14 CFR Part 25 — Airworthiness Standards: Transport Category Airplanes

Federal Aviation Administration, DOT Pt. 25

(c) Diagrams of structural access plates Subpart B—Flight and information needed to gain access for in- G ENERAL spections when access plates are not pro- vided.

25.21 Proof of compliance.

(d) Details for the application of special in- 25.23 Load distribution limits.

spection techniques including radiographic 25.25 Weight limits.

and ultrasonic testing where such processes 25.27 Center of gravity limits.

are specified by the applicant.

25.29 Empty weight and corresponding cen- (e) Information needed to apply protective ter of gravity.

treatments to the structure after inspection.

25.31 Removable ballast.

(f) All data relative to structural fasteners 25.33 Propeller speed and pitch limits.

such as identification, discard recommenda- tions, and torque values.

P ERFORMANCE (g) A list of special tools needed.

25.101 General.

(h) In addition, for level 4 airplanes, the 25.103 Stall speed.

following information must be furnished— 25.105 Takeoff.

(1) Electrical loads applicable to the var- 25.107 Takeoff speeds.

ious systems; 25.109 Accelerate-stop distance.

(2) Methods of balancing control surfaces; 25.111 Takeoff path.

(3) Identification of primary and secondary 25.113 Takeoff distance and takeoff run.

structures; and 25.115 Takeoff flight path.

(4) Special repair methods applicable to 25.117 Climb: general.

the airplane.

25.119 Landing climb: All-engines-operating.

A23.4 Airworthiness limitations section. 25.121 Climb: One-engine-inoperative.

25.123 En route flight paths.

The Instructions for Continued Airworthi- 25.125 Landing.

ness must contain a section titled Airworthi- ness Limitations that is segregated and C ONTROLLABILITY AND M ANEUVERABILITY clearly distinguishable from the rest of the 25.143 General.

document. This section must set forth each 25.145 Longitudinal control.

mandatory replacement time, structural in- 25.147 Directional and lateral control.

spection interval, and related structural in- spection procedure required for type certifi- 25.149 Minimum control speed.

cation. If the Instructions for Continued Air- T RIM worthiness consist of multiple documents, the section required by this paragraph must 25.161 Trim.

be included in the principal manual. This section must contain a legible statement in S TABILITY a prominent location that reads ‘‘The Air- 25.171 General.

worthiness Limitations section is FAA ap- 25.173 Static longitudinal stability.

proved and specifies maintenance required 25.175 Demonstration of static longitudinal under §§ 43.16 and 91.403 of Title 14 of the Code stability.

of Federal Regulations unless an alternative 25.177 Static lateral-directional stability.

program has been FAA approved.’’ 25.181 Dynamic stability.

PART 25—AIRWORTHINESS STAND- S TALLS

ARDS: TRANSPORT CATEGORY

25.201 Stall demonstration.

AIRPLANES

25.203 Stall characteristics.

25.207 Stall warning.

S PECIAL F EDERAL A VIATION R EGULATION N O .

G ROUND AND W ATER H ANDLING C HARACTERISTICS S PECIAL F EDERAL A VIATION R EGULATION N O .

25.231 Longitudinal stability and control.

25.233 Directional stability and control.

25.235 Taxiing condition.

Subpart A—General 25.237 Wind velocities.

Sec.

25.239 Spray characteristics, control, and 25.1 Applicability. stability on water.

25.2 Special retroactive requirements.

M ISCELLANEOUS F LIGHT R EQUIREMENTS 25.3 Special provisions for ETOPS type de- sign approvals. 25.251 Vibration and buffeting.

25.4 Definitions.

25.253 High-speed characteristics.

25.5 Incorporations by reference. 25.255 Out-of-trim characteristics.

14 CFR Ch. I (1–1–25 Edition) Pt. 25

25.495 Turning.

Subpart C—Structure 25.497 Tail-wheel yawing.

G ENERAL 25.499 Nose-wheel yaw and steering.

25.503 Pivoting.

25.301 Loads.

25.507 Reversed braking.

25.302 Interaction of systems and struc- 25.509 Towing loads.

tures.

25.511 Ground load: unsymmetrical loads on 25.303 Factor of safety.

multiple-wheel units.

25.305 Strength and deformation.

25.519 Jacking and tie-down provisions.

25.307 Proof of structure.

WATER L OADS FLIGHT L OADS 25.521 General.

25.321 General.

25.523 Design weights and center of gravity positions.

F LIGHT M ANEUVER AND G UST C ONDITIONS 25.525 Application of loads.

25.331 Symmetric maneuvering conditions.

25.527 Hull and main float load factors.

25.333 Flight maneuvering envelope.

25.529 Hull and main float landing condi- 25.335 Design airspeeds.

tions.

25.337 Limit maneuvering load factors.

25.531 Hull and main float takeoff condi- 25.341 Gust and turbulence loads.

tion.

25.343 Design fuel and oil loads.

25.533 Hull and main float bottom pressures.

25.345 High lift devices.

25.535 Auxiliary float loads.

25.349 Rolling conditions.

25.537 Seawing loads.

25.351 Yaw maneuver conditions.

25.353 Rudder control reversal conditions.

E MERGENCY L ANDING C ONDITIONS 25.561 General.

S UPPLEMENTARY C ONDITIONS 25.562 Emergency landing dynamic condi- 25.361 Engine and auxiliary power unit tions.

torque.

25.563 Structural ditching provisions.

25.362 Engine failure loads.

25.363 Side load on engine and auxiliary F ATIGUE E VALUATION power unit mounts.

25.571 Damage-tolerance and fatigue evalua- 25.365 Pressurized compartment loads.

tion of structure.

25.367 Unsymmetrical loads due to engine failure.

L IGHTNING P ROTECTION 25.371 Gyroscopic loads.

25.581 Lightning protection.

25.373 Speed control devices.

CONTROL S URFACE AND S YSTEM L OADS Subpart D—Design and Construction 25.391 Control surface loads: General.

G ENERAL 25.393 Loads parallel to hinge line.

25.601 General.

25.395 Control system.

25.603 Materials.

25.397 Control system loads.

25.605 Fabrication methods.

25.399 Dual control system.

25.607 Fasteners.

25.405 Secondary control system.

25.609 Protection of structure.

25.407 Trim tab effects.

25.611 Accessibility provisions.

25.409 Tabs.

25.613 Material strength properties and ma- 25.415 Ground gust conditions.

terial design values.

25.427 Unsymmetrical loads.

25.619 Special factors.

25.445 Auxiliary aerodynamic surfaces.

25.621 Casting factors.

25.457 Wing flaps.

25.623 Bearing factors.

25.459 Special devices.

25.625 Fitting factors.

G ROUND L OADS 25.629 Aeroelastic stability requirements.

25.631 Bird strike damage.

25.471 General.

25.473 Landing load conditions and assump- C ONTROL S URFACES tions.

25.651 Proof of strength.

25.477 Landing gear arrangement.

25.655 Installation.

25.479 Level landing conditions.

25.657 Hinges.

25.481 Tail-down landing conditions.

25.483 One-gear landing conditions.

C ONTROL S YSTEMS 25.485 Side load conditions.

25.487 Rebound landing condition. 25.671 General.

25.489 Ground handling conditions. 25.672 Stability augmentation and auto- 25.491 Taxi, takeoff and landing roll. matic and power-operated systems.

25.493 Braked roll conditions. 25.675 Stops.

Federal Aviation Administration, DOT Pt. 25

25.677 Trim systems. 25.833 Combustion heating systems.

25.679 Control system gust locks.

P RESSURIZATION 25.681 Limit load static tests.

25.683 Operation tests.

25.841 Pressurized cabins.

25.685 Control system details.

25.843 Tests for pressurized cabins.

25.689 Cable systems.

25.693 Joints. F IRE P ROTECTION 25.697 Lift and drag devices, controls.

25.851 Fire extinguishers.

25.699 Lift and drag device indicator.

25.853 Compartment interiors.

25.701 Flap and slat interconnection.

25.854 Lavatory fire protection.

25.703 Takeoff warning system.

25.855 Cargo or baggage compartments.

25.856 Thermal/Acoustic insulation mate- L ANDING G EAR rials.

25.721 General.

25.857 Cargo compartment classification.

25.723 Shock absorption tests.

25.858 Cargo or baggage compartment 25.725–25.727 [Reserved] smoke or fire detection systems.

25.729 Retracting mechanism.

25.859 Combustion heater fire protection.

25.731 Wheels.

25.863 Flammable fluid fire protection.

25.733 Tires.

25.865 Fire protection of flight controls, en- 25.735 Brakes and braking systems.

gine mounts, and other flight structure.

25.737 Skis.

25.867 Fire protection: other components.

25.869 Fire protection: systems.

F LOATS AND H ULLS M ISCELLANEOUS 25.751 Main float buoyancy.

25.753 Main float design.

25.871 Leveling means.

25.755 Hulls.

25.875 Reinforcement near propellers.

25.899 Electrical bonding and protection P ERSONNEL AND C ARGO A CCOMMODATIONS against static electricity.

25.771 Pilot compartment.

25.772 Pilot compartment doors. Subpart E—Powerplant 25.773 Pilot compartment view.

G ENERAL 25.775 Windshields and windows.

