tests, the maximum ultimate external load is specified to be applied at the sling-load-line to rotorcraft vertical axis (Z axis) angles up to 30°, except for the forward direction. The 30° angle may be reduced, if impossible to obtain due to physical constraints, or operating limitations. If the angle is reduced, appropriate placards and flight manual changes are required.

(12) For load Classes B and C, an external releasing system is mandated which requires an approved primary quick release device to be installed on one of the pilot’s primary controls. The quick release device (typically installed on the cyclic stick) is designed and located to allow the pilot to accomplish load release without hazardously limiting his ability to control the rotorcraft during emergency situations. A manual (backup) mechanical quick release device is also required. This control must also be readily accessible to the pilot or another designated crew member, such as a hoist operator. For Class B and C cargo applications, a sufficient amount of slack should be provided in the control cable to permit cargo hook movement without tripping the hook release.

(13) For Load Class D, an emergency release system is specified by § 133.45(e)(4) which requires two distinct actions for load release. This is intended for the phases of flight that the load is carried (and/or retrieved) externally. This release can be operated by the pilot from a primary control or, after a command is given by the pilot, by a dedicated crewmember from a remote location. Two distinct actions are required for the primary release to provide a higher level of safety for Class D human external loads. If the manual backup device is a cable cutter, it should be properly secured but readily accessible to the dedicated crewmember intended to use them.

(14) For Class D (human) load applications, to ensure personnel safety, the emergency release system design and associated placarding should be given special consideration. As stated previously, electrical release designs should be reviewed by the Rotorcraft Directorate prior to approval.

(15) For the majority of Class D applications, an approved single or multiple personnel carrier or container is required. The carrier or container may be previously approved or may be approved as part of the certification process. In any case, the single or multiple personnel carrier or container should be substantiated for the allowable ultimate load as determined under paragraphs c(2), (3), (4), (5), (6), (7), (8), and (9) above. The personnel carrier or container should be placarded for this capacity and show the proper internal arrangement and/or location of the intended occupants.

Page D - 133 AC 29-2C 9/30/99 Some exceptions may exist that are certifiable under Class D that involve the technique of “Rappelling” from a rotorcraft. Rotorcraft load-combination D allows for such applications by definition (reference § 1.1). Other types of human cargo devices can be applied for under the Class D external load combination definition. An example is external carriage of personnel in a conveyance rigidly attached to the rotorcraft (e.g., cage, pod, secured litter or strap harness/seat arrangement).

(16) The personnel carrier or container should be easily and readily ingressed or egressed. Appropriate placards are required to provide ingress and egress instructions. For door latch fail-safety, more than two fastener or closure devices are recommended. Direct visual inspectability of the latch device by both crew and passengers is recommended to ensure it is fastened and secured. Any fabric, if used, should be durable and should meet the flammability standards of safety belts as stated in TSO C-22. Sharp corners and edges should be avoided, and padding should be used when necessary to protect the carrier and container occupants.

(17) The U.S. Coast Guard has three containers or devices that are used with rotorcraft for emergency rescue work. These devices and their National Stock Numbers are listed below. These devices have not been FAA/AUTHORITY approved; however, applications which involve them may be submitted for approval.

National Stock No. Title 6530-00-042-6131 Stokes litter (one person) 1670-00-HR0-7970 Rescue basket 1680-090-511-2712 Rescue sling (one person) NOTE: The rescue sling is a “collar” device that requires a person to exert some effort to remain in the collar. This sling should only be used in conjunction with properly written instructions and with personnel trained in the proper use of the sling.

(18) Flight test verification work that thoroughly checks out the operational envelope should be accomplished with every device approved for external cargo carriage (especially rotorcraft load combination D which includes external human cargo). The flight test program should show that all aspects of the applied for operations are safe, uncomplicated and can be conducted by an average flight crew under the most critical service environment and, in the case of human external cargo, under the pressures of an emergency scenario.

AC 29.865A. § 29.865 (Amendment 29-30) EXTERNAL LOAD ATTACHING MEANS .

a. Explanation . Amendment 29-30 added two requirements to § 29.865: (1) Section 29.865(a) is clarified to allow use of a design factor less than 2.5g’s, for rotorcraft load combinations A, B, and C non-human external cargo applications provided the lower load factor is not likely to be exceeded by virtue of the Page D - 134 2/12/03 AC 29-2C, Chg 1 rotorcraft characteristics and capability. That is the rotorcraft design factors may be used for the cargo device system.

(2) Section 29.865(d) was added to clarify and specify the fatigue requirements for the external cargo attaching means. The “rotorcraft” standard is contained in § 29.571, paragraph AC 29.571.

b. Procedures .

(1) For § 29.865(a), if a design limit load factor less than 2.5g’s is requested, the applicant should provide a rational analysis and/or a flight operations data base that clearly shows that the load factor requested is unlikely to be exceeded in service.

Note: § 29.337(b) requires use of 2.0 g’s as a minimum.

(2) For § 29.865(d), all failures of the cargo attaching means (and the associated critical components) that are likely to be hazardous to the rotorcraft should be identified by an acceptable means such as an FMEA. The critical components associated with these failure modes should then be analyzed and/or tested to ensure that the likelihood of a fatigue failure or occurrence is acceptably minimized. In the majority of cases an analysis using the methods of AC 27 MG-11, “Fatigue Evaluation of Rotorcraft Structure”, will be sufficient. Any component’s airworthiness limitations and/or mandatory inspections should be identified by this analysis, approved, and placed in the airworthiness limitations section of the maintenance manual or Instructions for Continued Airworthiness. See paragraph AC 29.1529 (§ 29.1529) for information on these manuals.

Page D - 135 AC 29-2C, Chg 1 2/12/03 FIGURE AC 29.865-1 SUMMARY OF PART 133 ROTORCRAFT LOAD COMBINATIONS CERTIFIABLE UNDER § 29.865 CLASS A Basic Definition and Intended Use Fixed External Cargo Container Is defined by § 1.1 as a load combination in which the external load cannot move freely, cannot be jettisoned, and does not extend below the landing gear. This category usually features multiple attachments (loadpaths) to the airframe. Typical example is a hard mounted cargo basket attached to the rotorcraft crosstubes which is used to carry cargo from point A to point B.

Typical Load Limits Certification limit is N X Maximum Substantiable External load. N is 2.5 per § 29.865 (See ZW ZW Procedure, paragraph (2)).

Quick Release Requirements None. Cargo and its container are not jettisonable.

Certification Requirements -- Considerations - For cargo only.

- Flight Manual Restrictions - § 133.47 requires a rotorcraft load combination flight manual supplement. Any flight envelope restrictions from § 29.865 should be a part of this supplement.

- Load limit placards are required by § 29.865(c).

- Flight envelope restriction placards may also be required for gross weight limitations, e.g., limitations, elimination of dangerous maneuvers, etc.

- Cargo tiedowns to prevent load shifting relative to airframe may be required.

- Effect of external cargo carrier and its maximum cargo weight on load paths, loads and fatigue of existing structure should be determined.

- TIA testing may be necessary to determine whether or not the system performs as intended and if placards and flight manual supplements are adequate.

- The applicant may elect to test the aerodynamic effect of several representative load shapes and include applicable information in the flight manual supplement. If such information is not in the RFM, then the operator may be required to obtain an operations approval under Part 133.

Page D - 136 9/30/99 AC 29-2C FIGURE AC 29.865-1 (continued) SUMMARY OF PART 133 ROTORCRAFT LOAD COMBINATIONS CERTIFIABLE UNDER § 29.865 CLASS B Basic Definition and Intended Use Single Point Suspension External Load Airborne Is defined by § 1.1 as a load combination in which the external load is jettisonable and is lifted free of land or water during the rotorcraft operation. The payload is typically suspended from a hook or a similar device. The hook may be attached to the rotorcraft structure or it may be attached to a movable hoist cable and the hoist itself attached to the rotorcraft. Typical use is to lift a cargo load until it is completely airborne and fly it from point A to point B. The load on the hoist may be stowed in the fuselage (in some cases) while being transported.

Typical Load Limits Certification limit load is N X Maximum Substantiable External load. N is 2.5 per § 29.865 ZW ZW (See Procedure, paragraph (2)). Load may be limited by hoist allowables (reference paragraph (3)).

Quick Release Requirements § 29.865(b)(1) requires that a primary quick release system control device be installed on a primary control. Also, a manual quick release system backup actuation device must be available and readily accessible.

Certification Requirements -- Considerations - For cargo only.

- Flight Manual Restrictions - § 133.47 requires a rotorcraft load combination flight manual supplement. Any flight envelope restrictions from § 29.865 should be a part of this supplement.

- Load limit placards are required by § 29.865(c).

- Flight envelope restriction placards may also be required.

- Certifiable external cargo load capacity may be further limited by §§ 133.41 and 133.43 - Quick release devices must be approved and be operable on a nonhazard basis by the pilot per § 29.865(b).

- Manual backup must be reliable but need not be overly sophisticated (cable cutters, axes, etc., used by crew members) - Effect of maximum suspended load and its attachment to rotorcraft structure on load paths, loads and fatigue of existing structure should be determined.

- TIA testing may be necessary to determine whether or not the system performs as intended and if placards and flight manual supplements are adequate.

Page D - 137 AC 29-2C 9/30/99 FIGURE AC 29.865-1 (continued) SUMMARY OF PART 133 ROTORCRAFT LOAD COMBINATIONS CERTIFIABLE UNDER § 29.865 CLASS C Basic Definition and Intended Use Single Point Suspension External Load Partially Airborne Is defined by § 1.1 as a load combination in which the external load is jettisonable and remains in contact with land or water during the rotorcraft operation. The payload is typically partially suspended by a net or cables from a cargo hook or a similar device. The cargo hook may be attached to the rotorcraft structure or may be attached to a movable hoist cable and the hoist itself attached to the rotorcraft. Typically used for stringing wire or laying cable where the payload is only partially suspended from the ground. (Note: Many applications combine both Category B and C operations because of obvious utility involved.)

Typical Load Limits Certification limit load is N X Maximum Substantiable External load. N is 2.5 per § 29.865 ZW ZW (See Procedure, paragraph (2)). Load may be limited by hoist allowables (reference paragraph (3)).

Quick Release Requirements § 29.865(b)(1) requires that a primary quick release system control device be installed on a primary control. Also, a manual quick release system backup actuation device must be available and readily accessible.

Certification Requirements -- Considerations - For cargo only.

- Flight Manual Restrictions - § 133.47 requires a rotorcraft load combination flight manual supplement. Any flight envelope restrictions from § 29.865 should be a part of this supplement.

- Load limit placards are required by § 29.865(c).

- Flight envelope restriction placards may also be required.

- Certifiable external cargo load capacity may be further limited by §§ 133.41 and 133.43 - Quick release devices must be approved and be operable on a nonhazard basis by the pilot per § 29.865(b).

- Manual backup must be reliable but need not be overly sophisticated (cable cutters, axes, etc., used by a crewmember) - Effect of maximum suspended load and its attachment to rotorcraft structure on load paths, loads and fatigue of existing structure should be determined.

- TIA testing may be necessary to determine whether or not the system performs as intended and if placards and flight manual supplements are adequate.

Page D - 138 9/30/99 AC 29-2C FIGURE AC 29.865-1 (continued) SUMMARY OF PART 133 ROTORCRAFT LOAD COMBINATIONS CERTIFIABLE UNDER § 29.865 CLASS D Basic Definition and Intended Use Single Point Suspension External Airborne Personnel Load Is defined by § 1.1, as a load combination in which the external load is other than Class A, B, or C and has been specifically approved by the Administrator for that operation. This load combination includes human cargo. For human cargo operations, the payload which typically consists of personnel and their containment device is suspended from a hook or a similar device during all or part of a flight. The hook may be rigidly attached to the rotorcraft or may be attached to a movable hoist cable and the hoist itself rigidly attached to the rotorcraft. Typical use is for transfer of personnel to a ship. Carrying devices may transport one or more persons. Typical carrying devices are vest and straps, baskets, life preservers with straps and attachment devices, cages, or a suspended container.

Typical Load Limits Certification limit load is N X Maximum Substantiable External load. N varies from ZW ZW 2.5 at max gross weight to 3.5 at minimum gross weight. (See Procedures (2)). Load is usually limited by hoist allowable or by personnel carrying device allowable (See Procedure (2), (3), and (10)).

Quick Release Requirements Section § 29.865(b) does not currently contain quick release requirements for Class D rotorcraft - load combinations, but § 133.45(e)(4) requires that a primary emergency release system control device (requiring two distinct actions) be installed on a primary control or be installed near a designated crew member’s station. Also, a manual quick-release system backup actuation device must be available and readily accessible.

Certification Requirements -- Considerations - For loads other than Class A, B, or C loads. Is used for external personnel loads.

- § 29.865 has not been revised to reflect this category’s requirements (it is currently covered by § 133.45(e)(4) only).

- Unless a public-use rotorcraft is being certified, only transport Category A rotorcraft are eligible to use this load category.

- Transport Category A rotorcraft must be certified for an OEI weight and altitude envelope which becomes the maximum envelope that can be used for Class D operations. This is currently required for a Class D rating by § 133.45(e)(1).

Page D - 139 AC 29-2C, Chg 7 2/4/2016 FIGURE AC 29.865-1 (continued) CLASS D (continued) - Personnel lifting devices must be approved separately or as part of the certification project.

- Devices must carry personnel internally or secure them safely in a harness or equivalent device.

- Flight Manual Restrictions - § 133.47 requires a rotorcraft load combination flight manual supplement. Any flight envelope restrictions from § 29.865 should be a part of this supplement.

- Load limit placards are required by § 29.865(c).

- Flight envelope restriction placards may also be required.

- Certifiable external load capacity is further limited by §§ 133.41, 133.43 and 133.45(e)(3), the load limit of the personnel carrying device.

- Quick release devices must be approved and be operable on a nonhazard basis by the pilot or a designated crewmember per §§ 133.44(c)(6) and 29.865(b).

- The lifting device must have an emergency release requiring two distinct actions § 133.45(e)(4).

- Manual backup must be accessible and reliable.

- Rotorcraft must be equipped to allow direct intercom among all crewmembers per § 133.45(e)(2). This may affect § 29.865 indirectly if human error or placarding could cause inadvertent load release or retention.

- Effect of maximum suspended load and its attachment to rotorcraft structure on load paths, loads and fatigue of existing structure should be determined.

- TIA testing may be necessary to determine whether or not the system performs as intended and if placards and flight manual supplements are adequate.

AC 29.865B. § 29.865 (Amendment 29-43) EXTERNAL LOADS.

a. Background . The standards for external load attaching means for transport and normal category rotorcraft were originally contained in Subpart D--Airworthiness Requirements of 14 CFR part 133, “Rotorcraft External-Load Operations."

Amendment 29-12, issued in 1977, added a new § 29.865, which moved these airworthiness standards from part 133 to part 29. An identical transfer occurred in 1977 for part 27. Amendment 29-26, issued in 1990, clarified the intent of Amendment 29-12 but did not change it substantively. Transport categories A and B and normal category rotorcraft were initially used under part 133 operations, and after Amendment 133-6, restricted category rotorcraft were also included under part 133 operations. The carriage of persons external to the rotorcraft for hire first came about when a part 29 operator, exempt from part 133, transferred harbor pilots to and from ships by a hoist and sling.

The exemption was granted to study the feasibility of passenger transfer outside of the cabin. Grant of the exemption was based, in part, on similar, prior operations that had Page D - 140 2/6/2016 AC 29-2C, Chg 7 been conducted in Europe and Africa, for hire, with helicopters approved by the appropriate authorities and, in part, on similar military and public helicopter operations, not for hire, in the U.S. Subsequently, Amendment 29-43, adopted in October of 1999, established requirements for human external cargo (HEC). This amendment also removed references to rotorcraft load combinations (RLCs), as RLC approval is granted through the part 133 operational authority. Amendment 133-9 provided for the limitations and conditions for transport of external loads other than class A, B, or C and the necessary, associated safety requirements. Part 29 has recently been changed to reflect RLC class D requirements by requiring one-engine inoperative (OEI) hover performance and procedures to be included in the rotorcraft flight manual (RFM) or rotorcraft flight manual supplement (RFMS) for those operations that require such information. In addition, the scope and thus the title of the standard have changed from "External load attaching means" to "External loads" to reflect the more comprehensive approach for external loads required to assure the proper level-of safety.

b. Explanation .

(1) This advisory material contains guidance for the certification of helicopter external load attaching means and load carrying systems to be used in conjunction with operating rules such as part 133, "Rotorcraft External Load Operations." Subpart D of part 133 contains supplemental airworthiness requirements. 14 CFR part 1 defines the four RLC classes that are operationally approved under part 133 operating rules and eligible for airworthiness certification under § 29.865. Also, section 29.25 on this AC addresses, in part, jettisonable external cargo. Part 133 does not define nor permit fast- rope operations (an operation in which a person is not attached to an external means by a PCDS). Therefore, the FAA cannot certify installations for the purpose of fast-roping.

(2) Section 29.865 provides a minimum level of safety for transport category rotorcraft designs to be used with operating rules such as part 133. Certain aspects of operations such as microwave tower and high-line wirework may also be regulated separately by other Federal agencies such as DOE, EPA, and OSHA or by other international entities. For applications that could come under multiple agency regulation (or regulation by other entities), special certification emphasis will be required by both the applicant and the approving authority to assure all relevant safety requirements are identified and met. Potential additional requirements are noted herein.

(3) Except for § 29.865(c)(6), the airworthiness standard for external loads (§ 29.865) does not discern the difference between a crewmember and a compensating passenger when either are carried external to the rotorcraft. Both are considered HEC.

c. Definitions .

(1) Backup Quick-Release Subsystem (BQRS). The secondary or "second choice" subsystem used to perform a normal or emergency jettison of external cargo.

Page D - 141 AC 29-2C, Chg 7 2/4/2016 (2) Cargo. The part of any RLC that is removable, changeable, and is attached to the rotorcraft by an approved means. For certification purposes, ”cargo” applies to HEC and NHEC.

(3) Cargo hook. A hook that can be rated for both HEC and NHEC. It is typically used by being fixed directly to a designated hardpoint on the rotorcraft.

(4) Dual actuation device (DAD). This is a sequential control that requires two distinct actions in series for actuation. One example is removal of a lock pin followed by a "then free" switch or lever activation for load release to occur.

(5) Emergency jettison (or complete load release). The intentional, instantaneous release of NHEC or HEC in a preset sequence by the quick release system (QRS) that is normally performed to achieve safer aircraft operation in an emergency.

(6) External fixture. A structure external to and in addition to the basic airframe that does not have true jettison capability and has no significant payload capability in addition to its own weight. An example is an agricultural spray- boom. These configurations are not "External Loads" approvable under § 29.865.

(7) External Load System. The entire installation related to the carriage of external loads to include not only the hoist or hook, but also the structural provisions and release systems. The PCDS is considered part of the external load system only when the system is approved for HEC.

(8) Hoist. A hoist is a device that exerts a vertical pull, usually through a cable and drum system (i.e., a pull that does not typically exceed a 30-degree cone measured around the z-rotorcraft axis).

(9) Hoist demonstration cycle (or "one cycle"). The complete extension and retraction of at least 95 percent of the actual cable length, or 100 percent of the cable length capable of being used in service (i.e., that would activate any extension or retraction limiting devices), whichever is greater.

(10) Hoist load-speed combinations. Some hoists are designed so that the extension and retraction speed slows as the load increases or nears the end of a cable extension. Other hoist designs maintain a constant speed as the load is varied. In the latter design, the load-speed combination simply means the variation in load at the constant design speed of the hoist.

(11) Human external cargo (HEC). A person(s) that at some point in the operation is carried external to the rotorcraft.

(12) Nonhuman external cargo (NHEC). Any external cargo operation that at no time involves a person(s) carried external to the rotorcraft.

Page D - 142 2/6/2016 AC 29-2C, Chg 7 (13) Normal jettison (or selective load release). The intentional release, normally at optimum jettison conditions, of an NHEC.

(14) Personnel carrying device system (PCDS). A device or system that has the structural capability and features needed to safely transport occupants external to the rotorcraft during HEC operations. A PCDS includes, but is not limited to, life safety harnesses, and rigid baskets or cages either attached to a hoist or cargo hook or mounted to the rotorcraft airframe (see TSO C167).

(15) Primary Quick-Release Subsystem (PQRS). The primary or "first choice" subsystem used to perform a normal or emergency jettison of external cargo.

(16) Quick-release system (QRS). The entire release system for jettisonable external cargo, (i.e., the sum total of both the primary and backup quick-release subsystems). The QRS consists of all components including the controls, the release devices, and everything in between.

(17) Rescue hook (or hook). A hook that can be rated for both HEC and NHEC.

It is typically used in conjunction with a hoist or equivalent system.

(18) Rotorcraft-load combination (RLC). The combination of a rotorcraft and an external-load, including the external-load attaching means. The requirements for the different classes of RLC are contained in § 133.41. RLC classes are outside the scope of FAA design approval (i.e., § 29.865) as well as the associated rotorcraft flight manual or rotorcraft flight manual supplement (i.e., § 29.1581). RLC classes are only designated within an operational approval (i.e., § 133.17) and the rotorcraft-load combination flight manual (i.e., § 133.47).

(19) Spider. A spider is a system of attaching a lowering cable or rope or a harness to a NHEC (or HEC) RLC to eliminate unwanted flight dynamics during operations. A spider usually has four or more legs (or load paths) that connect to various points of a PCDS to equalize loading and prevent spinning, twisting, or other undesirable flight dynamics.

(20) True jettison capability. The ability to safely release an external load using an approved QRS in 30 seconds or less.

Note: In all cases, a PQRS should release the external load in less than 5 seconds.

Many PQRSs will release the external load in milliseconds, once the activation device is triggered. However, a manual BQRS such as a set of cable cutters could take as much as 30 seconds to release the external load. The 30 seconds would be measured starting from the time the release command is given and ending when the external load is cut loose.

Page D - 143 AC 29-2C, Chg 7 2/4/2016 (21) True payload capability. The ability of an external device or tank to carry a significant payload in addition to its own weight. If little or no payload can be carried, the external device or tank is an external fixture (see definition).

(22) Type inspection authorization (TIA). This is FAA Form 8110-1, used for authorizing official ground inspections and flight tests necessary to fulfill the requirements for type certification or supplemental type certification. Order 8110.4 contains the criteria for TIA issuance.

(23) Winch. A device that can employ a cable and drum or other means to exert a horizontal (i.e., x-rotorcraft axis) pull. However, in designs utilizing a winch to perform a hoist function by use of a 90-degree cable direction change device (such as a pulley or pulley system), the winch system is considered a hoist and must comply with the hoist requirements.

d. Procedures . The following certification procedures are provided in the most general form. Where there are significant differences between the cargo types, the differences are highlighted.

(1) General Compliance Procedures for § 29.865 . The applicant should clearly identify both the applicable cargo types (NHEC or HEC) and the jettison capability (jettisonable or non-jettisonable) for which application is being made. The structural loads and operating envelopes for each applicable cargo type should be determined and used to formulate the flight manual supplement and basic loads report. The applicant should show by analysis, test, or both, that the rotorcraft structure, the external load attachment means, and the PCDS, if applicable, meet the specific requirements of § 29.865 and the other relevant requirements of part 29 for the proposed operating envelope.

Note: The approved maximum internal gross weight should never be exceeded for any approved HEC configuration (or simultaneous NHEC and HEC configuration).

(2) Reliability of the external load system, including QRS .

(i) The hoist, QRS, and rescue hook system should be reliable for all phases of flight and the applicable configurations for those phases (i.e., operating, stowed, or unstowed), for which approval is sought. The hoist should be disabled (or an overriding, fail-safe mechanical safety device such as either a flagged removable shear pin or a load-lowering brake should be utilized) to prevent inadvertent load unspooling or release during any extended flight phases in which hoist operation is not intended. Loss of hoist operational control should also be considered.

(ii) Any failure mode of the external load system (including QRS, hook and attachments to the rotorcraft) leading to a loss of the HEC should be considered a Catastrophic event in accordance with the safety objectives in section 29.1309 of this AC. Uncontrolled high speed descent of the hoist cable would fall into this category.

Page D - 144 2/6/2016 AC 29-2C, Chg 7 Failures leading to the loss of NHEC should be evaluated and considered to be at least a Major event in accordance with the safety objectives in section 29.1309 of this AC.

(iii) The reliability of the system should be demonstrated by completion and approval of the following: (A) A functional hazard assessment (FHA) to determine the hazard severity of failures associated with the external load system. The effect of the flailing cable after a load release should be considered.

(B) A fault tree analysis (FTA) or equivalent to verify the hazard classification of the FHA has been met.

(C) A system safety assessment (SSA) to demonstrate compliance to the applicable certification requirements.

(D) An analysis of non-redundant external load system components that make up the primary load path (e.g., beam, cable, hook), to demonstrate compliance to the applicable structural requirements.

(E) A repetitive test of all functional devices that cycles these devices under critical structural conditions, operational conditions, or a combination of both at least 10 times each for NHEC and 30 times for HEC. This is applicable to both primary and backup subsystems. It is assumed that only one hoist cycle will typically occur per flight. This rationale has been used to determine the 10 demonstration cycles for NHEC applications and 30 demonstration cycles for HEC applications. However, if a particular application requires more than one hoist cycle per flight, then the number of demonstration cycles should be increased accordingly by multiplying the test cycles with the intended higher cycle number per flight. These repetitive tests may be conducted on the rotorcraft or by using a bench simulation that accurately replicates the rotorcraft installation.

(F) An environmental qualification for the proposed operating environment.

This review includes consideration of low and high temperatures (typically -40F to +150F), altitudes to 12,000 feet, humidity, salt spray, sand and dust, vibration, shock, rain, fungus, and acceleration. The FAA has accepted the results from the appropriate rotorcraft sections of RTCA Document DO-160 for high and low temperature and vibrations. The environmental qualification will address icing for those external load systems installed on rotorcraft approved for flight into icing conditions.

(G) Qualification of the hoist itself to the appropriate EMI and lightning threat levels specified for NHEC or HEC, as applicable. This qualification can occur separately or as part of the entire onboard QRS.

(3) Testing .

Page D - 145 AC 29-2C, Chg 7 2/4/2016 (i) Hoist system load-speed combination ground tests. The load versus speed combinations of the hoist should be demonstrated on the ground (either using an accurate engineering mock-up or a rotorcraft) by showing repeatability of the no load- speed combination, the 50 percent load-speed combination, the 75 percent load-speed combination, and the 100 percent (i.e., system rated limit) load-speed combination. If more than one operational speed range exists, the preceding tests should be performed at the most critical speed.

(A) At least 1/10 of the hoist demonstration cycles (see definition) should include the maximum aft angular displacement of the load from the vertical, applied for under § 29.865(a).

(B) A minimum of six consecutive, complete operation cycles should be conducted at the system's 100 percent (i.e., system limit rated) load-speed combination.

(C) In addition, the demonstration should cover all normal and emergency modes of intended operation and should include operation of all control devices such as limit switches, braking devices, and overload sensors in the system.

(D) All quick disconnect devices and cable cutters should be demonstrated at 0 percent, 25 percent, 50 percent, 75 percent, and 100 percent of system limit load or at the most critical percent.

Note: Some hoist designs have built-in cable tensioning devices that function at the no load-speed combination, as well as at other load-speed combinations. This device should work during the no load-speed and other load-speed cable-cutting demonstrations.

(E) Any devices or methods used to increase the mechanical advantage of the hoist should also be demonstrated.

(F) During a portion of each demonstration cycle, the hoist should be operated from each station from which it can be controlled.

(ii) Hoist and rescue hook systems or cargo hook systems flight test. An in- flight demonstration test of the hoist system should be conducted for helicopters designed to carry NHEC or HEC. The rotorcraft should be flown to the extremes of the applicable maneuver flight envelope and to all conditions that are critical to strength, maneuverability, stability, and control, or any other factor affecting airworthiness.

Unless a lesser load is determined to be more critical for either dynamic stability or other reasons, the maximum hoist system rated load or, if less, the maximum load requested for approval (and the associated limit load data placards) should be used for these tests. The minimum hoist system load (or zero load) should also be demonstrated in these tests.

Page D - 146 2/6/2016 AC 29-2C, Chg 7 (iii) Section 29.865(d) Flight test Verification Work. Flight test verification work that thoroughly examines the operational envelope should be conducted with the external cargo carriage device for which approval is requested (especially those that involve HEC).

The flight test program should show that all aspects of the operations applied for are safe, uncomplicated, and can be conducted by a qualified flight crew under the most critical service environment and, in the case of HEC, under emergency condition. Flight tests should be conducted for the simulated representative NHEC and HEC loads to demonstrate their in-flight handling and separation characteristics. Each placard, marking, and flight manual supplement should be validated during TIA flight testing.

(A) General. Flight testing or an equivalent combination of analysis, ground tests, and flight tests should be conducted under the critical combinations of configurations and operating conditions for which basic type certification approval is sought. The critical load condition of the intended cargo (e.g. rocks, lumber, radio towers, HEC, etc.) may be defined by a heavy weight and low area cargo or a low weight and high area cargo. The effects of these load conditions should be evaluated throughout the operational aspects of cargo loading, take-off, cruise up to maximum allowable speed with cargo, jettison, and landing. The helicopter handling with different cable conditions should include lateral transitions and quick stops up to the helicopter approved low airspeed limitations. Additional combinations of external load and operating conditions may be subsequently approved under relevant operational requirements as long as the structural limits and reliability considerations of the basic certification approval are not exceeded (i.e., equivalent safety is maintained). The qualification flight test of this subparagraph is intended to be accomplished primarily by analysis or bench testing. However, at least one in-flight, limit load drop test should be conducted for the critical load case. If one critical load case cannot be clearly identified, then more than one drop test might be necessary. Also, in-flight tests for the minimum load case (i.e., typically the cable hook itself) with the load trailing both in the minimum and maximum cable length configurations should be conducted. Any safety-of-flight limitations should be documented and placed in the RFM or RFMS. In certain low-gross weight, jettisonable HEC configurations, the PCDS may act as a trailing airfoil that could result in entangling the PCDS and the rotorcraft. These configurations should be assessed on a case-by-case basis by analysis or flight test to assure any safety-of-flight limitations are clearly identified and placed in the RFM or RFMS (also see PCDS).

(B) Separation characteristics of jettisonable external loads. For all jettisonable RLC of any applicable cargo type, satisfactory post-jettison separation characteristics of all loads should meet the minimum criteria that follow: ( 1 ) Separate functioning of the PQRS and BQRS resulting in a complete, immediate release of the external load without interference by the rotorcraft or external load system.

( 2 ) No damage to the helicopter during or following actuation of the QRS and load jettisoning.

Page D - 147 AC 29-2C, Chg 7 2/4/2016 ( 3 ) A jettison trajectory clear of the helicopter.

( 4 ) No inherent instability of the jettisonable (or just jettisoned) HEC or NHEC while in proximity to the helicopter.

( 5 ) No adverse or uncontrollable helicopter reactions at the time of jettison.

( 6 ) Stability and control characteristics after jettison within the originally approved limits.

( 7 ) No adverse degradation on helicopter performance characteristics after jettison.

(C) Jettison requirements for jettisonable external loads. For representative cargo types (low, medium, and high density loads on long and short lines), emergency and normal jettison procedures should be demonstrated (by a combination of analysis, ground tests, and flight tests) at sufficient combinations of flight conditions to establish a jettison envelope that should be placed in the flight manual.

(D) QRS demonstration. Repetitive jettison demonstrations, which may be accomplished during ground or flight tests, should be conducted that use the PQRS.

The BQRS should be utilized at least once.

(E) QRS reliability (i.e., failure modes) affecting flight performance. The FHA of the QRS (see paragraph d.(2) of this guidance) should show that any single system failure will not result in unsatisfactory flight characteristics, including any QRS failures resulting in asymmetric loading conditions.

(F) Flight test weight and CG locations. All flight tests should be conducted at the extreme or critical combinations of weight and longitudinal and lateral CG conditions within the applied for flight envelope. Typically, the two load conditions would be a heavy weight and low area cargo, and a low weight and high area cargo.

The rotorcraft should remain within approved weight and CG limits both with the external load applied and after jettison of the load.

(G) Jettison Envelopes. Emergency and normal jettison demonstrations should be performed at sufficient airspeeds and descent rates to establish any restrictions for satisfactory separation characteristics. Both the maximum and minimum airspeed limits and maximum descent rate for safe separation should be determined.

The sideslip envelope as a function of airspeed should be determined.

(H) Altitude. Emergency and normal jettison demonstrations should be performed at altitudes consistent with the approvable operational envelope and with the maneuvering requirements necessary to overcome any adverse effects of the jettison.

Page D - 148 2/6/2016 AC 29-2C, Chg 7 (I) Attitude. Emergency and normal jettison demonstrations should be performed from all attitudes appropriate to normal and emergency operational usage.

Where the attitudes of HEC or NHEC with respect to the helicopter may be varied, the most critical attitude should be demonstrated. This demonstration would normally be accomplished by bench testing.

(4) Rotorcraft Flight Manual (RFM) and Rotorcraft Flight Manual Supplement (RFMS) : (i) General.

(A) Present appropriate flight manual procedures and limitations for all HEC operations.

( 1 ) Approval of an external loads equipment design in accordance with § 29.865 does not provide approval to conduct external loads operations. Therefore, the following shall be included as a limitation in the RFM or RFMS:  The external load equipment certification approval does not constitute operational approval; operational approval for external load operations must be granted by the local Aviation Authority ( 2 ) The RFM or RFMS approved through the certification authority should not contain references to RLC classes, as this implies the design meets the operational rules.

(B) For non-HEC designs, the following limitation should be included within the RFM or RFMS:  The external load system does not comply with the 14 CFR part 29 certification requirements for Human External Cargo (HEC).

(C) The RFM or RFMS may contain wording to clarify whether the external load system meets the certification requirements for lifting an external load free of land or water and whether the load is jettisonable.

(D) The RFM or RFMS should contain emergency procedures detailing the steps to be taken by the flight crew during emergencies such as an engine failure, hoist failure, flight director or autopilot failure, etc.

(E) The RFM or RFMS normal procedures should explain the required procedures to conduct a safe external load operation. Such information may include the methods for attachment and normal release of the external load.

(ii) HEC installations.

Page D - 149 AC 29-2C, Chg 7 2/4/2016 (A) For HEC installations, the following additional limitations should be included in the RFM or RFMS.

( 1 ) The external load system meets the 14 CFR part 29 certification requirements for Human External Cargo (HEC).

( 2 ) Operation of the external load equipment with HEC requires the use of a Personnel Carrying Device Systems (PCDS), which must be approved by the local Aviation Authority. TSO-C167 provides one such acceptable means of approval.

(B) Crewmember communications.

(1) The flight manual should clearly define the method of communication between the flight crew and the HEC. These instructions and manuals should be validated during TIA flight testing.

(2) If the external load system does not include equipment to allow direct intercommunication among required crewmembers and external occupants, the following limitation may be included within the limitations section of the RFM or RFMS:  This external load system does not include equipment to allow direct intercommunication among required crewmembers and external occupants. Operating this external load equipment with HEC must not be authorized unless equipment to allow direct intercommunication among required crewmembers and external occupants is approved by the local Aviation Authority.

(iii) Additional RFM or RFMS requirements are contained within each applicable paragraph of this guidance.

(5) Continued airworthiness .

(i) Instructions for Continued Airworthiness. Maintenance manuals (and RFM supplements) developed by applicants for external load applications should be presented for approval and should include all appropriate inspection and maintenance procedures. The applicant should provide sufficient data and other information to establish the frequency, extent, and methods of inspection of critical structure, systems, and components. 14CFR 29.1529 and Appendix A to part 29 requires this information be included in the maintenance manual. For example, maintenance requirements for sensitive QRS squibs should be carefully determined, documented, approved during certification, and included as specific mandatory scheduled maintenance requirements that may require either "daily" or "pre-flight" checks (especially for HEC applications).

(ii) Hoist system continued airworthiness. The design life of the hoist system and any limited life components should be clearly identified, and the Airworthiness Limitations Section of the maintenance manual should include these requirements. For Page D - 150 2/6/2016 AC 29-2C, Chg 7 STCs, a maintenance manual supplement should be provided that includes these requirements.

Note: Design lives of hoist and cable systems are typically between 5,000 to 8,000 cycles. Some hoist systems have usage time meters installed. Others may have cycle counters installed. Cycle counters should be considered for HEC operations and high load or other operations that may cause low-cycle fatigue failures.

(6) Section 29.865(a) Static Structural and § 29.865(f) Fatigue Substantiation Procedures . The following static structural substantiation methods and fatigue substantiation procedures should be used: (i) Critical basic load determination. The critical basic loads and corresponding flight envelope are determined by statically substantiating the gross weight range limits, the corresponding vertical limit load factors (N ) and safety factors ZW applicable for the type of external load for which the application is being made.

Note: In cases where NHEC or HEC can have more than one shape, center-of-gravity, center-of-lift, or be carried at more than one distance in flight from the rotorcraft attachment, a critical configuration for certification purposes may not be determinable. If such a critical configuration can be determined, it may be examined for approval as a "worst case" to satisfy a particular certification criterion or several criteria, as appropriate. If such a critical configuration can not be determined, the extreme points of the operational – external load configuration envelope should be examined with consideration given to any other points within the envelope which experience or other rational would indicate should be investigated.

(ii) Vertical Limit and Ultimate Load Factors. The basic N is converted to ZW ultimate load by multiplying the maximum vertical limit load by the appropriate safety factor. (For restricted category approvals, see section AC 29 MG 5 of this AC.) This ultimate load is used to substantiate all existing structure affected by, and all added structure associated with, the load carrying device, its attachments and its cargo.

Casting factors, fitting factors, and other dynamic load factors should be applied where appropriate.

(A) NHEC applications. In most cases, it is acceptable to perform a standard static analysis to show compliance. A vertical limit load factor (N ) of 2.5 g is ZW typical for heavy gross weight NHEC hauling configurations (see section 29.337 of this AC). This vertical load factor should be applied to the maximum external load for which application is being made, together with a minimum safety factor of 1.5.

(B) HEC applications.

Page D - 151 AC 29-2C, Chg 7 2/4/2016 ( 1 ) If a safety factor 3.0 or more is used, it is acceptable to perform a standard static analysis to show compliance. The safety factor should be applied to the yield strength of the weakest component in the system (QRS, PCDS, and attachment load path). If a safety factor of less than 3.0 is used, both an analysis and a full-scale ultimate load test of relevant parts of the system should be submitted.

( 2 ) Since HEC applications typically involve lower gross weight configurations, a higher vertical limit load factor is required to assure that limit load is not exceeded in service. The applicant should use either the conservative value of 3.5 g or an analytically derived maximum vertical limit load factor for the requested operating envelope. Linear interpolation between the vertical load factors of maximum and minimum design weights may be used. However, in no case may the vertical limit load factor be less than 2.5 g for any RLC application for HEC.

( 3 ) For the purpose of structural analysis or test, assume a 223-pound (101 Kg) man as the minimum weight of each occupant carried as HEC.

Note: If the HEC is engaged in work tasks that employ devices of significant added weight (heavy backpacks, tools, fire extinguishers, etc.), the total weight of the 223-pound man and equipment should be assumed in the structural analysis or test.

(iii) Critical Structural Case. For applications involving more than one RLC class or cargo type, the structural substantiation is required only for the most critical case. The most critical case should be determined by rational analysis.

(iv) Jettisonable Loads. For the substantiating analyses or tests of all jettisonable RLC external loads, including HEC, the maximum external load should be applied at the maximum angle that can be achieved in service but not less than 30 degrees. The angle should be measured from the sling-load-line to rotorcraft vertical axis (z axis) and may be in any direction that can be achieved in service. The 30-degree angle may be reduced in some or all directions if it is impossible to obtain due to physical constraints or operating limitations. The maximum allowable cable angle should be determined and approved. The angle approval should be based on structural requirements, mechanical interference limits, and flight handling characteristics over the most critical conditions and combinations of conditions in the approved flight envelope.

(v) Hoist system limit load.

Note: In cases where hoist cables or long line cables are utilized, a new dynamic system is established. Characteristics of the system should be evaluated to assure that either no hazardous failure modes exist or that they are acceptably minimized. For example, the cable or long line may exhibit a natural frequency that could be excited by sources internal to the overall structural system (i.e., the rotorcraft) or by sources external to the system. Another example is the loading Page D - 152 2/6/2016 AC 29-2C, Chg 7 effect of the cable acting as a spring between the rotorcraft and the suspended external load.

(A) Determine the basic loads that result in the failure or unspooling of the hoist or its installation, respectively.

Note: This determination should be based on static strength and any significant dynamic load magnification factors.

(B) Select the lower of the two values as the ultimate load of the hoist system installation.

(C) Divide the selected ultimate load by 1.5 to determine the true structural limit load of the system.

(D) Determine the manufacturer's approved (or applicants applied for) "limit design safety factor." Divide this factor into the true structural limit load (from (C) above) to determine the hoist system's working (or placarded) limit load.

(E) Compare the system's derived limit load to the applied one "g" payload multiplied by the maximum downward vertical load factor (N ) to determine the ZWMAX critical payload's limit value.

(F) The critical limit payload should be equal to or less than the system's derived limit load for the installation to be approvable.

(vi) Fatigue Substantiation Procedures (A) Fatigue evaluation of NHEC applications. Any critical components of the suspended system and their attachments (e.g., the cargo hook, or bolted or pinned truss attachments), the failure of which could result in a hazard to the rotorcraft, should include an acceptable fatigue analysis.

(B) Fatigue evaluation of HEC applications. The entire external load system, including the PCDS, should be reviewed on a component-by-component basis to determine which, if any, components are fatigue critical. These components should be analyzed or tested to assure their fatigue life limits are properly determined and placed in the limited life section of the maintenance manual.

(7) Section 29.865(b) and § 29.865(c) Procedures for Quick Release Systems .

For jettisonable RLCs of any applicable cargo type, both a primary quick-release system (PQRS) and a back-up quick-release system (BQRS) are required. Features that should be considered are: Page D - 153 AC 29-2C, Chg 7 2/4/2016 (i) The PQRS, BQRS and their load release devices and subsystems (such as electronically actuated guillotines) should be separate (i.e., physically, systematically, and functionally redundant).

(ii) The controls for the PQRS should be installed on one of the pilot's primary controls, or in an equivalently accessible location. The use of an "equivalent accessible location" should be reviewed on a case-by-case basis and used only where equivalent safety is clearly maintained.

(iii) The controls for the BQRS may be less sophisticated than that of the PQRS. For instance, manual cable cutters are acceptable provided they are listed in the flight manual as a required device and have a dedicated, placarded storage location.

(iv) The PQRS should release the external load in less than 5 seconds. The BQRS should release the external load in less than 30 seconds. This time interval begins the moment an emergency is declared and ends when the load is released.

(v) Each quick-release device should be designed and located to allow the pilot or a crewmember to accomplish external cargo release without hazardously limiting the ability to control the rotorcraft during emergency situations. The flight manual should reflect the requirement for a crewmember and the related functions.

(vi) Section 29.865(c)(1) QRS Requirements for Jettisonable HEC Operations.

(A) For jettisonable HEC operations, both the PQRS and BQRS are required to have a DAD for external cargo release. The DAD must be designed to require two actions with a definite change of direction of movement, such as opening a switch or pushbutton cover followed by a definite change of direction in order to activate the release switch or pushbutton. Any possibility of opening the switch cover and inadvertently releasing the load with a single motion is not acceptable. An additional level of safety may also be provided through the use of Advisory and Caution messages. For example, an advisory “ON” message might be illuminated when the pilot energizes (but not arms) the system with a master switch. A cautionary “ARMED” message would then illuminate when the pilot opens the switch guard. In this case, a possible unwanted flip of the switch guard would be immediately recognized by the crew. The switch design should be evaluated by ground or flight test. The RFM or RFMS should contain a clear description of DAD functionality, to include the associated safety features, normal and emergency procedures, and applicable advisory and caution messages.

(B) The DAD is intended for emergency use during the phases of flight that the HEC is carried or retrieved. The DAD can be used for both NHEC and HEC operations. However, because it can be used for HEC, the instructions for continued airworthiness should be carefully reviewed and documented. The DAD can be operated Page D - 154 2/6/2016 AC 29-2C, Chg 7 by the pilot from a primary control or, after a command is given by the pilot, by a crewmember from a remote location. Additional safety precautions (such as a lockwire) should be considered for remote hoist console in the cabin. Any emergency release function provided by a remote hoist console must also be designed to protect against inadvertent activation during the hoist operation. If the backup DAD is a cable cutter, it should be properly secured, placarded and readily accessible to the crewmember intended to use it.

(vii) Section 29.865(b)(3)(ii) Electromagnetic Interference. Protection of the QRS against potential internal and external sources of electromagnetic interference (EMI) and lightning is required. This is necessary to prevent inadvertent load release from sources such as lightning strikes, stray electromagnetic signals, and static electricity.

(A) Jettisonable NHEC systems should be able to absorb a minimum of 20 volts per meter (i.e., CAT U) radio-frequency (RF) field strength per RTCA Document DO-160.

(B) Jettisonable HEC systems should be able to absorb a minimum of 200 volts per meter (i.e., CAT Y) RF field strength per RTCA Document DO-160.

Notes: ( 1 )These RF field threat levels may need to be increased for certain special applications such as microwave tower and high voltage high line repairs.

Separate criteria for special applications under multi-agency regulation (such as IEEE or OSHA standards) should also be addressed, as applicable, during certification. When necessary, the issue paper process can be used to establish a practicable level of safety for specific high voltage or other special application conditions. The helicopter HIRF safety assessment should consider the effects on helicopter flight safety due to a HIRF-induced failure or malfunction of external load systems, such as uncommanded hoist winch activation without ability to jettison, or uncommanded load jettison. The appropriate failure effect classification should be assigned based on this assessment, and compliance demonstrated according to § 29.1317 and AC 20-158. This requirement should not be limited to the cable cutter devices or load jettison subsystems only. In some designs, uncommanded load release or hoist winch activation could also result from a failure of the command and control circuits of the system.

( 2 ) An approved standard rotorcraft test, which includes the full HIRF frequency and amplitude external and internal environments, on the QRS and any applicable PCDS, or the entire rotorcraft including the QRS and any applicable PCDS, could be substituted for the jettisonable NHEC and HEC systems tests as long as the RF field strengths directly on the QRS and PCDS are shown to equal or exceed those defined by paragraphs d.(7)(vii)(A) and d.(7)(vii)(B) for NHEC and HEC respectively.

( 3 )The EMI levels specified in paragraphs d.(7)(vii)(A) and d.(7)(vii)(B) of this guidance are total EMI levels to be applied to the QRS (and affected QRS component) boundary. The total EMI level applied should include the effects of both Page D - 155 AC 29-2C, Chg 7 2/4/2016 external EMI sources and internal EMI sources. All aspects of internally generated EMI should be carefully considered including peaks that could occur from time-to-time due to any combination of on-board systems being operated. For example, special attention should be given to EMI from hoist operations that involve the switching of very high currents. Those currents can generate significant voltages in closely spaced wiring that, if allowed to reach some squib designs, could activate the device. Shielding, bonding, and grounding of wiring associated with operation of the hoist and the quick-release mechanism should be clearly and adequately evaluated in design and certification.

When recognized good practices for such installation are applied, an analysis may be sufficient to highlight that the maximum possible pulse generated into the squib circuit will have an energy contents' orders of magnitude below the squib no fire energy. If insufficient data is available for installation and/or squib no fire energy, this evaluation may require testing. One acceptable test method to demonstrate adequacy of QRS shielding, bonding, and grounding would be to actuate the hoist under maximum load together with likely critical combinations of other aircraft electrical loads and demonstrate that the test squibs (that are more EMI sensitive than the squibs specified for use in the QRS) do not inadvertently operate during the test.

(viii) Section 27.865(c)(3) Placarding and marking. For jettisonable HEC applications, the associated QRS placarding should be given special consideration to assure the proper level of occupant safety.

(8) Cargo hooks or equivalent devices and their related systems . All cargo hooks or equivalent devices should be approved to acceptable aircraft industry standards.

The applicant should present these standards, and any related manufacturer’s certificates of production or qualification as part of the approval package.

(i) General. Cargo hook systems should have the same reliability goals and should be functionally demonstrated under critical loads for NHEC and HEC, as appropriate. All engagement and release modes should be demonstrated. If the hook is used as a quick-release device, then release of critical loads should be demonstrated under conditions that simulate maximum allowable bank angles and speeds and any other critical operating conditions. Demonstration of any re-latching features and any safety or warning devices should also be conducted. Demonstration of actual in-flight emergency quick-release capability may not be necessary if the quick-release capability can be acceptably simulated by other means.

Notes: Cargo hook manufacturers specify particular shapes, sizes, and cross sections for lifting eyes to assure compatibility with their hook design (e.g., Breeze Eastern Service Bulletin CAB-100-41). Experience has shown that, under certain conditions, a load may inadvertently hang up because of improper geometry at the hook-to-eye interface that will not allow the eye to slide off an open hook as intended.

For both NHEC and HEC designs, the phenomena of hook dynamic roll out (inadvertent opening of the hook latch and subsequent release of the load) Page D - 156 2/6/2016 AC 29-2C, Chg 7 should be considered to assure that QRS reliability goals are not compromised. This is of particular concern for HEC applications. Hook dynamic roll-out occurs during certain ground handling and flight conditions that may allow the lifting eye to work its way out of the hook.

Hook dynamic roll-out typically occurs when either the RLCs sling or harness is not properly attached to the hook, is blown by down draft, is dragged along the ground or through water; or is otherwise placed into the dangerous hook- to-eye configuration.

The potential for hook dynamic roll-out can be minimized in design by specifying particular hook-and-eye shape and cross-section combinations.

For non-jettisonable RLCs, a pin can be used to lock the hook keeper in place during operations.

Some cargo hook systems may employ two or more cargo hooks for safety.

These systems are approvable. However, loss of any load by a single hook should be shown to not result in loss of control of the rotorcraft. In a dual hook system, if the hook itself is the quick-release device (i.e., if a single release point does not exist in the load path between the rotorcraft and the dual hooks), the pilot should have a dual PQRS that includes selectable, co- located individual quick releases that are independent for each hook used. A BQRS should also be present for each hook. For cargo hook systems with more than two hooks, either a single release point should be present in the load path between the rotorcraft and the multiple hook system, or multiple PQRS and BQRSs should be present.

(ii) Jettisonable cargo hook systems. For jettisonable applications, each cargo hook - (A) Should have a sufficient amount of slack in the control cable to permit cargo hook movement without tripping the hook release.

(B) Should be shown to be reliable.

(C) For HEC systems, unless the cargo hook is to be the primary quick-release device, each cargo hook should be designed so that operationally induced loads cannot inadvertently release the load. For example, a simple cargo hook should have a one-way, spring-loaded gate (i.e., "snap hook") that allows load attachment going into the gate but does not allow the gate to open (and subsequently lose the HEC) when an operationally induced load is applied in the opposite direction.

For HEC applications, cargo hooks that also serve as a quick-release device should be carefully reviewed to assure they are reliable.

(iii) Other load release types . In some current configurations, such as those used for high line (or long line) operations, a load release may be present that is not on Page D - 157 AC 29-2C, Chg 7 2/4/2016 the rotorcraft but is a remote release system. The long lines allow the pilot to not release the line itself during repetitive loading operations. Examples are a tension release device that lets out line under an operationally induced load, a personal rope cutter or dedicated switches at the pilot controls. The release of the load through the secondary hook on a long line presents additional risks due to the possibility of the long line to impact the tail or main rotor after a release due to its elasticity. These devices are acceptable if the: (A) off-rotorcraft release is considered a "third release". This type of release is not a substitute for a required release (i.e., PQRS or BQRS); (B) cargo hook release and the long line release are placed on the primary controls in a way to avoid confusion during operation. One example of compliance would be to place the cargo hook release on the cyclic and the long line release on the collective to avoid any possible confusion in the operation; (C) RFM or RFMS includes a description of the new control in the cockpit and its function and an RFM or RFMS note to the pilot is included indicating that the helicopter hook emergency release procedures are fully applicable; (D) release meets the same certification reliability criteria adopted for the helicopter certified hook, as well as other relevant requirements of § 29.865 and the methods of this guidance or equivalent methods; and (E) release has no operational or failure modes that would affect continued safe flight and landing under any operations, critical failure modes, conditions, or combination of either. The following points should be considered: ( 1 ) The long line should not be of an elastic material allowing spring up when unloaded or elevated dynamics when loaded.

( 2 ) The long line should have a residual weight allowing its release from the helicopter hook when the long line is unloaded.

( 3 ) The RFM or RFMS should include all operating procedures to ensure that the long line does not impact the rotors after cargo release and during unloaded flight phases.

( 4 ) The hook should be designed to minimize inadvertent activation.

An example may be a protective device (cage) around the locking mechanism of the long line hook.

( 5 ) A means should be provided to prevent fouling of cables in the event of rotation of the external load. An example may be the inclusion of a swivel or slipring.

Page D - 158 2/6/2016 AC 29-2C, Chg 7 ( 6 ) Installation of long line provided with electrical wiring to control the hook will generally represent a new electromagnetic coupling path from external area to the internal systems that may not have been considered for type certification. As such, the impact of this installation on the coupling to helicopter systems, due to direct connection or cross talk to wirings, shall be addressed in accordance with § 29.1317, 29.1316 and 29.610 as a minimum.

(9) Cable (i) Cable attachment . Either the cable should be positively attached to the hoist drum and the attachment should have ultimate load capability or equivalent means should be provided to minimize the possibility of inadvertent, complete cable unspooling.

(ii) Cable length and marking . A length of cable nearest the cable's attachment to the hoist drum should be visually marked to indicate to the operator that the cable is near full extension. The length of cable to be marked is a function of the maximum extension speed of the system and the operator's reaction time needed to prevent cable run out. It should be determined during certification demonstration tests.

In no case should the length be less than 3-1/2 drum circumferences.

(iii) Cable stops . Means should be present to automatically stop cable movement quickly when the system's extension and retraction operational limits are reached.

(10) Section 29.865(c)(2) PCDS . For all HEC applications, an approved PCDS is required. The PCDS may be either previously approved or is required to be approved during certification. A PCDS may also be approved after the external load system certification. Including a limitation in the RFM or RFMS requiring an approved PCDS will permit more flexibility for operators to utilize their own approved PCDS (see paragraph d.(4) of this guidance). In either case, its compatibility with the helicopter should be approved.

Note: PCDS designs can vary from simple single occupant lifesaving devices to relatively complex multiple occupant cages or gondolas. The purpose of the PCDS is to provide a minimum acceptable level of safety for personnel being transported outside the rotorcraft. The personnel being transported may be healthy or injured, conscious or unconscious.

(i) TSO C167 is an approved minimum performance specification for HEC body harnesses.

(ii) Static strength . The PCDS should be substantiated for the allowable ultimate load and loading conditions as determined under paragraph d.(6) above.

Page D - 159 AC 29-2C, Chg 7 2/4/2016 (iii) Fatigue . The PCDS should be substantiated for fatigue as determined under paragraph d.(6) above.

(iv) Personnel safety . For each PCDS design, the applicant should submit a design evaluation that assures the necessary level of personnel safety is provided. As a minimum, the following should be evaluated.

(A) The PCDS should be easily and readily ingressed or egressed.

(B) It should be placarded for proper capacity, internal arrangement and location of occupants, and ingress and egress instructions.

(C) For door latch fail-safety, more than one fastener or closure device should be used. The latch device design should provide direct visual inspectability to assure it is fastened and secured.

(D) Any fabric used should be durable and should be at least flame- resistant.

(E) Safety harnesses and belts should meet TSO C167 requirements.

(F) Occupant retention devices and related design safety features should be used as necessary. In simple designs, rounded corners and edges with adequate strapping (or other means of HEC retention relative to the PCDS) and head supports or pads may be all the safety features that are necessary. However, in more complex PCDS designs, safety features such as seat belts, handholds, shoulder harnesses, placards, or other personnel safety standards may be required.

(v) EMI and lightning protection . All essential, affected components of the PCDS, such as intercommunication equipment, should be protected against RF field strengths to a minimum of RTCA Document DO-160 Category Y.

(vi) Instructions for continued airworthiness . All instructions and documents necessary for continued airworthiness, normal operations, and emergency operations should be completed, reviewed, and approved during the certification process. There should be clear instructions to describe when the PCDS is no longer serviceable and should be replaced in part or as a whole due to wear, impact damage, fraying of fibers, other forms of degradation. In addition, any life limitations resulting from compliance with paragraphs d.(10)(ii) and (iii) should be provided.

(vii) Flotation devices . PCDSs that are intended to have a dual role as floatation devices or life preservers should meet the requirements of TSO-C13f, "Life Preservers." Also, any PCDS design to be used in the water should have a floatation kit. The kit should support the weight of the maximum number of occupants and the PCDS in the water and minimize the possibility of the occupants floating face down.

Page D - 160 2/6/2016 AC 29-2C, Chg 7 (viii) Flight testing . It should be shown by flight tests that the device is safely controllable and maneuverable during all requested flight regimes without requiring exceptional piloting skill. The flight tests should entail the PCDS weighted to the most critical weight. Litters and other types of PCDS designs may spin, twist or otherwise respond unacceptably in flight. These designs should be structurally restrained with devices such as a spider, a harness, or an equivalent device to minimize undesirable flight dynamics.

(ix) Medical design considerations . The PCDS should be designed to the maximum practicable extent and placarded to maximize the HECs protection from medical considerations such as blocked air passages induced by improper body configuration and excessive loss of body heat during operations. Injured or water soaked persons may be exposed to high body heat loss from sources such as rotor wash and the airstream. PCDS occupant safety from transit induced medical considerations can be greatly increased by proper design.

(x) Hoist operator safety device . When hoisting operations require the presence of a hoist operator on board, appropriate provisions shall be provided to allow the hoist operator to perform his task safely. These provisions shall include an appropriate hoist operator restraint system. This safety device is typically composed of a safety harness and a strap attached to the cabin used to adequately restrain the hoist operator inside the cabin while operating the hoist. For certification approval, the hoist operator safety device should comply with § 29.561(b)(3) for personnel safety. The applicant should submit a design evaluation that assures the necessary level of personnel safety is provided. The safety harness and strap should meet TSO C22g, TSO C114, or TSO C167 requirements; however, as a minimum, the following should be evaluated: (A) The strap attaching point on the body harness should not be located in the back of the hoist operator due to safety issues in case of a fall or crash.

(B) The safety device should be designed to be adjustable so that the strap is tightened behind the hoist operator.

(C) The strap might include a QRS including a DAD to allow hoist operator detaching from cabin in emergency conditions (e.g., crash).

(D) The safety device should be easily and readily ingressed or egressed.

(E) It should be placarded for proper capacity and lifetime limitation.

(F) Any fabric used should be durable and should be at least flame resistant.

(11) Section 29.865(c)(4) Intercom Systems for HEC Operations . For all HEC operations, the rotorcraft is required to be equipped for, or otherwise allow direct Page D - 161 AC 29-2C, Chg 7 2/4/2016 intercommunication under any operational conditions among crewmembers and the HEC. An intercommunications system may also be approved after the external load system certification as operators are standardizing communications equipment within their fleet for safety benefits. In lieu of an intercommunication system being provided as part of the external load system, this onus may be moved onto the operator through a limitation may be placed in the RFM or RFMS (see paragraph d.(4) of this guidance).

(12) Section 29.865(c)(6) HEC Operations Requiring Category A Rotorcraft . For jettisonable HEC operations, it may be required by Operations Requirements, § 133.45(e), that the rotorcraft meet the category A engine isolation requirements of part 29 and that the rotorcraft have OEI hover performance capability OGE in its approved, jettisonable HEC weight, altitude, and temperature envelope.

(i) In determining OEI hover performance, dynamic engine failures should be considered. Each hover verification test should begin from a stabilized hover at the maximum OEI hover weight, at the requested IGE or OGE skid or wheel height, and with all engines operating. At this point the critical engine should be failed and the aircraft should remain in a stabilized hover condition without exceeding any rotor limits or engine limits for the operating engine(s). As with all performance testing, engine power should be limited to minimum specification power. Engine failures may be simulated by rapidly moving the throttle to idle provided a ‘needle split’ is obtained between the rotor and engine RPM.

(ii) Pilot reaction times used following the engine failure to maintain the stabilized hover flight condition will depend on the environment in which such operations are planned. For HEC operations planned for day visual meteorological conditions (VMC), normal pilot reaction times are acceptable. Recent certification projects for HEC operations planned to take place overwater at night while in VMC or in instrument meteorological conditions (IMC) required greater pilot reaction times. When hovering OGE or IGE at maximum OEI hover weight, an engine failure should not result in an altitude loss of more than 10 percent or four (4) feet, whichever is greater, of the altitude established at the time of engine failure. In either case, sufficient power margin should be available from the operating engine(s) to regain the altitude lost during the dynamic engine failure and then to transition to forward flight.

(iii) Consideration should also be given to the time required to recover (reel in and bring aboard) the HEC and then to transition to forward flight. This time increment may limit the use of short duration OEI power ratings. For example, for a helicopter that sustains an engine failure at a height of 40 feet, the time required to stabilize the hover at the original hover height, recover the HEC (given the hoist speed limitations), and then transition to forward flight (with minimal altitude loss) would likely preclude the use of 30-second engine ratings and may encroach upon the 2 ½ -minute ratings. Such encroachment into the 2 ½ - ratings is not acceptable. In accordance with the intent of OEI power ratings, recovery and transition procedures shall not utilize more than one iteration of either a 30-second / 2-minute power combination or a 2½-minute rating.

Page D - 162 2/6/2016 AC 29-2C, Chg 7 (iv) For helicopters that incorporate engine driven generators, the hoist should remain operational following an engine or generator failure. A hoist should not be powered from a bus that is automatically shed following the loss of an engine or generator. Maximum two-engine generator loads should be established so that when one engine or generator fails, the remaining generator can assume the entire rotorcraft electrical load (including the maximum hoist electrical load) without exceeding approved limitations.

(v) The RFM or RFMS should contain information that describes the expected altitude loss, any special recovery techniques, and the time increment used for recovery of the external load when establishing maximum weights and wheel or skid heights.

Engine failure procedures that permit safe recovery of the HEC and safe fly-away capability shall be included in the RFM or RFMS The OEI hover chart should be placed in the performance section of the RFM or RFMS. Allowable altitude extrapolation for the hover data should not exceed 2000 feet. The following statement (modified appropriately) should be included to indicate to operators which charts are applicable to these HEC operations.

• For external load equipment installations specifically intended for HEC applications requiring the use of Transport Category A rotorcraft, OEI hover performance data, and procedures may be found in Figures/Charts/Section XX - YY… (13) Section 29.865(e) External Loads Placards and Markings . Placards and markings should be installed next to the external load attaching means, in a clearly noticeable location, that state the primary operational limitations - specifically including the maximum authorized external load. Not all operational limitations need be stated on the placard (or equivalent markings) only those clearly necessary for immediate reference in operations. Other more detailed operational limitations of lesser immediate importance should be stated either directly in the RFM or in a RFMS.

(14) Other Considerations .

(i) Agricultural Installation (AI). AIs can be approved for either jettisonable or non-jettisonable NHEC or HEC operations as long as they meet relevant certification and operations requirements and follow appropriate compliance methods. However, most current AI designs are external fixtures (see definition) - not external loads.

External fixtures are not approvable as jettisonable external cargo because they do not have a true payload (see definition), true jettison capability (see definition), or a complete QRS. Many AI designs can dump their solid or liquid chemical loads by use of a "purge port" release over a relatively long time period (i.e., greater than 30 seconds).

This is not considered true jettison capability (see definition) since the external load is not released by a QRS and since the release time span is typically greater than 30 seconds (see paragraph c.(21)). Thus, these types of AIs should be approved as a non-jettisonable external load. However, other designs that have the entire AI (or significant portions thereof) attached to the rotorcraft, that have short time frame jettison Page D - 163 AC 29-2C, Chg 7 2/4/2016 (or release) capability provided by a QRS that meets the definitions herein and that have no post-jettison characteristics that would endanger continued safe flight and landing may be approved as a jettisonable external load. For example, if all the relevant criteria are properly met, a jettisonable fluid load can be approved as a NHEC external cargo. Section AC 29 MG 5 of this AC discusses other AI certification methodology.

(ii) External Tanks. External tank configurations that have true payload (see definition) and true jettison capability (see definition) should be approved as jettisonable NHEC. External tank configurations that have a true payload capability but do not have true jettison capability should be approved as non-jettisonable NHEC. An external tank that has neither a true payload capability nor true jettison capability is an external fixture; it should not be approved as an external load under § 29.865. If an external tank is to be jettisoned in flight, it should have a QRS that is approved for the maximum jettisonable external tank payload and is either inoperable or is otherwise rendered reliable to minimize inadvertent jettisons above the maximum jettisonable external tank payload.

(iii) Logging Operations. These operations are very susceptible to low-cycle fatigue because of the large loads and relatively high load cycles that are common to this industry. It is recommended that load-measuring devices (such as load cells) be used to assure that no unrecorded overloads occur and to assure that cycles producing high fatigue damage are properly considered. Cycle counters are recommended to assure acceptable cumulative fatigue damage levels are identifiable and are not exceeded. As either a supplementary method or alternate method, maintenance instructions should be considered to assure proper cycle counting and load recording during operations.

(15) Noise Certification . 14 CFR part 36 is the noise certification standard.

Section 36.1(a)(4) specifically exempts helicopters that are designed exclusively for agricultural work, carrying firefighting materials, or external loads activity from the noise standards. Section 21.93(b)(4) also contains specific information regarding external loads and what configurations do not constitute an acoustical change.

Page D - 164 5/1/2014 AC 29-2C, Chg 4 SUBPART D - DESIGN AND CONSTRUCTION MISCELLANEOUS AC 29.871. § 29.871 LEVELING MARKS.

a. Explanation . Reference marks are required for leveling the rotorcraft on the ground. These marks are necessary for accurate determination of weight and balance effects, particularly after modifications to the basic rotorcraft.

b. Procedures .

(1) Reference marks are sometimes provided in pairs, one high in the cabin and one low. The plumb weight is suspended from the high mark by an appropriate mechanical attachment, and the lower mark is used to level the rotorcraft by centering the plumb weight. The lower reference mark should be a raised or depressed target symbol and shall be applied to a permanent structural component or permanently attached plate in a readily accessible location. Seat tracks, floors, or door sills which are attached with permanent fasteners are typical locations.

(2) Horizontal reference marks for support of bubble levels may also be used, particularly for smaller rotorcraft.

(3) Proper reference should be made to identify the leveling marks or points on the rotorcraft. Design provisions should be made to ensure these locations are not obscured by equipment, fairings, repair, or rework.

AC 29.873. § 29.873 BALLAST PROVISIONS.

a. Explanation .

(1) This rule requires that ballast provisions prevent inadvertent ballast shifting while in flight or as a result of a landing. Shifting of the ballast may cause a hazardous change in the center of gravity thereby affecting rotorcraft controllability.

(2) Other rules noted here allow removable and fixed ballast and require markings or placards to prevent overloading the ballast installation.

(i) Section 29.29 specifies that the rotorcraft empty weight will include any fixed ballast. Section 29.31 allows the use of removable ballast to comply with the flight requirements. However, ballast may not be adjusted (moved, reduced or increased) in flight.

(ii) Section 29.1541 requires conspicuous and durable markings or placards. Section 29.1557 requires placards stating allowable maximum weight, distributed loading, if necessary, and other appropriate limitations for ballast installation.

Page D - 165 AC 29-2C 9/30/99 (3) Section 29.1583(c) concerns Rotorcraft Flight Manual instructions and information about removable ballast or loading information. The instructions must be included in the operating limitations section of the flight manual to allow ready observance of the limitations.

b. Procedures .

(1) The ballast installation may be substantiated by analysis or by static test.

The design ultimate load may be derived from flight, landing, or minor crash conditions load factor specified in the rules. Substantiation by analysis will require use of the fitting factor prescribed by § 29.625 where appropriate. If static tests are to be conducted, a test plan should be prepared, submitted for evaluation and agreed upon prior to the test.

(2) Ballast installations in the aft part of the fuselage and tail boom may be subject to significant landing condition angular inertia load factors as well as the usual linear load factors.

(3) Substantiation methods and procedures acceptable for the airframe substantiation may be used for the ballast installation as well.

(4) Removable ballast will require attention to assure the ballast is secured easily and properly and will remain secured under the appropriate ballast design load factor requirements. The flight manual instructions should be evaluated for compliance with § 29.1583(c) by flight test and airframe personnel.

(5) The installation must be designed and placarded or marked for the maximum allowable ballast load and for other appropriate loading limits. Normally compliance with § 29.1541 is accomplished with a drawing review by airframe personnel along with an EMDO compliance and conformity inspection. An additional compliance inspection by airframe personnel can be conducted if desired.

AC 29.877. § 29.877 ICE PROTECTION.

NOTE: § 29.877 was removed and replaced by § 29.1419 in Amendment 29-21. This material is retained since this is one way to show compliance with § 29.877 .

a. Background .

(1) In March 1984, the FAA/AUTHORITY for the first time certificated a rotorcraft for flight into known icing conditions. Several other manufacturers are pursuing designs for icing flight capability with certification planned for 1985 or 1986.

2) Most rotorcraft icing technology has been developed for military rotorcraft.

The only U.S. military rotorcraft equipped and approved for flight into icing conditions is Page D - 166 9/30/99 AC 29-2C liquid water content (LWC) does not exceed 1.0gm/m and outside air temperature (OAT) is not below -20° C.

(3) Many rotorcraft operators have voiced a high priority on obtaining rotorcraft approved for operation in icing conditions.

(4) The icing characteristics envelope of FAR Part 25, Appendix C, has served as a satisfactory design criteria for fixed-wing operations for two decades. The envelope, as presented, extends to 22,000 feet with possible extents to 30,000 feet but does not present icing severity as a function of altitude. At the time the envelope was derived, it was assumed that all transport category airplanes would operate to at least 22,000 feet. For present state-of-the-art rotorcraft, this assumption is not valid. As such, an altitude limited icing envelope based on the same data used to derive the Part 25, Appendix C, and the Part 29, Appendix C, envelopes is presented as an alternate to the full icing envelope. In addition, a second icing envelope which effectively characterizes supercooled clouds from ground level to 10,000 feet is presented as a second alternative to the Part 29, Appendix C, envelope. The second altitude limited envelope described in reference 386d(2) was derived from recent additional airborne measurements.

b. Explanation .

(1) General .

(i) The discussion in this paragraph pertains generally to certifications to the full icing envelope of Part 29, Appendix C, within the altitude limitations of the rotorcraft or to one of the altitude limited icing envelopes based on a 10,000-foot pressure altitude limit. The actual icing envelope considered may be further restricted based on the actual pressure attitude envelope for which certification is requested. It envisions certification with full ice protection systems (rotor blades, windshields, engine inlets, stabilizer surfaces, etc.). With the exception of pilot controllable variables such as altitude and airspeed, limited certification (either in terms of icing envelope or protection capability) is not envisaged at this time due to the difficulty in forecasting the severity of icing conditions, relating the effects of the forecasted conditions to the type of aircraft, and relating the effects of reported icing among various types of aircraft, particularly between fixed and rotary-wing aircraft. In addition, with a limited protection capability, viable escape options may not be operationally available if limitations are exceeded.

(ii) The discussion in this paragraph, regarding rotor blade ice protection, is oriented primarily toward electrothermal rotor deicing systems, since these have the most widespread acceptance and projected use within the industry. Also, most of the testing and research into rotorcraft ice protection to date has been conducted with this type of system. Research is continuing with other types of systems such as anti-icing fluid systems, and information will be added to address certification of these as necessary. It should also be noted that most of the rotorcraft icing experience Page D - 167 AC 29-2C 9/30/99 accumulated to date has been on rotorcraft with symmetrical airfoil sections. The application of this experience to rotorcraft with asymmetrical airfoils should be carefully evaluated. Limited experience has been gained during development and qualification testing of the Army Blackhawk on asymmetrical airfoil icing characteristics. The most prominent difference appears to be a more rapid degradation of airfoil performance.

Rapidity of performance degradation is also dependent upon severity of the icing condition (primarily a function of liquid water content) and ice shape (primarily a function of OAT and median volumetric droplet diameter (MVD)).

(iii) The effects of ice can vary considerably from rotorcraft to rotorcraft.

Experience gained for a rotor system with an identical blade profile could provide valuable information but should be used cautiously when applied to another rotorcraft.

Assumptions cannot necessarily be made based on icing test results from another rotorcraft. Particular care should be exercised when drawing from fixed-wing icing experience as the widely different and varying conditions seen by the rotor blades make many comparisons with fixed-wing results invalid. Likewise, icing effects on rotor blades vary significantly from those on other parts of the rotorcraft. This is due to changing blade velocity as compared with the constant velocity of the remaining parts.

(2) Reference Material . Prior to commencement of efforts to design and certify a rotorcraft, the references listed in paragraph AC 29.877d should be reviewed. FAA Technical Report ADS-4, Engineering Summary of Airframe Icing Technical Data, December 1963, although somewhat dated, is recommended for basic aircraft icing protection system design information.

(3) Objective . The objective of icing certification is to verify that throughout the approved envelope, the rotorcraft can operate safely in icing conditions expected to be encountered in service (i.e., Appendix C of Part 29 or one of the altitude limited icing envelopes presented herein). This will entail determining that no icing limitations exist or defining what the limitations are, as well as establishing the adequacy of the ice warning means (or system) and the ice protection system. A limiting condition may manifest itself in one of several areas such as handling qualities, performance, autorotation, asymmetric shedding from the rotors, visibility through the windshield, etc.

Prior to flight tests in icing conditions, sufficient analyses should have been conducted to determine the design points for the particular item of the rotorcraft being analyzed (windshield, engine inlet, rotor blades, etc.). After the analyses are reviewed and found adequate, tests should be conducted to confirm that the analyses are valid and that the rotorcraft can operate safely in any supercooled cloud icing condition defined by Part 29, Appendix C, or one of the altitude limited icing envelopes. References 386d(1) and (3) may be useful in determining the design points and extrapolation of test data to the desired design points.

(4) Planning . For best utilization of both the applicant’s and the FAA/AUTHORITY’s resources, the applicant should submit a certification plan at the start of the design and development effort. The certification plan should describe all Page D - 168 9/30/99 AC 29-2C efforts intended to lead to certification and should include the following basic information: Rotorcraft and systems description.

Ice protection systems description.

Certification checklist.

Description of analyses or tests planned to demonstrate compliance.

Projected schedules of design, analyses, testing, and reporting efforts.

Methods of test - artificial vs. natural.

Methods of control of variables.

Data acquisition instrumentation.

Data reduction procedures.

(5) Environment .

(i) Definitions .

(A) Supercooled Clouds . Clouds containing water droplets (below 32° F) that have remained in the liquid state. Supercooled water droplets will freeze upon impact with another object. Water droplets have been observed in the liquid state at ambient temperatures as low as -60° F. The rate of ice accretion on an aircraft component is dependent upon many factors such as droplet size, cloud liquid water content, ambient temperature, and component size, shape, and velocity.

(B) Ice Crystal Clouds . Glaciated clouds existing usually at very cold temperatures where moisture has frozen to the solid or crystal state.

(C) Mixed Conditions . Partially glaciated clouds at ambient temperatures below 32° F containing a mixture of ice crystals and supercooled water droplets.

(D) Freezing Rain and Freezing Drizzle . Precipitation existing within clouds or below clouds at ambient temperatures below 32° F where rain droplets remain in the supercooled liquid state.

(E) Sleet . Precipitation of transparent or translucent pellets of ice which have a diameter of 5mm or less.

(F) Hail . Solid precipitation in the form of balls or pieces of ice (hail stones) with diameters ranging from 5mm to more than 50mm.

(ii) Appendix C of Part 29 defines the supercooled cloud environment necessary for certification of rotorcraft in icing except that the pressure altitude limitation is that of the rotorcraft or that selected by the applicant, provided the remaining altitude envelope is operationally practical. Due to air traffic system compatibility constraints, approval of a maximum altitude less than 10,000 feet pressure altitude should be discouraged. However, there are operations where a lower maximum altitude has no Page D - 169 AC 29-2C 9/30/99 effect on the air traffic system and would still be operationally useful. Figures 3 and 6 of Appendix C, Part 29, relate the variation of average LWC as a function of cloud horizontal extent. These relationships should be used for design assessment of the most critical combinations of conditions as a function of en route distance. This, in combination with a capability to hold in icing conditions for 30 minutes at the destination, is commensurate with policies previously established for fixed-wing aircraft. Figures 3 and 6 should be used in conjunction with the altitude limited criteria of figures AC 29.877-1 through AC 29.877-4 herein. The new criteria of figure AC 29.877-5 includes “duration” (horizontal extent) as the third dimension. It is emphasized that LWC extremes expressed in Part 29, Appendix C, criteria and the alternate envelopes represent the maximum average values to be anticipated within an exceedance probability of 99.9 percent. Transient, instantaneous peak values of much higher LWC have been observed. These instantaneous peak values appear to be of little significance to the design of protected and unprotected surfaces; however, these high values, if encountered, may induce shedding of ice from some unprotected surfaces. This is due to radical changes in the rate of release of latent heat and resultant changes in the structural properties and adhesion force of ice.

(iii) A recent analysis performed at the FAA Technical Center concludes that the aircraft icing environment below 10,000 feet is not as severe in terms of LWC and OAT as that depicted in Part 29, Appendix C, envelope. This AC presents two different altitude limited envelopes that may be employed by those applicants who elect to certify with a 10,000-foot pressure altitude limit. One of these altitude limited envelopes is based upon the same data that were used to derive the design criteria of the Part 29, Appendix C (figures AC 29.877-1 thru 4), while the other is based upon a recently established characterization of supercool clouds below 10,000 feet (figure AC 29.877-5). The applicant may select either of the approaches to altitude limitation. At the present time, applicants have not consistently selected one or the other. If experience shows a unanimous preference for one or the other, the one not used will be deleted in a future revision. The data used to derive these limited envelopes cannot be used to further define icing conditions between 10,000 feet and 22,000 feet; hence, above 10,000 feet, the Part 29, Appendix C, envelopes should be used. It should be noted that the engine inlets should still meet the icing requirements of § 29.1093. The limited icing envelopes may be used on an equivalent safety basis to show compliance with the intent of § 29.1093 if the altitude limit established for the rotorcraft is not greater than 10,000 feet.

(iv) Significant effects can result from various combinations of parameters.

For example, most rapid ice accumulations occur at the high values of liquid water content, and the greatest impingement area occurs at the high values of droplet size.

Most critical ice shapes are a function of each of these parameters in addition to airspeed, surface temperature, and surface contour. Care should be taken to explore the entire specified ranges of these parameters during the design, development, and certification efforts.

Page D - 170 9/30/99 AC 29-2C (v) Mixed conditions (i.e., a combination of ice crystals and supercooled water droplets) and freezing rain or freezing drizzle are not addressed in the Part 29 environmental criteria but can present more severe icing conditions than those defined.

Although the probability of encountering freezing rain is relatively low, mixed conditions commonly occur in supercooled cloud formations. Little data have been gathered on the effects of encountering mixed conditions (see reference AC 29.877d(7)). There are no criteria for certification in mixed conditions or freezing rain at present. In addition to the hazards of operating any aircraft in icing, certain aspects of rotorcraft icing (relatively low altitude operation, asymmetric shedding with resulting vibration, and ice damage or ingestion) warrant a caution notice in the RFM advising that the rotorcraft is not certified for operation in freezing rain or freezing drizzle. Avoidance procedures (e.g., climb or descent) may also be useful.

(6) Flight Test Prerequisites .

(i) The prototype rotorcraft should be capable of IFR and IMC flight.

(ii) Sufficient analyses should be developed, submitted, and accepted by FAA/AUTHORITY to show that the rotorcraft is capable of safely operating to the selected design points of both the continuous maximum and intermittent maximum conditions of Part 29, Appendix C, or one of the altitude limited icing envelopes. A detailed failure modes and effects analysis (FMEA) should be performed.

(iii) Specific attention should be given to (1) assuring that the selected design condition(s) of atmospheric and rotorcraft flight envelopes have been identified; (2) qualification and design of ice protection systems and components; and (3) component installation and ice formation effects upon basic rotorcraft structural properties and handling qualities. These assurances can be established from analyses, bench test, and/or dry air flight tests or simulated icing tests, as appropriate prior to flight tests in natural icing.

(iv) The applicant should assess rotor blade stability with ice deposits to assure that dynamic instability will not occur in icing conditions. This assessment may be accomplished by analysis including consideration of failure of the most critical segment of the rotor blade ice protection system. It also may be accomplished by experimental means such as attaching dummy ice shapes to the blades and using a whirl stand or wind tunnel.

c. Procedures .

(1) Compliance .

(i) In general, compliance can be established when there is reasonable assurance that while operating in the specified icing environment (1) the engine(s) will not flameout or experience significant power losses or damage; (2) stress levels are not reached with ice accumulations that can endanger the rotorcraft or cause serious Page D - 171 AC 29-2C 9/30/99 reductions in component life; (3) the handling qualities, performance, visibility, and systems operation are defined and are not deteriorated unacceptably; (4) inlet, vent or drain blockage (such as fuel vent, engine, or transmission cooler) is not excessive; and (5) autorotation characteristics are acceptable with maximum ice accretion between de - ice cycles. Assessment of performance loss should include not only the drag and weight of the ice itself but electrical or other load demands of the ice protection system and any performance changes resulting from modified rotor blade contours.

(ii) It is emphasized that ice formations (shape, weight, etc.) vary significantly under varying conditions of outside air temperature (OAT), liquid water content (LWC), median volume diameter (MVD), airspeed, attitude, and rotor RPM. The most critical conditions should be defined by means of analyses or test and verified by test. Performance changes under these various conditions should be determined and found acceptable.

(iii) Laboratory, icing tunnel, ground spray rig, and airborne icing tanker tests are all very useful in developing an ice protection capability, but none of these, either individually or collectively, can satisfy the full requirements for certification. None can presently duplicate the combinations of liquid water content, droplet size, flow field, and random shedding patterns found in natural icing conditions. Airborne tankers hold considerable promise of being able to fulfill certification requirements (in addition to the advantage of being able to produce an icing environment on demand rather than having to wait for it to occur in nature), but tankers have not been able to generate droplet sizes that cover the complete envelope for certification. Many improvements have been made in some tankers in recent years; however, large droplet sizes have typically been a problem. Also, the size of existing tanker clouds is not of sufficient cross section to immerse the entire rotorcraft. There are also solar radiation and relative humidity effects to be considered and correlated with natural icing when using a tanker. The tanker should be able to immerse the entire rotor system as a minimum and should have a means of controlling and changing the cloud characteristics uniformly and repeatably. Until an artificial method has been successfully demonstrated and accepted, icing certification should include flight tests in natural icing conditions.

(iv) Flight testing in natural icing conditions also has limitations.

Reference AC 29.877d(16) contains information that may be useful in planning natural icing flight tests. The key limitation of natural icing flight tests is being able to find the combinations of conditions that comprise critical design points. This is especially true of those points falling near the 99.9 percentile of exceedence probability; e.g., high LWC at low OAT with large MVD. It is emphasized that some more severe design points, however, may exist within the atmospheric icing envelope rather than near the edges or corners of the envelope. This does not mean that natural icing tests must be conducted at all the selected design conditions. Natural icing tests should be conducted in conditions as close to design points as possible and sufficient correlation shown with the analyses to assure that the rotorcraft can operate safely throughout the design envelope.

Page D - 172 9/30/99 AC 29-2C (v) Certification flight testing should be extensive enough to provide reasonable assurance that either induced or random ice shedding does not present a problem. The most likely indication of a problem if it exists will be ice impact on the airframe or rotor imbalance resulting in vibration. The following should be considered sufficient for rejection: (A) Vibrations sufficient to make the instruments difficult to read accurately.

(B) Vibrations sufficient to exceed the structural or fatigue limits of any rotorcraft part such as blade, mast, or transmission components.

(C) Ice impact damage to essential parts, such as the tail rotor, that could create a flight hazard. Cosmetic, nonstructure flaws that do not exceed wear and tear characteristics or maintenance criteria are acceptable. Any ice shedding effects that require immediate maintenance action are unacceptable.

(vi) There should be a means identified or provided for determining the formation of ice on critical parts of the rotorcraft which can be met by a reliable and safe natural warning or an ice detection system. A system utilizing OAT must include an accurate OAT measurement since the onset of icing can occur in a very narrow temperature band requiring sensitive and accurate OAT measurement. OAT accuracy should be relative to the true temperature of the air mass. Total system accuracy should be ±0.5° C in the -5.0° to +5.0° C range and ±1° C throughout the remaining temperature range. The location of the sensor has been shown to be very critical and, in effect, there can be a position error or other errors induced by ice formations or solar radiation. If the system measures liquid water content, consideration should be given to the fact that the actual LWC fluctuates considerably as the rotorcraft passes through an icing environment. A warning system displaying or utilizing a peak or average LWC value (rather than an instantaneous readout) should include sufficient conservatism to provide a margin of safety. The value of an LWC detecting system lies in its utility as a warning that ice is being encountered. The actual magnitude of LWC in combination with OAT and MVD can be used to indicate the icing severity level. The U.S. Army is currently developing an advanced ice detection system for potential application to rotorcraft.

(2) Instrumentation and Data Collection .