25.777 Cockpit controls.

25.901 Installation.

25.779 Motion and effect of cockpit controls.

25.903 Engines.

25.781 Cockpit control knob shape.

25.904 Automatic takeoff thrust control sys- 25.783 Fuselage doors.

tem (ATTCS).

25.785 Seats, berths, safety belts, and har- 25.905 Propellers.

nesses.

25.907 Propeller vibration and fatigue.

25.787 Stowage compartments.

25.925 Propeller clearance.

25.789 Retention of items of mass in pas- 25.929 Propeller deicing.

senger and crew compartments and gal- 25.933 Reversing systems.

leys.

25.934 Turbojet engine thrust reverser sys- 25.791 Passenger information signs and plac- tem tests.

ards.

25.937 Turbopropeller-drag limiting sys- 25.793 Floor surfaces.

tems.

25.795 Security considerations.

25.939 Turbine engine operating characteris- tics.

MERGENCY P ROVISIONS E 25.941 Inlet, engine, and exhaust compat- 25.801 Ditching.

ibility.

25.803 Emergency evacuation. 25.943 Negative acceleration.

25.807 Emergency exits. 25.945 Thrust or power augmentation sys- 25.809 Emergency exit arrangement. tem.

25.810 Emergency egress assist means and F UEL S YSTEM escape routes.

25.811 Emergency exit marking.

25.951 General.

25.812 Emergency lighting.

25.952 Fuel system analysis and test.

25.813 Emergency exit access.

25.953 Fuel system independence.

25.815 Width of aisle.

25.954 Fuel system lightning protection.

25.817 Maximum number of seats abreast.

25.955 Fuel flow.

25.819 Lower deck service compartments 25.957 Flow between interconnected tanks.

(including galleys).

25.959 Unusable fuel supply.

25.820 Lavatory doors.

25.961 Fuel system hot weather operation.

25.963 Fuel tanks: general.

V ENTILATION AND H EATING 25.965 Fuel tank tests.

25.831 Ventilation. 25.967 Fuel tank installations.

25.832 Cabin ozone concentration. 25.969 Fuel tank expansion space.

14 CFR Ch. I (1–1–25 Edition) Pt. 25

25.971 Fuel tank sump. POWERPLANT F IRE P ROTECTION 25.973 Fuel tank filler connection.

25.1181 Designated fire zones; regions in- 25.975 Fuel tank vents and carburetor vapor cluded.

vents.

25.1182 Nacelle areas behind firewalls, and 25.977 Fuel tank outlet.

engine pod attaching structures con- 25.979 Pressure fueling system.

taining flammable fluid lines.

25.981 Fuel tank explosion prevention.

25.1183 Flammable fluid-carrying compo- nents.

F UEL S YSTEM C OMPONENTS 25.1185 Flammable fluids.

25.991 Fuel pumps.

25.1187 Drainage and ventilation of fire 25.993 Fuel system lines and fittings.

zones.

25.994 Fuel system components.

25.1189 Shutoff means.

25.995 Fuel valves.

25.1191 Firewalls.

25.997 Fuel strainer or filter.

25.1192 Engine accessory section diaphragm.

25.999 Fuel system drains.

25.1193 Cowling and nacelle skin.

25.1001 Fuel jettisoning system. 25.1195 Fire extinguishing systems.

25.1197 Fire extinguishing agents.

O IL S YSTEM 25.1199 Extinguishing agent containers.

25.1201 Fire extinguishing system materials.

25.1011 General.

25.1203 Fire detector system.

25.1013 Oil tanks.

25.1207 Compliance.

25.1015 Oil tank tests.

25.1017 Oil lines and fittings.

Subpart F—Equipment 25.1019 Oil strainer or filter.

25.1021 Oil system drains.

G ENERAL 25.1023 Oil radiators.

25.1301 Function and installation.

25.1025 Oil valves.

25.1302 Installed systems and equipment for 25.1027 Propeller feathering system.

use by the flightcrew.

COOLING 25.1303 Flight and navigation instruments.

25.1305 Powerplant instruments.

25.1041 General.

25.1307 Miscellaneous equipment.

25.1043 Cooling tests.

25.1309 Equipment, systems, and installa- 25.1045 Cooling test procedures.

tions.

25.1310 Power source capacity and distribu- I NDUCTION SYSTEM tion.

25.1091 Air induction.

25.1316 Electrical and electronic system 25.1093 Induction system icing protection.

lightning protection.

25.1101 Carburetor air preheater design.

25.1317 High-intensity Radiated Fields 25.1103 Induction system ducts and air duct (HIRF) Protection.

systems.

25.1105 Induction system screens. INSTRUMENTS : I NSTALLATION 25.1107 Inter-coolers and after-coolers.

25.1321 Arrangement and visibility.

25.1322 Flightcrew alerting.

EXHAUST S YSTEM 25.1323 Airspeed indicating system.

25.1121 General.

25.1324 Angle of attack system.

25.1123 Exhaust piping.

25.1325 Static pressure systems.

25.1125 Exhaust heat exchangers.

25.1326 Pitot heat indication systems.

25.1127 Exhaust driven turbo-superchargers.

25.1327 Magnetic direction indicator.

25.1329 Flight guidance system.

P OWERPLANT C ONTROLS AND A CCESSORIES 25.1331 Instruments using a power supply.

25.1333 Instrument systems.

25.1141 Powerplant controls: general.

25.1337 Powerplant instruments.

25.1142 Auxiliary power unit controls.

25.1143 Engine controls.

E LECTRICAL S YSTEMS AND E QUIPMENT 25.1145 Ignition switches.

25.1147 Mixture controls.

25.1351 General.

25.1149 Propeller speed and pitch controls.

25.1353 Electrical equipment and installa- 25.1153 Propeller feathering controls.

tions.

25.1155 Reverse thrust and propeller pitch 25.1355 Distribution system.

settings below the flight regime.

25.1357 Circuit protective devices.

25.1157 Carburetor air temperature controls.

25.1360 Precautions against injury.

25.1159 Supercharger controls. 25.1362 Electrical supplies for emergency 25.1161 Fuel jettisoning system controls. conditions.

25.1163 Powerplant accessories. 25.1363 Electrical system tests.

25.1165 Engine ignition systems. 25.1365 Electrical appliances, motors, and 25.1167 Accessory gearboxes. transformers.

Federal Aviation Administration, DOT Pt. 25

L IGHTS 25.1513 Minimum control speed.

25.1515 Landing gear speeds.

25.1381 Instrument lights.

25.1516 Other speed limitations.

25.1383 Landing lights.

25.1517 Rough air speed, V RA.

25.1385 Position light system installation.

25.1519 Weight, center of gravity, and 25.1387 Position light system dihedral an- weight distribution.

gles.

25.1521 Powerplant limitations.

25.1389 Position light distribution and in- 25.1522 Auxiliary power unit limitations.

tensities.

25.1523 Minimum flight crew.

25.1391 Minimum intensities in the hori- 25.1525 Kinds of operation.

zontal plane of forward and rear position 25.1527 Ambient air temperature and oper- lights.

ating altitude.

25.1393 Minimum intensities in any vertical 25.1529 Instructions for Continued Air- plane of forward and rear position lights.

25.1395 Maximum intensities in overlapping worthiness.

beams of forward and rear position 25.1531 Maneuvering flight load factors.

lights. 25.1533 Additional operating limitations.

25.1397 Color specifications.

25.1535 ETOPS approval.

25.1399 Riding light.

MARKINGS AND P LACARDS 25.1401 Anticollision light system.

25.1403 Wing icing detection lights.

25.1541 General.

25.1543 Instrument markings: general.

SAFETY EQUIPMENT 25.1545 Airspeed limitation information.

25.1411 General.

25.1547 Magnetic direction indicator.

25.1415 Ditching equipment.

25.1549 Powerplant and auxiliary power unit 25.1419 Ice protection.

instruments.

25.1420 Supercooled large drop icing condi- 25.1551 Oil quantity indication.

tions.

25.1553 Fuel quantity indicator.

25.1421 Megaphones.

25.1555 Control markings.

25.1423 Public address system.

25.1557 Miscellaneous markings and plac- ards.

MISCELLANEOUS E QUIPMENT 25.1561 Safety equipment.

25.1431 Electronic equipment.

25.1563 Airspeed placard.

25.1433 Vacuum systems.

25.1435 Hydraulic systems.

A IRPLANE F LIGHT M ANUAL 25.1438 Pressurization and pneumatic sys- 25.1581 General.

tems.

25.1583 Operating limitations.

25.1439 Protective breathing equipment.

25.1585 Operating procedures.

25.1441 Oxygen equipment and supply.

25.1587 Performance information.

25.1443 Minimum mass flow of supplemental oxygen.

Subpart H—Electrical Wiring 25.1445 Equipment standards for the oxygen Interconnection Systems (EWIS) distributing system.

25.1447 Equipment standards for oxygen dis- 25.1701 Definition.

pensing units.

25.1703 Function and installation: EWIS.

25.1449 Means for determining use of oxy- 25.1705 Systems and functions: EWIS.

gen.

25.1707 System separation: EWIS.

25.1450 Chemical oxygen generators.

25.1709 System safety: EWIS.

25.1453 Protection of oxygen equipment 25.1711 Component identification: EWIS.

from rupture.

25.1713 Fire protection: EWIS.

25.1455 Draining of fluids subject to freez- 25.1715 Electrical bonding and protection ing.

against static electricity: EWIS.

25.1457 Cockpit voice recorders.

25.1717 Circuit protective devices: EWIS.

25.1459 Flight data recorders.

25.1719 Accessibility provisions: EWIS.

25.1461 Equipment containing high energy 25.1721 Protection of EWIS.

rotors.

25.1723 Flammable fluid fire protection: EWIS.

Subpart G—Operating Limitations and 25.1725 Powerplants: EWIS.

Information 25.1727 Flammable fluid shutoff means: 25.1501 General.

EWIS.

25.1729 Instructions for Continued Air- O PERATING L IMITATIONS worthiness: EWIS.

25.1503 Airspeed limitations: general. 25.1731 Powerplant and APU fire detector 25.1505 Maximum operating limit speed. system: EWIS.

25.1507 Maneuvering speed. 25.1733 Fire detector systems, general: 25.1511 Flap extended speed. EWIS.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, SFAR No. 13

for the modification in lieu of Part 4a or Subpart I—Special Federal Aviation Part 4b as in effect on September 1, 1953: And Regulations provided further, That each specific modifica- 25.1801 SFAR No. 111—Lavatory Oxygen tion must be accomplished in accordance Systems. with all of the provisions contained in the PPENDIX A TO P ART 25 A elected rules relating to the particular modi- A PPENDIX B TO P ART 25 fication.

A PPENDIX C TO P ART 25 3. Specific conditions for approval. An appli- A PPENDIX D TO P ART 25 cant for any approval of the following spe- A PPENDIX E TO P ART 25 cific changes shall comply with section 2 of PPENDIX F TO P ART 25 A this regulation as modified by the applicable A PPENDIX G TO P ART 25 [R ESERVED ] provisions of this section.

A PPENDIX H TO P ART 25—I NSTRUCTIONS FOR (a) Increase in take-off power limitation— C ONTINUED A IRWORTHINESS 1,200 to 1,350 horsepower. The engine take-off APPENDIX I TO P ART 25—I NSTALLATION OF AN power limitation for the airplane may be in- A UTOMATIC T AKEOFF T HRUST C ONTROL creased to more than 1,200 horsepower but S YSTEM (ATTCS) not to more than 1,350 horsepower per engine A PPENDIX J TO P ART 25—E MERGENCY E VACU - if the increase in power does not adversely ATION affect the flight characteristics of the air- A PPENDIX K TO P ART 25—E XTENDED O PER - plane.

ATIONS (ETOPS) (b) Increase in take-off power limitation to A PPENDIX L TO P ART 25—HIRF E NVIRON - more than 1,350 horsepower. The engine take- MENTS AND E QUIPMENT HIRF T EST L EV - off power limitation for the airplane may be ELS increased to more than 1,350 horsepower per A PPENDIX M TO P ART 25—F UEL T ANK S YSTEM engine if compliance is shown with the flight F LAMMABILITY R EDUCTION M EANS characteristics and ground handling require- A PPENDIX N TO P ART 25—F UEL TANK F LAM - ments of Part 4b.

MABILITY E XPOSURE AND R ELIABILITY (c) Installation of engines of not more than A NALYSIS 1,830 cubic inches displacement and not having A PPENDIX O TO P ART 25—S UPERCOOLED L ARGE a certificated take-off rating of more than 1,350 D ROP I CING C ONDITIONS horsepower. Engines of not more than 1,830 A UTHORITY : 49 U.S.C. 106(f), 106(g), 40113, cubic inches displacement and not having a 44701, 44702 and 44704; Pub. L. 115–254, 132 Stat certificated take-off rating of more than 3281 (49 U.S.C. 44903 note).

1,350 horsepower which necessitate a major modification of redesign of the engine instal- S OURCE : Docket No. 5066, 29 FR 18291, Dec.

lation may be installed, if the engine fire 24, 1964, unless otherwise noted.

prevention and fire protection are equivalent to that on the prior engine installation.

S PECIAL F EDERAL A VIATION R EGULATION (d) Installation of engines of more than 1,830 N O. 13 cubic inches displacement or having certificated 1. Applicability. Contrary provisions of the take-off rating of more than 1,350 horsepower.

Civil Air Regulations regarding certification Engines of more than 1,830 cubic inches dis- notwithstanding, this regulation shall pro- placement or having certificated take-off vide the basis for approval by the Adminis- rating of more than 1,350 horsepower may be trator of modifications of individual Douglas installed if compliance is shown with the en- DC–3 and Lockheed L–18 airplanes subse- gine installation requirements of Part 4b: quent to the effective date of this regulation. Provided, That where literal compliance with 2. General modifications. Except as modified the engine installation requirements of Part in sections 3 and 4 of this regulation, an ap- 4b is extremely difficult to accomplish and plicant for approval of modifications to a would not contribute materially to the ob- DC–3 or L–18 airplane which result in jective sought, and the Administrator finds changes in design or in changes to approved that the experience with the DC–3 or L–18 limitations shall show that the modifica- airplanes justifies it, he is authorized to ac- tions were accomplished in accordance with cept such measures of compliance as he finds the rules of either Part 4a or Part 4b in ef- will effectively accomplish the basic objec- fect on September 1, 1953, which are applica- tive.

ble to the modification being made: Provided, 4. Establishment of new maximum certificated That an applicant may elect to accomplish a weights. An applicant for approval of new modification in accordance with the rules of maximum certificated weights shall apply Part 4b in effect on the date of application for an amendment of the airworthiness cer- tificate of the airplane and shall show that 1 the weights sought have been established, It is not intended to waive compliance and the appropriate manual material ob- with such airworthiness requirements as are tained, as provided in this section.

included in the operating parts of the Civil Air Regulations for specific types of oper- N OTE : Transport category performance re- ation. quirements result in the establishment of

Federal Aviation Administration, DOT Pt. 25, SFAR No. 109

maximum certificated weights for various S PECIAL F EDERAL A VIATION R EGULATION altitudes.

N O. 109 (a) Weights–25,200 to 26,900 for the DC–3 and 1. Applicability. Contrary provisions of 14 18,500 to 19,500 for the L–18. New maximum CFR parts 21, 25, and 119 of this chapter not- certificated weights of more than 25,200 but withstanding, an applicant is entitled to an not more than 26,900 pounds for DC–3 and amended type certificate or supplemental more than 18,500 but not more than 19,500 type certificate in the transport category, if pounds for L–18 airplanes may be established the applicant complies with all applicable in accordance with the transport category provisions of this SFAR.

performance requirements of either Part 4a or Part 4b, if the airplane at the new max- Operations imum weights can meet the structural re- 2. General.

quirements of the elected part.

(a) The passenger capacity may not exceed (b) Weights of more than 26,900 for the DC–3 60. If more than 60 passenger seats are in- and 19,500 for the L–18. New maximum certifi- stalled, then: cated weights of more than 26,900 pounds for (1) If the extra seats are not suitable for DC–3 and 19,500 pounds for L–18 airplanes occupancy during taxi, takeoff and landing, shall be established in accordance with the each extra seat must be clearly marked (e.g., structural performance, flight characteris- a placard on the top of an armrest, or a placard sewn into the top of the back cush- tics, and ground handling requirements of ion) that the seat is not to be occupied dur- Part 4b: Provided, That where literal compli- ing taxi, takeoff and landing.

ance with the structural requirements of (2) If the extra seats are suitable for occu- Part 4b is extremely difficult to accomplish pancy during taxi, takeoff and landing ( i.e., and would not contribute materially to the meet all the strength and passenger injury objective sought, and the Administrator criteria in part 25), then a note must be in- finds that the experience with the DC–3 or L– cluded in the Limitations Section of the Air- 18 airplanes justifies it, he is authorized to plane Flight Manual that there are extra accept such measures of compliance as he seats installed but that the number of pas- finds will effectively accomplish the basic sengers on the airplane must not exceed 60.

objective.

Additionally, there must be a placard in- (c) Airplane flight manual-performance oper- stalled adjacent to each door that can be ating information. An approved airplane flight used as a passenger boarding door that states manual shall be provided for each DC–3 and that the maximum passenger capacity is 60.

L–18 airplane which has had new maximum The placard must be clearly legible to pas- certificated weights established under this sengers entering the airplane.

section. The airplane flight manual shall (b) For airplanes outfitted with interior contain the applicable performance informa- doors under paragraph 10 of this SFAR, the airplane flight manual (AFM) must include tion prescribed in that part of the regula- an appropriate limitation that the airplane tions under which the new certificated must be staffed with at least the following weights were established and such additional number of flight attendants who meet the re- information as may be necessary to enable quirements of 14 CFR 91.533(b): the application of the take-off, en route, and (1) The number of flight attendants re- landing limitations prescribed for transport quired by § 91.533(a)(1) and (2) of this chapter, category airplanes in the operating parts of and the Civil Air Regulations.