(i) Instrumentation proposed for certification tests, including flight strain surveys, should be reviewed as early as possible in the program to establish that it will provide the necessary data. The need for accurate OAT measurement previously noted for operation in icing also applies to the certificated configuration. Mechanical devices such as the rotating multicylinder and rotating disc have been used for measuring ice accretion rate which is relatable by calibration to average LWC and MVD. More recently, hybrid mechanical/electronic LWC measuring devices have been used.

Devices that rely on ice accretion as a signal source are subject to the Ludlam limit (the Page D - 173 AC 29-2C 9/30/99 limits whereby latent heat of fusion is not totally absorbed, thus resulting in incomplete freezing of the moisture and some inaccuracy in the indication). The Ludlam limit is a function of various parameters including OAT, airspeed, LWC, and MVD. The Ludlam limit may vary from one device to another. (See references AC 29.877d(8) and AC 29.877d(9)(i) for further information). Gelatin slides, soot and oil slides, and more recently, laser nephelometers, have been used to measure droplet size. Other calibrated devices intended for measurement of LWC should be used.

Reference AC 29.877d(16) describes several of these devices. Photographic coverage of critical areas may be necessary to ascertain that ice protection systems are functioning properly and that there are no runback problems. (The term “runback” refers to liquid water that has not been evaporated by surface de-ice equipment and flows back to an unheated area subject to freezing.) Reference AC 29.877d(19) highlights use of video techniques and equipment for this purpose. Some systems will require acceptable calibration techniques and data.

(ii) Gelatin, soot, and oil slides provide data that can be used to estimate MVD at discrete intervals while laser nephelometer data can provide time histories of MVD droplet size distributions. Gelatin slide data should be taken frequently during test flights to properly characterize the cloud. Laser nephelometer data have been found to be highly dependent upon knowledge of the equipment and calibration. Proper calibration, maintenance, and data processing techniques should be utilized and demonstrated. Additional information on the subject may be found in Reference AC 29.877d(18).

(iii) Structural instrumentation requirements should also be established as early as possible in the program. Flight strain measurements are strongly recommended in assessing the ice imposed stress on the rotorcraft. The flight strain measurements should determine the effect on fatigue life due to ice accumulation for such items as main rotor blades, main rotor hub components, rotating and fixed controls, horizontal stabilizer, tail rotor, etc. The subsequent proper operation of retractable devices such as landing gear should be demonstrated with representative ice accretion. In addition, the static and fatigue strength of the blade with heater mat must be substantiated. Any effect of the heater mat on fatigue strength of the blades must be considered.

(3) Additional Considerations . The following are items to consider in an icing certification program. They are not intended to be all-inclusive, and the possibility of widely differing characteristics and critical areas among various rotorcraft in icing should be considered.

(i) The rotorcraft should be shown by analysis and confirmed by either simulated or natural icing tests to be capable of holding for 30 minutes in the design conditions of the continuous maximum icing envelope at the most critical weight, CG, and altitude with a fully functional ice protection system. For those applicants who elect to certify their rotorcraft to the new supercooled cloud characterization of figure AC 29.877-5, the rotorcraft should be shown by analysis and confirmed by either Page D - 174 9/30/99 AC 29-2C simulated or natural icing tests to be capable of holding for 30 minutes in the design conditions of the icing envelopes up to a maximum of 0.8 grams per cubic meter of LWC at the most critical weight, CG, and altitude.

(ii) A single ice protection system and power source may be considered acceptable provided that after any single failure of the ice protection system, the rotorcraft can be shown by analysis and/or test to be capable of safe operation (no hazard) for 15 minutes following failure recognition in the continuous icing envelope used as the basis for certification within the same icing limits used for the 30-minute hold criteria. During this 15-minute period the rotorcraft may exhibit degraded characteristics. Pilot controllable operating limitations such as airspeed may be used to satisfy this continued safe flight criteria. For purposes of determining performance and handling qualities degradation, ice protection system failure need not be considered to occur simultaneously with engine failure unless ice protection system operation is dependent upon engine operation.

(iii) Although current airborne weather radar technology systems may be useful in avoiding potential icing conditions by detecting precipitation, the use of weather radar is not an FAA/AUTHORITY requirement for icing certification.

(iv) If the ice protection is not operating continuously, there must be a means to advise the crew when the rotorcraft is in icing conditions in order that the system may be activated.

(v) No autorotational performance data is required for rotorcraft which have Category A powerplant installations. All rotorcraft certified for flight in icing conditions must be capable of full autorotational landings with the ice protection system operating. Autorotational entry, steady state, and flare entry flying qualities and performance should be evaluated with an ice load. Since the Category A en route performance can vary as the ice protection system operates, a mean value of cyclic torque is acceptable provided at no time does the rate of climb fall below zero. The rotorcraft is assumed to be clear prior to takeoff, and therefore the takeoff performance is not degraded. The landing performance can be based on the in-flight assessment of overall performance degradation. Items such as fuel burns can be used as part of the in-flight performance degradation determination. Regardless of the methods used to determine performance degradation, it must be easily used by the crew. The hover performance should be addressed for the termination of a flight after an icing encounter.

The engines must be protected from the adverse effects of ice. When ice does accumulate on the inlets, screens, etc., it must be accounted for in performance, engine operating characteristics, and inlet distortion.

(vi) The handling qualities of the rotorcraft must be substantiated if ice can accumulate on any surface. When ice can accumulate on unprotected surfaces the rotorcraft must exhibit satisfactory VFR/IFR handling qualities. In addition, following the failure of the de-ice system, the rotorcraft must be safely controllable for 15 minutes, i.e., the rotorcraft must be free from excessive and rapid divergence. Artificial ice Page D - 175 AC 29-2C 9/30/99 shapes may be acceptable for acquisition of flight test data necessary for handling qualities and performance evaluations and demonstrations.

(vii) Items such as fuel tank vents, cooling vents, antennas, etc., must be substantiated for maximum icing effects.

(viii) The ice protection system should be sufficiently reliable to perform its intended function in accordance with the requirements of § 29.1309. These requirements may in some instances be met by the use of sound engineering judgment during design and compliance demonstrations. In many instances, use of good design practices, failure modes and effects analysis, and similarity analyses combined with good judgment will be adequate. In some instances the need for reliability analyses may be desirable. Additional information pertaining to reliability is contained in paragraph AC 29.1309 (§ 29.1309).

(ix) The subject of lightning must be addressed. The criteria applied on rotorcraft with ice protection systems are that “the rotorcraft must be protected in such a manner to minimize lightning risk.” The general rules of § 29.1309(a), (b), and (c) are applicable to assure adequate lightning protection.

(x) Ice protection of pitot-static sources, windshields, inlets, exposed control linkages, etc., must be considered.

(xi) The impact of ice protection system failure, complete and partial, and achieving adequate warning thereof must be assessed.

(xii) The impact of delayed application of ice protection systems should be assessed. Hazardous conditions should not be apparent. Any rotorcraft characteristic changes resulting should be covered in cautionary material in the rotorcraft flight manual.

(xiii) Possible droop stop malfunction with ice accumulation and its potential hazard to the rotorcraft, its occupants, and ground personnel must be assessed.

(xiv) Possible ice shedding hazards to ground personnel or equipment in proximity to turning rotors following flight in icing conditions should be given much consideration.

(4) Flight Manual . Areas of the flight manual which may require inputs are: (i) Operating limitations including approved types of operation and prohibiting operation in freezing rain or freezing drizzle conditions. Avoidance procedures may also be useful.

Page D - 176 9/30/99 AC 29-2C (ii) Normal Operating Procedures. Information on the ice detection means or system and ice protection system and its capabilities.

(iii) Emergency Operating Procedures. Operating procedures containing essential information particularly with system failure.

(iv) Caution Notes. These caution notes should advise or address: (A) Against inducing asymmetric shedding with rapid control inputs or rotor speed changes, except possibly as a last resort. Rotor speed changes appear to be more effective than control inputs in removing ice from the rotor blades of some rotorcraft.

(B) Loss in range, climb rate, and hover capability following prolonged operation in icing.

(C) The need for clean blade surfaces and use of approved cleaning solvents or ground deicing/anti-icing agents prior to starting rotors.

(D) Changes in autorotational characteristics resulting from formations.

(E) If the rotorcraft has been certificated for flight in supercooled clouds and falling and blowing snow, flight in other conditions such as freezing rain, freezing drizzle, sleet, hail, and combinations of these conditions with supercooled clouds should be avoided.

(F) The potential hazards to ground personnel, passengers deplaning, and equipment in proximity to turning rotors following flight in icing conditions.

d. Icing References .

(1) FAA Technical Report ADS-4, Engineering Summary of Airframe Icing Technical Data, December 1963.

(2) A New Characterization of Supercooled Clouds Below 10,000 Feet DOT/FAA/CT-83/22, June 1983.

(3) Advisory Circular 20-73, Aircraft Ice Protection, 21 April 71.

(4) Advisory Circular 91-51, Airplane De-ice and Anti-ice Systems, 9/15/77.

(5) FAA Report RD-77-76, Engineering Summary of Powerplant Icing Technical Data, July 1977.

(6) United States Army Aviation Engineering Flight Activity Reports: Page D - 177 AC 29-2C 9/30/99 (i) Natural Icing Tests, UH-1H Helicopter, Final Report, June 1974, USAASTA Project No. 74-31.

(ii) Artificial Icing Tests, UH-1H Helicopter, Part 1, Final Report, January 1974, USAASTA Project No. 73-04-4.

(iii) Artificial Icing Tests, UH-1H Helicopter, Part II, Heated Glass Windshield, Final Report, USAASTA Project No. 73-04-4.

(iv) Artificial Icing Tests, Lockheed Advanced Ice Protection System Installed on a UH-1H Helicopter, Final Report, June 1975, USAAEFA Project No. 74-13.

(v) Artificial and Natural Icing Tests for Qualification of the UH-1H, Kit A Aircraft, Letter Report, USAAEFA Project No. 78-21-1.

(vi) Microphysical Properties of Artificial and Natural Clouds and Their Effects on UH-1H Helicopter Icing, Report USAAEFA Project No. 78-21-2.

(vii) Helicopter Icing Spray System (HISS) Nozzle Improvement Evaluation, Final Report, September 1981, USAAEFA Project No. 79-002-2.

(viii) Artificial and Natural Icing Tests of the YCH-4TD, Final Report, May 1981, USAAEFA Project No. 79-07.

(ix) Limited Artificial Icing Tests of the OV-ID, Letter Report, July 1981, USAAEFA Project No. 80-16, (Limited Distribution).

(x) JUH-IH Ice Phobic Coating Tests, Final Report, July 1980, USAAEFA Project No. 79-02.

(xi) Artificial and Natural Icing Tests, Production UH-60A Helicopter, Final Report, June 1980, USAAEFA Project No. 79-19.

(xii) Helicopter Icing Spray System (HISS) Evaluation and Improvements, Letter Report, June 1981, USAAEFA Project NO. 80-04.

(xiii) Artificial Icing Test of CH-47C Helicopter with Fiberglass Rotor Blades, Final Report, July 1979, USAAEFA Project No. 78-18.

(xiv) Limited Artificial and Natural Icing Tests, Production UH-60A Helicopter (Reevaluation), Final Report, August 1981, USAAEFA Project No. 80-14.

(7) Further Icing Experiments on an Unheated Nonrotating Cylinder, National Research Council, Canada Report LTR-LT-105, dated November 1979, by J.R. Stallabrass and P.F. Hearty.

Page D - 178 9/30/99 AC 29-2C (8) Ludlam, F.H., Heat Economy of a Rimed Cylinder, Quarterly Journal, Royal Meteorological Society, Vol. 77, 1951.

(9) U.S. Army AMRDL Reports: (i) USAAMRDL TR 73-38, Ice Protection Investigation For Advanced Rotary Wing Aircraft, J.B. Werner, August 1973, AD 7711182.

(ii) Werner, J.B., The Development of an Advanced Anti-Icing/Deicing Capability for U.S. Army Helicopters, Volume 1, Design Criteria and Technology Considerations, USAAMRDL - TR-75-34A, Eustis Directorate, U.S. Army Air Mobility R&D Laboratory, November 1975, AD A019044.

(iii) Werner, J.B., The Development of an Advanced Anti-Icing/Deicing Capability for U.S. Army Helicopters, Volume 2, Ice Protection System Application to the UH-1H Helicopter, USAAMRDL - TR-75-34B, Eustis Directorate, U.S. Army Air Mobility Research and Development Laboratory, November 1975, AD A019049.

(iv) USAAMRDL-TR-76-32, Ottawa Spray Rig Tests of an Ice Protection System Applied to the UH-1H Helicopter, November 1976, AD A0034458.

(v) USARTL-TR-78-48, Icing Tests of a UH-1H Helicopter with an Electrothermal Ice Protection System Under Simulated and Natural Icing Conditions, April 1979.

(vi) USAAMRDL-TR77-36, Final Report, Natural Icing Flights and Additional Simulated Icing Tests of a UH-1H Helicopter Incorporating an Electrothermal Ice Protection System, July 1978, AD A059704.

(10) Technical Feasibility Test of Ice Phobic Coatings for Rain Erosion in Simulated Flight Conditions, U.S. Army Test and Evaluation Command, Final Report, 4-AI-192-IPS-001, August 1980.

(11) Technical Feasibility Test of Ice Phobic Coatings in Simulated Icing Flight Conditions, U.S. Army TECOM, Final Report, 4-CO-160-000-048, September 1980.

(12) Aircraft Icing, NASA Conference Publication 2086, FAA-RD-78-109, July 1978.

(13) Helicopter Icing Review, FAA Technical Center, Final Report, FAA-CT-80-210, September 1980.

(14) National Icing Facilities Requirements Investigation, Final Report, FAA Technical Center, FAA-CT-81-35, March 1981.

(15) Aircraft Icing, AGARD Advisory Report No. 127, November 1978.

Page D - 179 AC 29-2C 9/30/99 (16) Rotorcraft Icing - Review and Prospects, AGARD Advisory Report, AR-166, September 1981.

(17) Advisory Circular 20-117 - Hazards Following Ground Deicing and Ground Operations in Conditions Conducive to Aircraft Icing, Dec. 17, 1982.

(18) Olson, W., Experimental Comparison of Icing Cloud Instruments, January 1983, NASA TM 83340.

(19) JUH-1H Redesigned pneumatic boot deicing system flight test evaluation.

Hayworth, L., Graham, M., to be published. USAAEFA Edwards AFB, California.

Project No. 834-13.

(20) An Appraisal of the Single Rotating Cylinder Method of Liquid Water Content Measurement, National Research Council Canada Report LTR-LT-92, dated November 1978, by J.R. Stallabrass.

Page D - 180 9/30/99 AC 29-2C LL w • 20 w 0::

~

ffi

~ 10 w I- ~ w ai ~ 0 0 2000 4000 6000 8000 10,000 PRESSURE ALTITUDE - FT FIGURE AC 29.877-1 CONTINUOUS ICING - TEMPERATURE VS ALTITUDE LIMITS Figures AC 29.877-1 through 4 represent one approach to a 10,000-foot altitude limit and Figure AC 29. 877-5 represent another. See Paragraph 386b( 5)(i ii) for a discussion of the individual application of the tvvo approaches Page D - 181 AC 29-2C 9/30/99 LL w ' 20 w 0:: :::) ~ 0:: w 0...

w I- I- z w CD <( -10 0 2000 4000 6000 8000 10,000 PRESSURE ALTITUDE - FT FIGURE AC 29.877-2 INTERMITTENT ICING - TEMPERATURE VS ALTITUDE LIMITS Figures AC 29.877-1 through 4 represent one approach to a 10,000-foot altitude limit and Figure AC 29.877-5 represents another. See Paragraph 386b(5)(iii) for a discussion of the individual application of the two approach es Page D - 182 9/30/99 AC 29-2C 1.0 SEA LEVEL TO 10,000 FT PRESSURE ALTITUDE 0:: w I- w 0.8 ~ () ai :::) "1~ () (f)

-

~~ ~ ;,z,,)'.

~ 0.6 0:: <Y A~ ~ ~..o ,..___ ' I- z w I- z () 04 0:: w ~

s

:::) 0.2 a _J

0-----------------------

0 10 20 30 40 50 MEAN EFFECTIVE DROP DIAMETER- MICRONS FIGURE AC 29.877-3 MEAN EFFECTIVE DROP DIAMETER- MICRONS Figures AC 29.877-1 through 4 represent one approach to a 10,000-foot altitude limit and Figure AC 29.877-5 represents an other. See Paragraph 386b(5)(iii) for discussion of the individual application of the two approach es Page D - 183 AC 29-2C 9/30/99 2.0 SEA LEVEL TO 10,000 FT PRESSURE ALTITUDE

ffi

ti:i

1.6 ~ AMBIENT TEMP () CD

a

(7) 1.2 ~ 0:: ~ w ~ 0.8 O () 0::

~

04 o :::) G _J 0 10 20 30 40 50 MEAN EFFECTIVE DROP DIAMETER - MICRONS FIGURE AC 29.877-4 INTERMITTENT ICING - LIQUID 'MTER CONTENT VS DROP DIAMETER Figures AC 29 .877-1 through 4 represent one approach to a 1 D ,DO □ -foot altitude Iimit and Figure AC 29.877-5 represents another. See Paragraph 386b(5)(iii) for a discussion of the individual application of the two approaches.

Page D - 184 9/30/99 AC 29-2C SUPERCOOLED CLOUDS GROUND LEVEL TO 10,000 FEET 2.0 I?' 1.74 E 0) t- 1.5 z w I-

/

z 0to-15 ° C () 0:: w 1.0 ~

s

. 66 :::) _J 0.5 .L1 MEDIAN VOLUME DIAMETER (um) FIGURE AC 29.877-5 Fig1irAs AC'. ?~ 877-1 through 4 rArresAnt onA fll')l')roi=lr.h tn i'l 10,000-foot i=!ltit11rlil limit and Figure AC 29.877-5 represents another. See Paragraph 386b(5)(iii) for dis:ussion of the i1 divi du al application of the tvvo approaches Page D - 185 9/30/99 AC 29-2C CHAPTER 2. PART 29 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT SUBPART E - POWERPLANT POWERPLANT - GENERAL AC 29.901. § 29.901 (Amendment 29-17) INSTALLATION .

a. Section 29.901(a) : (1) Explanation . Paragraph (a) provides a definition of areas of rotorcraft for which safety requirements are set forth under the general title, SUBPART E - POWERPLANT. This subpart includes not only major propulsive elements and power transmissive components but also powerplant controls and instruments, safety devices, including fire protection and other devices to protect personnel, and critical flight structure in event of fires.

(2) Procedures . To ensure that no certification aspect is overlooked in establishing compliance, certification engineers should make at least an informal breakdown of all components of the rotorcraft, assigning responsibility to powerplant certification engineers of all items within the above definition. While this procedure is usually straightforward, the following items of FAA/AUTHORITY powerplant responsibility are listed to minimize questions regarding authority and responsibility.

(i) Drive system components . All parts of the transmission, clutches, shafting, including the driveshafts (masts) of main and auxiliary rotors, powerplant cooling components, and powerplant instrumentation requirements under §§ 29.1305, 29.1337, 29.1543, 29.1549, 29.1551, 29.1553, 29.1555, and 29.1583.

NOTE: The division of responsibility between FAA/AUTHORITY airframe engineers and FAA/AUTHORITY powerplant engineers (in accordance with FAA/AUTHORITY practice) regarding the driveshaft is at the flange or spline interface between the driveshaft and the rotor hub. Rotor hubs, controls, blades, and associated components are the airframe engineers’ responsibility. (Industry practice may not agree with this concept.)

(ii) Engines, except for mount structure.

(iii) Auxiliary power units, except for mount structure.

(iv) Combustion heaters, except for downstream ventilation air ducting, mixing, and distribution systems and for electrical aspects of controls and safety devices.

Page E - 1 AC 29-2C 9/30/99 (v) Water/alcohol or other fluid power augmentation systems.

(vi) Engine induction systems including induction icing and snow ingestion, and exhaust systems, including exhaust shrouds and drains.

(vii) All fuel systems, including those serving engines, auxiliary power units, combustion heaters, power augmentation systems, etc., and vents and drains for those systems.

(viii) Oil systems for engines, auxiliary power units, rotor drive transmissions, and gearboxes, including grease lubricated gears and bearings of the drive system.

(ix) Cooling aspects of engines, rotordrive transmissions and gearboxes, and auxiliary power units (APU). Electrical generating equipment and hydraulic component cooling may be the responsibility of the systems and equipment engineer provided agreement is established among responsible personnel.

(x) Rotor brakes, except hydraulic, electrical, and structural aspects of nonrotating brake components.

(xi) Fire protection, including firewalls, fire extinguisher systems, fire detector systems, flammable fluid lines, fittings, and shutoff valves. The powerplant engineer has responsibility for evaluating compliance with §§ 29.861 and 29.863 as they pertain to fuel and oil systems.

(xii) Engine and transmission cowling and covering, including latches.

(xiii) Powerplant flexible controls (reference § 29.1141(c)).

(xiv) Powerplant accessories.

(xv) Pneumatic systems (engine bleed air) within the engine or APU compartments, including shut-off valves and engine isolation features of bleed systems.

(xvi) Powerplant aspects of instrument markings and powerplant aspects of flight manuals, including limitations, normal and emergency procedures, engine performance; powerplant aspects of maintenance manuals, with emphasis on the limitations section of the manual and verification of the limitations established under § 29.1521.

b. Section 29.901(b) : (1) Explanation . Paragraph (b) requires compliance with the engine manufacturers’ approved installation instructions and any applicable provisions of this subpart that the powerplant installation must be installed in a manner to ensure Page E - 2 9/30/99 AC 29-2C continued safe operation, that accessibility for inspection and maintenance is provided, that appropriate electrical connections (ground connections) are provided, and that allowance is provided for thermal expansion of turbine engines.

(2) Procedures .

(i) Engine Installation . Compliance with most of the detail requirements in the engine installation manual can be established by test or by design features and arrangements negotiated between the rotorcraft manufacturer and the FAA/AUTHORITY powerplant engineer. Some aspects, usually involving inlet and/or exhaust distortion limitations, vibration limitations, and aircraft/engine interface items may require direct assistance and information from the engine manufacturer to determine that compliance with the installation manual exists. Fuel control/engine/rotor system torsional matching is usually a developmental problem to be worked out before presentation of the rotorcraft to the FAA/AUTHORITY; however, final flight tests for surge or stall, torsional stability, and acceleration/deceleration schedules may require direct coordination among FAA/AUTHORITY installation engineers, engine manufacturers’ representatives, and the FAA/AUTHORITY engine certification engineers. These items are addressed specifically under § 29.939. Reciprocating, carburetor-equipped engines usually require a particular carburetor configuration to achieve adequate engine cooling. This configuration, identified as a “carburetor parts list,” must be approved for the engine under Part 33 and should be listed with the engine on the type data sheet for the rotorcraft.

(ii) Arrangement and Construction . Each item of the powerplant area of responsibility should be shown to be suitable for its intended purpose and installed to operate satisfactorily and safely between normal inspections and overhauls.

Accessories mounted on engine or transmission drive pads should be determined to be compatible with the pad limits including fit and speed range, overhang moment loads, running torque, and static torque. This latter term pertains to protection of the engine or transmission, which drives the accessory, from damage to be expected from malfunction of the accessory. This protection is usually supplied by providing a shear section in the accessory drive shaft designed to fail before exceeding the static torque limit of the engine or transmission driving component. Note that when evaluating the strength of the mechanical shear section, material allowables quoted in materials handbooks should not be used since these are minimum strength values. Shear sections should consider maximum strength values to be expected which are on the order of 130 percent of the minimum strength values. Also, it should be verified that design data for shear sections are dimensioned to limit the maximum diameter as well as the minimum diameter. Installation of starter-generators may also require verification that horsepower extraction limits are not exceeded. Special flightcrew instructions in the flight manual to monitor generator load or to disconnect electrically loaded items to protect accessory or engine-transmission pad limits should be avoided.

Environmental qualification requires consideration or protection against adverse effects of heat, sand or dust, humidity and rain, salt-laden atmosphere, and extremes of cold weather. Accessories such as generators, pumps, etc., are subjected to many of these Page E - 3 AC 29-2C 9/30/99 aspects during the individual qualification tests; however, satisfactory overall integrated system performance under these adverse conditions should be verified. Cold weather testing should include verification that lubricating oils and greases function properly and that engine starting procedures are safe and do not impose excessive loads on accessories, engines, or drive system components. Powerplant engineers should coordinate compliance efforts in this area with the system engineer’s investigations of compliance with §§ 29.1301 and 29.1309. Full-scale rotorcraft operations in cold weather should be required. Performance tests are required at the minimum temperature to be certified. Propulsion systems may usually be evaluated at this time.

Cold soak or overnight exposure to cold weather is appropriate followed by starting and pretakeoff procedures in accordance with the flight manual. Attention should be given to the practicality of important mandatory inspection procedures as affected by cold weather.

(iii) Accessibility . Accessibility for maintenance should be reviewed.

Typically, some maintenance activities must involve disassembly or removal of adjacent components. This should be avoided if repetitive activity can jeopardize the performance of critical or safety-related equipment. Verify that easy access exists to items such as oil system sight gauges or dip sticks, filler ports and drain valves for engines, auxiliary propulsion units, transmissions, fuel tanks and filters, etc.

(iv) Electrical (Grounding) . Electrical interconnections to prevent difference of potential should be provided in the form of grounding straps or wires sized to carry the currents to be expected. Verify that the attachments for these grounding devices are not compromised by paint or zinc chromate which will tend to electrically insulate the engine or component. Note that engine mount structure should not be accepted as a grounding device since electrical current will cause corrosion at attachment points.

(v) Thermal Expansion . Axial and radial expansion of turbine engines is usually not a problem unless redundant mount arrangements are used. Special expansion provisions are usually required if engine components other than mounting points are attached to bulkheads, firewalls, other engines, or drive system components.

Engine output shaft axial or bending loads due to thermal expansion and to deflection of supports under ground or flight loads should be checked. Other components of concern are compressor inlet flanges, exhaust ducts, and rigid fluid or air lines between aircraft structure and the engine. The engine installation data will provide limit loads to be considered for parts of the engine which normally are attached to airframe components.

c. Section 29.901(c) : (1) Explanation . Paragraph (c) requires, with notable exceptions, a detailed failure modes and effects analysis (FMEA) of the various powerplant systems and components to establish that anticipated failures will not jeopardize the safe operation of the rotorcraft. Alternative methods such as top-down analysis may also be used.

Page E - 4 9/30/99 AC 29-2C Exceptions include engine rotor discs and structural elements for which the probability of failure can be shown to be “extremely remote.” Items in this latter case would include all components of the rotor drive system evaluated under § 29.571 provided that the reliability of any item or system exempted under § 29.901(c)(1) is not jeopardized by the failure of other systems/components which themselves may be less reliable than “extremely remote.” Items of consideration here would include, but not be limited to, powerplant cooling systems, probable maintenance errors, deterioration/failure of seals and other time/temperature/weather sensitive nonmetallics, high energy fragment impact damage of nearby dynamic components, etc. Some items in these categories are addressed by specific rules in this subpart which override consideration under § 29.901(c). For example, § 29.927 sets forth specific tests to demonstrate acceptable safety levels in event of overtorque, overspeed, and transmission lubrication system failures. Further consideration of failures in these areas (under § 29.901(c)(1)) probably would be inappropriate. It would not, however, be appropriate to assume that an engine certified under Part 33, an auxiliary power unit qualified under TSO C-77, or other components qualified under various TSO’s or military specifications would not be subject to failure. As a general rule, any component or system whose failure is “probable” and the failure, in conjunction with probable combinations of failures, significantly degrades safe operation and/or impairs the capability of the crew to operate the rotorcraft safely constitutes an apparent noncompliance unless it is compensated for by alternate components, systems, or if appropriate, special operating procedures which essentially restore a safe level of operation of the rotorcraft. Normally, safe “continued” flight is intended; however, for the special case of the single-engine rotorcraft, safe entry into autorotation after engine failure is an acceptable means of compliance provided that other coincidental or associated failures or malfunctions do not jeopardize this maneuver.

(2) Procedures .

(i) The general techniques of AC 25.1309-1, System Design Analysis, present an acceptable means of evaluating the powerplant systems/components for compliance. However, the quantitative assessments of the probability classifications in AC 25.1309-1 have not been universally adopted for powerplant systems and components. Other procedural techniques in AC 25.1309-1 may be impractical for powerplant systems. This does not preclude using a similar but simplified methodology in conjunction with conservative engineering judgment to arrive at a determination of compliance or identification of noncompliance aspects, using the following as a guide (extracted from AC 25.1309-1). Develop a matrix of all applicable powerplant components/systems which includes: (A) Possible modes of failure, including malfunctions and damage from external sources.

(B) The probability of multiple failures and undetected failures.

Page E - 5 AC 29-2C 9/30/99 (C) The resulting effects of the rotorcraft and occupants, considering the stage of flight and operating conditions, and (D) The crew warning cues, corrective action required, and the capability of detecting faults.

(ii) Prepare an item-by-item, system-by-system FMEA. The analysis to identify failure conditions should be qualitative. An assessment of the probability of a failure condition can be qualitative or quantitative. An analysis may range from a simple report which interprets test results or presents a comparison between two similar systems to a fault/failure analysis which may (or may not) include numerical probability data. An analysis may make use of previous service experience from comparable installations in other aircraft.

(iii) Powerplant engineers normally find that believable statistical failure data on powerplant components are not readily available. Therefore, the simpler form of analysis involving assumption of failure with either benign results or dependence on alternate or redundant systems/components becomes the most feasible method of finding compliance. Repetitive inspections and preflight checks are a significant part of this finding, particularly if the backup system/component is used or checked routinely in the operation of the rotorcraft.

d. Section 29.901(d) : (1) Explanation . This paragraph provides a generalized basis for requiring compliance with any rules in this Part applicable to safe installation and operation of auxiliary power units (APU’s). The wording of the rule is generalized to permit (and require) a detailed review of this Part to identify any existing rule related to this type of equipment. Generally, any rule related to engines and their installation, support systems, and fire protection should be considered to be applicable to APU’s. This review may result in a designation of “nonapplicable” to certain engine-related rules if limitations such as “ground-use-only” are applied or if the APU serves only nonessential services. Any questionable aspects or interpretation/policy involved in establishing the applicable rules should be coordinated with the FAA Aircraft Certification Office.

Notwithstanding the generalization discussed above, a number of specific rules in subparts E and F include reference to APU’s in their applicability. The presence of these references should not be interpreted as excluding applicability of other appropriate rules as discussed above. In addition, the APU itself must be shown to be safe and reliable. Normally, this aspect is satisfied by showing that the APU model is included in the qualified parts list of TSO-C77a. This TSO also requires establishment (by the APU manufacturer) of limitations and installation data peculiar to the model APU. A showing of compliance with these data for the APU installed in the rotorcraft will be expected.

(2) Procedures .

Page E - 6 9/30/1999 AC 29-2C (i) Verify that the Model APU is listed as qualified to TSO-C77(a) or other suitable specifications. Note that TSO qualification is not regulatory but simply defines an acceptable base qualification standard. Other standards may be acceptable or deviations from the TSO may be acceptable if evaluated and found not pertinent to the planned installation.

(ii) Review the installation data provided for the APU and determine that the installation is in compliance. Exceptions may be taken as discussed above. Note that the TSO provides different qualification standards for “essential” and “nonessential” service APU’s. However, it does not distinguish between “flight-use” and “ground-use-only” APU’s. Some deviations to the TSO may be authorized based on this aspect; i.e., operation during negative “g” conditions.

(iii) Review Part 29, especially subparts E and F for all rules related to engines, engine support/service systems, intakes, exhausts, instrumentation, fire protection, pneumatic systems, etc., for applicability to installation and operation of the APU. Develop and accomplish a compliance program for the rules identified by this review following policy and procedures used for engines with exceptions which may be justified as discussed above.

(iv) For reference, the following rules specifically refer to APU’s. Some comments regarding compliance are offered.

(A) Section 29.1041, Cooling. APU installation data should define limits to be substantiated.

(B) Section 29.1091, Air Induction. Note the requirements of paragraph (f).

(C) Section 29.1103, Induction System Ducts. Note the special requirements of paragraphs (a), (e), and (f).

(D) Section 29.1121, Exhaust Systems.

(E) Section 29.1142, Controls.

(F) Section 29.1181, Designated Fire Zones.

(G) Section 29.1191, Firewalls. Firewall construction should be provided to completely separate the APU from other parts of the rotorcraft.

(H) Section 29.1195, Fire Extinguishers. Note that only one adequate discharge is required.

(I) Section 29.1203, Fire Detector Systems. Detectors are required for each fire zone which would include APU installations.

Page E - 7 AC 29-2C, Chg 3 9/30/2008 (J) Section 29.1305, Powerplant Instruments. TSO-C77(a) specifies provisions for measuring gas temperature, rotor RPM, and any other parameter necessary for safe operation of the APU.

(K) Section 29.1337, Powerplant Instruments.

(v) Additional comments . APU fuel sources which tap into engine fuel systems should be carefully designed and arranged to minimize the probability that an APU fuel line failure will jeopardize continued normal engine operation. If the APU provides essential services, it should be provided with an independent fuel system.

Also, engine fuel systems which operate at negative pressures should not be tapped for APU fuel source since air leaks back through the APU fuel control or small leaks in the APU fuel system likely will fail the engine.

AC 29.901A. § 29.901 (Amendment 29-26) INSTALLATION .

a. Explanation . Amendment 29-26 changes § 29.901(b)(2) to require a satisfactory determination that the rotorcraft can operate safely throughout adverse environmental conditions such as high altitude and temperature extremes. This amendment was needed to provide consistent application of the environmental qualification aspects of the installation. This amendment also added a new paragraph § 29.901(b)(6) to require design precautions to minimize the potential for incorrect assembly of components and equipment essential to safe operation.

b. Procedures . All of the policy material pertaining to this section remains in effect with the addition of design precautions. Design precautions should be taken to minimize the possibility of improper assembly of the components essential to the safe operation of the rotorcraft. Fluid lines, electrical connectors, control linkages, etc., should be designed so that they cannot be incorrectly assembled. This can be achieved by incorporating different sizes, lengths, and types of connectors, wires, fluid lines, and mounting methods. The applicant should perform a detailed maintenance assessment to clearly define the maintenance requirements, reliability, and serviceability of the drive system design. The applicant should consider all design qualification tests and service history data, if available. A review of accident data supports the importance of this assessment. Some applicants have utilized drive system vibration monitoring to verify continuing safe operation of their drive system.

AC 29.901B. § 29.901 (Amendment 29-36) INSTALLATION .

a. Explanation . Prior to Amendment 29-36, paragraph (c) exempted engine rotor disc failures (engine rotorburst) from consideration as a failure that could jeopardize the safe operation of the rotorcraft. Amendment 29-36 removes this exclusion. Therefore, engine rotor disc failures should be considered as a failure that would jeopardize the safe operation of the rotorcraft.

b. Procedures . The method of compliance for this section is unchanged.

Page E - 8 9/30/99 AC 29-2C AC 29.903. § 29.903 (Amendment 29-12) ENGINES .

a. Explanation . While paragraph (a) of this section requires engines to be type certificated under Part 33 of this chapter, engines certificated under other approved certification rules (CAR Part 13 and § 21.29 for imported engines) are also eligible. The fact that a component, system, or arrangement for which Part 29 standards exist is approved as a part of a certificated engine should not, except when specifically stated in Part 29, relieve an applicant of the necessity for compliance with Part 29. Even if the component, system, or arrangement supplied as a part of a certificated engine does meet the Part 29 standard, the possibility that subsequent changes to these components, systems, or arrangements by the engine manufacturer could negate compliance with Part 29 must be considered. For example, an engine may initially be equipped by the engine manufacturer with an oil tank filler cap that meets the Category A requirements of § 29.1013(c)(2) but is subsequently changed to a simpler and less expensive cap complying with § 33.71(c)(4). Continued monitoring of the engine configuration by the rotorcraft certification team would be needed to preclude an occurrence of noncompliance.

b. Procedures .

(1) Category A; Engine Isolation . This rule is one of the most significant safety rules in Subpart E of Part 29. Compliance involves a very extensive and rigorous evaluation not only of essentially all systems of the rotorcraft, but of the controls, both flight and powerplant, instruments, cockpit arrangement, cockpit switches, and operating procedures. A complete failure modes and effects analysis is involved.

Section 29.903(b)(1) should be rigorously applied to rotorcraft engine control arrangements which utilize governors responding to main rotor speed to modulate power rather than power levers preset to produce equal or less than limit power.

Section 29.903(b)(2) precludes “immediate action by any crewmember for continued safe operation.” This should be interpreted as requiring all powerplant systems to operate safely and continuously without crew attention (except to maintain flight using primary flight controls) in event of an engine failure from any cause, including fire. The collective is considered a primary flight control and not a powerplant control even though collective movement affects engine operation. No adjustment to powerplant controls or configuration can be allowed for certification purposes for performance credit or for safety. The time increment associated with “immediate” action may vary among different designs; however, it must not be less than that required to established engine-out flight profiles and climb rates associated with Category A performance.

During critical takeoff flight regimes, flight translation to at least published takeoff safety speed is needed before crew attention can be mandated to modulate powerplant controls or change aircraft configuration (i.e., landing gear, power lever or rotorspeed governor setting, etc.) to achieve published flight performance. This does not mean crew action is prohibited--only that no credit for crew action can be allowed for any resulting improved performance in the performance section of the flight manual.

(2) Category A; Control of Engine Rotation .

Page E - 9 AC 29-2C 9/30/99 (i) Means for stopping any engine in flight is to be considered unless it is shown that after critical failure of the engine, or components/accessories driven by the engine (not including rotor drive system components), no hazard results from rotation during the coast-down period. If continued rotation occurs, no hazard should result due to rotation during the period that the rotation is expected to continue. (Consider unbalanced rotors, bearing failures, accessory failures, lack of lubrication to other engine rotors, etc.) Note that after emergency engine shutdown, coast-down and continued rotation speed can be influenced by ram air flow into the compressor and, for multiengine rotorcraft, drag through the freewheeling unit.

(ii) A requirement exists for Category A rotorcraft to incorporate a means for restarting any engine individually in flight. Compliance is usually obtained during official flight tests and/or applicant tests in accordance with an approved test plan by requiring actual engine air-start demonstrations to define an acceptable restart envelope. These air-starts should be conducted at various altitudes, ambient temperatures, and fuel temperatures using the fuel type most critical, unless the applicant can show that this parameter is not pertinent. Other concerns involve the pilot station arrangement for flight controls and engine starting controls; i.e., verify that the engine start can be accomplished without jeopardizing continued safe operation of the rotorcraft, considering the pilot workload for the preexisting one-engine-inoperative situation, the location of the restart system controls, availability of a second pilot, etc.

Also, verify that the emergency/malfunction instruction sections of the RFM present a detailed definition of the approved restart envelope and detailed instructions for the restart, including eligible ambient atmospheric conditions, prestart arrangement of fuel, electrical and pneumatic systems (as applicable), delay time between start attempts (to allow for waste fuel drainage), starter duty cycle (if different from ground start duty cycle), and prestart situation analysis (i.e., Should a restart be attempted in view of the cause for initial shutdown? Is inlet system ice ingestion a possibility? Is reignition of fuel in the engine nacelle a possibility? Is sufficient restart time available? Is power available and is altitude sufficient to maintain terrain clearance?). Although restart capability from an all-engines-out flight condition is not required, special instructions for restarting from this situation should also be included commensurate with the system capability to accomplish the starts.

(3) Although restart capability is required for only Category A rotorcraft, the applicant should be encouraged to provide air start instructions in accordance with the above criteria for both single and multiengine Category B rotorcraft, including all-engine-out instructions if reasonable and practicable.

c. Turbine Engine Installation .

(1) Explanation . The certification of turbine engines and particularly the qualification of turbine rotors assume that the limitations established during these certifications will be accurately and rigorously observed during ground and flight operations in an aircraft. This paragraph is intended to promote this concept.

Page E - 10 5/1/2014 AC 29-2C, Chg 4 (2) Procedures . Primary engine limitations in the form of time, gas temperature, torque, and rotational speed and their corresponding allowable transient values are defined in the approved engine installation manual. The rotorcraft manufacturer must provide reliable, accurate means to assure that these limitations are not exceeded. These means may be in the form of automatic limiters or by crew monitoring of appropriately marked instruments. The FAA/AUTHORITY powerplant certification engineer and the rotorcraft manufacturer’s staff should verify these aspects by: (i) Evaluating all applicable instrument, indicator, or warning devices, including transmitters, and limiting devices, if any, for system tolerances.

(ii) Closely reviewing the component qualification reports of items in c(2)(i) above to verify that these devices are properly qualified and that any deviations are acceptable.

(iii) Assuring that maintenance data are provided for functional checks and calibration of instruments and devices which are used to monitor or protect critical turbine rotor limitations. Preflight checks for automatic limiter devices may be appropriate.

(iv) Verifying that instrument markings are clear and relatively simple, that corresponding flight manual instructions and descriptions are straightforward and complete, and that instruments are located and orientated to minimize the probability of reading error.

AC 29.903A. §29.903 (Amendment 29-26) ENGINES.

a. Explanation . Amendment 29-26 adds § 29.903(a) that requires reciprocating engines used in rotorcraft to be certified in accordance with the rotorcraft engine testing requirements in § 33.49(d). This change is incorporated to ensure that certification requirements are not overlooked when reciprocating engines are installed in rotorcraft to be certified under Part 29 requirements. Section 29.903(b)(2) was revised to identify and clarify crew action; i.e., normal pilot action allowable with primary flight controls, in determining if adequate powerplant systems isolation is provided. This change 0eliminates any possible confusion that may exist regarding the acceptability of modifying optimum flight control manipulation to protect engine parameters.

Section 29.903(c)(3) was added and requires engine restart capability to be available throughout the flight envelope appropriate to the rotorcraft. This will avoid the concept that an in-flight engine restart envelope constitutes acceptable compliance with this rule.

b. Procedures .

(1) Engine type certification . All engines installed in rotorcraft should have a type certificate. The specific certification requirements for installation of reciprocating engines in rotorcraft are found in Part 33. Engines certificated under other approved Page E - 11 AC 29-2C, Chg 4 5/1/2014 certification rules (CAR Part 13 and FAR § 21.29, for imported engines) are also eligible. If a component, system, or arrangement is certified under Part 33 or other requirement, the applicant is not relieved of the necessity to comply with the requirements of Part 29. If the component, system, or arrangement, supplied as a part of a certificated engine, meets the Part 33 and Part 29 requirements, subsequent changes to these components, systems, or arrangements could negate compliance with Part 29. For example, an engine may initially be equipped by the engine manufacturer with an oil tank filler cap that complies with the Category A requirements of § 29.1013(c)(2) but is subsequently changed to a simpler and less expensive cap that complies with § 33.71(c)(4). The airframe manufacturer should ensure that the requirements of § 29.1013(c)(2) are maintained.

(2) Category A: control of engine rotation . Section 29.903(c)(3) requires an engine restart capability which is appropriate to the rotorcraft. The minimum envelope for the restart capability should be equal to or better than the rotorcraft takeoff/landing maximum altitude and temperature limits. Compliance is usually shown by conducting actual in-flight restarts during flight tests and/or other tests in accordance with an approved test plan. Restarts should be conducted at various altitudes, ambient temperatures, and fuel temperatures using the fuel type most critical, unless the applicant can show that this parameter is not pertinent. Other concerns involve the pilot station arrangement for flight controls and engine starting controls. It should be verified that the engine start can be accomplished without jeopardizing continued safe operation of the rotorcraft. Pilot workload for a preexisting one-engine-inoperative situation, the location of the restart system controls, and the availability of a second pilot should be considered. The emergency/malfunction instruction sections of the rotorcraft flight manual (RFM) should present a detailed definition of the approved restart envelope and detailed instructions for the restart. Eligible ambient atmospheric conditions, prestart requirements (to allow for waste fuel drainage), starter duty cycle (if different from the ground start duty cycle), and prestart situation analysis should be included. The prestart situation analysis should consider the following questions:  Should a restart be attempted in view of the cause for initial shutdown?

 Is inlet system ice ingestion a possibility?

 Is reignition of fuel in the engine nacelle a possibility?

 Is sufficient restart time available?

 Is power available?

 Is altitude sufficient to maintain terrain clearance?

Although restart capability from an all-engines-out flight condition is not required, special instructions for restarting from this situation should also be included commensurate with the system capability to accomplish the starts.

AC 29.903B. § 29.903 (Amendment 29-31) ENGINES.

a. Explanation . Amendment 29-31 clarified the requirements for control of engine rotation and in-flight restart of engines. Section 29.903(c)(1) was changed by adding Page E - 12 5/1/2014 AC 29-2C, Chg 4 the word “or” at the end of the paragraph, which provided an option on how to protect the engine stopping system from fire.

b. Procedures . All of the policy material pertaining to this section remains in effect.

Additionally, the new § 29.903(e) requires that any engine should have a restart capability that has been demonstrated throughout a flight envelope to be certificated for the rotorcraft. However, § 29.903(e)(2) does not require in-flight demonstration of restart capability for single-engine rotorcraft or for all-engine shutdown of multi-engine rotorcraft. In the past, engine relight capability for single engine rotorcraft has been demonstrated on the ground taking into account altitude effects, warm engine characteristics, depleted battery, etc. The minimum restart envelope for category A rotorcraft is discussed in section 29.903A of this AC. The restart capability can consider windmilling of the engine as part of this restart capability; however, most rotorcraft airspeeds and the locations of the engines do not support engine windmilling up to start speeds. Only electrical power requirements were considered for restarting; however, other factors that may affect this capability are permitted to be considered. Engine restart capability following an in-flight shutdown of the engine in single-engine rotorcraft, or all engines in a multi-engine rotorcraft, is the primary requirement, and the means of providing this capability is left to the applicant. To minimize any potential height loss following the failure of one or more engines, engine restart should be available at the earliest opportunity. The engine certification should be checked to ensure that the flight manual instructions for in-flight restart are consistent with any specific engine restart requirements.

AC 29.903C. § 29.903 (Amendment 29-36) TURBINE ENGINE INSTALLATION.

a. Explanation . Amendment 29-36 revises § 29.903(d) to require that design precautions should be taken to minimize hazards to the rotorcraft in the event of an engine failure. This advisory material sets forth a method of compliance with the requirements of §§ 29.901, 29.903(b)(1), and 29.903(d)(1) of the Federal Aviation Regulations (FAR) pertaining to design precautions taken to minimize the hazards to rotorcraft in the event of uncontained engine rotor (compressor and turbine) failure. It is for guidance and to provide a method of compliance that has been found acceptable.

As with all AC material, it is not mandatory and does not constitute a regulation.

b. Procedures. Although turbine engine manufacturers are making efforts to reduce the probability of uncontained rotor failures, service experience shows that such failures continue to occur. Failures have resulted in high velocity fragment penetration of fuel tanks, adjacent structures, fuselage, system components and other engines of the rotorcraft. Since it is unlikely that uncontained rotor failures can be completely eliminated, rotorcraft design precautions should be taken to minimize the hazard from such events. These design precautions should recognize rotorcraft design features that may differ significantly from that of an airplane, particularly regarding an engine location and its proximity to another engine or to other systems and components.

Page E - 13 AC 29-2C, Chg 4 5/1/2014 (1) Uncontained gas turbine engine rotor failure statistics for rotorcraft are presented in the Society of Automotive Engineers (SAE) Report No.’s AIR 4003 (period 1976-83) and AIR 4770 (period 1984-89).

(2) The statistics in the SAE studies indicate the existence of some failure modes not readily apparent or predictable by failure analysis methods. Because of the variety of uncontained rotor failures, it is difficult to analyze all possible failure modes and to provide protection to all areas. However, design considerations outlined in this AC provide guidelines for achieving the desired objective of minimizing the hazard to rotorcraft from uncontained rotor failures. These guidelines, therefore, assume a rotor failure will occur and that analysis of the effects or evaluation of this failure is necessary.

These guidelines are based on service experience and tests but are not necessarily the only means available to the designer.

c. Definitions .

(1) Minimize . Reduce to the least possible amount by means that can be shown to be both technically feasible and economically justifiable.

(2) Separation . Positioning of redundant critical structure, systems, or system components within the impact area such that the distance between the components minimizes the potential impact hazard. Redundant critical components should be separated within the spread angles of a rotor by a distance at least equal to either a ½ unbladed disk (hub, impeller) sector, or a 1/3 bladed disk (hub, impeller) sector with 1/3 blade height, with each rotating about its center of gravity (CG), whichever is greater (see figure AC 29.903C-6).

(3) Isolation . A means to limit system damage so as to maintain partial or full system function after the system has been damaged by fragments. Limiting the loss of hydraulic fluid by the use of check valves to retain the capability to operate flight controls is an example of “isolation.” System damage is confined allowing the retention of critical system functions.

(4) Rotor .

(i) Rotor means the rotating components of the engine and APU that analysis, test results, and/or experience has shown can be released during uncontained failure with sufficient energy to hazard the rotorcraft.

(ii) The engine or APU manufacturer should define those components that constitute the rotor for each engine and APU type design. Typical rotors have included, as a minimum, disks, hubs, drums, seals, impellers, and spacers.

(5) Uncontained Engine or APU Failure (or Rotorburst) . For the purposes of rotorcraft evaluations in accordance with this AC, uncontained failure of a turbine engine is any failure which results in the escape of rotor fragments from the engine or Page E - 14 9/30/99 AC 29-2C could create a hazard to the rotorcraft. Rotor failures of concern are those in which released fragments have sufficient energy to create a hazard to the rotorcraft.

Uncontained failures of APU’s which are “ground operable only” are not considered hazardous to the rotorcraft.

(6) Critical Component (System) . A critical component is any component or system whose failure or malfunction would contribute to or cause a failure condition that would prevent the continued safe flight and landing of the rotorcraft. These components (systems) should be considered on an individual basis and in relation to other components (systems) that could be degraded or rendered inoperative by the same fragment or by other fragments during any uncontained failure event.

(7) Fragment Spread Angle . The fragment spread angle is the angle measured, fore and aft, from the center of the plane of rotation of the disk (hub, impeller) or other rotor component initiating at the engine or APU shaft centerline or axis of rotation (see figure AC 29.903C-1). The width of the fragment should be considered in defining the path of the fragment envelope’s maximum dimension.

(8) Ignition Source . Any component that could precipitate a fire or explosion.

This includes existing ignition sources and potential ignition sources due to damage or fault from an uncontained rotor failure. Potential ignition sources include hot fragments, damage or faults that produce sparking, arcing, or overheating above the auto-ignition temperature of the fuel. Existing ignition sources include items such as unprotected engine or APU surfaces with temperature greater than the auto-ignition temperature of the fuel or any other flammable fluid.

d. Safety Assessment .

(1) Procedure . Assess the potential hazard to the rotorcraft using the following procedure: (i) Minimizing Rotorburst Hazard . The rotorburst hazard should be reduced to the lowest level that can be shown to be both technically feasible and economically justifiable. The extent of minimization that is possible will vary from new or amended certification projects and from design to design. Thus the effort to minimize must be determined uniquely for each certification project. Design precautions and techniques such as location, separation, isolation, redundancy, shielding, containment and/or other appropriate considerations should be employed, documented, agreed to by the certifying authority, and placed in the type data file. A discussion of these methods and techniques follows.

(ii) Geometric Layout and Safety Analysis . The applicant should prepare a preliminary geometric layout and safety analysis for a minimum rotorburst hazard configuration determination early in the design process and present the results to the certification authority no later than when the initial design is complete. Early contact and coordination with the certifying authority will minimize the need for design Page E - 15 AC 29-2C 9/30/99 modification later in the certification process. The hazard analysis should follow the guidelines indicated in paragraphs AC 29.901c(2) and AC 29.903Cd(6). Geometric layouts and analysis should be used to evaluate and identify engine rotorburst hazards to critical systems, powerplants, and structural components from uncontained rotor fragments, and to determine any actions which may be necessary to further minimize the hazard. Calculated geometric risk quantities may be used in accordance with paragraph d(4) following, to define the rotorcraft configuration with the minimum physical rotorburst hazard.

(2) Engine and APU Failure Model . The safety analysis should be made using the following engine and APU failure model, unless for the particular engine/APU type concerned, relevant service experience, design data, test results or other evidence justify the use of a different model. In particular, a suitable failure model may be provided by the engine/APU manufacturer. This may show that one or more of the considerations below do not need to be addressed.

(i) Single One-Third Disc Fragment . It should be assumed that the one-third disc fragment has the maximum dimension corresponding to one-third of the disc with one-third blade height and a fragment spread angle of ±3°. Where energy considerations are relevant, the mass should be assumed to be one-third of the bladed disc mass and its energy--the translational energy (i.e., neglecting rotational energy) of the sector (see figure AC 29.903C-2).

(ii) Intermediate Fragments. It should be assumed that the intermediate fragment has a maximum dimension corresponding to one third or the disc radius with one-third blade height and a fragment spread angle of ±5°. Where energy considerations are relevant, the mass should be assumed to be 1/30th of the bladed disc mass and its energy--the translational energy (neglecting rotational energy) of the piece traveling at rim speed (see figure AC 29.903C-3).

(iii) Alternative Engine Failure Model . For the purpose of the analysis, as an alternative to the engine failure model of paragraphs d(2)(i) and d(2)(ii) above, the use of a single one-third piece of disc having a fragment spread angle of ±5° would be acceptable, provided that the objectives of the analysis are satisfied.

(iv) Small Fragments . It should be assumed that small fragments have a maximum dimension corresponding to the tip half of the blade airfoil and a fragment spread angle of ±15°. Where energy considerations are relevant, the mass should be assumed to be corresponding to the above fragment dimensions and the energy is the translational energy (neglecting rotational energy) of the fragment traveling at the speed of its CG location. The effects of multiple small fragments should be considered during this assessment.

(v) Critical Engine Speed . Where energy considerations are relevant, the uncontained rotor event should be assumed to occur at the engine shaft speed for the maximum rating appropriate to the flight phase (exclusive of OEI ratings), unless the Page E - 16 9/30/99 AC 29-2C most probable mode of failure would be expected to result in the engine rotor reaching a red line speed or a design burst speed. For APU’s, use the maximum rating appropriate to the flight phase or the speed resulting from a failure of any one of the normal engine control systems.

(vi) APU Failure Model . Service experience has shown that some APU rotor failures produced fragments having significant energy to have been expelled through the APU tailpipe. For the analysis, the applicable APU service history and test results should be considered in addition to the failure model as discussed in paragraph d(2) above for certification of APU installations near critical items. In addition, the APU installer needs to address the rotorcraft hazard associated with APU debris exiting the tailpipe. Applicable service history or test results provided by the APU manufacturer may be used to define the tailpipe debris size, mass, and energy. The uncontained APU rotor failure model is dependent upon the design/analysis, test results and service experience.

(A) For APU’s in which rotor integrity and blade containment have been demonstrated in accordance with TSO-C77a/JAR APU, i.e., without specific containment testing, paragraphs d(2)(i), d(2)(ii), and d(2)(iv) or paragraph d(2)(iii) and d(2)(iv) apply. If shielding of critical airframe components is proposed, the energy level that should be considered is that of the tri-hub failure released at the critical speed as defined in paragraph d(2)(v). The shield and airframe mounting point(s) should be shown to be effective at containing both primary and secondary debris at angles specified by the failure model.

(B) For APU rotor stages qualified as contained in accordance with the TSO, an objective review of the APU location should be made to ensure the hazard is minimized in the event of an uncontained APU rotor failure. Historical data shows that in-service uncontained failures have occurred on APU rotor stages qualified as contained per the TSO. These failure modes have included bi-hub and overspeed failure resulting in some fragments missing the containment ring. In order to address these hazards, the installer should use the small fragment failure model, or substantiated in-service data supplied by the APU manufacturer. Analytical substantiation for the shielding system if proposed is acceptable for showing compliance.

(3) Engine/APU Rotorburst Data . The engine or APU manufacturer should provide the required engine data to accomplish the evaluation and analysis necessary to minimize the rotorburst hazard such as: (i) Engine failure model (range of fragment sizes, spread angles and energy).

(ii) Engine rotorburst probability assessment.

(iii) List of components constituting the rotors.

Page E - 17 AC 29-2C 9/30/99 (4) Fragment Impact Risks . FAA/AUTHORITY research and development studies have shown that, for rotorcraft conventional configurations (one main rotor and one tail rotor), the main and tail rotorblades have minimal risks from a rotorburst, and thus, they require no special protection. However, unique main and tail rotor blade configurations should be carefully reviewed. Certain zones of the tail rotor drive shaft and other critical parts which may be necessary for continued safe flight and landing may not have natural, minimal risk from uncontained rotor fragments.

(5) Engine Service History/Design .

(i) For the purpose of a gross assessment of the vulnerability of the rotorcraft to an uncontained rotorburst, it must be taken that an uncontained engine rotor failure (burst) will occur. However, in determining the overall risk to the rotorcraft, engine service history and engine design features should be included in showing compliance with § 29.903 to minimize the hazard from uncontained rotor failures. This is extremely important since the engine design and/or the service history may provide valuable information in assessing the potential for a rotorburst occurring and this should be considered in the overall safety analysis.

(ii) Information contained in the recent SAE studies should be considered in this evaluation (see paragraph b(1) above).

(6) Certification Data File . A report, including all geometric layouts, that details all the aspects of minimizing the engine rotorburst hazards to the rotorcraft should be prepared by the applicant and submitted to the certification authority. Items which should be included in this report are the identification of all hazardous failures that could result from engine rotor failure strikes and their consequences (i.e., an FMEA or equivalent analysis) and the design precautions and features taken to minimize the identified hazards that could result from rotor failure fragment strikes. Thus an analysis that lists all the critical components; quantifies and ranks their associated rotorburst hazard; and clearly shows the minimization of that quantified, ranked hazard to the “maximum practicable extent” should be generated and agreed upon during certification. Critical components should all be identified and their rotorburst hazard quantified, ranked, and minimized where necessary. Design features in which the design precautions of this guidance material are not accomplished should be identified along with the alternate means used to minimize the hazard. To adequately address minimizing the hazards, all rotorcraft design disciplines should be involved in the applicant’s compliance efforts and report preparation.

e. Design Considerations. Practical design precautions should be used to minimize the damage that can be caused by uncontained engine and APU rotor debris.

The following design considerations are recommended: (1) Consider the location of the engine and APU rotors relative to critical components, or areas of the rotorcraft such as: Page E - 18 9/30/99 AC 29-2C (i) Opposite Engine - Protection of the opposite engine from damage from 1/3 disc rotor fragments may not be feasible. Protection of the opposite engine from other fragments may be provided by locating critical components, such as engine accessories essential for proper engine operation (e.g. high pressure fuel lines, engine controls and wiring, etc.), in areas where inherent shielding is provided by the fuselage, engine, or other structure.

(ii) Engine Controls - Controls for the remaining engine(s) that pass through the uncontained engine failure zone should be separated/protected to the maximum extent practicable.

(iii) Primary Structure of the Fuselage .

(iv) Flight Crew - The flight crew is considered a critical component.

(v) Fuel system components, piping and tanks, including fuel tank access panels (NOTE: Spilled fuel into the engine or APU compartments, on engine cases or on other critical components or areas could create a fire hazard.)

(vi) Critical control systems, such as primary and secondary flight controls, electrical power cables, systems and wiring, hydraulic systems, engines control systems, flammable fluid shut-off valves, and the associated actuation wiring or cables.

(vii) Engine and APU fire extinguisher systems including electrical wiring and fire extinguishing agent plumbing to engine and APU compartments.

(viii) Instrumentation necessary for continued safe flight and landing.

(ix) Transmission and rotor drive shafts.

(2) Location of Critical Systems and Components . The following design practices have been used to minimize hazards to critical components: (i) Locate, if possible, critical components or systems outside the likely debris impact areas.

(ii) Duplicate and separate critical components or systems if located in debris impact areas or provide suitable protection.

(iii) Protection of critical systems and components can be provided by using airframe structure where shown to be suitable.

Page E - 19 AC 29-2C 9/30/99 (iv) Locate fluid shutoffs so that flammable fluids can be isolated in the event of damage to the system. Design and locate the shut-off actuation means in protected areas or outside debris impact areas.

(v) Minimize the flammable fluid spillage which could contact an ignition source.

(vi) For airframe structural elements, provide redundant designs or crack stoppers to limit the subsequent tearing which could be caused by uncontained rotor fragments.

(vii) Consider the likely damage caused by multiple fragments.

(viii) Fuel tanks should not be located in impact areas. However, if necessitated by the basic configuration requirements of the rotorcraft type to locate fuel tanks in impact areas, then the engine rotorburst hazard should be minimized by use of design features such as minimization of hazardous fuel spillage (that could contact an ignition source by drainage or migration); by drainage of leaked fuel quickly and safely into the airstream; by proper ventilation of potential spillage areas; by use of shielding; by use of explosion suppression devices (i.e., explosion resistant foam or inert gases); and by minimization of potential fuel ignition sources or by other methods to reduce the hazard.

(ix) The rotor integrity or containment capability demonstrated during APU evaluation to TSO-C77a, or JAR-APU should be considered for installation certification.

(x) The flight data recorder, cockpit voice recorder, and emergency locator transmitter, if required, should be located outside the impact zone when practical.

(xi) Items such as human factors, pilot reaction time, and correct critical system status indication in the pilot compartment after an uncontained engine failure has occurred should be considered in design to permit continued safe flight and landing.

(3) Rotorcraft Modifications . Modifications made to rotorcraft certified to this rule should be assessed with the considerations of this AC. These modifications include but are not limited to re-engining installations (including conversion from reciprocating to turbine powered), APU installations, fuselage stretch, and auxiliary fuel tank installations. Auxiliary fuel tank(s) should be located as much as practical so as to minimize the risk that this tank(s) will be hit by rotor failure fragments. The need to remain within the approved CG limits of the aircraft will of necessity limit the degree to which the risk may be minimized.

f. Protective Measures . The following list is provided for consideration as some measures which may be used to minimize effects of a rotorburst: Page E - 20 9/30/99 AC 29-2C (1) Powerplant Containment .

(i) Engine Rotor Fragment Containment . It should be clearly understood that containment of rotor fragments is not a requirement. However, it is one of many options which may be used to minimize the hazards of an engine rotorburst.

Containment structures (either around the engine, or APU, or on the rotorcraft) that have been demonstrated to provide containment should be accepted as minimizing the hazard defined by the rotor failure model for that particular rotor component. Contained rotor in-service failures may be used to augment any design or test data. Containment material stretch and geometric deformation should be considered in conjunction with fragment energies and trajectories in defining the hazards to adjacent critical components such as structures, system components, fluid lines, and control systems.

Data obtained during containment system testing along with analytical data and service experience should be used for this evaluation.

(ii) APU Containment . Rotor integrity or containment capability demonstrated during APU TSO evaluation should be considered for installation certification. If rotor containment option was shown by analysis or rig test an objective review of the APU location should be made to ensure the hazard is minimized in the event of an uncontained APU rotor failure.

(2) Shields and Deflectors . When shields, deflection devices, or intervening rotorcraft structure are used to protect critical systems or components, the adequacy of the protection should be shown by testing or analysis supported by test data, using the impact area, fragment mass, and fragment energies based on the definitions stated herein. Analytical methods used to compute protective armor or shielding thicknesses and energy absorption requirements should reflect established methods, acceptable to the certifying authority, that are supported by adequate test evidence. Protective armor, shielding, or deflectors that stop, slow down, or redirect uncontained fragments redistribute absorbed energy into the airframe. The resulting loads are significant for large fragments and should be considered as basic load cases for structural analysis purposes (reference § 29.301). These structural loads should be defined and approved as ultimate loads acting alone. The protective devices and their supporting airframe structures should be able to absorb or deflect the fragment energies defined herein and still continue safe flight and landing. If hazardous, the deflected fragment trajectories and residual energies should also be considered.

(3) Isolation or Redundancy .

(i) Other Engines - Although other engines may be considered critical, engine isolation from rotorburst on multi-engine rotorcraft is not mandatory. Other methods of minimizing the risk to the engine(s) may be acceptable.

(ii) Other Critical Components - Isolation or redundancy of other critical components, the failure of which would not allow continued safe flight and landing should be evaluated relative to the risk of occurrence and where the risk is deemed Page E - 21 AC 29-2C 9/30/99 unacceptable isolation or shielding or other means of reducing the risk should be incorporated.

(4) Composite Materials . If containment devices, shields, or deflectors are chosen by the applicant to be wholly or partially made from composites; they should comply with the structural requirements of AC 20-107A, “Composite Aircraft Structure,” and paragraph AC 29 MG 8, “Substantiation of Composite Rotorcraft Structure,” (which includes glass transition temperature considerations). Glass transition temperature considerations are critical for proper certification of composite or composite hybrid structures used in temperature zones that reach or exceed 200° to 250°F (93° to 121°C) for significant time periods. Hot fragment containment is typically accommodated in such protective devices by use of metal-composite hybrid designs that use the metal component’s properties to absorb the fragment heat load after the entire hybrid structure has absorbed the fragment’s impact load. These devices should comply with §§ 29.609 and 29.1529 to ensure continued airworthiness.

Page E - 22 9/30/99 AC 29-2C , 113RD AND LARGER BLADED MASS ~ (SEE FIGURE 2) ±3·--1 I

---~

±5· .._.....___ N.r. 1130 BLADED MASS l,;;t;, (INTERMEDIATE) (SEE FIGURE 3) ..____ SMALL FRAGMENTS (SEE PARAGRAPH 58(4 )) FRAGMENT SPREAD ANGLE IS THE ANGLE MEASURED FORE AND AFT, FROM THE CENTER OF THE PLANE OF ROTATION INITIATING AT THE ENGINE OR APU SHAFT CENTERLINE NOTE 1) THE POSSIBILITY OF TUR SINE MOVEMENT SHOULD BE CONSIDERED.

2) ALL ROTORS ARE CONSIDERED TO BE FULLY BLADED FOR CALCULATING MASS.

3) FAILURE OF EACH ROTOR STAGE SHOULD BE CONSIDERED.

FIGURE AC 29.903C-1 ESTIMATED PATH OF FRAGMENTS Page E - 23 AC 29-2C 9/30/99 MAXIMUM DIMENSION

--

-

/ 1/3 b

~/ ~ ✓ ~-------El:l-CG ____ ~r'\

Ir Where R = disc radius b = blade length The CG is taken to Ii e on the maximum dimension as sh own.

FIGURE AC 29.903C-2. SINGLE ONE-THIRD DISC FRAGMENT Page E - 24 9/30/99 AC 29-2C ,,_---....

/ ., ,/ 11/3 (R • bl

/

'

/

'

I

I

\

~

\

b I

Where R = disc radius b = blade length Maximum dimension= 1/3 (R + b) Mass assumed to be 1/ 30th of bladed disc CG is taken to lie on the disc rim FIGURE AC 29.903C-3. INTERMEDIATE AND SMALL PIECES OF DEBRIS Page E - 25 AC 29-2C 9/30/99 BE AT TO FRAGMENT THE TAKEN PLATFORM) 1/3 AT LENGTH & MASS ASSUMED LIE THE ROOT OF TO CG VELOCITY MASS AIRFOIL AIRFOIL X = THE BLADE TAKEN IS OF (LESS CG CENTERLINE FRAGMENT GEOMETRIC FRAGMENT 1/3 WHERE CG DEFINITION X GEOMETRIC FRAGMENT - BLADE

T

J_

1/3X ' Tl ► t-..) ~ (j) C ;:o m () <D <D 0 (>) () Page E - 26 9/30/99 AC 29-2C ROTATIONAL KE=½ lo, TRANS LAT 10 NAL KE= ½mv i=' z w a: w ~ >- (!)

50 a: w z w ~ w z S:'.

240 L 300 360

0 120 180 ROTATIONAL TRANSLATIONAL FRAGMENT SIZE 6' FIGURE AC 29.903C-5 DISTRIBUTION OF TRANSLATIONAL AND ROTATIONAL KINETIC ENERGY OF ROTOR-COMPONENT FRAGMENTS AS A FUNCTION OF FRAGMENT SIZES Page E - 27 AC 29-2C 9/30/99 CG of Fragment Becomes Cen ter of Rot ation of Fragment

For Separati on Dist an ce Calculati on s: a

1/3 Rot or w ith 1/ 3 Blade Heig ht FIGURE 29.903C-6 CR OSS SECTION THROUGH AI RCRAFT AT PLANE OF ROTATION OF THE ENGI NE a SK FRAG MENT Page E - 28 2/12/03 AC 29-2C, Chg 1 AC 29.907. § 29.907 ENGINE VIBRATION .

a. Explanation . This very generalized requirement is authority to require substantiation of the effects of vibration on any part of the engine or the rotorcraft. In normal certification practice, the vibration effects of concern to the powerplant engineer are the vibratory loads or stresses in the engine and in the rotor drive system. Vibration effects on the rotor drive system are of concern if the corresponding loads or stresses result in fatigue damage. This aspect, however, is adequately addressed in § 29.571.

Vibration effects on the engine are usually categorized as “installation vibration” and “torsional vibration.” Methods of evaluation and limitations of these vibrations are established by the engine manufacturer.

b. Procedures . Review Order 8110.9, Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopter and other Power Transmission Systems. Note that the mechanical coupling of the engines to the rotor drive system creates, for torsional vibration considerations, one, rather complicated, drive system which responds to any forced or resonant frequency. Antinodes or nodes and frequencies may exist in the engine shaft which are absent when the engine is operated on a test stand; therefore, the vibration investigation conducted under Part 33 is not conclusive with respect to torsionals. As noted in Order 8110.9, the engine manufacturers’ assistance is necessary to find compliance. Section 29.571 was amended by Amendment 29-13 to include “rotor drive systems between the engines and the rotor hubs” as part of the flight structure. This rule supplements § 29.907 and requires coordination with the structures certification engineer to avoid duplication of effort by the rotorcraft manufacturer.

Advisory Circular 27 MG-11, Fatigue Evaluation of Rotorcraft Structure, which provides acceptable methods of compliance with § 29.571, may also be used to find compliance with § 29.907. In addition to basic drive system components such as main and auxiliary rotor drive shafts, the vibratory evaluation should include couplings, gear teeth, gear cases and splines, and should consider, where appropriate, low cycle fatigue associated with ground-air-ground cycles.

AC 29.908. § 29.908 (Amendment 29-13) COOLING FANS .

a. Section 29.908(a) : (1) Explanation . This paragraph applies to Category A rotorcraft and is intended to require that powerplant area cooling fans be designed and installed to enable continued safe operation of the rotorcraft after failure of a cooling fan blade. The phrase “except that the loss of cooling need not be considered” at the end of this paragraph is intended to make clear that for the purposes of this section, the FAA/AUTHORITY is concerned only with the fragmentation effect of a fan blade failure (reference Preamble Item 3-64 of Amendment 29-12).

(2) Procedures . If a fan shroud is provided, the applicant may demonstrate that the shroud configuration and strength are adequate to contain a failed fan blade and any other fan blades, guide vanes, etc., which can be expected to fail sequentially to the Page E - 29 AC 29-2C 9/30/99 initial blade failure. The demonstration can be facilitated by making a saw slot at the root of a blade sufficiently deep to weaken the blade retention strength and create a failure while the fan is rotating at the maximum speed established for the test. If the fan is driven by the rotor drive system, the test speed should be equal to or above the maximum transient speed to be expected with the rotor system. If the fan is driven by other means; i.e., bleed air turbine, hydraulic motor, engine N turbine, etc., the rotational speed for the blade failure demonstration should be based on a critical analysis of speed regimes to be expected. Containment is not required if the fan is located so that blade failure (and any sequential fan component failure) will not jeopardize safety. This may be shown by test or analysis. Segment shielding would likely be involved.

b. Section 29.908(b) : (1) Explanation . This paragraph applies to Category B rotorcraft and is intended to provide safety to the rotorcraft in the event of an assumed cooling fan blade failure or to prescribe a test to show that the cooling fan blade retention means is sufficient that blade failure is not a consideration.

(2) Procedures .

(i) The applicant may select § 29.908(b)(1), (b)(2), or (b)(3) to show compliance with this section. If § 29.908(b)(1) is selected, follow the procedures outlined above for Category A rotorcraft.

(ii) Section 29.909(b)(2) may be selected; however, without containment, damage to any component or structure in the plane of the fan rotor or any other trajectory to be expected should not cause the loss of any function essential to a controlled landing.

(iii) If § 29.908(b)(3) is selected, a spin test at 122.5 percent of the maximum speed associated with either engine terminal speed or an overspeed limiting device would be acceptable to show compliance. No failure should occur, and distortion should not result in fan element contact with housings or other adjacent components.

(Note: 150 percent of the centrifugal force is achieved at 122.5 percent of the rotational speed.)

AC 29.908A. § 29.908 (Amendment 29-26) COOLING FANS .

a. Explanation . Amendment 29-26 requires that cooling fans be designed and installed to enable continued safe flight and adequate cooling of the rotorcraft following a fan blade failure. Compliance with the previous requirements could have resulted in hazards to the rotorcraft with the loss of cooling air to critical powerplant components. A new section was also added to the rule for cooling fans, which are not part of the powerplant installation and therefore not subject to the fatigue evaluation under Page E - 30 9/30/99 AC 29-2C § 29.571. It should be determined that no cooling fan blade resonant conditions exist within the operating limits of the rotorcraft unless a fatigue evaluation is conducted.

b. Procedures . Neither mechanical damage nor loss of cooling air should prevent “continued safe flight.” The definition of “continued safe flight” is contained in Appendix 1 of AC 20-136 and is quoted as follows: Continued safe flight and landing . This phrase means that the aircraft is capable of safely aborting or continuing a takeoff; continuing controlled flight and landing, possibly using emergency procedures but without requiring exceptional pilot skill or strength.

Some aircraft damage may occur as a result of the failure condition or upon landing.

For airplanes, the safe landing must be accomplished at a suitable airport. For rotorcraft, this means maintaining the ability of the rotorcraft to cope with adverse operating conditions and to land safely at a suitable site.

The FAA/AUTHORITY has determined that for Category A rotorcraft the phrase, “continued safe flight” means that the rotorcraft retains the capability to return and land safely at the point of departure or continue and land safely at the original intended destination or a suitable alternate site.

(1) This section is intended to ensure that a cooling fan blade failure will not jeopardize safety of the rotorcraft. Three ways to show compliance with this section are as follows: (i) A demonstration should be conducted to show that at the maximum fan speed to be expected, a failed blade will be contained within a housing or shroud which is included in the proposed type design and is designated as the containment shield; (ii) It should be shown that the installed cooling fan is located such that a blade failure will not jeopardize the safety of the rotorcraft or its ability to continue safe flight (Category A) or land safely (Category B); or, (iii) It should be shown that the cooling fan blades can withstand an ultimate load 1.5 times the maximum centrifugal force that may be expected in service.

The maximum centrifugal forces will occur at the maximum cooling fan rotational speeds. The maximum fan rotational speeds may be related to an overspeed limiting device or to the maximum transient speed to be expected from analysis or test of the engine, system, or component which drives the fan. The maximum rotational speed will be as follows: (A) For fans driven directly by the engine: ( 1 ) The terminal engine rotational speed that will occur under uncontrolled conditions; such as output shaft disconnect; or Page E - 31 AC 29-2C 9/30/99 ( 2 ) The maximum engine rotational speed that would be controlled by a reliable, approved engine overspeed limiting device.

(B) For fans driven by the rotor drive system, the maximum rotor drive system rotational speed to be expected in service including transients. (Note: Capability to withstand the ultimate load of 1.5 times the centrifugal force means that no failure would occur and distortion should not result in fan element contact with housings or other adjacent components during the 122.5 percent spin test which equates to 150 percent centrifugal force.)

(2) Fatigue . If the cooling fan is not included in the fatigue evaluation under § 29.571, it should be shown that the cooling fan blades are not operating at resonant conditions within the normal operating limits of the rotorcraft.

Page E - 32 9/30/99 AC 29-2C SUBPART E - POWERPLANT ROTOR DRIVE SYSTEM AC 29.917. § 29.917 (Amendment 29-12) DESIGN .

a. Section 29.917(a) General : (1) Explanation . This paragraph sets forth a definition of the rotor drive system and its associated components. The intent of this paragraph is to clarify and/or establish the identification of components to be considered in other rules which are applicable to the rotor drive system.

(2) Procedures . Coordinate with other certification personnel to ensure that other rules pertaining to rotor drive systems are properly addressed.

b. Section 29.917(b) Arrangement : (1) Explanation .

(i) Section 29.917(b)(1) pertains to multiengine rotorcraft and requires the drive system arrangement to be such that the rotors will continue to be driven by the remaining engines in order to ensure that lift and control to be expected from the rotors are available if an engine fails.

(ii) Section 29.917(b)(2) pertains to single-engine rotorcraft and is similar to the requirement of paragraph AC 29.917b(1)(i) except that it requires each rotor necessary for operation and control to be driven by the main rotor(s) after disengagement of the engine from the main and auxiliary rotors.

(iii) Section 29.917(b)(3) is intended to require a design which allows the rotor system to be protected from the torsional drag of an inoperative engine.

(iv) Section 29.917(b)(4) pertains to optional torque limiting means (shear sections or clutches) and prohibits these devices from being located in the cross-shafting system between rotors.

(v) Section 29.917(b)(5) is intended to ensure that the design prevents rotors from contacting each other if intermeshing is possible.

(vi) Section 29.917(b)(6) is intended to ensure that locking devices are installed to keep rotors in proper phase if dephasing is possible.

(2) Procedures .

Page E - 33 AC 29-2C 9/30/99 (i) Section 29.917(b)(1) is normally complied with by cross-shafting between rotors, usually via one or more transmissions or gear boxes, to optimize the mechanical simplicity and weight aspects. Individual engine input arrangements are required.

(ii) Section 29.917(b)(2) may be complied with by cross-shafting between rotors. Usually this involves driving the antitorque rotor via a drive shaft from the main transmission.

(iii) Section 29.917(b)(3) may be complied with by installing “free-wheel” or “one-way” clutches in the engine output shaft or transmission input quill. Note that the output section of “free power turbine” engines is not an acceptable method of compliance.

(iv) Section 29.917(b)(4). Any torque limiting devices in the rotor system should be located in the engine output or transmission input quill to ensure that any disconnect from overtorque does not preclude continued normal function and relation of the rotors.

(v) Section 29.917(b)(5). Phase control of intermeshing rotors should utilize positive mechanical drive components. Deflections in both shafting (torsional) and rotors (blade chordwise bending) should be considered in establishing compliance.

(vi) Section 29.917(b)(6). Reconnection of dephased rotors should employ positive mechanical locking pins with secure locking methods.

AC 29.917A. § 29.917 (Amendment 29-40) DESIGN .

a. Explanation . Amendment 29-40 introduces a new § 29.917(b). The previous § 29.917(b) has been redesignated as § 29.917(c). Section 29.917(a) sets forth a definition of the rotor drive system and its associated components and § 29.917(b) requires a design assessment to be performed. The intent of this paragraph (b) is to identify the critical components and to establish and/or clarify their design integrity to show that the basic airworthiness requirements, which are applicable to the rotor drive system, will be met.

b. Procedures .

(1) Section 29.917(a) General . The method of compliance for this section is unchanged.

(2) Section 29.917(b) Design Assessment . A design assessment of the rotor drive system should be carried out in order to substantiate that the system is of a safe design and that compensating provisions are made available to prevent failures classified as hazardous and catastrophic in the sense specified in paragraph (c) below.

In carrying out the design assessment, the results of the certification ground and flight Page E - 34 9/30/2008 AC 29-2C, Chg 3 testing (including any failures or degradation) should be taken into consideration.

Previous service experience with similar designs should also be taken into account (see also § 29.601(a)).

c. Definitions . For the purposes of this assessment, failure conditions may be classified according to the severity of their effects as follows: (1) Minor . Failure conditions which would not significantly reduce rotorcraft safety, and which involve crew actions that are well within their capabilities. Minor failure conditions may include, for example, a slight reduction in safety margins or functional capabilities, a slight increase in crew workload, such as routine flight plan changes, or some inconvenience to occupants.

(2) Major . Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be, for example, a significant reduction in safety margins or functional capabilities, a significant increase in crew workload or in conditions impairing crew efficiency, or discomfort to occupants, possibly including injuries.

(3) Hazardous . Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be- - (i) A large reduction in safety margins or functional capabilities; (ii) Physical distress or higher workload such that the flight crew cannot be relied upon to perform their tasks accurately or completely; (iii) Serious or fatal injury to a relatively small number of the occupants; (iv) Loss of ability to continue safe flight to a suitable landing site.

(4) Catastrophic . Failure conditions which would prevent a safe landing.

(5) Minimize . Reduce to the least possible amount by means that can be shown to be both technically feasible and economically justifiable.

(6) Health Monitoring . A Vibration Health Monitoring System (VHM) is used to acquire and process helicopter drive system vibration signals.

(i) The principal purpose of a VHM is to increase the likelihood of detection of incipient faults in the rotor drive system that could prevent continued safe flight and safe landing by providing timely warning of potential failures to the pilot and maintenance personnel.

Page E – 35 AC 29-2C, Chg 3 9/30/2008 (ii) VHM data can be used to improve the helicopter's monitoring practices to mitigate a major or a hazardous/severe failure. The VHM data can also be used to improve maintenance.

(iii) A VHM system can be used to monitor all or some critical components of the rotor drive systems. Critical components include the input driveshaft to the main gearbox from the engine, bearings, tail rotor drive shaft, hanger bearings, one way clutch, main rotor mast and tail rotor mast. The supplier should state the component coverage and the fault detection capability for all affected components. The health monitoring effectiveness should be validated by tests or analysis or both.

(iv) Typically, a VHM system consists of sensors (e.g., accelerometers and tachometer), signal acquisition, signal processing, data management, VHM alert generation and management, a pilot interface, and a maintenance interface.

(v) Signal Processing: The helicopter's rotors, its drive systems and engines are a mixture of complex and simple mechanical elements. Therefore, the sensors and signal processing and analysis techniques utilized should reflect the complexity of the mechanical elements and their vibratory modes.

(vi) AC 29 MG-15 provides airworthiness approval guidance for rotorcraft health usage monitoring systems. This guidance can be used for incorporating VHM.

d. Failure Analysis .

(1) The first stage of the design assessment should be the Failure Analysis, by which all the hazardous and catastrophic failure modes are identified. The failure analysis may consist of a structured, inductive bottom-up analysis, which is used to evaluate the effects of failures on the system and on the aircraft for each possible item or component failure. When properly formatted it will aid in identifying latent failures and the possible causes of each failure mode. The failure analysis should take into consideration all reasonably conceivable failure modes in accordance with the following: (i) Each item/component function(s).

(ii) Item/component failure modes and their causes.

(iii) The most critical operational phase/mode associated with the failure mode.

(iv) The effects of the failure mode on the item/component under analysis, the secondary effects on the rotor drive system and on the rotors, on other systems and on the rotorcraft. Combined effects of failures should be analyzed where a primary failure is likely to result in a secondary failure.

Page E – 36 9/30/2008 AC 29-2C, Chg 3 (v) The safety device or health monitoring means by which occurring or incipient failure modes are detected, or their effects mitigated. The analysis should consider the safety system failure.

(vi) The compensating provision(s) made available to circumvent or mitigate the effect of the failure mode (see also paragraph (1) below).

(vii) The failure condition severity classification according to the definitions given in paragraph (c) above.

(2) When deemed necessary for particular system failures of interest, the above analysis may be supplemented by a structured, deductive top-down analysis, which is used to determine which failure modes contribute to the system failure of interest.

(3) Dormant failure modes should be analyzed in conjunction with at least one other failure mode for the specific component or an interfacing component. This latter failure mode should be selected to represent a failure combination with potential worst- case consequences.

(4) When significant doubt exists as to the effects of a failure, these effects may be required to be verified by tests.

e. Evaluation of Hazardous and Catastrophic Failures .

(1) The second stage of the design assessment is to summarize the hazardous and catastrophic failures and appropriately substantiate the compensating provisions that are made available to minimize the likelihood of their occurrence. Those failure conditions that are more severe should have a lower likelihood of occurrence associated with them than those that are less severe. The applicant should obtain early concurrence of the cognizant certificating authority with the compensating provisions for each hazardous or catastrophic failure.

(2) Compensating provisions may be selected from one or more of those listed below, but not necessarily limited to this list.

(i) Design features; i.e., safety factors, part-derating criteria, redundancies, etc.

(ii) A high level of integrity: All parts with catastrophic failure modes and critical characteristics are to be identified as Critical Parts and be subject to a Critical Parts Plan (see AC 29.602.). Where a high level of integrity is used as a compensating provision, parts with a hazardous failure mode which would prevent continued safe flight may be included in a Critical Parts Plan or subjected to other enhancements to the normal control procedures for parts.

Page E – 37 AC 29-2C, Chg 3 9/30/2008 (iii) Fatigue tolerance evaluation.

(iv) Flight limitations.

(v) Emergency procedures.

(vi) An inspection or check that would detect the failure mode or evidence of conditions that could cause the failure mode.

(vii) A preventive maintenance action to minimize the likelihood of occurrence of the failure mode, including replacement actions and verification of serviceability of items which may be subject to a dormant failure mode.

(viii) Special assembly procedures or functional tests for the avoidance of assembly errors which could be safety critical.

(ix) Safety devices or use of vibration health monitoring systems are recommended in addition to those provisions identified in paragraphs e.(2)(vi) and e.(2)(vii).

[The remainder of this page intentionally left blank.]