(2) At least one flight attendant if the air- (d) Performance operating limitations. Each plane model was originally certified for 75 airplane for which new maximum certifi- passengers or more.

cated weights are established in accordance (c) The AFM must include appropriate lim- with paragraphs (a) or (b) of this section itation(s) to require a preflight passenger shall be considered a transport category air- briefing describing the appropriate functions plane for the purpose of complying with the to be performed by the passengers and the performance operating limitations applica- relevant features of the airplane to ensure ble to the operations in which it is utilized.

the safety of the passengers and crew.

5. Reference. Unless otherwise provided, all (d) The airplane may not be offered for common carriage or operated for hire. The references in this regulation to Part 4a and operating limitations section of the AFM Part 4b are those parts of the Civil Air Regu- must be revised to prohibit any operations lations in effect on September 1, 1953.

involving the carriage of persons or property This regulation supersedes Special Civil for compensation or hire. The operators may Air Regulation SR–398 and shall remain ef- receive remuneration to the extent con- fective until superseded or rescinded by the sistent with parts 125 and 91, subpart F, of Board.

this chapter.

[19 FR 5039, Aug. 11, 1954. Redesignated at 29 (e) A placard stating that ‘‘Operations in- FR 19099, Dec. 30, 1964] volving the carriage of persons or property

14 CFR Ch. I (1–1–25 Edition) Pt. 25, SFAR No. 109

for compensation or hire are prohibited,’’ Motor Vehicle Safety Standards (FMVSS) must be located in the area of the Airworthi- part 571.214, section S6.13.5 (49 CFR 571.214).

ness Certificate holder at the entrance to the The TTI must be less than 85, as defined in flightdeck. 49 CFR part 572, subpart F. Torso contact (f) For passenger capacities of 45 to 60 pas- during rebound is acceptable and need not be sengers, analysis must be submitted that measured.

demonstrates that the airplane can be evacu- (4) Pelvis. If the pelvis of an ATD at any ated in less than 90 seconds under the condi- seat place impacts seat and/or adjacent tions specified in § 25.803 and appendix J to structure during testing, pelvic lateral accel- part 25. eration injury criteria must be substantiated (g) In order for any airplane certified under by dynamic test or by rational analysis this SFAR to be placed in part 135 or part 121 based on previous test(s) of a similar seat in- operations, the airplane must be brought stallation. Pelvic lateral acceleration may back into full compliance with the applica- not exceed 130g. Pelvic acceleration data ble operational part.

must be processed as defined in FMVSS part 571.214, section S6.13.5 (49 CFR 571.214).

Equipment and Design (5) Body-to-Wall/Furnishing Contact. If the seat is installed aft of a structure—such as 3. General. Unless otherwise noted, compli- an interior wall or furnishing that may con- ance is required with the applicable certifi- tact the pelvis, upper arm, chest, or head of cation basis for the airplane. Some provi- an occupant seated next to the structure— sions of this SFAR impose alternative re- the structure or a conservative representa- quirements to certain airworthiness stand- tion of the structure and its stiffness must ards that do not apply to airplanes certifi- be included in the tests. It is recommended, cated to earlier standards. Those airplanes but not required, that the contact surface of with an earlier certification basis are not re- the actual structure be covered with at least quired to comply with those alternative re- two inches of energy absorbing protective quirements.

padding (foam or equivalent) such as 4. Occupant Protection.

Ensolite.

(a) Firm Handhold. In lieu of the require- (6) Shoulder Strap Loads. Where upper torso ments of § 25.785(j), there must be means pro- straps (shoulder straps) are used for sofa oc- vided to enable persons to steady themselves cupants, the tension loads in individual in moderately rough air while occupying straps may not exceed 1,750 pounds. If dual aisles that are along the cabin sidewall, or straps are used for restraining the upper where practicable, bordered by seats (seat torso, the total strap tension loads may not backs providing a 25-pound minimum break- exceed 2,000 pounds.

away force are an acceptable means of com- (7) Occupant Retention. All side-facing seats pliance).

require end closures or other means to pre- (b) Injury criteria for multiple occupancy vent the ATD’s pelvis from translating be- side-facing seats. The following require- yond the end of the seat at any time during ments are only applicable to airplanes that testing.

are subject to § 25.562.

(8) Test Parameters.

(1) Existing Criteria. All injury protection (i) All seat positions need to be occupied by criteria of § 25.562(c)(1) through (c)(6) apply to ATDs for the longitudinal tests.

the occupants of side-facing seating. The (ii) A minimum of one longitudinal test, Head Injury Criterion (HIC) assessments are conducted in accordance with the conditions only required for head contact with the seat specified in § 25.562(b)(2), is required to assess and/or adjacent structures.

the injury criteria as follows. Note that if a (2) Body-to-Body Contact. Contact between seat is installed aft of structure (such as an the head, pelvis, torso or shoulder area of interior wall or furnishing) that does not one Anthropomorphic Test Dummy (ATD) have a homogeneous surface, an additional with the head, pelvis, torso or shoulder area test or tests may be required to demonstrate of the ATD in the adjacent seat is not al- that the injury criteria are met for the area lowed during the tests conducted in accord- which an occupant could contact. For exam- ance with § 25.562(b)(1) and (b)(2). Contact ple, different yaw angles could result in dif- during rebound is allowed.

ferent injury considerations and may require (3) Thoracic Trauma. If the torso of an ATD separate tests to evaluate.

at the forward-most seat place impacts the seat and/or adjacent structure during test- (A) For configurations without structure ing, compliance with the Thoracic Trauma (such as a wall or bulkhead) installed di- Index (TTI) injury criterion must be substan- rectly forward of the forward seat place, Hy- tiated by dynamic test or by rational anal- brid II ATDs or equivalent must be in all ysis based on previous test(s) of a similar seat places.

seat installation. TTI data must be acquired (B) For configurations with structure (such with a Side Impact Dummy (SID), as defined as a wall or bulkhead) installed directly for- by 49 CFR part 572, subpart F, or an equiva- ward of the forward seat place, a side impact lent ATD or a more appropriate ATD and dummy or equivalent ATD or more appro- must be processed as defined in Federal priate ATD must be in the forward seat place

Federal Aviation Administration, DOT Pt. 25, SFAR No. 109

and a Hybrid II ATD or equivalent must be the maximum allowable after modification in all other seat places. when exits are deactivated, whichever is less.

(C) The test may be conducted with or (d) A distance of more than 60 feet between without deformed floor.

adjacent passenger emergency exits on the (D) The test must be conducted with either same side of the same deck of the fuselage, no yaw or 10 degrees yaw for evaluating oc- as measured parallel to the airplane’s longi- cupant injury. Deviating from the no yaw tudinal axis between the nearest exit edges, condition may not result in the critical area is allowed only once on each side of the fuse- of contact not being evaluated. The upper lage.

torso restraint straps, where installed, must 8. Emergency Exit Signs. In lieu of the re- remain on the occupant’s shoulder during quirements of § 25.811(d)(1) and (2) a single the impact condition of § 25.562(b)(2).

sign at each exit may be installed provided: (c) For the vertical test, conducted in ac- (a) The sign can be read from the aisle cordance with the conditions specified in while directly facing the exit, and § 25.562(b)(1), Hybrid II ATDs or equivalent (b) The sign can be read from the aisle ad- must be used in all seat positions.

jacent to the passenger seat that is farthest 5. Direct View. In lieu of the requirements from the exit and that does not have an in- of § 25.785(h)(2), to the extent practical with- tervening bulkhead/divider or exit.

out compromising proximity to a required 9. Emergency Lighting.

floor level emergency exit, the majority of installed flight attendant seats must be lo- (a) Exit Signs. In lieu of the requirements of cated to face the cabin area for which the § 25.812(b)(1), for airplanes that have a pas- flight attendant is responsible.

senger seating configuration, excluding pilot 6. Passenger Information Signs. Compliance seats, of 19 seats or less, the emergency exit with § 25.791 is required except that for signs required by § 25.811(d)(1), (2), and (3) § 25.791(a), when smoking is to be prohibited, must have red letters at least 1-inch high on notification to the passengers may be pro- a white background at least 2 inches high.

vided by a single placard so stating, to be These signs may be internally electrically il- conspicuously located inside the passenger luminated, or self illuminated by other than compartment, easily visible to all persons electrical means, with an initial brightness entering the cabin in the immediate vicinity of at least 160 microlamberts. The color may of each passenger entry door.

be reversed in the case of a sign that is self- 7. Distance Between Exits. For an airplane illuminated by other than electrical means.

that is required to comply with § 25.807(f)(4), (b) Floor Proximity Escape Path Marking. In in effect as of July 24, 1989, which has more lieu of the requirements of § 25.812(e)(1), for than one passenger emergency exit on each cabin seating compartments that do not side of the fuselage, no passenger emergency have the main cabin aisle entering and exit may be more than 60 feet from any adja- exiting the compartment, the following are cent passenger emergency exit on the same applicable: side of the same deck of the fuselage, as (1) After a passenger leaves any passenger measured parallel to the airplane’s longitu- seat in the compartment, he/she must be dinal axis between the nearest exit edges, able to exit the compartment to the main unless the following conditions are met: cabin aisle using only markings and visual (a) Each passenger seat must be located features not more that 4 feet above the cabin within 30 feet from the nearest exit on each floor, and side of the fuselage, as measured parallel to (2) Proceed to the exits using the marking the airplane’s longitudinal axis, between the system necessary to accomplish the actions nearest exit edge and the front of the seat in § 25.812(e)(1) and (e)(2).

bottom cushion.

(c) Transverse Separation of the Fuselage. In (b) The number of passenger seats located the event of a transverse separation of the between two adjacent pairs of emergency fuselage, compliance must be shown with exits (commonly referred to as a passenger § 25.812(l) except as follows: zone) or between a pair of exits and a bulk- (1) For each airplane type originally type head or a compartment door (commonly re- certificated with a maximum passenger seat- ferred to as a ‘‘dead-end zone’’), may not ex- ing capacity of 9 or less, not more than 50 ceed the following: percent of all electrically illuminated emer- (1) For zones between two pairs of exits, 50 gency lights required by § 25.812 may be ren- percent of the combined rated capacity of dered inoperative in addition to the lights the two pairs of emergency exits.

that are directly damaged by the separation.

(2) For zones between one pair of exits and a bulkhead, 40 percent of the rated capacity (2) For each airplane type originally type of the pair of emergency exits. certificated with a maximum passenger seat- (c) The total number of passenger seats in ing capacity of 10 to 19, not more than 33 per- the airplane may not exceed 33 percent of the cent of all electrically illuminated emer- maximum seating capacity for the airplane gency lights required by § 25.812 may be ren- model using the exit ratings listed in dered inoperative in addition to the lights § 25.807(g) for the original certified exits or that are directly damaged by the separation.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, SFAR No. 109

10. Interior doors. In lieu of the require- 14. Cooktops. Each cooktop must be de- ments of § 25.813(e), interior doors may be in- signed and installed to minimize any poten- stalled between passenger seats and exits, tial threat to the airplane, passengers, and provided the following requirements are met. crew. Compliance with this requirement (a) Each door between any passenger seat, must be found in accordance with the fol- occupiable for taxi, takeoff, and landing, and lowing criteria: (a) Means, such as conspicuous burner-on any emergency exit must have a means to indicators, physical barriers, or handholds, signal to the flightcrew, at the flightdeck, must be installed to minimize the potential that the door is in the open position for taxi, for inadvertent personnel contact with hot takeoff and landing.

surfaces of both the cooktop and cookware.

(b) Appropriate procedures/limitations Conditions of turbulence must be considered.

must be established to ensure that any such (b) Sufficient design means must be in- door is in the open configuration for takeoff cluded to restrain cookware while in place and landing.

on the cooktop, as well as representative (c) Each door between any passenger seat contents, e.g., soup, sauces, etc., from the ef- and any exit must have dual means to retain fects of flight loads and turbulence. Re- it in the open position, each of which is capa- straints must be provided to preclude haz- ble of reacting the inertia loads specified in ardous movement of cookware and contents.

§ 25.561.

These restraints must accommodate any (d) Doors installed across a longitudinal cookware that is identified for use with the aisle must translate laterally to open and cooktop. Restraints must be designed to be close, e.g., pocket doors.

easily utilized and effective in service. The (e) Each door between any passenger seat cookware restraint system should also be de- and any exit must be frangible in either di- signed so that it will not be easily disabled, rection.

thus rendering it unusable. Placarding must (f) Each door between any passenger seat be installed which prohibits the use of and any exit must be operable from either cookware that cannot be accommodated by side, and if a locking mechanism is installed, the restraint system.

it must be capable of being unlocked from ei- (c) Placarding must be installed which pro- ther side without the use of special tools.

hibits the use of cooktops (i.e., power on any 11. Width of Aisle. Compliance is required burner) during taxi, takeoff, and landing.

with § 25.815, except that aisle width may be (d) Means must be provided to address the reduced to 0 inches between passenger seats possibility of a fire occurring on or in the during in-flight operations only, provided immediate vicinity of the cooktop. Two ac- that the applicant demonstrates that all ceptable means of complying with this re- areas of the cabin are easily accessible by a quirement are as follows: crew member in the event of an emergency (1) Placarding must be installed that pro- (e.g., in-flight fire, decompression). Addition- hibits any burner from being powered when ally, instructions must be provided at each the cooktop is unattended. (N OTE : This passenger seat for restoring the aisle width would prohibit a single person from cooking required by § 25.815. Procedures must be es- on the cooktop and intermittently serving tablished and documented in the AFM to en- food to passengers while any burner is pow- sure that the required aisle widths are pro- ered.) A fire detector must be installed in the vided during taxi, takeoff, and landing.

vicinity of the cooktop which provides an au- 12. Materials for Compartment Interiors. dible warning in the passenger cabin, and a Compliance is required with the applicable fire extinguisher of appropriate size and ex- provisions of § 25.853, except that compliance tinguishing agent must be installed in the with appendix F, parts IV and V, to part 25, immediate vicinity of the cooktop. Access to need not be demonstrated if it can be shown the extinguisher may not be blocked by a by test or a combination of test and analysis fire on or around the cooktop.

that the maximum time for evacuation of all (2) An automatic, thermally activated fire occupants does not exceed 45 seconds under suppression system must be installed to ex- the conditions specified in appendix J to part tinguish a fire at the cooktop and imme- 25. diately adjacent surfaces. The agent used in 13. Fire Detection. For airplanes with a type the system must be an approved total flood- certificated passenger capacity of 20 or more, ing agent suitable for use in an occupied there must be means that meet the require- area. The fire suppression system must have ments of § 25.858(a) through (d) to signal the a manual override. The automatic activation flightcrew in the event of a fire in any iso- of the fire suppression system must also lated room not occupiable for taxi, takeoff automatically shut off power to the cooktop.

and landing, which can be closed off from the (e) The surfaces of the galley surrounding rest of the cabin by a door. The indication the cooktop which would be exposed to a fire must identify the compartment where the on the cooktop surface or in cookware on the fire is located. This does not apply to lava- cooktop must be constructed of materials tories, which continue to be governed by that comply with the flammability require- § 25.854. ments of part III of appendix F to part 25.

Federal Aviation Administration, DOT § 25.3 This requirement is in addition to the flam- Subpart A—General mability requirements typically required of the materials in these galley surfaces. Dur- § 25.1 Applicability.

ing the selection of these materials, consid- (a) This part prescribes airworthiness eration must also be given to ensure that the standards for the issue of type certifi- flammability characteristics of the mate- rials will not be adversely affected by the use cates, and changes to those certifi- of cleaning agents and utensils used to re- cates, for transport category airplanes.

move cooking stains.

(b) Each person who applies under (f) The cooktop must be ventilated with a Part 21 for such a certificate or change system independent of the airplane cabin and must show compliance with the appli- cargo ventilation system. Procedures and cable requirements in this part.

time intervals must be established to inspect and clean or replace the ventilation system § 25.2 Special retroactive require- to prevent a fire hazard from the accumula- ments.

tion of flammable oils and be included in the The following special retroactive re- instructions for continued airworthiness.

The ventilation system ducting must be pro- quirements are applicable to an air- tected by a flame arrestor. [N OTE : The appli- plane for which the regulations ref- cant may find additional useful information erenced in the type certificate predate in Society of Automotive Engineers, Aero- the sections specified below— space Recommended Practice 85, Rev. E, en- (a) Irrespective of the date of applica- titled ‘‘Air Conditioning Systems for Sub- tion, each applicant for a supplemental sonic Airplanes,’’ dated August 1, 1991.]

type certificate (or an amendment to a (g) Means must be provided to contain type certificate) involving an increase spilled foods or fluids in a manner that will in passenger seating capacity to a total prevent the creation of a slipping hazard to occupants and will not lead to the loss of greater than that for which the air- structural strength due to airplane corro- plane has been type certificated must sion.

show that the airplane concerned (h) Cooktop installations must provide meets the requirements of: adequate space for the user to immediately (1) Sections 25.721(d), 25.783(g), escape a hazardous cooktop condition.