Page E – 37.1 9/30/2008 AC 29-2C, Chg 3 AC 29.921. § 29.921 ROTOR BRAKE .

a. Background . Rotor brake safety requirements are intended not only to prevent adverse effects on aircraft performance due to brake drag but also to minimize the possibility of fires. These fires, caused by friction from a dragging rotor brake, have occurred both in flight and during ground operation with extremely hazardous consequences.

[The remainder of this page intentionally left blank.]

[Section AC 29.921 continued on next page.]

Page E – 37.2 AC 29-2C 9/30/1999 b. General . This rule requires (1) that any limitations on the use of the rotor brake must be established, and (2) that the control for the brake must be guarded to prevent inadvertent operation.

c. Limitations .

(1) The limitations on the use of the rotor brake should first be defined by the applicant and will normally consist of merely the maximum rotor speed eligible for application of the brake. In some installations, limitations associated with engine operation may be specified. For example, some “free power section” type turbine engines can be safely operated within certain low limits with the rotor brake engaged, while other engines cannot tolerate this condition. At least one manufacturer has included a maximum rotor speed for emergency rotor brake application. This is considered an enhancing safety consideration and is recommended.

(2) Control guard mechanisms to prevent inadvertent operation may be conventional. A cockpit evaluation should be conducted by flight test personnel to affirm the function of the guard and the brake, and that markings, if any, are adequate and that both latched and unlatched positions of the control do not interfere with other cockpit functions.

d. General qualification aspects should include: (1) The 400 applications required by § 29.923(j) conducted as a part of the § 29.923 endurance test.

(2) Torsional vibration measurements of the loads in the brake components and the rotor drive system during a critical brake engagement procedure, with appropriate consideration in the fatigue evaluation for these components. Brake engagements should be conducted with and without collective control displacement as authorized by the flight manual or a training manual.

(3) Brake component temperature measurements during a critical brake application in conjunction with an evaluation of the general brake compartment for compliance with §§ 29.863 and 29.1183.

(4) Placards, decals, and flight manual limitations and instructions appropriate to operate the rotor brake safely.

(5) An evaluation for hazardous failure modes as required by § 29.901(c). If the brake hydraulic system is integral with the rotorcraft hydraulic system, failure modes of pressure regulators and control valves, including valve leakage, will be of interest.

Mechanical cams, calipers, and levers may be prone to seize or fail to release the brake due, in part, to corrosion and lack of lubrication to be expected when brake components encounter high temperature cycling Page E - 38 9/30/99 AC 29-2C NOTE: Most rotor brakes include nonmetallic pucks or liners, usually included in nonrotating brake components, which are subject to wear in proportion to the number of applications. Replacement of these pucks during the § 29.923 endurance test has been found acceptable provided the reason for replacement is simply wear and not because of any change in brake loading, disk temperature, or vibratory characteristics which can be expected in service. Verify that the maintenance manual includes a routine check for excessive puck or liner wear.

e. Other comments . Rotor brakes may be added to the basic design as a postcertification program without necessarily reconducting the complete § 29.923 endurance test provided: (1) Steady and vibratory stresses in brake components, the rotor drive system, and in the rotor system itself are determined and shown to be acceptable.

NOTE: Moments, stresses, etc., from brake operation apply loads to the drive system in the reverse direction to normal powered flight. Advise the airframe engineer to require evaluation of chordwise bending loads in the hub and blade components of the main rotor system.

(2) The 400 brake engagements of § 29.923(j) should be accomplished with a complete rotor and rotor drive system, followed by disassembly sufficient to verify that all components subject to loads from the brake remain serviceable. Since this test may be so short as not to cause appreciable wear patterns to appear, special pretest coatings such as black oxide or Du-Lite may be needed on gear teeth and bearing races to distinguish and evaluate the contact patterns. Information on maximum deceleration rates should be supplied to the manufacturer of the engines to be used in the rotorcraft for evaluation of the acceptability of backloading or motoring of turbines, fuel control components, torque meters, etc.

AC 29.923. § 29.923 (Amendment 29-17) ROTOR DRIVE SYSTEM AND CONTROL MECHANISM TESTS .

a. Explanation .

(1) This rule is intended to require demonstration that the rotor drive system, as defined in § 29.917(a), is capable of normal operation within the limitations proposed, without hazard of failure from excessive wear or deterioration due to mechanical loads.

The basic test is not designed and should not be expected to demonstrate safety from oscillatory stresses normally investigated under §§ 29.571 and 29.907, although any data generated by these tests, which are in fact applicable to showing compliance with §§ 29.571 and 29.907, may be used. Some variations in the endurance test plan to generate data applicable to the vibration substantiative effort or other qualification aspects may be acceptable if the basic requirements of the endurance test are preserved.

Page E - 39 AC 29-2C 9/30/99 (2) The construction of this rule is such that a series of runs, each at least (but not limited to) 10 hours in length must be repeated 20 times, for a total of (at least) 200 hours of test, not including time required to adjust power or to stabilize operating conditions for those conditions that require stabilization. Extension of the total test beyond 200 hours (or extension of test runs beyond 10 hours) will occur if qualification for the 2½-minute one-engine-inoperative (OEI) optional rating is proposed by the applicant. The 30-minute OEI rating qualification test will extend the test beyond 200 hours for rotorcraft equipped with three or more engines. Also, compliance with § 29.923(g) may result in extended endurance tests if dynamic or malfunction conditions exist which adversely affect the endurance tolerance of the rotor drive system.

Section 29.923(a) should be interpreted as requiring test runs or cycles to be repeated in essentially the same sequence, although more than 10 hours may be needed to complete a run or cycle. This section also requires the test to be conducted “on the rotorcraft.” This means a rotorcraft in conformity to the design for which approval is requested. However, many nonconformity features, such as doors, some cowling and instrumentation, fuel tanks (alternate external fuel supply may be utilized), interior features, fire detectors, extinguishers, inlet ducts, exhaust baffles, etc., may be acceptable provided each item is technically considered and found to be unimportant to the test results. Any significant deviations from the conformed rotorcraft configuration should be coordinated with the cognizant FAA/AUTHORITY engineering staff and if found acceptable, documented as such. The restraint (tie-down) arrangement used during the test will necessarily be arranged to react rotor thrust loads in lateral as well as vertical directions. However, the restraint should permit normal deflections due to rotor thrust in the engine and drive system support arrangement.

(3) Safety cables may be installed normal to the tail boom at the tail rotor gearbox location; however, restraint may be provided to keep airframe deflections from exceeding those expected in normal and accelerated flight.

(4) The test torque requirements of § 29.923(a)(3)(i) mean the torque values for which approval is requested but not to exceed the values approved for the respective limits for the engine to be used. However, an applicant should be allowed to qualify the rotor drive system for torque values higher than those for which approval is requested if the engines actually used are capable of the torque and can be shown by an output shaft torsional investigation to be equivalent or conservative with respect to torsional vibration to the engines proposed for the initial certification configuration. Variations in rotational speed from the certification values should not be allowed except where careful evaluation of vibration aspects, bearing loads, centrifugal stiffening effects, and torque variations are conducted.

(5) The rotor configuration required by § 29.923(a)(3)(ii) is intended to assure that lift, torque, and vibration loads to be expected in service are introduced into the endurance test, although the presence of the vibration aspects does not normally satisfy the vibration evaluations required by §§ 29.571 and 29.907. In fact, vibration modes may be changed and amplified by the tie-down restraints and the increased thrust to be Page E - 40 9/30/99 AC 29-2C expected from in-ground effects on the rotor system. These effects, although unquantified, are intended as a normal part of endurance testing. Preproduction rotor blades have been successfully used in endurance tests but only after specific investigations of blade properties such as stiffness, inertia and inertia distribution, thrust and blade bending, and torsional frequency response have been carefully compared to assure validity of the test. The endurance test includes testing of the rotor control mechanism. Conformity of the rotors may be very significant to this aspect of the test.

b. Procedures .

(1) Section 29.923(b)(1) prescribes the takeoff portion of the endurance test.

This test involves a series of 5-minute repetitive runs at the torque and at the engine/rotor rotational speed selected by the applicant for the takeoff limit for the rotorcraft. These values of torque (manifold pressure, for reciprocating engines) and RPM should correspond to the red radials on the corresponding powerplant instruments, except on installations where uncompensated engine governor “droop” results in a higher rotational speed for lower powers. The requirement in this section for declutching the engine may be difficult to achieve if engine deceleration and rotor system deceleration rates are similar. In some cases, the engine fuel control deceleration schedule may be adjusted to achieve clutch disengagement, otherwise, an engine shaft brake mechanism may be needed.

(2) The torque and speed requirements for the optional 2½-minute one-engine-inoperative (OEI) tests should be interpreted as described above for the takeoff runs. If the test is conducted during warm ambient conditions, excessive engine gas temperatures may be required to achieve the torque and speed conditions required by this part of the test. Minor adjustments in the run schedule may be allowed to take advantage of cooler nighttime ambient temperatures. Addition of water/alcohol systems to increase engine hot-day power may be appropriate in some instances. Liquid nitrogen spray into engine inlets has also been used to depress inlet temperatures sufficiently to obtain test conditions.

(3) In § 29.923(c), (d), (e), and (f), the torque requirements should be interpreted as above; i.e., the run should be made with maximum continuous torque or percentage thereof, as specified by the subparagraph, and the rotational speed should be maximum continuous for paragraph (d) and the lowest permissible “power-on” speed for paragraphs (e) and (f). Rotor control cycling must be accomplished during the “maximum continuous” portion of the endurance run. The controls of concern are the flight controls; i.e., cyclic and directional controls for rotorcraft with tail rotor and single main rotor. The collective control is normally used to set power and is not involved in control cycling. During control cycling the controls may be cycled from stop to stop, or a limited travel may be accepted if the travel produces the maximum fore and aft, left and right, and yaw thrust components of the rotors as measured in flight. One method of determining the required control displacement is to measure main rotor mast bending in level forward flight at maximum continuous power for the forward control displacement limit, and in level rearward flight at maximum continuous power (or the power Page E - 41 AC 29-2C 9/30/99 associated with the maximum rearward flight speed to be expected) for the aft control displacement limit. Using the same mast bending instrumentation, with the rotorcraft in the ground tie-down situation, and with collective control set for maximum continuous power, displace the cyclic fore and aft to obtain the same mast bending as measured in flight. Similar measurements and control displacements may be used for sideward thrust components. Yaw control displacement should consider maneuver requirements in conjunction with sideward flight. Critical gross weight and center of gravity should be used to establish test conditions. These same procedures may be used to establish limited control positions required to comply with § 29.923(i) except that typical flight conditions to be used would be stabilized level flight at maximum continuous power, climb at maximum continuous power, and hovering, including stabilized sideward and rearward flight. Note that for § 29.923(i)(1) vertical thrust is required. Depending on the mast angle and center of gravity, this condition may not necessarily involve zero mast bending loads. Vertical thrust may be used during the takeoff run, including the runs at 2½-minute power and the overspeed run of § 29.923(h). One-engine-inoperative runs (§ 29.923(k)) should be conducted with the cyclic set for maximum forward thrust. For these runs and any run that does not specify the position for the yaw control, that control should be set to react main rotor torque.

(4) Section § 29.923(g) provides for introducing special tests into the endurance tests to demonstrate that the transmission and drive system can tolerate certain engine malfunctions to be expected. This was originally directed at demonstrating safety in the event of spark plug or magneto failures of reciprocating engines. Turbine engines normally do not exhibit failure modes suitable for substantiation by endurance testing; however, severe or abusive operating conditions which must be expected to occur in service should be defined and included in this test.

Conditions or phenomena to be considered should include but not be limited to moderate engine surge, abusive clutch engagements, torque mismatching, anticipated control mishandling, and so forth. Alternatively, repeating the takeoff run of § 29.923(b) may be appropriate. It is not intended that the special testing for 2½-minute power be repeated if a rerun of the takeoff power run is required by § 29.923(g).

(5) Section 29.923(h) requires overspeed testing at the torque which will produce maximum continuous power and at the maximum rotational speed to be expected. Normally this would be the maximum transient, power-on rotor speed available with speed controls operating. Special control adjustments for test purposes may be needed to achieve the required test conditions.

(6) Section 29.923(i) requires stabilized flight control displacement according to a prescribed schedule. The control displacement should be the same as derived to show compliance with § 29.923(c)(2).

(7) Section 29.923(j) requires 400 clutch and brake engagements. These tests are prescribed to establish a level of reliability of clutch and brake components installed as a part of the rotor drive system of rotorcraft. The clutch tests apply to all clutches installed to comply with § 29.917(b)(3), and each such clutch must be tested. A rotor Page E - 42 9/30/99 AC 29-2C brake is not required for certification, although a brake of some type may be installed temporarily to facilitate conducting the clutch testing required by this section. Clutch disengagement is also required by this section, thus, malfunction of the disengagement feature would be a basis for discontinuance. Some rotorcraft configurations (those with single-spool turbine engines or reciprocating engines) include an additional clutch to decouple the engine from the drive system to facilitate engine starting. These clutches should also be exercised at least 400 times during this test.

(8) Section 29.923(k) sets forth the optional tests to be conducted if a 30-minute OEI rating is requested. It may be noted that the time for conducting this test replaces time deducted from the run of § 29.923(f). Flight control positions should be set for level flight or climb, whichever produces the maximum forward thrust component, and the antitorque system control should be set to react the maximum rotor torque. The torque and rotational speed values should be the maximum for which approval is requested.

(9) Section 29.923(m) normally is satisfied by the requirements of §§ 29.571 and 29.907.

(10) Section 29.923(n) requires special tests for rotor drive systems designed to operate at two or more gear ratios. Depending on the limitations and instructions proposed for operating at other gear ratios, additional tests (beyond the normal 200-hour schedule) or substitutions into the basic test should be conducted to qualify the rotor drive system for operations at other gear ratios. The length of testing, torque and speed requirements, overspeed tests, and control positions for these tests should parallel the requirements of the basic endurance test.

(11) Section 29.923(o) requires the rotor drive system and rotor control mechanism to be in a serviceable condition at the end of the test. Verification of this requirement requires a complete disassembly and examination of the entire rotor drive system and rotor control mechanism. The disassembly itself should be closely monitored for evidence of adequate breakaway torque on all bolted fasteners. Samples of lubrication from oil sumps and filters should be retained for spectrographic analysis, and seals should be examined for possible damage due to test requirements. Care should be taken to differentiate between seal damage and bearing damage due to disassembly procedures so that the direct results of the test may be properly considered. Close visual observation of each tooth on each gear is necessary to affirm proper load/contact patterns and absence of excessive surface stress or scrubbing motions. Bearings should be examined to verify that ball or roller paths are within limits, bearing cages are undamaged, and bearing balls or rollers and their races are free from pitting. Any evidence of bearing races turning or spinning in respective housing or bores probably indicates design or fit deficiencies. The applicant should have available wear limits data which include items such as distance across pins and tooth profile limits for gears. Many of these items require special, close tolerance inspection equipment and trained inspectors to determine compliance. In some instances bearings, clutches, oil pumps, etc., should be returned to the original manufacturer for a finding of Page E - 43 AC 29-2C 9/30/99 serviceability. Localized overheating, usually exhibited by discolorations is an indication of an unsatisfactory condition. Should any of the items discussed above or other defects appear such that the component is unserviceable, a redesign which includes recognizable improvements should be required before authorizing a retest. To simply “try again” in hopes of success should not be accepted.

(12) This section also prohibits intervening disassembly which might affect test results. Generally, this simply means no disassembly whatsoever. However, some very limited disassembly can usually be conducted provided care is used to assure that items such as critical fastener torques or gear backlash controls are not disturbed.

c. Additional Test Considerations .

(1) Pressure Lubricated Gearboxes . The endurance test hardware can be adjusted/modified to sustain high-limit oil temperature and low-limit oil pressure in order to provide a basis for approval of the values listed as limits. A minimum of 20 hours at maximum continuous torque and maximum continuous rotational speed should be involved in the test. Other parameters such as minimum oil temperature and maximum oil pressure may more appropriately be evaluated by bench test. The significant points here are effects of extremely high oil pressure (due to the high viscosity of cold oil) on any positive displacement oil pump, on filters for possible collapse, on oil coolers for possible rupture due to internal pressure, seals, bypass valves, and most important, adequate lubrication of gears, bearings, etc., under conditions of minimal oil flow.

Normally, an operation restriction against exceeding idle power/speed conditions until significant warm-up occurs is prescribed. Individual component qualification tests may provide data to meet some of these aspects.

(2) The existing endurance test schedule does not necessarily provide for any asymmetric power inputs from multiengine drive system arrangements. For this situation, the drive system should at least be subjectively evaluated for possible hazards or excessive loads to be expected from asymmetric torque inputs and additional testing prescribed under the authority of § 29.923(g). The extent and severity of these tests should be established in consideration of the design peculiarities, the recommended operating procedures, and any OEI tests included in this test schedule.

(3) Accessory Drives . Normally, all accessory drives on a gearbox will be loaded during the endurance test. Electrical load banks or other suitable methods may be used to assure that the generator drives are loaded and thus properly qualified.

Hydraulic pumps may be loaded by resetting hydraulic system relief valves to maintain limit pressure (load) continuously. If this condition is excessively severe, a method of load cycling may be appropriate. Note that accessory loads reduce the power available to the main rotor. Also, tail rotor loads are, insofar as the transmission is concerned, another large accessory. Care should be taken to assure that in-flight unloading of these accessory drives, including the tail rotor does not subject the main gearbox to loads significantly beyond those qualified by endurance tests.

Page E - 44 9/30/99 AC 29-2C (4) Gearbox Oil Tanks . Normally, gearbox oil is contained in an integral cast sump which, for other reasons, has sufficient strength to obviate the need for pressure tests. However, a subjective evaluation should be made to assure that detail design features such as sight gauges, filler caps, etc., offer adequate strength.

AC 29.923A. § 29.923 (Amendment 29-26) ROTOR DRIVE SYSTEM AND CONTROL MECHANISM TESTS .

a. Explanation . Amendment 29-26 includes additional endurance test criteria for a new continuous OEI rating, and clarifies the torque and RPM relation intended for the various power ratings involved in the tests prescribed by this section.

(1) Section 29.923(a)(1) was amended to require that the test cycle be extended beyond 10 hours if OEI rating tests are included in the test program. This change was needed to maintain the cycle aspect of the test if OEI ratings are included.

(2) Section 29.923(a)(3) was amended to include rotational speed as a part of the test because the term “torque” by itself does not adequately define the test requirements.

(3) Section 29.923(b)(2), (f), and (k) were amended to add the test requirements for the new continuous OEI rating and retain, as an alternate, the 30-minute OEI rating tests for those applicants who may request this rating. This change provided a regulatory test basis for qualifying the rotor drive system for optional OEI ratings.

(4) Section 29.923(g) was amended to remove the inference that the 2½-minute OEI runs should be repeated if the takeoff run is reconducted. Under these circumstances, additional testing for the 2½-minute rating is unnecessary.

b. Procedures .

(1) The construction of this amendment is such that a series of runs, each at least (but not limited to) 10 hours in length, should be repeated 20 times for a total of at least 200 hours of test. The time required to adjust power or to stabilize operating conditions for those conditions that require stabilization is not included.

Figure AC 29.923A-1 shows a graphic representation of the 10-hour test cycle.

Extension of the total test time beyond 200 hours (or extension of test runs beyond 10 hours) will occur if qualification for the 2½-minute, 30-minute, or continuous OEI optional ratings is proposed by the applicant for rotorcraft equipped with two or more engines. Also, compliance with § 29.923(g) may result in extended endurance tests if dynamic or malfunction conditions exist which adversely affect the endurance tolerance of the rotor drive system. Section 29.923(a) should be interpreted as requiring test runs or cycles to be repeated in essentially the same sequence, although more than 10 hours may be needed to complete a run or cycle. This section also requires the test to be conducted “on the rotorcraft.” This means a rotorcraft that is in conformity to the Page E - 45 AC 29-2C 9/30/99 type design for which approval is requested. However, many nonconforming features, such as doors, some cowling and instrumentation, fuel tanks (alternate external fuel supply may be utilized), interior features, fire detectors, extinguishers, inlet ducts, exhaust baffles, etc., may be acceptable provided each item is technically considered and found to have no impact on the test results. Any significant deviations from the conformed rotorcraft type design should be coordinated with the cognizant FAA/AUTHORITY engineering staff and, if found acceptable, properly documented.

The restraint (tie-down) arrangement used during the test should be arranged to react rotor thrust loads in lateral as well as vertical directions. The restraint should permit normal deflections due to rotor thrust in the engine and drive system support arrangement.

(2) The test torque and speed requirements of § 29.923(a)(3)(i) refer to the torque/speed combination (or power) values for which approval is requested. The requested torque/speed combination should not exceed the limits approved for the respective engine(s) to be used. An applicant may qualify the rotor drive system for torque values higher than those approved for the engine.

(3) In §§ 29.923(c), (d), (e), and (f), the torque requirements should be interpreted as above; i.e., the run should be made with maximum continuous torque or percentage thereof, as specified by the subparagraph; and the rotational speed should be maximum continuous for paragraph (d) and the lowest permissable “power-on” speed for paragraphs (e) and (f). Rotor control cycling should be accomplished during the “maximum continuous” portion of the endurance run. The controls of concern are the flight controls (cyclic and directional controls for typical rotorcraft). The collective control is normally used to set power and is not involved in control cycling. During control cycling, the controls may be cycled from stop to stop; or a limited travel may be accepted if the travel produces the maximum fore and aft, left and right, and yaw thrust components of the rotors as measured in flight. The frequency for cycling the controls is defined in §§ 29.923(c)(1), (2), and (3), and specified in Note 3 of figure AC 29.923A-1. One method of determining the required control displacement is to measure main rotor mast bending in level forward flight at maximum continuous power for the forward control displacement limit, and in level rearward flight at maximum continuous power (or the power associated with the maximum rearward flight speed to be expected) for the aft control displacement limit. Using the same mast bending instrumentation, with the rotorcraft in the ground tie-down situation, and with collective control set for maximum continuous power, displace the cyclic fore and aft to obtain the same mast bending as measured in flight. Similar measurements and control displacements may be used for sideward thrust components. Yaw control displacement should consider maneuver requirements in conjunction with sideward flight. Critical gross weight and center of gravity should be used to establish flight test conditions. These same procedures may be used to establish limited control positions required to comply with § 29.923(i), except that typical flight conditions to be used would be stabilized level flight at maximum continuous power, climb at maximum continuous power, hover, and stabilized sideward and rearward flight. Note that for § 29.923(i)(1), vertical thrust is required. Depending on the mast angle and center of gravity, this Page E - 46 9/30/99 AC 29-2C condition may not necessarily involve zero mast bending loads. Vertical thrust may be used during the takeoff run, including the runs at 2½-minute power and the overspeed run of § 29.923(h). One-engine-inoperative runs (§ 29.923(k)) should be conducted with the cyclic set for maximum forward thrust. For these runs and any run that does not specify the position for the yaw control, that control should be set to react to main rotor torque.

(4) Section 29.923(k) sets forth the optional tests to be conducted if a 30-minute or a continuous OEI power rating is requested. Flight control positions should be set for level flight or climb (whichever produces the maximum forward thrust component) and the anti-torque system control should be set to react the maximum rotor torque. The torque and rotational speed values should be the maximum for which approval is requested.

Page E - 47 AC 29-2C 9/30/99 :: O.E.I. RUN 30 MIN O.E.I. OR CONT O.E.I.

EACH ENGINE 29.923 (k) NOTE 6 OVERSPEEDATM.C. TORQUE 29.923 (h) "' MALFUNCTION RUN OR REPEATT.O. RUN 29.923 Wl NOTE 60% M.C.