25.785(c), 25.803(c)(2) through (9), 25.803 (i) A means to shut off power to the (d) and (e), 25.807 (a), (c), and (d), 25.809 cooktop must be provided at the galley con- (f) and (h), 25.811, 25.812, 25.813 (a), (b), taining the cooktop and in the cockpit. If ad- and (c), 25.815, 25.817, 25.853 (a) and (b), ditional switches are introduced in the cock- 25.855(a), 25.993(f), and 25.1359(c) in ef- pit, revisions to smoke or fire emergency procedures of the AFM will be required. fect on October 24, 1967, and (j) If the cooktop is required to have a lid (2) Sections 25.803(b) and 25.803(c)(1) to enclose the cooktop there must be a in effect on April 23, 1969.

means to automatically shut off power to (b) Irrespective of the date of applica- the cooktop when the lid is closed.

tion, each applicant for a supplemental 15. Hand-Held Fire Extinguishers.

type certificate (or an amendment to a (a) For airplanes that were originally type type certificate) for an airplane manu- certificated with more than 60 passengers, factured after October 16, 1987, must the number of hand-held fire extinguishers show that the airplane meets the re- must be the greater of— quirements of § 25.807(c)(7) in effect on (1) That provided in accordance with the July 24, 1989.

requirements of § 25.851, or (2) A number equal to the number of origi- (c) Compliance with subsequent revi- nally type certificated exit pairs, regardless sions to the sections specified in para- of whether the exits are deactivated for the graph (a) or (b) of this section may be proposed configuration.

elected or may be required in accord- (b) Extinguishers must be evenly distrib- ance with § 21.101(a) of this chapter.

uted throughout the cabin. These extin- [Amdt. 25–72, 55 FR 29773, July 20, 1990, as guishers are in addition to those required by amended by Amdt. 25–99, 65 FR 36266, June 7, paragraph 14 of this SFAR, unless it can be 2000] shown that the cooktop was installed in the immediate vicinity of the original exits.

§ 25.3 Special provisions for ETOPS 16. Security. The requirements of § 25.795 are type design approvals.

not applicable to airplanes approved in ac- cordance with this SFAR.

(a) Applicability. This section applies to an applicant for ETOPS type design [Doc. No. FAA–2007–28250, 74 FR 21541, May 8, 2009] approval of an airplane: 14 CFR Ch. I (1–1–25 Edition) § 25.4 (1) That has an existing type certifi- (ii) A physical discomfort or a sig- cate on February 15, 2007; or nificant increase in flightcrew work- (2) For which an application for an load or in conditions impairing the ef- original type certificate was submitted ficiency of the flightcrew, before February 15, 2007.

(iii) Physical distress to passengers (b) Airplanes with two engines. (1) For or cabin crew, possibly including inju- ETOPS type design approval of an air- ries, or plane up to and including 180 minutes, (iv) An effect of similar severity.

an applicant must comply with (2) Hazardous failure condition means § 25.1535, except that it need not comply a failure condition that would reduce with the following provisions of Appen- the capability of the airplane or the dix K, K25.1.4, of this part: ability of the flightcrew to cope with (i) K25.1.4(a), fuel system pressure adverse operating conditions, to the ex- and flow requirements; tent that there would be— (ii) K25.1.4(a)(3), low fuel alerting; and (i) A large reduction in safety mar- (iii) K25.1.4(c), engine oil tank design. gins or functional capabilities, (2) For ETOPS type design approval (ii) Physical distress or excessive of an airplane beyond 180 minutes an workload such that the flightcrew can- applicant must comply with § 25.1535.

not be relied upon to perform their (c) Airplanes with more than two en- tasks accurately or completely, or gines. An applicant for ETOPS type de- (iii) Serious or fatal injuries to a rel- sign approval must comply with atively small number of persons other § 25.1535 for an airplane manufactured than the flightcrew.

on or after February 17, 2015, except (3) Catastrophic failure condition that, for an airplane configured for a means a failure condition that would three person flight crew, the applicant result in multiple fatalities, usually need not comply with Appendix K, with the loss of the airplane.

K25.1.4(a)(3), of this part, low fuel alert- (c) For purposes of this part, the fol- ing.

lowing failure conditions in order of [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, decreasing probability apply: 2007] (1) Probable failure condition means a failure condition that is anticipated to § 25.4 Definitions.

occur one or more times during the en- (a) For the purposes of this part, the tire operational life of each airplane of following general definitions apply: a given type.

(1) Certification maintenance require- (2) Remote failure condition means a ment means a required scheduled main- failure condition that is not antici- tenance task established during the de- pated to occur to each airplane of a sign certification of the airplane sys- given type during its entire operational tems as an airworthiness limitation of life, but which may occur several times the type certificate or supplemental during the total operational life of a type certificate.

number of airplanes of a given type.

(2) Significant latent failure is a latent failure that, in combination with one (3) Extremely remote failure condition or more specific failures or events, means a failure condition that is not would result in a hazardous or cata- anticipated to occur to each airplane of strophic failure condition.

a given type during its entire oper- (b) For purposes of this part, the fol- ational life, but which may occur a few lowing failure conditions, in order of times during the total operational life increasing severity, apply: of all airplanes of a given type.

(1) Major failure condition means a (4) Extremely improbable failure condi- failure condition that would reduce the tion means a failure condition that is capability of the airplane or the ability not anticipated to occur during the of the flightcrew to cope with adverse total operational life of all airplanes of operating conditions, to the extent a given type.

that there would be— (i) A significant reduction in safety [Doc. No. FAA–2022–1544, 89 FR 68731, Aug. 27, margins or functional capabilities, 2024] Federal Aviation Administration, DOT § 25.21 cannot be reasonably inferred from § 25.5 Incorporations by reference.

combinations investigated.

(a) The materials listed in this sec- (b) [Reserved] tion are incorporated by reference in (c) The controllability, stability, the corresponding sections noted.

trim, and stalling characteristics of These incorporations by reference were the airplane must be shown for each al- approved by the Director of the Federal titude up to the maximum expected in Register in accordance with 5 U.S.C.

operation.

552(a) and 1 CFR part 51. These mate- rials are incorporated as they exist on (d) Parameters critical for the test the date of the approval, and notice of being conducted, such as weight, load- any change in these materials will be ing (center of gravity and inertia), air- published in the F EDERAL R EGISTER .

speed, power, and wind, must be main- The materials are available for pur- tained within acceptable tolerances of chase at the corresponding addresses the critical values during flight test- noted below, and all are available for ing.

inspection at the National Archives (e) If compliance with the flight and Records Administration (NARA).

characteristics requirements is depend- For information on the availability of ent upon a stability augmentation sys- this material at NARA, call 202–741– tem or upon any other automatic or 6030, or go to: http://www.archives.gov/ power-operated system, compliance federal-register/cfr/ibr-locations.html.

must be shown with §§ 25.671 and 25.672.

(b) The following materials are avail- (f) In meeting the requirements of able for purchase from the following §§ 25.105(d), 25.125, 25.233, and 25.237, the address: The National Technical Infor- wind velocity must be measured at a mation Services (NTIS), Springfield, height of 10 meters above the surface, Virginia 22166.

or corrected for the difference between (1) Fuel Tank Flammability Assess- the height at which the wind velocity ment Method User’s Manual, dated is measured and the 10-meter height.

May 2008, document number DOT/FAA/ (g) The requirements of this subpart AR–05/8, IBR approved for § 25.981 and associated with icing conditions apply Appendix N. It can also be obtained at only if the applicant is seeking certifi- the following Web site: http:// cation for flight in icing conditions.

www.fire.tc.faa.gov/systems/fueltank/ (1) Paragraphs (g)(3) and (4) of this FTFAM.stm.

section apply only to airplanes with (2) [Reserved] one or both of the following attributes: [73 FR 42494, July 21, 2008, as amended by (i) Maximum takeoff gross weight is Doc. No. FAA–2018–0119, Amdt. 21–101, 83 FR less than 60,000 lbs; or 9169, Mar. 5, 2018] (ii) The airplane is equipped with re- versible flight controls.

Subpart B—Flight (2) Each requirement of this subpart, except §§ 25.121(a), 25.123(c), 25.143(b)(1) G ENERAL and (2), 25.149, 25.201(c)(2), 25.239, and § 25.21 Proof of compliance. 25.251(b) through (e), must be met in the icing conditions specified in Appen- (a) Each requirement of this subpart dix C of this part. Section 25.207(c) and must be met at each appropriate com- (d) must be met in the landing configu- bination of weight and center of grav- ration in the icing conditions specified ity within the range of loading condi- in Appendix C, but need not be met for tions for which certification is re- other configurations. Compliance must quested. This must be shown— be shown using the ice accretions de- (1) By tests upon an airplane of the fined in part II of Appendix C of this type for which certification is re- part, assuming normal operation of the quested, or by calculations based on, airplane and its ice protection system and equal in accuracy to, the results of in accordance with the operating limi- testing; and (2) By systematic investigation of tations and operating procedures estab- each probable combination of weight lished by the applicant and provided in and center of gravity, if compliance the airplane flight manual.

14 CFR Ch. I (1–1–25 Edition) § 25.23 (3) If the applicant does not seek cer- spanwise) that could be inadvertently tification for flight in all icing condi- exceeded, these limits and the cor- tions defined in Appendix O of this responding weight and center of grav- part, each requirement of this subpart, ity combinations must be established.

except §§ 25.105, 25.107, 25.109, 25.111, (b) The load distribution limits may 25.113, 25.115, 25.121, 25.123, 25.143(b)(1), not exceed— (b)(2), and (c)(1), 25.149, 25.201(c)(2), (1) The selected limits; 25.207(c), (d), and (e)(1), 25.239, and (2) The limits at which the structure 25.251(b) through (e), must be met in is proven; or the Appendix O icing conditions for (3) The limits at which compliance which certification is not sought in with each applicable flight require- order to allow a safe exit from those ment of this subpart is shown.

conditions. Compliance must be shown using the ice accretions defined in part § 25.25 Weight limits.

II, paragraphs (b) and (d) of Appendix (a) Maximum weights. Maximum O, assuming normal operation of the weights corresponding to the airplane airplane and its ice protection system operating conditions (such as ramp, in accordance with the operating limi- ground or water taxi, takeoff, en route, tations and operating procedures estab- and landing), environmental conditions lished by the applicant and provided in (such as altitude and temperature), and the airplane flight manual.

loading conditions (such as zero fuel (4) If the applicant seeks certifi- weight, center of gravity position and cation for flight in any portion of the weight distribution) must be estab- icing conditions of Appendix O of this lished so that they are not more than— part, each requirement of this subpart, (1) The highest weight selected by except §§ 25.121(a), 25.123(c), 25.143(b)(1) the applicant for the particular condi- and (2), 25.149, 25.201(c)(2), 25.239, and tions; or 25.251(b) through (e), must be met in the Appendix O icing conditions for (2) The highest weight at which com- which certification is sought. Section pliance with each applicable structural 25.207(c) and (d) must be met in the loading and flight requirement is landing configuration in the Appendix shown, except that for airplanes O icing conditions for which certifi- equipped with standby power rocket cation is sought, but need not be met engines the maximum weight must not for other configurations. Compliance be more than the highest weight estab- must be shown using the ice accretions lished in accordance with appendix E of defined in part II, paragraphs (c) and this part; or (d) of Appendix O, assuming normal op- (3) The highest weight at which com- eration of the airplane and its ice pro- pliance is shown with the certification tection system in accordance with the requirements of Part 36 of this chapter.

operating limitations and operating (b) Minimum weight. The minimum procedures established by the applicant weight (the lowest weight at which and provided in the airplane flight compliance with each applicable re- manual.

quirement of this part is shown) must be established so that it is not less [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as than— amended by Amdt. 25–23, 35 FR 5671, Apr. 8, 1970; Amdt. 25–42, 43 FR 2320, Jan. 16, 1978; (1) The lowest weight selected by the Amdt. 25–72, 55 FR 29774, July 20, 1990; Amdt.

applicant; 25–121, 72 FR 44665, Aug. 8, 2007 Amdt. 25–135, (2) The design minimum weight (the 76 FR 74654, Dec. 1, 2011; Amdt. 25–140, 79 FR lowest weight at which compliance 65524, Nov. 4, 2014] with each structural loading condition of this part is shown); or § 25.23 Load distribution limits.

(3) The lowest weight at which com- (a) Ranges of weights and centers of pliance with each applicable flight re- gravity within which the airplane may quirement is shown.

be safely operated must be established.

If a weight and center of gravity com- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as bination is allowable only within cer- amended by Amdt. 25–23, 35 FR 5671, Apr. 8, tain load distribution limits (such as 1970; Amdt. 25–63, 53 FR 16365, May 6, 1988] Federal Aviation Administration, DOT § 25.101 (c) The means used to limit the low § 25.27 Center of gravity limits.

pitch position of the propeller blades The extreme forward and the extreme must be set so that the engine does not aft center of gravity limitations must exceed 103 percent of the maximum al- be established for each practicably sep- lowable engine rpm or 99 percent of an arable operating condition. No such approved maximum overspeed, which- limit may lie beyond— ever is greater, with— (a) The extremes selected by the ap- (1) The propeller blades at the low plicant; pitch limit and governor inoperative; (b) The extremes within which the (2) The airplane stationary under structure is proven; or standard atmospheric conditions with (c) The extremes within which com- no wind; and pliance with each applicable flight re- (3) The engines operating at the take- quirement is shown.

off manifold pressure limit for recipro- cating engine powered airplanes or the § 25.29 Empty weight and cor- maximum takeoff torque limit for tur- responding center of gravity.

bopropeller engine-powered airplanes.

(a) The empty weight and cor- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as responding center of gravity must be amended by Amdt. 25–57, 49 FR 6848, Feb. 23, determined by weighing the airplane 1984; Amdt. 25–72, 55 FR 29774, July 20, 1990] with— (1) Fixed ballast; P ERFORMANCE (2) Unusable fuel determined under § 25.959; and § 25.101 General.

(3) Full operating fluids, including— (a) Unless otherwise prescribed, air- (i) Oil; planes must meet the applicable per- (ii) Hydraulic fluid; and formance requirements of this subpart (iii) Other fluids required for normal for ambient atmospheric conditions operation of airplane systems, except and still air.

potable water, lavatory precharge (b) The performance, as affected by water, and fluids intended for injection engine power or thrust, must be based in the engine.

on the following relative humidities; (b) The condition of the airplane at (1) For turbine engine powered air- the time of determining empty weight planes, a relative humidity of— must be one that is well defined and (i) 80 percent, at and below standard can be easily repeated.

temperatures; and [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (ii) 34 percent, at and above standard amended by Amdt. 25–42, 43 FR 2320, Jan. 16, temperatures plus 50 ° F.

1978; Amdt. 25–72, 55 FR 29774, July 20, 1990] Between these two temperatures, the relative humidity must vary linearly.

§ 25.31 Removable ballast.

(2) For reciprocating engine powered Removable ballast may be used on airplanes, a relative humidity of 80 per- showing compliance with the flight re- cent in a standard atmosphere. Engine quirements of this subpart.

power corrections for vapor pressure must be made in accordance with the § 25.33 Propeller speed and pitch lim- following table: its.

Vapor Specific humidity Density ratio (a) The propeller speed and pitch Altitude pressure e w (Lb. moisture r / s = H ( ft. ) must be limited to values that will en- (In. Hg.) per lb. dry air) 0.0023769 sure— 0 0.403 0.00849 0.99508 (1) Safe operation under normal oper- 1,000 .354 .00773 .96672 ating conditions; and 2,000 .311 .00703 .93895 3,000 .272 .00638 .91178 (2) Compliance with the performance 4,000 .238 .00578 .88514 requirements of §§ 25.101 through 25.125.

5,000 .207 .00523 .85910 (b) There must be a propeller speed 6,000 .1805 .00472 .83361 7,000 .1566 .00425 .80870 limiting means at the governor. It 8,000 .1356 .00382 .78434 must limit the maximum possible gov- 9,000 .1172 .00343 .76053 erned engine speed to a value not ex- 10,000 .1010 .00307 .73722 ceeding the maximum allowable r.p.m. 15,000 .0463 .001710 .62868 14 CFR Ch. I (1–1–25 Edition) § 25.103 Vapor Specific humidity Density ratio 25.125, respectively, must be deter- Altitude pressure e w (Lb. moisture r / s = H ( ft. ) mined with all the airplane wheel (In. Hg.) per lb. dry air) 0.0023769 brake assemblies at the fully worn 20,000 .01978 .000896 .53263 limit of their allowable wear range.

25,000 .00778 .000436 .44806 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55466, Dec. 20, (c) The performance must correspond 1976; Amdt. 25–92, 63 FR 8318, Feb. 18, 1998] to the propulsive thrust available under the particular ambient atmos- § 25.103 Stall speed.

pheric conditions, the particular flight condition, and the relative humidity (a) The reference stall speed, V , is a SR specified in paragraph (b) of this sec- calibrated airspeed defined by the ap- tion. The available propulsive thrust plicant. V may not be less than a 1-g SR must correspond to engine power or stall speed. V is expressed as: SR thrust, not exceeding the approved

power or thrust less— V

CL MAX

V ≥

(1) Installation losses; and SR

n

(2) The power or equivalent thrust ZW absorbed by the accessories and serv- where: ices appropriate to the particular am- V = Calibrated airspeed obtained when CL MAX bient atmospheric conditions and the the load factor-corrected lift coefficient particular flight condition.

(d) Unless otherwise prescribed, the ⎛ ⎞ n W ZW applicant must select the takeoff, en ⎜ ⎟ route, approach, and landing configura- qS ⎝ ⎠ tions for the airplane.

is first a maximum during the maneuver (e) The airplane configurations may prescribed in paragraph (c) of this section. In vary with weight, altitude, and tem- addition, when the maneuver is limited by a perature, to the extent they are com- device that abruptly pushes the nose down at a selected angle of attack (e.g., a stick push- patible with the operating procedures er), V may not be less than the speed ex- required by paragraph (f) of this sec- CL MAX isting at the instant the device operates; tion.

n = Load factor normal to the flight path ZW (f) Unless otherwise prescribed, in de- at V CL MAX termining the accelerate-stop dis- W = Airplane gross weight; tances, takeoff flight paths, takeoff S = Aerodynamic reference wing area; and distances, and landing distances, q = Dynamic pressure.

changes in the airplane’s configura- (b) V is determined with: CLMAX tion, speed, power, and thrust, must be (1) Engines idling, or, if that result- made in accordance with procedures es- ant thrust causes an appreciable de- tablished by the applicant for oper- crease in stall speed, not more than ation in service.

zero thrust at the stall speed; (g) Procedures for the execution of (2) Propeller pitch controls (if appli- balked landings and missed approaches cable) in the takeoff position; associated with the conditions pre- (3) The airplane in other respects scribed in §§ 25.119 and 25.121(d) must be (such as flaps, landing gear, and ice ac- established.