W 00 29.923 (f) <.:,W z:a; NOTE 4

-

....

~~~ :::) N 00 0 f- ct: z <( :::> ~ WwO 00 M (t'.N :a; o.. I wW+ :::>~ <D ~a::g 0~ 00 • N Io 80% M.C.

>- W LL ' <( 00 _J 0 29.923 (e) Ww w 0 _, ~ct: <::(}<( f- :::> ct: ct'.f-' (:, "' o=>o LL ct:ot::;, 90% M.C.

OI 29.923 (d) b~ ct:~ ...

M M w MAXIMUM CONTINUOUS RUN f- 29.923 (C) z N N ~ w w f- f- 0 z z

\

C : I z 1 ........

:::> ~ ct: LL ' LL

-

'= 0

w :!:: ~ I • <(

-- 0

. .

. ' ' ' . . .

' ' . .

' Q) iii iii (.) (.) (.) (.) .!!1 '§ 0 0 :!1 f-C :a; :a; :a; :a; ....

~ ~ -;j< -;j< -;j< LL <°5 :a; :a; 0 0 0 00 <D "°!

"' M N Page E - 48 9/30/99 AC 29-2C =ll:: t0 - ~ ~ 0

=ll:: ~ m m z z (j) (j) - z z m m t0 t0 ~ -- t0 t0 ~ z z 0 m m

- ~ -

• ; "

.

- ROTOR

•

NOTE 1 REQUESTED OF .

IS 40 I - 29.923A-2

"

IF

~ ' AC

!3 5

f0 s

I

=ll:: ~ m z (j) z m t0 ~ t0 ~ z 0 m

-

'

-

TIME - MINUTES FIGURE

•

~- ~ ~ MINUTE OEI RATING 20 DRIVE SYSTEM ENDURANCE TEST TAKEOFF RUN PORTION ' 21/2 ~ • I ~

•

• NOTE 2 • . . .

- - -

Idle Flt. Idle Gr.

T.O.

2 1/2 MIN. OE Page E - 49 AC 29-2C 9/30/99 TEST ROTOR RATING OF OEI IF ENDURANCE PORTION MINUTE RUN SYSTEM SECOND/2 REQUESTED TAKEOFF DRIVE IS ~ ~ m z (j) z m (>) 0 (f) m () QO t0 ~ z 0 m ..

29.923A-3 MINUTES AC - NOTE ◄ 'It ~ m z (j) z m (>) 0 (f) m () QO t0 ~ z 0 m TIME FIGURE ;:.; NOTE * - OEI OEI T.O.

IDLE CONT.

MIN SEC GR.

OR MIN Page E - 50 9/30/99 AC 29-2C Figures AC 29.923A-1, AC 29.923A-2, and AC 29.923A-3 Notes 1. If the 2½-minute OEI rating is requested, the following should be conducted for each engine: Demonstrate 2½-minute OEI power twice or 5 minutes per cycle for a total of 100 minutes at 2½-minute OEI power during the 200-hour endurance test. See figure AC 29.923A-2 for a graphic description of the takeoff run for a two-engine rotorcraft using the 2½-minute OEI rating (Refer to § 29.923(b)(2)). If the 30-second/2-minute OEI rating is requested, the following should be conducted for each engine: Demonstrate 30-second OEI power followed by 2-minute OEI power. After 2 minutes reduce and stabilize power to 30-minute or continuous OEI level. Once the power is stabilized reapply 2 minute OEI power. This should result in 4½ minutes per cycle for a total of 10 minutes at 30-second OEI power and 80 minutes at 2-minute OEI power during the 200-hour endurance test. See figure AC 29.923A-3 for a graphic description of the takeoff run for a two-engine rotorcraft (refer to § 29.923(b)(3)). If either the 2½-minute or 30-second/2-minute OEI ratings are demonstrated, the takeoff run portion of figure AC 29.923A-1 will be longer than 1 hour as shown in figure AC 29.923A-2 or figure AC 29.923A-3, respectively.

2. Apply the rotor brake during the first minute of the 5-minute idle period. Conduct 400 brake applications during the 200-hour endurance test (§ 29.923(j)).

3. During the maximum continuous run, cycle the rotor controls 15 times per hour: (Refer to §§ 29.923(c)(1) - (3)). The cyclic control should be cycled through maximum vertical thrust, maximum forward, maximum left, maximum right, and maximum rearward thrusts. The pedal controls should be cycled through maximum right, neutral, and maximum left positions. Each maximum cyclic and pedal control position should be held for at least 10 seconds. During the remainder of the test, set the yaw control to react to the main rotor torque, and set the flight controls to achieve: Condition Portion of Test max vertical thrust 20 percent max forward thrust 50 percent max left thrust 10 percent max right thrust 10 percent max rearward thrust 10 percent Page E - 51 AC 29-2C, Chg 4 5/1/2014 Figures AC 29.923A-1, AC 29.923A-2, and AC 29.923A-3 Notes (continued) 4. The 60 percent maximum continuous run is 2 hours (Refer to § 21 29.923(f)), unless either 30-minute OEI or continuous OEI power is requested. In that case, the 60 percent maximum continuous run is 1 hour.

5. A 1-hour malfunction run (if deemed necessary) or the takeoff run is repeated (without OEI portions). Refer to § 29.923(g).

6. The OEI run defined in § 29.923(k) is not required unless an OEI power rating is requested. If a 30-minute OEI power rating is requested, each engine in sequence should be run at the 30-minute OEI condition for 30 minutes. If a continuous OEI power rating is requested, each engine in sequence should be run at the continuous OEI condition for 1 hour. The total OEI run time may exceed the 1 hour shown in figure AC 29.923A-1.

AC 29.923B. § 29.923 (Amendment 29-31) ROTOR DRIVE SYSTEM AND CONTROL MECHANISM TESTS.

a. Explanation . Amendment 29-31 added § 29.923(p) that defines qualification tests for lubricants used in the rotor drive system and control mechanisms. Section 29.923(p) contains a requirement for a portion of the system qualification tests to be accomplished with specific lubricating oil temperatures and pressures. It also provides requirements for the qualification of additional or alternate lubricants by equivalent testing or by using comparative analysis of lubricant specifications and rotor drive and control system characteristics.

b. Procedures .

(1) When performing the ten-hour endurance test cycles prescribed in § 29.923(a), a minimum of three of the required test cycles should be accomplished with the main transmission and, if applicable, for gearboxes that are used for rotor phasing with the lubricating oil temperature set to the maximum operating temperature for which approval is requested. The oil temperature should be measured at the temperature sensor location that is used for flight crew indication of the oil temperature. For main transmission and gearboxes that are pressure lubricated, the testing should include setting the oil pressure to the minimum operating oil pressure for which approval is requested. The oil pressure should be measured at the pressure sensor location that is used for flight crew indication of the oil pressure. If approval of OEI ratings is sought, the ten hour test cycles must be extended to include the OEI endurance testing prescribed in § 29.923(b)(2) and (k).

(2) To be approved for use in the main transmission and gearboxes, lubricants must meet the specifications of lubricants that were used to satisfy the test requirements in § 29.923. Alternate or additional lubricants may be approved provided they are qualified by testing that is considered equivalent to the endurance testing Page E –52 5/1/2014 AC 29-2C, Chg 4 required by § 29.923. A comparative analysis may be used to approve alternate or additional lubricants in lieu of testing, provided the analysis shows that the lubricant properties, specifications, and applications are equal to or better than those of a lubricant that has been previously approved for use in the main transmission or gearbox.

AC 29.923C. § 29.923 (Amendment 29-34 and 29-40) ROTOR DRIVE SYSTEM AND CONTROL MECHANISM TESTS.

a. Explanation . This paragraph, AC 29.923C, reflects changes made by Amendment 29-34 and 29-40. Amendment 29-34 added § 29.923(b)(3) that defines qualification tests for 30-second/2-minute OEI ratings. This new paragraph also allows for the 30-second/2-minute OEI portion of the endurance test to be accomplished on a representative bench test stand using the drive system components which can be adversely affected by these tests. Amendment 29–40 doubles the endurance test time for 2-minute OEI for each power section.

b. Procedures .

(1) For accomplishment of the endurance test for 30-second/2-minute OEI, § 29.923(b)(3) requires that immediately following one of the 5-minute power on takeoff runs of § 29.923(b)(1) each engine must simulate a power failure and each engine providing power after the failure must apply the maximum torque and maximum speed for use with 30-second OEI power. This power level should be maintained for at least 30-seconds. The 30-second OEI power should then be followed by an application of the maximum torque and maximum speed for 2-minute OEI power for at least 2 minutes. After the 2-minute OEI power application the power should be reduced to and stabilized at 30-minute or continuous OEI, (whichever rating the rotorcraft will be certified with). After the power has been stabilized, the maximum torque and maximum speed for use with 2-minute OEI power should be reapplied for at least 2 minutes.

Figure AC 29.923A-1 shows a graphic representation of the ten-hour test cycle with the 30-second/2-minute OEI segment included for each engine presented in figure AC 29.923A-3. This figure shows the OEI test segment being accomplished immediately following a 5 minute takeoff run. The OEI test segment can be accomplished after any of the 5-minute takeoff run segments. Section 29.923(b)(3) also requires that one of the 30-second/2-minute OEI segments for each engine be accomplished from the flight idle condition.

(2) Additionally, due to the damage inflicted on the engines and the ensuing cost caused by operating the engine at these powers, the 30-second/2-minute portion of the endurance test can be accomplished on a bench test rig found to be representative of the rotorcraft. The representative bench test rig should have the ability to generate the torques, speeds, torsional vibration frequency, and engine acceleration rate generated by the actual installation. The power should have the same method/path of application as that used on the rotorcraft. The test rig should be configured with the same components used for conducting the endurance test on the rotorcraft except that Page E – 53 AC 29-2C, Chg 4 5/1/2014 the test components not affected by asymmetric power application may not have to be installed (i.e., if a separate combining gearbox is used it may not be necessary to have the main transmission installed on the bench test rig).

(3) When conducting the bench test for 30-second/2-minute OEI, it is not necessary to reaccomplish the takeoff portion of the endurance test. The simulated power failure and application of 30-second/2-minute OEI power by the remaining power section should be accomplished after the input power has stabilized at takeoff power.

The takeoff portion of the endurance test should be accomplished on the rotorcraft.

Page E –54 7/6/2012 AC 29-2C, Chg 4 AC 29.927. § 29.927 (Amendment 29-13) ADDITIONAL TESTS.

a. Section 29.927(a) : (1) Explanation . This paragraph is authority to require any special tests or investigations to establish that the rotor drive system is safe.

(2) Procedures . The certification engineer should review the design of the rotor drive system and its installation and intended operation for features or conditions that may not be adequately qualified in the tests prescribed by this Part. Additional qualification test programs should be developed and accomplished to ensure safe operation of the system. Items of interest would include poorly defined load paths associated with redundant design features, flight deflections of structure, mounting arrangements which may not be properly qualified by ground tests, and special or unusual operating procedures which are anticipated by the applicant.

b. Section 29.927(b): (1) Explanation . This paragraph prescribes testing to qualify the rotor drive system for the power excursions to be expected with governor-controlled engines wherein the engine power changes automatically to maintain rotorspeed at preselected values. At high collective flight control displacements, the normal rotor speed droop will result in governor-controlled engines automatically accelerating to maximum fuel flow or to any other power, speed, temperature, or torque limiting device, regardless of crew action or artificially established limitations reflected by instrument markings. This high power condition can occur typically during a normal landing when the crew applies high collective to cushion ground contact or, for multiengine rotorcraft, during any flight regime when an engine fails and the corresponding loss of power results in drooping the rotor speed. Special tests are prescribed by this section to provide assurance that the rotor drive system can safely sustain these conditions. The tests of this section should be conducted without intervening disassembly, and all rotor drive system components should be in serviceable condition after the test. It is permissible but not required that these tests be performed on the same specimen of the rotor drive system used to show compliance with § 29.923.

(2) Procedures . These tests should be conducted on a ground-test rotorcraft conformed to the type design configuration similar to that required for endurance testing under § 29.923. Cyclic and collective control may be set to simulate vertical lift and antitorque control set and/or adjusted to react to main rotor torque. Rotation speed should be maximum normal for the test condition; i.e., for the all-engines test under § 29.927(b)(1), use the maximum RPM for takeoff power. For the one-engine-inoperative (OEI) test of § 29.927(b)(2), RPM droop, if any, that would occur in service, may be allowed. Since the OEI test of § 29.927(b)(2) usually requires the remaining engine(s) to produce power not usually available under normal atmospheric conditions, some supplemental method, such as refrigerating and/or ramming inlet air, or overfueling the engine, may be required. Alternatively, bench testing (transmission Page E – 54.1 AC 29-2C, Chg 4 7/6/2012 test rig) of the rotor drive system (using only the components subject to the higher OEI power, if desired) may be appropriate providing close simulation of the rotor drive system installation environment is achieved. Overtesting, to compensate for inadequacies in the bench test setup may be negotiated with the FAA/AUTHORITY approval office. Note that compliance with § 29.903(b) requires that the remaining engine(s) be capable of continued safe operation under the same conditions as dictated by this test. The engine manufacturer may have already conducted tests adequate to substantiate this requirement. If not, his assistance in testing and the subsequent serviceability finding is imperative.

c. Section 29.927(c) : (i) This section prescribes a test to demonstrate that any failure resulting in the loss of lubrication pressure to the rotor drive primary oil system will not impair the capability of the rotorcraft to operate under autorotative conditions for 15 minutes.

(ii) The regulation is intended to apply to pressurized lubrication systems and has not been applied to splash lubricated gearboxes since historically their design has not been as critical or complex when compared to pressurized systems. The likelihood of loss of lubrication is significantly greater for transmissions that use pressure lubrication and external cooling. This is due to the increased complexity of the lubrication system and the external components that circulate oil outside the gearbox. A pressure lubrication system is more commonly used in the rotorcraft’s main transmission but may also be used in auxiliary transmissions or gearboxes.

(iii) The lubricating system has two primary functions. The first is to provide lubricating oil to contacting or rubbing surfaces and thus reduce friction losses. The second is to dissipate heat energy generated by friction of meshing gears and bearings, thus maintaining surface and material temperature. Accordingly, a loss of lubrication leads to increased friction between components and increased component surface temperatures. With increased component surface temperatures, component surface hardness can be lost, resulting in the inability of the component to carry or transmit loads. Thermal expansion in transmission components can eventually lead to the mechanical failure of bearings, journals, gears, shafts, and clutches that are subjected to high loads and rotational speeds. A loss of lubrication may result from internal and external failures. Failures include, but are not limited to, oil lines, fittings, seal plugs, sealing gaskets, valves, pumps, oil filters, oil coolers, accessory pads, etc. A leak caused by a crack in the transmission outer case need not be considered as a source of a loss of lubrication provided the outer case has been structurally substantiated to satisfy the requirements of §§ 29.307, 29.923(m), and 29.571.

(2) Procedures . Conventionally, a bench test (transmission test rig) is used to demonstrate compliance with this rule. Since this is essentially a durability test of the transmission to operate with residual oil, typically the worst case failure (i.e., the undrainable oil or the oil remaining after a severe pressure leak, whichever results in a greater loss of oil in the transmission’s normal lubrication system) is used as a critical Page E –54.2 7/6/2012 AC 29-2C, Chg 4 entry point for the test. The transmission should be stabilized at the torque associated with maximum continuous power (reacted as appropriate at main mast and tail rotor output quills) at a normal main rotor speed, oil temperature that is at the highest limit for continuous operation, and oil pressure that is within the normal operating range. A vertical load should be applied at the mast, equal to the gross weight of the rotorcraft at 1g. Once the transmission oil temperature is stabilized, simulate the worst case failure in the normal use lubrication system. Upon illumination of the low oil pressure warning device (required by § 29.1305), reduce input torque to simulate an autorotation and continue transmission operation for 15 minutes. To complete the test, apply an input torque to the transmission for approximately 10 seconds to simulate a minimum power landing. A successful demonstration may involve limited damage to the transmission, provided it is determined that the autorotative capabilities of the rotorcraft were not significantly impaired.

d. Section 29.927(d) : (1) Explanation . This test is intended to demonstrate that overspeed conditions which may result from control failure or control misapplication will not incur damage to the rotor drive system. Specific conditions for conducting the test are provided in § 29.927(d)(1), (d)(2), and (d)(3).

(2) Procedures . The test may be conducted on a rotorcraft configured for the endurance tests prescribed by § 29.923. Turbine engines involved in the test may require fuel control rerigging or operation on the manual fuel control system, if available, to achieve test requirements.

Note: Some equivalent safety findings have been issued based on limiting the test speed to that permitted by an independent, reliable overspeed trip device, thus avoiding permanent damage to yokes, engines, etc., involved but not subject to evaluation under this rule.

(3) With collective control set for minimum rotor pitch for smooth operation, the cyclic control positioned for vertical lift, and the antitorque control set in flat pitch, add power to achieve 120 percent of maximum continuous speed and hold this condition for 30 seconds. Deceleration and operation between overspeed runs should be as described in the rule. Acceleration and deceleration must be at maximum rates available to the configuration.

e. Section 29.927(e) : (1) Explanation . This paragraph sets forth conditions to be normally employed during the overtorque and overspeed tests of this section and authorizes certain exceptions with criteria for justification.

(2) Procedures . None.

Page E - 55 AC 29-2C, Chg 4 7/6/2012 AC 29.927A. § 29.927 (Amendment 29-26) ADDITIONAL TESTS.

a. Section 29.927(c) : (1) Explanation .

(i) Amendment 29-26 revised the rotor drive system loss of lubrication test requirements for Category A rotorcraft in § 29.927(c). This requires testing to show that any failures that result in a loss of lubrication in any normal use lubrication system, unless the failures are extremely remote, will not prevent continued safe flight for at least 30 minutes after the flight crew recognizes the loss of lubricant failure.

(ii) The introductory phrase to this amendment to the regulation, “unless such failures are extremely remote” has caused confusion. The NPRM did not contain this expression and the only change documented in the preamble to the final rule (53 FR 34204) explains that the final rule was revised in response to a public comment that the proposed regulation could be interpreted to “preclude credit for auxiliary lubrication systems or to require consideration of lubricant failures to self-lubricated bearings.” This was not intended and the final rule was “revised to eliminate this possible ambiguity.” The phrase, “unless such failures are extremely remote,” was introduced to resolve the public comment to convey that the applicant does not have to consider failures that may exist in the auxiliary lubrication system prior to performing the loss of lubrication testing. Under the current regulation, the extremely remote language in the final rule means that testing to demonstrate at least 30 minutes continued flight capability (for Category A), following loss of lubrication in the normal lubrication system, is not required if the failures leading to that loss of lubrication condition are determined to be extremely remote. While this compliance approach is allowed, it may not be achievable due, in part, to the unforeseen variables and complexity associated with predicting potential lubrication failure modes and their associated criticality and frequency of occurrence. This includes considering lubrication failures that may result from improper transmission maintenance and servicing. The expected compliance approach has been to assume a failure in the normal lubrication system leading to rapid loss of lubrication and to rely on an auxiliary lubrication system or the robustness of the transmission components to accomplish at least 30 minutes of operation (for Category A) at the prescribed conditions. With this approach, the normal and auxiliary systems must be independent in order to preclude common loss of lubrication failure points and possible cross contamination. Compliance with § 29.1309 would only apply to any electrical and software design aspects of the normal and auxiliary lubrication systems.

The auxiliary lubrication system must also be designed, constructed, and functionally tested to show that it can perform its intended function.

(iii) The regulation is intended to apply to pressurized lubrication systems and has not been applied to splash lubricated gearboxes since historically their design has not been as critical or complex when compared to pressurized systems. The likelihood of loss of lubrication is significantly greater for transmissions that use pressure lubrication and external cooling. This is due to the increased complexity of the Page E –56 7/6/2012 AC 29-2C, Chg 4 lubrication system and the external components that circulate oil outside the gearbox. A pressure lubrication system is more commonly used in the rotorcraft’s main transmission but may also be used in auxiliary transmissions or gearboxes.

(iv) The lubricating system has two primary functions. The first is to provide lubricating oil to contacting or rubbing surfaces and thus reduce friction losses. The second is to dissipate heat energy generated by friction of meshing gears and bearings, thus maintaining surface and material temperature. Accordingly, a loss of lubrication leads to increased friction between components and increased component surface temperatures. With increased component surface temperatures, component surface hardness can be lost resulting in the inability of the component to carry or transmit loads. Thermal expansion in transmission components can eventually lead to the mechanical failure of bearings, journals, gears, shafts, and clutches that are subjected to high loads and rotational speeds. A loss of lubrication may result from internal and external failures. Failures include, but are not limited to, oil lines, fittings, seal plugs, sealing gaskets, valves, pumps, oil filters, oil coolers, accessory pads, etc. A leak caused by a crack in the transmission outer case need not be considered as a source of a loss of lubrication provided the outer case has been structurally substantiated to satisfy the requirements of §§ 29.307, 29.923(m), and 29.571.

(v) The intent of the rule change for Category A rotorcraft was to assure that these rotorcraft have significant continued flight capability after the loss of lubricant to any single transmission in order to optimize eventual landing opportunities. Extending the bench testing beyond 30 minutes, although not required, is considered highly desirable. Accomplishing this would further improve the capability of the rotorcraft to reach a suitable landing location in order to improve occupant safety when operating in remote geographic areas that include harsh environmental conditions. Indefinite flight with a lubrication system failure is not expected. However, it may be acceptable to include a time interval in the emergency procedures. That time interval should be reduced sufficiently when compared to the bench test demonstration to allow for an adequate safety margin.

(2) Procedures .

(i) Section 29.927(c)(1) prescribes a test to demonstrate that the effects of a loss of lubrication will not prevent continued safe powered operation for category A rotorcraft for at least 30 minutes after illumination of the low oil pressure warning device (required by § 29.1305). For category B rotorcraft, § 29.927(c)(2) prescribes the tests for safe operation under autorotative conditions must continue for at least 15 minutes.

(ii) An acceptable means of demonstrating compliance with this rule is through the use of a bench test (transmission test rig). Since this is essentially a durability test of the transmission to operate with residual oil, typically the worst case failure (i.e., the undrainable oil or the oil remaining after a severe pressure leak, whichever results in a greater loss of oil in the transmission’s normal use lubrication system) is used as a critical entry point for the test; see paragraph a.(2)(iii).

Page E - 56.1 AC 29-2C, Chg 4 7/6/2012 (iii) The transmission should be stabilized at the torque associated with maximum continuous power (reacted as appropriate at the main mast and tail rotor output quills) at a normal main rotor mast speed, oil temperature that is at the highest limit for continuous operation, and oil pressure that is within the normal operating range.

A vertical load should be applied at the mast, equal to the gross weight of the rotorcraft at 1g. Once the transmission oil temperature is stabilized, simulate the worst case failure in the normal use lubrication system. Upon illumination of the low oil pressure warning device (required by § 29.1305), reduce the input torque for category A rotorcraft to the minimum torque necessary to sustain flight and continue the test for at least 30 minutes at the maximum gross weight and the most efficient flight conditions. To complete the test, apply an input torque to the transmission for approximately 25 seconds to simulate an autorotation. The last 10 seconds (of the 25 seconds) should be at the torque required for a minimum power landing. A successful demonstration may involve limited damage to the transmission, provided it is determined that the autorotative capabilities of the rotorcraft were not significantly impaired. For category B rotorcraft, upon illumination of the low oil pressure warning device, reduce the input torque to simulate an autorotation and continue transmission operation for 15 minutes. To complete the test, apply an input torque to the transmission for approximately 10 seconds to simulate a minimum power landing. A successful demonstration may involve limited damage to the transmission provided it is determined that the autorotative capabilities of the rotorcraft were not significantly impaired. If compliance with category A requirements is demonstrated, category B requirements will have been met.

b. Section 29.927(d) : (1) Explanation . The revision to paragraph (d) includes a requirement to demonstrate that overspeed conditions, which may result from an engine control device failure or other event such as a control misapplication, will not result in damage to the rotor drive system.

(2) Procedures . The overspeed endurance cycle and overspeed conditions to be demonstrated are defined in this section. The test may be conducted on a rotorcraft configured for the endurance tests prescribed by § 29.923. Turbine engines involved in the test may require fuel control re-rigging or operation on a manual fuel control system to meet test requirements. Fifty overspeed runs of 30 ±3 seconds must be run on the rotor drive system. The overspeed runs must be alternated with stabilizing runs of 1 to 5 minutes duration each at 60 to 80 percent of maximum continuous speed.

(i) The maximum speed to be demonstrated during the power on overspeed test is: (A) The higher of: Page E –56.2 7/6/2012 AC 29-2C, Chg 4 ( 1 ) The speed to be expected from an engine control device failure; or, ( 2 ) 105 percent of the maximum rotational speed to be expected in service, including transients.

(B) The maximum speed allowed by a speed limiting device if the device is installed independent of the engine controls and is shown to be reliable.

(ii) From the stabilizing run condition, increase power to achieve the maximum speed established from (i) above. Set the collective for minimum blade pitch for smooth operation. The cyclic control should be positioned for vertical lift, and the anti-torque control should be set in flat pitch. Hold this condition for 30 seconds, then decelerate to the stabilizing run condition.

(iii) The acceleration and deceleration described above should be accomplished in 10 seconds or less except where it can be shown that the certified engine acceleration or deceleration rate exceeds 10 seconds. The time required for acceleration and deceleration may not be deducted from the 30 second overspeed period.

Note: Some equivalent safety findings have been issued based upon limiting the test speed to that permitted by an independent, reliable overspeed trip device. This has been done to avoid permanent damage to rotors, yokes, engines, etc., which are involved, but not under evaluation by this test.

c. Section 29.927(f) : (1) Explanation . Amendment 29-26 also added a new paragraph (f), which requires that the overtorque, lubrication system failure, and overspeed tests required by § 29.927(b), (c), and (d), respectively, be conducted without intervening disassembly during the individual test. After each test, a teardown inspection is performed, and except for the components used in the lubrication system failure test, the components are required to be in serviceable (return to service) condition.

(2) Procedures . None.

Page E - 57 AC 29-2C, Chg 4 5/1/2014 AC 29.931. § 29.931 (Amendment 29-12) SHAFTING CRITICAL SPEEDS.

a. Explanation .

(1) At certain speeds, rotating shafts tend to vibrate violently in a transverse direction. These speeds are variously known as “critical speeds,” “whirling speeds,” or [Section AC 27.931 continued on next page.]

Page E –58 9/30/99 AC 29-2C “whipping speeds.” The vibration results from the unbalance of the rotating system and can be shown to reach destructive values with only minimal unbalance. The nature of this phenomena is that as shaft rotational speed increases, residual unbalance in the shaft gives rise to centrifugal forces. These forces cause the shaft to rotate in a bent or bowed configuration with the centrifugal force induced bending loads being balanced by coriolis and elastic forces in the shaft. As shaft rotational speed increases, the centrifugal forces increase to the point at which they exceed the elastic forces in the shaft, and divergence occurs. This point in the speed range is called the critical speed.

At shaft speeds above the critical speed, a 180° phase change occurs; the shaft’s mass center moves toward the center of rotation and the amplitude of vibration diminishes with further increases in shaft speed.

(2) The most prominent design option is to operate the shafting subcritical; i.e., below the first critical speed, with adequate margins from critical speed at the maximum allowable speed, including transients. However, another option, that of supercritical shaft operation; i.e., operating above the first or even higher critical speeds with adequate margins between any critical speed for the normal operating speed range.

This latter portion requires some form of fixed system damping to permit safe transition through the critical speed range and to avoid excessive nonsyncronous vibrations or instability in the critical speed mode at suboperating frequency.

(3) A review of typical design practices and drive system arrangements discloses several types of shaft support and loading: (i) Main rotor/mast/transmission assemblies rigidly mounted to the airframe; (ii) Main rotor/mast/transmission assemblies compliantly mounted to the airframe; (iii) Main rotor supported through a bearing arrangement by a rigid nonrotating structure with a coaxial torque shaft driving the rotor; (iv) Cross-shafting, interconnect shafting, tail rotor drive shafting which are generally supported by gearboxes at each end and by hanger bearings at semispan; (v) Engine to transmission shafting which, for compliant pylons, incorporate flexible or geared coupling, to accommodate the misalignment and chucking; and (vi) Tail rotor/mast/gearbox supported on the tailboom or near the upper extremity of a vertical fin.

(4) With regard to compliant pylon mountings, recent developments in vibration control have led to rotor isolation wherein the fuselage is isolated from the rotor and Page E - 59 AC 29-2C 9/30/99 transmission, resulting in improved vibration and system reliability. Rotor isolation systems typically entail the installation of isolation devices at the transmission-airframe interface. The crux of rotor isolation is providing adequate, low-frequency isolation without excessive relative displacement or loss of mechanical stability. Rotor isolation affects shaft critical speeds in the following ways: (i) First, the transmission mounting configuration, system stiffness, and tuning requirements may result in different fore-and-aft and lateral natural frequencies, imposing additional analytical requirements. For compliant mounting, the response while transitioning through the fundamental or rocking modes is generally controlled by dampers or elastomeric elements.

(ii) Second, the relatively high displacements permitted by the isolation system, depending on configuration, may result in variations in shaft misalignment and length thus adding further complexity to the analytical prediction of critical speeds.

b. Procedures .

(1) Subcritical Shafting Designs . Three basic methods of qualification may be considered, with the required margins relative to the degree of assurance provided: (i) Analytical .

(A) Simplistic model(s) as shown in figures AC 29.931-1 and AC 29.931-2; 35 percent margin shown above maximum operating speed.

(B) Detailed model, taking into account significant variations in shaft stiffness, mass distribution, cone adapters, support bearing stiffness, support structure; 20 percent margin shown above maximum operating speed.

(ii) Analytical supported by tests . Analysis supported by shake test (rotating or nonrotating) or by bench test, where appropriate adjustments are made for differences between the bench and the aircraft; 15 percent margin shown above maximum operating speed.

(iii) Whirl test on the aircraft .

(A) For all cases, it should be shown that, under maximum permissible unbalance and at the maximum operating speed, the shafting and support structure has acceptable clearance and does not have excessive vibration.

(B) For compliant pylon mountings, damping of the rigid body rocking modes, which are often transitioned during run-up to normal speed (and which are not critical flexing modes), may be verified by analysis, laboratory tests, or ground run-up with the rotor at maximum permissible unbalance. Damping on the order of 5 percent equivalent viscous damping is generally acceptable.

Page E - 60 9/30/99 AC 29-2C (C) For tail rotor masts, the analysis should include fixed system structural response including tailboom, fixed control surfaces, and vertical fin. The frequency analysis will then contain both fixed system and rotating system modes. An energy analysis can then be used to identify whether the modes are predominantly fixed system or rotating system modes. Systems with up to 35 percent energy in the rotating system have been operated in the field without significant problems. For this type of shafting installations, it is advisable to avoid fixed system modes at multiples of shaft speed, particularly where highly nonisotropic mountings exist.

(2) Supercritical Shafting Design . Another facet occasionally encountered with shafting is the concept of normally operating at speeds above the critical speed, commonly referred to as “supercritical operation.” To function properly, suitable dampers must be installed to enable the shaft to pass safely through the lower critical speed up to the operating speed, and speed controls should be devised so as to avoid any tendency to operate continuously at any critical speed. Accurate balancing of the rotating components will also decrease the energy to be dissipated into the damping device during transition thereby increasing its serviceability and reliability. It should be noted that damper design and locations become more complex as selected operating speed increases through the third or fourth critical frequency. Multiple node points will exist where dampers will not be effective. Production specimen testing at high speed/high torque conditions should include checks for shaft straightness until experience verifies that shaft deflecting is not significant. For system utilizing squeeze film dampers at the support bearings, variations in oil pressure, flow restrictions, and the effects of bearing preload should be evaluated. The effects of shaft and unbalance and the proximity of the damper to bottoming under maximum unbalance should be evaluated.

(3) If the shafting configuration of the rotorcraft includes universal joints or misalignment couplings, a velocity differential will exist across the joint which creates sinusoidal torques and bending moments at both shafts at multiples of the rotation speed. To avoid amplification of these torques and bending moments, the design should preclude coincidence of critical speeds and multiples of normal speeds.

(4) Note that failure considerations required under § 29.901(d) may result in abnormal rotational speed and torque excursions. Resulting encounters with critical speeds should not create hazards.

(5) Order 8110.9, Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopters and Other Power Transmission Systems, also addresses this subject. This document is distributed to section level and above in all Regional Aircraft Certification Offices.

Page E - 61 AC 29-2C 9/30/99 k M + 0.23m Wcr = first critical speed, RAD/SEC k = shaft spring rate, LB/IN = 3EI/L

/] ◄ 41--- L ---•I

E = m odu I us of elasticity ...__ -- r = moment on inertia ----::_-..1'

- - ~

-, I ...., M = mass of weight, LB-SEC /IN m = mass of shaft, LB-SEC /IN Fl GURE AC 29.931-1. CANTILEVERED SHAFT, Fl RST CRITICAL SPEED wcr = a/,~

~ ~

W , = first critical speed-RAD/SEC u = mass per unit Iength E = Young's modulus L = length between supports = inertia of shaft a = a numerical constant for first critical speed, a= (pi) = 9.87 The numerical constant (a) for higher critical whirl modes or other shaft support systems may be derived from standard texts on this subject FIGURE AC 29.931-2. SHAFT BETWEEN SUPPORT BEARINGS, FIRST CRITICAL SPEED Page E - 62 5/1/2014 AC 29-2C, Chg 4 AC 29.935. § 29.935 SHAFTING JOINTS.

a. Explanation . This rule requires the design of shafting joints to include provisions for lubrication when such lubrication is necessary for operation.

b. Procedures . Review the design of the rotor drive system for universal joints, slip joints (splines), and other shaft couplings. Lubrication access points (Zerk fittings) should be required unless the design incorporates alternate provisions for lubrication acceptable to the FAA/AUTHORITY.

AC 29.939. § 29.939 (Amendment 29-12) TURBINE ENGINE OPERATING CHARACTERISTICS.

a. Explanation . This section provides guidance for evaluation of engine operation, engine inlet airflow distortion, and engine to drive system torsional stability. A satisfactory rotorcraft design for all three items should be established by the manufacturer early in the development program since changes in design to satisfy these requirements are typically very expensive and will adversely impact other basic design features. Introduction of full authority digital engine control (FADEC) controls has increased the complexity in evaluating engine operating characteristics with the need to investigate FADEC degraded or failure modes. The certification test plan should address the engine control being used. In addition, where manufacturing tolerances could affect engine handling and rotor governing, tests should be performed considering the worst tolerances regarding the tests to be conducted. The results of these evaluations are used in part to verify that the requirements for the engine installation Instruction Manual mandated by § 33.5(a) are satisfied.

b. Procedures .

(1) Turbine engine operation .

(i) Explanation . Smooth, stable operation of turbine engines is essential to safety and control of rotorcraft. This can be adversely affected by rotorcraft maneuvers, turbulence, high altitude, temperature, airspeed, and installation features such as the engine air inlet duct, exhaust duct, and the location with respect to other airframe items which induce or influence air flow through the engine. Powerplant control displacement rate can also be a factor, although most modern engines incorporate internal protection for this aspect. The engine’s tolerance to these factors is reflected as the “stall margin” which is established by the engine manufacturer through design and test. However, this stall margin is applicable only to an engine with a specified inlet and exhaust and at specified altitude, temperature, and effective airspeed. Typically, the specified engine inlet duct is a symmetrical bellmouth and the exhaust is a short straight duct of specified diameter and length. The stall margin, even under the above test conditions, usually varies with engine power, acceleration or deceleration, compressor air bleed, and accessory power extraction.

Page E - 63 AC 29-2C, Chg 4 5/1/2014 (ii) Procedures . The official flight test plan should include requirements to investigate the engine operating characteristics for stall, surge, flameout, acceleration and deceleration response, and transient response (within approved limits) throughout the operating range of the rotorcraft. The results must show that no adverse characteristics are present, to a hazardous degree, during normal and emergency operation within the range of operating limitations of the rotorcraft and engine. Test configurations should encompass the critical engine power settings and design or rigging tolerances expected to be seen in service. In addition, normal and degraded engine operating control modes should be evaluated where applicable. Test conditions should include maximum airspeed-sideslip combinations, power recoveries, hover with wind from all directions including tailwinds and other maneuvers appropriate to the certificated operating envelope of the rotorcraft. In addition, recirculation of exhaust gases during hover or rearward flight can be critical for engine operation and should be evaluated. Particular attention should be given to flight and operating conditions that can be judged critical from review of data on engine inlet pressure and temperature distribution patterns and engine stall margin data if available. High altitude has typically been critical for these tests unless other critical areas of the flight envelope are identified. In addition, during rearward flight at high altitude, results may indicate unacceptable thermal distortions in the inlet due to reingestion. Stall, surge, or flameout which may be hazardous (i.e., causes loss of engine function, loss of control, severe torsional shock through the rotor drive system, or otherwise damages the rotorcraft) is unacceptable. The flight test program should include: (A) Normal operation (hover, forward stabilized flight, collective inputs, rearward flight).

( 1 ) Checks in hover. Hover with wind from all azimuths to the allowable wind limits, including tailwinds, should be evaluated. This evaluation should include maneuvers with rapid power changes. Recirculation of exhaust gases can be critical. Particular attention should be given to flight and operating conditions that can be judged critical from review of data on engine inlet pressure and temperature distribution patterns and engine stall margin data. Operating at high density altitudes, especially engine operation during low-speed rearward flight, is more likely to result in unacceptable engine inlet thermal distortion due to reingestion of exhaust gas.

Behavior of the compressor control bleed air valves, if any, must be carefully checked by collective oscillations around the shut off or open bleed air valves operating points.

To mitigate the possible risk of power loss, a safe test build up should be considered, such as a build up in actual wind conditions before progressing to rearward or sideward flight.

( 2 ) Checks in forward stabilized flight.

(i) Behavior of the engine must be checked during level flight, climb, and descent at various power settings in order to verify the: Page E - 64 5/1/2014 AC 29-2C, Chg 4 (A) Governing stability (e.g., absence of engine parameters oscillations).

(B) Variation of rotor speed with the requested power.

(C) Engine matching on multi-engine helicopters.

(D) Accuracy of the engine parameters available to the pilot.

(E) Opening and closing thresholds of the bleed valve, if any.

(F) Effect of sideslip.

(ii) Stabilized flight conditions should be long enough to allow the engine to reach its thermal stabilization (typically 2 or 3 minutes) with the aim of: ( A ) Assessing that the engine cycle (TOT and turbine inlet temperature) as installed is consistent with the engine certificated type design.

( B ) Evaluating installation effects.

( 3 ) Checks of engine behavior during collective inputs. Before conducting the below collective increase inputs and collective decrease inputs tests, the tested engines should be precisely defined, especially the gas generator turbine nozzle and free turbine nozzle sections (key point for the stall characteristics). Additionally, the acceleration controllers’ settings (key points for acceleration and deceleration) should be identified.

(i) Collective increase inputs.

( A ) The collective increase inputs will verify: ( a ) The transient behavior of the engine, the acceleration, and the minimum N speed achieved.

R ( b ) The values that the engines could reach in transients depending of the collective inputs.

( c ) The effects of the collective anticipation, if any.

( d ) The rotorcraft handling qualities (amplitude of variation of pitch, roll and yaw, cross coupling, etc.) during the transients.

( B ) The flight test techniques consist of increasing collective R from the limit of desynchronization (N2/N needles are still matched) up to the collective pitch corresponding to the MCP. Depending on the altitude (speed of collective input Page E – 64.1 AC 29-2C, Chg 4 5/1/2014 should be reduced when altitude increases) and the rotorcraft category (speed of collective input should be reduced when rotorcraft max weight increases), the duration of the maneuver should be from 1 to 3 seconds at low altitude to approximately 3 to 5 seconds at higher altitude.

( C ) Some collective inputs should be made from desynchronized conditions (N > N2) but should be carefully done as those maneuvers R generally lead to larger reductions of N and high free wheel or engine constraints.

R ( D ) On multi-engine helicopters, collective inputs at lower speeds should be done with one engine at idle position to check the behavior of the other engines in case of an engine failure.

( E ) Droop of N2/N out of the green arc is permitted during the R test, provided it is acceptable to the engine or airframe manufacturer for the purpose of the test. The goal is to check for generally satisfactory acceleration (no surge, no extreme N2/N droop, no significant overshoot). Any N2/N droop must not result in a R R condition that requires exceptional piloting skill, alertness, or strength to maintain safe flight.

( F ) During those collective inputs, Nr below the minimum Nr should be avoided by lowering the collective. The reasons for poor engine acceleration should be determined.

( ii ) Collective decrease inputs.

( A ) Collective decrease inputs will verify: ( a ) Combustion stability (i.e. that there is no flame-out).

( b ) The engine behavior in transients and in particular the maximum free turbine speed, minimum oil pressure, etc.

( c ) The rotorcraft handling qualities (variation of pitch and roll and yaw, cross coupling, etc.) during the transients.

( B ) Rate of collective lowering should be built up from 5 seconds down to under 3 seconds. Regarding N2 overspeed, a quick stop type maneuver that leads to an initial acceleration of the N should be envisaged.

R ( 4 ) Compressor stall investigation.

(i) The turbine engine installation should not suffer from compressor stalls anywhere in the flight envelope. Previous tests as described above include some of the required areas, but additional testing is necessary. In stabilized conditions, the collective pitch is moved in slow and small oscillations in: Page E – 64.2 5/1/2014 AC 29-2C, Chg 4 ( A ) Level flight with or without sideslip at different power settings.

( B ) Pull up or in turn.

( C ) Autorotation.

( D ) Descent at different rates.

( ii ) The maneuvers described in paragraphs b.(1)(ii)(A)(4)(i)(A) through (D) above check the engine susceptibility to air flow distortions. The test technique involves oscillations of the collective at various rates around the identified thresholds of the compressor control bleed valves.

(B) Degraded modes: Governor failures.

( 1 ) The specific degraded mode testing for a hydromechanically controlled engine is different than for a digitally controlled engine. However, the basic principles for both types of engine controls include: (i) Review of the safety analysis, especially the FMEA.

( ii ) Determine the failures to be checked in flight, based on prior analysis or testing.

( iii ) For each failure to be checked in flight: ( A ) Evaluate the behavior of the engine when the failure occurs.

( B ) Determine the acceptability of the procedure of identification of the degraded engine.

( C ) Assess the human factor aspects of the machine interface (cautions, warning, etc.).

( D ) During the use of the degraded engine, consider the workload impact related to engine response in adjusting NR or power for approach and landing.

( E ) Evaluate the adequacy of the RFM procedures.

( 2 ) Helicopters equipped with FADECs are becoming increasingly integrated into the helicopter systems for engine control sensor inputs. For example, N2/N values available could depend on parameters such as airspeed, altitude, or R temperature as provided by the basic helicopter systems and sent to the FADEC. A Page E – 64.3 AC 29-2C, Chg 4 5/1/2014 concern is the effects of degradation, discrepancies or failures in the basic helicopter systems on engine governing laws and power available. Therefore, a review of the FMEA before beginning engine failure or degraded mode evaluations should be performed.

( 3 ) It is recommended to have special test equipment installed on the test engine in order to simulate failures to the FADEC or digital engine control. Test risk mitigation should include the evaluation of possible malfunctions of this special test equipment.

( 4 ) As a minimum, the following typical failures or degraded modes, including engine control reversion that is based on safety analysis and ground test results, should be evaluated in flight: (i) Failure causing the engine to increase power.

( iii ) Failure causing the engine to decrease power.

( iii ) Failure fixing the engine power at a static value.

( iv ) Failure introducing oscillations.

( v ) Failure introducing lower acceleration or deceleration.

( vi ) Failure causing the N2/N to increase or decrease.

R (2) Vibration .

(i) Explanation . Engine airflow patterns are deflected or distorted by the presence of airframe inlet hardware, cowling, fuselage panels, and, to a degree, in almost all flight regimes. Additional items such as airframe installed particle separators, deflectors for snow, ice, or sand protection, and obstructions forward of the engine inlet, such as a hoist kit, could affect the engine air flow patterns. The rotating elements of the engine, particularly the compressor blades, will be subjected to a cyclically varying air flow as these elements move into and out of areas of deflected airflow to the engine.

A corresponding aerodynamic load will be imposed on these engine elements. Since this loading is also cyclic, the possibility of critical frequency coupling with an engine component shall be investigated.

(ii) Procedure . Typically, this evaluation would involve installation in the engine inlet of a special multiple probe, total pressure sensing system, and flight testing which largely follows that prescribed for evaluation of engine operating characteristics as described above. Data from these tests can be reduced to create a pressure map at the compressor inlet face which, in conjunction with compressor speeds, may be used to determine the frequencies and relative amplitudes of the cyclic air loading imposed on the engine compressor blades. The engine manufacturer either supplies the sensing Page E – 64.4 5/1/2014 AC 29-2C, Chg 4 probe or specifies its design and performance. Also, the engine manufacturer may evaluate the test results or publish acceptance criteria. A wave analysis may be involved in identifying higher order excitations. Engine exhaust ducts which include bends, noise suppressors, or other obstructions may require an evaluation similar to that discussed above for the engine inlet. The engine manufacturer should be consulted for instructions or approval of this aspect. High performance engines may also require an engine inlet temperature survey. Details of instrumentation and acceptance criteria should be provided by the engine manufacturer. Engines equipped with only centrifugal compressors are less likely to encounter frequency coupling and may not require this investigation. The engine manufacturer’s recommendations should be followed in these cases.

(3) Torsional Stability .

(i) Explanation . Governor-controlled engines installed in rotorcraft are subject to a fuel control resonant feedback condition which could be divergent if not properly designed or compensated. This condition occurs when the response frequency of the governor on the engine is coincident with or close to a low order natural torsional frequency of the rotorcraft’s rotor drive system. Typically, these frequencies appear in the 3 to 5 cycles per second (CPS) range. The manufacturer usually resolves torsional instability problems by introducing damping into the engine governor or fuel control.

Provisions for this change must be supplied by or approved by the engine manufacturer.

The final configuration may be a compromise between a lightly damped control, which will allow a positive but slow convergence of drive system torsional oscillations, and a highly damped control which exhibits excessive rotor speed droop or overspeed following rotorcraft collective control displacement.

(ii) Procedures . A ground and flight test program should be devised to evaluate the torsional response of the engine and drive system combination presented by the applicant. Instrumentation to record drive system torsionals should be applied to all major branches of the drive system. Engine parameters such as torque, RPM fuel manifold or nozzle pressure, compressor discharge pressure, and governor lever position should be recorded simultaneously with drive system parameters. The test program should include ground tie-down operation and flight operation across a range of engine power and rotor speeds while injecting control inputs as close to the first order drive system natural frequency as possible. Mechanical methods of making these inputs are not usually necessary if the desired frequency is in the 3 to 5 CPS range and the instrumentation readout confirms that the drive system was actually excited torsionally at its natural frequency. Control inputs should include collective, antitorque, and throttle. Also, cyclic inputs may be important on tandem rotor rotorcraft. The acceptance criteria may be dependent on several items. Among these are rotor and drive system fatigue loading, engine power response characteristics, limitations established by the engine manufacturer, etc. The acceptance criteria are usually stated as a percent damping (minimum). Typically, 1 percent of critical equivalent viscous damping (or greater) is acceptable. In effect, this means that the free vibration response to a control input damps to ½ amplitude in 11 cycles or less.

Page E - 65 AC 29-2C 9/30/99 SUBPART E - POWERPLANT FUEL SYSTEM AC 29.951. § 29.951 (Amendment 29-12) FUEL SYSTEM - GENERAL.

a. Explanation .

(1) The term “fuel system” means a system which includes all components required to deliver fuel from the tank(s) to the engine(s). This includes, but is not limited to, all components provided to contain, convey, drain, filter, shutoff, pump, jettison, meter, and distribute fuel to the engines.

(2) Paragraph (a) of this section is a general statement of the performance requirements for fuel systems and constitutes authority to require fuel systems to be adequate notwithstanding compliance with detail requirements listed in §§ 29.953 through 29.999 of this subpart.

(3) Paragraph (b) of this section requires fuel systems to be designed so that air will not enter the system under any operating conditions by either arranging the system so that no fuel pump can draw fuel from more than one tank or by other acceptable means.

(4) Paragraph (c) of this section sets forth a fuel system performance requirement intended to ensure that ice to be expected in fuel when operating in cold weather will not prevent the fuel system from supplying adequate fuel to the engines.

Although fuel system filters and strainers are the items in the fuel system most susceptible to clogging from ice particles in the fuel, this paragraph requires that the entire fuel system be shown to be capable of delivering fuel, initially contaminated with ice, to the engine(s).

b. Procedures .

(1) For paragraph (a), the applicant should show compliance with the fuel system requirements of this subpart, except that if unusual fuel system arrangements or requirements exist which are not adequately addressed by these subparts, this paragraph may be used as authority to require special tests, analysis, or system performance needed for proper engine functioning.

(2) For paragraph (b), review the fuel system design with special attention to fuel tank selector valves, crossfeed systems, and multiple tank outlet arrangements to ensure that no allowable fuel system configuration will permit air to enter the system.

For questionable situations, the applicant should conduct ground or flight tests, as necessary, to verify compliance with this section.

Page E - 66 9/30/99 AC 29-2C (3) Paragraph (c) provides for sustained satisfactory operation of the fuel system with cold fuel initially contaminated with water. Since ice in the fuel system is not considered to be an emergency condition but, rather, is an expected service encounter, compliance would not involve the imposition of special rotorcraft limitations.

Flight manual instructions such as land as soon as practicable, reduce altitude to some value less than otherwise permitted, reduce power, turn on boost pumps, etc., are not appropriate in demonstrating compliance. Some methods of fuel system ice protection which have been used to show compliance follow.

(i) Fuel heater . Usually these devices are fuel-to-engine oil heat exchangers and are normally located to protect the fuel filter from blockage by ice in the fuel. The adequacy of these devices should be established. Usually this involves generation of a heat balance between heat gained by fuel and heat lost by oil using performance data provided by the manufacturers of the fuel-oil heater, the oil cooler, the heat rejected by the engine to the oil, etc. A minimum oil temperature associated with the adequacy of the fuel heater may need to be established, marked on the oil temperature gauge, and verified to be maintained during critical flight conditions. Other unprotected parts of the fuel system remain to be evaluated and substantiated for compliance with this requirement.

(ii) Oversized fuel filter . This method may only substantiate the fuel filter and, as with the fuel heater method, is incomplete without evaluation of the remainder of the fuel system. An icing test of the filter should be accomplished. Fuel preparation procedures and method of testing should follow the applicable portion of SAE Aerospace Recommended Practice (ARP) No. 1401. A satisfactory configuration is achieved when a filter is demonstrated to have the capacity to continue to provide the filtration function, without bypassing, when subjected to fuel contaminated by ice to the degree required by this rule. Usually, a delta pressure caution signal for the filter is needed to alert the flightcrew that progressive filter blockage is in progress. The caution device setting should be established by test which demonstrates that after illumination of the caution signal sufficient filter capacity exists to enable completion of the flight.

Fuel pressure should not fall below established limits because of ice accumulation on the filter.

(iii) Anti-ice additives . This method utilizes the properties of ethylene glycol to reduce the freezing temperature of water in the fuel. It has the advantage over other methods of protecting all components in the fuel system from ice blockage.

Compliance with the rule by this method involves the following.

(A) Eligible additives . PFA-55MB (Phillips Petroleum Co.) and additives per specification MIL-I-27868, Revision D, or earlier. Later versions of this specification do not require glycerin, which may be needed to protect fuel tank coatings.

(B) Compatibility . Both engine fuel system and aircraft fuel system should be verified to be chemically compatible with the additive at the maximum concentration to be expected in the fuel system. Usually, information on eligible system materials can Page E - 67 AC 29-2C 9/30/99 be obtained from the engine manufacturer for the engine fuel system and from the additive manufacturer for aircraft fuel system materials.

(C) Adding or blending the additive to the fuel . These additives do not mix well with the fuel and indiscriminate dumping of additive into the tank will not only fail to protect the system from ice accumulation but likely will damage nonmetallic components in the system. Some fuels may have additive premixed in the fuel. If other fuels are to be eligible, a method for blending additive into the fuel during refueling must be devised and demonstrated to be effective.

(D) Placards should be added near the fuel filler opening to note that fuel must contain the approved anti-ice additive within the minimum and maximum allowed concentration.

(E) The FAA/AUTHORITY-approved flight manual should contain necessary information to attain satisfactory blending of the additive and procedures to allow the operator to check the blend in the fuel tank.

(iv) Fuel system protection (other than filters) . If the fuel heater method or oversize filter method (paragraphs b(3)(i) and b(3)(ii)) is proposed, the remainder of the fuel system should be shown to be free from obstruction by fuel ice. This may be shown by testing the system with ice contaminated fuel (prepared as suggested for filter tests) or, in many cases, by selecting fuel system components which by test or by previous experience are known to be free of ice collection tendencies. Tank outlet screens (or tank-mounted pump inlet screens) may be the significant fuel system feature for further evaluation. In some instances, fuel turbulence due to pump motions may be sufficient to keep the screen clear of ice. In other instances, small screen bypass openings (approximately one-fourth inch in diameter) located outside the predominant fuel flow path have been found satisfactory.

NOTE: Advisory Circular (AC) 20-29 contains information regarding compliance with the fuel ice protection requirements of Part 25. The information in this AC is largely valid except for references to the quantity of water to be expected in fuel and the amount of additive required to ensure freedom from fuel ice hazards.

AC 29.952. § 29.952 (Amendment 29-35) FUEL SYSTEM CRASH RESISTANCE .

a. Explanation .

(1) Section 29.952 provides safety standards that minimize postcrash fire (PCF) in a survivable impact. The rule contains comprehensive crash resistant fuel system (CRFS) design and test criteria that significantly minimize fuel leaks, creation of potential ignition sources, and the occurrence of PCF. Section 29.952 accomplishes this for survivable impacts by- Page E - 68 2/12/03 AC 29-2C, Chg 1 (i) Providing comprehensive criteria to minimize fuel leaks and potential ignition sources; (ii) Requiring increased crash load factors for fuel cells in and behind occupied areas to ensure the static, ultimate strength necessary for impact energy absorption, structural integrity, fuel containment, and occupant safety; (iii) Maintaining the load factors of § 29.561 for fuel cells in other areas (particularly underfloor cells) to ensure leak-tight fuel cell deformation in energy absorbing underfloor structure without unduly crushing or penetrating the occupiable volume; and (iv) Requiring a 50 ft. dynamic vertical impact (drop) test to measure fuel tank structural and fuel containment integrity.

(2) Section 29.952 applies to all fuel systems (including auxiliary propulsion unit (APU) systems).

(3) Some similarities exist among the fire protection requirements of §§ 29.863, 29.1337(a)(2), and 29.952. The requirements in each standard are not mutually exclusive. Overlapping requirements should be certified simultaneously.

(4) The use of bladders is not mandated as this would unduly dictate design.

However, in the majority of cases, their use is necessary to meet the test requirements of § 29.952. If a design does not use bladders, the application should be treated as a new and unusual design feature that should be thoroughly coordinated with the Airworthiness Authority for technical policy to insure adequate safety. Experience has shown that bladders with wall thicknesses from 0.03 to 0.018 inches typically meet the § 29.952 test requirements.

b. Related Material . Documents shown below may be obtained from The Naval Publications and Forms Center, 5801 Tabor Avenue, Philadelphia, Pennsylvania 19120-5094, ATTN: Customer Service (NPODS).

(1) Military Specification, MIL-T-27422B, Amendment 1, April 13, 1971, Tank, Fuel, Crash-resistant Aircraft . (canceled 6-7-97 without replacement) (2) Military Standard, MIL-STD-1290 (AV), January 25, 1974, Light Fixed and Rotary Wing Aircraft Crashworthiness.

(3) Military Standard, MIL-H-83796, August 1, 1974, Hose Assembly, Rubber, Lightweight, Medium Pressure, General Specification for.

(4) Military Specification, MIL-V-27393 (USAF), July 12, 1960, Valve, Safety, Fuel Cell Fitting, Crash Resistant, General Specification, for.

Page E - 69 AC 29-2C 9/30/99 (5) Military Specification, MIL-H-25579 (USAF).

(6) Military Specification, MIL-H-38360.

(7) U.S. Army Publication USARTL-TR-79-22E, “Aircraft Crash Survival Design Guide, Volume V---Aircraft Postcrash Survival”, dated January 1989.

NOTE : Section 4, “Postcrash Fire Protection” of Volume V of the Design Guide is the modern update to MIL-STD-1290. Section 4 contains a comprehensive design guide for military CRFS designs that may be useful for civil CRFS designs.

c. Conceptual Definitions .

(1) Survivable Impact . An impact (crash) where human tolerance acceleration limits are not exceeded in any of the principal rotorcraft axes, where the structure and structural volume surrounding occupants are sufficiently intact during and after impact to constitute a livable volume and permit survival, and where an item of mass does not become unrestrained and create an occupant hazard. “Livable volume” relates to the ability of an airframe to maintain a protective shell around occupants during a crash and to minimize threats, such as accelerations, applied to the occupiable portion of the aircraft during otherwise survivable impacts. In lieu of a more rational, approved criteria, the load factors of § 29.952(b)(1) constitute the structural human survivability accelerations limits.

(2) Postcrash Fire (PCF) . A fire occurring immediately after and as a direct result of an impact. The fire is either the result of fuel released from a leaking fuel system reaching an existing or a crash-induced ignition source, a crash-induced ignition source internal to an undamaged or damaged fuel system, or a combination. PCF's have an intensity range from the minimum of a small local flame to the maximum of an instantaneous massive fire or fireball (explosion).

(3) Fuel Tank or Cell . A reservoir that contains fuel and may consist of a hard shell (of a composite, metal, or hybrid construction) with either a laced-in, snapped in, or otherwise attached semirigid or flexible rubber matrix bladder (or liner), spray-on bladder, or no bladder. The hard shell may be either the airframe (integral tank) or a separate rigid tank attached to the airframe. The device has inlets and outlets for fuel transfer and internal pressure control.

(4) Ignition Source . An ignition source that when wet with fuel or in contact with fuel vapor would cause a PCF.

(5) Major Fuel System Component . A fuel system part with enough mass, installation location hazard or a combination to be structurally considered in a crash.

Structural consideration is required when crash-induced relative motion can occur between the part and its surrounding structure from inertial impact forces, airframe deformation forces, or for other reasons.

Page E - 70 9/30/99 AC 29-2C (6) Drip Fence . A physical barrier that interrupts liquid flow on the underside of a surface, such as a fuel cell, and allows it to drip nonhazardously to an external drain.

(7) Flow Diverter . A physical barrier that interrupts or diverts the flow of a liquid.

(8) Frangible Attachment or Fitting . An attachment or fitting containing a part that is designed and constructed to fail at a predetermined location and load.

(9) Deformable Attachment or Fitting . An attachment or fitting containing a part that is designed and constructed to deform at a predetermined location and load to a predetermined final configuration.

(10) Self-Sealing Breakaway Fuel Fitting . A fuel-carrying in-line, line-to-firewall, bulkhead or line-to-tank connection that breaks in half and self-seals when subjected to forces greater than or equal to the unit’s design breakaway force.

Each half self-seals using a spring-loaded valve (e.g., trap door or equivalent means) that is normally open but is released and closed upon fitting separation. Fitting breakaway force is typically controlled by a frangible metal ring (or series of circumferential tabs) that connects the two fitting halves. Normal, fuel-tight integrity is maintained by “O” rings held under pressure by the rigid, frangible connecting ring (or tabs). When broken open, a small amount of fuel (usually less than 8 ounces) is released. This is the fuel trapped in the coupling space between the two spring-loaded valves. Once failed each coupling half may leak slightly. Typically, this leak rate should be less than 5 drops per minute per coupling half.

(11) Crash Resistant Flexible Fuel Cell Bladder . Flexible, rubberized material, usually with fibers (i.e., rubber “resin” and natural or synthetic fiber) in both the 0° (warp) and 90° (fill) directions that is used as a liner in a rigid shell or integral tank. The material acts as a membrane because, when unsupported, it can only carry pure tension loads. Therefore, it must be uniformly supported by rigid structure (reference § 29.967) so that the liner carries only compressive fluid loads and the surrounding shell structure carries the fluid-induced shear, tension, and bending loads transmitted through the liner or bladder. The material is usually secured (e.g., laced, snapped, etc.) into its surrounding structure at key locations to maintain its intended conformal shape. In many designs, lightweight spacers, such as structural foam, are used between the liner and the airframe to maintain the liners intended conformal shape and to transmit fluid loads to the airframe. The material is either qualified under TSO-C80, “Flexible Fuel and Oil Cell Material,” or qualified during certification.

Sections 29.952 and 29.963(b) have increased the minimum puncture resistance qualification requirement for liner material (See TSO-C80, paragraph 16.0) from 15 to 370 pounds.

Page E - 71 AC 29-2C 9/30/99 (12) Crash Resistant Fuel System (CRFS) . A fuel system designed and approved in accordance with § 29.952 that either prevents a PCF or delays the start of a severe PCF long enough to allow escape.

(13) As Far as Practicable . “As Far as Practicable” means that within the major constraints of the applicant’s design (e.g., aerodynamic shape, space, volume, major structural relocation, etc.), this standard’s criteria should be met. The level of practicability is much higher in a new design project than in a modification project. The engineering decisions, evaluations, and trade studies that determine the maximum level of practicability should be documented and approved.

(14) Fireproof . Defined in § 1.1, “General Definitions” and in AC 20-135, “Powerplant Installation and Propulsion System Component Fire Protection Test Methods, Standards and Criteria” dated February 6, 1990.

d. Procedures .

(1) Section 29.952 should be applied to all fuel system installations. Any major design change should be reevaluated for compliance with the CRFS requirements. It should be noted that most standard materials and processes are acceptable for crash resistant fuel system construction; however, magnesium, magnesium alloys, and cadmium plated parts (when exposed to fuel) are not recommended, because of their inherent ability to create or contribute to a post crash fire. Section 29.952(a) requires each tank, or the most critical tank (if clearly identified by rational analysis) to be drop tested. The tank is filled 80 percent with water and the remaining 20 percent is filled with air (or, in the case of a flexible fuel cell, the air may be evacuated by hand and the cell resealed). The tank openings, except for the vents, are closed with plugs (or other suitable means) so that they remain watertight. The vents are left open to simulate natural venting. Otherwise, the tank is flight configured. The test tanks are installed in their surrounding structure and dropped from a height of 50 feet on a nondeformable surface (e.g., concrete or equivalent). To be considered a valid test, the tank must impact horizontally ±10°. The 50-foot distance is measured between the nondeformable surface and the bottom of the tank. The ±10° attitude requirement can be ensured by using lightweight cord or a light sling to balance the tank assembly horizontally prior to being dropped. MIL-T-27422B shows a typical test setup.

Tank attitude at impact should be verified by photography or equivalent means. The nondeformable floor surface should be covered by a thin plastic sheet so that any leakage is readily detected. The tank water should be tinted with dye to make leakage and seepage sources easy to identify. The tank (except for the vent openings) should be wrapped in light plastic sheet to ensure that minor leakage or seepage (and its source) is detected. Minor spillage through the open vents during the drop test is allowed. The dye should not significantly affect the water’s viscosity or other physical properties that may reduce or eliminate any leakage from the drop test. The nondeforming drop test surface should be carefully reviewed. Concrete is acceptable.

A fixed and uniformly supported steel plate (loaded only in uniform compression without any springback) is acceptable. Floors or floor coverings such as dirt, clay, wood, or Page E - 72 9/30/99 AC 29-2C sand are not acceptable. Selection of the critical fuel tank is important. Factors such as size, fuel cell design and construction, and material(s) should be accounted for when selecting the critical tank. The applicant may elect to drop only a bare fuel cell, not a surrounding structural airframe segment with a fuel cell installed. If so, the applicant must show that puncture hazards to the fuel cell have been eliminated.

(i) If the applicant elects to perform the drop test with surrounding aircraft structure, the cell should be enclosed in enough surrounding structure (production or simulated) so that the airframe/fuel tank interaction during the 50-foot drop is realistically evaluated. This allows the fuel-tight integrity of the “as installed” fuel cell to be evaluated and may provide protection in some designs due to the energy absorption of the surrounding airframe when crushed by impact. This provides realistic testing of fuel cell rupture points caused by installation design features, projections, excessive deformation and local tearout of fittings, joints, or lacings. The amount of actual (or simulated) structure included in the test requires engineering evaluation, risk assessment, and detailed analysis and may require subassembly (e.g., joint) tests for proper determination. Typically, the structure surrounding and extending 1 foot forward and aft of the fuel cell is adequate. This structure has a high probability of causing crash-induced fuel cell leakage. Each application should be examined individually to include all potential structural hazards. If the surrounding structure is clearly shown not to be a contributing hazard for the drop test, and if the applicant elects to do so, the fuel cell may be conservatively dropped alone. This determination should be carefully made by a detailed engineering evaluation. The evaluation should use standard, finite element-based programs (e.g., ‘KRASH”, NASTRAN, etc.) or similar programs submitted during certification, subassembly or component tests. Elimination of the surrounding structure for the drop test configuration is not trivial. If elimination is applied for, the data should clearly and conclusively show that the surrounding structure is not an impact hazard. In any case, the drop height is a constant 50 feet. The work that determines the test article configuration should be summarized, documented, and approved.

(ii) If the drop test is used to show partial compliance with the underfloor fuel cell load factors of § 29.952(b)(3), test plans should be approved. Minor spillage from the open vents is allowed. Full compliance to these load factors should be shown by static analysis and/or tests. The intent is to provide a fuel cell that is fuel tight and does not unduly crush the occupiable volume or overly stiffen energy absorbing underfloor structure under vertical impact.

(iii) Immediately after the drop test, the tank should be placed in the same axial orientation from which it was dropped and visually examined for leakage. Minor spillage from the open vents is allowed. After 15 minutes, the tank should be reexamined and any new leakage or seepage sources noted and recorded. Any evidence of fluid on the plastic floor cover or tank wrapping sheet should be noted and recorded. Any fluid leakage or seepage constitutes a test failure. This procedure should be repeated immediately with the tank inverted and the vents plugged. The inversion procedure will identify any leak sources on the upper surfaces.

Page E - 73 AC 29-2C 9/30/99 (2) Section 29.952(b) provides three sets of static load factors for design and static analysis of fuel tanks, other fuel system components of significant mass and their installations. “Installation” is structurally defined as the fuel cell’s attachment to the airframe and any additional local (point design) airframe structure affected significantly by fuel cell crash loads (i.e., that would fail or deform to the extent that a fuel spill or a ballistic hazard would occur in a survivable impact). Section 29.952(d) significantly limits the amount of local airframe structure to be considered. The provision of load factors by zone ensures the fuel-tight integrity necessary to minimize PCF in a survivable impact. Unless explicitly shown by both analysis and test that the probability -9 of fuel leakage in a survivable impact is 1 x 10 or less, each tank and its installation must be designed and analyzed to one set of these load factors. Also, as stated and explained in the advisory material for § 29.561, the load factors specified by § 29.561(d) are for the airframe structure surrounding the fuel cell only. The fuel cells themselves (and any fuel system components of significant mass in the underfloor area) and their attachments to the surrounding airframe structure are subject to the load factors of § 29.952(b)(3).

(i) Section 29.952(b)(1) provides load factors for the design and static analysis of fuel cells and their attachments inside the cabin volume. These load factors are provided to prevent crash-induced fuel cell ballistics hazards to and fuel spills (that may cause a PCF) directly on occupants from local structural failures in a survivable impact.

(ii) Section 29.952(b)(2) provides load factors for design and static analysis of fuel cells and their attachments located above or behind the cabin volume.

These load factors are provided to prevent injury or death from a fuel cell behind or above the occupied volume that is loosened by impact and to prevent fuel spills (which may cause a PCF) in a survivable impact.

(iii) Section 29.952(b)(3) provides load factors identical to those of § 29.561 for design and static analysis of fuel cells and attachments located in areas other than inside, behind, or above the cabin volume. Since many fuel cells are located under the cabin floor, these load factors provide fuel-tight structural protection in a survivable impact.

(iv) For some crash resistant semi-rigid bladder and flexible liner fuel cell installations, the 50-foot drop test (reference § 29.952(a)) can (with some additional rational analysis) simultaneously satisfy both the drop test requirement and the vertical down load factor (-N ) requirement of § 29.952(b)(3) for the fuel cell itself and its Z installation. This approach reduces the certification burden.

(v) For applicants that seek to substantiate the -N load factor Z requirement of § 29.952(b)(3) using the 50-foot drop test, additional substantiation is required for § 29.952(b)(3) (as is currently practiced) for the fuel cell under the loading of the remaining three load factors and the remaining rotorcraft structure under the Page E - 74 9/30/99 AC 29-2C loading of all four load factors. In some cases, substantiation of the remaining three load factors can be further simplified by a successful drop test if the fuel cell is symmetric (i.e., structurally equivalent in all four directions).

(3) Section 29.952(c) requires self-sealing breakaway fuel fittings at all fuel tank-to-line connections, tank-to-tank interconnects, and other points (e.g., fuel lines penetrating firewalls or bulkheads) where a reasonable probability (as determined by engineering evaluation, service history, analysis, test or a combination) of impact-induced hazardous relative motion exists that may cause fuel leakage to an ignition source and create a PCF during a survivable impact. In some coupling installations (such as fuel line-to-fuel tank connections), the tank coupling half should be sufficiently recessed into the tank or otherwise protected so that hazardous relative motion (of the fuel cell relative to its surroundings) following an impact-induced coupling failure does not cause a tearout or deformation of the tank half of the separated coupling that would release fuel. The only exceptions are either - (i) Installations that use equivalent devices such as extensible lines (hoses with enough slack or stretch to absorb relative motion without leakage) or motion absorbing fittings (rotational or linearly extensible joints); or (ii) Installations that conclusively show by a combination of experience, tests, and analysis to have a probability of fuel loss to an ignition source in a survivable -9 crash of 1 x 10 or less.

(4) Section 29.952(c)(1) specifies the basic design features required for self-sealing breakaway couplings.

(5) Section 29.952(c)(1)(i) defines the design load (strength) conditions necessary to separate a breakaway coupling. These loads should be determined from analysis and/or test, reference paragraph d(6). The minimum ultimate failure load (strength) is the load that fails the weakest component in a fluid-carrying line based on that component’s ultimate strength. This load comes from local deformation between the coupling and its surrounding structure during a worst-case survivable impact. A failure test of three specimens of the weakest component in each line that contains a coupling should be conducted in the critical loading mode. (If a single critical loading mode cannot be clearly identified, each of the three most critical loading modes should be tested.) The three specimen test results should be averaged. The average value is then used to size the breakaway fuel coupling. [For standard specification (i.e., “off the shelf”) hardware, equivalent testing may have already been accomplished and, if no other mitigating circumstances in the design and installation exist, need not be repeated.] To assure separation of the coupling prior to fuel line failure and to prevent inadvertent actuation, the design load that separates the coupling should be between 25 and 50 percent of the minimum ultimate failure load (strength) of the line’s weakest component. The critical loads should be compared to the normal service loads calculated and measured at the coupling location to insure unintended service failures do not occur. Typically this criterion is readily satisfied by the natural design because Page E - 75 AC 29-2C, Chg 1 2/12/03 working loads are much less than crash-induced loads. A separation load less than 300 pounds should not be used regardless of the line size. The minimum 300-pound load is necessary to prevent ground maintenance failures. A fatigue analysis and/or test (reference paragraph d(10)) should be performed to ensure the installation is either a safe-life design or has a conservative, mandatory replacement time. The simplified method of paragraph e(2)(i) of AC 27 MG-11 may normally be used because of the low ratio of working-load-to-crash-induced failure load. However, since fatigue failures have occurred in service, all fatigue sources (especially high-cycle vibratory sources) should be evaluated. Fracture critical materials should be avoided, and damage tolerant materials utilized. Also, if airframe deformation due to flight loads is significant, its effect on the couplings should be checked to ensure that static or low-cycle fatigue failures do not occur prior to the part’s intended retirement life. Large flight load deformations are not usually present in rotorcraft.

(6) Section 29.952(c)(1)(ii) requires a self-sealing breakaway coupling to separate when the minimum breakaway load (reference paragraph d(5) and § 29.952(c)(1)(i)) is met or exceeded in a survivable impact. The loading modes (each of which produces a breakaway load) are determined by analyzing and/or testing the surrounding structure to determine the probable impact forces and directions. The modes usually occurring are tension, bending, shear, compression, or a combination (reference figure AC 29.952–1). The coupling should be designed and tested to separate at the lowest ultimate impact load (lowest critical mode) as long as the minimum working load criterion of § 29.952(c)(1)(i) is also satisfied. Each breakaway coupling design should be tested in accordance with the following (reference MIL-STD-1290) or equivalent procedures. It should be noted that the ratio of the ultimate failure load of the weakest component in the fuel line and the normal service load (i.e., the peak load or approved clipped peak load experienced during a typical flight) of that component should be as high as possible and still meet the other load criteria of this section. Typically, this ratio should not be less than 5.

(i) Static Tests . Each breakaway coupling design should be subjected to tension and shear loads to verify and establish the design load required for separation, nature of separation, leakage during valve actuation, general valve functioning, and leakage following valve actuation. The rate of load application should not be greater than 20 inches per minute. Tests to be used where applicable are shown in figure AC 29.952-1.

(ii) Dynamic Tests . Each breakaway coupling design should be proof-tested under dynamic loading conditions. The couplings should be tested in the three most likely anticipated modes of separation as defined in paragraph d(5). The test configurations should be similar to those shown in figure AC 29.952-1. The load should be applied in less than 0.005 second, and the velocity change experienced by the loading jig should be 36 ±3 feet per second.

(7) Section 29.952(c)(1)(iii) requires that breakaway couplings be visually inspectable to determine that the coupling is locked together (fuel-tight) and remains Page E - 76 2/12/03 AC 29-2C, Chg 1 open during normal operations. Visual means (such as, an axial misalignment between the two coupling halves, a designed-in visual indicator, a combination or other acceptable criteria) should be considered and specified in the maintenance manual rejection criteria for operational inspections. Inspectability and phased inspection requirements should be evaluated. Special inspections after severe maneuvers or hard landings should be required.

(8) Section 29.952(c)(1)(iv) requires breakaway couplings to have design provisions that prevent uncoupling or unintended closing by operational shocks, vibrations, or accelerations. These provisions depend on both the coupling’s design and installation location. The structural environment should be defined, analyzed, and compared with coupling specifications and certification data so that inadvertent decoupling or closing does not occur. A phased inspection requirement should be considered.

(9) Section 29.952(c)(1)(v) requires a coupling design to not release more than its entrapped fuel quantity when the coupling has separated and each end is sealed off.

The entrapped fuel is determined by the coupling design and is essentially the fuel trapped between the seals when separation occurs (See breakaway coupling definition).

This is usually less than 8 ounces of fuel per coupling. Most coupling designs will leak slightly after separation. This is acceptable but the leak rate should be 5 drops per minute, or less, per coupling half. Specifications defining the entrapped volume of fuel should be approved. If the coupling is not approved or manufactured to an acceptable military or civil specification, the qualification testing of d(6) should be conducted.

(10) Section 29.952(c)(2) requires that each breakaway coupling or equivalent device either in a single fuel feed line or a complex fuel feed system (e.g. a multiple feed line or multitank cross feed system) be designed, tested, installed, inspected, maintained, or a combination, so that the probability of inadvertent fuel shutoff in flight is -5 1 x 10 , or less, as required by § 29.955(a). This should be determined by reliability and failure analysis, other analysis, tests, or a combination and should be documented and approved. Continued airworthiness should be ensured by phased inspections, specific component replacement schedules, or a combination. This section also requires each coupling or equivalent device to meet the fatigue requirements of § 29.571 to prevent leakage. (See the fatigue discussion in paragraph d(5).) The typical method of compliance with § 29.571 used for rotor system parts may not be necessary to meet § 29.952(c)(2). An S-N curve may not need to be generated using full-scale specimen fatigue tests if the conservative method of paragraph e(2)(i) of AC 27 MG-11, “Fatigue Evaluation of Rotorcraft Structure” can be applied successfully.

(11) Section 29.952(c)(3) requires that an equivalent device, used instead of a breakaway coupling, not produce a load, during or after a survivable impact, on the fuel line to which it attaches greater than 25-50 percent of the ultimate load (strength) of the line’s weakest component. This minimizes crash-induced fuel spills that may cause a PCF. The ultimate strength of the weakest component should be determined by Page E - 77 AC 29-2C 9/30/99 analysis and/or tests. At least three specimens of the component should be tested to failure in the critical loading mode and the results averaged. [For standard specification (i.e., “off the shelf”) hardware, equivalent testing may have already been accomplished and, if no other mitigating circumstances in the design and installation exist, need not be repeated.] The average value is then used to size the equivalent device. Each equivalent device must meet the fatigue requirements of § 29.571 to prevent fatigue-induced leakage. Equivalent devices should be statically and dynamically tested in an identical manner (where feasible) to breakaway couplings (reference paragraph d(6)). All fuel hoses and hose assemblies (whether or not they are used in lieu of breakaway fittings) should meet the following (reference MIL-STD-1290) or equivalent requirements. Any stretchable hoses used as equivalent devices should be able to elongate a minimum of 20 percent without leaking fuel. All other hoses used as equivalent devices should have a minimum of 20-30 percent slack. It should be noted that the ratio of the ultimate failure load of the weakest component in the fuel line and the normal service load (i.e., the peak or approved clipped peak load experienced during a typical flight) of that component should be as high as possible and still meet the other load criteria of this section. Typically, this ratio should not be less than 5.

(i) All hose assemblies should meet or exceed the cut resistance, tensile strength, and hose-fitting pullout strength criteria of MIL-H-25579 (USAF), MIL-H-38360, or equivalent standards.

(ii) Hoses should neither pull out of their end fittings nor should the end fittings break at less than the minimum loads shown in figure AC 29.952-3 when the assemblies are tested as described in d(11)(iii) below. In addition to the strength requirements, the hose assemblies should be capable of elongating to a minimum of 20 to 30 percent by stretch, slack, or a combination without fluid spillage.

(iii) Hose assemblies should be subjected to pure tension loads and to loads applied at a 90° angle to the longitudinal axis of the end fitting, as shown in figure AC 29.952-2. Loads should be applied at a constant rate not exceeding 20 inches per minute.

(12) Section 29.952(d) requires frangible or deformable structural attachments to be used to install fuel tanks and other major system components to each other and to the airframe when crash-induced hazardous relative motion could cause local rupture and tearout of the component, spill fuel to an ignition source, and create a PCF. If it can -9 be conclusively determined that the probability of fuel spillage is 1 x 10 or less, no further action is required. Typically, frangible designs are much easier to certify than deformable designs because the scatter in failure loads is much less. Also, some standard frangible military hardware (e.g., frangible bolts) is readily available. This is not so for deformable designs. Each frangible or deformable structural attachment and its installation should be reviewed to insure that, after an impact failure (i.e., separation or deformation), it does not become a puncture or tear-out hazard and cause fuel spillage.

Page E - 78 2/12/03 AC 29-2C, Chg 1 (13) Section 29.952(d)(1) defines the impact design load conditions necessary to deform a deformable attachment or to separate a frangible attachment. These loads should be determined from analysis and/or test (reference paragraph d(14)), and verified during certification. All impact loading modes (tension, bending, compression, shear, and a combination) should be analyzed and the minimum critical frangible or deformable design load determined, based on the ultimate strength of the attachment’s weakest component. The critical load should be compared to the normal service loads calculated and measured at the attachment’s location to insure unintended service failures do not occur. (Normally, this criterion is readily satisfied because working loads are much less than impact loads.) A fatigue check should be conducted to ensure that the attachments meet the requirements of § 29.571. Typically, this can be accomplished using the simplified method of paragraph e(2)(i) of AC 27 MG-11 because of the low ratio of working-load-to-crash-induced failure load. However, because of service history, all fatigue sources (especially high cycle vibratory sources) should be reviewed. The standard method of compliance with § 29.571 used for rotor system parts may not be necessary to meet § 29.952(d)(3). An S-N curve may not need to be generated using full-scale specimen fatigue tests, if the conservative method of paragraph e(2)(i) of AC 27 MG-11 can be applied successfully. Fracture critical materials should be avoided and ductile, damage tolerant materials utilized. Phased inspections to ensure continued airworthiness should be considered. Special inspections after severe maneuvers or hard landings should be required. A breakaway or deformation load less than 300 pounds (based on maintenance considerations) is not permitted. If airframe deformation due to flight loads is significant, its effect should be checked to ensure that a static failure or low cycle fatigue failure does not occur. Large flight load deflections are not usually present in rotorcraft.

(14) Section 29.952(d)(2) requires a frangible or locally deformable attachment to function when the minimum breakaway or deformation load (reference § 29.952(d)(1)) is met or exceeded in a survivable impact. The minimum breakaway or deformation load is the load that either breaks or deforms each of the frangible or deformable attachment(s) of each fuel cell, fuel line, or other critical fuel system component to the airframe. Each breakaway/deformation load must be between 25 percent to 50 percent of the load which would cause failure (i.e., impact induced tearout and subsequent fuel leakage) of the attachment to fuel cell, fuel line, or other critical component interface. This is necessary in some installations to prevent tearout of the structural attachment from the fuel cell component to which it is attached and the resultant fuel leakage in a survivable impact. The primary loading modes (each of which will produce a breakaway or deformation load) must all be considered to determine the minimum load. This is done by analyzing the surrounding structure (reference paragraph d(13)) to determine the three most probable impact failure forces and their directions. The attachment should then be tested to insure it breaks or deforms at the lowest ultimate crash (impact) load as long as the minimum working load criterion of § 29.952(d)(1) is also satisfied. It should be noted that the ratio of the ultimate failure load of the weakest component in the frangible or deformable component’s load path and the normal service load (i.e., the peak load or approved clipped peak load experienced during a typical flight) of that component should be as Page E - 79 AC 29-2C, Chg 1 2/12/03 high as possible and still meet the other load criteria of this section. Typically this ratio should not be less than 5. The following certification tests (reference MIL-STD-1290) or equivalent should be conducted on each frangible or deformable attachment design.

(i) Static Tests . Each frangible or deformable device should be tested in the three most likely anticipated modes of failure as defined in paragraph d(13). Test loads should be applied at a constant rate not exceeding 20 inches per minute until failure occurs.

(ii) Dynamic Tests . Each frangible or deformable attachment should be tested under dynamic loading conditions. The attachment should be tested in the three most likely failure modes as determined in paragraph d(13). The test load should be applied in less than 0.005 second, and the velocity change experienced by the loading jig should be 36 ±3 feet per second. It should be noted that the dynamic load pulse is a ramp function starting at either 0 or some small test fixture preload and reaching the previously determined failure load in 0.005 seconds. The velocity change of the test jig is also a ramp function starting at 0 and reaching a final velocity of 36 ± 3 ft./sec. in 0.005 seconds. These ramps functions simulate the dynamic conditions of a survivable impact under which the frangible/deformable attachment must perform its intended function.

(15) Section 29.952(d)(3) requires a frangible or locally deformable attachment to meet the fatigue requirements of § 29.571 to eliminate premature fatigue failure. The simplified method of AC 27 MG-11 may be used. Because of service history, all fatigue sources (especially high-cycle vibratory sources) should be reviewed. Fracture critical materials should be avoided and ductile, damage tolerant materials utilized.

(16) Section 29.952(e) requires that, as far as practicable, fuel and fuel containment devices be adequately separated from occupiable areas and potential ignition sources. Several generic categories of ignition sources and potential PCF-producing contact scenarios exist. The intent of the section is to define all possible leak and ignition sources that could be activated in a survivable impact and to provide design features to eliminate or minimize them such that the occurrence of PCF is minimized and escape time is maximized. Adequate separation should be accomplished by a thorough design review, potential PCF hazard analysis, and detailed design trade studies. The resultant findings should be documented and approved. The following PCF hazards and any other such hazards should be documented, minimized by design to the maximum practicable extent, and their resolution documented and FAA/AUTHORITY approved. Conditions to be reviewed should include, but are not limited to, the following: (i) High temperature ignition sources .

(A) Tank fillers or overboard fuel drains should not be located adjacent to engine intakes or exhausts so that fuel vapors could be ingested and ignited.

Page E - 80 9/30/99 AC 29-2C (B) Fuel lines should not be located in any occupiable area unless they are shrouded or otherwise designed to prevent spillage and subsequent ignition during and immediately following a survivable impact.

(C) Fuel tanks should not be located in or immediately adjacent to engine compartments, engine induction or exhaust areas, heaters, bleed air ducts, hot air-conditioning ducts, or any other hot surface.

(D) Fuel lines should be kept to a minimum in the engine compartment.

Fluid lines should not be located immediately adjacent to engine exhaust areas, heaters, bleed air ducts, hot air-conditioning ducts, or any other hot surface.

(E) Fuel lines should not be located where they can readily spill, spray, or mist onto hot surfaces or into engine induction or exhaust areas. These locations should be determined for each aircraft design by considering probable structural deformation hazards in relation to the fuel system.

(ii) Electrical ignition sources .

(A) Fuel tanks and lines should not be located in electrical compartments.

(B) Electrical components and wiring should be separated from fuel lines and vent openings kept to a minimum in fuel areas.

(C) Electrical wiring should be hermetically sealed, and equipment should be explosion proofed in areas where they are immersed in or otherwise directly subjected to fuel and vapors and should meet § 29.1309 or should be otherwise protected such that ignition is extremely improbable.

(D) Electrical sensor lines that penetrate fuel tank walls should be protected from abrasion or guillotine cutting during a survivable impact by use of potting, rubber plugs or grommets, or other equivalent means and should be designed with sufficient local slack, or equivalent means, to prevent both the wires and their protective mountings from being cut by or torn from fuel tank walls by local deformation.

(E) Electrical wires should be designed with sufficient slack or equivalent means to accommodate structural deformation without creating an ignition source.

(F) Electrical wires that could be subjected to severe local abrasion, cutting, or other damage during a survivable impact should be protected locally by nonconductive shields or shrouds.

(G) Electrical wires that are not sufficiently separated from heat or ignition sources to avoid potential contact during a survivable impact should be locally shrouded with a nonconductive fireproof shroud.

Page E - 81 AC 29-2C 9/30/99 (iii) Friction spark, chemical, and electrostatic ignition sources . Fuel lines and tanks should be designed and located to eliminate fuel or fuel vapor ignition from potential mechanical friction spark ignition sources, chemical ignition sources, and electrostatic ignition sources having a high probability of being activated or created during a survivable impact.

(iv) Separation of fuel tanks and occupiable areas . Fuel tanks should be located as far as practicable from all occupiable areas. This minimizes potential PCF sources in occupiable areas and the potential for occupant saturation with fuel on impact. The design should be reviewed to minimize these potential hazards. Fuel tanks should also be removed, as far as practicable, from other potentially hazardous areas such as engine compartments, electrical compartments, under heavy masses (e.g., transmissions, engines, etc.), over landing gear, and other probable areas of significant impact damage, including rollover and skidding damage.

(v) Fuel Line Shielding . Areas of the fuel line system where the probability of spilled fuel reaching potential ignition sources or occupiable areas is greater than extremely improbable should be shielded with drainable fireproof shrouds.

Shrouds should be drainable to allow periodic inspections for internal fuel leaks. The design should be reviewed to ensure these criteria are met.

(vi) Flow Diverters and Drain Holes .

(A) Drainage holes should be located in all fuel tank compartments to prevent the accumulation of spilled fuel within the aircraft. Holes should be large enough to prevent clogging by typical debris and to prevent fluid accumulation from surface tension force blockage.

(B) Drip fences and drainage troughs should be used to prevent gravity-induced flow of spilled fuels from reaching any ignition sources such as hot engine areas, electrical compartments, or other potential hot spots. Drip fences and troughs are also necessary to prevent PCF by routing spilled fuel around ignition sources to drainage holes to minimize fuel accumulation inside the fuselage. Recurring inspection requirements to ensure holes and troughs remain airworthy should be identified. These criteria should be met, as far as practicable, for all postcrash attitudes.

This is readily accomplished for the standard landing attitude, but is more difficult for other abnormal attitudes. However, the design should be thoroughly reviewed to insure maximum compliance without adversely impacting other safety and design criteria such as aerodynamic smoothness.

(vii) Fuel Drain System . The fuel drain system and its attachments to the airframe should be designed and constructed, as far as practicable, to be crash resistant. The following and other appropriate means should be considered for a crash resistant design. Tank drains should be recessed or otherwise protected so that they are minimally damaged by impact. Attachment of fuel drains to the airframe should be made with either frangible fasteners or equivalent means to prevent impact induced Page E - 82 9/30/99 AC 29-2C tearout and leakage. The number of drains should be minimized by design techniques such as those that avoid low points in the lines. Drain lines should be made of ductile materials or otherwise designed to provide impact tolerance. Drain line connections, fittings, and other components should be designed to meet the fatigue requirements of § 29.571 and § 29.952(d)(3). This ensures that unintended partial or full fatigue failures do not occur in normal operations that, if undetected, could compromise the CRFS’s intended level-of-safety for the mitigation of post crash fire in a survivable impact. Drain valves should be designed to have positive locking provisions in the closed position in accordance with § 29.999(b)(2).

(17) Section 29.952(f) specifies that fuel tanks, fuel lines, electrical wires, and electrical devices must be designed and constructed, as far as practicable, to be crash resistant. Typical mechanical design criteria necessary to minimize fuel spillage sources, ignition sources, and their mutual contact in a survivable impact (i.e., provide crash resistance) are stated by the following subparagraphs. These mechanical design criteria should be incorporated in each design to the maximum practicable extent.

Compliance is accomplished and assessed by a thorough design review and potential PCF hazard analysis with findings and solutions that are documented and approved.

Any additional PCF hazards that are identified should be documented, included, addressed equally, and eliminated to the maximum practicable extent. Engineering evaluation, analysis, and tests are all required to determine the maximum level of practicability.

(i) They should not initiate or contribute to a post crash fire in an otherwise survivable impact. A hazard analysis should show which components are critical in this regard and should be assessed in detail for hazard elimination purposes.

(ii) Fuel and electrical lines and components should be located away from each other, away from probable crash impact areas, and away from areas where structural deformation or large objects (such as engines or transmissions) may, by crushing or penetration, cause fuel spillage or create an electrical ignition source, or both.

(iii) Fuel and electrical lines and components should be located separately and away from areas where impact and severing by rotor blades during a survivable impact are probable.

(iv) Fuel and electrical lines and components should be in no danger of being punctured or severed during a survivable impact by locally stiff vertical understructure such as a collapsed landing gear strut.

(v) Fuel and electrical lines and components should be routed separately in areas of maximum protection, such as along heavier structural members, and away from areas where significant damage is probable.

Page E - 83 AC 29-2C 9/30/99 (vi) Fuel and electrical lines and components running through hazardous areas or directly through structure, such as a bulkhead, should be locally separated and protected from over-extension, severe abrasion and guillotine cutting by frangible panels, suitable clearance, rubber grommets, braided armor shielding (which should be nonconductive for electrical lines), or other equivalent means.

(vii) Fuel lines routed directly to instruments, transducers, or other equivalent devices should be crash resistant, in accordance with § 29.1337(a)(2), to minimize leakage in case of line rupture induced during a survivable impact.

(viii) Electrical wires routed directly into electrical boxes or instruments should be designed with sufficient local slack and locally routed in the least probable damage direction and zone, or otherwise protected to minimize the probability of damage-induced arcing.

(ix) Fuel lines routed directly into fuel tanks or other fuel system components should be locally routed in the least probable damage direction and zone, or otherwise protected, to minimize the probability of damage-induced fuel leaks.

(x) Fuel pumps mounted inside fuel tanks should be rigidly attached to the fuel tank only. If the pump is airframe mounted and has structural significance, it should have a frangible or deformable attachment (reference paragraph d(12)).

Electrical boost pumps, if used, should be installed with a minimum of 6 inches of slack wire at the pump connection. The pump wires should be shrouded to prevent cutting in a survivable impact. Nonsparking, breakaway wire disconnects or other equivalent means may be used in lieu of the 6 inches of slack wire.

(xi) Fuel filters and strainers, to the maximum practicable extent, should not be located in or adjacent to the engine intake or exhausts and should retain the smallest practicable quantity of fuel.

(xii) The number of fuel valves should be kept to a minimum. If electrically operated valves are used, they should be installed with a minimum of 6 inches of slack in the electrical lines, unless protected by equivalent means (reference 17(i)). The valves should be installed with the maximum amount of protection and separation of the electrical wires from the remainder of the valve assembly.

(xiii) Fuel quantity indicators mounted in or on fuel tanks should be selected, designed, and installed to provide the minimum puncture or tear hazard to the fuel tank in a survivable impact.

(xiv) Fuel tank and bladder enclosures should have smooth, regular shapes that avoid sharp edges and corners. Minimum concave and convex radius design criteria should be developed and adhered to. Magnesium should not be used in fuel cells, and any cadmium-plated parts should not be exposed to fuel.

Page E - 84 9/30/99 AC 29-2C (xv) Any shielding of electrical wires from abrasion, cutting, or overextension must be nonconductive.

(xvi) All fuel line installations not containing breakaway couplings should be reviewed to insure that they will not be overtensioned in a survivable impact, that they are properly grouped and properly exit fuel tanks, firewalls, and bulkheads in the area of least probable damage, and that their number and lengths are safely minimized.

(xvii) Crash resistance guidance for other basic components is contained in related paragraphs such as AC 29.963 (§ 29.963, bladders and liners), AC 29.973 (§ 29.973, fuel tank filler connections) and AC 29.975 (§ 29.975, fuel tank vents).

(18) Section 29.952(g) requires rigid or semirigid fuel tank or bladder walls of any material construction to be both impact and tear resistant. This minimizes a PCF from impact-induced rupture and tear.

(i) A rigid tank or bladder can resist fluid pressure loads as a flat plate in bending. A semirigid tank can resist fluid pressure loads partially as a flat plate in bending and partially as a membrane in tension. Flexible liners are exempt from the requirements of § 29.952(g) since an unsupported flexible liner can resist only pure tension loads acting as a membrane (i.e., it has negligible bending strength). The rigid shell structure required by § 29.967(a)(3) that surrounds the flexible liner (membrane) carries the crash-induced impact and tear loads; whereas, the flexible liner is only significantly loaded in tension if the shell structure is penetrated by a sharp object on impact.

(ii) For metallic tanks, rigid or semirigid composite tanks (resin matrix), semirigid bladder designs (rubber matrix), metal-composite hybrid designs, and all other tank designs, impact and tear resistance should be shown by analysis and tests.

(iii) Designs using resin matrix composites should be subjected to the composite structure substantiation guidance of AC 20-107A, Composite Aircraft Structure, dated April 25, 1984, and paragraph AC 29 MG 8. Designs using rubber matrix composites are subject to the standard substantiation requirements for these devices, such as TSO-C80.

(iv) One set of crash resistance tests that constitutes an acceptable method of substantiation to the requirements of § 29.952(g) for all tank designs regardless of the materials used are those specified in paragraphs 4.6.5.1 (Constant Rate Tear); 4.6.5.2 (Impact Penetration); 4.6.5.3 (Impact Tear); 4.6.5.4 (Panel Strength Calibration); and 4.6.5.5 (Fitting Strength) of MIL-T-27422B, “Military Specification; Tank, Fuel, Crash-Resistant Aircraft.” These test requirements, or equivalent means, should be applied for and discussed early in certification. If the MIL-T-27422B tests are selected, severity differences between military combat requirements and the civil environment should be accounted for by reducing the MIL-T-27422B requirements, as follows: Page E - 85 AC 29-2C 9/30/99 (A) Constant Rate Tear . The minimum energy for complete separation should be 200 foot-pounds (reference 4.6.5.1).

(B) Impact Penetration . The drop height of a 5-pound chisel should be reduced to 8.0 feet (reference 4.6.5.2).

(C) Impact Tear . The drop height of a 5-pound chisel should be reduced to 8.0 feet and the average tear criteria should not exceed 1.0 inch (reference 4.6.5.3).

(19) Section 29.952(g) also requires that all fuel tank designs (regardless of the materials utilized and whether or not a flexible liner of any type is used) for each tank or the most critical tank be analyzed and tested to the criteria of § 29.952 d(18)(iv), or equivalent.

(20) Any type of flexible liner or bladder used in any type of fuel tank construction (integral, hard shell, etc.) must meet the strength and puncture resistance requirements of § 29.963(b). Section 29.963(b) contains the new puncture resistance requirement for flexible liners and other liner material certification requirements.

Unlined, bladderless fuel tanks are also required to meet this requirement. Most unlined, rigid fuel cell designs should readily exceed the 370-pound minimum puncture force requirement because of overriding design requirements and material characteristics, such as stiffness and ductility.

NOTE: TSO-C80, “Flexible Fuel and Oil Cell Material,” is referenced in the advisory material for § 29.963(b) and contains the detailed qualification requirements for these materials. The current puncture resistance test of TSO-C80, paragraph 16.0, states that the force required to puncture the bladder material must be greater than or equal to 15 pounds (e.g., screwdriver test). Section 29.963(b) has increased the TSO paragraph 16.0 puncture force value to be greater than or equal to 370 pounds. This is for fuel cell bladder or liner material only. Oil cell material puncture force requirements are not changed.

e. Typical Examples of Loading Modes and Test Setups for CRFS Components .

The following figures, which are referred to periodically in the advisory circular, show typical examples of test setups for CRFS components such as breakaway fuel fittings, hoses, hose end fittings, and hose assemblies.

Page E - 86 9/30/99 AC 29-2C LOAD

t

+

LOAD STATIC SHEAR TEST STATIC TENSION TEST LOAD

t

LOAD

t

LOAD STATIC BENDING STATIC SHEAR TEST (TENSION-SHEAR) TEST (TANK-TO-TANK COUPLING) FIGURE AC 29.952-1 STATIC TENSION AND SHEAR LOADING MODES Page E - 87 AC 29-2C 9/30/99 ).

., TENSION TESTS 90-DEGREE TESTS FIGURE AC 29.952-2 HOSE ASSEMBLY TESTS Page E - 88 9/30/99 AC 29-2C Load (lbl Tension Load (lbl Bendin Hose End Minimum Minimum Minimum Minimum Fitting Average Individual Average Individual Fitting Type Size Load• Load Load* Load STRAIGHT -4 600 475 425 400 Tension= -6 700 575 425 400

i

-8 900 650 650 600

,JL -10 1450 1175 675 625

-12 1775 1475 950 850

Bending= t

-16 2125 1825 1425 1300

~-

-20 2375 2075 1550 1425 -4 600 475 425 400 9Q EI BQW -6 700 575 425 400 Tension= -8 900 650 450 400

i

~ -10 1450 1175 425

-12 1775 1475 500

Bending= t

-16 2125 1825 775 700

Jr

-20 2375 2075 1100 1000 • Average of at least 3 tests.

FIGURE AC 29.952-3 MINIMUM AVERAGE AND INDIVIDUAL LOADS FOR HOSE AND HOSE-END FITTING COMBINATIONS Page E - 89 AC 29-2C 9/30/99 AIR FRAME STRUCTURE -TANK HOSE END COUPLING ,..___ METAL TANK FITTING FLEX HO - BREAKAWAY VALVE ITEM LOWEST FAILURE LOAD (LB)* FAILURE MODE Flex Hose 3000 Tension Breakage Pul Out of End Fitting Flex Hose 1500 Tali< Fitting 7500 Pul Out of Tank Hose End Coll)ling Break (Bending) Breakawoy Valve 2500 Pul Out of Tank Fitting Breakawoy Valve Break at mmgillle Hot More Than 1500 = 750 Seciion Hot Less Than 1500 = 375 *Loads may or may not be representative; values are for explanatory purposes only.

FIGURE AC 29.952-4 TYPICAL METHOD OF BREAKAWAY FUEL FITTING LOAD CALCULATIONS (TANK INSTALLATION USED AS EXAMPLE ONLY; BASIC TECH NI QUE APPLICABLE TO OTHER CONFIGURATIONS) Page E - 90 9/30/99 AC 29-2C TANK WALL AIRCRAFT STRUCTURE~ MET ALT ANK FITTING CRITICAL FLANGE AREA A. FRANGIBLE BOLT AIRCRAFT STRUCTURE AIRCRAFT STRUCTURE 4000 SHEAR TANK FITTING 3000 PULLOUT OF TANK FLANGE 5000 SHEAR FRANGIBLE BOLT NOT MORE THAN NOT LESS THAN BREAK (TENSION-SHEAR) 3000 = 1500 3000 = 750 FIGURE AC 29.952-5 TYPICAL METHODS OF FRANGIBLE OR DEFORMABLE ATTACHMENT LOAD CALCULATIONS EXAMPLE 1, FRANGIBLE BOLTS.

Page E - 91 AC 29-2C 9/30/99 ~AIRFRAME RIGID BULKHEAD GROMMET ---FRANGIBLE BAFFLE

~

LOWEST FAILURE LOAD LB• FAILURE MODE ITEM RIG ID BULl-0-l EAD BEARING FLEX HOSE 3000 TENSION BREAKAGE FLEX HOSE 1500 PULLOUT OF END FITTING END FITTING 1750 BENDING FRANGIBLE BAFFLE NOT MORE THAN NOT LESS THAN BEARING 1500 = 750 1500 = 375 'VALUES ARE SHOWN FOR EXPLANATORY PURPOSES ONLY FIGURE AC 29.952-6 TYPICAL METHODS OF FRANGIBLE OR DEFORMABLE ATTACHMENT LOAD CALCULATIONS EXAMPLE 2, FRANGIBLE BAFFLE.

Page E - 92 9/30/99 AC 29-2C AC 29.953. § 29.953 FUEL SYSTEM INDEPENDENCE .

a. Explanation .

(1) Section 29.953(a)(1) stipulates that fuel systems for Category A rotorcraft must meet the requirements of § 29.903(b) engine isolation.

(2) Section 29.953(a)(2) specifies independent fuel feed systems for each engine for Category A rotorcraft unless other provisions are made to meet the § 29.903(b) engine isolation requirement.

(3) Section 29.953(b) specifies independent fuel feed systems for each engine for Category B rotorcraft, except that separate fuel tanks are not required.

b. Procedures .

(1) The purpose of § 29.953(a) is to ensure an independent fuel supply system for each engine. Multiengine Category B rotorcraft do not require separate fuel tanks, as are intended for Category A.

(2) The assessment to ensure compliance with § 29.903(b), engine isolation, should include consideration of component failure, malfunction, and damage. For multiengine Category B rotorcraft, leakage of the fuel cell could be excluded from consideration since § 29.953(b) explicitly states that separate fuel tanks are not required for this category rotorcraft.

NOTE: Of interest is that § 29.903(c), engine isolation for normal category airplanes, also excludes the fuel tank from consideration if only one tank is used.

(3) Consideration of fuel tank leakage under § 29.903(b) has dictated separate fuel tanks for Category A rotorcraft, but the regulation leaves the door open for unique designs by the expression, “Unless other provisions are made…,” in § 29.953(a)(2).

Separate tanks are intended for Category A as evidenced by the identical fuel system independence requirements for multiengine Category B rotorcraft, except that separate tanks are specifically not required.

(4) A common supply tank, with individual “collector” tanks for each engine for Category A rotorcraft, has been allowed under § 29.953 provided that the capacity of the collector tanks will allow 20 minutes of maximum allowable en route OEI power.

(5) The fuel system independence regulations are not intended to preclude single-point fueling designs.

(i) For multiengine Category B rotorcraft, the assessment of an independent fuel supply system for each engine would begin at the fuel supply pickup point within the tank and continue to the engine fuel inlet at the engine.

Page E - 93 AC 29-2C 9/30/99 (ii) For Category A rotorcraft, the assessment would begin with the tanks and continue to the engine fuel inlet.

(6) If supply line crossfeed capability is included as a feature, care must be exercised to ensure that the opening of the crossfeed does not jeopardize the continued safe operation of more than one engine. For example, if the crossfeed valve is automatically operated by a low pressure signal in the supply line for one engine, the possibility that fuel line leakage could cause opening of the crossfeed and jeopardize the continued safe operation of both engines should be considered. Similarly, opening the crossfeed valve with a suction lift system following engine or system malfunction should not allow air into the fuel supply line of the remaining engine.

(7) The independent fuel supply system requirement for each engine is for normal fuel system operations. Care should be exercised to ensure that flight manual procedures do not authorize normal usage of fuel system configurations which may violate the engine isolation principle. For example, routine fuel balance procedures should not allow usage of a common supply line if a failure can jeopardize the continued safe operation of more than one engine.

(8) Fuel system designs which allow the continued safe operation of all engines under expected fuel system component failure conditions (for example, a failed boost pump) by using common fuel flow paths under failure conditions are not prohibited.

(9) For APU’s which perform a required in-flight function, a separate, independent fuel system complying with the corresponding engine fuel system rules should be provided. Other APU’s (which do not perform a required in-flight function) may be supplied with fuel from a tee connection to a main engine fuel supply. The fuel shutoff valve for the APU should be located as close as possible to the APU system’s connection to the main engine fuel system and a checkvalve should be included in the APU fuel system to prevent reverse-flow if negative pressure exists momentarily in the main engine fuel system. Maximum fuel demand of the APU will not jeopardize compliance with § 29.955.

AC 29.954. § 29.954 (Amendment 29-26) FUEL SYSTEM LIGHTNING PROTECTION .

a. Background . During the initial development and promulgation of the standards concerning the airworthiness of rotorcraft, it was not deemed necessary to specify design features that would protect the rotorcraft from the meteorological phenomenon of lightning. This was due, in part, to the fact that rotorcraft were primarily operated in a VFR and non-icing environment. Also, a prudent pilot avoided thunderstorms where the possibility of encountering severe weather and a lightning strike was much greater. The construction, design, and operating environment of civil rotorcraft have changed markedly within the past two decades. Many rotorcraft are now authorized to fly IFR in all types of weather environment. One transport design has been approved for flight Page E - 94 9/30/99 AC 29-2C into known icing conditions. Additionally, many rotorcraft now use the same advanced technologies in structures and systems as do airplanes. Because of these facts the possibility of a lightning strike encounter to the rotorcraft has been greatly increased. If the fuel system of the rotorcraft has not been properly designed and constructed, a fuel vapor ignition may occur. This occurrence generally results in a catastrophe to the rotorcraft. To prevent such a catastrophe and provide a level of safety equivalent to transport category airplanes, a specific rule for the lightning protection of transport category rotorcraft fuel systems was adopted in Amendment 29–26.

b. Explanation .

(1) This regulation requires that the rotorcraft’s fuel system be designed and constructed so that an ignition of fuel vapor will not occur when the rotorcraft is involved in a lightning strike. For the purposes of this regulation the fuel system is comprised of the fuel tank with all its associated plumbing and any other areas of the rotorcraft likely to have fuel vapor present (such as sumps and drains for the tank itself). Externally mounted fuel tanks are also considered to be part of the “fuel system.” (2) Other associated installations such as electrical wiring in the fuel tanks which could provide a source of ignition due to an indirect or induced effect should also be considered.

c. Procedures .

(1) The current revision of Advisory Circular 20-53 provides guidance on an acceptable method and procedure to be utilized to demonstrate that the design and construction of the fuel system is compliant with § 29.954.

(2) FAA Report No. DOT/FAA/CT-89/22 contains additional information regarding the lightning environment. Also contained in this report are design and test techniques which provide for a design that will be adequately protected from fuel vapor ignition when the rotorcraft encounters the lightning environment. This report is available to the public by order from the National Technical Information Service, Springfield, VA 22161.

AC 29.955. § 29.955 (Amendment 29-2) FUEL FLOW .

a. Explanation .

(1) Section 29.955 is intended to ensure adequate fuel flow to the engine(s) at maximum power under the intended aircraft operating conditions and maneuvers. In ensuring adequate fuel flow, both hot and cold fuel would normally be evaluated for the suction lift system, whereas cold fuel is usually more critical for the boosted pressure system.

(2) In showing adequate fuel flow, the rule provides that- - Page E - 95 AC 29-2C 9/30/99 (i) The fuel be supplied within the appropriate engine fuel pressure range; (ii) The test be conducted with minimum fuel onboard, consistent with test safety; (iii) For pump systems, fuel flow requirements be satisfied with the critical airframe furnished pump inoperative; and (iv) The fuel flowmeter, if installed, must be blocked such that fuel must flow through the meter or its bypass.

(3) Section 29.955(b) specifies that if an engine can be supplied with fuel from more than one tank, the fuel system must feed promptly when fuel becomes low in one tank and another tank is selected.

b. Procedures .

(1) Testing (including bench tests) has been the accepted method to show compliance with § 29.955(a). Analytical techniques may be used to adjust the system test results to various fuel conditions and flows or to account for minor modifications to a system. A purely analytical approach is not generally acceptable.

(i) Methods to adjust the test data for different fuel properties and flows should be verified by limited testing.

(ii) If a suction lift system is used and hot fuel verification is involved (reference § 29.961) testing is appropriate.

(2) Demonstrating that the system is capable of providing “…100 percent of the fuel flow required under the intended operating conditions…” will depend on the particular system design, whether boosted or suction lift, Category A or Category B, and whether single or multiengine. Some of the factors to be evaluated are as follows: (i) Acceleration fuel flow requirements may exceed those for steady-state operation. For example, if on a cold day, engine torque is the limiting parameter, the steady-state fuel flow demand corresponding to that torque may be exceeded during engine acceleration to that power.

(ii) For single-engine rotorcraft and for multiengine rotorcraft with all engines operating, some margin should be included to account for possible inadvertent overtorque.

Page E - 96 9/30/99 AC 29-2C NOTE: Notice of Proposed Rulemaking (NPRM) No. 84-19 proposes to include this consideration as a firm requirement (reference 49 FR 46670; dated November 27, 1984).

(iii) For multiengine rotorcraft, adequate fuel flow under OEI conditions should be ensured.

(A) For Category A systems, evaluation of § 29.903(b) should ensure that following the failure of one engine, lack of fuel flow will not jeopardize the safe operation of the remaining engine(s). Since governor-controlled engines will automatically accelerate to some limit if power demand is high, and since immediate crew action is not presumed under § 29.903(b), compliance with § 29.955 would include adequate fuel flow to the cold day maximum OEI torque to be expected (reference § 29.927(b)(2)).

(B) A proposed revision to § 29.955 (reference NPRM No. 84-19) would require that fuel flow for multiengine Category B rotorcraft be adequate for the § 29.927(b)(2) OEI overtorque condition.

(C) Following an engine failure, the remaining engine(s) may accelerate to the gas producer speed topping limit fuel flow, rather than to the fuel flow for the steady-state OEI power value. This consideration would be most important for suction lift systems which may be critical with hot fuel at altitude.

(3) The critical fuel system configuration should be evaluated.

(i) For pump fed (boosted) systems, fuel flow requirements should be satisfied with the critical airframe furnished pump inoperative.

(ii) If on multiengine rotorcraft it is acceptable to operate following an engine failure in more than one fuel system configuration (for example, if crossfeed is an acceptable mode), then the supplying of multiple engines through common components may be more critical than the OEI condition.

(4) Adverse transient and steady-state maneuver loads should be considered since the g-loading experienced may tend to decrease the engine fuel inlet pressure below allowable limits.

(5) The fuel should be delivered to the engine inlet within the limits specified in the engine type certificate. The method of specifying these fuel inlet pressure requirements varies with the engine model. Some of these include: (i) Specification of a gage pressure as a function of altitude for suction system operation. The particular fuel and fuel temperature for demonstrating the criteria may be specified in the engine documents. Other approved fuels, fuel temperatures, and boost-pump-on operation are considered satisfactory if the demonstration with the specified fuel is successful.

Page E - 97 AC 29-2C 9/30/99 (ii) Specification of a maximum allowable vapor-to-liquid ratio for hot fuel, and minimum absolute pressure as a function of altitude for cold fuels.

(iii) Specification of a fuel inlet pressure relative to the true vapor pressure of the fuel, in combination with a maximum allowable vapor-to-liquid ratio.

(iv) Specification of separate pressure limits for boost-on and suction lift operation.

(v) Specification of special limits for emergency use or emergency fuels.

(6) For those systems which specify a minimum V/L ratio, the methods provided in Aerospace Recommended Practice (ARP) 492 published by the Society of Automotive Engineers are acceptable in evaluating test results.

(7) Since the lower quantity of fuel in the tank will reduce the hydrostatic head and thus the fuel inlet pressure, § 29.955(a)(2) specifies that the quantity of fuel in the tank should be minimum.

(8) Section 29.955(a)(3) specifies that each main and emergency pump be evaluated. If it can be determined which pump and flow path is critical, only that configuration would be tested. Similarly, for suction fuel systems, the critical flow paths and flow requirements should be evaluated. If pumps are required to supply the necessary fuel, § 29.1305 would require a fuel pressure indicator and § 29.1549 would require a red radial at the minimum safe operating fuel pressure for any fuel or fuel usage condition. This pressure limit should be used to determine compliance with § 29.955(a)(1) for all operations.

(9) Section 29.955(a)(4) specifies that the fuel flowmeter, if installed, be “blocked” in showing compliance with the fuel flow requirements. Consideration of flowmeter component failure or malfunction would most often be more appropriate than blockage.

(i) If the flowmeter is completely blocked in assessing compliance, then a bypass would be dictated, and the provision for “flow through the meter” following blockage would not be a viable alternative. It is not the intent of the rule to arbitrarily preclude flowmeter installations without a bypass system.

(ii) Section 29.1337(c) clarifies that if the malfunction of a metering component severely restricts fuel flow, a bypass would be required. An example of a malfunction to be considered would be a locked rotor on a rotating element design.

(iii) NPRM No. 84-19 proposes to clarify the intent of § 29.955 by requiring that proper fuel flow be ensured with fuel flow transmitter component failure, rather than with transmitter blockage as specified in the existing rule.

Page E - 98 9/30/99 AC 29-2C (10) Section 29.955(b) requires the fuel system to feed promptly when fuel becomes low in one tank and another tank is selected. This requirement is important because momentary fuel flow interruption must be expected to result in complete power failure and, for single engine rotorcraft, an emergency landing.

AC 29.955A. § 29.955 (Amendment 29-26) FUEL FLOW .

a. Explanation . Amendment 29-26 adds new requirements for test conditions to ensure that adequate fuel flow is available to the engine in critical combinations of adverse conditions that may be expected during operation of the rotorcraft. The amendment also requires a correlation between fuel filter blockage and the fuel filter warning device required by § 29.1305(a)(17). Design and performance standards for auxiliary fuel tank and transfer tank fuel systems are provided. These changes were made to ensure that all parameters associated with fuel supply to the engine are adequately addressed.

b. Procedures .

(1) Section 29.955 is intended to ensure adequate fuel flow to the engine(s) during all operating conditions of the rotorcraft. This includes the fuel flows necessary to operate the engine(s) under the test conditions required by § 29.927. Testing (including bench or rig tests) has been the accepted method of showing compliance with this section although analytical techniques may be used to adjust system test results to various fuel flow conditions or to account for minor modifications to a system.

Analytical methods that are used to adjust the test results should be verified with limited testing. It should be shown during compliance testing that the fuel pressure, at the engine to airframe interface, will be within the limits specified by the engine manufacturer. The fuel pressure at this point should be maintained within limits specified by the engine manufacturer during all critical maneuvers and accelerations.

All of the following conditions should be met during compliance testing unless it can be shown that combinations of the conditions are not possible.

(i) The fuel quantity in the tank(s) in use during the test may not exceed the unusable fuel quantity established under § 29.959, plus the minimum quantity required to conduct the test.

(ii) During the compliance test, the rotorcraft should be maneuvered to create the most critical fuel pressure head between the fuel tank outlet and the engine to airframe interface (engine fuel inlet).

(iii) For boost pump fed systems, it should be determined which pump (primary or secondary) would create the most critical restriction if it failed. The critical pump should then be installed to create the critical restriction, either by actual or simulated failure.

Page E - 99 AC 29-2C 9/30/99 (iv) Various combinations of engine power demand, electrical power available, and motive flow requirements for ejector pumps, will have an effect upon the fuel flow and pressure available at the engine to airframe interface. Adequate fuel pressure should be available to the engine with the most critical combination of these parameters.

(v) Critical values of fuel properties that may adversely affect fuel flow and/or fuel pressure should be applied. This includes alternate types of fuel if certification with alternate fuels is requested. At the minimum, the fuel that will create the highest vapor to liquid ratio should be used during hot fuel tests (§ 29.961). The most viscous fuel should be used during cold fuel tests.

(vi) The fuel filter, required by § 29.997, should be partially blocked to simulate the maximum contamination allowable. The blockage should be sufficient to activate the impending bypass indicator that is required by § 29.1305(a)(17).

(2) Unique Conditions . The phrase, “…Provide the engine with at least 100 percent of the fuel required under all operating and maneuvering conditions…” (§ 29.955(a)), includes unique flight conditions within the operational envelope of the rotorcraft. Critical conditions of fuel flow to the engine(s) may exist under the following conditions (and others identified by the applicant); therefore, they should be evaluated and tested if applicable: (i) In a single engine rotorcraft, a rapid acceleration to maximum power (torque) that will be requested for certification may be a critical condition. In this case the fuel flow required during the transient may exceed the fuel flow required for steady state at the maximum power condition.

(ii) In multiengine rotorcraft, a rapid acceleration to the maximum OEI power rating that will be requested may be a critical condition. The fuel flow during the transient may be higher than that required at the steady state OEI condition.

(3) If auxiliary fuel pumps (boost pumps) are used to supply fuel to the engines, and ejector pumps are used for cross-feed or other inter-tank fuel distribution systems, a test should be run that will place the maximum fuel demand on the auxiliary pump(s).

(4) In some multiengine rotorcraft, a single pump may be required to provide fuel flow to all engines in the event of an auxiliary pump failure. If this is the case, a test should be conducted with a simulated (or actual) failed auxiliary pump. If the functional auxiliary pump is designed to provide motive flow for cross-feed systems, the most critical condition of fuel flow demand should be tested.

(5) Transient and steady state maneuver loads (g-loading) may affect the fuel pressure at the engine to airframe interface. This effect should be considered and then tested, if appropriate.

Page E - 100 9/30/99 AC 29-2C (6) The methods of specifying the engine inlet fuel pressure requirements are sometimes related to fuel temperature and altitude. Therefore, it is necessary to explore the extremes of the envelope to assure compliance rather than attempting to select one critical condition. For instance, the increase in fuel viscosity at cold temperatures may increase system pressure drop and offset a slight drop in required fuel flow. In this case, critical fuel inlet conditions may not be experienced at maximum engine fuel flow.

(7) A conservative demonstration would consider the maximum allowable fuel viscosity in combination with the maximum fuel flow. Otherwise, several test points may be required.

Page E - 101 AC 29-2C 9/30/99 t:- I ~, 'v I '-> I ~, FUELFL - - - - L OW (W) LIMIT , --- f /

T

w -------

::::, W" f Ct: f- OAT--+ FIGURE AC 29.955A-1 FUEL FLOW NOTES (1) Point A on figure AC 29.955A-1 is the highest fuel flow within aircraft limitations, but the system pressure drop is not expected to be maximum because of the low kinematic fuel viscosity (2) Point Bis the maximum flow at cold temperatures but as the fuel temperature is further reduced, the fuel viscosity in creases very rapidly (3) Point C represents the maximum viscosity of the fuel, but the fuel flow is somewhat reduced from point B. The maximum system pressure drops and, therefore, minimum fuel inlet pressure may occur between points Band C depending on the specific relationship of fuel viscosity to required fuel flow.

Page E - 102 623/23 AC 29 - 2C, Chg 9 AC 29.955B. § 29.955 (AMENDMENT 29 - 59 ) FUEL FLOW.

a. Explanation . Amendment 29 - 59 revised § 29.955(a)(7) to correct a reference from § 29.1305(a)(17) to § 29.1305(a)(18). When § 29.1305 was updated at Am endment 29 - 40, the numbering sequence was changed. The fuel filter contamination warning, § 29.1305(a)(17), was referenced in § 29.955(a)(7) but was not updated at Amendment 29 - 40 to account for the change in numbering sequence.

b. Procedures . The policy pertaining to the procedures in this section remains unchanged, however the reference to the § 29.1305 fuel filter contamination indicator has been corrected.

Page E - 102.1 9/30/99 AC 29-2C AC 29.957. § 29.957 FLOW BETWEEN INTERCONNECTED TANKS .

a. Section 29.957(a) : (1) Explanation . This paragraph sets forth a design requirement that prohibits approval of a fuel tank interconnect arrangement wherein gravity or acceleration-induced flow between tanks will result in overflow through a tank vent.

(2) Procedures . The design of the vent for the receiving tank should be sufficiently elevated to preclude gravity or flight accelerations from causing overflow through the vent. A flight test may be needed to determine the effectiveness of the arrangement. Check valves in the vent system to prevent overflow should be discouraged because of reliability aspects.

b. Section 29.957(b) : (1) Explanation . For fuel system arrangements which permit fuel to be pumped from one tank to another, design precautions to prevent structural damage to the receiving tank in the event of overfilling are required as well as a design means to warn the crew before overflow through the vents occurs.

(2) Procedures . The design of the receiving tank should have large vent lines or a recirculation line back to the original tank to prevent overfilling of the receiving tank.

Alternatively, a float switch may be used to de-energize the transfer pump, providing that faults in the system do not adversely affect safety. A float switch may be used to warn the crew that overfilling of the receiving tank is impending. If a float switch is used, review the system reliability requirements of § 29.901(c).

AC 29.959. § 29.959 UNUSABLE FUEL SUPPLY .

a. Explanation . This rule requires the applicant to establish a value for unusable fuel for each tank. This value for unusable fuel may be selected by the applicant to facilitate compliance with § 29.1337(b)(1) provided the amount is equal to or greater than the actual unusable fuel. The actual unusable fuel is the amount of fuel in the tank when, in the critical flight attitude, evidence of system or engine malfunction occurs, or in the case of transfer tanks, when flow to the receiving tank is interrupted.

b. Procedures .

(1) The unusable fuel for each tank can be determined by flight tests which involve flight in the critical attitude or maneuver until indication of a malfunction. For boosted systems, the “first evidence of malfunction” may be a pressure fluctuation to below the fuel pressure minimum redline, engine power fluctuation, or boost pump failure warning indication. For suction lift systems, the indication may be a low fuel pressure warning light. In some instances, particularly for suction-lift systems, special test instrumentation for fuel pressure is required, and, since an accurate measurement Page E - 103 AC 29-2C 9/30/99 of the remaining fuel in the tank should be obtained, a method to close off flow from that tank would be needed. For transfer tanks, or tanks which are limited to use only during cruise flight, the flight regimes usually can be limited to level flight or hover at the c.g.

condition which, by inspection, would create the maximum unusable fuel. For tanks for general use, the flight regimes should also include takeoff and landing using steady pitch attitudes to be expected, as well as hover and level flight conditions. The possible adverse effects of extreme lateral c.g. should be considered.

(2) Normally, these tests are conducted with all equipment (pumps, ejectors, etc.) operating as prescribed by the design. However, values for unusable fuel with pump failures, if significantly different, should also be determined and listed in the flight manual. These values for unusable fuel need not be considered in the empty weight of the aircraft.

c. While the procedures of paragraph b(1) are acceptable, fuel exhaustion during critical flight test conditions must be expected. To minimize this possible flight test hazard, the applicant may in many cases, utilize analysis and/or ground tests involving normally available flight test data on aircraft attitudes, tank configuration studies, and critical flight condition studies to determine unusable fuel. Any questionable results, however, should be resolved by actual flight test or introduction of conservatism into the finding.

AC 29.961. § 29.961 FUEL SYSTEM HOT WEATHER OPERATION .

a. Explanation .

(1) Section 29.961 specifies that a hot fuel test be conducted on suction lift systems, and on other fuel systems conducive to vapor formation, to ensure that the system is free from vapor lock at a fuel temperature of 110° F under critical operating conditions.

(2) Pressure boosted systems would not ordinarily require hot fuel tests unless - - (i) There are high points in the fuel system which would allow accumulation of vapor; or (ii) The engine fuel inlet pressure is negative relative to tank pressure because of low boost pump pressure or high fuel system pressure losses (but still within fuel pressure limits).

(iii) The airframe boost pump is not actually submerged such that a portion of the system is suction lift.

(3) Boosted system vapor lock difficulties, at relatively low system flows compared to pump capacity, have occurred in at least two instances.

Page E - 104 9/30/99 AC 29-2C (i) If the fuel pump is a positive displacement type with an internal bypass and the pump capacity significantly exceeds system demand, excessive recirculation within the pump may significantly raise the local fuel temperature resulting in pump cavitation.

(ii) Parallel pump systems, where one supplies the majority of the fuel while the other “deadhead” pump supplies only a negligible amount of fuel, may experience vapor lock and cavitation of the deadhead pump due to excessive recirculation of fuel as described in a(3)(i).

(4) The requirement to use 110° F fuel is a carryover from the recodification of CAR Part 6, although the use of hotter fuel at the same Reid Vapor Pressure would tend more toward vapor formation.

(5) The term “vapor lock” means a change in normal engine operation as a result of the formation of fuel vapor-air mixtures in the fuel feed system.

(6) Section 29.961(b) and (c) inappropriately specify a particular flight condition, weight, and power spectrum which may not be critical. Hence, a demonstration of compliance to the specifics of § 29.961(c) will probably be inadequate for compliance with § 29.961(a)(2). NPRM No. 84–19 proposes to revise § 29.961 to delete these unnecessary, detailed regulations with a simple requirement to show satisfactory operation under critical operating conditions with hot fuel. The guidance which follows should be sufficient to establish compliance with § 29.961, in total, without regard to the misleading specifics of § 29.961(b) and (c).

b. Procedures .

(1) The fuel type to be used should be that with the highest true vapor pressure (TVP) at the 110° F condition.

(2) The fuel should be heated as rapidly as possible since the longer fuel is heated the more vaporization occurs resulting in unconservative test results. Likewise, heating the fuel above the target temperature, then allowing it to cool will “weather” the fuel excessively resulting in a reduction in Reid Vapor Pressure and unconservative testing.

(3) If the test is performed at cool ambients, the fuel lines, tanks, etc., may have to be insulated to ensure that the fuel inlet temperature is approximately the same as would be experienced on a hot day. This should be verified by instrumenting the fuel temperature at the engine inlet.

(4) The fuel level should be the lowest consistent with test safety. The reference to full fuel tanks in § 29.961(c)(2) is misleading because: Page E - 105 AC 29-2C 9/30/99 (i) Section 29.955(a)(2) would require adequate fuel flow under low fuel level conditions.

(ii) The provision of § 29.961(a)(2) to verify satisfactory hot fuel operation “under critical operating conditions” would mean verification at maximum rate of climb and maximum fuel suction head. The maximum fuel suction head would occur with lowest fuel level.

(5) The flight tests to the service ceiling should include maximum power climbs to selected intermediate altitudes where various maneuvers including the following are performed: (i) Low power descent with rapid transition to takeoff power.

(ii) Turns and cyclic pull-ups with load factors comparable to the flight strain survey.

(iii) For multiengine rotorcraft with 30-minute and/or 2.5-minute OEI power ratings, conduct a rapid single-engine acceleration from low power to engine topping power followed by cruise at 30 minute OEI power.

(6) The flight test maneuvers should be repeated at the service ceiling.

(7) Except for transients and descents, the power available used should correspond to a 100° F sea level day lapsed 3.6° F/1,000-foot pressure altitude.

(8) Engine operation throughout the test should be normal; i.e., no surge, stall, flameout, etc., and the engine fuel inlet requirements should not be violated.

(9) Alternative tests on appropriate test rigs may be conducted ensuring proper simulation of altitude, ambient temperature, fuel temperature, fuel flow, and load factors.

AC 29.961A. § 29.961 (Amendment 29-26) FUEL SYSTEM HOT WEATHER OPERATION .

a. Explanation . Amendment 29-26 simplifies and restates the fuel system hot weather certification requirements. This eliminates detail requirements in the existing rule which were, to some extent, redundant, or not necessarily critical for some rotorcraft. The phrase, “including, if applicable, the engine operating conditions defined by §§ 29.927(b)(1) and (b)(2),” was added to ensure that certain critical certification aspects are properly considered.

b. Procedures . This paragraph specifies that all suction lift systems and any other fuel system that may be conducive to vapor formation, show satisfactory engine fuel inlet conditions (within criteria established by the engine manufacturer) when using the fuel with the highest true vapor pressure (TVP) at 110° F fuel temperature. Engine Page E - 106 9/30/99 AC 29-2C operating conditions should include those defined by §§ 29.927(b)(1) and 29.927(b)(2).

Compliance can be shown by analysis, testing, or a combination of both.

AC 29.963. § 29.963 FUEL TANKS: GENERAL .

a. Explanation . Section 29.963(a) sets forth general requirements for fuel tank structural aspects. Paragraph (b) requires design features to react forces defined by § 29.561 without leaking fuel. Paragraph (c) requires that whenever flexible fuel tank liners are used, they must be FAA/AUTHORITY approved. Paragraph (d) requires that integral fuel tank interiors be inspectable and repairable.

b. Procedures .

(1) For compliance with § 29.963(a), the tests of § 29.965 are normally adequate if performed in conjunction with the reliability test of § 21.35 or other simulation tests.

(2) For compliance with § 29.963 (b), a structural analysis is usually required to show adequate strength under the loads of § 29.561. Testing, if proposed, may also be an acceptable method of compliance.

(3) For compliance with § 29.963(c), prior FAA/AUTHORITY approvals should be reviewed to ensure compatibility with current project requirements. Also, if a new approval is required as part of the project, then analysis and/or tests should be conducted as appropriate to ensure compliance.

(4) For compliance with § 29.963(d), a review of the design data and/or a visual inspection of any prototype available for inspectability and repairability considerations is usually sufficient to determine compliance. Features such as inspection ports and access panels are typical methods of compliance.

AC 29.963A. § 29.963 (Amendment 29-26) FUEL TANKS: GENERAL .

a. Explanation . Amendment 29-26 adds § 29.963(e) that requires designs and tests to ensure that no exposed surface inside a fuel tank would, under normal or malfunction conditions, constitute an ignition source. It also sets forth standards for the design and qualification of fuel tanks located in personnel compartments. These requirements are needed to ensure freedom from the hazards of fuel tank internal explosions and to ensure that fuel tanks installed in passenger compartments present no hazards to the personnel or to the rotorcraft.

b. Procedures . Section 29.963(e) requires the temperature of any exposed surface inside a fuel tank to be at least 50° F lower than the lowest auto-ignition temperature of the fuel or fuel vapors in the tank (reference paragraph AC 29.1185b(3), § 29.1185). For compliance with § 29.963(e), the internal component surface temperatures can be determined by flight or laboratory tests. The most critical flight Page E - 107 AC 29-2C 9/30/99 conditions are established with sensitive temperature and pressure measuring equipment. This equipment is installed inside the tanks and in the ventilation air spaces.

AC 29.963B. § 29.963 (Amendment 29-35) FUEL TANKS: GENERAL .

a. Explanation . Amendment 29-35 adds a new paragraph (b) that includes the requirements previously contained in paragraph (c) that each flexible fuel tank bladder or liner be either FAA/AUTHORITY approved or be suitable for each particular installation. In addition, the new paragraph (b) adds the requirement that the fuel tank bladder or liner be puncture resistant by meeting the TSO C80, paragraph 16.0, screwdriver test requirements, using a new crash resistance based minimum puncture force of 370 lbs. The requirements previously contained in paragraph (b) are replaced by the crash resistant fuel system requirements of § 29.952 (including load factors). A new paragraph (e) is also added. Paragraph (e) requires that each fuel tank installed in a personnel compartment be isolated by fume-proof and fuel-proof enclosures that are drained and vented to the exterior of the rotorcraft. Further, the design and construction of the enclosures must provide the necessary protection for the tank, must be crash resistant by meeting the applicable criteria of the new Crash Resistant Fuel System requirements of § 29.952, and must be adequate to withstand the loads and abrasions to be expected in personnel compartments.

b. Procedures .

(1) Paragraph (b) . The procedures for paragraph (c) prior to Amendment 29-35 still apply to new paragraph (b). In addition, to comply with the added puncture resistance requirement under new paragraph (b), the requirements of § 29.952(g) must be met. Paragraph AC 29.952 gives the detailed compliance procedures for § 29.952(g). The compliance procedures for § 29.952(g) also provide compliance for puncture resistance under § 29.963(b).

(2) Paragraph (e) . Compliance with paragraph (e) can be shown by conducting a thorough design review of each fuel tank and its enclosure that is installed in a personnel compartment to ensure the regulatory criteria are met. (All fuel drains and vents should also be reviewed to ensure that they meet applicable § 29.952 requirements.) A basic static loads analysis followed by a stress analysis is typically used to determine that the enclosure protects the fuel tank and provides the crash resistance level necessary for occupant survival in an otherwise survivable impact. The applicable emergency load factors are typically used to design the enclosure.

(Section 29.952 contains the corresponding load factors for fuel cells and their attachments.) The emergency load factors are typically adequate for all loading conditions encountered by the enclosure in service. The typical design approach is to design the enclosure to crush at a rate approximately the same as the crush rate of the fuel tank and to ensure that all puncture hazards (such as sharp projections either enhanced or created by impact that would penetrate the fuel tank) are minimized in design. (See paragraph AC 29.952 guidance material for details.) The design of the Page E - 108 9/30/99 AC 29-2C enclosure should also be reviewed for overall durability and resistance to all reasonable occupant abuses that could cause a hazard to the integrity of the enclosure, the fuel tank, its vents and its drains.

AC 29.965. § 29.965 (Amendment 29-13) FUEL TANK TESTS .

a. Explanation .

(1) This section prescribes the fuel tank structural tests to be accomplished without failure or leakage.

(2) Section 29.965(b) prescribes pressure testing for conventional metal tanks, integral tanks, and for nonmetallic tanks with walls that are not supported by the rotorcraft structure.

(3) Section 29.965(c) prescribes pressure testing for nonmetallic tanks with walls supported by the rotorcraft structure.

(4) Section 29.965(d) prescribes slosh and vibration testing for tanks with large unsupported or unstiffened flat areas.

b. Pressure Tests .

(1) Each conventional metal tank, integral tank, and each nonmetallic tank without supporting rotorcraft structure should be subjected to pressures of at least 3.5 PSI gage.

(i) If the pressures developed during maximum limit acceleration or emergency deceleration with a full tank exceeds the 3.5 PSI value, a hydrostatic pressure test (or equivalent) should be used to duplicate these acceleration loads as far as possible.

(ii) Pressures need not exceed 3.5 PSI on surfaces not exposed to the acceleration loading.

(iii) Section 29.337 gives the value for the maximum limit acceleration.

(2) Section 29.965(c) applies to nonmetallic tanks with walls supported by the rotorcraft structure. Section 29.965(c)(1) does not require that the tank alone be capable of withstanding 2.0 PSI. Rather, the tank may be mounted in the supporting structure and subjected to the testing of § 29.965(c)(2).

(3) Pressure tests may be conducted by slowly applying a controlled, gauged air source to the tank with sealed vents and fluid entrances and exits. The air pressure source should then be positively sealed and the tank should retain the prescribed pressure.

Page E - 109 AC 29-2C 9/30/99 (4) Tank and surrounding structure should be carefully examined during and after pressure testing to ensure that there is no damage.

(5) If the prescribed 3.5 PSI or 2.0 PSI, depending on the type of tank, will be exceeded on some surfaces during maximum limit acceleration loading, hydrostatic testing may be preferred. High density fluids have been used to apply the acceleration loads to lower surfaces with supplemental air pressure used above the liquid surface to provide the appropriate pressure on upper surfaces.

(6) For fuel tanks in those areas designated by § 29.967(f), the pressure tests may be designed such that compliance with that paragraph also demonstrates compliance with § 29.965 pressure test requirements.

(c) Slosh and Vibration Tests .

(1) The test requirements of § 29.965(d) are very specific and require little explanation.

(2) There is not an absolute value of what constitutes “large” unsupported or unstiffened flat areas. However, it has generally been considered that any fuel tank with less than a 10-gallon capacity, constructed with a simple, wide, flat geometric shape and using metal (in metal tanks) of 0.05-inch thickness or greater would not require tests in accordance with § 29.965(d). Using this basis, a 14- by 14- inch properly constructed tank would not require vibration and slosh tests.

(3) If the tank construction is of a metal or integral design which can be shown to be similar to previously approved tanks with acceptable service history, the vibration and slosh tests may not be required. Similarity would entail comparing the construction technique; i.e., similar panel size, similar sealing methods, skin and angle thickness, similar loads, etc.

(4) For fuel tanks located in a sponson or stub wing, the entire sponson or wing should be rocked and vibrated unless it can be determined that a certain portion of the tanks is critical. In this case a fixture should be developed such that the portion of the tank being tested is rocked about a pivot point which would produce the same amplitudes of motion for the portion of the tank being tested, as if the whole sponson or wing was being tested. Structural loads in conjunction with these tests have not been required.

(5) The amplitude of vibration specified in the regulation is double amplitude (peak-to-peak). Vibration amplitudes less than one thirty-second of an inch should be justified by instrumented tests of the tank installed in the aircraft.

(6) The vibration and slosh procedures listed in Military Specification MIL-T-6396 have been accepted to show compliance with § 29.965(d).

Page E - 110 9/30/99 AC 29-2C (7) After all tests have been conducted, the tanks should be leak checked using test fluid conforming to Federal Specification TT-S-735 type III or equivalent.

AC 29.967. § 29.967 FUEL TANK INSTALLATION .

a. Section 29.967(a) : (1) Explanation . This paragraph sets forth a series of detail requirements for fuel tanks intended to ensure that tank leakage or failure is unlikely. These requirements pertain primarily to proper support of the tank and protection against chafing.

(2) Procedures . For conventional metal tanks, the support devices, commonly called “cradles,” should be designed with wide flanges or cap strips at the contact area with the tank to distribute the loads in the tank material. To prevent chafing, install nonmetallic padding, treated to eliminate absorption of fuel between the tank and the support structure. Cork strips sealed with shellac and bonded to the support structure have been found suitable. Fuel cell sealant material should be applied over rivet heads and in corners. Bladder cells must be designed to fit accurately in the cell cavity in order to avoid fluid loads in the bladder itself. The interior of the cavity should be smooth to avoid damage to the bladder cells.

b. Section 29.967(b) : (1) Explanation . This paragraph requires the design to provide ventilation and drainage of spaces adjacent to fuel tanks to avoid accumulation of fuel or fumes to be expected from minor leakage of fuel tanks. This is needed to minimize the possibility of fire or explosion in these spaces. An exception to this requirement is allowed for bladder cells installed in a closed compartment. For this configuration, ventilation may be limited to that provided by compartment drains if the ventilation is adequate to maintain proper pressure relationship between the bladder cell and cell compartment air spaces.

(2) Procedures . With the assumption that fuel tank leakage will occur, require the tank compartments to be provided with drains at any low point. These drains should conduct fuel clear of the rotorcraft and should be three-eighths of an inch or larger in diameter to minimize clogging. As with any drain intended to function in flight, verification that reverse flow will not occur due to pressure differentials at each end of the drain is appropriate. Ventilation for these tanks should involve openings in the compartment walls such that in-flight slipstream and/or rotor downwash will rapidly and continuously purge the tank compartment of fuel fumes. Openings should not be located so the fumes or fuel can reenter the rotorcraft. For flexible tank liner configurations (bladder cells), no specific ventilation is required if the cell is located in a compartment which is closed, except for drain holes. Note that a cell leak may be expected to produce fumes in the compartment airspace which are flammable; thus, Page E - 111 AC 29-2C 9/30/99 items installed in bladder tank cavities shall not create a hazard during either normal or malfunction conditions. The vent system for the interior of the cell must be adequate to ensure that the bladder cell interior pressure is always positive or at least neutral with respect to any other airspace in the cell compartment to prevent collapse of the bladder cell. Drainage of the cell compartment should meet the criteria discussed above.

(3) A light mesh or string network hung between the bladder cell and its compartment walls is recommended to provide seepage channels to facilitate fuel leakage to the low-point compartment drains.

c. Section 29.967(c) : (1) Explanation . This paragraph requires a measure of protection for fuel tanks from adverse effects of a fire in a fire zone.

(2) Procedures . Verify that a firewall meeting the requirements of § 29.1191(e) effectively separates any fuel tank from any engine. To minimize hazards of heat transfer to a fuel tank through a firewall during an engine compartment fire, verify that at least one-half inch of clear airspace exists between the tank and the firewall.

d. Section 29.967(d) : (1) Explanation . This paragraph is intended to prevent hazards to integral fuel tanks to be expected by impingement of flames or products of combustion from an engine compartment fire.

(2) Procedures . Review the design for relative positions of engine compartments and integral fuel tanks to estimate the flowpath of fire or heat from an engine compartment fire. Consider autorotation for single-engine rotorcraft and, for multiengine rotorcraft, low power descent as power-on flight in this evaluation. If questionable compliance exists, clear indication of the flow impingement patterns may be identified by ejecting a dye from engine compartment openings during flight.

e. Section 29.967(e) : (1) Explanation . This paragraph is primarily intended to provide a standard for installing fuel tanks in personnel compartments. The primary safety concern is to isolate fuel or fumes from personnel in event of a leak in the tank.

(2) Procedures . Assume a leak in the tank and determine that, through the use of additional walls, bulkheads, enclosures, etc., that fuel and fumes will be safely drained and/or purged to the exterior of the rotorcraft. Note that, in order to perform their intended function, the enclosure material and structure should withstand the mechanical stresses and abrasions to be expected from crew and passenger activities within the compartment.

Page E - 112 9/30/99 AC 29-2C f. Section 29.967(f) : (1) Explanation . This paragraph is intended to require the design to prevent fuel tank or tank support failure when exposed to the minor crash loads of § 29.561 if such failure could result in fuel entering personnel compartments or fire hazard areas.

(2) Procedures . If a review of the design indicates that tanks are in or adjacent to passenger compartments, or are adjacent to combustion heaters or engines (including APU’s), further evaluation of the structural integrity of the tank and its support features must be accomplished. Normally, this involves a quantitative analysis of the tank support structure to confirm that it can sustain the minor crash loads plus one or more pressure tests to simulate the fluid loads on the tank interior to be expected when the minor crash loads are applied. This latter requirement may be, in many cases, satisfied by the qualification requirements of §§ 29.963 and 29.965. Pressure tests tend to overstress upper surfaces of a tank in order to achieve the required stress in the lower surfaces. To minimize this, some applicants have filled the tank to be tested with high density fluids and applied only supplemental pressure to the airspace at the top of the tank. High density fluids are available from the petroleum industry.

AC 29.967A. § 29.967 (Amendment 29-35) FUEL TANK INSTALLATION .

a. Explanation . Amendment 29-35 removes paragraph (e) from § 29.967 and places the identical criteria in a new paragraph (e) to § 29.963. This was done to make §§ 29.963 and 29.967 parallel with §§ 27.963 and 27.967.

b. Procedures . The procedures specified in paragraph AC 29.967, subsection (e) now apply under paragraph AC 29.963B. Thus there is no change in the certification requirements or the compliance methodology, only a change in their location in the FARs and Advisory Material, respectively.

AC 29.969. § 29.969 FUEL TANK EXPANSION SPACE .

a. Explanation .

(1) Space must be provided in each fuel tank system to allow for expansion of the fuel as a result of a fuel temperature increase. The space provided for this purpose must have a minimum volume equal to 2 percent of the tank capacity.

(2) The fuel tank filling provisions must be designed to prevent inadvertent filling of the fuel tank expansion space when fueling the rotorcraft in the normal ground attitude on level ground.

b. Procedures .

(1) Fuel tanks with interconnected vents need not have provisions for fuel expansion in each tank if equivalent expansion provisions are available in another area.

Page E - 113 AC 29-2C 9/30/99 (2) The fuel filler ports should be located below the designated fuel expansion space height to assure that the fuel expansion space cannot be inadvertently filled with fuel. For pressure refueling systems, compliance with this section may be shown with the means provided to comply with § 29.979(b).

(3) Each fuel tank expansion space must comply with the venting requirements of § 29.975.

(4) For multiengine rotorcraft using a single expansion tank to satisfy the requirements of this regulation, the effect of blockage or failure of any vent from this common tank must be considered with respect to compliance with the applicable engine isolation requirements.

AC 29.969A. § 29.969 (Amendment 29-26) FUEL TANK EXPANSION SPACE .

a. Explanation . Amendment 29-26 was issued so that properly interconnected fuel tanks will not be required to have an expansion space for each tank if adequate expansion space is otherwise provided. This amendment eliminates unnecessary design requirements when simpler designs have been proven to be satisfactory.

b. Procedures . Methods of compliance are not changed with this amendment.

AC 29.971. § 29.971 (Amendment 29-12) FUEL TANK SUMP .

a. Explanation .

(1) Each fuel tank should be provided with a drainable sump which is located at the lowest point in the tank with the rotorcraft at ground attitude in order to allow drainage of possibly hazardous accumulations of water from the system.

(2) The minimum required sump capacity, 0.10 percent of the tank capacity or one-sixteenth of a gallon, whichever is greater, should be effective at any normal attitude and located such that the sump contents cannot escape from the tank outlet opening.

(i) Combined interconnected tanks can be treated as a single tank and utilize only one sump if that sump can be located to allow effective trapment and drainage of the potential combined water accumulation.

(ii) The requirement that sump contents not be allowed to escape through the tank outlet opening is intended to ensure that water, or other impurities which may precipitate from the fuel in the tank(s), does not enter the fuel feed system.

Page E - 114 9/30/99 AC 29-2C (3) Section 29.971(c) would ensure that the fuel tank design and installation allows drainage of hazardous quantities of water to the sump with the rotorcraft in the ground attitude.

(4) Section 29.971(d) would ensure that not only are possibly hazardous accumulations trapped, but also that they are drainable with the rotorcraft in the ground attitude.

(5) Proposed Amendments (Notice 84-19) to §§ 29.971(c) and 29.999(a) would require that the tank sumps be designed or arranged to collect water and be drainable in any ground attitude to be expected in service. This proposed provision would require consideration of the effectiveness of the sumps and drains at the sloped landing limits as well as at normal ground attitude.

b. Procedures .

(1) Demonstration of compliance with the minimum sump capacity requirements may be shown by analysis, test, or a combination of both depending on the complexity of the fuel system design.

(2) If minimum sump capacity is to be established by tests, the following procedure has been accepted.

(i) Fuel the aircraft tanks to ensure that all sumps are filled, that any transfer pumps are immersed, and that the fuel level is above the fuel feed pickup point in the tank(s).

(ii) Use the normal fuel feed provisions to remove fuel from the system.

The fuel inlet line at the engine/airframe interface may be disconnected and the fuel pumped overboard. If an engine-supplied suction lift pump is the normal feed mechanism, a suction lift pump of approximately the same capability may be substituted to avoid operating the engine.

(iii) Determine the most critical ground attitude to be expected in service from such considerations as uneven terrain, slope landing limits, etc. The critical attitude for each tank will be that for which the maximum amount of fuel can be withdrawn from the tank using the rotorcraft’s fuel supply system.

(iv) Using a rotorcraft with a fuel system which conforms to the final design specification, position the rotorcraft to the critical attitude for the tank to be tested using leveling jacks, actual terrain of a predetermined slope, or other similar means.

(v) Using the rotorcraft’s fuel supply system, pump fuel from the tank being tested until the supply system will no longer withdraw fuel. This can be done without the rotorcraft engine actually running unless an engine driven pump is an essential component of the fuel supply system. Caution should be exercised if an Page E - 115 AC 29-2C 9/30/99 engine is to be run to fuel exhaustion since engine surge at the pump cavitation point can result in damaging torsional loads in the transmission drive system.

(vi) When no more fuel can be removed from the tank with the rotorcraft fuel supply system, return the rotorcraft to a normal ground attitude. Completely drain the sump of the tank or tanks being tested into a container and measure the volume drained from each sump. The volume measured must satisfy the minimum capacity requirements of paragraph AC 29.971(a)(2).

(3) If, in the above procedure, a known quantity of fuel is added to initially empty tanks and the total fuel removed (pumped overboard and drained) is recorded, the data may also be used to show compliance with §§ 29.971(d) and 29.999(a).

AC 29.971A. § 29.971 (Amendment 29-26) FUEL TANK SUMP .

a. Explanation . Amendment 29-26 requires that fuel tank sump designs be arranged so that drainage from the sump area will be effective with the rotorcraft parked in any allowable ground attitude in lieu of “normal” attitude as previously required.

b. Procedures . All of the policy material pertaining to this section remains in effect with the extra requirement that each fuel tank should be provided with a drainable sump which is located at the lowest point in the tank with the rotorcraft at “any” ground attitude in order to allow drainage of possibly hazardous accumulations of water from the system. This provision requires consideration of the effectiveness of the sumps and drains at the sloped landing limits as well as at normal ground attitude.

AC 29.973. § 29.973 FUEL TANK FILLER CONNECTION .

a. Explanation .

(1) Fuel tank filler connections must be designed so that no fuel can enter into any part of the rotorcraft other than the fuel tank during fueling operations. Spilled fuel must be considered as well as fuel entering the fuel filler port.

(2) A recessed filler connection that can retain appreciable quantities of fuel should have a drain that discharges clear of the rotorcraft.

(3) Section 29.1557(c)(1) prescribes the marking of the filler.

(4) The filler cap must be fuel-tight under the pressures expected in normal operation.

(5) For Category A rotorcraft, the filler cap or cover must warn if the cap is not fully locked or seated. An improperly locked and seated fuel cap should be evident on the preflight inspections.

Page E - 116 9/30/99 AC 29-2C (6) The parallel Part 23 and 25 requirements specify that, except for pressure refueling connection points, the filling point must have a provision for electrically bonding the aircraft to ground fueling equipment. Though not specifically required by Part 29, rotorcraft manufacturers have included this provision in recognition that the same potential hazard exists for possible discharge of sparks between the fuel dispensing nozzle and the aircraft as would exist for airplanes.

(7) A proposed rule (Notice 84-19) would add a fuel system lightning protection requirement for rotorcraft. The potential for fuel vapor ignition near the filler cap would be a primary concern. (NASA publication 1008, Lightning Protection of Aircraft , and the user’s manual to AC 20–53A, Protection of Aircraft Fuel Systems Against Fuel Vapor Ignition Due to Lightning, provide further information.)

b. Procedures . Compliance with the requirements of this paragraph can normally be demonstrated by analysis and physical inspection of the fuel filler connection design.

Testing is not normally required.

AC 29.973A. § 29.973 (Amendment 29-35) FUEL TANK FILLER CONNECTION .

a. Explanation .

(1) Amendment 29-35 revised the requirements for fuel tank filler connections.

paragraph (a) is revised to require that all fuel tank filler connections be made crash resistant in accordance with the requirements of § 29.952(f) and its associated advisory material (reference paragraph AC 29.952).

(2) Paragraph (a)(3) is revised to require that all filler caps remain fuel tight under fuel pressures induced during a survivable impact.

(3) Paragraph (b) is revised to require that all transport category rotorcraft (not just Category A as currently required) have a filler cap cover or filler cap that warns when the cap is not fully locked or seated on the filler connection. This change ensures that a loose filler cap will not allow spilled fuel and cause a postcrash fire in an otherwise survivable accident.

b. Procedures .

(1) The compliance procedures for general paragraph (a) are those of § 29.952(f) and those described herein for the three subparagraphs to (a).

(2) The compliance procedures for (a)(1) and (a)(2) can normally be demonstrated by analysis and physical inspection of the fuel filler design. Testing is not normally required.

(3) The compliance procedures for (a)(3) are as follows: The fuel tank filler connection must be shown to be leak free under the worst case fuel pressures (due to Page E - 117 AC 29-2C 9/30/99 combination of static pressure and sloshing induced head) from both normal operations and from a survivable impact. The worst case loads from these two conditions must be determined. In most cases the load resulting from a survivable impact will prevail. For the survivable impact, normally the worst case combined pressure loading occurs at the time of impact at the fuselage that places the filler tube neck (at the vicinity of the filler cap connection) in a vertical or near vertical attitude. Once the critical load case is determined by analysis, test, or a combination; the fuel tank filler connection (or an approved mockup) can be tested for sealing capability by applying a fluid such as water at the critical pressure at the critical attitude of the tube (with the cap inverted) for a period of at least 5 minutes. If no significant leakage occurs, then compliance has been shown. Significant leakage is defined as leakage in excess of 10 drops per minute at any time during or after the 5-minute test.

(4) Compliance procedures for paragraph (b) are as follows: Visual means, such as placards and alignment marks, and mechanical means, such as detents and locking slots, must both be provided. This is necessary to give both a clear visual and mechanical indication that a filler cap or a filler cap cover is properly installed and fuel tight after each removal and replacement. Visual indications such as alignment marks, that show proper installation should be easily read from a distance of at least 5 feet by anyone making a routine inspection or check.

AC 29.975. § 29.975 FUEL TANK VENTS AND CARBURETOR VAPOR VENTS .

a. Explanation . This section sets forth design requirements that address siphoning of fuel, pressure differentials, moisture accumulation, fumes in personnel compartments, and carburetor vapor vents.

b. Procedures . The design of the vent for the fuel system should be adequate to preclude problems associated with this section. Analysis and/or flight testing may be required to demonstrate this adequacy depending upon the fuel system design. If flight testing is required, the following flight test procedure is one method of verifying proper vent system operation.

(1) Using a rotorcraft with a fuel tank and vent system which conforms to production design specifications, install differential pressure instrumentation to measure the difference between the gas pressure inside each fuel tank expansion space and the air pressure in the cavity or area surrounding the outside of the fuel tank.

(2) Conduct ground and flight tests, recording the differential pressures between the inside and the outside of the fuel tanks. The following conditions should be evaluated: (i) Refueling and defueling (if applicable).

(ii) Level flight to V .

NE Page E - 118 9/30/99 AC 29-2C (iii) Maximum rate of ascent and descent.

(3) Compare the measured differential pressure values with the maximum allowable for the fuel tank design being evaluated. For flexible, bladder-type fuel cells, the pressure inside the tank should not be significantly less than the surrounding pressure to avoid the possibility of collapsing the bladder.

AC 29.975A. § 29.975 (Amendment 29-26) FUEL TANK VENTS AND CARBURETOR VAPOR VENTS .

a. Explanation . Amendment 29-26 adds § 29.975(a)(7) which requires that fuel tank vent systems be designed to minimize fuel spillage and subsequent fire hazards in the event of rollover of the rotorcraft during landing or ground operations.

b. Procedures . All of the policy material pertaining to this section remains in effect with the added requirement that the fuel tank vent system design should minimize spillage of fuel in the vicinity of a potential ignition source in the event of rollover during landing or ground operation.

AC 29.975B. § 29.975 (Amendment 29-35) FUEL TANK VENTS AND CARBURETOR VAPOR VENTS .

a. Explanation . In addition to the current requirements, Amendment 29-35 revises paragraph (a)(7) to add the requirement that the venting system be designed to minimize fuel spillage through the vents to an ignition source in the event of a fully or partially inverted rotorcraft fuselage attitude following a survivable impact. (A survivable impact is defined in paragraph AC 29.952.) Since rotor action on impact and other impact dynamics have been found in numerous cases to cause rollovers or other unusual postcrash attitudes, compliance with this paragraph would significantly mitigate the postcrash fire hazard by minimizing fuel spills through vents to ignition sources when the postcrash attitude of the rotorcraft would allow gravity and/or post impact sloshing induced fuel spills through a normally open fuel vent.

b. Procedures .

(1) In addition to the compliance procedures for the previous amendment; installation of design features, such as gravity activated shuttle valves in the vent lines (that are normally open but close under certain predictable, postcrash scenarios that are generated by involvement in a survivable impact that results in either an inverted or partially inverted fuselage attitude) must be accomplished.

(2) Once selected, the design feature chosen for compliance should be shown to function effectively without significant leakage by either full-scale and/or bench tests that apply the total pressure forces that correspond to a 100 percent full, 50 percent full, and 5 percent full fuel load applied to the device in a worst case survivable impact. (If a critical fuel level can be clearly identified, then only that fuel level and the corresponding Page E - 119 AC 29-2C 9/30/99 critical total pressure load need be utilized for certification approval.) The total pressure forces should be determined and applied in a manner that simulates the magnitude and rate of load onset (due to a combination of gravity and sloshing) that would occur in otherwise survivable impacts that would involve rollover attitudes of 45 degrees (or the minimum spillage roll angle), 90 degrees (rotorcraft on its side), and 180 degrees (rotorcraft fully inverted). (In some designs, the 45-degree attitude may not be the correct initial roll angle at which fuel spillage through a given vent would begin to occur due to the placement of the vents on the fuselage. For these cases, the minimum angle should be determined by analysis.)

(3) Once all test conditions are defined, these tests should be conducted with all structural deformation present in the test set up that is necessary to simulate the actual structural deformation either in or applied to the vent line or system in a worst case survivable impact. The structural deformation to be applied can be determined by rational analysis, analysis, test, or a combination. Significant leakage is defined as leakage of 10 drops per minute, or less, after all testing is complete. The criteria of 10 drops per minute, or less, corresponds to the criteria of 5 drops per minute, or less, per breakaway coupling half (i.e., a total of 10 drops per minute, or less, for the entire separated coupling) specified in the advisory material for § 29.952 (reference paragraph AC 29.952).

AC 29.977. § 29.977 (Amendment 29-12) FUEL TANK OUTLET .

a. Explanation .

(1) This section prescribes a fuel strainer for the fuel tank outlet (suction lift system) or for the booster pump (boosted systems) for both reciprocating and turbine engine installations.

(2) This requirement ensures that relatively large, loose objects which may be present in the fuel tank do not interfere with fuel system operation. The provisions of § 29.997 should ensure protection from smaller contaminants which may occur in service.

b. Procedures .

(1) Section 29.977(a) specifies an 8- to 16-mesh-per-inch strainer for reciprocating engine installations, and a strainer which will prevent passage of any object which could restrict fuel flow or damage any fuel system component for turbine installations.

(2) In addition to the requirement of § 29.977(a), the flow area of the strainer should be at least five times the area of the outlet line. Furthermore, the diameter of the strainer must be at least that of the fuel tank outlet line.

(3) Each finger strainer should be accessible for inspection and cleaning.

Page E - 120 9/30/99 AC 29-2C (4) Compliance with § 29.977 is usually verified by inspection, and testing is not required. The ice protection provisions of § 29.951(c) are applicable to the strainer at the fuel outlet, and testing to show compliance with that provision may be required.

AC 29.979. § 29.979 (Amendment 29-12) PRESSURE REFUELING AND FUELING PROVISIONS BELOW FUEL LEVEL .

a. Explanation .

(1) Each fueling system that has the fueling connection below the fuel level in the tanks must prevent the loss of fuel if the fuel entry valve malfunctions.

(2) For pressure refueling systems, a back-up limiting device must be provided in addition to the primary means for limiting the amount of fuel in the tank.

(3) Components of the pressure fueling and defueling systems must be able to withstand an ultimate load that is 2.0 times the maximum pressure (positive or negative) most likely to occur during fueling or defueling. This requirement provides a level of structural integrity for the pressure fueling and defueling system components in the event a system malfunction occurs, which would result in an overpressurization of the fuel system. The fuel tanks and vents are not included in this requirement.

b. Procedure .

(1) Designs which have the pressure refueling and fueling provisions below the fuel level in each tank must demonstrate that when there is a malfunction of the fuel entry valve, no hazardous quantity of fuel will be lost. Generally, any amount of fuel loss in excess of 8 ounces is considered to be hazardous. Any amount of fuel that can come in contact with an ignition source is hazardous and unacceptable. Compliance should be demonstrated by test and supported by a failure mode and effects analysis.

(2) For pressure refueling systems, one of the most hazardous failure modes is an undetected overpressurization of the fuel tank which could lead to a number of potential fuel system failures. The pressure refueling system must contain a device which insures that fuel tank capacity cannot be exceeded. This device can operate on a differential pressure principle or can sense fluid level. A back-up limiting device is required in case of failure of the primary limiting device. Compliance must be demonstrated by test. A failure mode and effects analysis should be performed which verifies that the failure of either the primary or back-up limiting device will not result in the failure of the other limiting device.

(3) The rotorcraft pressure fueling and defueling systems must be designed to withstand an ultimate internal pressure load that is twice the maximum pressure that is likely to occur during fueling or defueling. The maximum pressure will include surges that could occur from the fueling source and/or from any single tank valve or Page E - 121 AC 29-2C 9/30/99 combination of valves being either intentionally or inadvertently closed. System substantiation may be demonstrated by analysis or test. The substantiation should include all components of the pressure fueling and defueling system except the fuel tank and the fuel tank vents. The rotorcraft defueling system must also be substantiated for a negative pressure application. If tests are conducted, the pressure measurements for both tests (positive and negative) will be made at the fueling connection and the test set-up should conform to the installed system.

Page E - 122 9/30/99 AC 29-2C SUBPART E - POWERPLANT FUEL SYSTEM COMPONENTS AC 29.991. § 29.991 FUEL PUMPS .

a. Explanation .

(1) Section 29.991, paragraph (a) provides a definition of the main pump(s) and § 29.991, paragraph (b) requires an “emergency pump(s).” The main pump(s) that is certified as part of the engine does not fall under § 29.991 requirements. The main pump(s) discussed under § 29.991 should therefore be considered “main aircraft pump(s).” (2) The main aircraft pump(s) consists of whatever pump(s) is required to meet engine or fuel system operation throughout the range of ambient temperature, fuel temperature, fuel pressure, altitude, and fuel types intended for the rotorcraft. If the main aircraft pump(s) is required to meet the above criteria, then an emergency pump(s) is required.

b. Procedures .

(1) Each pump classified as a main aircraft pump, which is also a positive displacement pump, must have provisions for a fuel bypass. An exception is made for fuel injection pumps used on certain reciprocating engines and for the positive displacement, high pressure, fuel pumps routinely used in turbine engines. The bypass may be accomplished via internal spring check valve and fuel passage, or by external plumbing and a check valve. High capacity positive displacement pumps with internal pressure relief and recirculation passages should be checked for overheating if they may be expected to operate continuously at or near 100 percent recirculation.

(2) Section 29.991, paragraph (b) specifies a requirement for “emergency” pumps to provide the necessary fuel after failure of any (one) main aircraft pump.

(Injection pumps and high pressure pumps used on turbine engines are exempt.) As stated in this rule, the “emergency” pump must be operated continuously or started automatically to assure continued normal operation of the engine. For some multiengine rotorcraft, another main aircraft pump may possibly be used as the required “emergency” pump. In this case, the dual role of this pump requires it to have capacity to feed two engines at the critical pressure/flow condition. Availability of fuel flow from this backup pump must be automatic and this function should be verified in the preflight check procedure. For Category A rotorcraft, a comprehensive fault analysis of the fuel system is mandatory to assure compliance with § 29.903, paragraph (b).

Page E - 123 AC 29-2C 9/30/99 (3) Section 29.991, paragraphs (c)(1)(i) and (ii) address the situation, usually associated with supercharged reciprocating engines, where fuel pressure must be modulated with respect to carburetor deck pressure. This is accomplished with interconnecting air lines from the carburetor intake (after the supercharger) to the pressure relief connection on the fuel pump(s). A similar connection from the carburetor intake to the vented side of the fuel pressure gauge is needed to obtain correct fuel pressure reading. These systems may require orifices and/or surge chambers to operate correctly.

(4) Section 29.991, paragraphs (c)(2) and (3) requires seal drains which drain safely. A drain impingement test is normally required to verify safe drainage. Use of a colored dye to simulate fuel discharge at the drain line exit or a fluid sensitive coating (Bon Ami) on the aircraft skins will facilitate evaluation of the safety aspects of drain impingement. Pump seal drain requirements would not be applicable for tank immersed pumps.

AC 29.991A. § 29.991 (Amendment 29-26) FUEL PUMPS .

a. Explanation . Amendment 29-26 revises § 29.991 to clarify fuel pump redundancy requirements. Redundancy for fuel pump failure includes consideration of both the pump and the pump motivating device.

(1) Section 29.991(a)(1) now stipulates that a single fuel pump failure should not jeopardize the capability of the fuel system from delivering the fuel necessary to satisfy the requirements of § 29.955. This stipulation excludes any fuel pumps that are approved as a part of the type certificated engine.

(2) Section 29.991(a)(2) expands the stipulation of § 29.991(a)(1) by including any component(s) required to drive the fuel pump (such as electric motors or generators for electric pumps). This section also stipulates that if the pump is engine driven, failure should not affect more than the engine served by the pump.

b. Procedures . The method of compliance for this section is unchanged.

AC 29.993. § 29.993 FUEL SYSTEM LINES AND FITTINGS .

a. Explanation . This rule outlines design requirements for fuel system lines.

b. Procedures .

(1) Compliance is usually obtained by employing routing and clamping as described in paragraph 709, Chapter 14, Section 2, of AC 43.13-1A and by monitoring the arrangement throughout the developmental and certification test period.

Requirements for approved flexible lines may be resolved by utilizing lines listed as TSO C53a approved for installation in either normal or high temperature areas as appropriate.

Page E - 124 9/30/99 AC 29-2C (2) Verify adequate clearance exists between lines and elements of the rotorcraft control system at extremes of control travel, including control deflections and, for flexible lines (hoses), possible variations in routing.

(3) Flexible lines inside fuel or oil tanks require special evaluation to assure that the external surfaces of these lines are compatible with the fluids involved and that fluid sloshing will not cause line failure. Lines inside tanks should be routed to avoid impingement by fuel or oil filler nozzles.

(4) Good design practice suggests that all flammable fluid lines should be routed to minimize the possibility of rupture in the event of a crash or from engine rotor disc failure.

AC 29.995. § 29.995 (Amendment 29-13) FUEL VALVES .

a. Explanation . This regulation requires that fuel valves be supported so that no loads resulting from their operation or from accelerated flight conditions are transmitted to the lines attached to the valve.

b. Procedures . Compliance with this rule is usually accomplished by designing the installation of the fuel valve so that the valve is supported by either primary or secondary airframe structure.

AC 29.997. § 29.997 (Amendment 29-10) FUEL STRAINER OR FILTER .

a. Explanation . This rule provides for a main in-line fuel filter designed to collect all fuel impurities which could adversely affect fuel system and engine components downstream of the filter. The rule also requires a sediment bowl and drain (or that the bowl be removable for drain purposes) to facilitate separation of contaminations, both solid and liquid, from the fuel.

b. Procedures .

(1) The filter should be mounted in a horizontal segment of the fuel line to facilitate proper action of the sediment bowl. If the filter is located above the fuel tank, it becomes necessary to activate a fuel boost pump to achieve positive drainage of the filter bowl. Without pump pressure, air may enter the fuel system during the filter draining operation and, for turbine engines, result in transient power surges or engine failure during subsequent engine operation. A flight manual note to require pump(s) to be “on” during filter draining would be appropriate.

(2) Section 29.997(d) sets forth a requirement for filter capacity and for filter mesh. The capacity requirement may be substantiated by showing that the filter, when partially blocked by fuel contaminates (to a degree corresponding to the indicator marking or setting required by § 29.1305(a)(17)), does not impair the ability of the fuel Page E - 125 AC 29-2C 9/30/99 system to deliver fuel at pressure and flow values established as minimum limitations for the engine. The filter mesh must be sized to prevent passage of particulate which cannot be tolerated by the engine. FAR Part 33 requires that the degree and type of filtration be established. This information should be the base for selecting the filter mesh. Although a test may be devised and conducted, data from the filter manufacturer usually are acceptable to verify compliance. Note that when the filter capacity is reached, continued flow of contaminated fuel may result in engine failure. A flight manual note regarding precautionary procedures is appropriate.

(3) FAR Part 33 (through Amendment 33-6) has an identical requirement for a fuel filter for engine fuel systems; however, it is not intended that two filters should be required.

AC 29.997A. § 29.997 (Amendment 29-26) FUEL STRAINER OR FILTER .

a. Explanation . Amendment 29-26 requires that a fuel strainer or filter should be installed between the fuel tank outlet and the first fuel system component that is susceptible to fuel contamination. Components that will be protected from contamination include but are not limited to fuel metering devices which control flow rate, fuel heaters, and positive displacement pumps. The amendment also requires a sediment bowl and drain (unless the bowl is readily removable for drain purposes) to facilitate separation of solid and liquid contaminants from the fuel.

b. Procedures .

(1) The fuel strainer or filter should be accessible for draining and cleaning. It should incorporate a screen or other element that is easily removable. It should be mounted so that its weight is not supported by the inlet or outlet connections of the strainer itself, unless it can be shown that adequate strength margins exist in the lines and connections.

(2) The fuel strainer or filter should have a sediment trap and drain (unless the trap is readily removable for drain purposes). The volume capacity of the sediment trap is specified in § 29.971(a) (0.10 percent of the tank capacity or 1/16 of a gallon).

(3) The fuel strainer or filter mesh should provide the filtration stipulated in the FAA/AUTHORITY-approved engine installation manual that is prepared for the type certificated engine (FAR Part 33).

(4) The fuel strainer or filter should have the capability to remove any contaminant that would jeopardize the flow of fuel that is necessary to meet the requirements of § 29.955. In addition, the strainer or filter should have a bypass system with an impending bypass indicator (Refer to § 29.1305(a)(17)). When the strainer or filter is partially blocked with contaminants, to the degree that the fuel flow requirements of § 29.955 can no longer be achieved, the impending bypass indicator should be activated. At this point, the strainer or filter should not yet be bypassing unfiltered fuel.

Page E - 126 9/30/99 AC 29-2C Although a test may be devised and conducted, data from the filter manufacturer usually are acceptable to verify compliance. Note that when the filter capacity is reached, continued flow of contaminated fuel may result in engine failure. A flight manual note regarding precautionary procedures is appropriate.

(5) Section 33.67(b) has an identical requirement for a fuel filter for engine fuel systems; however, it is not intended that two filters should be required.

Page E - 127 6/23/23 AC 29 - 2C, Chg 9 AC 29.997B. § 29.997 (AMENDMENT 29 - 59 ) FUEL STRAINER OR F ILTER.

a. Explanation . Although amendment 29 - 59 did not revise the regulatory text of § 29.997, it did change the numbering sequence of § 29.1305. The AC material for § 29.997 references the fuel filter contamination warning requirement in § 29.1305(a)(17). Beginning with Amendment 29 - 40 this reference should be § 29.1305(a)(18).

b. Procedures . The policy pertaining to the procedures in this se ction remains unchanged, however the reference to the § 29.1305 fuel contamination indicator has been corrected.

Page E - 127.1

9/30/99 AC 29-2C Although a test may be devised and conducted, data from the filter manufacturer usually are acceptable to verify compliance. Note that when the filter capacity is reached, continued flow of contaminated fuel may result in engine failure. A flight manual note regarding precautionary procedures is appropriate.

(5) Section 33.67(b) has an identical requirement for a fuel filter for engine fuel systems; however, it is not intended that two filters should be required.

AC 29.999. § 29.999 (Amendment 29-12) FUEL SYSTEM DRAINS .

a. Explanation . This regulation provides for fuel system drains and defines the requirements which the system must meet.

b. Procedures .

(1) The location and function of the fuel system drains are an integral part of any fuel system. There may be several drains required dependent upon the fuel system design. Each fuel tank sump and certain types of fuel strainers or filters require a means to drain (reference §§ 29.971 and 29.997).

(2) Selection of the location and orientation of the drain discharge in the design phase is important to assure that there is no impingement on any part of the rotorcraft.

To show compliance with the requirement may require tests dependent upon whether the applicant has a previously approved design which is similar, or if the system is a new design for which no previous experience is available.

(3) The location of the drain valve should be selected so that the requirements for accessibility, ease of operation, and protection are met.

(4) Advisory Circular 20-119 provides an acceptable means, but not the only means, of compliance with the requirement for positive locking of fuel drain valves in the closed position.

(5) The fuel drain installation on aircraft with retractable landing gear will be satisfactory if recessed within the outside surface of the aircraft.

AC 29.999A. § 29.999 (Amendment 29-26) FUEL SYSTEM DRAINS .

a. Explanation . Amendment 29-26 adds the requirement that fuel system drains be effective with the rotorcraft in any allowable ground attitude including uneven terrain.

In addition, the change amended § 29.999(b)(2) to require fuel drains have a means to ensure positive closure as contrasted to positive locking when in the “off” position. This will accommodate designs featuring spring-loaded drain closures that have been found to be satisfactory.

Page E - 127.2 AC 29-2C 9/30/99 b. Procedures . All of the policy material pertaining to this section remains in effect. Additionally, selection of the location and orientation of the fuel drain discharge in the design phase is important to assure that there is no impingement upon any part of the rotorcraft. The location and orientation should also ensure effective fuel drainage when the rotorcraft is parked on uneven terrain. To show compliance with the requirement, tests may be required, dependent upon whether the applicant has a previously approved design that is similar, or the system is a new design for which no previous experience is available.

AC 29.1001. § 29.1001 (Amendment 29-26) FUEL JETTISONING .

a. Explanation . Amendment 29-26 adds § 29.1001 to set forth the certification requirements for a fuel jettisoning system if it is installed in the rotorcraft.

b. Procedures . In showing compliance with the requirements of § 29.1001, the following guidance is provided.

(1) The fuel jettison system should be demonstrated to be safe in all normal flight regimes. Takeoff, hover, and in-ground-effect maneuvers may be excluded if appropriate limitations are prescribed.

(2) The fuel jettison system, and its operation, should be shown to be free from fire hazard. If possible, the fuel should discharge clear of any part of the rotorcraft; however, it should be shown that any fumes or fuel, that do impinge upon the rotorcraft in the form of a fine mist, does not form droplets that run along the exterior structure and enter any part of the rotorcraft (wheel wells, cargo area, tail boom, etc.). It should also be shown that jettisoned fuel is not ingested by the engines or the auxiliary power unit (APU). This demonstration can be conducted by jettisoning a glycol based, dye colored fluid and noting the pattern displayed on a dye sensitive coating applied to the rotorcraft exterior. The demonstration should be conducted over all flight regimes in which system operation is permitted. The demonstration should also take into account the maximum rate of descent and all airspeeds where fuel impingement upon the fuselage would most likely occur. Rotorcraft controllability should remain satisfactory throughout the fuel jettisoning operation and should also be demonstrated.

(3) The requirements in § 29.1001(c) were established to prevent complete fuel depletion and provide the capability to effect continued safe flight and landing.

(4) The controls for the fuel jettison system should be designed so that a “minimum” flight crew can perform the jettison operation and be able at any time to stop the jettison process or begin it again. These design requirements give the flight crew the capability and flexibility to manage their on-board resources.

(5) The requirements of § 29.901(c) are intended to emphasize that no single failure or malfunction or probable combination of failures of the fuel jettisoning system will jeopardize the safe operation of the rotorcraft.

Page E - 128 9/30/99 AC 29-2C (6) If the rotorcraft has an auxiliary fuel tank, an auxiliary fuel jettisoning system may be installed to jettison the additional fuel provided the jettisoning system has separate and independent controls and it also meets all of the requirements of this section.

Page E - 129 AC 29-2C, Chg 4 5/1/2014 SUBPART E - POWERPLANT OIL SYSTEM .

AC 29.1011. § 29.1011 ENGINES: GENERAL.

a. Explanation .

(1) The oil system provided for each installed engine should provide all of the lubrication required by the engine and supply it at a temperature which is within the operating temperature limits established for that engine when it was certified.

(2) The usable oil capacity of each oil system should be sufficient to provide oil to the engine at the maximum oil consumption limit of the engine under critical operating conditions. All circulating requirements and operating temperature limits for the oil should be met.

b. Procedures .

(1) There are three basic engine oil supply and cooling system concepts that are used. There are self-contained systems (a complete system certified with the engine), systems that have both engine and airframe components, and systems that are totally supported by airframe components. Any one of these three concepts can be used to meet the requirement of having an independent oil system for each engine.

(2) Oil tank capacity is primarily determined by the engine’s oil consumption rate. Other factors which should be considered when sizing the oil supply system are the endurance of the rotorcraft under critical operating conditions, and the amount of oil circulating in the system to maintain proper cooling. Adequacy of the engine oil supply system can be shown by analysis supported by engine oil consumption and cooling system data. For reciprocating engines, the ratio of one gallon of oil for each 40 gallons of fuel can be used; however, an oil-fuel ratio lower than 1:40 can be used if properly substantiated by oil consumption data on the engine.

(3) The engine oil cooling requirements are defined in §§ 29.1041 through 29.1049. The design of the engine oil cooling system will be influenced by hot day conditions, by the engine heat rejection rate, and other oil system operating data provided by the engine manufacturer. Sizing of the oil cooler will depend upon the engine data and whether the oil cooler will also be used for main transmission oil cooling. Oil cooler size should be kept as small as possible due to its effect on rotorcraft structure, but in all cases, adequate cooling should be demonstrated throughout the operating envelope of the rotorcraft.

Page E - 130 9/30/99 AC 29-2C AC 29.1013. § 29.1013 (Amendment 29-10) OIL TANKS .

a. Explanation . This regulation identifies the requirements that each oil tank must meet. It also specifies that the oil tank installation must meet the installation requirements of § 29.967.

b. Procedures .

(1) The oil tanks usually are constructed of aluminum, aluminum alloy, or stainless steel and are of such a design to permit installation in the aircraft as close to the engine as the design allows. The choice of materials will generally be determined by the selected location of the tank. The tank envelope or outline will generally be determined by the location within the structure of the rotorcraft.

(2) The design of the tank is required to meet the expansion space requirements as specified in the regulation for the particular installation. This is generally accomplished by locating the filler cap in such a manner that the expansion space cannot be inadvertently filled with the rotorcraft in normal ground attitude.

(3) The tank is required to be properly vented and the vent requirements are identified in the regulation.

(4) Unless alternate means are provided, it is good design practice to locate the oil tank with respect to the engine so that when the rotorcraft is in its normal ground attitude, a positive head to the oil pump inlet is provided.

(5) Sections of the regulation address specific requirements when Category A certification is requested.

(6) The designer should be aware of the requirements associated with the location of the oil tank outlet and the marking requirements specified in § 29.1557(c)(2).

(7) Flexible oil tank liners may be used; however, they must be approved or shown to be suitable for the particular installation.

(8) An “external oil system” which is defined as being those components, lines, etc., of an oil system which are outside the engine and not supplied as part of a certificated engine. The components of such a system which are within the fire zone and required to be fire resistant. Those outside the fire zone need not be fire resistant.

AC 29.1015. § 29.1015 (Amendment 29-10) OIL TANK TESTS .

a. Explanation . This regulation defines the tests that must be accomplished to show compliance for rotorcraft oil tanks.

Page E - 131 AC 29-2C 9/30/99 (1) The oil tank should be designed and installed so that it can withstand, without failure, any vibration, inertia, and fluid loads to which it may be subjected in operation.

(2) The installation should meet the requirements of § 29.965 except that for pressurized tanks used with turbine engines, the test pressure may not be less than 5 PSI plus the maximum operating pressure of the tank. For all other tanks, the test pressure may not be less than 5 PSI.

b. Procedures . The pressure tests require that 5 PSI plus operating pressure but in any case no less than 5 PSI be used to substantiate the oil tank. To accomplish these tests, the various tank openings are sealed. An adapter fitting is fabricated by which regulated, pressurized air is introduced into the tank. This air pressure is measured by means of a calibrated air pressure gauge. Any of several methods to determine whether the tank is leaking may be used. As an example, if the tank is relatively small, emergence in a tank of water may be used. Other means such as applying soapy water to the joints are also satisfactory. In any respect, the leak check using test fluid conforming to Federal Specification TT-S-735, Type III, may also be used.

AC 29.1017. § 29.1017 OIL LINES AND FITTINGS .

a. Explanation . This regulation outlines the certification requirements for oil lines and fittings.

b. Procedures . The oil system lines and fittings are required to meet the requirements of § 29.993; therefore, the routing and clamping described in paragraph 709, Chapter 14, Section 2, of AC 43.13-1A may be utilized as guidance for the system design. An evaluation carried out through the development and certification test period will usually surface any problems of interference and/or vibration.

(1) When flexible hoses are used in the lubrication system they must be substantiated. Hoses listed in TSO C53a may be used which would preclude certain substantiation requirements.

(2) Location of the breather lines and discharge should be carefully evaluated to determine that the requirements of this paragraph are followed.

(3) The routing of fluid lines should be such that drooping lines and fluid traps which are undrainable are avoided.

AC 29.1019. § 29.1019 (Amendment 29-10) OIL STRAINER OR FILTER .

a. Explanation . This regulation defines the requirements for the engine oil system strainer or filter. If a strainer or filter which meets the requirements of this paragraph is Page E - 132 9/30/99 AC 29-2C incorporated as part of the type certificated engine, an additional airframe filter is not required.

b. Procedures . This paragraph requires an oil strainer or filter through which all of the oil flows for each turbine engine installation. The strainer or filter should be sized to allow oil flow at the flow rates and within the pressure limits as specified in the engine requirements. The effect of oil at the minimum temperature for which certification is sought should be accounted for.

(1) For each oil strainer or filter required by § 29.1019(a) which has a bypass, the bypass should be sized to allow oil flow at the normal rate through the oil system with the filtration means completely blocked.

(2) For each oil strainer or filter installed per this rule, the capacity must be such that the oil flow and pressure are within the operating limits established for the engine. The mesh requirements are determined by the engine specification for the filtration of particle size and density.

(3) Section 29.1019(a)(3) requires an indicator that will show when the contaminant level of the filtration system, as specified in § 29.1019(a)(2), has been reached. The indicator should signal a contaminant level which has not caused the filter to go into a bypass condition. Consideration should also be given so that the contaminant level at which the indicator is activated is such that the filter would not bypass during a flight time based on full fuel at a cruise condition with the lubricant contaminated to the degree used to show compliance with § 29.1019(a)(2).

(4) An evaluation of the construction and location of the bypass associated with the strainer or filter should be accomplished. The appropriate installation of the filter based on this evaluation would preclude the release of the collected contaminants in the bypass oil flow.

(5) If an oil strainer or filter installed in compliance with this regulation does not have a bypass, there must be a means to connect it to the warning system required in § 29.1305(a)(18). This warning should indicate to the pilot the contamination before it reaches the capacity established in § 29.1019(a)(2). Section 29.1019(b) covers the blocked oil filter requirements associated with reciprocating engine installations. The lubrication system should be such that the normal oil flow will occur with the filter completely blocked.

AC 29.1019A. § 29.1019 (Amendment 29-26) OIL STRAINER OR FILTER .

a. Explanation . Amendment 29-26 relaxes the requirements of § 29.1019(a)(3) from requiring an indicator to indicate the contamination level of oil filters. The rule change allows acceptance of a “means to indicate” the contaminate level to allow a wider range of acceptable methods of compliance.

Page E - 133 AC 29-2C 9/30/99 b. Procedures . Unless the filter is located at the oil tank outlet, § 29.1019(a)(3) requires that the oil strainer or filter have the means to indicate when the contaminant level of the filtration system, as specified in § 29.1019(a)(2), has been reached. If the indicator is installed, it should signal a contaminant level that will allow completion of the flight before the filter reaches a bypass condition. The indicator may be a pop-out button or other maintenance cue that is checked on each preflight inspection.

Page E - 134 6/23/23 AC 29-2C, Chg 9 AC 29.1019B. § 29.1019 (AMENDMENT 29 - 59 ) OIL STRAINER OR FI LTER.

a. Explanation . Amendment 29-59 revised § 29.1019(a)(5) to correct a reference from § 29.1305(a)(18) to § 29.1305(a)(19). When § 29.1305 was updated at Amendment 29-40, the numbering sequence was changed. The oil filter contamination warning, § 29.1305(a)(18), was referenced in § 29.1019(a)(5) but was not updated at Amendment 29-40 to account for the change in numbering sequence.

b. . The policy pertaining to the procedures in this section remains Procedures unchanged, however the reference to the § 29.1305 oil filter contamination indicator has been corrected.

Page E - 134.1

AC 29-2C 9/30/99 AC 29.1021. § 29.1021 OIL SYSTEM DRAINS .

a. Explanation . This regulation requires provisions be provided for safe drainage of the entire oil systems and defines certain requirements for assuring that no inadvertent oil flow occurs from the system provided.

Procedures b. . The design of the oil system must provide a means for safe drainage of the entire oil system. This may require one or more drains dependent upon the design of the system. If a valve is used for this function, it must provide a means for a positive lock in the closed position. The method by which the lock is accomplished may be manual or automatic.

AC 29.1023. § 29.1023 OIL RADIATORS .

Explanation a. . This regulation defines the installation requirements to be considered for oil system radiators.

Procedures. b.

(1) The primary concern with respect to oil radiators is that they are sized to provide the required heat rejection and to provide adequate fluid flow within the prescribed pressure limits.

(2) The structural design of the radiator must consider the system oil pressure requirements and the service involvement of the intended application. The selection of the location of the radiator can have a significant bearing on its ability to withstand the vibration and inertia loads.

(3) If the system design incorporates an air duct to direct the airflow, the effects of a fire as defined in this regulation must be considered.

AC 29.1025. § 29.1025 OIL VALVES .

a. Explanation . This regulation identifies the requirements which oil system valves must meet. In addition to the items specified in this rule, this regulation specifies compliance with the requirements of § 29.1189.

Procedures b. . The closing of the oil shutoffs may not preclude a safe autorotation. Compliance with this requirement is best accomplished in the design Page E - 134.2 9/30/99 AC 29-2C phase. This can be accommodated by proper orientation of the valve and/or system plumbing routing. Another means is to design adequate entrapment of lubricants to provide for the autorotation state. The design of the oil shutoff valve must consider the stop or index provisions of this rule. The installation must be such that the loads specified in the rule are addressed.

AC 29.1027. § 29.1027 (Amendment 29-26) TRANSMISSION AND GEARBOXES: GENERAL .

a. Explanation . Amendment 29-26 adds a new § 29.1027. This new section provides the regulations for rotorcraft transmission and gearbox lubrication systems. It incorporates lubrication system requirements that were removed from § 29.1011 and adds additional lubrication system requirements that were derived from existing engine-oil system requirements. These additional requirements have been adjusted or modified to reflect the needs of transmissions and gearboxes. Transmission and gearbox lubrication system regulations are similar to those for engines; therefore, reference is made to the engine lubrication sections as applicable.

b. Procedures .

(1) The pressure lubrication systems for rotorcraft transmissions and gearboxes should comply with the same requirements as the engine lubrication systems stipulated in § 29.1013 (except §§ 29.1013(b)(1), 29.1015, 29.1017, 29.1021, and 29.1337(d)).

These sections provide the requirements for oil tanks, tank tests, oil lines and fittings, and oil system drains.

(2) Each pressure lubrication system for rotorcraft transmissions and gearboxes should have an oil strainer or filter. The strainer or filter should: (i) Remove any contaminants from the lubricant that may damage the transmission, gearbox, or other drive system component and any contaminants that may impede the lubricant flow to a hazardous degree.

(ii) Be equipped with a means to indicate that the bypass system (required by § 29.1027(b)) is at the point of opening, due to the collection of contaminants on the strainer or filter; and, (iii) Be equipped with a bypass system that will permit lubricant to continue to flow at the normal rate if the strainer or filter is completely blocked. In addition, the bypass system should be designed so that contaminants, that have collected on the filter, will not enter the bypass flow path when the system is in the bypass mode.

(3) Section 29.1027(b)(2) requires a screen at the outlet of each lubricant tank or sump that supplies lubrication to rotor drive systems and rotor drive system components. The screen should remove any object that might obstruct the flow of Page E - 135 AC 29-2C 9/30/99 lubricant to the filter required by § 29.1027(b)(1). The requirements of § 29.1027(b)(1) do not apply to the tank outlet screen.

(4) Splash-type lubrication systems for rotor drive system gearboxes should comply with §§ 29.1021 and 29.1337(d).

Page E - 136 9/30/99 AC 29-2C SUBPART E - POWERPLANT COOLING AC 29.1041. § 29.1041 COOLING - GENERAL .

a. Background .

(1) Few substantive changes have been made to the cooling provision requirements, §§ 29.1041 through 29.1049, since the rules were defined in the Civil Air Regulations, Part 7, effective August 1, 1956. Testing procedures utilized have not precisely followed those rigorously set forth in §§ 29.1045 through 29.1049 as industry and the FAA/AUTHORITY have recognized the need to vary procedures slightly to accomplish the practical test objectives.

(2) In the paragraphs which follow, the cooling regulations will be explained, and in some instances where the regulations provide specific procedures, “alternative procedures” which have been found acceptable in achieving the rule objectives will be presented. The intent of providing those alternative procedures is not to promulgate new regulations, but rather to provide recognized, accepted procedures for compliance with the objective of the current standards.

b. Explanation .

(1) The rotorcraft design should provide for cooling to maintain the temperatures of all powerplant, auxiliary power unit, and power transmission components and fluids within the limitations established for these items.

(2) Cooling provisions should be adequate for shutdown and for water, ground, and flight operating conditions.

(3) The adequacy of the cooling provisions should be demonstrated by flight testing.

c. Procedures .

(1) Detailed procedures for the demonstration of climb, takeoff and climb, and hover cooling are given in §§ 29.1045 through 29.1049. Other test conditions and procedures necessary to demonstrate adequate cooling for water, ground, flight, and shutdown conditions must be negotiated between the applicant and the FAA/AUTHORITY certification engineer. A cooling test proposal which defines the agreed test points and procedures should be prepared well in advance of the official certification testing.

(2) The test conditions selected, in addition to those in §§ 29.1045 through 29.1049, would typically include cruise at various airspeeds and altitudes, shutdown Page E - 137 AC 29-2C 9/30/99 after prolonged hover, and sling load cooling if applicable. One test condition which should be examined, particularly with regard to transmission cooling, is the point of highest multiengine mechanical power at the maximum ambient temperature. This is identified as test point “A” in figure AC 29.1041-1. The selection of test points should be tempered with engineering judgment and based on results from similar aircraft, if such data are available.

(3) In showing compliance with the cooling requirements, the applicant should not be required to exceed rotorcraft established limits (gross weight, drive system torque, measured gas temperature, etc.), aircraft power required, or power available.

The applicant may elect, however, to exceed these limits in order to minimize test points by conservative testing, or to anticipate future growth (increased gross weight, etc.).

(4) The need for a comprehensive cooling test plan prior to certification testing cannot be overemphasized. Highly derated engine installations, the relationship of power required to power available, the use of bleed air devices which would increase the measured gas temperature while aircraft power required remains the same, auxiliary cooling provisions, and the increase in engine temperatures with engine deterioration are factors which could affect the selection of cooling demonstration test points. The following paragraphs will provide some general guidance, but the cooling test plan is the key to a successful program.

Page E - 138 9/30/99 AC 29-2C MUL TI EN GI NE POW ER AVAIL AB LE 20K I- 15K ~ ::::; w z 0 (f) :::) 2' I- X i== _J ~ w 0:: 10K :::) (f) (f) w 0:: 0...

SK POW ER FI GU RE AC 29 .1 041-1 ADCH I ON AL COO LING TE ST POINT Page E - 139 AC 29-2C 9/30/99 AC 29.1043. § 29.1043 (Amendment 29-15) COOLING TESTS .

a. Explanation .

(1) Section 29.1043(a) requires that certain ambient temperature correction factors be applied unless testing is accomplished at the maximum ambient atmospheric temperature prescribed.

(2) No corrected temperatures may exceed established limits.

(3) The statement of § 29.1043(a)(4) which requires that test procedures be in accordance with §§ 29.1045 through 29.1049 does not limit testing to the conditions prescribed in those sections. Section 29.1041(a) and (b) provide the basis for examination of other operating and shutdown conditions.

(4) The maximum ambient atmospheric temperature must be at least 100° F at sea level, lapsed to altitude at a rate of 3.6° F per 1,000 feet pressure altitude. The applicant may select a lower maximum ambient atmospheric temperature for winterization installations.

(5) Unless a more rational correction applies, the temperature data (except for cylinder barrels) are to be corrected by adding the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum component or fluid temperature recorded during the cooling test.

(6) Cylinder barrel temperature data are corrected in a similar manner to other components except 0.7 times the difference between the maximum ambient atmospheric temperature and the ambient temperature at the first occurrence of the maximum cylinder barrel temperature is applied.

b. Procedures.

(1) Seldom is testing actually accomplished at the maximum required ambient temperature of at least 100° F at sea level lapsed 3.6° F per 1,000 feet pressure altitude. Component and fluid temperatures must therefore be corrected to derive the item temperature that would have been reached if the test day had matched exactly the maximum ambient temperature day. The applicant may select a higher maximum ambient temperature for cooling certification than the 100° F sea level hot day prescribed. Provisions are also made for selecting a maximum ambient temperature less than the 100° F sea level hot day for winterization installations not intended to function at the hot day conditions.

(2) When cooling test ambient conditions are cooler than the selected or prescribed hot day conditions, the applicant may take advantage of cooling air or fluid Page E - 140 9/30/99 AC 29-2C flows that would exist at hot day conditions. For example, thermostatically controlled oil cooler flow could be set for hot day conditions.

(3) The component and fluid temperature correction factor to be applied when test ambients do not correspond to the hot day conditions is commonly called the “degree-for-degree correction.” It may be possible to justify, and the regulation allows the application of a more rational, less conservative correction factor. A correction factor other than degree-for-degree should be based on engineering test data.

(4) No corrected temperatures may exceed established limits. In order to maintain temperatures within established limits, the applicant may be willing to accept lesser performance than the full capability of a device. For example, a starter/generator capable of cooling under test cell conditions to 200 amperes continuous load may be limited to a lesser value, perhaps to 150 amperes, when installed in the aircraft due to cooling considerations. This continuous load for cooling must be equal to or greater than the allowable continuous load designated on aircraft instruments.

c. Thermal Limit Correction .

(1) An important correction factor which is not discussed in the regulations, but is frequently necessary to show the cooling adequacy required by § 29.1041, is the thermal limit correction factor. This factor is sometimes required if, at test day conditions, the engine measured gas temperature does not correspond to that which would have occurred on a minimum specification engine at hot day conditions.

(2) The correction factor would not apply to those components not affected by changes in measured gas temperature (MGT) at a constant power. Typical items expected to be affected by changes in the MGT at constant power would be engine oil temperature, thermocouple harnesses, or other fluid, component, or ambient temperatures in the vicinity of the engine hot-section or exhaust gases. Other items remote from the hot-section, perhaps the starter-generator or fuel control, would not be expected to be influenced by MGT variations; however, the items affected and the magnitude of the factor to be applied should be established by testing.

(3) There are several acceptable methods for establishing the appropriate thermal limit correction factor during development testing. The general idea is to establish a stabilized flight condition, typically ground-run or IGE hover, and to vary the measured gas temperature at approximately fixed power and OAT conditions. This may be accomplished by utilizing engine anti-ice bleed air, customer bleed air, or by ingesting warmer than ambient air (either an external source or the engine bleed air) into the engine inlet. Care should be used in ingesting warmer than ambient air to assure that the warm air is diffused in order to avoid possible engine surge.

(i) If it is not possible to attain a suitable variation in MGT by these methods, an acceptable, but more conservative thermal limit correction may be Page E - 141 AC 29-2C 9/30/99 obtained by allowing both shaft horsepower and MGT to vary at a stabilized flight condition and OAT.

(ii) The component temperature is plotted as a function of MGT, and the thermal limit correction from any test day MGT for any flight condition, to the MGT that would have existed with minimum specification engines on a hot day, is then applied to derive the final measured component temperature.

(4) In certain rare instances, it may not be required that the correction factor be applied to the full thermal limit capability of the engine. Consider the following example for the hot day hover IGE cooling test point at sea level.

Corresponding Power (SHP) MGT (°C) Drive System Limit 900 -- - Twin-Engine Hot Day Power Available 1,050 750 Hot Day Power Required at Maximum G.W. 850 650 Engine Maximum Allowable MGT (Instrument Marking) --- 765 Test Day (90° F OAT) Parameters 850 600 (i) Notice that the installed hot day power available MGT from the engine performance program, is 15° C cooler than the limit MGT (750° vs. 765° C), thus the engine has 15° C “field margin” which would allow the engine temperature to gradually increase 15° C to maintain a given power as engine life is utilized. Secondly, the measured gas temperature corresponding to hot day power required at maximum gross weight, is less than that corresponding to either the drive system limit or twin-engine hot day power available. Thus, the thermal limit correction could be applied from the test day MGT, 600° C, to the power required MGT plus the field margin, 650° C plus 15° C, rather than applying the correction factor to the full thermal capability of the engine, 765° C.

(ii) Care should be used in applying this relieving method, because as the hover altitude changes, the maximum gross weight and power required (and the associated MGT) will vary. The data must be corrected to at least the maximum MGT for a minimum specification engine that can occur in service at the flight condition under investigation.

AC 29.1043A. § 29.1043 (Amendment 29-26) COOLING TESTS .

a. Explanation . Amendment 29-26 adds a new paragraph to § 29.1043(a)(5), to define “stabilization” as it pertains to powerplant systems cooling tests.

Procedures b. . All of the policy material pertaining to this section remains in effect with additional information that “stabilized temperatures” are achieved when the rate of change is less than 2° F per minute.

Page E - 142 9/30/99 AC 29-2C AC 29.1045. § 29.1045 CLIMB COOLING TEST PROCEDURES .

a. Objective . The objective of the regulation is to verify, for Category A and for Category B rotorcraft described, that cooling provisions are adequate for a one-engine-inoperative (OEI) climb or descent initiated from a multiengine cruise at the critical altitude with stabilized component temperatures. The specific flight conditions and powers are described in the regulation.

b. Explanation .

(1) This regulation specifies climb or descent cooling with OEI for Category A rotorcraft and for Category B rotorcraft with Category A powerplant isolation and fireproof or isolated structure, controls, etc., which are essential for controlled flight and landing. For the Category B machine described, the testing should be accomplished at the steady rate of climb or descent established under § 29.67(b), i.e., at the best OEI rate of climb (or descent) and the remaining engine at maximum continuous power or 30-minute power, whichever is applicable.

(2) The engine whose shutdown has the most adverse effect on the cooling conditions for the remaining engine(s) and powerplant components should be inoperative.

(3) The regulation provides that the climb cooling test may be conducted in conjunction with the takeoff cooling test of § 29.1047. This possible combining of tests applies only to § 29.1047(a), since § 29.1047(b) is a multiengine climb and not related to the OEI climb procedures of § 29.1045.

c. Procedures .

(1) The OEI climb cooling test point begins from a multiengine cruise, with stabilized fluid and component temperatures, 1,000 feet below either the all-engine-critical altitude or the maximum altitude at which the rate of climb is 150 FPM, whichever is the lowest altitude. If the minimum altitude derived is less than sea level, the climb should begin from a twin engine cruise with stabilized fluid and component temperatures at the minimum practical altitude.

(i) The all-engine-critical altitude is the maximum altitude at which, for the ambient conditions prescribed, it is possible to maintain the multiengine specified power. For example, if for multiengine operations, the transmission maximum continuous torque can be maintained on the hot day to a maximum altitude of 10,000 feet above which power would have to be reduced because of gas temperature or other limitations, then 10,000 feet is the all-engine-critical altitude. Point ”A” in figure AC 29.1045-1 illustrates the all-engine-critical altitude.

Page E - 143 AC 29-2C 9/30/99 (ii) The 150 FPM climb criteria should be based on multiengine operation at maximum continuous power available at hot day conditions at maximum gross weight.

(iii) Fluid and component temperatures are considered stabilized when the rate of change is less than 2° F per minute.

(2) The OEI climb power to be utilized is 30-minute OEI hot day power available (if approval of 30-minute power on the aircraft is requested), followed by maximum continuous hot day power available. If 30-minute OEI power approval is not requested, the power to be utilized would be maximum continuous hot day power available.

(i) Rotorcraft for which approval of a continuous OEI power rating is requested would use the power available on a hot day at the maximum continuous OEI rating following the 30-minute OEI climb phase (or for the entire climb if approval of 30-minute OEI power is not requested).

(ii) If the maximum continuous OEI approval is not requested, then the highest hot day power available approved for continuous usage from the remaining engine(s) under OEI conditions would be used following the 30-minute OEI climb phase (or for the entire climb if approval of 30-minute OEI power is not requested).

(3) In order to achieve representative test results, the rotorcraft climb rate and airspeed should approximate those which would occur on a hot day. This is accomplished by adjusting rotorcraft gross weight as required to produce the desired climb rate based on published or predicted climb performance data. The possible adverse effects of climb fuselage attitude on cooling air duct entrances should be considered in the selection of center-of-gravity of the test aircraft.

(4) The OEI climb should be continued for at least 5 minutes after the occurrence of the highest temperature recorded or until the maximum certification altitude is reached. Generally, temperatures would be expected to peak a short time after the climb begins since component and fluid temperatures are stabilized prior to entry to the climb phase.

(5) For Category B rotorcraft, defined in § 29.1045(a)(2) without a positive OEI rate of climb, the descent should begin from a hot day maximum continuous power multiengine cruise, with stabilized fluid and component temperatures, at the all-engine-critical altitude.

(6) The descent should conclude at either the maximum altitude at which level flight can be maintained with one engine inoperative or at the minimum practical altitude, whichever is higher.

Page E - 144 9/30/99 AC 29-2C (7) The OEI powers available to be utilized during the descent would be the same as those prescribed previously for OEI climb cooling. OEI operation should continue until component and fluid temperatures stabilize.

(8) The airspeeds utilized in the climb and descents should be representative of normal speeds unless cooling provisions are sensitive to rotorcraft airspeed, in which case the airspeeds most critical for cooling should be used. In no case, however, should it be required that the selected airspeeds exceed the speeds established under §§ 29.67(a)(2) and 29.67(b).

Page E - 145 AC 29-2C 9/30/99 MULTI ENGINE POWER AVAILABLE 20K 15K w :::) I- I- _J ~ w CRITICAL ALTITUDE 0:: 10K

---------

:::) (f) (f) w 0:: 0...

t::I 5K 2:

~

~

~

POWER Fl GURE AC 29.1045-1 ALL-ENGINE CRITICAL ALTITUDE ADDITIONAL COOLING TEST POINT Page E - 146 9/30/99 AC 29-2C AC 29.1047. § 29.1047 (Amendment 29-1) TAKEOFF COOLING TEST PROCEDURES .

a. Objective .

(1) For Category A rotorcraft, the objective is to verify satisfactory takeoff and OEI climb cooling for the Category A takeoff profile defined in aircraft performance §§ 29.59(c) and 29.67(a) following a prolonged hover.

(2) For Category B rotorcraft, the objective is to verify satisfactory cooling for the takeoff and subsequent climb for the Category B takeoff defined in performance §§ 29.63 and 29.65(a) following a prolonged hover.

b. Procedure - Category A .

(1) The rotorcraft is hovered in-ground-effect (IGE) at the power required to hover on the test day at the maximum Category A takeoff gross weight for the hot day, until temperatures stabilize.

(i) Alternate Procedure . If the test day OAT is high, it may not be possible to hover IGE at the prescribed gross weight without entering the takeoff range on the measured gas temperature (MGT) indicator. Since operations in the takeoff range are allowed only for 5 minutes and the typical stabilization time is 20 to 35 minutes, it is permissible to reduce the initial aircraft gross weight so the initial MGT will be at least at the MCP limit, but will not be in the takeoff range for more than 5 minutes; and (ii) The fuel burn during the anticipated 20 to 35 minute stabilization period may cause the aircraft to leave the prescribed hover IGE condition unless power is reduced or additional weight is added by fluid transfer or other methods. It is permissible to reduce power to maintain the IGE hover for this phase of testing rather than attempt special weight control procedures.

(2) After temperatures have stabilized, an OEI climb is initiated from the lowest practicable altitude.

(i) Multiengine power may be used from the stabilized IGE hover to the CDP before OEI operations for cooling verification begin.

(ii) Actual shutdown of the simulated failed engine may not be necessary if the applicant can show that cooling of the remaining engine, fluids, and components is not affected by operation of the “failed” engine at idle power.

(iii) The power utilized at the initiation of the OEI climb would be the same as for establishing the takeoff climbout path of § 29.59, typically 2.5-minute OEI hot-day power available.

Page E - 147 AC 29-2C 9/30/99 (3) After the time period for which the power is used in establishing the takeoff climbout path has expired, OEI power is changed to that used in meeting the steady rate of climb (150 FPM, 1,000 feet above the takeoff surface of § 29.67(a)(2)).

(i) The power to be used for this phase is 30-minute OEI hot-day power available, if approval of this power rating for performance is requested.

(ii) If 30-minute OEI approval is not requested, the highest hot-day power available approved for continuous usage under OEI conditions would be utilized.

(4) Climb at the OEI power used in meeting § 29.67(a)(2) would continue for at least- - (i) Thirty minutes if 30–minute OEI power is used; or (ii) Five minutes after the occurrence of the highest temperature recorded, if other than 30-minute OEI is used.

(5) Unlike § 29.1045, the procedure set forth in § 29.1047 for Category A rotorcraft does not specifically require continuation of the OEI climb beyond the 30-minute duration allotted for 30-minute OEI power usage.

c. Procedure - Category B .

(1) The rotorcraft is hovered IGE until temperatures stabilize at the power required on the test day to hover IGE at the maximum Category B takeoff gross weight for the hot day.

(i) Alternate Procedure . If the test day OAT is high, it may not be possible to hover IGE at the prescribed gross weight without entering the takeoff range on the MGT indicator. Since operation in the takeoff range is allowed only for 5 minutes and the typical stabilization time is 20 to 35 minutes, it is permissible to reduce the initial aircraft gross weight so the initial MGT will be at least at the MCP limit, but will not be in the takeoff range for more than 5 minutes; and (ii) The fuel burn during the anticipated 20 to 35 minute stabilization period may cause the aircraft to leave the prescribed hover IGE condition unless power is reduced or additional weight is added by fluid transfer or other methods. It is permissible to reduce power to maintain the IGE hover for this phase of testing rather than attempt special weight control procedures.

(2) After temperatures have stabilized in hover IGE, a multiengine climb is initiated at hot-day takeoff power available from the lowest practicable altitude.

Section 29.1047(b)(3) requires only that takeoff power be maintained for the same time interval as used in determining the takeoff flight path under § 29.63. This time interval Page E - 148 9/30/99 AC 29-2C could be less than the 5 minutes for which takeoff power is approved. Unless the applicant can show that the time interval used in § 29.63 provides more conservative results, or unless additional testing is proposed, the full 5 minutes allowed for takeoff power should be used to assure that the most critical condition has been surveyed.

(3) After the use of takeoff power for the appropriate time interval, the power should be reduced to multiengine maximum continuous hot-day power available and the climb continued until at least 5 minutes after the occurrence of the highest temperature recorded.

(4) The airspeeds utilized in the climb should be representative of normal speeds unless cooling provisions are sensitive to rotorcraft airspeed, in which case the airspeed most critical for cooling should be used. The airspeed need not exceed the speed for best rate of climb with maximum continuous power available.

AC 29.1049. § 29.1049 HOVERING COOLING TEST PROCEDURES .

a. Objective . The objective is to verify satisfactory hover IGE cooling at sea level and at the hover ceiling for hot-day conditions.

b. Explanation . The rule provides for a hover IGE cooling check in still air at sea level and at the hover ceiling at maximum continuous power. Still air is interpreted as a wind speed of 5 knots or less.

c. Procedures .

(1) The aircraft should be hovered IGE at the maximum certificated hover weight or at the IGE hover weight corresponding to hot-day maximum continuous power available, whichever is less.

(i) The power utilized would normally be hot-day maximum continuous power available and the initial gross weight would be selected as required to achieve hover IGE on the test day.

(ii) After initiation of the hover, special weight control procedures need not be implemented in attempting to maintain hover IGE as fuel burn-off occurs. The power may be gradually reduced to maintain the IGE hover condition.

(2) The hover test is to continue until at least 5 minutes after the occurrence of the highest temperature recorded.

(3) Section 29.1049 also requires a hover IGE at the maximum continuous power available at the altitude resulting in zero rate of climb.

Page E - 149 AC 29-2C 9/30/99 (i) Often, compliance is illustrated by extrapolating component cooling margins from sea level test results and from selected altitude test site results to the altitude resulting in zero rate of climb.

(ii) Considerable engineering judgment must be exercised in utilizing the extrapolation method described. In general, if test data is extrapolated more than 2,000 feet to the hover ceiling from the highest altitude site selected and the resulting component margin is less than 5° F, additional verification at altitude may be required.

Page E - 150 9/30/99 AC 29-2C SUBPART E - POWERPLANT INDUCTION SYSTEM AC 29.1091. § 29.1091 (Amendment 29-17) AIR INDUCTION.

a. Explanation .

(1) The air induction system for each engine and auxiliary power unit must supply the air required under the operating conditions for which certification is requested. For reciprocating engine installations, the system must provide air that is suitable for proper fuel metering and mixture distribution. This should be shown with the induction system valves in any position.

(2) The intake system shall be designed such that a backfire flame will not constitute a fire hazard within the engine accessory compartment or within other areas of the powerplant compartment.

(3) Each reciprocating engine must have an alternate air source which must be located to prevent entrance of rain, ice, or other foreign matter.

(4) For rotorcraft powered by turbine engines and rotorcraft incorporating auxiliary power units, there must be means to prevent leakage of hazardous amounts of flammable fluids from entering the engine or auxiliary power unit intake system.

(5) Also, the air ducts must be located or protected to minimize the ingestion of foreign matter during takeoff, landing, and taxiing.

b. Procedures .

(1) For turbine-engine installation, the induction system should supply air of suitable quality to meet the installation requirements of the engine manufacturer. The installation requirements should be met throughout the operating envelope of the rotorcraft. In addition, the design and location of the air induction system should prevent accumulations of rain or hail, either external or internal to the induction system, that could adversely affect engine operation.

(2) The inlet design should account for the prevention of hazardous fluids entering the engine. Some designs will have inlet ducts which are free from any fluid lines; however, other designs may route the engine inlet air through a compartment which has flammable fluid lines. When this condition exists, test demonstrations of critical leakage during operation have been used to substantiate the installation. The fluid leakage may not have an adverse effect on engine operation.

(3) The air induction system design should also account for and minimize the possibility of foreign matter ingestion during takeoff, landing, and taxiing.

Page E - 151 AC 29-2C, Chg 4 5/1/2014 (4) For reciprocating engine installations, the induction system should supply air of suitable quality and quantity to the combustion system of the engine. The condition of this air at the entering face of the carburetor is extremely important. For proper operation, it is essential that the airflow be smooth and uniform, clean, and unrestricted throughout the very wide range of horsepower expected from the engine.

AC 29.1093. § 29.1093 (Amendment 29-22) INDUCTION SYSTEM ICING PROTECTION.

a. Reciprocating Engines . No advisory material is presented here for reciprocating engines since it is unlikely that these types will be used in transport rotorcraft.

b. Turbine Engines - Ice Protection .

(1) Explanation .

(i) This rule requires turbine engines and turbine-engine inlets to perform satisfactorily in atmospheric icing conditions defined in Appendix C of part 29. On an equivalent safety basis, the limited icing envelopes described in section 29.1419 of this AC may be used to show compliance with the intent of the regulation if the rotorcraft is limited to not greater than a 10,000-foot pressure altitude for all operations. If operations are permitted above 10,000 feet, the Appendix C, part 29, envelope must be used from 10,000 feet to the service ceiling or 22,000 feet. These possible equivalent safety approaches are not discussed herein. Compliance with the induction system icing protection rule is required regardless of flight manual limitations or restrictions against flight into atmospheric icing conditions.

(ii) In showing compliance with § 29.1093(b)(1)(i), the FAA/AUTHORITY has accepted the concept of limited exposure associated with escape from inadvertent ice encounters.

(A) It is presumed that there will be a flight manual limitation against flight into icing conditions, and that the engine induction system will be reevaluated if total aircraft ice protection certification is requested. Under this concept, the rotorcraft is assumed to fly directly through the icing environment (i.e., direct sequential penetration and straight line exit from both the continuous maximum and intermittent maximum icing clouds). Thus, the duration of exposure to the icing environment could be calculated by knowing the aircraft flight speed and cloud horizontal extent. A range of engine power and rotorcraft airspeeds should be evaluated to encompass the operating envelope of the rotorcraft.

(B) When this limited exposure concept is used, the aircraft type certificate data sheet should clearly specify that the engine induction system must be reevaluated if certification to the general ice protection regulation, § 29.1419, is requested. This Page E - 152 5/1/2014 AC 29-2C, Chg 4 direct penetration and exit approach is inappropriate for aircraft for which full icing clearance is requested (reference § 29.1419).

(iii) Engine induction system continuous icing protection would be necessary for aircraft for which full-icing clearance is requested (reference § 29.1419(d)). The approach is much preferred for all programs in order to reduce the scope of any eventual total aircraft icing program effort and to increase the safety level in conducting the rotorcraft natural icing tests. Since some rotorcraft have been FAA/AUTHORITY certificated to operate in icing conditions, applicants may request full-icing clearance and, as a result, must demonstrate that the engine induction system will operate in a continuous icing environment.

(iv) It is noted in section 29.1419 of this AC that some natural icing tests are required to show compliance with the overall rotorcraft ice protection requirements. It is not required that the engine induction system be evaluated as a part of that natural icing test if adequate verification has been shown by tunnel testing, analysis, or other means to assure satisfactory operation in an extended continuous icing environment. If, however, subsequent rotorcraft natural icing testing shows unanticipated detrimental engine inlet effects, the inlet ice protection system should be reexamined.

(v) The regulation specifies the examination of flight idling conditions. This requirement is normally associated with a low-power letdown at the minimum practical forward airspeed. Alternatively, evaluation of the minimum power and minimum airspeed combination specified in the rotorcraft flight manual (RFM) for operation in visible moisture when below 41°F (+5°C) will accomplish the intent of the idling requirement.

(vi) An acceptable approach to a finding of compliance would be a combination of analysis of the performance of the ice protection system, which covers the range of the applicable icing flight envelope (maximum altitude, minimum temperature, etc., of the basic rotorcraft) supported and validated by tests. Ideally, these tests would be conducted in natural atmospheric ice with special instrumentation for droplet size and liquid water content. In practice, however, natural icing testing may pose unacceptably severe problems since rotorcraft may not have the range and speed to reasonably find icing clouds and may not be equipped with the airframe and rotor ice protection needed for safety during the testing.

(vii) Problems with analysis emerge if engine inlets incorporate screens, turning vanes, sideward or upward openings, and edge or lip configurations, which deviate from the airfoil shapes assumed in most of the analytical procedures described in current technical literature. The applicant should recognize that if meaningful analytical methods are not available, extensive testing with significant conservatism or possibly design changes may be required. Inlet screens in particular, if not adequately heated, fall in this category and can only be accepted if shown by very conservative ice testing to not significantly impede airflow to the engine.

Page E - 153 AC 29-2C, Chg 4 5/1/2014 (2) Procedures.

(i) Review section 29.1419 of this AC, ADS-4, Report No. FAA-RD-77-76, and the current version of AC 20-73. These data provide extensive description and methodology for evaluation of ice protection systems, however, as noted above, these data generally apply to near straight line droplet trajectory with impingement onto conventional airfoil shaped inlets. As such, the applicability of these data to rotorcraft engine inlet ducts is limited and may require extensive adjustment to accommodate the different inflow trajectories and shapes of rotorcraft.

(ii) An analysis, appropriate to the configuration; i.e., heated or unheated impingement surfaces, should be prepared. To be acceptable, this analysis should show the inlet to be adequately protected by heat, or if unheated, to show that the inlet with ice accretions as predicted, will provide adequate airflow to the engine throughout the flight envelope of the rotorcraft.

(A) For heated surfaces, ADS-4 and Report No. FAA-RD-77-76 provide detailed suggestions on heat transfer analysis particularly applicable to bleed air heated inlet lips formed in airfoil shapes. These data are limited in applicability and may not be useful for analyzing engine inlet water droplet trajectories to be expected at low airspeed and high engine airflow. Actual icing tests may be needed to derive the impingement patterns for these conditions.

( 1 ) Acceptability criteria for heated inlet ducts usually require sufficient heat to evaporate the water to be expected in a “continuous maximum” icing cloud and to anti-ice the duct during flight in “intermittent maximum” icing clouds, provided the run-back and refreeze to be expected does not cause additional airflow disruption or damage to the engine. Full-scale inlet icing tests with the engine installed and operating should be conducted to verify the analysis. Engine power changes, which may be expected in service should be included in the testing. Wind tunnels equipped for icing tests probably are the most useful means of conducting these tests if natural icing tests are impractical. The rotor downwash effect should be considered to the extent possible by adjusting the inflow angle in the tunnel.

( 2 ) The power loss (bleed air, generator load, etc.) attributable to the heating requirements will affect the performance of the rotorcraft. Normally, this may be accounted for by specifying a gross weight incremental deduction from the flight manual performance data for flight into visible moisture below 41°F (+5°C).

( 3 ) Special evaluation of the possibility of ice ingestion damage to the engine should be made for heated systems, which considers the ice ingestion to be expected when the anti-ice system is actuated after a delay of 1 minute for the pilot to recognize that the rotorcraft has encountered ice. This time delay may be reduced if the crew is provided adequate distinctive cues to alert them that the rotorcraft has encountered icing conditions.

Page E - 154 5/1/2014 AC 29-2C, Chg 4 (B) For unheated inlets, an acceptable method for showing compliance would include an extensive, detailed analysis (which shows that ice accretions on and in the inlet do not obstruct adequate airflow to the engine) and tests as necessary to validate the analysis. The analysis of ice accretion becomes even more questionable since the unheated inlet involves ice buildups which themselves progressively change shape during icing exposure.

( 1 ) Flight testing with an instrumented rotorcraft in natural ice to verify the analysis is desirable; however, wind tunnel tests as discussed above may be used.

Since unheated inlets typically continue to accrete ice as a function of exposure, both the analysis and the test should realistically consider the actual exposure to be expected in service. This should not be less than penetration of the continuous maximum icing cloud followed immediately by exposure to the intermittent maximum cloud for rotorcraft not certified for icing. Engine power changes which may be expected in service should be included in the testing, and a warm-up period at the conclusion of the icing exposure should be shown for some selected test points to evaluate potential ice breakaway and ingestion.

( 2 ) For the non-icing certified rotorcraft using the limited icing exposure concept for inlet certification, some conservatism should be applied to account for the fact that inlet icing may occur without airframe icing, and that the escape procedure from this unapproved operating condition is not defined. A demonstration of 30-minute hold capability in the continuous maximum cloud would be acceptable. Alternatively, if positive cues (perhaps a carefully located ice detector) of potential inlet icing are provided to the crew, the time increment could be reduced to recognition plus 15 minutes (15-minute escape time after recognition is consistent with the single ice protection system failure recognition and escape guidance for aircraft ice protection systems in section 29.1499 of this AC). It should not be assumed that airframe icing will always be available as a cue to potential inlet icing. The main rotor, for example, may not show icing indications above 25°F (-4°C), whereas some inlets may ice critically near 32°F (0°C) ambient. A reduction of the acceptable 30-minute exposure should not be based on observation of ice accretions on protruding components which are likely to be changed. For example, a limited exposure inlet icing program which reduces the inlet icing exposure time based on crew recognition of icing on the windshield wipers may be invalidated at a later date if a new windscreen deletes the wipers.

(iii) Inlet capability during IGE hover in icing conditions has not generally been considered for rotorcraft not certified for icing. However, the FAA/AUTHORITY is aware that some inlets may ice at zero airspeed near 32°F (0°C) with no indications of airframe icing in the field of view of the crew. This special concern of operating within RFM limitations, and yet placing the induction system in jeopardy, may be addressed in several ways. If the induction system ice protection scheme is not dependent on airspeed for proper function, the issue may be addressed by tunnel testing with inlet airflows approximating hover with no particular attention to tunnel windspeed. For protection schemes which may be sensitive to airspeed (external screens have shown Page E - 155 AC 29-2C, Chg 4 5/1/2014 this tendency), actual hover demonstration at or near zero speed tunnel conditions may be appropriate. Icing detectors located to indicate induction system icing in hover may be an option to a hover icing protection demonstration. On an external screened configuration, the FAA/AUTHORITY has accepted a satisfactory IGE hover demonstration of 30 minutes at the critical ambient temperature (i.e., ambient consistent with no airframe icing but potential inlet icing), 0.6 grams/meter LWC and 40 micron droplet size as an adequate response to this concern.

(iv) For aircraft requesting full icing approval, or for those electing to show continuous induction system icing protection, the forward flight icing exposure would not be less than that time required to stabilize any ice accretions observed during repeated cycles of the continuous maximum followed by intermittent maximum cloud exposure.

Typically, any ice accretions resulting from these repeated cycles would be expected to stabilize in less than 30 minutes. The 30-minute hold capability in the continuous maximum icing environment could thus be assured without special testing by careful selection of the test points for this repeated cycle.

(v) A rotorcraft requesting full icing approval should also have hover capability in the icing environment. Intermittent maximum icing conditions are not likely to exist near ground level and a satisfactory demonstration could involve the ability to hover indefinitely in the continuous maximum icing environment. Alternatively, carefully worded RFM limitations to restrict hover time may be acceptable if the system is not capable of indefinite exposure. Hover capability verification may not involve zero airspeed demonstration if the inlet protection system is insensitive to rotorcraft airspeed.

(vi) The engine(s) must be installed or protected to avoid engine damage from ice ingestion due to ice accretion in the inlet or on other parts of the rotorcraft, including the rotors, which may break away to enter the inlet. If screens or bypass arrangements are provided for these purposes, they should be included in the icing tests and shown by test or rational analysis to effectively protect the engine.

(vii) For unheated inlets, significant ice accumulations to be expected on the inlet may adversely affect the engine stall margin, acceleration characteristics, duct loss, etc. Dry air flight tests to evaluate these aspects can be accomplished by affixing ice shapes to the inlet. These shapes should closely match the actual ice shapes defined by test or analysis.

c. Turbine Engines - Snow Protection.

Explanation. (1) (i) Section 29.1093(b)(1)(ii) provides that the turbine engine and its air inlet system operate satisfactorily within the limitations established for the rotorcraft, in both falling and blowing snow. The section does not provide the definition of falling and blowing snow.

Page E - 156 5/1/2014 AC 29-2C, Chg 4 (ii) Since the regulation provides for certification “within the limitations established for the rotorcraft,” the FAA/AUTHORITY can accept a restriction against snow operations in the limitations section of the RFM in lieu of demonstration of compliance to the Full Falling & Blowing Snow Conditions defined below in paragraph c.(2)(i).

(A) If an applicant elects not to demonstrate compliance to the FULL falling and blowing snow conditions (i.e., seeks a restriction against snow operations), it can either: ( 1 ) demonstrate that the aircraft turbine engine(s) and its inlet system(s) will operate satisfactorily in the Inadvertent Falling & Blowing Snow Conditions, defined below in paragraph c.(2)(ii), with a restriction for snow operations.

This approach will not require that a minimum operational temperature limit of 41°F (+5°C) be included in the flight manual; or ( 2 ) include a flight manual limitation for minimum operational temperature of 41°F (+5°C).

(B) If no restriction on snow operations appears in the RFM, it is presumed that the aircraft may operate in snow at the pilot’s discretion.

(2) Guidance .

(i) Engine induction system operation in falling and blowing snow can be approved without restriction if normal operations under the following conditions are demonstrated: FULL FALLING & BLOWING SNOW CONDITIONS Visibility: ¼-mile or less as limited by snow.

Temperature: 25°F (-4°C) to 34°F (+1°C) [28°F (-2°C) to 34°F (+1°C) desired], unless other temperatures are deemed critical.

Operations: Ground operations - 20 minutes.

IGE hover - 5 minutes.

Level flight - 1 hour.

Descent and landing.

(ii) Demonstration to the below Inadvertent Falling & Blowing Snow Conditions with a flight manual limitation that prohibits flight into falling and blowing snow is acceptable. This approach will not require that a minimum operational temperature limitation of 41°F (+5°C) be included in the flight manual.

Page E - 157 AC 29-2C, Chg 4 5/1/2014 INADVERTENT FALLING AND BLOWING SNOW CONDITIONS Visibility: 1 mile or less as limited by snow.

Temperature: 25°F (-4°C) to 34°F (+1°C) [28°F (-2°C) to 34°F (+1°C) desired], unless other temperatures are deemed critical.

Operations: Ground operations - 5 minutes.

IGE hover - 1 minutes.

Level flight - 10 minutes.

Descent and landing.

(iii) RFM visibility restrictions for falling and blowing snow operations are not appropriate.

(iv) Time limitations, other than possibly for ground and hover operations, are not appropriate.

(v) Artificially produced snow should not be used as the sole means of showing compliance.

(3) Guidance Rationale.

(i) The test conditions specified--visibility, temperature, and operations--are based on previous certification programs, previous guidance, and on research by the FAA technical center and others.

Visibility (A) . The test visibility defined, ¼-mile (Full Falling & Blowing Snow Conditions) visibility or less as limited by snow, represents a heavy snowstorm and is the maximum likely to be encountered in service. Rotorcraft, which have been certified to the ¼-mile visibility test criteria, have not shown engine inlet snow-related service difficulties. It is important to note that the visibility specified is a test parameter rather than an operational limitation to be imposed on the rotorcraft after the tests are completed.

(B) Temperature .

1 ( ) The ambient temperature specified is conducive to wet snow ow tends to accumulate on unheated surfaces subject to conditions. Wet sn impingement.

( ) Colder ambients, more conducive to dry snow conditions, may be 2 critical for some indu ction systems. Colder exterior surfaces may be bypassed, and the snow crystals may stick to partially heated interior surfaces where partial melting and refreezing may occur.

Page E - 158 5/1/2014 AC 29-2C, Chg 4 ( ) Company development testing or experience with very similar type 3 induction systems may be adequate to determine the critical ambient conditions for certification testing.

(C) Operations .

( ) Ground running, taxiing, and IGE hover operations are generally 1 the most critical sin ce the rotorcraft may be operating in recirculating snow. Twenty-five minutes under these extreme conditions would seem a reasonable maximum, both from the view of pilot stress and the maximum expected taxi time prior to takeoff in bad weather.

( 2 ) One hour of level flight operation under ¼-mile visibility snow conditions should provide ample opportunity for hazardous accumulations to begin to build.

( 3 ) The descent and landing will provide an engine power change, an induction system airflow change, and a variation in the external airflow pattern near the induction system entrance. The initiation of the descent and final flare for landing may also produce additional airframe vibration transmitted to the induction system. These power, airflow, and vibration changes may provide an opportunity for any level flight accumulations to be ingested into the engine. Hazardous accumulations are not acceptable during or after any test phase.

(ii) Visibility may fluctuate rapidly in snowstorms. It is affected by the presence of fog or ice crystals, is not crew measured or controlled, and is difficult to estimate. A visibility operational limitation based on snow, therefore, is not appropriate.

(iii) Since during cruise in snow conditions the aircraft is likely to be in and out of heavy snowfall, it is not practical for the crew to account for the time spent in snow in level flight conditions. Thus, it is not appropriate to include time limitations in the RFM for level flight snow operations.

(iv) A practical ground and IGE hover time limitation of less than 25 minutes in recirculating snow may be considered. The expected action at the expiration of this specified time period would be shut down and inspection of the inlet system or transition to a safe flight condition where demonstration has shown that moisture accumulations will not intensify or shed and cause engine operational problems.

(v) Artificially produced snow is an excellent development tool and has been successfully used to indicate potential problem areas in induction systems. These devices are usually restricted to use for hover and ground evaluations, and the snow pellets produced by these machines are not sufficiently similar to natural snowflakes to justify the use of artificial snow as the sole basis of certification.

(4) Procedures .

Page E - 159 AC 29-2C, Chg 4 5/1/2014 (i) Satisfactory demonstration of the test conditions requires that the engine, induction system, and proximate cowling surfaces remain free of excessive snow, ice, or water accumulation. Excessive accumulation is defined as accumulation that may cause engine instability, damage, or significant loss of engine power. If a questionable amount of snow or moisture accumulates in the inlet, the applicant may elect to demonstrate that this amount in the form of snow or water and ice, as appropriate, can be ingested by the engine without incurring surge, flameout, or damage.

(ii) The conditions specified assume actual flight demonstration in natural snow. The ground operations and IGE hover test conditions assume operation in recirculating snow. Blowing snow, resulting from rotor airflow recirculation, can be expected to be more severe than natural blowing snow if the rotorcraft continues to move slowly over freshly fallen snow. Thus, the blowing snow operational capability is usually demonstrated by the taxi and hover operations in recirculating snow.

(iii) For VFR rotorcraft, the airspeeds for the level flight test condition should include the maximum consistent with the visibility conditions. For IFR operations, the airspeed should be the maximum cruise speed or the maximum speed specified for snow operations in the flight manual limitations, unless other airspeeds are deemed more critical. It is recognized that many rotorcraft initially certified VFR are later IFR certified with a resulting possible increase in airspeed in snow conditions. This factor should be considered if IFR certification is anticipated.

(iv) The visibility specified assumes that visual measurements are made in falling snow in the absence of fog or recirculating snow by an observer at the test site outside the tests rotorcraft’s area of influence. An accepted equation for relating this 0.7734 measured visibility to snow concentration is V = 374.9/C where C is the snow concentration (grams/meter ) and V is the visibility (meters).

(A) This equation can be reasonably applied to all snowflake type classifications and is credited to J.R. Stallabrass, National Research Council of Canada.

(B) Other equations may be applied if they are shown to be accurate for the particular snowflake types for the test program.

(v) The snow concentration corresponding to the visibility prescribed, ¼-mile or less, will be extremely difficult to locate in nature. Data from Ottawa, Canada, research indicate that fewer than 4 percent of the snowstorms encountered there meet the 0.91 grams/m concentration associated with the ¼-mile visibility. Furthermore, the likelihood that the desired concentration will exist for the duration of the testing is even more remote. Because of these testing realities, it is very likely that exact target test conditions will not be achieved. Those involved in certification must exercise good judgment in accepting alternate approaches.

Page E - 160 5/1/2014 AC 29-2C, Chg 4 (vi) For some engine induction systems, it may become apparent by inspecting for moisture accumulations that ground and IGE hover operations in recirculating snow are much more severe than the level flight test. In this instance, it is reasonable to accept prolonged IGE operations in recirculating snow and to accept durations of less than 1-hour level flight in ¼-mile or less visibility. Best efforts should be made to ensure that at least some level flight time is accomplished at ¼-mile or less visibility to ensure that the spectrum is covered.

(vii) It should be determined that the visibility established at the test sight is limited by snow and not by fog or poor lighting (twilight) conditions.

(viii) The concentration of snow approaching the inlet in severe recirculation will far exceed the quantity encountered in the natural snowfall. Recirculation is necessarily a qualitative judgment by the test pilot. The snow concentration at the inlets during recirculation would vary for different rotorcraft types and would be dependent on rotor characteristics, power setting, and inlet location. For test purposes, recirculation should be the highest snow concentration attainable in the maneuver, or that corresponding to the lowest visibility at which (in the pilot’s judgment) control of the rotorcraft is possible in the IGE condition. The visibility specification of ¼-mile or less outside of the recirculation influence becomes inconsequential provided that fresh, loose snow is continually experienced during the ground operation and IGE hover testing phase. However, since it is intended that the test phases be accomplished sequentially to ensure that transition to takeoff and other transients are considered, the conditions at takeoff, level flight, and descent and landing should approximate the ¼-mile visibility criteria.

d. Turbine Engines - Ground Icing.

(1) Explanation . This requirement addresses the situation where extended ground operation in icing exposes the rotorcraft and its engine inlet to icing (ground fog) conditions which may have different droplet impingement patterns and involve different or less effective means of ice protection. Note that the requirement is effective at Amendment 10 and is applicable regardless of any desire to prohibit dispatch into icing conditions.

Procedure (2) . Since this condition assumes zero airspeed, wind tunnel testing may be inappropriate unless conservative extrapolation of low speed tunnel data can be determined to be valid. For protection schemes which are dependent primarily on airspeed for proper functions (external screens have shown this tendency), it may be necessary to verify adequate ground operation protection capability by very low speed tunnels or by the use of outside facilities such as the Canadian National Research Council’s spray rig at Ottawa, Canada. For heated systems or for internal bypass schemes, tunnel speed may not be important, and adequate demonstration may be accomplished at higher tunnel speeds provided that internal inlet airflows and heat available are properly considered. Testing should proximate the regulatory test conditions and be continued for 30 minutes using engine power and control Page E – 160.1 AC 29-2C, Chg 4 5/1/2014 manipulation as normally accepted during taxiway operations, followed by an acceleration to takeoff power. The test time may be shortened if de-ice or anti-ice protection is adequate or if stabilization of ice build-up is affirmed. The induction system should be in condition for safe flight at the conclusion of the test.

AC 29.1093A. § 29.1093 (Amendment 29-26) INDUCTION SYSTEM ICING PROTECTION.

a. Explanation . Amendment 29-26 clarifies that the phrase, “within the limitations established for the rotorcraft” applies only to the requirement in § 29.1093(b)(1)(ii) for demonstrating flight in falling and blowing snow.

Procedures b. . All of the policy material for this section remains in effect with the update that turbine engines and turbine engine inlets should perform satisfactorily in atmospheric icing conditions defined in Appendix C of part 29. In addition to section 29.1093 of this AC, the following procedures should be followed: (1) A “serious loss of power” in this section has been interpreted to be any power loss that requires immediate pilot action. In addition, the term “adverse effect on engine operation” in § 29.1093(b)(1)(ii) has been interpreted to be an effect that would prevent the engine from achieving rated aircraft flight manual performance (takeoff, climb, etc.).

This term also includes effects on the engine induction system characteristics to an acceptable level established by the engine manufacturer (inlet distortion, etc.).

(2) The applicant should show that rotorcraft prohibited from flight into falling and blowing snow can exit inadvertent entrance into those conditions without adverse effect upon the operating characteristics of the engine or the rotorcraft. This requires that the engine(s) and its inlet system be shown to operate satisfactorily throughout the flight power range of the engine and within the operating limitations of the rotorcraft during operation in the Inadvertent 1 mile visibility falling and blowing snow conditions defined herein.

(3) For unrestricted flight capability into Full snow conditions, both falling and blowing, the applicant should show that each engine, and its inlet system, will operate satisfactorily throughout the flight power range of the engine and within the operating limitations of the rotorcraft. The applicant should show that any build-up or accumulation of snow will not reduce or block the flow of inlet air to the engine. Any accumulations that become dislodged should not affect engine operation.

(4) If a design is not satisfactorily demonstrated to either the Full or the Inadvertent snow conditions, a limitation must be included within the flight manual prohibiting flight in temperatures below 41°F (+5°C).

Page E – 160.2 5/1/2014 AC 29-2C, Chg 4 AC 29.1101. § 29.1101 CARBURETOR AIR PREHEATER DESIGN.

a. Explanation . Each carburetor air preheater must be designed and constructed to: (1) Ensure ventilation of the preheater when the engine is operated in cold air.

(2) Allow inspection of the exhaust manifold that it surrounds.

(3) Allow inspection of critical parts of the preheater itself.

b. Procedures . Although carburetors of some design and fuel injections are free from icing difficulties, the most common remedy is to preheat the air supply entering the carburetor. In this way, sufficient heat is added to replace the heat lost due to vaporization of fuel, and the mixing chamber temperature cannot drop to the freezing point of water. The air preheater is essentially a tube or jacket through which the exhaust of one or more cylinders is passed with the air flowing over the heated surface raised to the required temperature before entering the carburetor. A control for adjusting the preheater valve is installed in the cockpit so that heat may be applied only when actually required to prevent ice formation.

AC 29.1103. § 29.1103 (Amendment 29-17) INDUCTION SYSTEM DUCTS AND AIR DUCT SYSTEMS.

a. § 29.1103(a) : (1) Explanation . This paragraph is intended to require the design of induction system ducts for engines and auxiliary power units to include fuel and water drains which are effective in the ground attitude and do not discharge into any location where the fuel drainage could be ignited to cause a fire hazard.

(2) Procedures . Determine that each induction duct is provided with at least one drain of sufficient size to minimize clogging and located at the low point of the duct with the rotorcraft in the ground attitude. Discharge from the drain should not create a hazard to the rotorcraft.

b. § 29.1130(b) : [Section AC 29.1103 continued on Page E – 163] Page E – 161 AC 29-2C, Chg 4 5/1/2014 [This page intentionally blank.]

Page E – 162 9/30/99 AC 29-2C (2) Procedures . The magnitude of the backfire to be considered is somewhat subjective; however, the rule can generally be satisfied by testing which involves inducing actual backfires in the engine. This can usually be accomplished by crossing ignition leads between cylinders to cause ignition when the intake valve is open. Tests should include both engine cranking and power-on regimes.

c. § 29.1103(c) : (1) Explanation . Induction ducts, particularly on reciprocating engines, involve connections with other ducts and with structure. Flexibility is required to prevent relative motion (expansion, structural deflections, etc.) from prestressing the duct.

(2) Procedures . Review the design for long runs of ducting between the engine and structural supports and between other connections or supports in the duct system.

Short segments of the duct constructed of bellows will usually provide the necessary flexibility.

d. § 29.1103(d) : (1) Explanation . The effectiveness of fire extinguisher systems is based, in part, on testing for agent concentration in the fire zone with the airflows to be expected.

Any duct failure (burnout) during an engine compartment fire may be expected to introduce air to dilute the agent concentration, or if the duct passes through a firewall, duct burnout could result in an opening in the firewall. Fireproof ducts, as specified by this rule, are needed to ensure the integrity of the firewalls and the effectiveness of the fire extinguisher system. Fire resistant ducts may be used if located totally within the fire zone.

Procedures (2) . Ducts within a fire zone are usually engine air induction ducts, air bypass ducts, or cooling air ducts. For ducts which penetrate the firewall or other fireproof construction such as fireproof cowling, verify that the duct is of fireproof construction. Other ducts may be only fire resistant. A duct constructed of material which has been accepted as firewall material would be considered as fireproof without further testing (unless the duct is subject to significant structural loads, in which case, fire testing may be necessary with the loads applied to the duct). The tests for “fireproof” and “fire resistant” qualification differ only in the time exposure; i.e., 15 minutes for “fireproof” and 5 minutes for “fire resistant.” If nonmetallics are used in duct construction intended for “fireproof” applications and the integrity of the test specimen is deteriorating towards the end of the 15-minute fire test period, assessment of the situation with respect to possible hazards if the engine fire were to exist beyond 15 minutes is appropriate. Duct burnout should not result in the possibility that fire could escape the fire zone and create hazardous conditions.

e. § 29.1103(e) : Page E - 163 AC 29-2C 9/30/99 Explanation . This rule requires additional fireproofing of the inlet duct of (1) auxiliary power units (APU’s) to ensure safe disposal or containment of hot gas reverse flow from the APU from entering any other compartment of the rotorcraft in which a hazard would be created. This rule could, in some designs, require fireproof construction of the inlet duct for the APU to extend upstream beyond the confines of the firewall provided in compliance with § 29.1191(b). The extent of the fireproofing is subjective and may require malfunction testing if no applicable information can be provided by the manufacturer of the APU. For ducting upstream of the fireproof section, materials selected need not be qualified for fire impingement; however, they must be shown to be suitable for the maximum normal heat conditions to be expected.

Procedures . Normally, fireproof ducting upstream of the APU to the contour (2) of the roto rcraft is acceptable for compliance. However, if this distance is less than 36 inches, the possibility of impingement of hot gases on the contour skin of the rotorcraft is required. Fireproofing of contour skin or duct relocation should be considered if the impingement area is a nonmetallic structure or is part of or close to fuel tanks. Other system air inlets in the impingement area should also be evaluated for possible hazards due to ingestion of hot gases in event of reverse flow from the APU.

f. § 29.1103(f) : (1) Explanation . APU inlet ducts subject to reverse flow of hot gases should be constructed of materials that will not absorb fuel or other flammable liquids to avoid induction duct inlet fires which may ignite by backfires or reverse flow from an APU.

(2) Procedures . Any nonmetallic duct material should be shown by test or by previous qualification to be sealed or otherwise free of tendencies to absorb flammable liquids. Tests, if necessary, should follow the guidelines for absorption qualification set forth in TSO’s or military specifications for fuel and oil tanks.

AC 29.1105. § 29.1105 INDUCTION SYSTEM SCREENS .

a. Explanation . This paragraph concerns reciprocating engine installations. If induction system screens are used, the following considerations apply.

(1) Each screen must be upstream of the carburetor.

(2) No screen may be in any part of the induction system that is the only passage through which air can reach the engine unless it can be deiced by heated air.

(3) No screen may be deiced by alcohol alone, and it must be impossible for fuel to strike any screen.

b. . Inlet screens in the engine induction system are generally Procedures provided t o prevent the entrance of foreign objects. The induction design may Page E - 164 9/30/99 AC 29-2C incorporate features which address the concerns identified above. Also, some designs incorporate an alternate air door which, with appropriate consideration, accounts for the requirements of this paragraph. The alternate air source should provide the required air to maintain flight and landing to a suitable landing site at appropriate airspeeds and gross weights.

AC 29.1107. § 29.1107 INTERCOOLERS AND AFTER-COOLERS .

a. Explanation . Each intercooler and after-cooler must be able to withstand the vibration, inertia, and air pressure loads to which it would be subjected in operation.

Procedures b. . In complying with this regulation, the various vibrations, inertia, and air pressure loads should be identified. The installation may be verified by either analysis or test appropriate for the design.

AC 29.1109. § 29.1109 CARBURETOR AIR COOLING .

a. Explanation . It must be shown under § 29.1043 that each installation using two-stage superchargers has means to maintain the air temperature at the carburetor inlet, at or below the maximum established value.

Procedures b. . When the powerplant installation design utilizes a supercharger installation, it should be shown by testing that the air temperature at the carburetor inlet does not exceed established values.

Page E - 165 AC 29-2C 9/30/99 SUBPART E - POWERPLANT EXHAUST SYSTEM AC 29.1121. § 29.1121 (Amendment 29-13) EXHAUST SYSTEM - GENERAL.

a. Explanation .

(1) This section addresses the arrangement of exhaust components and the protection against hazardous conditions which exist with hot exhaust gases for powerplant and auxiliary power unit installations.

(2) The objective is to ensure safe disposal of exhaust gases without fire hazard or physical impairment to any occupant.

b. Procedures.

(1) During the certification process, carbon monoxide levels should be monitored in the personnel compartments to verify that the gas levels are well within the acceptable range. The conditions under which the measurements are taken should be representative of the normal operating limitations of the rotorcraft. This paragraph is not applicable to gas turbine-engine-powered rotorcraft.

(2) Exhaust system surfaces hot enough to ignite flammable fluids or vapors must meet the isolation or shielding requirements of this section in addition to the requirements of §§ 29.1183 and 29.1185. Good design practice suggests that the isolation and shielding features incorporated would continue to be effective under the emergency landing conditions specified in § 29.561.

(3) Compliance with the § 29.1121(c) fireproof requirements can be accomplished by demonstrating that the material or component will withstand a 2000° F ± 50° F flame for 15 minutes while still fulfilling its design purpose. This testing should accurately simulate, as near as practicable, the operating environment of the material or component in service. In addition to the fireproof requirements, the requirements of § 29.1191 must be met.

(4) Compliance with § 29.1121(d) can be accomplished by locating the vents and drains where fumes and fluids cannot interact with the hot exhaust gases. Drains should discharge positively and be a minimum of 0.25 inches in diameter. No drain may discharge where it will cause a fire hazard. This can be demonstrated by discharging a colored liquid through the drain system in flight and on the ground. The dye should not impinge on any ignition source.

(5) It should be demonstrated that exhaust gases are discharged in such a manner that they do not cause distortion or glare seriously affecting the pilot’s visibility Page E - 166 9/30/99 AC 29-2C at night. One method of compliance would be a night flight evaluation at critical azimuth and variable wind conditions to verify that no degradation exists.

(6) Hot spots that can occur on exhaust system components should be eliminated by providing deflectors and/or adequate ventilation. Exhaust shrouds can either be ventilated or insulated to keep the temperatures low enough so that ignition of flammable vapors or fluids cannot occur under normal operation or under the emergency landing conditions specified in § 29.561.

(7) Compliance with § 29.1121(h) can be accomplished by ensuring that the drain will not discharge where it might cause a fire hazard. This can be demonstrated by discharging a colored liquid through the drain system in flight and on the ground.

The dye should not impinge on any ignition source.

AC 29.1123. § 29.1123 EXHAUST PIPING .

a. Explanation . This section contains the following requirements that must be met for proper certification of exhaust piping on engines, auxiliary propulsion units (APU), and other similar devices.

(1) § 29.1123(a) requires that the piping be heat and corrosion resistant so that it performs its intended function during its operational life (either the life of the rotorcraft or a specified limited life) without significant metal corrosion, metal erosion, or creation of hazardous hot spots. The piping system should be designed, have an installation design, or a combination that allows performance of its function without thermal expansion (thermal strain) induced structural failures, such as ruptures caused by operating temperature excursions and by overpressurization during its operational life.

(2) § 29.1123(b) requires that the piping must be supported to withstand the vibration and loading environment (including inertia loads) to which it will be subjected in service.

(3) § 29.1123(c) requires that piping that connects to components between which relative motion exists in service must have the necessary flexibility and structural integrity to withstand the relative motion without exceeding limit load (at the maximum operating temperature) of the piping, or creating unintended loads (or load paths) on the components to which the piping connects.

b. Procedures . Exhaust piping is typically certified by analysis and installation tests conducted during the basic certification process, including flight tests, as follows: (1) For compliance with § 29.1123(a), because of its durability in the hot exhaust environment, exhaust piping is typically made from stainless steel or alloy steel of the appropriate structurally and thermally derived wall thickness. Hot aircraft exhaust gases are very corrosive; thus, proper material selection and corrosion protective design should be performed and validated during certification. Advisory Circular Page E - 167 AC 29-2C 9/30/99 (AC) 43-4, “Corrosion Control For Aircraft” contains a detailed discussion of exhaust gas corrosion problems. Analysis and/or verification tests of the exhaust system should be conducted. This work is necessary to ensure thermal and structural integrity; to ensure that thermal expansion does not cause a structural overload or failure; and, to ensure that exhaust piping does not contact (or come close to) ambient temperature materials (such as structure or system components). Hot exhaust piping in contact with (or close to) ambient temperature materials can either create a fire hazard or cause an unintended strength reduction. To ensure that thermal expansion analyses and tests are properly conducted, the maximum in-service temperature excursion should be properly defined. The maximum temperature excursion should be based on the maximum temperature of the piping and exhaust gases, as affected by the insulatory characteristics of the piping’s enclosure, and as affected by a worst case hot day. The worst case temperature environment used for analysis can be verified by a temperature survey. If run on cooler days, the survey can be adjusted for the worst case hot day environment using methods identical to those used for engine cooling tests (reference paragraph AC 29.1043, Cooling Tests). The piping should be designed to expand freely so that thermal expansion (thermal strain) induced loads on the piping and its restraint system are minimized. If thermal expansion induced loads (in conjunction with deflection induced loads and exhaust flow loads, discussed in b(4)) are significant relative to limit load of any item in the load path, then a fatigue check on the critical design point(s) should be performed. The fatigue check should establish a safe life or an approved limited life for the critical component(s) in the system. An accurate analytical fatigue check on exhaust piping may be difficult to perform because of erosion, corrosion, etc., in service; therefore, phased inspections should be considered to ensure the exhaust piping’s continued airworthiness.

(2) For compliance with § 29.1123(b), exhaust piping should be properly supported so that the maximum loads anticipated in-service are properly distributed and reacted, and, as previously discussed, so that thermal expansion induced loading is minimized. Typically the worst case static design load conditions are either the inertia loads from an emergency impact (reference § 29.561) or the combined loading from thermal expansion, in-flight deflections and internal exhaust gas flow (See paragraph b(4)). It should be noted that several combinations of these loads should be examined to determine the critical combination. The piping should be supported and restrained such that critical frequencies are avoided and the induced vibration environment’s effect is minimized. Flight test vibration surveys may be necessary, in some cases, to properly define or validate the critical modes and environment and their effect on the exhaust piping design. Operating modes such as ground idle, flight idle, 40 percent and 80 percent of maximum continuous power, maximum continuous power, OEI power settings and other power settings should be investigated to determine their vibratory effect on the exhaust gas piping system. The strength reduction of the piping materials at operating temperature (and at worst case temperature) should be properly considered in the design and structural substantiation. MIL-HDBK-5D contains material allowables versus temperature data for a wide variety of metallic engineering materials.

Page E - 168 9/30/99 AC 29-2C (3) For compliance with § 29.1123(c), the piping and its restraint system should be designed to minimize loading induced on the piping by the relative motion (in-service deflections) of the components to which the system attaches. Isolation of significant deflection induced loading (if required based on analysis and strain surveys) by use of flexible joints or other equivalent devices or designs should be considered. Any such in-line device used to reduce deflection loading should be fireproof and leak free when performing its intended function.

(4) For critical load case determination, the expansion-induced thermal loading should be added in with mechanical relative-motion induced loads and internal exhaust gas flow loads to provide total critical loads for both a proper static and a proper fatigue structural substantiation. The critical combined static load should be compared with the emergency impact loads of § 29.561(paragraph b(2)) to determine the critical design load case for static strength substantiation.

(5) It should be noted that the majority of the exhaust piping verification testing required for certification can be accomplished during the rotor drive system tie down testing of § 29.923.

AC 29.1125. § 29.1125 (Amendment 29-12) EXHAUST HEAT EXCHANGERS .

a. Explanation . This section applies only to rotorcraft powered by reciprocating engine(s) or equipped with reciprocating auxiliary propulsion units (APU). This regulation states the certification requirements for exhaust heat exchangers (EHE’s) which are summarized as follows: (1) § 29.1125(a) requires that each EHE be constructed and installed to withstand vibration, inertia and other operational loads.

(2) § 29.1125(a)(1) requires that each EHE be able to operate continuously at the highest anticipation service temperature.

(3) § 29.1125(a)(1) requires that each EHE be corrosion resistant to exhaust gases and other corrosion sources.

(4) § 29.1125(a)(2) requires that each EHE have provisions for inspecting its critical parts and areas.

(5) § 29.1125(a)(3) requires that each EHE have cooling provisions where it is subjected to hot exhaust gases.

(6) § 29.1125(a)(4) requires that each EHE muff design eliminate stagnation areas or liquid traps that would contribute to ignition of leaked flammable fluids.

(7) § 29.1125(b) requires that each EHE used to heat ventilating air for occupants- - Page E - 169 AC 29-2C 9/30/99 (i) Either have a secondary heat exchanger between the primary EHE and the ventilating air system; or (ii) Have other equivalent means to prevent harmful contamination of ventilating air.

b. Procedures . EHE’s and their installations are typically certified by analysis and installation tests conducted during the basic certification process, including flight tests or simulated flight tests, as follows: (1) Because of their durability in the hot exhaust environment, EHE’s are usually constructed from stainless steel or alloy steel of the appropriate structurally and thermally derived wall thickness. The EHE and its system should be designed to expand freely to minimize thermal expansion (thermal strain) induced loads on the EHE and its restraint system. If thermal expansion induced loads (in conjunction with deflection induced loads and exhaust flow loads) are significant relative to the limit load of the EHE or its attachments, a fatigue check on critical design point(s) should be performed. The fatigue check should establish a safe life or an approved limited life for the critical component(s) in the EHE system.

(2) EHE’s should be properly supported so that the maximum loads anticipated in service are properly distributed and reacted and so that thermal-expansion-induced loading is minimized. Typically, the worst-case static design load conditions are either the emergency impact loads acting alone (reference § 29.561), or the critical combination of loads from thermal expansion, in-flight deflections and internal exhaust gas flow. Several combinations of these loads should be examined to determine the critical combination. The EHE should be supported and restrained so that critical frequencies are avoided and the induced vibration environment is minimized. Flight tests or bench tests, such as vibration surveys conducted during rotor system endurance testing, may be necessary in some cases, to properly define or validate the vibration environment and EHE’s critical modes and their effect on EHE design.

Operating modes such as ground idle, flight idle, 40 percent and 80 percent of maximum continuous power, maximum continuous power, OEI power settings, and other critical power settings should be investigated to determine their vibratory effect on the EHE system. The strength reduction of EHE materials at operating temperature and at critical temperatures should be properly considered in EHE design and structural substantiation (MIL-HDBK-5D contains material allowables versus temperature data for a wide variety of metallic engineering materials). The EHE and its restraint system should be designed to minimize loads induced by the relative motion (in-service deflections) of the components to which the EHE attaches. Isolation of significant-deflection-induced loading (as required, based on analysis and strain surveys) by use of flexible joints, other equivalent flexible devices, or designs should be considered. Any such in-line device used to reduce deflection loading should meet applicable certification requirements and be leak-free.

Page E - 170 9/30/99 AC 29-2C (3) Expansion analysis and verification tests of the EHE should be conducted to ensure its thermal (and structural) integrity and to ensure that thermal expansion does not cause the EHE to contact (or come close to) ambient temperature aircraft materials, structure or system components and either create a fire hazard or an unintended reduction in strength. To ensure that expansion analyses and tests are properly conducted, the maximum in-service temperature excursion should be properly defined.

The maximum temperature excursion should be based on the maximum temperatures of the EHE and exhaust gases, as affected by the insulatory characteristics of the EHE’s enclosure, and as affected by a worst-case hot day. The worst-case temperature environment used for analysis can be verified by a temperature survey which, when run on cooler days, can be adjusted to the worst-case hot day environment using methods identical to those used for engine cooling tests (reference paragraph AC 29.1043, Cooling Tests).

(4) Hot aircraft exhaust gases are very corrosive; thus, proper material selection and corrosion protection design should be performed and validated during certification. Advisory Circular (AC) 43-4, “Corrosion Control For Aircraft” contains a detailed discussion of exhaust gas corrosion problems. The in-service corrosive environment should be identified and characterized as thoroughly as possible by chemical analysis, tests and service experience. Once defined, appropriate design techniques and materials should be selected. Certification tests may be required to ensure proper substantiation. Phased inspections and inspectability should be considered (reference (4)).

(5) The EHE’s design should be reviewed for inspectability to ensure that structural and thermal integrity is maintained over the intended life of the EHE. Also, if the design review is not conclusive relative to inspectability, a tear down inspection should be conducted.

(6) Each EHE design should be reviewed, analyzed, and tested to ensure that cooling provisions are adequate where EHE surfaces are subjected to hot exhaust gases. This is necessary to prevent hazardous hot spots or a burn through which may cause a fire and contaminate the occupied environment.

(7) Each EHE design should be reviewed, analyzed, and tested to ensure that stagnation areas and liquid traps do not exist. This can be done using bench flow tests.

These stagnant areas and traps could become ignition sources if wetted with a leaking flammable fluid. A review of potential leaking flammable fluid hazards should be conducted and appropriate preventative measures such as drains and drip fences installed to ensure they are routed away from EHE’s.

(8) Each EHE design which will be used to heat ventilating air for occupants should be reviewed to ensure that the EHE is a double walled system, (i.e., it would require failure of two EHE surfaces to allow toxic exhaust gases to intermix with cabin ventilating air). Each EHE wall should be designed with equal thermal and structural resistance since a single undetected inner wall failure would subject the outer wall to the Page E - 171 AC 29-2C 9/30/99 primary heat load. Also, inspectability provisions should be provided or means identified to ensure that inner wall failures can be detected in service. Any equivalent means which is applied for must clearly provide an equivalent level of safety to a double walled EHE.

Page E - 172 9/30/99 AC 29-2C SUBPART E - POWERPLANT POWERPLANT CONTROLS AND ACCESSORIES AC 29.1141. §. 29.1141 (Amendment 29-13) POWERPLANT CONTROLS: GENERAL .

a. Explanation .

(1) Section 29.1141(a) References §§ 29.777 and 29.1555. The detailed compliance procedures for powerplant control arrangement and markings are found in these sections.

(2) Section 29.1141(b) requires that controls be located and/or shielded such that normal movement of cockpit personnel will not cause inadvertent control movements.

(3) Section 29.1141(c) requires that each flexible control (push-pull cables) be properly approved.

(4) Section 29.1141(d) requires that each control maintain its set position without movement from an inadvertent source such as vibration or control system loads.

This is required so that constant flightcrew attention is not necessary.

(5) Section 29.1141(e) requires that each control be able to withstand operating loads without excessive deflection. Excessive deflection is interpreted to be that deflection that would cause erratic movement, lack of crispness, or premature failure.

(6) Section 29.1141(f) specifies acceptable open/close positions for manual valves to prevent power failure due to improper control valve positioning.

Power-assisted valves should have means to indicate to the flightcrew that the valve is either in the fully open or fully closed position or that the valve is moving between these two positions.

(7) The control system is subject to evaluation under § 29.901(c); i.e., for turbine installations, no single failure or malfunction, or probable combination thereof, of any powerplant control system should cause the failure of any powerplant function necessary for safety. One acceptable way to determine this is by use of a failure modes and effects analysis (FMEA).

b. Procedures .

(1) For compliance with § 29.1141(a), review the procedures for paragraph AC 29.1555. Evaluation by the flight test pilot during the official flight test program is appropriate.

Page E - 173 AC 29-2C 9/30/99 (2) Compliance with § 29.1141(b) is normally evaluated during the flight test program and documented in the flight test report.

(3) Compliance with § 29.1141(c) may be accomplished by qualifying the control to MIL-C-7958, “Controls, Push-Pull, Flexible, and Rigid,” or other approved standards or by previous approval in a similar function, installation, or arrangement.

(4) Compliance with § 29.1141(d) may be shown during the flight test program by monitoring the means to prevent control creep. This device or arrangement should be effective without crew attention and should not impose undue control displacement loads or interfere with accurate settings.

(5) Compliance with § 29.1141(e) may be shown by an appropriate structural analysis and/or a witnessed static load test using the factors specified under § 29.397 unless a lower value can be shown to be applicable. Operation tests and design details described in §§ 29.683 and 29.685 should also be considered.

(6) Compliance with § 29.1141(f)(1) may be accomplished by installing manual valves which have positive stops in the fully open and closed positions. The fuel valves, however, may have an arrangement to facilitate the capability of switching to different fuel tanks if suitable indexing is provided. Compliance with § 29.1141(f)(2) may be accomplished by installing a device which displays to the flightcrew one indication with valve fully open and another with the valve fully closed. Alternatively, an indication could be given when the valve is moving from fully open to fully closed with the indication ceasing when the valve position corresponds to the selected switch position (open or closed). An example would be a light that is “off” when the valve is fully open or fully closed and illuminates while the valve is transitioning.

AC 29.1142. § 29.1142 (Amendment 29.17) AUXILIARY POWER UNIT CONTROLS .

a. Explanation .

(1) This section addresses control requirements for any APU installed in a rotorcraft.

(2) The requirement for starting, stopping, and emergency shutdown of the APU from the flight deck is primarily to control APU operation in the event of improper operation or malfunction which could affect the safety of the aircraft.

b. Procedure .

(1) The requirements of this section apply to all APU installations in rotorcraft without regard to whether or not the APU is to be operated on the ground only, or operated in flight and on the ground.

Page E - 174 9/30/99 AC 29-2C (2) The APU installation must provide sufficient controls to the flight crew to enable them to control the operation of the APU under normal and emergency conditions.

(3) Compliance can be shown by both demonstration and a failure analysis.

AC 29.1143. § 29.1143 (Amendment 29-12) ENGINE CONTROLS.

a. Explanation . This section prescribes safety standards applicable to arrangement and operation of the engine controls.

(1) Section 29.1143(a) requires a separate throttle for each engine.

(2) Section 29.1143(b) requires a throttle arrangement for control of all engines be achieved by: (i) Separate control of each engine.

(ii) Simultaneous control of all engines.

(3) Section 29.1143(c) requires that immediate actuation at the engine control should be provided by any given input at the cockpit throttle control.

(4) Section 29.1143(d) requires that each fluid injection system control (e.g., water-alcohol) other than the fuel system control must reside in the throttle controls.

This does not preclude the injection system pump from having a control located separately from the throttle.

(5) Section 29.1143(e) requires that power or thrust controls (that have fuel shut-off features) provide a means to prevent inadvertent movement to the shut-off - position. This means should- (i) Provide a positive lock or stop at the idle position; and (ii) Require a separate and distinct operation to place the control in the shut-off position.

Procedures. b.

(1) Certification data submitted by the applicant should be reviewed to ensure that all the design features stated in § 29.1143 exist.

(2) Proper engine control functioning (to verify the design features of § 29.1143) should be verified as part of the type inspection authorization (TIA) for the certification project.

Page E - 175 AC 29-2C 9/30/99 (3) Compliance with § 29.1143(e)(1) has been shown successfully in the past by use of idle detents (mechanical or electrical/mechanical such as a solenoid).

(4) In the past, compliance with § 29.1143(e)(ii) has been achieved by use of a switch or button to displace the idle stop or by use of distinct offsets in throttle motion to allow movement from the idle stop to shutoff.

AC 29.1143A. § 29.1143 (Amendment 29-26) ENGINE CONTROLS .

a. Explanation . Amendment 29–26 revises § 29.1143 by replacing the terms “throttle control” and “thrust control” with the more general term “power control.” The changes should preclude misconceptions regarding engine control arrangements when governor-controlled turboshaft engines are employed in rotorcraft.

Procedures. The means of compliance for this section is unchanged. b.

AC 29.1143B. § 29.1143 (Amendment 29-34) ENGINE CONTROLS .

a. Explanation . Amendment 29-34 introduced the option of using 30-second/2-minute OEI power ratings to multiengine rotorcraft. This amendment revises § 29.1143 by adding the requirement for automatic control of 30-second OEI limits in the new § 29.1143(e). Automatic control of the 30-second OEI limits are required to prevent exceedances of the remaining power sections after the precautionary shutdown of one engine. The use of 30-second OEI power must be limited to emergency use only during flight conditions where one engine has failed or has been shutdown for precautionary reasons. During this critical stage of flight crew attention should not be focused on powerplant instruments to avoid limit exceedances.

Procedures b. . The automatic controls used to prevent 30-second OEI limit exceedances can be installed on the airframe or the engine. The applicant should demonstrate that 30-second OEI limits that can affect the continued safe operation of the drive system or engine such as gas generator speed, measured gas temperature, torque, etc., cannot be exceeded. It should also be shown that these devices do not restrict the ability to achieve the full 30-second OEI limits. The operation of these limit devices can be demonstrated on the aircraft or if possible by using bench tests.

AC 29.1145. § 29.1145 (Amendment 29-13) IGNITION SWITCHES .

a. Explanation .

(1) This section addresses the arrangement and protection of ignition switches for reciprocating engines or for turbine engines which require continuous ignition.

(2) The objective is to provide a means to shut off all ignition quickly, if required, while at the same time providing protection against inadvertent ignition switch operation.

Page E - 176 9/30/99 AC 29-2C (3) Section 29.1145(b) does not specifically state that turbine engines not requiring continuous ignition are excluded from the rule, but no benefit is realized by the capability of shutting off all ignition to these engines.

b. Procedures.

(1) Section 29.1145(b) is self-explanatory in specifying that a means be available to shut off all ignition quickly by the grouping of switches or by a master ignition switch control. A “T” arrangement or split rocker switches are possible configurations. A master ignition control, if utilized, would need to be carefully evaluated if rotorcraft performance credit is given for engine isolation.

(2) Each group of ignition switches and the master ignition control should have a means to prevent inadvertent operation. “Guarded” switches are the usual means of showing compliance.

AC 29.1147. § 29.1147 MIXTURE CONTROLS .

a. Explanation . This section addresses the arrangement of fuel mixture controls, if installed. Major manual adjustment of the fuel mixture to optimize performance is not normally allowed due to the possibility of engine failure or detonation if significant misadjustment occurs. If “best-power” with respect to fuel mixture is desired, normal practice is to utilize engines with automatic mixture controls, in which case the lever in the cockpit reverts to merely an engine shutdown device. In any case, manual adjustment of the mixture, except for intentional shutdown, should not be prescribed without positive means of ascertaining that the resulting fuel-air mixture is within the range associated with safe engine operation. Some manual mixture adjustment may be acceptable for more efficient engine operation if suitable stops or automatic means are provided to prevent inadvertent engine shutdown with mixture movement or engine malfunction with flight condition changes.

(1) Section 29.1147(a) requires (if mixture controls exist) that controls be arranged to allow: (i) Separate control of each engine.

(ii) Simultaneous control of all engines.

(2) Section 29.1147(b) requires that each intermediate position of the mixture controls corresponding to a normal operating setting be identifiable by both feel and sight.

Procedures. b.

Page E - 177 AC 29-2C 9/30/99 (1) Certification data submitted by the applicant should be reviewed to ensure that the design features stated in § 29.1147 exist.

(2) Proper mixture control functioning (to verify the design features of § 29.1147) should be verified as part of the TIA for the certification project.

(3) Compliance is typically shown by use of a side-by-side arrangement of the controls, provided that the arrangement is compatible with other controls and considering that crew attention to the primary flight controls may be a full-time, “hands-on” operation.

AC 29.1151. § 29.1151 ROTOR BRAKE CONTROLS .

a. Explanation .

(1) Paragraph (a) of § 29.1151 is intended to require design features which, for all practicable purposes, prevent brake application in flight even under conditions of reasonably expected crew error or confusion.

(2) Paragraph (b) of § 29.1151 would require warning devices to alert the crew if the brake has not been completely released.

b. Background . Inadvertent or undetected application of the rotor brake is expected to result in excessive heat and fire in the rotor brake area. Rotor brake components are usually located integral with, or in close proximity to, rotor drive system components and, in many cases, close to critical hydraulic main rotor control system components. Fires in these areas would be extremely hazardous.

Methods of Compliance. c.

(1) For paragraph (a) literal compliance can be achieved by lock-out devices sensitive to the higher RPM. ranges of the main rotor or other flight parameters, hydraulic bypass or lockout devices controlled by flyweight governor systems, etc. The guard required by § 29.921 does not, in itself, provide compliance with this requirement.

For some designs, if careful evaluation of the overall control, including location, guard mechanism, control manipulation requirements, accessibility, etc., provides an extremely high degree of assurance that inadvertent application will not occur, compliance may be assumed. Also, if brake application does occur, annunciation appears, and no immediate hazard to flight operation exists, compliance may be assumed.

(2) Warning devices supplied to comply with this rule should provide a signal at any time the rotor brake is engaged, including partial engagement. Typically, micro-switches installed to close a circuit to a cockpit warning (red) light when the brake puck moves out of the retract position will provide compliance, provided the designer gives full consideration to the vibration, temperature, moisture, and other environmental Page E - 178 9/30/99 AC 29-2C considerations appropriate to configuration. Other methods such as system pressure switches, brake handle position indicators, etc., may not provide the warning required by this rule.

AC 29.1157. § 29.1157 CARBURETOR AIR TEMPERATURE CONTROLS .

a. Explanation .

(1) This section addresses the air temperature control for carburetor equipped reciprocating engines.

(2) For rotorcraft which have more than one such engine installed, a separate carburetor air temperature control must be provided for each engine.

b. Procedure.

(1) The engine air induction system should incorporate a means for the prevention and elimination of ice accumulations by preheating the air prior to its entry into the carburetor.

(2) Manually operated push/pull systems have been used which operate a flapper valve inside the air induction system. One such system for each engine is one method of compliance.

AC 29.1159. § 29.1159 SUPERCHARGER CONTROLS .

a. Explanation .

(1) This section addresses the accessibility to supercharger controls in the cockpit, if installed.

(2) These controls must be located so they are easily reached by the pilots or, if the rotorcraft is so configured, by a flight engineer.

. b. Procedure (1) The location and shape of the controls should be conveniently accessible and sufficiently unique to preclude inadvertent actuation of the wrong control.

(2) Compliance is typically shown by a cockpit evaluation.

Page E - 179 AC 29-2C 9/30/99 AC 29.1163. § 29.1163 (Amendment 29-26) POWERPLANT ACCESSORIES .

a. Explanation .

(1) This section addresses the interface requirements for powerplant accessories which are mounted on the engine or rotor drive system components.

(2) Areas which should be addressed include structural loads imposed upon the engine case and isolation between the accessory and engine oil systems. Electrical equipment isolation from flammable fluids or vapors should be addressed as well as the effect of an accessory failure on the continued operation of the engine and drive system components.

b. Procedures .

(1) Accessories installed and certified by the engine manufacturer can be mounted on the engine without additional justification.

(2) Any accessory to be mounted on the engine, which was not certificated with the engine and does not meet the engine installation design manual requirements, should have a structural analysis showing the mounting of that accessory on the engine will not induce loads into the engine case which are higher than the original design loads.

(3) When the accessory is mounted and operating on the engine, it should not be possible to contaminate either the engine or accessory oil systems. This contamination can take the form of debris following a failure, airborne dirt or water, or any other substance that would impair proper operation of the engine or accessory.

Compliance with these requirements can be accomplished by a combination of test and analysis. The design interface should be such that when the equipment is operating, there are no high/low pressure differentials between the components which would induce fluid transfer between components resulting in a low fluid level in one component and an overfill condition in the other component. Where this potential exists, an analysis and/or test should be used to demonstrate compliance.

(4) Engine mounted accessories which are subject to arcing and sparking must be isolated from all flammable fluids or vapors to minimize the probability of fire. This can be accomplished by isolating the electrical equipment from the flammable fumes or vapors or by isolating the flammable fumes or vapors from the potential ignition source.

Compliance can be shown by analysis.

(5) A failure mode and effect analysis should be submitted which shows that a failure of any engine mounted and driven accessory will not interfere with the continued operation of the engine. If a hazard is created by the continued rotation of an engine driven accessory after a failure or malfunction, provisions to stop its rotation or eliminate Page E - 180 9/30/99 AC 29-2C the hazard must be provided. The effectiveness of this device should be demonstrated by test.

(6) The main transmission and rotor drive system should be protected from excessive torque loads and damage imposed upon them by accessory drives. One method which has been used is a torque limiting device (i.e., shear section of main rotor drive shaft). The effectiveness of any protection device should be demonstrated by test.

AC 29.1165. § 29.1165 (Amendment 29–12) ENGINE IGNITION SYSTEMS .

a. Explanation .

(1) This section defines the design requirements for battery, generator, and magneto ignition systems installed in either reciprocating or turbine engine powered rotorcraft.

(2) The requirements specify common failure modes of batteries, generators, and installed wiring which must be considered in the design process and provides for crew warning of malfunctions.

b. Procedures .

(1) In a battery ignition system, a generator should be available to supply current to the engine ignition system if the battery fails. The generator power should be switched over automatically with an appropriate warning to the crew. The automatic switchover can be accomplished by a low voltage sensor which activates a relay that simultaneously activates a caution light in the cockpit.

(2) An electrical load analysis should be conducted to insure that the capacity of the batteries and generator is large enough to meet the worst-case demands in the system. If there are other electrical system components installed which draw from the same source, the analysis should show that there is sufficient electrical power available from either the battery or the generator to operate all components simultaneously.

(3) The requirements of § 29.1165(c)(1) through (3), should be demonstrated by test. A proposed test plan should be coordinated with the FAA/AUTHORITY prior to conducting the testing.

(4) Compliance with the requirements of § 29.1165(d) can be shown by a failure mode and effect analysis.

(5) The requirements of § 29.1165(e) and (f) are self-explanatory.

Page E - 181 AC 29-2C 9/30/99 SUBPART E - POWERPLANT POWERPLANT FIRE PROTECTION AC 29.1181. § 29.1181 (Amendment 29–26) DESIGNATED FIRE ZONES: REGIONS INCLUDED .

a. Explanation . A designated fire zone is a zone on a rotorcraft within which it is assumed (based on past operational experience) that a severe fire (see definitions) will occur sometime in the service life of each rotorcraft; therefore, proper protection must be provided for each new or modified unit by meeting the requirements of §§ 29.1183 through 29.1203. Some common examples of designated fire zones are: (1) For reciprocating engines: (i) The power section.

(ii) The accessory section.

(iii) The complete powerplant compartment, if there is no isolation between the power and accessory sections.

(2) Any auxiliary power unit (APU) compartment.

(3) Any fuel burning heater or other combustion equipment installation described under § 29.859.

(4) For Turbine Engines: (i) The compressor section.

(ii) The accessory section.

(iii) The combustor turbine and tailpipe section unless they- - (A) Do not contain lines and components carrying flammable fluids or gases; and (B) Are isolated from the designated fire zone prescribed in § 29.1181(a)(6) by a firewall that meets § 29.1191.

(5) Any other essential or non-essential device or system (such as spray rigs using flammable fluids) capable of leaking flammable fluid or gas and creating a severe fire.

Page E - 182 9/30/99 AC 29-2C b. Definition. Severe fire. See definition in paragraph AC 29.859.

c. Procedures . A FAA/AUTHORITY/applicant design review should be conducted early during certification to identify all designated fire zones and to define the detailed method-of-compliance to be used to meet the requirements of §§ 29.1183 through 29.1203. If significant design changes are made the design change and the method-of-compliance should be re-reviewed to insure they properly support the certification requirements.

AC 29.1183. § 29.1183 (Amendment 29-22) LINES, FITTINGS, AND COMPONENTS .

a. Explanation . This section requires that any line, fitting or other component of a flammable fluid, fuel or flammable gas system which carries, conveys or contains the fluid or gas in any area subject to engine fire conditions (i.e., a severe fire) must be at least fire resistant (reference § 1.1 for definition of fire resistant and see paragraph AC 29.859 which defines a severe fire). An exception is for flammable fluid tanks and supports which are part of and attached to the engine or are in a designated fire zone. These items are required to either be fireproof (see § 1.1 for definition of fireproof and see paragraph AC 29.859 which defines a severe fire) or to be enclosed by a fireproof shield, unless fire damage to any non-fireproof part (e.g., secondary line or valve support) will not cause leakage of a flammable gas, flammable fluid or otherwise prevent continued safe flight and landing of the rotorcraft. All such components must be shielded, located, otherwise protected, or a combination to safeguard against the ignition of leaking flammable fluids or gases. Integral oil sumps of less than 25 quarts capacity on a reciprocating engine need not be fireproof or enclosed by a fireproof shield; however, they should be fire resistant. Most integral sumps in this category are, by natural design and material selection, fire resistant.

Exemptions to the preceding requirements are as follows: (1) Lines, fittings and components already approved under Part 33 as part of the engine itself; (2) Vent and drain lines (and their fittings) whose failure will not result in or add to an operational fire hazard. In addition, all flammable fluid drains and vents must discharge clear of the induction system air inlet and other obvious ignition hazards.

b. . A detailed review of the design should be conducted to identify Procedures and quantify all lines, fittings, and other components which carry flammable fluids and/or gases and are in areas subject to engine fire conditions such as engine compartments and other fire zones. Once these items are identified the design means of fire protection should be selected and validated, as necessary, during certification. For materials and devices that cannot be qualified as fireproof or fire resistant by similarity or by known material standards, testing to severe fire conditions (see definition, AC 20-135, and AC 23-2 for detailed requirements) should be conducted on full-scale specimens or representative samples to establish their fireproof or fire resistance capabilities. Exceptions to these standards (as provided in the regulatory section) Page E - 183 AC 29-2C 9/30/99 should be reviewed and approved/disapproved on a case-by-case basis during certification. Also, operational fire hazards from drains, vents, and other similar sources should be identified and eliminated during certification.

AC 29.1185. § 29.1185 FLAMMABLE FLUIDS.

a. Explanation . This section requires that fuel, flammable fluid or vapor tanks, reservoirs or collectors be sufficiently isolated from engines, engine compartments, and other designated fire zones so that hazardous heat transfer from these areas to fuel, flammable fluid, and vapor tanks, reservoirs or collectors is prevented in either normal or emergency service.

Definitions b. .

(1) Fuel or Flammable Fluid Collector . Any device such as a large valve, accumulator, or pump that contains a significant amount of flammable fluid, fuel, or vapor (e.g., the volume equal to 10 ounces or more of fluid).

Flammable Fluid or Vapor Tank (2) . Any fuel, flammable fluid or vapor tank, reservoir or collector.

Sufficiently Isolated (3) . Fuel, flammable fluids, or vapors in a tank, reservoir, or collector are insulated, removed, otherwise protected or a combination such that their worst case temperatures (the worst case measured or calculated surface temperature of their containers) in either normal or emergency service is always 50° F or more away from the autoignition temperature of the fuel, flammable fluid, or vapor in question.

Minimum Autoignition Temperature (4) . The temperature at a given vapor pressure at or above which liquid fuel or fuel vapor will self combust. When determining the minimum design value of autoignition temperature which will occur in either normal or emergency operations, the critical, in-service combination of vapor pressure and fuel temperature should first be determined.

Hazardous Heat Transfer (5) . A total incident heat flux (a combination of conduction, convection, and radiation, as applicable) from or in an engine compartment or any other designated fire zone which would raise the temperature level of a flammable fluid or fuel, their vapors, or the surface temperature of their containers to within 50° F or less of the minimum in-service autoignition temperature. Typically, the most critical heat transfer case to be considered is emergency service where a severe fire (see definition) is assumed to occur in each engine compartment and each designated fire zone on a case-by-case basis.

. See definition in paragraph AC 29.859. (6) Severe Fire Procedures. c.

Page E - 184 9/30/99 AC 29-2C (1) The fuel, flammable fluid, and vapor system designs should be reviewed early in the certification process to insure that all fuel or flammable fluid or vapor tanks are properly identified and isolated from engines, engine compartments, and other designated fire zones during both normal and emergency operations such as in-flight engine compartment or other fire zone fires. In some cases fuel or flammable fluid components must be located in an engine compartment or other designated fire zone.

In these cases, an equivalent safety finding (which considers the design, construction, materials, fuel lines, fittings, and controls used in the system, or system segment, contained in the engine compartment or other designated fire zone) should be undertaken as a part of the normal certification process. If the level of safety provided is equivalent to that provided by removing the system or system segment from the engine compartment or designated fire zone, then the design should be accepted. For fuel tanks only, isolation is required by regulation to be achieved by use of either a firewall (reference paragraph AC 29.1191 for Firewall Requirements) or by use of a shroud. A shroud if used should be fireproof (see § 1.1 for definition and the definition of a Severe Fire for further details) and should be drainable (or otherwise inspectable) to insure the fuel tank is not leaking in service. For other flammable fluid or vapor tanks, the regulations allow either the identical treatment previously described for fuel tanks (i.e., firewalls or shrouds) or, alternatively, use of an equivalent safety finding. Regulations require that the equivalent safety finding be based on system design, tank materials, tank supports, and flammable fluid system connectors, lines, and controls. In all cases the flammable fluids, fuels, and vapors should be sufficiently isolated from hazardous heat fluxes during both normal and emergency operations to prevent autoignition.

(2) In addition, the regulations require at least ½-inch of clear airspace between each flammable fluid or vapor tank, and each firewall or shroud that isolates the system, unless equivalent means (such as fireproof insulation) are used to prevent hazardous heat transfer from each engine compartment or other fire zone to the flammable fluid or vapor mass (or its container surface) at the fluid or vapor’s minimum autoignition temperature. If in-service structural deflections are significant, they must be taken into account when certifying the ½-inch minimum clear airspace requirement. For example, if a ½-inch clearance exists on the ground but in some normal and emergency flight conditions (e.g., autorotation) the ½ inch is reduced to ¼ inch at a critical time (in-flight engine fire), then the design (static) configuration should have at least a ½ plus ¼ equals 3/4-inch static clear airspace to insure the regulation’s intent is met.

Alternatively, fireproof insulation or additional stiffeners could be used to insure the regulation’s intent is met (i.e., the thermal equivalent of ½ clearance is maintained at all times). Any material used as insulation on or used adjacent to flammable fluid or vapor tank, should be certified as chemically compatible with the flammable fluid or vapor and to be non-absorbent in case of fuel or vapor leaks. Otherwise, the material should either be treated for compatibility and non-absorbency or not accepted.

Page E - 185 AC 29-2C 9/30/99 AC 29.1187. § 29.1187 DRAINAGE AND VENTILATION OF FIRE ZONES .

a. Explanation . To insure that any component malfunction which results in fuel, flammable fluid or vapor leaks is safely drained or vented overboard and to insure that a fire hazard is not created during either normal or emergency service, there should be complete, rapid drainage and ventilation capability present for each part of the rotorcraft powerplant installation and any other designated fire zone which utilizes flammable fluid or vapor carrying components. As a minimum, the routing, drainage, and ventilation system should accomplish the following: (1) It should be effective under normal and emergency operating conditions.

(2) It should be designed and arranged so that no discharged fluid or vapor will create a fire hazard under normal and emergency operating conditions.

(3) It should prevent accumulation of hazardous fluids and vapors in any engine compartments and other designated fire zones.

b. Definitions . Drip Fence . A physical barrier that interrupts the flow of a liquid on the underside of a surface, such as a fuel tank, and allows any leaked liquid to drip from the surface away from a hazardous locations to a safe external drain.

Procedures c. . The design of flammable fluid and gas systems running through engine compartments and other designated fire zones should have a thorough hazard analysis performed early during certification. The analysis should be updated periodically as design changes dictate. The hazard analysis should identify and quantify all normal and emergency service failures that could result in leakage of fuel, flammable fluids and vapors. Once these potential hazards are identified and quantified, appropriate design features, such as drains, drip fences and vents, that minimize or eliminate the hazard should be provided. These means should be analyzed and/or tested, as necessary, to insure that their size, flow capacity, and other design parameters are adequate to rapidly remove hazardous fluids and vapors safely away from the rotorcraft under normal and emergency flight conditions. Typically a venting or draining system should be designed to a 3-to-1 flow capacity margin over the probable worst case leak to which it could be subjected. Adverse effects such as clogging and surface tension flow reduction should be accounted for in design. Testing, including flight testing, using inert fluids or vapors may be necessary for proper design certification. In some instances it may be appropriate to include ventilation and drainage tests when the aircraft is parked.

AC 29.1189. § 29.1189 (Amendment 29-26) SHUTOFF MEANS .

a. Explanation .

(1) This section establishes the requirements for controlling hazardous quantities of flammable fluids which flow into, within, or through designated fire zones.

Page E - 186 9/30/99 AC 29-2C (2) When any shutoff valve is operated, any equipment, including a remaining engine, which is essential for continued flight, cannot be affected.

b. Procedures .

(1) Combustible fluid supply lines which pass into, within, or through a firewall into the fire zone must incorporate shutoff valves. This requirement does not apply to lines, fittings, and components which were certified with and are part of the engine.

These requirements do not apply to oil systems for Category B rotorcraft with reciprocating engines with less than 500 cubic inches displacement or to any other installation where all components, including the oil tanks, are fireproof or are located in an area that will not be affected by an engine fire.

(2) Eight fluid ounces or less of a combustible fluid is not considered hazardous and no more than this amount should be present after activating the shutoff valve.

(3) Engine isolation is to be maintained when incorporating shutoff valves into engine fuel and lubrication lines. The design must insure that when one engine is shut down or fails and the fuel and lubrication fluid shutoff valves are activated, the remaining good engine is not affected in any way, and the rotorcraft can continue safe flight to a landing. This should be demonstrated by test.

(4) Each shutoff valve located in a fire zone should be fireproof. If the shutoff valve is located outside of the fire zone, then it should be at least fire resistant or protected so that it will function under a worst case fire condition within a fire zone. This should be demonstrated by test.

(5) Except for ground-use-only auxiliary power unit installations, the flammable fluid shutoff to all engine installations must be protected from inadvertent operation.

Where electrical shutoffs are used, the switches must be guarded or require double actions. If the shutoffs are mechanically activated, the design of the knob and the location of the lever must be such that inadvertent actuation cannot occur. It must be possible to reopen the shutoff valve in flight after it has been closed and this should be demonstrated by test.

AC 29.1191. § 29.1191 (Amendment 29-3) FIREWALLS .

a. Explanation . This section states the certification requirements for the proper certification of fireproof protective devices such as firewalls, shrouds, or equivalent.

These devices are necessary to isolate each engine (including combustor, turbine, and tailpipe sections of turbine engines and auxiliary propulsion units (APU); each APU; each combustion heater; each unit of combustion equipment; or each high temperature device (or source) from personnel compartments and critical components (not already protected under § 29.1191). The isolation of these fire zones is necessary to prevent the spread of fire, prevent or minimize thermal injuries and fatalities, and prevent Page E - 187 AC 29-2C 9/30/99 damage to critical components that are essential to a controlled landing. Even though § 29.1191(b) implicitly excludes APU’s, combustion heaters, and other combustion equipment that are not used in flight; they should be protected by fireproof enclosures, because of § 29.901(d) and the requirements of the relevant parts of §§ 29.1183 through 29.1203. This is because, even if the device is rendered inoperative in flight, it typically contains residual heat, fuel, fumes and potential ignition sources (i.e., “potential hazards”). Each fireproof protective device must, by regulation, meet the following criteria: (1) Its design and location must take into account the probable fire path from each fire zone or source considering factors such as internal airflow, external air flow, and gravity.

(2) It must be constructed so that no hazardous quantity of air, fumes, fluids, or flame can propagate through it to unprotected parts of the rotorcraft.

(3) Its openings (e.g., shaftholes, lineholes, etc.) must be sealed with close fitting fireproof grommets, bushings, bearings, firewall fittings, or equivalent that prevent burn through and leakage of hazardous fumes or fluids from the fire zone.

(4) It must be fireproof (see definition).

(5) It must be either corrosion resistant or otherwise safely protected from corrosion.

b. Definitions .

(1) Fireproof Protective Device . A fireproof protective device is a device such as a firewall, shroud, enclosure, or equivalent used to isolate a heat or potential fire source (severe fire) from personnel compartments and from critical aircraft components which are essential for a controlled landing.

(2) Fireproof . Fireproof is defined in § 1.1 “General Definitions.” (3) Controlled Landing . A landing which is survivable (i.e., does not fatally injure all occupants) but may produce an unairworthy, partially salvageable, or unsalvageable rotorcraft.

(4) Severe Fire . See definition in paragraph AC 29.859.

c. Procedures . Fireproof protective devices are typically certified by analysis, tests, or a combination conducted during the certification process, including flight tests or simulated flight tests, as follows: (1) Fireproof protective devices should be provided wherever a hazard exists which requires isolation from a severe fire (see definition) to avoid fires in personnel Page E - 188 9/30/99 AC 29-2C compartments and to avoid thermal damage to critical components (such as structural elements, controls, rotor mechanisms, and system components) that are necessary for a controlled landing. A thorough hazard analysis should be conducted during certification to identify, define and quantify in order of severity (i.e., maximum temperature, hot exposed area, etc.) all thermal hazards or zones that require fireproof protection in a given design. Engines (including the combustor, turbine, and tailpipe sections of turbine engines), APU’s, combustion heaters, and combustion devices are required by regulation to be isolated. Other high temperature devices may also require isolation because of local hot spots (which occur during normal operations or from failure modes) that can thermally injure occupants or cause spontaneous combustion of surroundings. A hazard analysis should identify these potential problems and provide proper certification solutions.

(2) Fireproof protective devices should be able to withstand at least 2000 ± 150° F for at least 15 minutes (reference AC 20-135). The fireproof protective device should allow the protected parts, subsystems or systems to perform their intended function for the duration of a severe fire (see definition). For firewalls, examples of flat, geometry materials undergoing uniform heat fluxes with material gauges that automatically meet the certification requirements are given in figure AC 29.1191–1. If firewalls are utilized that involve other materials, significant geometric changes, or significantly non-uniform heat fluxes, then automatic compliance may not be assured. In such cases the fireproof protective devices should be analyzed and, in some cases, tested in accordance with AC 23–2 to ensure proper certification.

For example, a curved protective surface may absorb a uniform incident heat flux unevenly and create a local hot spot that exceeds 2050° F that burns through in less than 15 minutes; whereas, a flat surface of equal thickness would not exceed 2050° F and would not burn through in less than 15 minutes. It should be noted that composite materials are not generally used for protective devices because of their inability to withstand high temperatures (i.e., exceedance of the glass transition temperature); however, some specially formulated composites have been previously certified as engine cowlings. Titanium is an acceptable material for fireproof protective devices such as firewalls. However, use of titanium should always be carefully considered and reviewed, because it can lose all structural ability and burn severely (self combust) above 1,050° F, under certain thermodynamic environments, and contribute to the fire instead of providing the intended fire protection. AC 33–4, “Design Considerations Concerning the Use of Titanium in Aircraft Turbine Engines” and MIL-HDBK-5D contain more detailed information on the unique thermal properties of titanium.

Page E - 189 AC 29-2C 9/30/99 FIGURE AC 29.1191-1 TABLE OF MATERIALS AND GAGES ACCEPTABLE FOR FIREPROOF PROTECTIVE DEVICES WITH FLAT (1) SURFACE GEOMETRIES (2) (3) MATERIAL MINIMUM THICKNESS Titanium Sheet .016 in Stainless Steel .015 in Mild Carbon Steel .018 in Terne Plate .018 in Monel Metal .018 in (4) Firewall Fittings .018 in (Steel or Copper Base) NOTES: (1) Assumes essentially flat vertical or horizontal surfaces undergoing a uniform heat flux. Any significant variation in either geometry or heat flux distribution should be examined in detail for adequate gauge thicknesses on a case-by-case basis.

(2) Must have corrosion protection if not inherent in the material itself.

(3) The minimum thickness is for thermal containment only. Structural integrity considerations may require thickness increases. MIL-HDBK-5D contains material allowable versus temperature data for common metallic materials.

(4) This is the minimum wall thickness measured at the smallest dimension (e.g., thread root or other location) of the part.

(5) Distortion of thin sheet materials and the subsequent gapping at lap joints or between rivets is difficult to predict; therefore, testing of the simulated installation is necessary to prove the integrity of the design. However, rivet pitches of 2 inches or less on non load-carrying titanium firewalls of .020 inch or steel firewalls of .018 inch are acceptable without further testing.

Page E - 190 9/30/99 AC 29-2C (3) The probable path of a fire (as affected by internal and external air flow during normal flight and autorotation, gravity, flame propagation paths, or other considerations) should be taken into account when performing the hazard analysis of item (1). Such a review will insure that fireproof protective devices are placed in the proper location for intercepting, blocking or containing a severe fire before occupants are injured and a controlled landing is prevented. If the probable path cannot be readily determined by inspection or analysis, testing using simulated airflows, rotorcraft attitudes, and dyed inert fluids or vapors can be used to aid in this determination.

(4) Each opening in a protective device should be sealed with close fitting sealing devices such as fireproof grommets, bushings, firewall fittings, rotating seals or equivalent that are at least as effective as the fireproof protective device itself. This is necessary to insure that no local breakdowns in protection occur. For materials not listed as acceptable in item (1), FAA/AUTHORITY standards and analysis and testing should be required in accordance with the definition of a severe fire for proper substantiation.

(5) Each protective device should be fireproof in order to withstand a severe fire (see definition). Unless designs and materials have been previously FAA/AUTHORITY approved (e.g., see Item 1), the protective device’s design and material selection should be tested to insure its fireproof thermal and structural integrity. A full-scale test of a structurally loaded article or a representative sample should be conducted to insure proper compliance is achieved. Also, the continued sealing ability of the protective device in its deformed state due to a hard controlled landing should be considered during certification (e.g., use of ductile materials). The corrosion environment should be defined and appropriate protection provided. Phased inspections should be specified, if necessary, to insure continued corrosion integrity. Certification tests for adequacy of corrosion protection should be conducted, using sample plates or by other equivalent means, as required.

AC 29.1193. § 29.1193 (Amendment 29-13) COWLING AND ENGINE COMPARTMENT COVERING.

a. Explanation .

(1) Section 29.1193(a) requires the cowling and engine compartment coverings to withstand structural loads experienced in flight.

(2) In order to prevent pooling of flammable fluids, § 29.1193(b) requires ventilation and complete drainage from the cowling and engine compartment as specified in § 29.1187.

(3) In § 29.1193(c), (d), and (e), clarification of fireproof requirements is provided along with interaction between the requirements of § 29.1191 for firewalls.

Page E - 191 AC 29-2C, Chg 4 5/1/2014 b. Procedures .

(1) Compliance with § 29.1193(a) can be shown by analyzing the cowling and engine compartment covering and determining that no structural degradation will occur under the highest loads experienced on the ground or in flight.

(2) Compliance with § 29.1193(b) can be accomplished by ensuring that the drain will discharge positively with no traps and is a minimum of 0.25 inches in diameter.

No drain may discharge where it might cause a fire hazard. This can be demonstrated by colored liquid flowing through the drain system while in flight. The dye should not impinge on any ignition source during any approved flight regime.

(3) Compliance with the fireproof requirements of § 29.1193(c), (d), and (e) can be accomplished by demonstrating that the material will withstand a 2,000° F ± 150°F flame for 15 minutes while still fulfilling its design purpose. This testing should accurately simulate, as near as practicable, the likely fire environment to prove the materials and components will provide the necessary fire containment when exposed to a fire situation in service. In addition to the fireproof requirements, the requirements of § 29.1191 must also be met. The primary objectives are: (i) To contain and isolate a fire and prevent other sources of fuel and/or oxygen from feeding the existing fire; and (ii) To ensure that components of the engine control system will function effectively to permit a safe landing and/or shutdown of the engine.

AC 29.1193A. § 29.1193 (Amendment 29-26) COWLING AND ENGINE COMPARTMENT COVERING.

a. Explanation. Amendment 29-26 adds a new § 29.1193(f) that requires redundant retention means for each panel, cowling, engine, or rotor drive system covering that can be opened or readily removed. Conventional fasteners for these devices are subject to frequent operation by maintenance personnel and have deteriorated, failed from wear or vibration, or been left unsecured after preflight inspections. Such a failure could be hazardous if a loose panel, cowling, or covering strikes, or is struck by, the rotors or by critical controls.

b. Procedures.

(1) Compliance with § 29.1193(f) can be accomplished by simulating, or actually failing, one or more of the retention devices or by structural analysis. If a failure of a single retention device can contribute to multiple failures, these multiple failures should be considered. It should be shown that the cowling or cover will not open, strike, or be struck by the rotor or other critical component.

Page E - 192 9/30/99 AC 29-2C (2) Consideration should be given to minimize the possibility of latches being improperly closed that could result in a cowl coming open in flight.

(3) The failure of one latching device should not cause the failure of another latching device.

(4) The consequences of “forgetting” to latch a cowl should be considered.

(5) The use of safety straps should be considered to minimize the impact of a latching device failure.

AC 29.1194. § 29.1194 (Amendment 29-3) OTHER SURFACES .

a. Explanation . This section states the fire resistance requirements for material surfaces near engine compartments and designated fire zones (other than tail surfaces not subject to heat, flames, or sparks emanating from a designated fire zone or engine compartment).

b. Definition .

(1) Other Surface . Any airframe, system, or powerplant component aft of and near an engine compartment, a designated fire zone, or another heat source which would receive a heat flux as a result of a fire in the engine compartment or fire zone that would require the component to be fire resistant.

(2) Fire Resistant . In accordance with § 1.1, is defined as follows: (i) Sheet metal or structural members with the capacity to withstand the heat associated with the fire at least as well as aluminum alloy in dimensions appropriate for the purpose for which they are used.

(ii) Fluid carrying lines, fluid system parts, wiring, air ducts, fittings and powerplant controls with the capacity to perform their intended functions under the heat and other conditions resulting from a fire.

(3) Fire . A fire in either an engine compartment or a designated fire zone is assumed to occur that produces a heat flux on a system, airframe or powerplant component aft of or near the fire. The effect of each such fire on other surfaces must be considered on a case-by-case basis to determine the critical case. Unless a more rationale definition is furnished and approved during certification, the fire in any engine compartment or designated fire zone should be assumed, for purposes of analysis, to be a severe fire (see definition in paragraph AC 29.859).

c. Procedures .

Page E - 193 AC 29-2C 9/30/99 (1) Other surfaces should be identified during certification by a design review and by a conservative, thorough hazard analysis based on an analytical estimate of the total heat flux (i.e., conduction, convection, and radiation in combination, as applicable) using the definition of a severe fire and of the resultant “other surface” temperature based on a single fire occurring in each engine compartment and designated fire zone, on a case-by-case basis. Once the other surfaces are identified and their severe fire induced maximum temperatures determined, their configuration and material selection should be reviewed on a case-by-case basis to determine either that they are fire resistant, that they can be made fire resistant (within the limits of practicability), or that it is impracticable to make them fire resistant. If the non-fire resistant other surfaces can be readily made fire resistant they should be. If it is impracticable to make them fire resistant, then they should be relocated, insulated, or a combination in order to reduce the total incident heat flux (and, thus, lower their surface temperature) so that they no longer need to be fire resistant. If insulation is used to shield a surface that is subjected to a significant temperature, it must be fire resistant.

(2) A partial validation of analytical heat flux models using the definition of a severe fire can sometimes be achieved during certification tests by using thermocouples or heat-sensitive stickers to measure in-flight temperature ranges and distributions on other surfaces from known thermal environments in engine compartments or other designated fire zones.

AC 29.1195. § 29.1195 (Amendment 29-17) FIRE EXTINGUISHING SYSTEMS .

a. Explanation . This section specifies the types of rotorcraft which must have fire extinguishing systems and the number of discharges. The types of tests and airflow conditions are also specified for demonstration of compliance.

b. Procedures .

(1) The requirements are applicable to each turbine engine powered rotorcraft, Category A reciprocating engine powered rotorcraft, and each Category B reciprocating engine powered rotorcraft with an engine of more than 1,500 cubic inches. There must be a fire extinguishing system for the designated fire zones defined in § 29.1181.

(2) A fire extinguishing system should dilute all of the atmosphere within and entering a compartment with sufficient inert agent that it will not support combustion, and continue the process long enough to extinguish the existing flame and either dissipate the vapors or eliminate the ignition sources. Conventional systems utilize perforated tubing or discharge nozzles to distribute a specific quantity of agent in approximately 2 seconds. HRD (high rate of discharge) systems utilize open end tubes to deliver a given quantity of agent within 1.35 seconds for CO and 1 second for all other agents. The HRD systems are recommended for use in compartments having high airflow where the required discharge rates can be more effectively provided by a HRD rather than a perforated tubing system. Tests indicate that unrestricted release through such an open end tube distribution system can be relied on for adequate Page E - 194 9/30/99 AC 29-2C distribution, provided the outlets are located properly. Although the discharge times given above are considered satisfactory, any reduction in discharge time below that specified would improve system effectiveness. However, consideration should be given to the time requirements for draining accumulated combustibles, dissipating combustible vapors and cooling or eliminating ignition sources to assure that the minimum agent concentration is maintained for a duration sufficient to prevent reignition of the combustibles.

(3) The possible variety of tankage and plumbing configurations to accomplish the result should be examined for each specific aircraft in order to achieve the optimum.

Systems can vary from tankage in a central location, which is directed through complex distribution systems to various hazards, to agent which is tanked adjacent to each hazard. Terminology generally accepted to define various arrangements is as follows: (i) Central System : A single supply of agent, centrally located, with valves to direct the agent to any protected zone or zones.

(ii) Individual System : A separate supply of agent for each protected zone or zones.

(4) The selection of the distribution system should be made with full cognizance of the hazards to be covered. The distributor system (i.e., discharge nozzles fed individually by lines from a central manifold) is the most efficient. The complexity of such a system, however, may show it in a less favorable light than the loop or ring system (i.e., orifices drilled in a distribution line, the loop being fed from one end, and the ring being fed from a point on a continuous circle) as far as weight, complexity of manufacture and types of hazard to be covered are concerned. For HRD systems, open feed lines are recommended. In high air flow zones, outlets should be located as far upstream as possible with the discharge directed across the air stream and slightly downstream such that a helical spray pattern is produced. In zones of low or negligible air flow, the outlet location is not critical but a location at the top-center of the zone with the agent directed downward is suggested.

(5) In a conventional CO system, all lines upstream of direction valves should be 4,000 PSI (27,600 kPa) burst and lines which are open should be 2,000 PSI (13,800 kPa) burst. Care should be taken to insure that all valving and/or equipment in the distribution line has an appropriate flow rate. Expansion of fittings, tee, etc., should be checked to insure that not over 150 percent of the inflow area exists downstream.

130 percent is accepted as the safest target value. If overexpansion occurs, snow will form and plug the lines. Because of high storage pressure, the orifice areas of a conventional CO system seem to act as the flow control with system flow losses as a minor effect. Because of this, distribution systems of 50 ft. (15.24m) or less can be satisfactorily computed by the following factor: (i) Line Area = .10 sq. in./lb CO /sec (142.2 mm /kgCO /sec).

2 2 Page E - 195 AC 29-2C 9/30/99 (ii) Orifice Area = .072 sq. in./lb CO /sec (102.4 mm /kgCO /sec) 2 2 (72 percent of equivalent line area).

(iii) Min. Orifice Size = 1/16 in. (1.6 mm) diameter.

Br, line and fitting losses (6) In low pressure systems such as “CB” and CH become a greater effect in the discharge rates and distribution than was true with CO .

Consideration should be given to the small I.D. of an AN line fitting with respect to the I.D. of the mating tube sizes. This may be done by extra pressure drop allowances, by enlarging these fittings, or by making special fittings. Within reasonable line lengths, however, area factors can be used with fair accuracy. (It is generally conceded that a system designed to these factors, especially a complex layout, should be carefully tested or analyzed for time of discharge and distribution.) These areas are as follows: (i) Line Area = .07 sq. in./lb agent/sec (99.6 mm /kg agent/sec).

(ii) Orifice Area = .05 sq. in./lb agent/sec (71.1 mm /kg agent/sec) (72 percent of equivalent line area).

(iii) Min. Orifice Area = 1/32 in. (.8 mm) diameter.

(7) For HRD systems of all types, feed line cross-sectional area is dependent upon the rate desired and upon system volume considerations. The minimum diameter of the feed line is established by the required rate; the maximum diameter of the feed line, and by the need for keeping the system volume at a minimum. Specifically, with the propelling gas in a system pressurized to 400 PSI (2760 kPa), the “volumetric efficiency” should be at least 0.50; that is, the original volume of the propelling gas in the system should be at least ½ the volume of the entire system, including that of the agent container. It is recommended that for HRD systems the feed lines be open. No nozzles or series of perforations are required. It is believed that the unrestricted release of the more volatile liquid agents, as well as carbon dioxide, can be relied upon for adequate distribution, provided the outlets are properly located. It is important that any such system be carefully tested for time of discharge, distribution, and minimum concentrations.

(8) From the basic definition, the system should be effective if the distribution of the agent floods the various portions of a compartment simultaneously and dilutes the incoming air. It is noted that the typical high flow compartment requires a greater proportion of its total agent discharged at the air inlet than does the conventional low air flow zone. All parts of the fire extinguisher system directed to any one powerplant installation should be discharged simultaneously. The theory behind the HRD type system is that with rapid discharge of the agent, the concentration necessary for extinguishment is reached more rapidly with correspondingly less time for dissipation or dilution of the agent by incoming air. The duration of this critical concentration necessary for extinguishment is believed to remain the same as for conventional systems.

Page E - 196 9/30/99 AC 29-2C (9) Detailed system configuration recommendations are not available for conventional systems; however, the recommendations on the configuration of HRD systems would probably apply equally well to all types. For HRD systems, it is recommended that feed lines be as short as possible, requiring that agent containers be as close as practical to the zones to be protected. Feed lines should be direct; the fewer fittings and turns, the better. Expansions and restrictions have adverse effects on rate; and it is probable that in a feed line with long rises or many changes or direction, quantities of propelling gas can get past a liquid agent, thus reducing the discharge rate and making the discharge sporadic and ineffective. Where such fittings, changes of direction, and long vertical rises are unavoidable, compensation in the form of additional agent may be necessary.

(10) A fixed “one shot” fire extinguisher system should be provided for the heater extinguisher system in order to extinguish the fire in the combustion chamber.

The regions surrounding the heater and combustion chamber must also be protected if these regions contain components with potential combustible leakage. No fire extinguishment is needed in cabin air passages.

AC 29.1197. § 29.1197 (Amendment 29-13) FIRE EXTINGUISHING AGENTS .

a. Explanation .

(1) Fire extinguishing agents used in rotorcraft fire extinguishing systems must be capable of extinguishing any fire in the area where the system is installed.

(2) The extinguishing agent must maintain its effectiveness after prolonged storage under the environmental conditions of the compartment in which it is stored.

(3) If a toxic extinguishing agent is used, the harmful concentration level of the fluid vapors must be determined and it must be shown that it is not possible for this concentration level to enter into any personnel compartment.

b. Procedures .

(1) The fire extinguishing system should dilute all of the atmosphere within and entering a compartment with sufficient inert agent so that combustion cannot be supported. The extinguishing process should continue for a duration sufficient in length to extinguish any existing flame. When a compartment is to be flooded with agent and there is a source of fresh air entering the compartment, the incoming air should be either shut off prior to the release of the agent or rendered inert by directing extinguishing agent into the air blast (preferably the former) or the quantity of agent should be increased to offset the incoming airflow.

Page E - 197 AC 29-2C 9/30/99 (2) There are a number of extinguishing agents which have been used on rotorcraft in the past. The following list identifies the agent and some advantages and disadvantages of each.

Agents Advantages Disadvantages Carbon Dioxide Safest agent to use from the Mental confusion and suffocation CO standpoint of toxicity and corrosion hazard to occupants if sufficient gas hazards. is discharged into personnel compartments. CO has an extremely large variation in vapor pressure with temperature which makes it necessary to use stronger (heavier) containers than are required for methyl bromide.

Methyl More effective for equal mass than Much more toxic than CO . Due to Bromide CO . Approx. 80 percent of this its toxic effects on humans, CH Br 2 3 CH Br agent by weight as compared to CO should not be used as a fire 3 2 is required. extinguisher agent in areas where harmful time concentrations can Less variation in vapor pressure than enter personnel compartment.

CO . Much lower container pressure required resulting in lighter Aluminum alloy material should not containers. Treated magnesium be used in methyl bromide systems alloys are satisfactory for use in due to serious corrosion and possible CH Br systems outside of the spontaneous ignition. Rapidly potential fire zones. corrodes aluminum, magnesium, and zinc.

Tubing systems should be vented at all times and steps should be taken to free the tubing of residual methyl bromide after each discharge.

Containers must be recharged at the extinguisher manufacturer’s plant or at a depot by specially trained personnel.

Bromo- Low vapor pressure compound - Toxic when burned.

chloro - 3 PSIA (20.7 Kpa) at 70° F (21.1°C).

methane (“CB”) One of the more effective agents.

CH BrCl Dibro - Low vapor pressure compound - Very toxic when burned.

modi- 14 PSIA (96.5 Kpa) at 70° F Page E - 198 9/30/99 AC 29-2C Agents Advantages Disadvantages fluoro - (21.1° C). One of the more effective methane agents.

CF Br Non-corrosive to aluminum, steel and brass.

Bromotri One of the more effective agents. High vapor pressure compound - fluoro - 220 PSIA (1517 Kpa) at 70° F methane Low toxicity in natural condition and (21.1° C).

CF Br when burned.

Least toxic of agents in burned Non-corrosive to aluminum, steel and condition except for CO .

brass.

Nitrogen N If a fuel tank inerting system using N 3 - 4 times quantity and rate of 2 2 is provided, use as extinguishing conventional agents required.

agent may be considered. N offers cooling not available with CF Br.

Note: The relative effectiveness of the various agents listed above is considerably Influenced by the type of system employed, high rate discharge or conventional; by the method of distribution, open end outlet, nozzle, or spray ring; and by the air flow conditions.

(3) The extinguishing agent must not be affected by the temperature extremes experienced in the compartments in which they are stored. The agent containers should be either “winterized” for extreme temperature operation or so located in the rotorcraft that they will not be subjected to extreme temperatures. Safe limits for unwinterized carbon dioxide cylinders are approximately 0° F (-18° C) to 140° F (60° C).

The cartridge detonators have a variable age-with-temperature limit. Contact should be made with the manufacturer for the latest information available for both installation and storage temperatures.

(4) It must be shown by test that the harmful level of toxic fluid or vapors cannot enter into any personnel compartment due to leakage or activation of the system during normal operation of the rotorcraft in flight or on the ground. The entire fire extinguishing system should be mocked-up or installed in the aircraft down to and including distribution tubing and outlets. The tests should be conducted under actual or simulated cruise conditions. The system should be discharged, and compliance verified by use of an appropriate method for measuring agent concentration.

Page E - 199 AC 29-2C 9/30/99 AC 29.1199. § 29.1199 (Amendment 29–13) EXTINGUISHING AGENT CONTAINERS .

a. Explanation .

(1) This section presents the requirements for fire extinguisher containers. The containers are subjected to high internal pressures for the propulsion of the agent as well as a wide range of external environmental temperatures.

(2) The containers must be adequately protected to preclude any adverse effect on the operation of the system from these external influences.

b. Procedures.

(1) Each extinguishing agent container must have a pressure relief valve which will open at a pressure that is below the burst pressure of the agent container. The pressure relief valve lines must be located and protected so that they cannot be clogged by dirt, ice, or other contaminates. Both the agent container burst pressure and the relief valve opening pressure limits should be verified by test. Agent containers which meet military specification, MIL-C-22284, requirements are acceptable.

(2) The containers should be located so that an indicator is readily visible to determine if the container has discharged or the charging pressure is below operating minimums. The number and size of agent containers should be adequate to obtain the established agent concentration and duration for the intended compartment. It is preferred that the agent supply containers and the flow control valves are not located in a fire zone.

(3) The brackets for mounting the containers and securing the discharge lines should be designed to withstand all loads to which they may be subjected due to recoil during discharge or any other applied load factor.

(4) The agent containers should be protected from extreme temperature excursions which could have an adverse effect upon the operation of the extinguishing system. Safe temperature limits for “unwinterized” carbon dioxide cylinders are approximately 0° F (-18° C) to 140° F (60° C). Safe limits for “CB” and CH Br spheres are approximately -65° F (-54° C) to 200° F (93° C). The cartridge detonators have a variable age-with-temperature limit and the manufacturer should be contacted for the latest information on installation and storage temperatures. Location of the container in the aircraft should take these temperature limits into consideration.

Page E - 200 9/30/99 AC 29-2C AC 29.1201. § 29.1201 FIRE EXTINGUISHING SYSTEM MATERIALS .

a. Explanation .

(1) Many different fire extinguishing agents are available for use in fire extinguishing systems. The choice of extinguishing agent should take into account the chemical reaction (if any) between the extinguishing agent and the materials utilized in the extinguishing system. If there are any incompatibilities, they should not create a hazard by creating volatile or toxic vapors or fumes which could feed a fire or cause injury to passengers, crew, or other personnel.

(2) The fire extinguishing components in an engine compartment must be fireproof to ensure operation in the event of a compartment fire.

b. Procedures .

(1) Compliance with the requirements of § 29.1201(a) can be demonstrated by analysis, test, or a combination of both.

(2) Certification data submitted by the applicant should contain a listing of the chemical ingredients of the extinguishing agent and the other materials in the extinguishing system. These data should also show that the chemical reaction (if any) of these materials, when combined, does not create a hazard.

(3) Where chemical compounds exist and the chemical reaction is not predictable when two different compounds are combined, actual tests may be necessary to determine the hazard potential.

(4) Analysis, test, or a combination of both may be used to demonstrate compliance with the fireproof requirement for all fire extinguishing components located within the engine compartment.

AC 29.1203. § 29.1203 (Amendment 29-40) FIRE DETECTOR SYSTEMS .

a. Explanation .

(1) Fire detection systems are required in turbine engine powered rotorcraft, Category A reciprocating engine powered rotorcraft and each Category B reciprocating engine powered rotorcraft where the engine displacement is greater than 900 cubic inches.

(2) This section specifies material, installation, and some operational requirements for fire detectors to ensure prompt detection of fire in the fire zones and other designated areas.

b. Procedures .

Page E - 201 AC 29-2C 9/30/99 (1) The detector system should be designed for highest reliability to detect a fire and not to give a false alarm. It is desirable that it only responds to a fire and misinterpretation with a lesser hazard should not be possible. Engine overtemperature, harmless exhaust leakage, and bleed air leakage should not be indicated by a fire detector system. A fire detection system should be reserved for a condition requiring immediate measures such as engine shutdown or fire extinguishing. There are three general types of detector-procedure systems that are commonly used: (i) A manual system utilizes warning lights to alert the pilot who then follows prescribed cockpit procedure as a countermeasure. A manual system is adequate for hazards in which a few seconds are not important.

(ii) There is also a semi-automatic system. Occasionally a rotorcraft becomes so complex that the emergency procedure exceeds reasonable expectations of the pilot. In such cases, psychology should be weighted against complexity, and “panic switches,” combining multiple procedure functions, should be provided to simplify the mental demands on the pilot. Speed is gained by such designs for hazards which may need it.

(iii) The detector of an automatic system automatically triggers the appropriate countermeasures and warns the pilot simultaneously. Such a system should be carefully evaluated to assure that the advantages outweigh the disadvantages and potential malfunctions.

(2) Fires, or dangerous fire conditions can be detected by means of various existing techniques. The following is a partial list of available detectors: (i) Radiation-sensing detectors.

(ii) Rate-of-temperature-rise detectors.

(iii) Overheat detectors.

(iv) Smoke detectors.

(v) CO detectors.

(vi) Combustible mixture detectors.

(vii) Fibre-optic detectors.

(viii) Ultraviolet.

(ix) Observation of crew or passengers.

Page E - 202 9/30/99 AC 29-2C (3) In many rotorcraft it is desirable to have a detection system which incorporates several of these different types of detectors. Radiation-sensing detectors are most useful where the materials present will burn brightly soon after ignition, such as in the powerplant accessory section. Rate of rise detectors are well-suited to compartments of normally low ambient temperatures and low rates of temperature rise where a fire would produce a high temperature differential and rapid temperature rise. It should be noted that under certain circumstances, where a relatively slow temperature increase occurs over a considerable period of time, a fire can occur without detection by rate of rise detectors. Overheat detectors should be used wherever the hazard is evidenced by temperatures exceeding a predicted, set value. Smoke detectors may be suited to low air flow areas where materials may burn slowly, or smolder. Fibre-optic detectors can be used to visually observe the existence of flame or smoke. The three major detector types used for fast detection of fires are the radiation-sensing, rate-of-rise, and overheat detectors. Radiation-sensing detectors are basically “volume” type which senses flame within a visible space. Overheat-fire detectors can be obtained in either “continuous” or “unit” type.

(4) The detector system should: (i) Indicate fire within 15 seconds after ignition, and show which engine compartment in which the fire is located.

(ii) Remain on for the duration of the fire.

(iii) Indicate when the fire is out.

(iv) Indicate re-ignition of the fire.

(v) Not by itself precipitate or add to the potential of any other hazards.

(vi) Not cause false warnings under any flight or ground operating condition.

A false fire detector indication could significantly increase crew workload, impair crew efficiency, or reduce safety margins and so is classified as a major failure condition. In consequence, such false fire detector indication should be shown to be improbable based on a probability assessment and service experience of the fire detector system.

If the probability of the fire detection system experiencing a false indication cannot be shown to be improbable, a secondary means of determining the validity of the fire indication should be provided.

(5) Additional features of the detection system are as follows: (i) A means should be incorporated so that operation of the system can be tested from the cockpit.

Page E - 203 AC 29-2C 9/30/99 (ii) Detector units should be of rugged construction, to resist maintenance handling, exposure to fuel, oil, dirt, water, cleaning agent, extreme temperatures, vibration, salt air, fungus, and altitude. Also, they should be light in weight, small, and compact, and readily adaptable to desired positions of mounting.

(iii) The detector system should operate on the rotorcraft electric system without inverters. The circuit should require minimum current unless indicating a fire or unless a monitoring system is in use.

(iv) Fixed temperature fire detectors should preferably be set at 100° F (37.7° C) to 150° F (65.6° C) above maximum safe ambient temperature, or higher when in compartments where extremely high rate of rise is normally encountered.

(v) Detector system components located within fire zones should be fire resistant.

(vi) Each detector system should actuate a light which indicates the location of the fire. If fire warning lights are used, they must be in the pilot’s normal field of view.

(vii) Two or more engines should not be dependent upon any one detector circuit. The installation of common zone detection equipment prevents the detection system from distinguishing between the engine installations, necessitating shutting down more than one engine.

(6) The sensing portion of the fire detection system should not extend outside of the coverage area into another fire zone. Detectors, with the exception of radiation– sensing detectors, should be located at points where the ventilation air leaves compartments. If a reverse-flow cooling system is used, detectors should be installed at locations which are outlets under both flight and ground operating conditions. Stagnant air spaces should be avoided and the number of ventilation air exits should be kept to a minimum. The ventilation requirements of § 29.1187(e) must also be taken into consideration. Compliance with these recommendations allows the effective placement of a minimum amount of detectors, and still ensures prompt detection of fire in those zones. Radiation-sensing detectors should be located such that any flame within the compartment is immediately sensed. This may or may not be where the ventilation air leaves the compartment.

(7) Fire detectors must be installed in designated fire zones, the combustor, turbine and tailpipe sections of turbine installations.

(i) Engine Power Section (Combustor, Turbine and Tailpipe): This zone is usually characterized by predictable hazard areas which facilitate proper detector location. It is recommended that coverage be provided for any ventilating air outlet as well as intermediate stations where leaking combustibles may be expected.

Page E - 204 9/30/99 AC 29-2C (ii) Compressor Compartment: This is usually a zone of relatively low air flow velocities, but wide geographical possibility for fires. When fire detectors other than radiation-sensing detectors are used, detection at air outlets provides the best protection, and intermediate detector locations are of value only when specific hazards are anticipated.

(iii) Accessory Bullet Nose: Where such a compartment is so equipped that it is a possible fire zone, its narrow confines permit sufficient coverage with one or more detectors at the outlets.

(iv) Heater Detector Location: An overheat detector should be placed in the hot air duct downstream of the heater. If the heater fuel system or exhaust system configuration is such that it is a fire hazard, the compartment surrounding the heater should also be examined as a possible fire zone.

(v) Auxiliary Power Unit Detector Location: The use of a combustion-driven auxiliary power unit creates another set of typical engine compartments defined and treated as above. Some units are so shrouded with fireproof material that these compartments exist only within the confines of the shroud. They are still, however, fire zones and must have a detection system.

Page E - 205 2/4/2016 AC 29-2C, Chg 7 CHAPTER 2. PART 29 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT SUBPART F - EQUIPMENT EQUIPMENT - GENERAL AC 29.1301. § 29.1301 FUNCTION AND INSTALLATION.

a. . It should be emphasized that this rule applies to each item of Explanation installed equipment which includes optional equipment as well as required equipment.

b. Procedures.

(1) Information regarding installation limitations and proper functioning is normally available from the equipment manufacturers in their installation and operations manuals. In addition, some other paragraphs in this AC include criteria for evaluating proper functioning of particular systems. (An example is paragraph AC 29 MG 1 for avionic equipment.)

(2) This general rule is quite specific in that it applies to each item of installed equipment. It should be emphasized, however, that even though a general rule is relevant, a rule that gives specific functional requirements for a particular system will prevail over a general rule. Therefore, if a rule exists that defines specific system functioning requirements its provisions should be used to evaluate the acceptability of the installed system and not the provisions of this general rule. It should also be understood that an interpretation of a general rule should not be used to lessen or increase the requirements of a specific rule. Section 29.1309 is another example of a general rule, and this discussion is appropriate when applying its provisions.

AC 29.1303. § 29.1303 (Amendment 29-24) FLIGHT AND NAVIGATION INSTRUMENTS.

a. Explanation . This rule lists the flight and navigation instruments that are required for VFR certification. Several additional rules to be consulted when determining the flight and navigation instrument installation design are § 29.1321, Arrangement and visibility; § 29.1331, Instruments using a power supply; and § 29.1333, Instrument systems. Additional information regarding the different instruments can be found in the Technical Standard Order (TSO) for each one. Compliance with the TSO requirements does not ensure compliance with the appropriate part 29 installation requirements.

Other considerations may also be found by reviewing the requirements of §§ 29.1323, 29.1327, 29.1335, 29.1381, 29.1543, 29.1545, and 29.1547, and Appendix B to part 29.

Paragraphs VIII.(a) and (b) of Appendix B to part 29 include IFR operation considerations for flight and navigation instruments. In addition, if the maximum allowable airspeed is dependent on conditions such as weight and altitude, that Page F - 1 AC 29-2C, Chg 7 2/4/2016 information is normally provided on tables, graphs, placards, or other means in the cockpit and in the rotorcraft flight manual.

(1) This guidance is being revise to clarify the requirements for the non-tumbling gyroscopic bank and pitch indicator required by § 29.1303(f) relative to the specified requirements in § 29.1303(g)(1). A review of § 29.1303 preamble material revealed that the intent of the regulation was to adopt similar levels of safety to transport category rotorcraft as that found in transport category airplanes. Therefore, the requirements of § 29.1303 for transport category rotorcraft are similar to those in § 25.1303 for transport category airplanes, with the exception that transport airplanes require a usable range of 360º in pitch and roll.

(2) For rotorcraft, loss of or misleading primary flight information (attitude, altitude, and airspeed) is considered a catastrophic failure condition in instrument meteorological conditions. For an attitude instrument to be usable, it must be capable of providing the pilot with reliable references to pitch and roll attitudes throughout the instrument’s operating range so a pilot can correctly recognize the extent of the unusual or extreme attitude and initiate an appropriate recovery maneuver. As indicated previously, a TSO approval does not ensure compliance with the part 29 installation requirements, including those requirements in § 29.1303(g)(1).

(i) The minimum usability requirements for the aircraft attitude systems is defined in § 29.1303(g)(1). The phrase in § 29.1303(g)(1) “…is usable through +/- 80 degrees of pitch and +/- 120 degrees of roll" means that the pilot must be able to quickly and accurately determine the aircraft’s pitch attitudes up to 80 degrees nose up and 80 degrees nose down. The ADI must also allow the pilot to quickly and accurately determine the aircraft’s roll attitude to 120 degrees of left and right roll.

(ii) The minimum usability requirement for the aircraft attitude system defined in § 29.1303(g)(1) applies to all attitude systems installed in the aircraft. Attitude systems that do not meet the minimum usability requirements can provide misleading information to the pilot.

(3) Operational rules require equipment with specific operational capabilities (for example, 14 CFR 91.205(d), 135.159(2), and 135.159(3)). Therefore, it is advisable to check the operational rules for the type of operation being considered to ensure the attitude indicator meets the requirements for that operation.

b. Procedures .

(1) The following instruments are flight instruments: (i) Airspeed indicator.

(ii) Sensitive altimeter.

Page F - 2 2/4/2016 AC 29-2C, Chg 7 (iii) Free-air temperature indicator.

(iv) Non-tumbling gyroscopic bank and pitch indicator.

(v) Gyroscopic rate-of-turn indicator with an integral slip-skid indicator.

(vi) Rate-of-climb (vertical speed) indicator.

(2) The remaining instruments are navigation instruments.

(3) If a speed warning device is required to be included as part of the rotorcraft design, it must meet the performance requirements given in the rule. In addition, the evaluation of the acceptability of the aural warning (i.e., this warning differs distinctively from aural warnings used for other purposes) should be accomplished by flight test personnel as part of their overall cockpit evaluation.

[Next section AC 29.1305 starts on Page F – 7] Page F - 3 AC 29-2C, Chg 7 2/4/2016 [This page intentionally blank.]

Page F - 4 2/4/2016 AC 29-2C, Chg 7 [This page intentionally blank.]

Page F - 5 AC 29-2C, Chg 7 2/4/2016 [This page intentionally blank.]

Page F - 6 2/4/2016 AC 29-2C, Chg 7 AC 29.1305. § 29.1305 (Amendment 29-10) POWERPLANT INSTRUMENTS.

a. Explanation . This section specifies those instruments which are required for reciprocating and turbine engine installations. It also provides instrumentation requirements for operating rotorcraft in Category A or Category B. These instruments will provide the pilot with essential data to determine operational status of critical components and select desired performance conditions.

b. FAR 29.1305(a)(4), (a)(6), and (a)(9) requires a warning device for low fuel, gearbox oil pressure, and transmission oil temperature. An indicator/gage is not acceptable for use as a warning device since the indicator/gage is not a primary instrument and therefore is not actively monitored.

c. There are advanced display systems that take advantage of microprocessor power by integrating the processing of several parameters. These systems have to date been referred to as Engine Caution Advisory Systems (ECAS) or as Integrated Instrument Display Systems (IIDS) and possibly other variations of these names. These systems typically integrate propulsion instruments, fuel quantity indication, and caution and warning system into a single display system. In traditional designs the powerplant instruments, fuel quantity display, and the caution and warning system are independent from each other. The integration of these systems/indicators eliminates their independence from one another and increases the probability of loss of more than one indicator/system as a result of a single fault or malfunction. Redundant design is generally applied to compensate for the loss of independence.

(1) This integration and resultant mitigation of independence can result in an increased opportunity for common mode failures. Approval of the compensating features is elevated in importance as it is this aspect that allows the concept to be acceptable and subsequently certifiable. The loss of all displayed information or erroneous information should be considered for determination of worst case criticality.

With this determination of criticality, the design can be evaluated to see that it meets the [Section AC 29.1305 continued on next page] Page F - 7 AC 29-2C 9/30/99 minimum associated level of design assurance. Additionally, due to space limitations, some systems employ “page over” features that may have some difficulty displaying the required information when needed and human factors aspects must be considered.

(2) The instrument display system must be investigated and found to be acceptable under both normal and emergency conditions, must perform its intended function under foreseeable operating conditions, and must be designed to minimize the hazards in the event of probable malfunction or failure.

(3) It must be shown that there is appropriate redundancy to provide adequate compensation for the loss of independence in the system. If a multi-page system is employed, it must be shown that needed information is displayed when required.

Specific issues that must be addressed to assure compliance with the minimum safety standards are as follows: (i) The level of most severe hazard must be determined.

(ii) Equivalent reliability and software design assurance to the determined criticality level must be shown.

(iii) Pretest capability must be provided for the warning and caution system to preclude an associated latent failure.

(iv) Human Factors (reference §§ 29.1321 and 29.1322).

d. Additional rules to be consulted when determining the powerplant instrument installation design are §§ 29.1321, Arrangement and visibility; 29.1337, Powerplant instruments; 29.1381, Instrument lights; 29.1543, Instrument markings: General; 29.1549, Powerplant instruments; 29.1551, Oil quantity indicator; and 29.1553, Fuel quantity indicator.

AC 29.1305A. § 29.1305 (Amendment 29-26) POWERPLANT INSTRUMENTS.

a. Explanation . Amendment 29-26 revises, edits, and adds new powerplant instrument requirements. Section 29.1305(a)(4) was revised to require a low fuel warning device for each tank that can be used to feed an engine. The amendment allows a longer time between warning actuation and fuel exhaustion and requires the low fuel warning device to be independent of the normal fuel quantity indication system.

Section 29.1305(a)(17) was changed to extend its application to all rotorcraft (not just those with turbine engines) and to require an indication to the crew of the degree of filter blockage as it relates to the fuel flow requirements in § 29.955. Section 29.1305(a)(19) was revised to require function indicators only for fuel heaters that can be selected or are controllable. A new paragraph (a)(20) was added to § 29.1305 that combined identical requirements for fuel pressure indicators in paragraphs (b)(2) and (c)(2) and modified the applicability of these requirements to only Page F - 8 9/30/99 AC 29-2C those fuel systems with devices or components that could adversely affect fuel pressure at the engine, if they fail. It also eliminates the requirement for fuel pressure indicators in fuel systems, such as suction or gravity feed systems, which do not incorporate pumps or filters. A new § 29.1305(a)(21) was added that requires a warning device to indicate to the flight crew the failure of any fuel pump that is required to supply adequate fuel flow to the engine according to § 29.955; such indication is not required for fuel pumps which are demonstrated to be only necessary for engine starting.

Section 29.1305(a)(22) adds a requirement for a warning or caution device to alert the flight crew when particles are detected by the chip detector required by § 29.1337(d). A new § 29.1305(a)(23) added a requirement for powerplant instruments or warning devices for auxiliary power units installed in rotorcraft.

b. Procedures . The requirement and purpose for each instrument is self-explanatory in the amendment. Other sections that should be considered when designing powerplant instruments are listed in paragraph AC 29.1305.

AC 29.1305B. 29.1305 (Amendment 29-34) POWERPLANT INSTRUMENTS .

a. Explanation . Amendment 29-34 added §§ 29.1305(a)(24) and 29.1305(a)(25) to provide for 30-second/2-minute OEI power ratings.

(1) Section 29.1305(a)(24) adds the requirement that a device or means be provided to alert the crew of the use of the 30-second and 2-minute OEI power level.

The crew should be alerted when the 30-second or 2-minute interval begins and when the time interval ends. The amount of time spent at the 30-second or 2-minute OEI power levels is at the crew’s discretion, unlike the other limits for 30-second OEI that are set by an automatic control required by § 29.1143. The purpose for providing the time interval alerts and automatically controlling the 30-second OEI limits is to free the crew from monitoring the engine instruments during critical phases of flight caused by the loss of an engine.

(2) Section 29.1305(a)(25) adds the requirement for a device to record the usage and the amount of time spent at the 30-second OEI power level and the amount of time spent at the 2-minute OEI power level. The information recorded by this device is for the use of the ground crew to determine if maintenance actions/inspections are to be conducted.

b. Procedures . For the purpose of complying with FAR 29.1305(a)(24) and 29.1305(a)(25), the 2-minute OEI power level is considered to be achieved whenever one or more of the operating limitations applicable to the next lower OEI power rating are exceeded. The 30-second OEI power level is considered to be achieved whenever one or more of the operating limitations applicable to the 2-minute OEI power rating are exceeded.

(1) A review of the method to meet the requirements of § 29.1305(a)(24) should be conducted by flight test personnel. A determination should be made as to whether Page F - 9 AC 29-2C 9/30/99 the method used to alert the crew of 30-second or 2-minute OEI power usage can be recognized and understood by the crew.

(2) To meet the requirements of Section 29.1305(a)(25), a device should be installed on the engine or the airframe to record the time and each usage of 30-second and 2-minute OEI power levels. The information on the time and usage of 30-second and 2-minute OEI power should be recoverable from the recording device by ground personnel. The device should not be capable of being reset in flight and should only be capable of being reset by ground personnel. Prior to each flight this device should be capable of being checked for proper operation and to determine if 30-second or 2-minute OEI power levels were used during the previous flight.

c. Integrated Display Systems . This advisory material is to provide guidance for compliance to Part 27 and Part 29 regulations as they apply to integrated display systems. The integration aspects of these systems require some additional issues to be addressed during certification. The term “must” in this advisory material is used in the sense of ensuring the applicability of these particular methods of compliance when the acceptable means of compliance described herein is used. This advisory material establishes an acceptable means, but not the only means of certifying an integrated display system.

(1) Definitions .

(i) Integrity . The term “integrity” for the purpose of this advisory material includes the hardware quality requirements, including reliability; as well as the software level requirements, as defined in DO178B.

(ii) Criticality . The term “criticality” refers to the five levels of criticality addressed in FAA Advisory Circulars AC 27-1B and AC 29-2C.

(2) Related documents .

(i) Federal Aviation Regulations (FARs) paragraphs 21.21, 29.1301, 29.1309, 29.1305, and 29.1322 (ii) Standards - Latest revision of RTCA/DO178 and RTCA/DO-160; SAE documents (iii) ARP4754 and ARP4761 (3) Background . A tendency to integrate functions/indications that have previously been independent is a result of technology advancement. Microprocessor driven systems have facilitated the ease of this integration. Integrated Instrument Display System (IIDS) or Engine Instrument and Caution Advisory System (EICAS) are examples of integrated display systems. IIDS, EICAS, or any other similar systems are defined as a combination of engine instruments ( previously independent indicators), Page F - 10 5/1/2014 AC 29-2C, Chg 4 fuel quantity indication, and caution/warning parameters, as a minimum, presented by a common display driven by a common processor.

(4) Discussion. This design philosophy does not result in the traditional requirement for individual display independence for failure/malfunction considerations.

This loss of independence means that a single failure could result in loss of most, if not all, instrument displays on the integrated display system. Redundancy of the integrated display system is often proposed to compensate for this lack of independence.

However, redundancy alone may not meet the integrity requirements since they are derived from the level of criticality associated with the loss or malfunction of instrument/parameter displays for flight operations that are dependent on these indications.

(5) Certification Approach . A two step procedure should be used to determine the adequate safety level for this type of system. The first step is to determine the level of criticality associated with the total loss/malfunction of these functions/indications or loss/malfunction of the critical parts of the display. This can be achieved through the use of a functional hazard assessment (FHA). This criticality assessment must be a product of failure/malfunction of the indication system and the flight operation that would represent the worst case for loss of this information. The second step is to determine that the design integrity of the system is at least equal to the assessed criticality level determined in step one.

(6) F unctional Hazard Assessment. The operational classifications to be considered when assessing the criticality are Cat A, Cat B, and IFR. The need for critical information varies with each of these different operational categories. An example would be the demand for OEI parameter information in the single engine Cat A operation. Another example is the loss of fuel quantity indication and fuel low level indication simultaneously in IFR flight conditions. The FHA should address not only loss of one type of indication, but combined loss of engine parameter indication, including total loss of display information, caution/warning, fuel quantity indication, and any other included display in combination with a particular flight operation. There are techniques to lessen the consequences of the failure/malfunction requirements for integrity, such as providing back-up displays for the information deemed critical for a particular operational consideration.

Summary (7) . The loss of all integrated display information for certain types of flight operations may have the highest level of criticality associated with it. The same may be true for malfunctions that result in misleading indications. These failures/malfunctions must be addressed by the commensurate design integrity level.

Lesser levels of criticality must also be addressed by the appropriate design integrity levels.

Page F - 11 AC 29-2C, Chg 4 5/1/2014 AC 29.1305C. § 29.1305 (Amendment 29-40) POWERPLANT INSTRUMENTS.

a. Explanation . Amendment 29-40 added section 29.1305(a)(6) to require an oil pressure indicator for each pressure-lubricated gearbox. Paragraphs (a)(6) through (a)(25), prior to this amendment, have been redesignated as paragraphs (a)(7) through (a)(26).

Procedures b. . In addition to providing an oil pressure indicator for each pressure-lubricated gearbox, the guidance material of the previous AC 29.1305 paragraphs continues to apply.

Page F - 12 6/23/23 AC 29 - 2C, Chg 9 AC 29.1305D. § 29.1305 (AMENDMENT 29 - 59 ) POWERPLANT INSTRUM ENTS.

a. Explanation . Amendment 29 - 59 revised § 29.1305 to include the use of technology such as a synthesized power indicator (SPI), and added § 29.1305(b)(4) to address One Engine Inoperat ive (OEI) Training Mode.

(1) Revised § § 29.1305(a)(5), (a)(11), and (a)(12) to incorporate flexibility for the use of a synthesized power indicator (SPI).

(2) Added § 29.1305(b)(4) to provide a requirement for when an applicant requests approval for OEI Training Mode. OEI Training Mode is intended to simulate an engine failure, provide an annunciation of that simulation, and indicate the actual power being provided by the engines to the flight crew.

b. Procedures . The purpose of the SPI display is to reduce pi lot workload by providing an indicator showing a single parameter that is closest to the applicable engine or transmission limit, versus monitoring individual engine parameters. The SPI display for reciprocating engines presents a normalized indication of the current engine manifold pressure and engine RPM. For turbine engines, the SPI display presents normalized indications of the engine power or torque, gas temperature, and engine compressor speed output. The normalized value is presented to the flight cr ew on the primary flight displays, along with a caption indicating the nearest engine limiting parameter (engine manifold pressure, RPM, torque, gas temperature indicator, engine compressor speed, as applicable) that is being used for the SPI display calcu lation.

The SPI display function is typically performed by the avionics suite using data provided by the engine FADEC.

Section 29.1305 establishes the required powerplant instruments and warning and caution annunciations that must be displayed within th e view of the flying pilot. These have traditionally been independent instruments and are needed not only to measure the performance output of the engines, but they collectively allow the flying pilot to continuously monitor the condition and health of th e engines.

The absence of continuously displayed engine manifold pressure and RPM for reciprocating engines or engine torque, gas temperature, and engine compressor speed parameters may affect the pilot’s ability to directly monitor the health and condi tion of each engine. In addition, the absence of individual gauges deprives the pilot of the ability to quickly assess trend or rate of change information and proximity to individual parameter limits, and to accurately compare engine - to - engine data. As a result, in addition to the requirements in § § 29.1305 and 29.1549, the SPI display must also meet the requirements of § 29.1321.

Page F - 12.1 AC 29 - 2C, Chg 9 6/23/23 The applicant’s description and justification of the SPI display should address the following: (1) The accuracy of the engine p ower and other engine indications to safely conduct Category B and, when applicable, Category A operations (power setting, NR control, piloting a given rating with precision).

(2) The flight crew must be able to assess how close the indicated parameter is rela tive to the limit, including the transitions from one limiting factor to another (engine manifold pressure to engine RPM or engine torque to gas temperature indicator to engine compressor speed), and the ability to quickly and accurately compare engine - to - engine data when engine limiting factors are different (Engine 1 Torque, Engine 2 Gas Temperature Indicator). See § § 29.1321 and 29.1549.

(3) Engine design features or characteristics that forewarn the crew prior to the parameter reaching the operating limit (the redline) should be evaluated.

(4) The SPI display must provide a timing countdown when operating with time limited power settings and as required by § 29.1305(a)(25).

(5) The minimum and maximum operating limits must be marked as required by § 29.1549.

(6) The S PI display should allow the pilot to identify indicator failures, engine failures, partial power loss conditions, and adverse trends, to meet the requirements in § § 29.771(a), 29.1309, 29.1321, and 29.1549. The SPI design should also address all normal and abnormal operations including start/shutdown and OEI conditions per § 29.771(a).

(7) Since the SPI is a system comprised of instruments requirement by subchapter C of Title 14, its hazards are among those that must assessed under the requirements of § 29.1309. When conducting this assessment, to meet the requirements in § 29.1301, the applicant must take into account that the crew will operate the rotorcraft, monitoring the engine status and health, without continuously accessing the powerplant display page. In addition, the assessment should assume the SPI will remain the only engine instrument available in case of loss of one display.

For guidance on § 29.1305(b)(4) OEI training mode, see Miscellaneous Guidance 22 (MG 22).

Page F - 12.2

AC 29-2C, Chg 4 5/1/2014 AC 29.1307. § 29.1307 (Amendment 29-12) MISCELLANEOUS EQUIPMENT.

a. Explanation. This rule provides a listing of several items of required miscellaneous equipment. Each item is self-explanatory, except for the one requiring a master switch arrangement for electrical circuits other than ignition. The purpose of a master switch arrangement is to allow rapid removal of all bus loads from sources of electrical power in an emergency situation. Requirements for radio communications are discussed more in AC 29.1431.

b. Procedures. When reviewing possible solutions to the master switch arrangement requirement, the following considerations should be included.

(1) System separation . Since wiring from each electrical system will be brought in close proximity to each other, extra care should be taken to maintain some separation. As examples, common connectors, common grounds, and common wire routing should be avoided.

Installation of switches (2) . The single switch should be avoided since it introduces the possibility of a single failure turning off the entire electrical system. One solution that is commonly used provides a close grouping of the switches such that the pilot can easily reach all switches and turn them all off with one action. This solution requires a cockpit evaluation to ensure the installation will be suitable for different hand sizes. Another solution involves a gang bar that can be moved with a single motion to turn off all sources. This solution has been found to be acceptable in several instances.

Other solutions should be evaluated on their own merits, and the primary emphasis should be on maintaining some minimum system separation and conducting a cockpit evaluation by flight test personnel.

Page F - 12.3 5/1/2014 AC 29-2C, Chg 4 AC 29.1309. § 29.1309 (Amendment 29-40) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.

a. Explanation.

(1) Applicability. Section 29.1309 of the Title 14 CFR regulations is intended as a general requirement that is applicable to any equipment or system as installed, in addition to specific systems requirements, except as indicated below.

(i) General . If a specific part 29 requirement exists which predefines systems safety aspects (e.g., redundancy level or criticality) for a specific type of equipment, system, or installation, then the specific part 29 requirement will take precedence (see section 29.1301b.(2) of this AC). This precedence does not preclude accomplishment of a system safety assessment, if necessary. For example, § 29.695 is a rule that predefines a required level of redundancy and an implied system reliability. However, a system safety assessment approach may still be required to show that the implied system reliability is met and to address assessment of the failure modes.

(ii) . Section 29.1309 does not apply to Section 29, Subparts B, C, and D Subparts B, C, and D for aspects such as the performance, flight characteristics, structural loads , and structural strength requirements. However, it does apply to any equipment and system on which compliance with the requirements of Subparts B, C, and D is based (e.g., Health Usage Monitoring System certified for maintenance credit and Stability Augmentation System).

(iii) Section 29, Subpart E .

(A) Section 29.1309 does not apply to the uninstalled type-certificated engine. However, it does apply to the equipment and systems associated with the engine installation (e.g., electrical power generation, engine displays, transducers, etc.)

on the rotorcraft (reference § 29.901).

(B) Section 29.1309 does not apply to the Rotor Drive Systems (reference § 29.917(b)).

(iv) Section 29, Subpart F .

(A) Section 29.1309 does not apply to stowed safety equipment such as life rafts, life preservers, and emergency floatation equipment. It also does not apply to safety belts, rotorcraft seats, and hand held fire extinguishers. However, it does apply to hazards to the rotorcraft, its occupants, and flight crew introduced by the installation or presence of this type of equipment or system (e.g., Electromagnetic-Interference considerations, fire hazards, and inadvertent deployment of emergency floatation equipment) approved as part of the type design.

Page F - 13 AC 29-2C, Chg 4 5/1/2014 (B) Section 29.1309 does not apply to the functional aspects of aircraft non-safety related equipment such as entertainment systems, hoists, Forward Looking Infrared systems (FLIR), or emergency medical equipment such as defibrillators, etc.

However, it does apply to hazards to the rotorcraft, its occupants, and flight crew introduced by the installation or presence of this type of equipment or system (e.g., Electromagnetic-Interference considerations, fire hazards, and failure of the electrical system fault protection scheme) approved as part of the type design.

(C) Section 29.1309 does not apply to the lighting characteristics (e.g., light intensity, color, and coverage) of the position lights, anti-collision lights, and riding lights. However, it does apply to hazards to the rotorcraft, its occupants, and flight crew introduced by the installation or presence of this type of equipment or system (e.g., Electromagnetic-Interference considerations, fire hazards, and pilot visibility impairment due to glare) approved as part of the type design.

(2) Background .

(i) This guidance is not intended to change the existing reliability requirements of the current §§ 29.1309(b) and (c). Neither is this guidance intended to impose an overly rigorous safety assessment process for those installations involving equipment or systems that are not considered to be complex or do not perform functions whose failure or malfunction result in higher level failure condition categories.

(ii) However, this section does provide guidance that describes a method of showing compliance that utilizes the latest system safety assessment techniques available in the aerospace industry. The safety assessment process described in this guidance can be easily adopted and applied with various degrees to equipment or systems, and installations ranging from simple systems that perform functions whose failure or malfunction have no safety impact, to the most critical and most complex systems.

b. Procedures .

(1) Definitions of Probability Classifications .

-5 (i) FREQUENT . (This is the lower part of the range 10 or greater previously applied to the term “PROBABLE”.) Frequent events may be expected to occur often during the operational life of each rotorcraft that are based on a probability on the order -3 of 10 or greater.

-5 (ii) REASONABLY PROBABLE . (This is the upper part of the range 10 or greater previously applied to the term “PROBABLE”.) Reasonably probable even ts may be expected to occur several times during the operational life of each rotorcraft that are -3 -5 to 10 . based on a probability on the order of between 10 Page F - 14 5/1/2014 AC 29-2C, Chg 4 (iii) REMOTE . (The term “REMOTE” is not related to the structural use of the -9 -5 term.) (This is the lower part of the range 10 to 10 previously applied to the term “IMPROBABLE”.) Remote events are expected to occur a few times during the total operational life of a random single rotorcraft of a particular type, but may occur several times during the total operational life of a number of rotorcraft of a particular type, that -5 -7 are based on a probability on the order of between 10 to 10 .

(iv) EXTREMELY REMOTE . (The term “REMOTE” is not related to the -9 -5 structural use of the term.) (This is the upper part of the range 10 to 10 previously applied to the term “IMPROBABLE”.) Extremely remote events are not expected to occur during the total operational life of a random single rotorcraft of a particular type, but may occur a few times during the total operational life of all rotorcraft of a particular -7 -9 to 10 . type, that are based on a probability on the order of between 10 (v) EXTREMELY IMPROBABLE . (Remains the same for both the three or the five classifications.) Extremely improbable events are so unlikely that they need not be considered to ever occur, unless engineering judgment would require their -9 or less is assigned to this classification. consideration. A probability on the order of 10 Note: The five probability terms defined in paragraph b.(1) above are intended to relate to an identified failure condition resulting from or contributed to by the improper operation or loss of a function or functions. These terms do not define the reliability of specific components or systems.

(2) Definitions of Failure Condition Classifications . Failure Conditions may be classified according to the severity of their effects into one of the following categories (reference Figure AC 29.1309-2).

NO EFFECT (i) . Failure Conditions that would have no effect on safety; for example, Failure Conditions that would not affect the operational capability of the rotorcraft or increase crew workload, however, could result in an inconvenience to the occupants, excluding the flight crew .

MINOR (ii) . Failure conditions which would not significantly reduce rotorcraft safety, and which would involve crew actions that are well within their capabilities.

Minor failure conditions may include, for example, a slight reduction in safety margins or functional capabilities, a slight increase in crew workload, such as, routine flight plan changes, or some physical discomfort to occupants.

(iii) MAJOR . Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be, for example, a significant reduction in safety margins or functional capabilities, a significant increase in crew work load or in conditions impairing crew efficiency, physical distress to occupants, possibly including injuries, or physical discomfort to the flight crew.

Page F - 15 AC 29-2C, Chg 4 5/1/2014 (iv) HAZARDOUS or SEVERE-MAJOR . Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be: (A) A large reduction in safety margins or functional capabilities; (B) Physical distress or excessive workload such that the flight crew’s ability is impaired to where they could not be relied on to perform their tasks accurately or completely; or, (C) Possible serious or fatal injury to a passenger or a cabin crew member, excluding the flight crew.

Note: Hazardous or Severe-Major failure conditions can include events that are manageable by the crew by use of proper procedures which, if not implemented correctly or in a timely manner, may result in a Catastrophic Event.

(v) CATASTROPHIC . Failure Conditions which would result in multiple fatalities to occupants, fatalities or incapacitation to the flight crew, or result in loss of rotorcraft.

(3) Safety Objective .

(i) The objective of § 29.1309 is to ensure an acceptable safety level for equipment and systems as installed on the rotorcraft. A logical and acceptable inverse relationship must exist between the Average Probability per Flight Hour and the severity of Failure Condition effects, as shown in Figure AC 29.1309-1, such that: (A) Failure Conditions with No Effect on safety may be more frequent than Reasonably Probable.

(B) Minor Failure Conditions may be Reasonably Probable.

(C) Major Failure Conditions must be no more frequent than Remote.

(D) Hazardous or Severe-Major Failure Conditions must be no more frequent than Extremely Remote.

(E) Catastrophic Failure Conditions must be Extremely Improbable.

(ii) The safety objectives associated with Failure Conditions are described in Figure AC 29.1309-2, and appear in Figure AC 29.1309-1, below.

Page F - 16 5/1/2014 AC 29-2C, Chg 4 Probable Unacceptable Probability Remote of Failure Extremely Condition Remote Acceptable Extremely Improbable Minor Major Hazardous Catastrophic Severity of Failure Condition Elfects FIGURE AC 29.1309-1 Relationship between Probability and Severity of Failure Condition Effects (iii) The safety objectives associated with Catastrophic Failure Conditions, may be satisfied by demonstrating that: (A) No single failure will result in a Catastrophic Failure Condition; and (B) Each Catastrophic Failure Condition is Extremely Improbable.

(iv) Failure conditions may be classified according to the severity of their effects to the rotorcraft, occupants, or flight crew as identified in the following Figure AC 29.1309-2.

Page F - 17 AC 29-2C, Chg 4 5/1/2014 Table for Failure Condition Categories and Probability Definitions Effect on No effect on Slight Significant Large reduction in Loss of rotorcraft operational reduction in reduction in functional rotorcraft capabilities or functional functional capabilities or safety capabilities capabilities safety margins (Note 4) or safety or safety margins margin Effect on Inconvenience Physical Physical Serious or fatal Multiple occupants discomfort distress, injury to a Fatalities excluding flight possibly passenger or a crew including cabin crew member (NOTE 2) injuries Effect on flight No effect on Slight Physical Physical distress or Fatalities or crew flight crew increase in discomfort or excessive workload incapacitation work load a significant impairs ability to which increase in perform tasks involve crew workload or accurately or actions well in conditions completely within crew impairing capabilities crew such as efficiency routine flight plan changes DO-178C E D C B A Software Level (Note 3) Failure Hazardous or Condition No Effect Minor Major Severe-Major Catastrophic Category Qualitative Reasonably Extremely Remote Extremely Probability Frequent Probable Remote Improbable -3 -5 -7 -9 Quantitative No probability ≤ 10 ≤ 10 ≤ 10 ≤ 10 (Note 1) Probability : requirement Note 1: A numerical probability range is provided here as reference. The applicant is not required to perform a quantitative analysis, or substantiate by such an analysis, that this numerical criterion has been met for Minor Failure Conditions.

Note 2: This is true if it can be shown that the given failure condition can be contained to a fatal injury of one occupant only.

Note 3: This is not intended to imply that the identified software levels are assigned a probability value, but instead, shows a correlation to the Failure Condition Category.

Note 4: Hazardous or Severe-Major failure conditions can include events that are manageable by the crew by use of proper procedures which, if not implemented correctly or in a timely manner, may result in a Catastrophic event.

FIGURE AC 29.1309-2 Failure Condition Categories and Probability Definitions (v) The following Figure AC 29.1309-3 provides conversion between the current application of the five failure condition categories as defined in Figure AC 29.1309-2 and the three failure condition categories contained in the present rules.

Page F - 18 5/1/2014 AC 29-2C, Chg 4 Present Rule Extremely Qualitative -------- Probable ------- - ----------Improbable-------- - Improbable Probability Classification - 5 - 5 - 9 Quantitative >10 < 10 < 10 Probability Classification Present Rule Non - essential Essential Critical Failure Condition Category – AC 29 2C & DO-178A Current Hazardous or Application of No Minor Major Severe - Catastrophic Failure Condition Effect Major Category FIGURE AC 29.1309-3: Failure Condition Conversion (4) Safety Assessment Process Overview .

(i) When showing compliance with § 29.1309, the considerations covered in this guidance should be addressed in a methodical and systematic manner, which ensures that the process and its findings are visible, and readily assimilated. This section of the AC (Safety Assessment Process Overview) is provided primarily for the use of applicants who are not familiar with the various methods and procedures generally used in the industry to conduct safety assessments. This guide and Figures AC 29.1309-4 and AC 29.1309-5 are not intended to be certification checklists, and they do not include all the information provided in this guidance. This guide is one method but not the only method for showing compliance. The safety assessment process is a structured method of general applicability for showing compliance with the regulation.

Other methodologies may be used to show compliance, but the safety assessment process is the only structured process that is defined and these defined objectives would be the logical criteria applied to any other means of showing compliance. More detailed guidance can be found in SAE Document ARP 4761 on how to perform a safety assessment. SAE Document ARP 4754A, invoked by AC 20-174, includes additional guidance on how the safety assessment process relates to the system development process.

(ii) The safety assessment process contains several parts that may be necessary as a whole or part depending on the criticality or complexity of the system under consideration. The rigor of assessment and analysis performed is also dependent on the system criticality or complexity. At the extremes, some systems may be simple enough such that the entire safety assessment can be performed by observation and compliance shown by a simple statement. Complex and higher criticality systems may require application of all the safety assessment elements to show compliance. Many states of varying complexity exist between these extremes that Page F - 19 AC 29-2C, Chg 4 5/1/2014 may use less than the entire safety assessment process, but require more than a simple statement to show compliance.

(iii) Elements of the safety assessment process are as follows: • Functional Hazard Assessment (FHA) • Preliminary System Safety Assessment (PSSA) •• Fault Tree Analysis (FTA) • System Safety Assessment (SSA) •• Failure Mode and Effect Analysis (FMEA) • Common Cause Analysis (CCA) •• Zonal Safety Analysis (ZSA) •• Particular Risk Assessment (PRA) •• Common Cause Analysis (CMA) (A) Define the system and its interfaces, and identify the functions that the system is to perform. Determine whether or not the system is complex, similar to systems used on other aircraft, or conventional. Where multiple systems and functions are to be evaluated, consider the relationships between multiple safety assessments.

(B) Identify and classify Failure Conditions. All relevant applicant- engineering organizations, such as systems, structures, propulsion, and flight test, should be involved in this process. This identification and classification may be done by conducting a Functional Hazard Assessment (FHA) which is a top down assessment.

The FHA starts at the rotorcraft operational level to identify possible hazards to the rotorcraft, its occupants, and flight crew. When classifying a function, consider for example, the loss of function, erroneous or misleading, or degradation of the function as possible hazards. The hazards are classified using the five failure condition categories found in Figure AC 29.1309-2. The identified failure conditions must be evaluated with respect to the intended operations and combinations of operations for which certification of the rotorcraft is sought (e.g., VFR, IFR, CAT A, IFR plus CAT A, etc.). The product of the FHA is a determination of the failure or malfunction condition category for each system function and identification of the source of the function. Lower level FHAs that focus on the system level hazards reveal detailed information about the specific part of a system that provides the function under evaluation. This more detailed information may narrow the system part to be considered for compliance with enhanced integrity or for additional analysis. However, it must be recognized that there is a limit to the Page F - 20 5/1/2014 AC 29-2C, Chg 4 number of levels that can be relevant in this process. Too many levels can result in the assessment becoming design-dependent rather than the FHA driving the design, as intended.

(C) The FHA is usually based on one of the following methods, as appropriate: ( ) If the system is not complex and its relevant attributes are similar to 1 those of systems us ed on other aircraft, the identification and classification may be derived from design, installation, and service experience appraisals that can be shown to be sufficiently similar.

2 ( ) If the system is complex, it is necessary to systematically postulate the effects on the safety of the aircraft and its occupants resulting from any possible failures, considered both individually and in combination with other failures or events.

Note: A system is considered to be complex when its operation, failure modes, or failure effects are difficult to comprehend without the aid of analytical methods.

(D) Choose the means to be used to determine compliance with Section 29.1309. The depth and scope of the analysis depends on the types of functions performed by the system, the severity of the associated Failure Conditions, and whether or not the system is complex (see Figure AC 29.1309-5). Once the failure conditions are determined to be Major Failure Conditions, experienced engineering and operational judgment, design and installation appraisals, and similarity consideration for previously installed systems may be acceptable. This may be sufficient in itself or in conjunction with qualitative analyses or selectively used quantitative analyses. For Catastrophic and Hazardous or Severe-Major Failure Conditions, a thorough safety assessment is necessary. The applicant should obtain early concurrence of the certification authority on the choice of an acceptable means of compliance.

(E) If an analysis is used as an acceptable means of showing compliance, it should typically include the following information: ( 1 ) A statement of the functions, boundaries, and interfaces of the system.

2 ( ) A list of the parts and equipment of which the system is comprised, including their performance specifications or design standards and development assurance levels if applicable. This list may reference other documents (e.g., Technical Standard Orders (TSOs), manufacturer's or military specifications, etc.).

3 ( ) The conclusions, including a statement of the Failure Conditions and their classifications and probabilities (expressed qualitatively or quantitatively, as appropriate) that show compliance with the requirements of § 29.1309.

Page F - 21 AC 29-2C, Chg 4 5/1/2014 ( 4 ) A description that establishes correctness and completeness and

-

traces the work leading to the conclusions. This description should include the basis for the classification of each Failure Condition (e.g., analysis or ground, flight, or simulator tests). It should also include a description of precautions taken against common-cause failures, provide any data such as component failure rates and their sources and applicability, support any assumptions made, and identify any required flight crew or ground crew actions, including any Candidate Certification Maintenance Requirements (CCMRs) which are also referred to as Certification Check Requirements (CCRs).

Certification Maintenance Requirements (CMRs) are not the preferred method for showing compliance to the rules; therefore, they must be reviewed by the FAA/AUTHORITY for approval.

(F) Certification Maintenance Requirements (CMRs) may be needed to help show compliance with § 29.1309 for significant latent failures. Rational methods, which usually involve quantitative analyses or relevant service experience data, should be used to determine CMR intervals. These intervals should have reasonable tolerances so that CMRs can be performed concurrently with other maintenance, inspection, or check procedures not required by design for compliance with § 29.1309.

Such tolerances are acceptable because the uncertainty described in the “Note” below is accounted for as discussed in this section. If CMRs are used, they and their intervals and tolerances, and any post-certification changes, or procedures provided in the type design for a rotorcraft owner or operator to make such changes, should be approved by the FAA/AUTHORITY having cognizance over the type design that relates to the system and its installation.

Note: It is recognized that, for various reasons, component failure rate data are not precise enough to enable accurate estimates of the probabilities of Failure Conditions. This results in some degree of uncertainty, as indicated by the expression “on the order of” in the descriptions of the quantitative probability terms that are provided in paragraph b.(1) above. When calculating the probability of each Failure Condition, this uncertainty should be accounted for in a way that does not compromise safety.

(G) Assess the analyses and conclusions of multiple safety assessments to ensure compliance with the requirements for all rotorcraft level Failure Conditions.

(H) Prepare compliance statements, maintenance requirements, and flight manual requirements.

(I) When assessing the acceptability of a failure condition using a quantitative analysis, the given numerical range should normally be interpreted to be the allowable risk for an hour of flight time based on a flight of mean duration for the rotorcraft type. However, when assessing a function which is used only at a specific time during a flight, the probability of the failure condition should be calculated for the specific period and expressed as the risk for the flight condition, takeoff, landing, etc., as appropriate. This is only true for those systems that cannot fail undetected when they Page F - 22 5/1/2014 AC 29-2C, Chg 4 are not being used. The probability of failure for those systems must be calculated from that value based on the extended period of not only the active operation, but also the exposure time for the latent failure.

Note: If a quantitative analysis is used to help show compliance with the regulations for equipment or systems, which are installed and only required for a specific operating condition for which the rotorcraft is thereby approved, credit is usually not allowed due to the fact that the operating condition does not always exist. If an applicant does request that such credit be allowed, they must obtain the approval of the FAA/AUTHORITY.

Page F - 23 AC 29-2C, Chg 4 5/1/2014 Certification Requirements Maintenance Requirements Flight Manual Requirements Rotorcraft Level Functional Assessment of Rotorcraft Hazard Assessment Level System Safety Assessment System Safety System Level FHA’s Assessments Analyses FIGURE AC 29.1309-4: Safety Assessment Process Overview Page F - 24 5/1/2014 AC 29-2C, Chg 4 Note: FHA may be based on a design and installation appraisal for these systems.

Conduct Functional Hazard Assessment Is There a No Safety Effect?

Yes Verify by Design Yes Is the and Installation Failure Condition Appraisal.

Minor ?

No Develop Similarity Is the System Yes Analysis and Obtain and Installation Approval from Similar to a Previo us FAA/Authority Design ?

No * Conduct Qualitative or Quantitative Analysis, or both.

* Note: Catastrophic and Hazardous/ Severe-major failure conditions will likely require both qualitative and quantitative analysis, depending on system complexity.

FIGURE AC 29.1309-5: Depth of Analysis Flowchart Page F - 25 AC 29-2C, Chg 4 5/1/2014 (5) Assessment Methods . Various methods for assessing the causes, severity, and probability of Failure Conditions are available to support experienced engineering and operational judgment. Some of these methods are structured. The various types of analysis are based on either inductive or deductive approaches. Probability assessments may be qualitative or quantitative. Descriptions of some types of analysis are provided below and in Document ARP 4761.

Design Appraisal (i) . This is a qualitative appraisal of the integrity and safety of the system design.

Installation Appraisal (ii) . This is a qualitative appraisal of the integrity and safety of the installation. Any deviations from normal, industry-accepted installation practices, such as clearances or tolerances, should be evaluated, especially when appraising modifications made after entry into service. Hazards introduced by the installation of the system should also be evaluated.

(iii) Failure Modes and Effects Analysis (FMEA) .

(A) This is a structured, inductive, bottom-up analysis, which is used to evaluate the effects on the system and the rotorcraft of each possible element or component failure. When properly formatted, it will aid in identifying latent failures and the possible causes of each failure mode. Document ARP 4761 provides methodology and detailed guidelines that may be used to perform this type of analysis. An FMEA could be a piece part FMEA or a functional FMEA. For modern microcircuit based line replaceable units (LRUs) and systems, an exhaustive piece part FMEA is not practically feasible with the present state of the art. In that context, a FMEA may be more functional than piece part oriented. A functional oriented FMEA can lead to uncertainties in the qualitative and quantitative aspects, which can be compensated for by more conservative assessment, such as: ( 1 ) assuming all failure modes result in the Failure Conditions of interest, ( 2) careful choice of system architecture, or ( 3) taking into account the experience lessons learned on the use of similar technology.

(B) Specific reliability numbers are not shown in § 29.1309. The necessary degree of reliability is a function of the criticality of the system under consideration. Acceptable sources of component failure rates are (1) military aircraft component failure or fault reports or handbooks, such as MIL-HDBK-217C; (2) operator or manufacturer component malfunction or defect records, such as airline component defect records on sufficiently similar component designs; and (3) laboratory life tests.

Page F - 26 5/1/2014 AC 29-2C, Chg 4 (iv) Fault Tree or Dependence Diagram Analysis . Structured, deductive, top- down analyses which are used to identify the conditions, failures, and events that would cause each defined Failure Condition. They are graphical methods of identifying the logical relationship between each particular Failure Condition and the primary element or component failures, other events, or combinations thereof that can cause it. A failure modes and effects analysis may be used as the source document for those primary failures or other events.

Markov Analysis (v) . A Markov model (chain) represents various system states and the relationships among them. The states can be either operational or non - operational. The transitions from one state to another are a function of the failure and repair rates. Markov analysis can be used as a replacement for fault tree and dependence diagram analysis, but it often leads to more complex representation, especially when the system has many states. It is recommended that Markov analysis be used when fault tree or dependence diagrams are not easily usable, namely to take into account complex transition states of systems which are difficult to represent and handle with classical fault tree or dependence diagram analysis.

Common Cause Analysis (CCA) (vi) . The acceptance of adequate probability of Failure Conditions is often derived from the assessment of multiple systems based on the assumption that failures are independent. Therefore, it is necessary to recognize that such independence may not exist in the practical sense and specific studies are necessary to ensure that independence can either be assured or deemed acceptable.

CCA identifies common causes for faults or malfunctions. Once identified, the source of the common cause is either eliminated from the design, the common cause or the system design is made to be tolerant of the related fault or malfunction, or the common cause or system design combination is found to be acceptable due to the failure or malfunction condition category it represents. The potential for failures or malfunctions due to common causes is inherent in designs that provide multiple functions reliant on common hardware or common software. This is also true for systems that produce related functions and share a common installation area. The installation area may represent a threat from several sources such as Electro-Magnetic Interference (EMI), mechanical hazards, and environmental influences. The Common Cause Analysis is sub-divided into three areas of study: (A) Zonal Safety Analysis (ZSA) . This analysis has the objective of ensuring that the equipment installations within each zone of the rotorcraft are at an adequate safety standard with respect to design and installation standards, interference between systems, and maintenance errors. The ZSA examines the physical zone of the rotorcraft, in which the system under consideration is installed, to ensure that the surrounding equipment or appliance installations do not compromise the system independence requirements. Mechanical failures that might generate fragments that could damage the system under evaluation are an example of the type of event that would be considered in a zonal analysis.

Page F - 27 AC 29-2C, Chg 4 5/1/2014 (B) Particular Risk Analysis . Particular risks are defined as those events or influences that are outside the systems concerned. Examples are fire, leaking fluids, bird strike, tire burst, high intensity radiated fields exposure, lightning, uncontained failure of high energy rotating machines, etc. Each risk should be the subject of a specific study to examine and document the simultaneous or cascading effects or influences that may violate independence.

(C) Common Mode Analysis (CMA) . This analysis is performed to confirm the assumed independence of the events that were considered in combination for a given Failure Condition. The effects of specification, design, implementation, installation, maintenance, and manufacturing errors, environmental factors other than those already considered in the particular risk analysis, and failures of system components should be considered. The CMA considers many aspects of design; one is design dissimilarity for both hardware and software. Without dissimilarity for redundant applications, there is a possibility that a hardware or software failure or malfunctions could occur in the same flight for all redundant subparts of a system with a Catastrophic or Hazardous or Severe-Major Failure Condition Category. Another major contribution from the CMA is the determination of which failure or malfunction combinations inputs to the Fault Tree Analysis (FTA) must be independent for events that are Catastrophic or Hazardous or Severe-Major. This analysis is iterative in nature as it will be employed early in design to identify possible common modes for failures or malfunctions and then is used after the design is complete to determine if the FTA goals have been met.

(6) Documentation . All laboratory, ground and flight tests, and failure analyses, must be documented in sufficient detail to show compliance with § 29.1309 and be included in the type design file. Section 21.31(a) provides the regulatory basis for requiring this documentation. If the applicant elects to use a numerical reliability and probability analysis, it must also be documented in sufficient detail.

Software (7) . RTCA Document DO-178C, “ Software Considerations in Airborne Systems and Equipment Certification ,” dated December 13, 2011, is the latest standard and is recommended to be used for qualification and subsequent approval of airborne software. See AC 20-115C for guidance on using DO-178C and earlier standards.

Airborne Electronic Hardware (AEH). For airborne complex custom micro- (8) coded electronic components assessed as level A, B, or C in the safety assessment, RTCA Document DO-254, “ Design Assurance Guidance for Airborne Electronic Hardware ,” dated April 19, 2000, provides certification guidance for qualification and subsequent approval of these components. Per DO-254, the component will need to be identified as either “simple” or “complex” and tagged to one of the five failure condition classifications. Based on these identifiers, DO-254 provides the guidance to show compliance to 29.1309. Although RTCA DO-254 applies specifically to complex custom micro-coded components, applicants are highly encouraged to apply DO-254 to other hardware components up through to the LRU and systems level.

Page F - 28 5/1/2014 AC 29-2C, Chg 4 (9) Environmental Qualification.

(i) Laboratory Tests .

(A) Environmental Standards. In order to assure that the components and systems under consideration will function properly when exposed to adverse environments, they should be tested in the laboratory under a simulated adverse environment. If a TSO exists and it is appropriate in environmental range and performance for an equipment installation, it is preferable the equipment be TSO approved. If there is no applicable TSO or an existing TSO does not provide for a sufficiently adverse environment, the latest revision of the Radio Technical Commission for Aeronautics (RTCA) document DO-160 is an acceptable environmental standard for laboratory qualification of aircraft equipment.

(B) Adverse environmental variables for all types of required and critical equipment include, but are not necessarily limited to temperature, humidity, vibration, shock, altitude, overpressure, and power source transients.

(C) For electrical and electronic equipment, adverse environmental variables include all of (b) above plus overvoltage and undervoltage. Electronic equipment should also be tested for electromagnetic interference (EMI). These tests should include both emission and susceptibility evaluations of both conducted and radiated EMI.

(D) Explosion Tests. Those items of electrical and electronic equipment that are to be located in areas subject to flammable fluids and vapors, as a result of any single probable malfunction or failure, including failure of couplings or lines should be tested as an ignition source. These tests consist of normal operation of the equipment in a physically contained explosive atmosphere. The explosion test procedure in the latest revision of DO-160 will satisfy this requirement. Section 29.863 of this AC provides further guidance on safety from explosion. If another standard is used that is at least as good as the latest revision of DO-160, it may also be accepted to satisfy this requirement.

(ii) Installed Environmental Tests . After the environmental ratings of the components and systems have been established, it should be assured that as installed, these ratings will not be exceeded. Normally, installed equipment need not be instrumented and tested in flight nor is it necessary to instrument the compartment or rack where the equipment is installed. Satisfactory environment and equipment compatibility are assured by selection of the proper environmental category of laboratory tests. The category is determined by the type of aircraft (reciprocating or turbine) and flight envelope (altitude and temperature). Exceptions to normal installations are (a) Alternator or generator cooling, where radiated and conducted heat is almost always uncertain, also cooling air temperatures and flow rates are uncertain; (b) Where flight tests reveal excessive instrument panel vibration. In this case, the panel should be instrumented, tested, and, if necessary, design improvements made; Page F - 29 AC 29-2C, Chg 4 5/1/2014 and (c) Any other cases where good engineering judgment and application of sound engineering principles indicate a high likelihood that the installed environment is more severe than the equipment is capable of operating within.

(A) Temperature Tests.

( ) Temperature tests may be accomplished by instrumenting the 1 installed equipmen t environment with a recorder that provides a permanent record of time, altitude, and temperature. The pertinent temperature should be recorded as the rotorcraft is operated throughout its altitude range, including ground operation. The maximum and minimum temperatures recorded should be corrected degree for degree to assure the equipment under test remains within its temperature rating while the rotorcraft operates throughout its approved ambient temperature envelope. (For generator or alternator cooling test procedures, refer to section 29.1351 of this AC.)

Section 29.1043(b) requires the maximum approved operating OAT to be at least 100° F for powerplant-mounted accessories such as starter generators, vacuum pumps, etc. Due to the impracticality of the 100° F hot day temperature limit, rotorcraft systems mounted on the powerplant are normally evaluated for at least 115° F hot day sea level conditions with corresponding 3.6° F/1,000-foot correction. The maximum hot day OAT at sea level must be specified in the rotorcraft flight manual. Section 29.1043(b) is the regulatory basis for the lapse rate of 3.6° F/1,000 feet. This lapse rate should be applied regardless of the hot day sea level temperature the applicant chooses to certify for operation.

( 2 ) The § 29.1043(b) maximum ambient temperature definition should not be confused with operating temperatures in closed areas. Closed equipment rack areas can easily reach temperatures of 140° F when sitting on the ramp in the southern United States in midsummer. Normally, proper selection of the altitude temperature category in the latest revision of DO-160 will assure compliance.

3 ( ) In some cases, the equipment manufacturer furnishes temperature limits for internal critical parts. For example, brushes, bearings, or field windings on DC generators. In these cases it is better to record the critical component temperature rather than equipment or equipment environment temperature.

4 ( ) The following will illustrate an acceptable high temperature evaluation method: T = Maximum outside air temperature at which temperature tests OAT MAX are conducted.

T = Maximum temperature to which the installed equipment has MAX been tested in the laboratory.

T = Maximum installed equipment temperature recorded during TEST MAX tests.

Page F - 30 5/1/2014 AC 29-2C, Chg 4 T = The high reference outside air temperature. It varies with ORH altitude starting at the highest sea level temperature at which rotorcraft operation is to be approved and decreases at 3.6° F/1,000 foot altitude. It can be no less than 100° F (reference § 29.1043(b)); however, it can be as high as the applicant wants.

T = Temperature margin between the maximum equipment H MAR temperature substantiated in the laboratory and the maximum installed equipment temperature when the rotorcraft is operating in the highest available OAT and approximately corrected at the altitude under consideration. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.

T = T - (T + (T - T )) H MAR MAX TEST MAX ORH OAT MAX Example #1 : Assume the applicant is seeking approval for rotorcraft operation at the lowest acceptable OAT, at sea level, of 100° F and T for Generator MAX Brush = 295° F at maximum load current throughout the altitude range. In-flight test data are: Measured Altitude (ft, MSL) Cylinder Temp (T ) OAT = T TEST MIN OAT MAX sea level 275°F 90°F 5,000 270°F 80°F 10,000 285°F 60°F 15,000 294°F 42°F 20,000 290°F 20°F First, T must be calculated for each altitude test point.

ORH @ sea level, T = 100° F ORH @ 5,000 ft., T = 100° F -5,000 ft., x 3.6°/1,000 ft. = 82° F ORH @ 10,000 ft., T = 100° F - 10,000 ft. x 3.6°/1,000 ft. = 64° F ORH @ 15,000 ft., T = 100° F - 15,000 ft. x 3.6°/1,000 ft. = 46° F ORH @ 20,000 ft., T = 100° F - 20,000 ft. x 3.6°/1,000 ft. = 28° F ORH Then at sea level: T = 295 - (275 + (100 - 90)) = 10° F H MAR At 5,000 feet: Page F - 31 AC 29-2C, Chg 4 5/1/2014 T = 295 - (270 + (82 - 80)) = 23° F H MAR At 10,000 feet: T = 295 - (285 +(64 - 60)) = 6° F H MAR At 15,000 feet: T = 295 - (294 + (46 - 42)) = -3° F H MAR At 20,000 feet: T = 295 - (290 + (28 -20)) = -3° F H MAR Since T comes out negative at the 15,000- and 20,000-foot points, the generator H MAR fails. It will be necessary for the applicant to reduce the maximum load current, improve cooling, or otherwise change the design to assure the generator is operating within its approved temperature limit of 295° F.

( 5 ) In most cases, the equipment is laboratory tested to minimum

-

temperatures as severe as that of the rotorcraft’s maximum certified altitude on a minimum temperature day. Therefore, unless equipment minimum temperature is affected by refrigeration or other temperature reducing environments, actual installed instrumented minimum temperature tests are unnecessary. If low temperature evaluation is necessary for the installed equipment, the following is an acceptable method: T = Minimum outside air temperature at which temperature tests are OAT MIN conducted.

T = Minimum temperature to which the installed equipment has been MIN tested in the laboratory.

T = Minimum installed equipment temperature recorded during TEST MIN tests.

T = The low reference outside air temperature. It varies with altitude ORL starting at the lowest sea level temperature at which rotorcraft operation is to be approved and decreases at 3.6° F/1,000-foot altitude.

T = Temperature margin between the minimum equipment 1 MAR temperature substantiated in the laboratory and the minimum installed equipment temperature. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.

Page F - 32 5/1/2014 AC 29-2C, Chg 4 T = -(T - (T + (T - T ))) 1 MAR MIN TEST MIN ORL OAT MIN Note: This equation assumes all temperatures are negative. It is necessary to place a (-) in front of the right side of the equation in order to convert the T value to 1 MAR the conventional positive answer for acceptance and a negative answer for rejection. Temperature in the 0 to 32° F range can be handled by conversion to the centigrade scale.

Example #2 : Assume the applicant is seeking a low temperature operating limit at sea level of -25° F. Assume the hydraulic control cylinder has been substantiated in the laboratory to operate at a cylinder temperature of -40° F. The in-flight test data are: Measured Altitude (ft, MSL) Cylinder Temp (T ) OAT = T TEST MIN OAT MAX sea level 0°F -25°F 5,000 -9°F -45°F 10,000 -21°F -59°F 15,000 -32°F -65°F 20,000 -40°F -69°F (T must be calculated for each altitude test point) orl @ sea level, T = -25° F ORL @ 5,000 ft., T = -25° F -5,000 ft., x 3.6°/1,000 ft. = -43° F ORL @ 10,000 ft., T = -25° F - 10,000 ft. x 3.6°/1,000 ft. = -61° F ORL @ 15,000 ft., T = -25° F - 15,000 ft. x 3.6°/1,000 ft. = -79° F* ORL @ 20,000 ft., T = -25° F - 20,000 ft. x 3.6°/1,000 ft. = -97° F* ORL *According to § 29.1043(b), the lowest temperature to be considered is -69.7° F.

Then at sea level: T = -(-40 -(0 + (-25 -(-25)))) = 40° F 1 MAR At 5,000 feet: T = -(-40 - (-9 + (-43 - (-45)))) = 33° F 1 MAR At 10,000 feet: T = -(-40 - (-21 + (-61 - (-59)))) = 17° F 1 MAR At 15,000 feet: Page F - 33 AC 29-2C, Chg 4 5/1/2014 T = -(-40 - (-32 + (-69.7 - (-65)))) = 3.3° F 1 MAR At 20,000 feet: T = -(-40 - (-40+ (-69.7 - (-69)))) = -0.7° F 1 MAR It can be seen that there is an acceptable margin at all altitudes up to and including 15,000 feet. However, at 20,000 feet, the margin is negative and the system fails.

(B) Vibration tests. Normally, installed vibration tests are not necessary for equipment qualified in accordance with the latest revision of RTCA document DO-160. This paper categorizes vibration tests according to installed rotorcraft equipment location such as fuselage, engine compartment, instrument panel, equipment rack, etc. However, installed equipment vibration tests may be necessary when it appears the equipment location environment may exceed the laboratory-tested equipment vibration limits.

(C) Altitude tests. If the equipment has been laboratory tested to the maximum certified altitude of the rotorcraft, installed altitude tests are unnecessary.

The installed equipment must be either laboratory tested or tested in the rotorcraft to the maximum certified altitude of the rotorcraft.

(iii) Lightning Strike Protection of Full Authority Digital Engine Controls (FADEC) .

(A) Explanation.

( 1 ) The following discussion is written specifically for FADEC with an alternate technology backup fuel control installed on rotorcraft with category A engine isolation. The requirement for increased consideration of lightning strike encounter effects on avionic equipment and systems has been brought about by the increased use of avionics to perform functions, the failure or malfunction of which could result in a hazard to the rotorcraft. The susceptibility of current high technology avionic systems is increased by the use of large scale integration, very large scale integration, and complementary metallic oxide silicon technologies, which exhibit a greatly reduced tolerance to large amplitude, low energy electrical transients as compared to conventional bipolar technology, and the reduced physical protection and electromagnetic shielding afforded aircraft avionic systems by the advanced technology composite airframe materials. Additionally, processor-based systems have the failure phenomenon of digital upset. A digital upset occurs when a system, perturbed by an electrical transient, ceases proper operation in accordance with its embedded software while suffering no apparent component or device damage.

2 ( ) Since elements of electrical and electronic engine subsystems are typically spread throughout much of the rotorcraft, transients caused by lightning are coupled into the subsystem interface cables and may damage the system or cause Page F - 34 5/1/2014 AC 29-2C, Chg 4 upset. Effective lightning protection must be designed and incorporated into these systems. Reliance upon redundancy as a means of protection against lightning effects is generally not adequate because lightning electromagnetic fields and structural IR voltages usually interact (to some extent) with all electrical wiring aboard a rotorcraft.

( 3 ) The testing and analysis outlined in this discussion are methods by which the FAA/AUTHORITY may be assured that when the rotorcraft experiences “the foreseeable operating condition” of a worst-case lightning strike encounter that the electronically controlled engines will continue to “perform their intended function” and therefore be in compliance with § 29.1309(a) as installed.

4 ( ) The definition of what constitutes a full authority engine control is not at this time clearly defined. However, it has been accepted in past certification that any control which relies upon the electronics for the function on which Civil Certification or Military Qualification is based (e.g., rotor speed governing) is a full authority control, regardless of the backup control mode provided. If engine certification or qualification can be achieved without the electronic control, which is subsequently added to achieve improved operational efficiency in the aircraft, the control is “supervisory.” However, if the controls are used in a multiengine rotorcraft, a common failure caused by a lightning strike could result in simultaneous failures, which would cause a reduction in power greater than the loss of one engine. This would also be considered “full authority.” Note: If OEI ratings are approved, cumulative loss of power from all engines should be limited to allow flight manual performance based on OEI ratings.

(B) Procedures. Although not a regulatory requirement, it is recommended that a formal written certification plan be used to assure regulatory compliance. The use of this plan is beneficial to both the applicant and the FAA/AUTHORITY because it identifies and defines an acceptable resolution to the critical issues early in the certification process. These are the usual steps to be followed when utilizing a certification plan: 1 ( ) Prepare a certification plan that describes the analytical procedures and the qualification tests to be utilized to demonstrate protection effectiveness. Test plans should describe the rotorcraft and FADEC system to be utilized, test drawing(s) as required, the method of installation that simulates the production installation, the lightning zone(s) applicable, the lightning simulation method(s), test voltage or current waveforms to be used, diagnostic methods, and the appropriate schedules and location(s) of proposed test(s).

2 ( ) Obtain FAA/AUTHORITY concurrence that the certification plan is adequate.

3 ( ) Obtain FAA/AUTHORITY detail part conformity of the test articles and installation conformity of applicable portions of the test setup.

Page F - 35 AC 29-2C, Chg 4 5/1/2014 ( 4 ) Schedule FAA/AUTHORITY witnessing of the test.

( 5 ) Submit a final test report describing all results and obtain FAA/AUTHORITY approval of the report.

(C) Definition of Environment. The SAE AE4L Committee report dated June 6, 1999, has been and remains acceptable criteria to define the worst-case lightning strike that may be encountered by the rotorcraft in service. An additional explanation of the lightning environment could be found in FAA Report DOT/FAA/CT-89/22, “Aircraft Lightning Protection Handbook.” This handbook will assist aircraft design, manufacturing, and certification organizations in protecting aircraft against the direct and indirect effects of lightning strikes, in compliance with 14 CFR regulations. It presents a comprehensive test criteria to provide the essential information for the in-flight lightning protection of all types of fixed wing, rotary wing, and powered lift aircraft of conventional, composite, and mixed construction and their electrical and fuel systems. The handbook contains chapters on the natural phenomenon of lightning, the interaction between the aircraft and the electrically charged atmosphere, the mechanism of the lightning strike, and the interaction with the airframe, wiring, and fuel system. Further chapters cover details of designing for optimum protection; the physics behind the voltages, currents, and electromagnetic fields developed by the strike; and shielding techniques and damage analysis. The handbook ends with discussion of test and analytical techniques for determining the adequacy of a given protection scheme. On September 7, 2011, FAA AC 20-136B, “Aircraft Electrical and Electronic System Lightning Protection,” was issued, superseding previous versions of that AC. For new designs and applications, this revised definition of the lightning environment should be used.

(D) Certification Plan. The following subjects are not intended to provide a complete list of the items that should be included in the certification plan, but rather highlight some of the areas that should receive consideration. The certification plan should address the total protection that is required to allow the FADEC to continue to operate properly when the rotorcraft experiences a worst-case lightning strike encounter.

( 1 ) Determination of Lightning Strike Attachments. Determine the locations on the rotorcraft where lightning strike attachment is likely to occur and the portions of the airframe through which currents may flow between attachments. The main and tail rotors are recognized as likely attachment points; however, consideration should be given to all possible attachment points. The swept stroke phenomenon may not exist for all lightning strike encounters due to the fact that the rotorcraft may be airborne with little or no airspeed.

( 2 ) Establish the Lightning Environment. Establish the components of the total lightning event to be considered. These are the currents and voltages that are described in the definition of the environment.

Page F - 36 5/1/2014 AC 29-2C, Chg 4 ( 3 ) Full-Level, Complete Vehicle Testing. In accordance with traditional FAA/AUTHORITY policy, the demonstration that the FADEC installed in a complete type design rotorcraft will continue to operate properly when exposed to a worst-case lightning strike is sufficient to demonstrate compliance with § 29.1309(a).

Because of the difficulties involved in utilizing this type of an approach, it is generally not used.

( ) Analytical Processes. A description should be given in the 4 certification plan of the analytical process and certification tests to be utilized to demonstrate protection effectiveness. Typically, the certification plan will include a combination of analysis and tests. (Analytical techniques are most often utilized to predict the levels of lightning-induced transients in interconnecting wiring.) In most cases, successful analyses are based upon well-defined geometrical or electrical parameters such as structural dimensions and materials resistivity. When electrical characteristics of structural materials are not well established, development tests are often utilized to obtain this data which is subsequently utilized in an analysis. In more complex structures or electrical and electronic system installations, it is sometimes difficult or impossible to define the problem in terms that can be analyzed. In these cases, development or verification testing is often relied upon. The purpose of the certification plan is to show how developmental tests, analyses, and verification tests are combined to demonstrate protection design adequacy. In certain cases, previously verified designs can be incorporated and their adequacy confirmed by reference to previous verifications. Such reference should also be incorporated in the certification plan.

(i) The verification testing should be conducted on a system that simulates as closely as possible the installed configuration. As few items as possible of actual hardware should be simulated.

( ii ) Unless the applicant has opted to follow the guidance in AC 20-136B as a means of showing compliance with § 29.1309(h), then the following or other acceptable guidance should be followed. The use of various analytical processes usually requires that the system component tolerance is established. The SAE AE4L Committee Report No. AE4L-81-2 is one example of an acceptable reference to be used for the testing accomplished to determine these tolerances. The testing which is performed to determine the tolerance level of the control computer should include a consideration for the occurrence of a non-recoverable digital upset. One method to provide this consideration is to have the unit powered and the processor operating normally under software control (usually this should be the exact software for which approval is sought) when the test is performed. If strike testing is used, then several shots should be made to develop enough data to provide a reasonable confidence level.

It is an acceptable procedure for the engine manufacturer, while he is obtaining his type certificate, to accomplish this bench testing to determine the level of tolerance of the FADEC system components to lightning encounter indirect effects. This approach has the advantages that the bench tests are not necessarily required to be repeated when the engine is installed in a different airframe. This recommendation is not meant to add Page F - 37 AC 29-2C, Chg 4 5/1/2014 a requirement to the engine manufacturer but to propose a more efficient method of certification. If this tolerance was not determined by the engine manufacturer, the applicant installing the FADEC in a rotorcraft would be expected to furnish this data.

( iii ) As with any analytical method, it is prudent to include a margin of safety to account for the uncertainties involved in the analytical and testing processes. Margins account for uncertainty in the compliance method. As confidence in the compliance method increases, the margin can decrease. A factor of two is an acceptable margin for systems with catastrophic failure conditions, if this margin is verified by aircraft test or by analysis supported by aircraft tests. For other verification methods, the margin should be agreed upon with the cognizant certification office.

( iv ) When an analysis has no associated full-scale vehicle testing to confirm the analysis, the analysis should be very rigorous. Additionally, it should be expected in this situation that this analysis indicates a very large margin of protection.

Many factors must be considered in determining what constitutes an acceptably large margin. The specific additional margin required should be based on an assessment of the inherent uncertainty of a given analysis.

( 5 ) Pass or Fail Criteria. The certification plan should address a pass or fail criteria for the testing and analysis to be performed. The following items should be satisfied to assure acceptable system performance: (i) No immediate crew action must be required.

( ii ) Automatic control of the engine cannot be lost for any appreciable period of time. The engine must not be allowed to be out of control for a period of time that will result in a hazard in a worst-case flight condition. Obviously, any rapid, uncontrolled divergence is not acceptable.

( iii ) No crew action should be required to reset the system. This is not to imply that the system cannot be designed with a manual reset, but the manual reset cannot be used to show compliance to recover from a digital upset.

( iv ) The resumption of engine control after an upset must be reasonably within the range that existed before the upset.

( v ) No critical data can be lost.

vi ( ) After the system recovers, if the performance of the system has been degraded in a noncritical manner which would reduce the capability of the rotorcraft or the ability of the pilot to cope with adverse operating conditions, then the crew must be alerted to this system degradation.

(E) System Installation Considerations. In most cases, the installation of the system components is a constituent part of the lightning protection. This is Page F - 38 5/1/2014 AC 29-2C, Chg 4 particularly true in the use of shielding techniques. If these installation features are required for adequate lightning protection, consideration should be given to ensure that their effectiveness is not derogated in service. ICAs be provided to the parties who service and operate the rotorcraft to allow them to take actions necessary to ensure the continued effectiveness of the system lightning protection.

(iv) Lightning Protection .

(A) Background. During the original design and development of rotorcraft and the development of regulations concerning these aircraft, little attention was given to protection from the meteorological phenomenon of lightning. This was, in part, because the early aircraft were constructed mostly of metal and had little, if any, dependence on advanced technology systems. Contemporary design transport category rotorcraft are utilizing the same advanced technology systems and materials as transport category airplanes. Because of this fact, a specific requirement was added by Amendment 29-24 for the consideration of lightning strike protection of required systems, equipment, and installations. The addition of paragraph (h) to § 29.1309 further defined the consideration required for the foreseeable operating condition of a lightning strike encounter on the rotorcraft.

(B) Procedures.

( 1 ) Section 29.1309(h) requires, when showing compliance with § 29.1309(a) and (b), that the effects of lightning strikes on the rotorcraft be considered.

Guidance to show compliance with § 29.1309(h) is available in a dedicated AC 20-136B.

2 ( ) Detailed means of compliance should be agreed with the authorities taking into account the effects on the rotorcraft and minimum considerations are as follows: (i) Any combination of analysis and testing should be agreed to by the authority.

ii ( ) Margins account for uncertainty in the compliance method. As confidence in the comp liance method increases, the margin can decrease. A factor of two is an acceptable margin for systems with catastrophic failure conditions, if this margin is verified by aircraft test or by analysis supported by aircraft tests. For other verification methods, the margin should be agreed upon by the cognizant certification office.

3 ( ) Flight and engine controls are examples of “critical” functions. With these critical functions defined, an analysis and testing should be performed to show compliance with § 29.1309(a) (i.e., equipment, systems, and installations performing critical functions should be designed and installed to ensure they continue to perform those intended functions subject to the conditions encountered in a worst-case lightning Page F - 39 AC 29-2C, Chg 4 5/1/2014 strike). Section 29.610 contains some methods which may be utilized for less complex mechanical systems; however, a great deal of difficulty will be experienced in trying to use these criteria to demonstrate that a very complex avionic system complies with § 29.1309(a). These avionic systems usually only require protection from indirect effects of lightning and therefore a method such as that outlined in paragraph b.(9)(iii) above is recommended. This method may be readily adapted to other avionic systems performing critical functions. Also, this identifies an acceptable quantification of the expected airborne environment. The next step involves expanding the fault/failure analysis to determine if the malfunctioning of several “essential” systems in relation to other systems would result in a hazard to a category B rotorcraft or preclude the continued safe flight and landing of a category A rotorcraft. If groups of functions are so identified, sufficient lightning protection should be provided to prevent a hazardous malfunction situation on category B rotorcraft or provide those conditions which prevent continued safe flight and landing on a category A rotorcraft are extremely improbable.

In performing this part of the analysis, attention should be given to the fact that many of the required equipment, systems, and installations may fail simultaneously with other required equipment, systems, and installations and result in a reduction of the capability of the rotorcraft but still not result in a catastrophe. An example of required equipment for which the simultaneous failure of all the required equipment is catastrophic is a failure which results in a total loss of attitude display for IFR certified rotorcraft operating in instrument meteorological conditions. Note that the analysis which is utilized to demonstrate that these failures are extremely improbable should have the encounter with a worst-case lightning strike as a given event (i.e., probability is unity). Additionally, for a category A rotorcraft, an autorotation is not considered continued safe flight and landing.

AC 29.1309A. § 29.1309 (Amendment 29-53) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.

a. Explanation . Amendment 29-53 removed § 29.1309(h) and incorporated a dedicated rule § 29.1316 to better address Electrical and Electronic System Lightning Protection.

Procedures b. . If the certification basis includes § 29.1309 (h), applicants should utilize the previous section of this guidance (section 29.1309 of this AC) to show compliance with protection from lightning effects. If the certification basis includes § 29.1316, applicants should show compliance with protection from lightning effects by referring to the applicable section of this AC (section 27.1316, Electrical and Electronic System Lightning Protection). The paragraphs other than (iv) of section 29.1309 of this AC are still applicable. This should be coordinated with the appropriate certification office early in the discussions so the path to show compliance with electrical and electronic system lightning protection is clear to the applicant and authority.

Page F - 40 6/23/23 AC 29 - 2C, Chg 9 AC 29.1309B. § 29.1309 (AMEN DMENT 29 - 59 ) EQUIPMENT, SYSTEMS , AND INSTALLATIONS.

a. Explanation . Amendment 29 - 59 revised § 29.1309 to address advances in technology and eliminate the distinction between Category A and Category B rotorcraft. The rule also:  Revised section § 29.1309 (a) to clarify the requirements for performing a failure analysis.

 Replaced section § 29.1309 (b) with language consistent with contemporary industry practices.

 Revised the alerting requirements of section § 29.1309 (c) to remove the term “ warning, ” which has been interpreted as requiring a red level alert, when the intent was to notify the crew of all required annunciations.

 Removed sections § 29.1309 (e), (f), and (g) dealing specifically with electrical systems, as they are covered by other regul atory sections.

(1) Applicability . Section 29.1309 is intended as a general requirement that is applicable to any equipment or system as installed, in addition to specific systems requirements, except as indicated below.

(i) General . If a specific part 29 requirement predefines systems safety aspects (e.g., redundancy level or criticality) for a specific type of equipment, system, or installation, then the specific part 29 requirement will take precedence. This precedence does not pre clude accomplishment of a system safety assessment, if necessary. For example, § 29.695 predefines a required level of redundancy and an implied level of system reliability. However, a system safety assessment approach may still be required to show that the implied system reliability is met and to address assessment of the failure modes.

Section 29, Subparts B, C, and D (ii) . Section 29.1309 does not apply to the design features required by Subparts B, C, and D of part 29 for aspects such as the performance, f light characteristics, structural loads, and structural strength. However, the requirements in § 29.1309 do apply to any equipment and system that the proposed design uses to comply with the requirements of Subparts B, C, and D (e.g., Health Usage Monitor ing System certified for maintenance credit and Stability Augmentation System).

(iii) Section 29, Subpart E .

Page F - 40.1 6/23/23 AC 29 - 2C, Chg 9 (A) Section 29.1309 does not apply to the type - certificated engine.

However, it does apply to the equipment and systems associated with the engine ins tallation (e.g., electrical power generation, engine displays, transducers, etc.) on the rotorcraft.

(B) Section 29.1309 does not apply to the Rotor Drive Systems (reference § 29.917(b)).

(iv) Section 29, Subpart F . Section 29.1309 does not apply to the loss of function of non - required equipment; however, it does apply to hazards introduced by the installation of that equipment on the rotorcraft and any adverse effects on other systems.

(2) Background.

(i) This guidance describes a method of showing compliance that utilizes the system safety assessment techniques available in the aerospace industry. The safety assessment process described in this guidance can be adopted and applied to equipment, systems, and installations ranging from simple systems to the most critical and most complex systems.

(ii) This guidance is not intended to impose an overly rigorous safety assessment process for those installations involving equipment or systems that are not considered to be complex.

Note: A system is considered to be complex when its operation, failure modes, or failure effects are difficult to comprehend without the aid of analytical methods.

b. Procedures .

(1) Definitions of Probability Classifications.

Note: the following terms were derived from Table 1 – “Failure Conditions Severity as Related to Probability Objectives and Assurance Levels” of SAE ARP 4761.

(i) FREQUENT . The occurrence of any frequent event may be no more often -3 than 10 times during the operational life of each rotorcraft.

Page F - 40.2 6/23/23 AC 29 - 2C, Chg 9 (ii) PROBABLE . Probable events may be expected to occur on the order of - 3 - 5 between 10 to 10 times during the opera tional life of each rotorcraft.

(iii) REMOTE . Remote events are not expected to occur a few times during the total operational life of a random single rotorcraft of a pa rticular type, but may occur several times during the total operational life of a number of rotorcraft of a particular - 5 - 7 type, that are based on a probability on the order of between 10 to 10 .

(iv) EXTREMELY REMOTE . Extremely remote events are not expected to occur during the total operational life of a random single rotorcraft of a particular type, but may occur a few times during the total operational life of all rotorcraft of a particular - 7 - 9 to 10 . type, that are based on a probability on the order of between 10 (v) EXTREMELY IMPROBABLE . Extremely improbable events are so unlikely that they need not be considered to ever occur, unless engineering judgment would - 9 require their consideration. A probability on the order of 10 or less is assigned to this classif ication.

Note: The five probability terms defined in paragraph b(1) above are intended to relate to an identified failure condition resulting from or contributed to by the improper operation or loss of a function or functions. These terms do not define t he reliability of specific components or systems.

(2) Definitions of Failure Condition Classifications . Failure Conditions may be classified according to the severity of their effects into one of the following categories (reference Figure AC 29.1309 - 2).

(i) NO EFFECT . Failure Conditions that would have no effect on safety. For example, Failure Conditions that would not affect the operational capability of the rotorcraft or increase crew workload; however, could result in an inconvenience to the occupants, excluding the flight crew .

MINOR (ii) . Failure Conditions that would not significantly reduce rotorcraft safety and th at would involve crew actions that are well within their capabilities. Minor Failure Conditions may include, for example, a slight reduction in safety margins or functional capabilities, a slight increase in crew workload, such as routine flight plan changes, or some physical discomfort to occupants.

(iii) MAJOR . Failure Conditions that would reduce the capability of the rotorcraft or the ability of t he crew to cope with adverse operating conditions to the extent that there would be, for example, a significant reduction in safety margins or functional capabilities, a significant increase in crew work load or in conditions impairing crew efficiency, phy sical distress to occupants, possibly including injuries, or physical discomfort to the flight crew.

Page F - 40.3 6/23/23 AC 29 - 2C, Chg 9 (iv) HAZARDOUS . Failure Conditions that would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to t he extent that there would be: (A) A large reduction in safety margins or functional capabilities; (B) Physical distress or excessive workload such that the flight crew’s ability is impaired to where they could not be relied on to perform their tasks accurately o r completely; or (C) Possible serious or fatal injury to a passenger or a cabin crew member, excluding the flight crew.

Note: Hazardous Failure Conditions can include events that are manageable by the crew by use of proper procedures which, if not implemented correctly or in a timely manner, may result in a Catastrophic Event.

CATASTROPHIC . Failure Conditions that would result in multiple (v) fatalities to occupa nts, fatalities or incapacitation to the flight crew, or loss of the rotorcraft.

(3) Safety Objective .

(i) The objective of § 29.1309 is to ensure an acceptable safety level for equipment and systems as installed on the rotorcraft. A logical and acceptable inverse relationship must exist between the Average Probability per Flight Hour and the se verity of Failure Condition effects, per the requirements of § 29.1309 and as shown in Figure AC 29.1309 - 1, such that: (A) Failure Conditions with No Effect on safety may be more frequent than Probable.

(B) Minor Failure Conditions may be Probable.

(C) Major Failure Conditions must be no more frequent than Remote.

(D) Hazardous Failure Conditions must be no more frequent than Extremely Remote.

(E) Catastrophic Failure Conditions must be Extremely Improbable.

Page F - 40.4 6/23/23 AC 29 - 2C, Chg 9 Probable Unacceptable Probability Remote of Failw-e Extre me ly Condition R em ote Acceptable Extre me ly Improbable M in or Ma jor Hazardous Cat as tr ophic Severity of Failw-e Condition Effect s FIGURE AC 29.1309 - 1 Relationship between Probability and Severity of Failure Condition Effects (ii) The safety objectives associated with Catastrophic Failure Conditions may be satisfied by demonstrating that: (A) No single failure will result in a Catastrophic Failure Cond ition; and (B) Each Catastrophic Failure Condition is Extremely Improbable.

(iii) Failure Conditions may be classified according to the severity of their effects to the rotorcraft, occupants, or flight crew as identified in the following Figure AC 29.1309 - 2.

Page F - 40.5 6/23/23 AC 29 - 2C, Chg 9 Table for Failure Condition Categories and Probability Definitions Effect on rotorcraft No effect on Slight Significant Large reduction in Loss of rotorcraft operational reduction in reduction in functional capabilities capabilities or functional functional or safety margins safety capabilities or capabilities or (Note 4) safety safety margin Effect on Inconvenience Physical Physical Serious or fatal injury Multiple Fatalities occupants discomfort distress, to a passenger or a excluding flight possibly cabin crew member crew including (Note 2) injuries Effect on flight No effect on Slight Physical Physical distress or Fatalities or crew flight crew in crease in discomfort or a excessive workload incapacitation work load that significant impairs ability to involves crew increase in perform tasks actions well workload or in accurately or within crew conditions completely capabilities impairing crew such as efficiency routine flight plan changes DO - 178C E D C B A Software Level (Note 3) Failure Condition Hazardous or Severe - Category Major No Effect Minor Major Catastrophic Qualitative Reasonably Extremely Remote Extremely Probability Frequent Probable Remote Improbable - 3 - 5 - 7 - 9 Quantitative No probability 10 (Note 1) 10 10 10 Probability requirement Note 1: A numerical probability range is provided here as reference. The applicant is not required to perform a quantitative analysis, or substantiate by such an analysis, that this numerical criterion has been met for Minor Failure Conditions .

Note 2: This is true if it can be shown that the given failure condition can be contained to a fatal injury of one occupant only.

Note 3: AC 20 - 174, Development of Civil Aircraft and Systems , provides guidance on using SAE ARP4754A as an acceptable method for establishing a development assurance process.

Note 4: Hazardous Failure Conditions can include events that are manageable by the crew by use of proper procedures which, if not implemented correctly or in a timely manner, may resu lt in a Catastrophic event.

FIGURE AC 29.1309 - 2 Failure Condition Categories and Probability Definitions Page F - 40.6 6/23/23 AC 29 - 2C, Chg 9 (4) Safety Assessment Process Overview.

(i) When showing compliance with § 29.1309, the considerations covered in this guidance should be addressed in a methodical and systematic manner. This section of the AC (Safety Assessment Process Overview) is provided primarily for the use of applicants who are not familiar with the various methods and procedures generally used in the industry to conduct safety a ssessments. This section and Figures AC 29.1309 - 4 and AC 29.1309 - 5 are not intended to be certification checklists, and they do not include all the information provided in this guidance. This section contains one method, but not the only method, for show ing compliance. The safety assessment process is a structured method of general applicability for showing compliance with the regulation.

Other methodologies may be used to show compliance, but the safety assessment process is the only structured process that is defined and these defined objectives would be the logical criteria applied to any other means of showing compliance. More detailed guidance can be found in SAE ARP 4761 on how to perform a safety assessment. SAE ARP 4754A, discussed AC 20 - 174, in cludes additional information on how the safety assessment process relates to the system development process.

(ii) The safety assessment process contains several parts that may be necessary as a whole or in part, depending on the criticality or complexity of th e system under consideration. The rigor of assessment and analysis performed is also dependent on the system criticality or complexity. At the extremes, some systems may be simple enough such that the entire safety assessment can be performed by observat ion and compliance shown by a simple statement. Complex and higher criticality systems may require application of all of the safety assessment elements to show compliance. Many states of varying complexity exist between these extremes that may use less t han the entire safety assessment process but require more than a simple statement to show compliance. For purposes of the FHA, an autorotation is not considered continued safe flight and landing.

(iii) Typical elements of the safety assessment process are as fo llows:  Functional Hazard Assessment (FHA)  Preliminary System Safety Assessment (PSSA) o Fault Tree Analysis (FTA)  System Safety Assessment (SSA) o Failure Mode and Effect Analysis (FMEA)  Common Cause Analysis (CCA) o Zonal Safety Analysis (ZSA) o Particular Risk Assessment (PRA) o Common Cause Analysis (CMA) Page F - 40.7 6/23/23 AC 29 - 2C, Chg 9 (A) Define the system and its interfaces, and identify the functions that the system is to perform. Determine whether or not the system is complex, similar to systems used on other aircraft, or conventional. Where multiple systems and functions are to be evaluated, consider the relationships between multiple safety assessments.

(B) Identify and classify Failure Conditions . All relevant applicant-engineering organizations, such as systems, structures, propulsion, and flight test, should be involved in this process. This identification and classification may be done by conducting a FHA, which is a top-down assessment. The FHA starts at the rotorcraft operational level to identify possible hazards to the rotorcraft, its occupants, and flight crew. When establishing the hazard classification of a function, consider for example, the known loss of the function, the erroneous behavior or misleading information provided by the function, and degradation of the function as possible hazards. The hazards are classified using the five failure condition categories found in Figure AC 29.1309-2. The identified Failure Conditions are then evaluated with respect to the intended operations and combinations of operations for which certification of the rotorcraft is sought (e.g., VFR, IFR, CAT A, IFR plus CAT A). SAE ARP 4761 provides guidance for determining when and under what conditions multiple failures should be considered.

(C) If the system is complex, it is necessary to systematically postulate the effects on the safety of the aircraft and its occupants resulting from any possible failures, considered both individually and in combination with other failures or events.

(D) Choose the means to be used to determine compliance with § 29.1309.

The depth and scope of the analysis depends on the types of functions performed by the system, the severity of the associated Failure Conditions, and whether the system is complex (see Figure AC 29.1309-5). Once the Failure Conditions are determined to be Major Failure Conditions, experienced engineering and operational judgment, design and installation appraisals, and similarity consideration for previously installed systems may be acceptable. This may be sufficient in itself or in conjunction with qualitative analyses or selectively used quantitative analyses. For Catastrophic and Hazardous Conditions, a thorough safety assessment is necessary. The applicant should obtain early concurrence from the FAA on the choice of an acceptable means of compliance.

(E) Include a description that establishes correctness and completeness and traces the work leading to the conclusions. This description should include the basis for the classification of each Failure Condition (e.g., analysis or ground, flight, or simulator tests). It should also include a description of precautions taken against common-cause failures, provide any data such as component failure rates and their sources and applicability, support any assumptions made, and identify any required flight crew or ground crew actions, including any Candidate Certification Maintenance Requirements (CCMRs), which are also referred to as Certification Check Requirements (CCRs).

Certification Maintenance Requirements (CMRs) are not the preferred method for showing compliance with the rules; therefore, they must be reviewed by the FAA for approval. Any CMRs must be included in the airworthiness limitations section of the instructions for continued airworthiness in the maintenance manual, as required by Page F - 40.8 6/23/23 AC 29 - 2C, Chg 9 § 29.1529 and Appendix A to Part 29.

Note: CMRs may be needed to help show compliance with § 29.1309 for significant latent failures. Rational methods, which usually involve quantitative analyses or relevant service experience data, should be used to determine CMR intervals. It is recognized that, for various reasons, component failure rate data is not precise enough to accurately estimate the probabilities of Failure Condit ions . This results in some degree of uncertainty, as indicated by the expression “ on the order of ” in the definition of the quantitative probability terms in paragraph b.(1) above. These intervals should have reasonable tolerances so that CMRs can be per formed concurrently with other maintenance, inspection, or check procedures not required by design for compliance with § 29.1309. Any tolerance ability should be built into the interval that is included in the airworthiness limitation section of the maint enance manual.

If CMRs are used, they and their intervals and tolerances, and any post - certification changes, or procedures provided in the type design for a rotorcraft owner or operator to make such changes, must be approved by the office having cognizan ce over the type design that relates to the system and its installation, as required by § § 21.31, 21.95, and 21.97.

(F) When assessing the acceptability of a failure condition using a quantitative analysis, the given numerical range should normally be interpre ted to be the allowable risk for an hour of flight time based on a flight of mean duration for the rotorcraft type.

However, when assessing a function which is used only at a specific time during a flight, the probability of the failure condition should be calculated for the specific period and expressed as the risk for the flight condition, takeoff, landing, etc., as appropriate. This is only true for those systems that cannot fail undetected when they are being used. The probability of failure for those s ystems should be calculated from that value based on the extended period of not only the active operation, but also the exposure time for the latent failure.

Note: If a quantitative analysis is used to help show compliance with the regulations for equipme nt or systems that, are installed and only required for a specific operating condition for which the rotorcraft is thereby approved, credit is usually not allowed due to the fact that the operating co ndition does not always exist. If an applicant does requ est that such credit be allowed, they must obtain the approval of the FAA.

Page F - 40.9 6/23/23 AC 29 - 2C, Chg 9 Cert.iflcaUon Maintenance Requir-emenls Requirements flight Manual Requirements Assessment of Rotorcnifl Rotorcreft Level functional Leven System Safety Hazard Assessment Assessment System Safety System Level FHAs Assessments Anal yses Note: FHA may bo based on a design and installation appraisal for these ayatema.

FIGURE AC 29.1309 - 4: Safet y Assessment Process Overview Page F - 40.10 6/23/23 AC 29 - 2C, Chg 9 Note: FHA -Y be b<atodoua dOlip lad lllatalladoa lllP!alAI l>r lbeao l)'IIOIIII, Conduct Functional Hazard Assessment Yes Verily by 0.sign and lnmllation Appraisal.

No Develop Sunilinty A1wysis and Obbin Approval from FAA No • Conduct Qual1t-tfrc or Quantiuti, ·t A.JW y1 l$ or both.

*NOTE : cawuophit faitutt conditiont w ill hkely require bolb qualitative and qu,ntnau,-. aoaJy,is, depending on sys ttm c:ompluiry ~ FIGURE AC 29.1309 - 5: Depth of Analysis Flowchart Page F - 40.11 6/23/23 AC 29 - 2C, Chg 9 (5) Assessment Methods . Various methods for assessing the causes, severity, and probability of Failure Conditions are available to support experienced engineering and operational judgment. Some of these methods are structured. The various types of analysis are based on either inductive or deductive approaches. Probability assessments may be qualitative or quantita tive. Descriptions of some types of analysis are provided below and in SAE ARP 4761.

(i) Design Appraisal . This is a qualitative appraisal of the integrity and safety of the system design.

(ii) . This is a qualitative appraisal of the integ rity and Installation Appraisal safety of the inst allation. Any deviations from normal, industry - accepted installation practices such as clearances or tolerances, should be evaluated, especially when appraising modifications made after entry into service. Hazards introduced by the installation of the system should also be evaluated.

(iii) Failure Modes and Effects Analysis (FMEA) .

(A) This is a structured, inductive, bottom - up analysis, which is used to evaluate the effects on the system and the rotorcraft of each possible eleme nt or component failure. When properly formatted, it will aid in identifying latent failures and the possible causes of each failure mode. SAE ARP 4761 provides methodology and detailed guidelines that may be used to perform this type of analysis. An FME A could be a piece part FMEA or a functional FMEA. For modern microcircuit based line replaceable units (LRUs) and systems, an exhaustive piece part FMEA may not be practically feasible given current technology. In that context, an FMEA may be more functio nal than piece part oriented. A functional oriented FMEA can lead to uncertainties in the qualitative and quantitative aspects, which can be compensated for by more conservative assessment, such as: 1. Assuming all failure modes result in the Failure Cond itions of interest, 2. Carefully choosing system architecture, or 3. Taking into account the experience lessons learned on the use of similar technology.

(B) Specific reliability numbers are not shown in § 29.1309. The necessary degree of reliability is a function of the criticality of the system under consideration. Acceptable sources of component failure rates include (1) military aircraft component failure or fault reports or handbooks, such as MIL - HDBK - 217C; (2) operator Page F - 40.12 6/23/23 AC 29 - 2C, Chg 9 or manufacturer component m alfunction or defect records, such as airline component defect records on sufficiently similar component designs; and (3) laboratory life tests.

Fault Tree or Dependence Diagram Analysis . This is a s tructured, (iv) deductive, top - down analysis that is used to identify the conditions, failures, and events that would cause each defined Failure Condition. It is a graphical methods of identifying the logical relationship between each particular Failure Condition and the primary element or component failures, ot her events, or combinations thereof that can cause it.

A failure modes and effects analysis may be used as the source document for those primary failures or other events.

Markov Analysis (v) . A Markov model (chain) represents various system - states and the rel ationships among them. The states can be either operational or non operational. The transitions from one state to another are a function of the failure and repair rates. Markov analysis can be used as a replacement for fault tree and dependence diagram an alysis, but it often leads to a more complex representation, especially when the system has many states. It is recommended that Markov analysis be used when fault tree or dependence diagrams are not easily usable, namely to take into account complex trans ition states of systems that are difficult to represent and handle with classical fault tree or dependence diagram analysis.

Common Cause Analysis (CCA) (vi) . The acceptance of adequate probability of Failure Conditions is often derived from the assessment of multiple systems based on the assumption that failures are independent. Therefore, it is necessary to recognize that such independence may not exist in the practical sense and specific studies are necessary to ensu re that independence can either be assured or deemed acceptable.

CCA identifies common causes for Failure Conditions . Once identified, the source of the common cause is either eliminated from the design, the common cause or the system design is made to be tolerant of the related fault or malfunction, or the common cause or system design combination is found to be acceptable due to the failure or malfunction condition category it represents. The potential for failures or malfunctions due to common causes i s inherent in designs that provide multiple functions reliant on common hardware or common software. This is also true for systems that produce related functions and share a common installation area. The installation area may represent a threat from severa l sources such as Electro - Magnetic Interference (EMI), mechanical hazards, and environmental influences. The CCA is sub - divided into three areas of study: Zonal Safety Analysis (ZSA) . This analysis has the objective of (A) ensuring that the equipment install ations within each zone of the rotorcraft are at an adequate safety standard with respect to design and installation standards, interference between systems, and maintenance errors. The ZSA examines the physical zone of the rotorcraft in which the system under consideration is installed to ensure that the surrounding equipment or appliance installations do not compromise the system Page F - 40.13 6/23/23 AC 29 - 2C, Chg 9 independence requirements. Mechanical failures that might generate fragments that could damage the system under evaluation are an example of the type of event that would be considered in a zonal analysis.

(B) Particular Risk Analysis . Particular risks are defined as those events or influences that are outside the systems concerned. Some examples are fire, leaking fluids, bird strike, tire burst, high intensity radiated fields exposure, lightning, and uncontained failure of high energy rotating machines. Each risk should be the subject of a specific study to examine and document the simultaneous or cascading effects or influences that may violate independence.

(C) Common Mode Analysis (CMA) . This analysis is performed to confirm the assumed independence of the events that were considered in combination for a given Failure Condition. The effects of specification, design, implementation, installation, maintenance and manufacturing errors, environmental factors other than those already considered in the particular risk analysis, and failures of system components should be considered. The CMA considers many aspects of design; one is design dissimilarity for both hardware and software. Without dissimilarity for redundant applications, there is a possibility that a hardware or software failure or malfunction could occur in the same flight for all redundant subparts of a system with a Catastrophic or Hazardous Condition Category. Another major contribution from the CMA is the determination of which failure or malfunction combinations inputs to the Fault Tree Analysis (FTA) must be independent for events that are Catastrophic or Hazardous.

This analysis is iterative in nature as it will be employed early in design to identify possible common modes for failures or malfunctions and then used after the design is complete to determine if the FTA goals have been met.

(6) Documentation . All laboratory, ground, and flight tests and failure analyses, must be documented in sufficient detail to show compliance with § 29.1309 and must be included in the type design file. Section 21.31(a) provides the regulatory basis for requiring this documentation. If the applicant elects to use a numerical reliability and probability analysis, it must also be documented in sufficient detail.

(7) . RTCA DO-178C, Software Considerations in Airborne Systems Software and Equipme nt Certification , dated December 13, 2011, is the latest standard and may be used for qualification and subsequent approval of airborne software. See AC 20 - 115C , Airborne Software Assurance, for guidance on using DO-178C and earlier standards.

(8) Airborne Electronic Hardware (AEH) . For airborne complex custom micro coded electronic components assessed as level A, B, or C in the safety assessment, RTCA DO-254, Design Assurance Guidance for Airborne Electronic Hardware , dated April 19, 2000, provides certification guidance for qualification and subsequent approval of these components. Per DO-254, the component will need to be identified as either “simple” or “complex” and tagged to one of the five failure condition classifications.

Based on these identifiers, DO-254 provides guidance to show compliance with Page F - 40.14 6/23/23 AC 29 - 2C, Chg 9 § 29.1309. Although RTCA DO - 254 applies specifically to complex custom micro - coded components, applicants are highly encouraged to apply DO - 254 to other hardware components up to the LRU and systems level.

(9) Environmental Qualification.

(i) Laboratory Tests.

(A) Environmental Standards . In order to ensure that the components and systems under consideration will function properly when exposed to adverse environments, they should be tested in the laboratory under a simulated adverse environment. If a TSO exists and is appropriate in environmental range and performance for an equipment installation, it is preferable the equipment be TSO approved. If there is no applicable TSO or an existing TSO does not provide for a sufficiently adverse environment, the latest revision of RTCA DO-160 is an acceptable environmental standard for laboratory qualification of aircraft equipment.

(B) Explosion Tests . Those items of electrical and electronic equipment that are to be located in areas subject to flammable fluids and vapors, as a result of any single probable malfunction or failure, including failure of couplings or lines should be tested as an ignition source. These tests consist of normal operation of the equipment in a physically contained explosive atmosphere. The explosion test procedure in the latest revision of DO-160 will satisfy this requirement. Section 29.863 of this AC provides further guidance on safety from explosion. Another standard that is equivalent to the latest revision of DO-160, it may also be accepted to satisfy this requirement.

(ii) Installed Environmental Tests . After the environmental ratings of the components and systems have been established, it should be ensured that as installed, these ratings will not be exceeded. Normally, installed equipment need not be instrumented and tested in flight nor is it necessary to instrument the compartment or rack where the equipment is installed. Satisfactory environment and equipment compatibility are ensured by selection of the proper environmental category of laboratory tests. The category is determined by the type of aircraft (reciprocating or turbine) and flight envelope (altitude and temperature). Exceptions to normal installations are (a) Alternator or generator cooling, where radiated and conducted heat is almost always uncertain, and cooling air temperatures and flow rates are uncertain; (b) Where flight tests reveal excessive instrument panel vibration. In this case, the panel should be instrumented, tested, and, if necessary, design improvements made; and (c) Any other case where good engineering judgment and application of sound engineering principles indicate a high likelihood that the installed environment is more severe than the equipment is capable of operating within.

Page F - 40.15 6/23/23 AC 29 - 2C, Chg 9 (A) Tem perature Tests .

Temperature tests may be accomplished by instrumenting the installed equipment environment with a recorder that provides a permanent record of time, altitude, and temperature. The pertinent temperature should be recorded as the rotorcraft is operated throughout its altitude range, including ground operation. The maximum and minimum temperatures recorded should be corrected degree for degree to ensure the equipment under test remains within its temperature rating while the rotorcraft opera tes throughout its approved ambient temperature envelope. (For generator or alternator cooling test procedures, refer to section 29.1351 of this AC.)

Section 29.1043(b) requires the maximum approved operating OAT be at least 100° F for powerplant - mounted accessories such as starter generators and vacuum pumps.

Due to the impracticality of the 100° F hot day temperature limit, rotorcraft systems mounted on the powerplant are normally evaluated for at least 115° F hot day sea level conditions with correspond ing 3.6° F/1,000 - foot correction. The maximum hot day OAT at sea level must be specified in the rotorcraft flight manual per § § 29.1309 and 29.1581(a)(2). Section 29.1043(b) is the regulatory basis for the lapse rate of 3.6° F/1,000 feet. This lapse rat e should be applied regardless of the hot day sea level temperature the applicant chooses to certify for operation.

The § 29.1043(b) maximum ambient temperature definition should not be confused with operating temperatures in closed areas. Closed equipme nt rack areas can easily reach temperatures of 140° F when sitting on the ramp in the southern United States in midsummer. Normally, proper selection of the altitude temperature category in the latest revision of DO - 160 will ensure compliance.

In some ca ses, the equipment manufacturer furnishes temperature limits for internal critical parts (for example, brushes, bearings, or field windings on DC generators). In these cases, it is better to record the critical component temperature rather than equipment o r equipment environment temperature.

The following illustrates an acceptable high temperature evaluation method: T OAT MAX = Maximum outside air temperature at which temperature tests are conducted.

T = Maximum temperature to which the installed equip ment has been tested MAX in the laboratory.

T TEST MAX = Maximum installed equipment temperature recorded during tests.

T = The high reference outside air temperature. It varies with altitude starting ORH at the highest sea level temperature at which rotorcra ft operation is to be Page F - 40.16 6/23/23 AC 29 - 2C, Chg 9 approved and decreases at3.6° F/1,000 foot altitude. It can be no less than 100° F (reference § 29.1043(b)); however, it can be as high as the applicant wants.

T H MAR = Temperature margin between the maximum equipment temperature substantiated in the laboratory and the maximum installed equipment temperature when the rotorcraft is operating in the highest available OAT and approximately corrected at the altitude under consideration. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.

T = T - (T + (T - T )) H MAR MAX TEST MAX ORH OAT MAX Example #1 : Assume the applicant is seeking approval for rotorcraft operatio n at the lowest acceptable OAT, at sea level, of 100° F and T MAX for Generator Brush = 295° F at maximum load current throughout the altitude range. In - flight test data are: Measured Altitude (ft, MSL) Cylinder Temp (T TEST MIN ) OAT = T OAT MAX sea leve l 275°F 90°F 5,000 270°F 80°F 10,000 285°F 60°F 15,000 294°F 42°F 20,000 290°F 20°F First, T must be calculated for each altitude test point.

ORH @ sea level, T ORH = 100° F @ 5,000 ft., T ORH = 100° F - 5,000 ft., x 3.6°/1,000 ft. = 82° F @ 10,000 ft., T ORH = 100° F - 10,000 ft. x 3.6°/1,000 ft. = 64° F @ 15,000 ft., T = 100° F - 15,000 ft. x 3.6°/1,000 ft. = 46° F ORH @ 20,000 ft., T = 100° F - 20,000 ft. x 3.6°/1,000 ft. = 28° F ORH Then at sea level: T = 295 - (275 + (100 - 90)) = 10° F H MAR At 5,0 00 feet: T H MAR = 295 - (270 + (82 - 80)) = 23° F Page F - 40.17 6/23/23 AC 29 - 2C, Chg 9 At 10,000 feet: T H MAR = 295 - (285 + (64 - 60)) = 6° F At 15,000 feet: T H MAR = 295 - (294 + (46 - 42)) = - 3° F At 20,000 feet: T H MAR = 295 - (290 + (28 - 20)) = - 3° F Since T H MAR comes out negative at the 15,000 - and 20,000 - foot points, the generator fails. It will be necessary for the applicant to reduce the maximum load current, improve cooling, or otherwise change the design to ensure the generator is operating within its appr oved temperature limit of 295° F.

5. In most cases, the equipment is laboratory tested to minimum temperatures as severe as that of the rotorcraft’s maximum certified altitude on a minimum temperature day. Therefore, unless equipment minimum temperature is affected by refrigeration or other temperature reducing environments, actual installed instrumented minimum temperature tests are unnecessary. If low temperature evaluation is necessary for the installed equipment, the following is an acceptable method : T OAT MIN = Minimum outside air temperature at which temperature tests are conducted.

T = Minimum temperature to which the installed equipment has been tested in MIN the laboratory.

T TEST MIN = Minimum installed equipment temperature recorded during tes ts.

T = The low reference outside air temperature. It varies with altitude starting ORL at the lowest sea level temperature at which rotorcraft operation is to be approved and decreases at 3.6° F/1,000 - foot altitude.

T1 MAR = Temperature margin between the minimum equipment temperature substantiated in the laboratory and the minimum installed equipment temperature. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.

Page F - 40.18 6/23/23 AC 29 - 2C, Chg 9 T = - (T - (T + (T - T ))) 1 MAR MIN TEST MIN ORL OAT MIN Note: This equation assumes all temperatures are negative. It is necessary to place a ( - ) in front of the right side of the equation in order to convert the T1 MAR value to the conventional positive answer for acceptance and a negative answer for rejection. Temperature in the 0 to 32° F range can be handled by conversion to the centigrade scale.

Example #2 : Assume the applicant is seeking a low temperature operating limit at sea level of - 25° F. Assume the hydraulic control c ylinder has been substantiated in the laboratory to operate at a cylinder temperature of - 40° F. The in - flight test data are: Measured Altitude (ft, MSL) Cylinder Temp (T TEST MIN ) OAT = T OAT MAX sea leve l 0 °F - 25 °F 5,000 - 9 °F - 45 °F 10,000 - 21 °F - 59 °F 15,000 - 32 °F - 65 °F 20,000 - 40 °F - 69 °F Page F - 40.19 AC 29 - 2C, Chg 9 6/23/23 (T must be calculated for each altitude test point) ORL @ sea level, T ORL = - 25° F @ 5,000 ft., T ORL = - 25° F - 5,000 ft., x 3.6°/1,000 ft. = - 43° F @ 10,000 ft., T = - 25° F - 10,000 ft. x 3.6°/1,000 ft. = - 61° F ORL @ 15,000 ft., T = - 25° F - 15,000 ft. x 3.6°/1,000 ft. = - 79° F* ORL @ 20,000 ft., T ORL = - 25° F - 20,000 ft. x 3.6°/1,000 ft. = - 97° F* *According to § 29.1043(b), the lowest temperature to be considered is - 69.7° F.

Then at sea level: T 1 MAR = - ( - 40 - (0 + ( - 25 - ( - 25)))) = 40° F At 5,000 feet: T 1 MAR = - ( - 40 - ( - 9 + ( - 43 - ( - 45)))) = 33° F At 10,000 feet: T = - ( - 40 - ( - 21 + ( - 61 - ( - 59)))) = 17° F 1 MAR At 15,000 feet: T = - ( - 40 - ( - 32 + ( - 69.7 - ( - 65)))) = 3.3° F 1 MAR At 20,000 feet: T 1 MAR = - ( - 40 - ( - 40+ ( - 69.7 - ( - 69)))) = - 0.7° F In this example, there is an acceptable margin at all altitudes up to and including 15,000 feet. However, at 20,000 feet, the margin is negative and the system fails.

(B) Vibration tests . Normally, installed vibration tests are not necessary for equipment qualified in accordance with the latest revision of RTCA DO - 160. This document categorizes vibration tests according to installed rotorcraft eq uipment location such as fuselage, engine compartment, instrument panel, and equipment rack.

However, installed equipment vibration tests may be necessary when it appears the equipment location environment may exceed the laboratory - tested equipment vibrati on limits.

Page F - 40.20 6/23/23 AC 29 - 2C, Chg 9 (C) Altitude tests . If the equipment has been laboratory tested to the maximum certified altitude of the rotorcraft, installed altitude tests are unnecessary. The installed equipment must be either laboratory tested or tested in the rotorcraft t o the maximum certified altitude of the rotorcraft. See introductory text section of § § 29.1309 and 29.1309(d).

(10) Lightning Strike Protection of Full Authority Digital Engine Controls (FADEC).

(i) Explanation.

(A) The following discussion is written specifically for FADEC with an alternate technology backup fuel control installed on part 29 rotorcraft that meet the category A engine isolation requirements. The requirement for increased consideration of lightning strike encounter effects on avionic equipment and systems has been brought about by the increased use of avionics to perform functions, the failure or malfunction of which could result in a hazard to the rotorcraft. The susceptibility of current high technology avionic systems is increased by the use of large scale integration, very large scale integration, and complementary metallic oxide silicon technologies, that exhibit a greatly reduced tolerance to large amplitude, low energy electrical transients as compared to conventional bipolar technology, and the reduced physical protection and electromagnetic shielding afforded aircraft avionic systems by the advanced technology composite airframe materials. Additionally, processor-based systems have the failure phenomenon of digital upset. A digital upset occurs when a system, perturbed by an electrical transient, ceases proper operation in accordance with its embedded software while suffering no apparent component or device damage.

(B) Since elements of electrical and electronic engine subsystems are typically spread throughout much of the rotorcraft, transients caused by lightning are coupled into the subsystem interface cables and may damage the system or cause upset. Effective lightning protection should be designed and incorporated into these systems. Reliance upon redundancy as a means of protection against lightning effects is generally not adequate because lightning electromagnetic fields and structural IR voltages usually interact (to some extent) with all electrical wiring aboard a rotorcraft.

(C) The testing and analysis outlined in this discussion are methods by which the FAA may be assured that when the rotorcraft experiences “the foreseeable operating condition” of a worst-case lightning strike encounter that the electronically controlled engines will continue to “perform their intended function” and therefore be in compliance with § 29.1309 as installed.

(D) The definition of what constitutes a full authority engine control is not at this time clearly defined. However, it has been accepted in past certification that any control that relies upon the electronics for the function on which Civil Certification or Page F - 40.21 6/3/23 AC 29-2C, Chg 9 Military Qualification is based (e.g., rotor speed governing) is a full authority control, regardless of the backup control mode provided. If engine certification or qualification can be achieved without the electronic control that is subsequently added to achieve improved operational efficiency in the aircraft, the control is “supervisory.” However, if the controls used in a multiengine rotorcraft have a common failure caused by a lightning strike that could result in simultaneous failures that would cause a reduction in power greater than the loss of one engine, this would also be considered “full authority.” Note: If OEI ratings are approved, cumulative loss of power from all engines must be limited to allow flight manual performance based on OEI ratings.

Procedures . Although not a regulatory requirement, it is recommended (ii) that a formal written certification plan be used to assure regulatory compliance. The use of this plan is beneficial to both the applicant and the FAA because it identifies and defines an acceptable resolution to critical issues early in the certification process.

These are the usual steps to be followed when utilizing a certification plan: (A) Prepare a certification plan that describes the analytical procedures and the qualification tests to be utilized to demonstrate protection effectiveness. Test plans should describe the rotorcraft and FADEC system to be utilized, test drawing(s) as required, the method of installation that simulates the production installation, the lightning zone(s) applicable, the lightning simulation method(s), test voltage or current waveforms to be used, diagnostic methods, and the appropriate schedules and location(s) of proposed test(s).

(B) Obtain FAA concurrence that the certification plan is adequate.

(C) Obtain FAA concurrence on detail part conformity of the test articles and installation conformity of applicable portions of the test setup.

(D) Schedule FAA witnessing of the test.

(E) Submit a final test report describing all results and obtain FAA approval of the report.

(iii) Definition of Environment . The SAE AE4L Committee report dated June 6, 1999, remains acceptable criteria to define the worst-case lightning strike that may be encountered by the rotorcraft in service. An additional explanation of the lightning environment can be found in FAA Report DOT/FAA/CT-89/22, Aircraft Lightning Protection Handbook . This handbook will assist aircraft design, manufacturing, and certification organizations in protecting aircraft against the direct and indirect effects of lightning strikes, in compliance with the FAA’s regulations. It presents comprehensive test criteria to provide the essential information for the in-flight lightning protection of all Page F - 40.22 6/23/23 AC 29 - 2C, Chg 9 types of fixed wing, rotary wing, and powered lift aircraft of conventional, composite, and mixed construction and their electrical and fuel systems. The handbook contains chapters on the natural phenomenon of lightning, the interaction between the aircraft and the electrically charged atmosphere, the mechanism of the lightning strike, and the interaction with the airframe, wiring, and fuel system. Further chapters cover de tails of designing for optimum protection; the physics behind the voltages, currents, and electromagnetic fields developed by the strike; and shielding techniques and damage analysis. The handbook ends with discussion of test and analytical techniques for determining the adequacy of a given protection scheme. For new designs and applications, the revised definitions of the lightning environment in AC 20 - 136B, Aircraft Electrical and Electronic System Lightning Protection , dated September 7, 2011, should be used.

Certification Plan . The following subjects are not intended to (iv) provide a compl ete list of the items that should be included in the certification plan, but rather highlight some of the areas that should receive consideration. The certification plan should address the total protection that is required to allow the FADEC to continue to operate properly when the rotorcraft experiences a worst - case lightning strike encounter.

(A) Determination of Lightning Strike Attachments . Determine the locations on th e rotorcraft where lightning strike attachment is likely to occur and the portions of the airframe through which currents may flow between attachments. The main and tail rotors are recognized as likely attachment points; however, consideration should be g iven to all possible attachment points. The swept stroke phenomenon may not exist for all lightning strike encounters due to the fact that the rotorcraft may be airborne with little or no airspeed.

(B) Establish the Lightning Environment . Establish the comp onents of the total lightning event to be considered. These are the currents and voltages that are described in the definition of the environment.

(C) Full - Level, Complete Vehicle Testing . In accordance with traditional FAA policy, the demonstration that the FADEC installed in a complete type design rotorcraft will continue to operate properly when exposed to a worst - case lightning strike is sufficient to demonstrate compliance with § 29. 1309(a). Because of the difficulties involved in utilizing this type of an approach, it is generally not used.

(D) Analytical Processes . A description should be given in the certification plan of the analytical process and certification tests to be utilized to demonstrate protection effectiveness. Typically, the certification plan will include a combination of analysis and tests. (Analytical techniques are most often utilized to predict the levels of lightning - induced transients in interconnecting wiring.) In most cases, successful analyses are based upon well - defined geometrical or electrical parameters such as structural dimensions and materials resistivity. When electrical characteristics of structural materials are not well established, development tes ts are Page F - 40.23 6/23/23 AC 29 - 2C, Chg 9 often utilized to obtain this data, which is subsequently utilized in an analysis. In more complex structures or electrical and electronic system installations, it is sometimes difficult or impossible to define the problem in terms that can be anal yzed. In these cases, development or verification testing is often relied upon. The purpose of the certification plan is to show how developmental tests, analyses, and verification tests are combined to demonstrate protection design adequacy. In certain cases, previously verified designs can be incorporated and their adequacy confirmed by reference to previous verifications. Such reference should also be incorporated in the certification plan.

1. The verification testing should be conducted on a system that simulates as closely as possible the installed configuration. As few items as possible of actual hardware should be simulated.

2. Unless the applicant has opted to follow the guidance in AC 20 - 136B as a means of showing compliance with § 29.1309(d), th en the following or other acceptable guidance should be followed. The use of various analytical processes usually requires that the system component tolerance is established. The SAE AE4L Committee Report No. AE4L - 81 - 2 is one example of an acceptable ref erence to the testing that will determine these tolerances. The testing performed to determine the tolerance level of the control computer should include a consideration for the occurrence of a non - recoverable digital upset. One method to provide this co nsideration is to have the unit powered and the processor operating normally under software control (usually this should be the exact software for which approval is sought) when the test is performed. If strike testing is used, then several shots should b e made to develop enough data to provide a reasonable confidence level. It is acceptable for the engine manufacturer, while it is obtaining the engine type certificate, to accomplish this bench testing to determine the level of tolerance of the FADEC syste m components to lightning encounter indirect effects. This approach has the advantages that the bench tests are not necessarily required to be repeated when the engine is installed in a different airframe. This recommendation is not meant to add a requir ement to the engine manufacturer but to propose a more efficient method of certification. If this tolerance was not determined by the engine manufacturer, the applicant installing the FADEC in a rotorcraft would be expected to furnish this data.

3. As with a ny analytical method, it is prudent to include a margin of safety to account for the uncertainties involved in the analytical and testing processes. Margins account for uncertainty in the compliance method. As confidence in the compliance method increases , the margin can decrease. A factor of two is an acceptable margin for systems with catastrophic Failure Conditions , if this margin is verified by aircraft test or by analysis supported by aircraft tests. For other verification methods, the margin should be agreed upon with the cognizant certification office.

4. When an analysis has no associated full - scale vehicle testing to confirm the analysis, the analysis should be very rigorous. Additionally, it Page F - 40.24 6/23/23 AC 29 - 2C, Chg 9 should be expected in this situation that this analysis indicates a very large margin of protection. Many factors need to be considered in determining what constitutes an acceptably large margin. The specific additional margin required should be based on an assessment of the inherent uncertainty of a given an alysis.

(E) Pass or Fail Criteria . The certification plan should address pass or fail criteria for the testing and analysis to be performed. The following items should be satisfied to assure acceptable system performance: 1. No immediate crew action mu st be needed.

2. Automatic control of the engine cannot be lost for any appreciable period of time. The engine should not be allowed to be out of control for a period of time that will result in a hazard in a worst - case flight condition. Any rapid, uncon trolled divergence would not be acceptable.

3. No crew action should be required to reset the system. This is not to imply that the system cannot be designed with a manual reset, but the manual reset cannot be used to show compliance to recover from a dig ital upset.

4. The resumption of engine control after an upset should be reasonably within the range that existed before the upset.

5. No critical data can be lost.

6. After the system recovers, if the performance of the system has been degraded in a noncr itical manner that would reduce the capability of the rotorcraft or the ability of the pilot to cope with adverse operating conditions, then an alert to the crew about this system degradation is needed.

(F) System Installation Considerations . In most cas es, the installation of the system components is a constituent part of the lightning protection. This is particularly true in the use of shielding techniques. If these installation features are required for adequate lightning protection, consideration sh ould be given to ensure that their effectiveness is not derogated in service. ICAs must be furnished to the aircraft owner on initial delivery, and thereafter made available to parties who service and operate the rotorcraft, to allow them to take actions necessary to ensure the continued effectiveness of the system lightning protection. See § 21.50(b).

Page F - 40.25 5/1/2014 AC 29-2C, Chg 4 AC 29.1316. § 29.1316 (Amendment 29-53) ELECTRICAL AND ELECTRONIC SYSTEM LIGHTNING PROTECTION.

a. Explanation .

(1) During the original design and development of rotorcraft and the development of regulations concerning these aircraft, little attention was given to protection from the meteorological phenomenon of lightning. This was, in part, because the early aircraft were constructed mostly of metal and had little, if any, dependence on advanced technology systems. Contemporary design of rotorcraft utilizes the same advanced technology systems and materials as airplanes. Because of this, a specific requirement was added by Amendment 29-53 establishing standards for lightning strike protection of required systems, equipment, and installations. Section 29.1316 replaces paragraph (d) of § 29.1309, which previously defined the consideration required for the foreseeable operating condition of a lightning strike encounter on the rotorcraft and systems.

Note: Aviation authorities in other countries may not have issued regulations or standards similar to § 29.1316 for lightning protection. As a result, authorities in other countries may use other means, such as special conditions, to establish requirements for lightning protection.

b. Procedures .

(1) Guidance for you to show compliance with the applicable sections of § 29.1316 is available in a dedicated AC 20-136B. AC 20-136B is available for download from the FAA Regulatory Guidance Library website at http://rgl.faa.gov/ .

(2) EASA Interpretative Material is introduced by a CRI referring to EUROCAE documents (ED) ED-113, ED-81, and FAA AC 20-136B.

Page F - 41 AC 29-2C, Chg 4 5/1/2014 AC 29.1317. § 29.1317 (Amendment 29-49) HIGH INTENSITY RADIATED FIELDS (HIRF) PROTECTION.

a. Explanation .

(1) Regulation amendments in 2007 added requirements for the protection of aircraft electrical and electronic systems from the effects of High Intensity Radio Frequency (HIRF) environment. This effort was due to technological advances in airframe and electronic systems design and a concurrent increase in the levels of radiated power in the aircraft environment. These changes have raised vulnerability to the electromagnetic environment of the electrical and electronic systems, which perform critical and essential functions. Prior to this regulatory requirement, the issuance of special conditions for products involving advanced electrical and electronic systems provided an adequate level of safety. The new regulation included a five-year period of relief from the new testing requirements by allowing an applicant to show that the system continues to comply with previously issued HIRF special conditions. Beginning December 1, 2012, data used to show compliance as part of a previously issued special condition is no longer accepted as a means of showing compliance with paragraph (a) of § 29.1317. All systems will be required to show compliance data to the appropriate paragraphs and sections of the regulation, based on the HIRF safety analysis.

(2) EASA has not yet adopted regulatory requirements for HIRF protection.

EASA is using special conditions invoking the JAA INT POL 27/29-1 Issue 3, dated January 10, 2003. The requirements addressed in the JAA INT POL are similar to those of § 29.1317(a)(b)(c).

b. Procedures .

(1) Guidance for you to show compliance with the applicable sections of § 29.1317 is available in AC 20-158. This AC is available for download from the FAA Regulatory Guidance Library website: http://rgl.faa.gov/ .

(2) EASA Interpretative Material provided in EURO CAE ED-107 and AMJ 20.1317 is considered by the FAA to be technically equivalent to the guidance material in AC 20-158 and may be used as a means of compliance with § 29.1317.

Page F - 42 2/4/2016 AC 29-2C, Chg 7 AC 29.1321. § 29.1321 (Amendment 29-21) ARRANGEMENT AND VISIBILITY.

a. Background . This section is the first in a series that concerns the installation of instruments. Other sections address specific requirements for individual instruments.

The instruments should be arranged so that pilots can see critical information without moving their heads, and without undue distraction. Additionally, for instrument flight, 14 CFR 29.1321(b) requires grouping of the attitude, altitude, airspeed, and compass indicators in the standard “T.” The design for the instrument location and arrangement, with respect to the pilot’s seat, should accommodate pilots from 5’2” to 6’0” in height.

th th Alternatively, a methodology to accommodate the 5 to the 95 percentile of male and female pilots may be used. Pilots within this height range should be able to see and, where necessary, reach and operate all of the displays.

b. Procedures .

(1) Flight, navigation, and powerplant instruments should be placed such that the pilot and copilot, if a required crewmember, can easily see and read these instruments when seated normally. Additionally, the instruments’ location should minimize unnecessary pilot head turning. The instruments necessary for safe operation (airspeed indicator, gyroscopic direction indicator, gyroscopic bank-and-pitch indicator, slip-skid indicator, altimeter, rate-of-climb indicator, master caution or warning (if installed), rotor tachometers, and the indicator most representative of engine power) should be installed immediately in front of pilots in their primary field of view (FOV).

(i) Primary FOV: Figure AC 29.1321-1 depicts the primary FOV for rotorcraft.

Display high priority information and the primary flight information in the primary FOV.

LT Horizontal Vertical Figure AC 29.1321-1 PRIMARY FIELDS OF VIEW (ii) Secondary FOV: Figure AC 29.1321-2 depicts the secondary fields of view. Signals and information pilots scan as part of monitoring the aircraft and its performance are located in the secondary FOV. The priority of the information and signals when locating information in the secondary field of view should be considered.

Page F - 43 AC 29-2C, Chg 7 2/4/2016 The farther away from the primary field of view a signal is located, the greater the chance a pilot will not notice it without a prompt. Additionally, the pilot’s scan workload to see information increases the farther away from the pilot’s primary field of view it is located. The secondary FOV is bound laterally to minimize the need for pilots to turn their heads to view the information. Examples of information typically located in the secondary FOV are: (A) Ancillary navigation information (moving maps, weather displays, etc).

(B) Secondary powerplant information, such as engine oil pressures and temperatures.

(C) Caution advisory and warning system (CAWS) panel (unless a master caution or warning indicator is not located in the primary FOV, in which case the CAWS panel should be as close to the pilots primary FOV as possible).

(D) Autopilot controls (E) Navigation controls.

UP LT Horizontal Vertical 30° C/L 30° NOM RT ON Figure AC 29.1321-2 SECONDARY FIELDS OF VIEW (2) The other powerplant instruments should be grouped together and visible to any appropriate crewmember. On multiengine rotorcraft, there should be no confusion regarding which engine an individual gauge represents. This is usually accomplished by mounting the engine gauges vertically in the center of the instrument panel. Identical gauges are placed next to each other and positioned from left to right in the same position and sequence as the corresponding engine location in the airframe.

Page F - 44 2/4/2016 AC 29-2C, Chg 7 (3) An evaluation should determine that vibration of the instrument panel does not exceed the tolerances of the instruments. The instrument manufacturer will usually provide data that indicate the level of vibration for which the instrument has been qualified. The flight test evaluation of the rotorcraft should explore and determine that the vibration of the instrument panel does not affect the readability of the instruments.

To meet these two criteria, it has been necessary in some installations to “shock mount” or otherwise isolate the instrument panel.

(4) The flight test evaluation should also determine that the flags or malfunction indicators of the instruments should be readily visible in all combinations of lighting for approved kinds of operations.

(5) For IFR-certified rotorcraft, there is an additional requirement that the airspeed, altimeter, attitude, and compass instrument be located in a standard “T” configuration in front of the pilot. This configuration is: airspeed - attitude - altitude | compass (AC 29-2, section AC 29 Appendix B further addresses IFR panel arrangement.)

(6) Geometric variation from a perfect “T” has been permitted. Each installation should be evaluated for suitability based on criteria such as panel size, ease of scan, and readability of the individual elements in the overall presentation. AC 25-11, Transport Category Airplane Electronic Display Systems, provides additional guidance for “glass cockpit” installation.

Page F – 44.1 AC 29-2C, Chg 7 2/4/2016 AC 29.1322. § 29.1322 (Amendment 29-12) WARNING, CAUTION, AND ADVISORY LIGHTS.

a. Explanation .

(1) Cockpit devices are color coded to symbolically represent various functions and varying levels of importance for flight crew operation. From early times, an attempt has been made to take full advantage of associations developed early in life as a result of continuous exposure to our daily environment.

(2) Military design specifications were the first to reference color-coding in cockpit design requirements. In the mid-1940s, the CAA initiated the first color-coding requirements for civil cockpit design. Color-coding standards for cockpit visual signals soon followed. MIL-STD-411, May 31, 1957, identified three separate categories of light signals: (i) Warning Light - indicates the existence of a hazardous condition which may require immediate corrective action.

(ii) Caution Light - serves to alert the operator to an impending dangerous condition requiring attention but not essential equipment, or attracts attention for routine purposes.

(iii) Advisory Light - indicates safe or normal configuration, condition of performance, or operation of essential equipment, or attracts attention for routine purposes.

(3) Examples of warning and caution signals were included in later versions of the military standard, and a few of those are shown below: [Section AC 29.1322 continued on next page.} Page F – 44.2 9/30/99 AC 29-2C Warning Signals Caution Signals Cabin Pressure Failure Trim Failure Fire Fuel Low Fuel System Failure Generator Inoperative Landing Gear Unsafe Defrosting Failure (4) Specific color designation for civil advisory lights was first addressed in Amendment 3 to the Rotorcraft Certification Rules (Parts 27 and 29) on January 19, 1968, with adoption of new §§ 27.1322 and 29.1322.

(5) In subsequent revision (Amendment 29-12), green lights were redesignated and additional colors introduced for flexibility in the requirement.

(6) Green signifies a safe operating condition and more specifically has come to signify landing gear extended and locked. Extensive use of green annunciators throughout the cockpit should generally be avoided due to possible confusion with the special use of green for landing gear. If green annunciators are physically and functionally removed from the landing gear operation, they may be found acceptable for a variety of “safe operating” applications. One such application is “all green for approach,” used in autopilot, flight director, and other navigation system displays.

(7) Other colors may be utilized as advisory lights in accordance with § 29.1322(d). Red and amber must not be used as advisory lights due to the possibility of introducing confusion into the cockpit. Obviously, yellow and pink annunciators should be avoided due to their similarity to amber and red. White and blue have been successfully utilized as advisory segments in past civil designs.

(8) The primary test for designation of color is: (i) Red - Is immediate action required: (ii) Amber - Is pilot action (other than immediate) required?

(iii) Green - Is safe operation indicated, and is the indication sufficiently distinct to prevent confusion with the landing gear down indication?

(iv) Other advisory lights - Is the meaning clear and distinct enough to prevent confusion with other annunciators? Do the colors which are utilized differ sufficiently from the colors specified in paragraphs (b)(1), (2), and (3) above?

(9) Annunciator lights should be visible during bright daylight conditions. This should include visibility in direct sunlight unless lights are located in such a manner that direct sunlight cannot impinge on them.

Page F - 45 AC 29-2C 9/30/99 (10) If dimming capability is provided, all annunciators, including master warning and caution, may be dimmable so long as the annunciation is clearly discernible for night operation at the lower lighting level. Undimmed annunciations have been found unacceptable for night operation due to disruption of cockpit vision at the high intensity. The dimming circuit should automatically revert to the high intensity setting when power is removed. Automatic dimming/brightening through the use of a photo cell is also acceptable, as are circuits which enable a dimming switch through a position light or other cockpit lighting controls.

(11) The use of flashing lights should be minimized. If a flashing feature is used, it should be controllable through pilot action so that flashing annunciation does not persist indefinitely. The indicator should be so designed that if it is energized and the flasher device fails, the light will illuminate and burn steadily.

(12) The activation of caution and warning lights should readily attract the attention of the appropriate crewmember while performing duties under both normal and high workload conditions.

b. Procedures .

(1) Red shall be reserved for annunciation of emergency conditions requiring immediate corrective action. Typical examples include fire, transmission oil pressure, engine failure, and battery overheat. The use of red for annunciators which do not require immediate action must be avoided. Use of red when it is not needed tends to lessen the impact of a red annunciator and the needed pilot association for immediate action. In evaluating cockpit annunciators for acceptability, the FAA/AUTHORITY should assure all annunciators which require immediate action are red and that only those requiring such action are red. If a master warning light is provided, it should be red, and should be powered by the same signal that powers any of the individual red warning signals. An aural warning may accompany visual warning signals to enhance pilot response. Care should be taken that any aural signal is sufficiently distinct from other aural warnings, such as low rotor RPM, to prevent confusion and to assure proper crew response. A means to deactivate and reset the master warning (visual and aural) is required. Resetting the master warning must not deactivate any individual warning signal.

(2) Amber shall be reserved for indicating malfunction or failure conditions which do not require immediate crew action to assure safe flight. Typical examples include door unlatched, inverter failure, generator failure, fuel filter clogged, and parking brake engaged. Amber should generally be utilized for malfunction and failure conditions which do not require immediate action. The key word here is “require.” Obviously, a pilot should perform corrective action for malfunction or failure conditions in a timely manner as soon as other cockpit priorities allow. The time increment associated with “immediate action” may vary with the system involved, the flight regime, and the aircraft; however, 15 seconds is a representative value in evaluating this term.

This by no means indicates that any red annunciator can be ignored for 15 seconds.

Page F - 46 9/30/99 AC 29-2C For red annunciators, some type of immediate pilot response is expected. If immediate pilot action is not required, the FAA/AUTHORITY should recommend the use of an amber designation. If a master caution light is provided in addition to a master warning light, the master caution annunciator should be amber, and should be powered by the same signal that powers any of the individual amber caution signals. Reset considerations for the master caution are the same as those detailed above for the master warning.

c. Annunciator Panel Design (1) Explanation.

(i) The annunciator panel design should be reviewed for the presence of failure modes that can cause illumination of multiple panel segments.

(ii) Many test circuits that are diode isolated are vulnerable to this condition. A typical sequence begins with the shorting of a test circuit diode. This failure is undetectable and goes unnoticed until an actual failure condition occurs which causes the associated panel segment to illuminate. At this time all panel segments connected to the test circuit will illuminate.

(iii) This configuration becomes a special problem when one or more of the panel segments are red. A red light calls for immediate action by the crew, and the crew does not have adequate information for immediate action when many false panel segments are illuminated.

(iv) If the design review indicates a problem, a redesign of the panel to eliminate the condition is considered to be the best solution and is highly encouraged.

(2) Procedures.

(i) An alternative to panel redesign might be the following: (A) Review the annunciator panel design and note which segments are red.

(B) Determine if cross reference information is available in the cockpit to allow elimination from consideration of any the red segments. (Example: Red low fuel pressure light and low fuel pressure gauge. Normal operation of the gauge would be a reason to assume the light did not cause the problem.)

(C) Where a cross reference is available, further design review of that function is not necessary; however, it may be appropriate to include procedural information in the emergency procedures section of the rotorcraft flight manual.

Page F - 47 AC 29-2C 9/30/99 (D) If cross references are not available for red segments, additional isolation should be incorporated into the annunciator design for those functions.

(ii) If cross referencing is not practical the following approach is encouraged.

(A) Review the annunciator panel design and note which segments are red.

(B) Determine if isolation diodes are checked during the application of battery or external power before starting the engines. (Example: Red low oil pressure light. If isolation diode is shorted, all panel segments will light as soon as battery or external power is applied.)

(C) When the isolation diode can be checked before starting engines, further design review is not necessary.

(D) If diodes are not automatically checked before starting, then additional isolation, should be considered.

(3) Annunciator Panel Arrangement . The annunciator panels should be arranged in a logical manner to reduce the crew’s time required to locate faults and to increase their efficiency in following Aircraft Flight Manual procedures. For example, engine annunciators on multiengine rotorcraft should be physically located on the panel to coincide with engine location (left or right) so that properly operating engines are not inadvertently shut down due to crew confusion over which engine has malfunctioned.

AC 29.1323. § 29.1323 (Amendment 29-3) AIRSPEED INDICATING SYSTEM .

a. Explanation .

(1) The accuracy of all flight test data concerned with the velocity of the rotorcraft is dependent on the calibration of the airspeed indicating system. For this reason, the airspeed system position error should be determined very early in the program.

(2) Since air density varies with altitude, the speed reading will only be correct under standard sea level conditions. However, in an actual installation, the indicator reading, even under standard sea level conditions, may differ from the calibrated airspeed because the static system does not sense true static pressure. This error in detection of static pressure is called position error. It is caused by the pressure field built up around the rotorcraft in flight. This pressure field will vary in intensity with dynamic pressure making the position error a function of calibrated airspeed. Since airspeed information is presented to the crew in terms of indicated airspeed, it is necessary to determine the position error for the rotorcraft to be flown safely.

Page F - 48 9/30/99 AC 29-2C b. Procedures .

(1) There are different methods to determine position error such as trailing bomb, airspeed course, boom system, and so forth. Each method has its own advantages and disadvantages, but will yield satisfactory results if done correctly. The airspeed system should be calibrated throughout the airspeed range of the rotorcraft and under the various flight conditions of cruise, climb, and autorotation standard. In addition, the effects of gross weight and center of gravity should be investigated.

(2) It may also be necessary to recalibrate the system with a change in external configuration if such a change may affect the airflow near the pitot or static sources.

(3) Additional information regarding position error is included in AC 29 Appendix B b(10) and should be considered if pursuing an IFR approval.

(4) Static system installation information is included in paragraph AC 29.1325.

Technical Standard Order (TSO) C16, Airspeed Tubes (Heated), gives minimum performance standards for pitot tubes, and pitot tubes qualified to this TSO normally allow for a satisfactory aircraft installation.

(5) The calibration requirements of the standard seem to be self-explanatory and are not discussed further in this paragraph.

AC 29.1323A. § 29.1323 (Amendment 29–24) AIRSPEED INDICATING SYSTEM .

a. Explanation . Amendment 29-24 to the regulations provides the requirements for Category A and Category B and defines the maximum allowable error for both.

b. Procedures . All of the policy material pertaining to this section remains in effect. In addition, calibration should be determined in level flight speeds of 20 knots and greater, and over an appropriate range of speeds for flight conditions of climb and autorotation; and takeoff. The takeoff calibration should be repeatable with respect to field lengths defined in the flight manual and avoidance of height-speed limiting envelope defined in § 29.79. Calibration errors, excluding instrument errors, may not exceed the following: (1) Category A - Three percent or 5 knots, whichever is greater, in level flight at speeds above 80 percent of takeoff safety speed; and 10 knots in climb at speeds from 10 knots below takeoff safety speed to 10 knots above V .

Y (2) Category B - Three percent or 5 knots, whichever is greater, in level flight at speeds above 80 percent of the climb-out speed attained at 50 feet when complying with § 29.63.

Page F - 49 AC 29-2C 9/30/99 AC 29.1325. § 29.1325 (Amendment 29-24) STATIC PRESSURE SYSTEMS .

a. Explanation .

(1) This section, in conjunction with § 29.1323, provides minimum performance standards for static pressure systems. The standard provides some relief when considering the icing environmental condition in that it allows the use of an alternate static port to account for the icing condition.

(2) The standard for the consideration of environmental conditions is § 29.1309(a).

(3) The standard for consideration of malfunction conditions is § 29.1309(b).

(4) For rotorcraft that will be approved for IFR operation, the provisions of Appendix B VIII(b)(5) of Part 29 as discussed in paragraph AC 29 Appendix B, should also be considered.

b. Procedures . The installation of the static system should consider the following: (1) Static lines should be initially routed upward immediately behind the static pressure port. This procedure will minimize the entry of moisture into the system when operating in rain or washing the rotorcraft.

(2) Drain(s) should be located at low points in the system. Line routing and clamping should allow for all moisture that does enter the system to be routed to the drain(s).

(3) If independent systems are provided, the placement of each system component should allow for maximum practicable separation of each system. As much as possible, one system should be on one side of the rotorcraft and the second system on the opposite side.

(4) Most static pressure ports that are provided for IFR operation are heated.

Before any tests are conducted, a program to qualify the heater on the port should normally be agreed upon through discussions between the FAA/AUTHORITY and the applicant. It is suggested that the requirements of TSO C16, Airspeed Tubes (Heated), be used as a guide for these discussions. If the ports are not to be heated, a comprehensive analysis should be prepared, and limited testing should be conducted to verify the analysis.

(5) Other static system considerations are included in paragraphs AC 29.1323 and AC 29 Appendix B.

Page F - 50 5/1/2014 AC 29-2C, Chg 4 AC 29.1327. § 29.1327 MAGNETIC DIRECTION INDICATOR.

a. Background. This section contains specific requirements regarding installation and functioning of a magnetic direction indicator. The magnetic direction indicator (commonly referred to as a compass) described by this paragraph is the unit required by § 29.1303(c) or the unit or system required for IFR operation by Appendix B VIII(a) to Part 29. Both of these indicators provide the pilot with an aircraft heading which is referenced to the earth’s magnetic field. The unit required by § 29.1303(c) is the indicator commonly referred to as a “whiskey compass.” This unit was given this designation because early units were constructed using alcohol as the medium in which the compass ball floats. This unit is generally approved as meeting the requirements of TSO-C7c. The indicator required by Appendix B to Part 29 is usually a system of units which meets the requirements of TSO-C6c.

b. Procedures. In showing compliance to § 29.1327(a), generally the magnetic indicator and its respective components will be tested to an appropriate standard such as RTCA DO 160B for use in a rotorcraft. If the unit functioned properly as described in the TSO during this testing, then no additional evaluation is generally required concerning vibration immunity. To determine the immunity of the indicator (system) from magnetic effects and its installed accuracy, a ground and flight test should be performed. This test should turn the rotorcraft a full 360° heading change in 45° increments. The indicator should not have an error in excess of 10° on any of the 45° increments. When performing these tests, the electrical equipment and systems should be functioning normally, and the effect of windshield heating (if installed) should be investigated. The results of the investigation may be used to construct the calibration placard which is required by § 29.1547. It should be noted that a calibration placard has not been traditionally required for slaved compass systems. Also, it should be emphasized that other aspects of the functioning and installation of these indicators should comply with the other general requirements (i.e., §§ 29.1301, 29.1309, 29.1555, etc.).

AC 29.1329. § 29.1329 (Amendment 29-24) AUTOMATIC PILOT SYSTEMS.

a. Explanation. The automatic flight control systems used on most modern rotorcraft often perform two different and distinct functions when viewed from a regulatory compliance aspect. These two functions are augmentation of the stability of the rotorcraft and a pilot aid in maintaining attitude, altitude, and airspeed, or in radio navigation tasks. The first function of stability is not covered in § 29.1329 but is included under § 29.672. The second function as a pilot aid is the automatic pilot function covered by this section. The following procedure discusses only those parts or systems which are installed as a pilot aid. MG 3 of this AC discusses the use of automatic systems for Category II approaches, and AC 29-2 Appendix B discusses the evaluation of stability augmentation systems.

b. Procedures.

Page F - 51 AC 29-2C, Chg 4 5/1/2014 (1) General.

(i) The automatic pilot system should be evaluated to demonstrate that it can perform its intended function of flying the rotorcraft and that it complies with the installation, operation, and malfunction requirements of § 29.1329. In demonstrating malfunctions of the autopilot system, generally servo actuator hardovers are the most critical malfunction. If this is the case and the autopilot system utilizes the same servos and servo amplifiers as the stability augmentation system (SAS) and the autopilot function cannot produce a more severe hardover than the SAS, then no additional consideration is required for this malfunction. An evaluation using the guidance in Appendix B of this AC would be sufficient.

(ii) There have been autopilots approved that require the use of a monitor since they cannot meet the hardover malfunction requirements. These approvals have involved a finding of equivalent safety that is beyond the scope of this guidance. Such findings of equivalent safety are made on a case-by-case basis. If an applicant is considering such a design, the applicant and the approving office should contact the Rotorcraft Standards Staff specialists for guidance.

(iii) The rule specifies that unless there is automatic synchronization, there should be some method to indicate the alignment of the actuating device to the pilot.

The intent of this requirement is to provide a means such that the pilot does not inadvertently engage the system into a hardover condition. One method of achieving this has been the use of servo force meters. These meters monitor the current into the servo motor and indicate to the pilot if a signal is being sent to the servo prior to system engagement.

(iv) Various autopilot systems have used a preflight test to ensure adequate reliability. The question that often arises is: Should the preflight test function be interlocked so the autopilot cannot be engaged if the preflight test has not been accomplished? The guidance used in the past to answer this question is: If the preflight test is simple and rapid enough that the pilot may reasonably be expected to perform such a test, then it is not required to be interlocked. If, however, the preflight test is very complicated and lengthy and a pilot who was pressed by a schedule might skip such a test, then this preflight test should be interlocked.

(v) Most of the autopilots that have been approved utilize series actuators or servos such as those required for a SAS. However, this does not preclude the approval of an autopilot that uses outer loop parallel actuation. This type of autopilot may be particularly helpful in a VFR aircraft.

(2) Cockpit controls . Evaluation of the cockpit controls should include the following items: (i) Location of the automatic pilot system controls are such that their operation is properly labeled and is readily accessible to the pilot(s).

Page F - 52 5/1/2014 AC 29-2C, Chg 4 (ii) Annunciator colors conform to the colors specified in § 29.1322 (reference section 29.1322 of this AC).

(iii) A determination is made that the controls, control labels, and placards are readable and discernible under all expected cockpit lighting conditions.

(iv) Motion and effect of the autopilot cockpit controls should conform with the requirements of § 29.779.

(v) Any disconnect of the autopilot should be annunciated.

c. Malfunction Evaluations . To preclude hazardous conditions that may result from any failure or malfunctioning of the autopilot the following failures should be evaluated.

This evaluation should also account for any hazards that also might be caused by inadvertent pilot action. The guidance in Appendix B of this AC should be used to determine the appropriate reaction times of the human pilot to an autopilot malfunction.

(1) Climb, cruise, and descent flight regimes .

(i) Recovery from malfunctions should simulate instrument conditions, or visual flight conditions, depending on the category of certification that is sought.

Justification should be provided for recognition (e.g., audio or visual warning, excess deviation alert, or acceleration cues in the case of a hardover). Continuous close monitoring of the flight attitude instruments by the pilot may not be relied upon as a reliable means for detecting low rate attitude deviations (typically < 3°/sec) and thus for determining the point at which slowover recognition occurs. In such cases, analysis should be employed to establish the recognition criteria for the particular helicopter and flight phase, and the acceptability of the recovery. For each flight regime, the maximum height loss recorded for all malfunction testing should be established. The applicant should ensure that sufficient data is generated to substantiate a height loss figure that can be used for an operational determination of a minimum use height, where appropriate.

(A) For cruise, the height loss is defined as the difference between the aircraft altitude at the time the failure is introduced, and the minimum altitude achieved in the recovery, taking into account appropriate pilot delays as discussed above.

(B) For a descent, the height loss should be determined as illustrated in Figure AC 29.1329-1. The evaluation should consider the maximum rate of descent approved for hands-off, IFR, operation.

(C) For approach without vertical guidance, the height loss should also be determined as illustrated in Figure AC 29.1329-1, but with consideration to the critical approach angle and nominal approach speed.

Page F - 53 AC 29-2C, Chg 4 5/1/2014 (D) For approach with vertical guidance, the height loss should be determined as discussed in paragraph (d) below.

(ii) The more critical of the following should be induced into the automatic pilot system.

(A) A signal about any axis equivalent to the cumulative effect of any single failure, including autotrim (if installed).

(B) The combined signals about all affected axes, if multiple axes failures can result from the malfunction of any single component.

1. Fa ilure Initiati on 2. Failure Recogn ition by pi lot 3. Init iation of Manual Recovery action by pil ot .

. ....

' · .

' · ....

....... .

....

·, .

. ·, ....

·< ....

....... .

·, ....

·, ....... ....

· -....: ....

....... .

·, ....

He i ght Loss ·, _. ....

....... . .

....... ....

....

- - ....

....

....

....

....

....

....

....

FIGURE AC 29.1329-1 HEIGHT LOSS METHOD Limit Loads . The simulated failure and the subsequent corrective action (2) should n ot create loads in excess of structural limits or result in dangerous dynamic conditions or hazardous deviations from the flight path. Additional guidance regarding the method of determining pilot recognition times and reasonable flight path deviation due to these simulated failures is contained in paragraph b.(6) of Appendix B of this AC.

Resultant flight loads outside the envelope of zero to 2g will be acceptable provided adequate analysis and flight test measurements are conducted to establish that no resultant aircraft load is beyond limit loads for the structure, including a critical assessment and consideration of the effects of structural loading parameter variations Page F - 54 5/1/2014 AC 29-2C, Chg 4 (i.e., center of gravity, load distribution, control system variations, maneuvering gradients, etc.). Analysis alone may be used to establish that limit loads are not exceeded where the aircraft loads are in the linear range of loading (i.e., aerodynamic coefficients for the flight condition are adequately established and no significant nonlinear air loadings exist). If significant nonlinear effects could exist, flight load survey measurements may be necessary to substantiate that the limit loads are not exceeded.

The power for climb should be the most critical of: (1) that used in the performance climb demonstrations; (2) that used in the longitudinal stability tests; or (3) that actually used for operational climb speeds.

(3) Maneuvering Flight . Malfunctions should also be induced into the automatic pilot system similar to paragraph c.(1) above. When corrective action is taken, the resultant loads and speeds should not exceed the values contained in paragraph c.(2) above. Maneuvering flight tests should include turns with the malfunction induced when maximum bank angles for normal operation of the system have been established and in the critical aircraft configuration and/or stages of flight likely to be encountered when using the automatic pilot. The altitude loss should be measured.

(4) Oscillatory Tests .

(i) An investigation should be made to determine the effects of an oscillatory signal of sufficient amplitude to saturate the servo amplifier of each device that can move a control. The investigation should cover the range of frequencies that can be induced by a malfunction of the automatic pilot system and systems functionally connected to it, including an open circuit in a feedback loop.

(ii) The results of this investigation should show that the peak loads imposed on the parts of the aircraft by the application of the oscillatory signal are within the limit loads for these parts.

(iii) The investigation may be accomplished largely through analysis with sufficient flight data to verify the analytical studies or largely through flight tests with analytical studies extending the flight data to the conditions which impose the highest percentage of limit load to the parts.

(iv) When flight tests are conducted in which the signal frequency is continuously swept through a range, the rate of frequency change should be slow enough to permit determining the amplitude of response of any part under steady frequency oscillation at any critical frequency within the test range.

(5) Recovery of Flight Control . To aid in recovery of the rotorcraft, after a malfunction occurs, one pilot should be able to physically overpower and disengage the autopilot with ease, and the autopilot should remain disengaged until further pilot action to reengage. The control to disconnect the autopilot should be easily available to the pilot who is now resisting the malfunctioning force of the autopilot. It is recommended that the disconnect button be placed on the cyclic control. It should be red and Page F - 55 AC 29-2C, Chg 4 5/1/2014 conspicuously marked “Autopilot Disconnect.” The pilot should be able to return the rotorcraft to its normal flight attitude under full manual control without exceeding the loads or speed limits defined in paragraph c.(2) above and without engaging in any dangerous maneuvers during recovery. The maximum servo authority used for these tests should not exceed those values shown to be within the structural limits for which the rotorcraft was designed. The maximum altitude loss experienced during these tests should be measured.

(6) External Interfaces. The autopilot system should have appropriate interlocks to its engagement to ensure it does not operate improperly as a result of information furnished by an external device or system. An example of this is the navigation receivers and the compass system. If for a particular mode of operation the autopilot uses signals from these systems, the autopilot should be interlocked from operating in those modes if invalid information is being received from that system.

d. Automatic Pilot Instrument Approach Approval.

(1) Throughout an approach, no signal or combination of signals simulating the cumulative effect of any single failure or malfunction in the automatic pilot system, except vertical gyro mechanical failures, should provide hazardous deviations from flight path or any degree of loss of control.

(2) The aircraft should be flown down the instrument landing system (ILS) in the configuration and at the approach speed specified by the applicant for approach.

Simulated autopilot malfunctions should be induced at critical points along the ILS, taking into consideration all possible variations in autopilot sensitivity and authority. The malfunctions should be induced in each axis. While the pilot may know the purpose of the flight, the pilot should not be informed when a malfunction is about to be or has been applied except through aircraft action, control movement, or other acceptable warning devices.

(3) An engine failure during an automatic ILS approach should not cause a lateral deviation of the aircraft from the flight path at a rate greater than 3° per second or produce hazardous attitudes.

(4) If approval is sought for ILS approaches initiated with one engine inoperative, the automatic pilot should be capable of conducting the approach.

(5) Deviations from the ILS flight profile should be evaluated as follows: (i) The rotorcraft should be instrumented so the following information is recorded–– (A) The path of the rotorcraft with respect to the normal glide path; Page F - 56 5/1/2014 AC 29-2C, Chg 4 (B) The point along the glide path when the simulated malfunction is induced; (C) The point where the pilot indicates recognition of the malfunction; and (D) The point along the path of the rotorcraft where recovery action is initiated.

(ii) Data obtained from the point of the indicated malfunction to the point where the rotorcraft has either again intersected the glide slope or is in level flight will define the deviation profile. When changes to the aircraft autopilot configuration are made during the approach and these changes alter the deviation profile, additional data should be obtained to define each of the applicable deviation profiles. An example of a deviation profile is shown in Figure AC 29.1329-2.

Page F - 57 AC 29-2C, Chg 4 5/1/2014 be er may Automatic ionkiov Provided) at . (rf er l 1ay configuration ece th.

Projected D Mode : No.

ull'f pa lransition R nce on orcrafl i .

are ol glide Fl rolorcraft nsta i lo : R of marker the le For autt F Configurat .

nt i ailong midd change Cllange .- po or of of ieNo r ~atbn ft.

De

one ~

changes Point Poinl an 200 lh -Made E: he T Gain T NO more 1.

2.

.

□ Na.

N , un i or ....

e tion a ~ Rotorcrafl Devial ProfiJe Config;uralfon

"

Chang ...

\ of Fault Configur olorcraft \ Point R .- ..,,,.

. 1 No _,..,,.

otorcraft No.

R in ....

a lt rr Deviation

' rofile

Te P ConfiBuration Fau

Z

-- FIGURE AC 29.1329-2 DEVIATION PROFILE Page F - 58 5/1/2014 AC 29-2C, Chg 4 (iii) Recoveries from malfunctions should simulate under-the-hood instrument conditions with an appropriate time delay between pilot recognition of the fault and initiation of the recovery at all altitudes down to 80 percent of the minimum decision altitude for which the applicant requests approval.

(iv) Recoveries from malfunctions at altitudes between 80 percent of the minimum decision altitude for which the applicant requests approval and the minimum altitude for which the applicant requests approval to operate the autopilot may be visual with no time delay between pilot recognition of fault and initiation of recovery.

(v) The minimum altitude at which the autopilot may be used should be determined as the altitude that results in the critical deviation profile becoming tangent with a minimum operational tolerance line. An example of this may be found in Figure AC 29.1329-3. The 29:1 slope of the minimum operational tolerance line provides a 1 percent gradient factor of safety over the 50:1 obstacle clearance line. An additional factor of safety is provided by measuring the 29:1 slope from the horizontal at a point 15 feet above the runway threshold. It is recognized that this minimum altitude will vary with glide slope angle. Information regarding these variations should be obtained and presented.

Page F - 59 AC 29-2C, Chg 4 5/1/2014

' ' . '. '.. · g~

er _ 1cal F It . P. n ~ M m

fo O per • of ·

.a,~ .

. '

.

' FIGURE AC 29.1329-3 DEVIATION PROFILE (vi) A malfunction of the autopilot during a coupled ILS approach should not place the aircraft in an attitude that would preclude conducting a satisfactory go-around or landing.

e. . The automatic pilot system should be installed and adjusted so Servo Authority the syst em tolerances established during certification tests can be maintained in normal operation. This may be ensured by conducting flight tests at the extremes of the tolerances. Those tests conducted to determine that the automatic pilot system will adequately control the aircraft should establish the lower limit. Those tests to determine that the automatic pilot will not impose dangerous loads or deviation from the flight path should be conducted at the upper limit. Appropriate aircraft loadings to produce the critical results should be used.

Page F - 60 5/1/2014 AC 29-2C, Chg 4 f. Rotorcraft Flight Manual Information. The following information should be placed in the rotorcraft flight manual: (1) In the Operating Limitations Section: Airspeed, Maximum Height Loss following AFCS malfunction for each phase of flight, Minimum Use Height where appropriate, and other applicable operating limitations for use of the autopilot.

(2) In the Operating Procedures Section: Normal operation information.

(3) In the Emergency Operation Procedures Section: (i) A statement of the downward flight path deviation in the cruise, climb, and descent configurations and the maneuvering flight configuration in accordance with paragraphs d(5)(iii) and d(5)(iv) above, if this deviation exceeds 100 feet.

(ii) True profiles of deviations below the glide slope or projected flare path for the critical conditions tested (see paragraphs d.(5)(iv) above and Figure AC 29.1329-2) and the deviation profile indicating the lowest altitude at which the autopilot can be used(see paragraph d.(5)(v) above, if applicable, and if this deviation exceeds 100 feet or excessive deviation for an ILS approach.

g. There should be a means of sequencing actions or interlocking engagement with sensor inputs to prevent autopilot initiated maneuvers that could result in hazardous operations due to: (1) Engagement of the autopilot; (2) Malfunctions of autopilot input or feedback signals that could result in unbounded output commands.

AC 29.1329A. § 29.1329 (Amendment 29-42) AUTOMATIC PILOT SYSTEM.

a. Explanation. Amendment 29-42 adds paragraph (f) to § 29.1329, which requires a means of indicating to the pilots the current mode of operation for those automatic pilot systems that can be coupled to the airborne navigation equipment.

b. Procedures. The policy material pertaining to section 29.1329 of this AC remains in effect with the following additions: (1) Mode annunciation must indicate the state of the system including mode change and disengagement. Mode annunciation should be presented in a manner compatible with flightcrew procedures and tasks and be consistent with other flight deck systems’ mode annunciations. Mode selector switch position or status is not acceptable as the sole means of mode annunciation. Modes and mode changes should be depicted in a manner that achieves flightcrew attention and awareness.

Page F – 60.1 AC 29-2C, Chg 4 5/1/2014 (2) Mode annunciations must effectively and unambiguously indicate the active and armed modes of operation. The mode annunciation should convey explicitly, and as simply as possible, what the system is doing (for active modes), what it will be doing (for armed modes), and target information (such as selected speed, heading, and altitude) for satisfactory flightcrew awareness. The pilot must be alerted to any deviation from the pilot-commanded target.

Page F – 60.2 6/23/23 AC 29 - 2C, Chg 9 AC 29.1329B. § 29.1329 (AMENDMENT 29 - 59 ) AUTOMATIC PILOT AN D FLIGHT GUIDANCE SYSTEM a. Explanation. Amendment 29-59 changed the section heading name to “Automatic pilot and flight guidance system” and combined the requirements for automatic pilot (14 CFR 29.1329) and flight director systems (14 CFR 29.1335) into one rule (14 CFR 29.1329). The amendment also removed 14 CFR 29.1335.

b. Procedures. This section of the AC is updated to include the guidance previously found in AC 29-2C Section 29.1335. Therefore, the guidance pertaining to section 29.1329 of this AC remains in effect with the following additions: (1) Three-cue flight directors for rotorcraft use the usual pitch and roll command cues with the third cue displayed on the left side of the attitude director indicator (ADI). These instruments can be used in either the two-axes or three-axes modes. In either mode, the lateral command cue controls the roll attitude, and the vertical command cue controls the pitch attitude. The rotorcraft attitude, controlled by the cyclic control, is changed to satisfy the flight director commands. The third cue, when displayed, commands collective pitch position and is used when an airspeed or pitch attitude mode and a vertical mode (altitude hold, glide slope, etc.) are selected.

(2) The general convention for flight director design is that each command bar is a “fly to” command. The motion of the flight director indicator is such to command a corresponding sense of control system motion. This is true of flight director pitch and roll commands and should hold true for additional commands such as collective pitch.

(3) Some consideration should be given to the collective, or third cue, display.

For example, if the collective symbol is selected as the fixed index, the command cue and collective pitch control should move in opposite directions when collective pitch changes are made. This configuration would constitute a conventional “fly to” indicator.

If the collective symbol is selected for the movable index, the direction of motion of the collective symbol will coincide with the direction of collective pitch changes. In this case the moving collective symbol does not comply with the “fly to” convention; however, this configuration has been approved by the FAA/ with special symbology, special background effects, and special color coding and has performed satisfactorily in service.

(4) The recommended display for a three-cue flight director incorporates the standard pitch and roll command symbols: either pitch and roll bars or the “V” bar display. The third cue, or collective symbol, should be located on the left side of the ADI.

The shape of the moving cue and the background display should be unique to avoid it being confused with a glide slope display or angle of attack display. One display uses a Page F - 60.2a 6/23/23 AC 29 - 2C, Chg 9 third cue, shaped like a small handle, to aid in identifying it as the collective pitch symbol.

(5) The color of the pitch and roll command indicators, the aircraft symbol, the background marking of the third cue, and third cue itself should be consistent. The optimum color scheme uses the same color for the aircraft symbol and the collective symbol. This i s usually fire orange. The command cues including the collective cue also should use the same color, usually yellow. The rationale for the different colors is that the aircraft symbol and the collective symbol (the same color) are moved toward their respec tive command cues. If the pitch command cue is above the center, the aircraft symbol is raised (nose pulled up) and, if the collective command cue is above the collective symbol, the collective pitch is raised, moving the collective symbol towards the comm and cue.

(6) If the ADI provides a monochromatic display, the collective pitch cue and its background markings must, per 14 CFR 29.1321, be distinctive to reduce the chance of them being confused with the glide slope indicator. This can be accomplished through the use of different shaped cues and background marks. A round cue with a chevron - shaped background marking has been satisfactory.

Page F - 60.2b 5/1/2014 AC 29-2C, Chg 4 AC 29.1331. § 29.1331 (Amendment 29-24) INSTRUMENTS USING A POWER SUPPLY.

a. Explanation . The rule concerns each flight instrument using a power supply that is installed in a Category A rotorcraft. A reference to paragraph AC 29.1303 will give a listing of the flight instruments that are specifically required for certification. The discussion included in this paragraph is directed toward electrical instruments since these are the type normally installed in Category A rotorcraft. It should be noted, however, that the rule is not restricted to electrical instruments. Further, the discussion provided here can be used to evaluate non-electrical applications.

b. Procedures .

(1) This requirement must be considered when designing the electrical distribution system for Category A rotorcraft. It states that each required flight instrument must have two independent sources of power and a means of selecting either source. The flight instruments required for certification are listed in paragraph AC 29.1303, and independent power sources are discussed in paragraph AC 29.1355b(4).

(2) Some older flight instruments may not have integral visual means to indicate that adequate power is being supplied to the instrument as required by the rule. For these instruments, external annunciation has been accepted that monitors the presence of adequate voltage at the power pin on the electrical disconnect that mates with the electrical connector on the back side of the instrument. The annunciator light should be located in close proximity to the indicator and placarded to identify its function. Note that the rule requires the voltage monitored to be within the approved limits for the instrument to be adequate. Since relay coils normally operate well outside the approved instrument voltage limits, the use of a relay contact that closes when the monitored voltage drops low enough to pull in or release a relay coil does not normally result in a satisfactory design to meet the regulatory requirement. Annunciator lights provided for this application are normally red.

(3) The power supply system requirements of this rule should be coordinated with the requirements of § 29.1355 (see paragraph AC 29.1355). Both rules concern equipment or systems that require two independent sources of electrical power.

Examples of faults in the distribution system to be considered include open feeders, shorted feeders, shorted busses, etc.

(4) Amendment 29-24 revised the regulation to further clarify the power adequacy indication requirement. The clarification provided was intended to make it easier to understand the meaning of adequate power in the event it was necessary to provide separate annunciation. The application of the rule in each form (before and after Amendment 29-24) should be the same.

Page F – 60.3 AC 29-2C, Chg 4 5/1/2014 [This page intentionally blank.]

Page F – 60.4 5/1/2014 AC 29-2C, Chg 4 AC 29.1333. § 29.1333 (Amendment 29-24) INSTRUMENT SYSTEMS.

a. Explanation .

(1) Electromechanical Displays. Prior to Amendment 29-24, this requirement was intended to apply to duplicate flight instruments required by any operating rule.

Due to the increased complexity of instrumentation that was available and being used, it was considered appropriate to amend the provisions of this requirement by adopting § 29.1333 in Amendment 29-24 to more appropriately consider the extreme range of operational environments to which rotorcraft were being routinely exposed. It is the intent of § 29.1333 to prevent degrading the first pilot’s instrument system, or the only pilot’s instrument system in a single-pilot-approved rotorcraft, by not permitting peripheral systems to be connected to it. In addition, equipment must not be connected to operating systems for the second pilot’s required instruments unless it is extremely improbable that failure of such additional equipment would affect that operating system.

Similar provisions are also included in VIII.(b)(5)(i) through (iii) of Appendix B to Part 29, Airworthiness Criteria for Helicopter Instrument Flight.

(2) Advanced Display Systems. The increased use of microprocessor technology in avionic systems has resulted in the use of computer-generated graphics to replace conventional electromechanical instruments. These displays may replace individual instruments or may integrate several flight critical parameters into single displays. For display of redundant information, “crosstalk” between the pilot and copilot displays and supporting systems has been allowed to provide detection and annunciation of faults or “miscompare” of critical flight information. A level of safety finding equivalent to that level of safety provided by § 29.1333 may be possible through the implementation of integration technology that will assure that failure of one system does not and can not adversely affect the other system.

b. Procedures.

(1) The provisions of the current rule are essentially self-explanatory.

(2) If the certification basis of the rotorcraft is prior to Amendment 29-24, the provisions are more precise; however, they only apply in the instance where duplicate instruments are required by the operating rules.

(3) If an IFR approval is part of the certification effort, then Part 29, Appendix B, applies, and the provisions of paragraph VIII(b)(5) are essentially the same as the current rule. If the certification basis of the rotorcraft is prior to Amendment 29-24, and an IFR approval is being added, the instrument systems should be carefully reviewed since their design may not have considered the provisions of the IFR rule.

Page F – 61 6/23/23 AC 29 - 2C, Chg 9 AC 29.1333A. § 29.1333 (AMENDMENT 29 - 59 ) INSTRUMENT SYSTEMS.

Explanation . Amendment 29 - 59 revised § 29.1333 by revising paragraph (a), which requires isolating the pilot instrument system from any other operating systems, to only apply to pneumatic systems. This change allows other types of pilot instrument systems to be connected to other systems on the rotorcraft and all ows the use of modern technology to monitor and display highly integrated information.

Page F - 61.1 AC 29-2C 9/30/99 AC 29.1335. § 29.1335 (Amendment 29-14) FLIGHT DIRECTOR SYSTEMS.

a. Explanation . This section prescribes the accepted display criteria for a rotorcraft three-cue flight director providing command guidance for pitch, roll, and power. Three-cue flight directors for rotorcraft use the usual pitch and roll command cues with the third cue displayed on the left side of the attitude director indicator (ADI).

These instruments can be used in either the two-axes or three-axes modes. In either mode, the lateral command cue controls the roll attitude, and the vertical command cue controls the pitch attitude. The rotorcraft attitude, controlled by the cyclic control, is changed to satisfy the flight director commands. The third cue, when displayed, commands collective pitch position and is used when an airspeed or pitch attitude mode and a vertical mode (altitude hold, glide slope, etc.) are selected.

(1) The general convention for flight director design is that each command bar is a “fly to” command. The motion of the flight director indicator is such to command a corresponding sense of control system motion. This is true of flight director pitch and roll commands and should hold true for additional commands such as collective pitch.

(2) Some consideration should be given to the collective, or third cue, display.

For example, if the collective symbol is selected as the fixed index, the command cue and collective pitch control should move in opposite directions when collective pitch changes are made. This configuration would constitute a conventional “fly to” indicator.

If the collective symbol is selected for the movable index, the direction of motion of the collective symbol will coincide with the direction of collective pitch changes. In this case the moving collective symbol does not comply with the “fly to” convention; however, this configuration has been approved by the FAA/AUTHORITY with special symbology, special background effects, and special color coding, and has performed satisfactorily in service.

b. Procedures . The recommended display for a three-cue flight director incorporates the standard pitch and roll command symbols, either pitch and roll bars or the “V” bar display. The third cue, or collective symbol, should be located on the left side of the ADI. The shape of the moving cue and the background display should be unique to avoid being confused with a glide slope display or angle of attack display.

One display uses a third cue, shaped like a small handle, to aid in identifying it as the collective pitch symbol.

(1) The color of the pitch and roll command indicators, the aircraft symbol, the background marking of the third cue, and third cue itself, should be consistent. The optimum color scheme uses the same color for the aircraft symbol and the collective symbol. This is usually fire orange. The command cues including the collective cue also should use the same color, usually yellow. The rationale for the different colors is that the aircraft symbol and the collective symbol (the same color) are moved toward their respective command cues. If the pitch command cue is above the center, the aircraft symbol is raised (nose pulled up) and, if the collective command cue is above Page F – 62 5/1/2014 AC 29-2C, Chg 4 the collective symbol, the collective pitch is raised, moving the collective symbol towards the command cue.

(2) If the attitude director indicator (ADI) provides a monochromatic display, the collective pitch cue and its background markings must be distinctive to reduce the chance of being confused with the glide slope indicator. This can be accomplished through the use of different shaped cues and background marks. A round cue with a chevron-shaped background marking has been satisfactory.

Page F – 63 6/23/23 AC 29 - 2C, Chg 9 AC 29.1335A. § 29.1335 (AMENDMENT 29 - 59 ) FLIGHT DIRECTOR SYSTEMS.

Explanation . Modern designs include both automatic pilot and flight director systems, referring to them now as automatic pilot and flight guidance system. Having these systems in separate rules that use different terminology has resulted in some confusion. Therefore, amendment 29 - 59 removed Section 29.1335 and incorporated its provisions into Section 29.1329.

Page F - 63.1

5/1/2014 AC 29-2C, Chg 4 AC 29.1337. § 29.1337 (Amendment 29-13) POWERPLANT INSTRUMENTS - (Paragraph (b) - FUEL QUANTITY INDICATOR).

a. Explanation . Section 29.1337(b) requires, in part, a means to indicate to the flight crew the quantity of usable fuel during flight in each tank. When two or more tanks are interconnected so that a failure of the system could cause fuel to become trapped in a fuel tank, the fuel quantity indicating system must provide the flight crew with the ability to determine the total effective amount of remaining usable fuel. Since the flight attitude of a rotorcraft may vary significantly with center of gravity (CG) and airspeed, a standard attitude for calibration of the fuel quantity gauge is needed. In addition, guidelines for gauge accuracy and comments regarding other fuel quantity gauging aspects are offered.

b. Procedures .

(1) Determine the rotorcraft pitch attitudes for most forward and most aft CG at a median gross weight and at an airspeed of 0.9 V or 0.9 V , whichever is less. The NE H mean attitude of the extremes defined above, further adjusted for lateral CG effects, if necessary, define the rotorcraft attitude for fuel gauge calibration.

(2) After establishing the calibration attitude, the requirements of § 29.1337(b) can be accomplished. The aircraft should be placed in the calibration attitude. Add fuel to the filler neck spillover level. Defuel the aircraft in increments corresponding to fuel gauge increment markings or at least 10 increments until gauge zero is obtained.

Precautions should be taken during this step to be sure that the fuel transmitter is sensing fuel level and not simply reflecting a physical “STOP” or end point in the system range. The fuel remaining in the tank below the “ZERO” mark must not be less than that amount determined by flight testing under § 29.959. (Otherwise, the zero point must be adjusted upward.) The gauging system accuracy is acceptable when it meets a tolerance of ±2 percent of the total useable fuel plus ±4 percent of the remaining usable fuel at any gauge reading, provided that the gauge indicates zero fuel with unusable fuel in accordance with § 29.959 in the tank. (For a 100-gallon tank, this formula would allow a ±6-gallon error at the full level, ±4-gallon error at 50-gallon level, converging to a ±2-gallon error at low fuel with the further provision that the zero mark accurately reflects unusable fuel.)

Page F – 63 .2 AC 29-2C, Chg 4 5/1/2014 (3) Certain other aspects of a fuel gauging system need attention in order to minimize fuel exhaustion incidents: (i) Gauge reading with the aircraft at ground attitude is frequently used by the crew in calculating range, weight and balance, and actual gross weight. Significant gauge errors in either direction during this reading can introduce hazards to the operation of the aircraft. If a calibration at this attitude indicates an unconservative error in excess of 6 percent of the gauge reading, corrective information should be applied adjacent to the fuel quantity gauge or be made available to the crew in other handbook data.

(ii) Flight during hover with maximum rearward wind may introduce significantly different fuel gauge readings. A check should be made to assure that the gauge is either accurate or at least does not read high (unconservative) in this attitude.

(iii) Consistent with the requirements of § 29.1337(b)(2), a separate fuel quantity indication is necessary for any interconnected fuel tank that has a flow control device, such as a fuel transfer pump or flapper valve, which could fail and trap fuel.

This requirement also applies to auxiliary fuel tanks. A sight gauge that is readable by the flight crew in flight may be acceptable for use with auxiliary fuel tanks.

(4) Fuel gauging system transmitters which are strictly volumetric measuring devices (float-actuated variable rheostats) introduce a gauge readout error of about 5 percent if calibrated with a fuel temperature of 0°C (32°F) and subjected to -55°C (-67°F) fuel or +55°C (131°F) fuel. This error may be minimized by calibrating the gauge with fuel temperature in the middle of the useful range (i.e., 15°C (59°F)).