(h) The procedures established under cretions) in the condition existing in paragraphs (f) and (g) of this section the test or performance standard in must— which V is being used; SR (1) Be able to be consistently exe- (4) The weight used when V is being SR cuted in service by crews of average used as a factor to determine compli- skill; ance with a required performance (2) Use methods or devices that are standard; safe and reliable; and (5) The center of gravity position (3) Include allowance for any time that results in the highest value of ref- delays, in the execution of the proce- erence stall speed; and dures, that may reasonably be expected (6) The airplane trimmed for straight in service. flight at a speed selected by the appli- (i) The accelerate-stop and landing cant, but not less than 1.13V and not SR distances prescribed in §§ 25.109 and greater than 1.3V .

SR Federal Aviation Administration, DOT § 25.107 (c) Starting from the stabilized trim (ii) At the option of the applicant, condition, apply the longitudinal con- grooved or porous friction course wet, trol to decelerate the airplane so that hard-surfaced runways.

the speed reduction does not exceed (2) Smooth water, in the case of sea- one knot per second. planes and amphibians; and (d) In addition to the requirements of (3) Smooth, dry snow, in the case of paragraph (a) of this section, when a skiplanes.

device that abruptly pushes the nose (d) The takeoff data must include, down at a selected angle of attack (e.g., within the established operational lim- a stick pusher) is installed, the ref- its of the airplane, the following oper- erence stall speed, V , may not be less ational correction factors: SR than 2 knots or 2 percent, whichever is (1) Not more than 50 percent of nomi- greater, above the speed at which the nal wind components along the takeoff device operates. path opposite to the direction of take- off, and not less than 150 percent of [Doc. No. 28404, 67 FR 70825, Nov. 26, 2002, as nominal wind components along the amended by Amdt. 25–121, 72 FR 44665, Aug. 8, takeoff path in the direction of takeoff.

2007] (2) Effective runway gradients.

§ 25.105 Takeoff.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) The takeoff speeds prescribed by amended by Amdt. 25–92, 63 FR 8318, Feb. 18, 1998; Amdt. 25–121, 72 FR 44665, Aug. 8, 2007; § 25.107, the accelerate-stop distance Amdt. 25–140, 79 FR 65525, Nov. 4, 2014] prescribed by § 25.109, the takeoff path prescribed by § 25.111, the takeoff dis- § 25.107 Takeoff speeds.

tance and takeoff run prescribed by (a) V must be established in relation § 25.113, and the net takeoff flight path 1 to V as follows: prescribed by § 25.115, must be deter- EF (1) V is the calibrated airspeed at mined in the selected configuration for EF which the critical engine is assumed to takeoff at each weight, altitude, and fail. V must be selected by the appli- ambient temperature within the oper- EF cant, but may not be less than V de- ational limits selected by the appli- MCG termined under § 25.149(e).

cant— (1) In non-icing conditions; and (2) V , in terms of calibrated air- (2) In icing conditions, if in the con- speed, is selected by the applicant; figuration used to show compliance however, V may not be less than V 1 EF with § 25.121(b), and with the most crit- plus the speed gained with critical en- ical of the takeoff ice accretion(s) de- gine inoperative during the time inter- fined in appendices C and O of this val between the instant at which the part, as applicable, in accordance with critical engine is failed, and the in- § 25.21(g): stant at which the pilot recognizes and (i) The stall speed at maximum take- reacts to the engine failure, as indi- off weight exceeds that in non-icing cated by the pilot’s initiation of the conditions by more than the greater of first action (e.g., applying brakes, re- 3 knots CAS or 3 percent of V ; or ducing thrust, deploying speed brakes) SR (ii) The degradation of the gradient to stop the airplane during accelerate- of climb determined in accordance with stop tests.

in terms of calibrated air- § 25.121(b) is greater than one-half of (b) V 2MIN, speed, may not be less than— the applicable actual-to-net takeoff (1) 1.13 V for— flight path gradient reduction defined SR (i) Two-engine and three-engine tur- in § 25.115(b).

bopropeller and reciprocating engine (b) No takeoff made to determine the powered airplanes; and data required by this section may re- (ii) Turbojet powered airplanes with- quire exceptional piloting skill or out provisions for obtaining a signifi- alertness.

(c) The takeoff data must be based cant reduction in the one-engine-inop- on— erative power-on stall speed; (1) In the case of land planes and am- (2) 1.08 V for— SR phibians: (i) Turbopropeller and reciprocating (i) Smooth, dry and wet, hard-sur- engine powered airplanes with more faced runways; and than three engines; and 14 CFR Ch. I (1–1–25 Edition) § 25.109 (ii) Turbojet powered airplanes with tained in accordance with this para- provisions for obtaining a significant graph, must be used to show compli- reduction in the one-engine-inoper- ance with both the one-engine-inoper- ative power-on stall speed; and ative and the all-engines-operating (3) 1.10 times V established under takeoff provisions.

MC § 25.149. (3) It must be shown that the one-en- (c) V , in terms of calibrated air- gine-inoperative takeoff distance, speed, must be selected by the appli- using a rotation speed of 5 knots less cant to provide at least the gradient of than V established in accordance with R climb required by § 25.121(b) but may paragraphs (e)(1) and (2) of this section, not be less than— does not exceed the corresponding one- (1) V ; engine-inoperative takeoff distance 2MIN (2) V plus the speed increment at- using the established V . The takeoff R R tained (in accordance with § 25.111(c)(2)) distances must be determined in ac- before reaching a height of 35 feet cordance with § 25.113(a)(1).

above the takeoff surface; and (4) Reasonably expected variations in (3) A speed that provides the maneu- service from the established takeoff vering capability specified in § 25.143(h). procedures for the operation of the air- (d) V is the calibrated airspeed at plane (such as over-rotation of the air- MU and above which the airplane can safe- plane and out-of-trim conditions) may ly lift off the ground, and con- tinue not result in unsafe flight characteris- the takeoff. V speeds must be se- tics or in marked increases in the MU lected by the applicant throughout the scheduled takeoff distances established range of thrust-to-weight ratios to be in accordance with § 25.113(a).

certificated. These speeds may be es- (f) V is the calibrated airspeed at LOF tablished from free air data if these which the airplane first becomes air- data are verified by ground takeoff borne.

tests.

(g) V , in terms of calibrated air- FTO (e) V in terms of calibrated air- R, speed, must be selected by the appli- speed, must be selected in accordance cant to provide at least the gradient of with the conditions of paragraphs (e)(1) climb required by § 25.121(c), but may through (4) of this section: not be less than— (1) V may not be less than— R (1) 1.18 V ; and SR (i) V ; 1 (2) A speed that provides the maneu- (ii) 105 percent of V ; MC vering capability specified in § 25.143(h).

(iii) The speed (determined in accord- (h) In determining the takeoff speeds ance with § 25.111(c)(2)) that allows V , V , and V for flight in icing condi- 1 R 2 reaching V before reaching a height of 2 tions, the values of V , V , and V MCG MC MU 35 feet above the takeoff surface; or determined for non-icing conditions (iv) A speed that, if the airplane is may be used.

rotated at its maximum practicable [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as rate, will result in a V of not less LOF amended by Amdt. 25–38, 41 FR 55466, Dec. 20, than — 1976; Amdt. 25–42, 43 FR 2320, Jan. 16, 1978; (A) 110 percent of V in the all-en- MU Amdt. 25–92, 63 FR 8318, Feb. 18, 1998; Amdt.

gines-operating condition, and 105 per- 25–94, 63 FR 8848, Feb. 23, 1998; Amdt. 25–108, cent of V determined at the thrust- MU 67 FR 70826, Nov. 26, 2002; Amdt. 25–121, 72 FR to-weight ratio corresponding to the 44665, Aug. 8, 2007; Amdt. 25–135, 76 FR 74654, Dec. 1, 2011] one-engine-inoperative condition; or (B) If the V attitude is limited by MU § 25.109 Accelerate-stop distance.

the geometry of the airplane ( i.e., tail contact with the runway), 108 percent (a) The accelerate-stop distance on a of V in the all-engines-operating dry runway is the greater of the fol- MU condition, and 104 percent of V deter- lowing distances: MU mined at the thrust-to-weight ratio (1) The sum of the distances nec- corresponding to the one-engine-inop- essary to— erative condition. (i) Accelerate the airplane from a (2) For any given set of conditions standing start with all engines oper- (such as weight, configuration, and ating to V for takeoff from a dry run- EF temperature), a single value of V ob- way; R, Federal Aviation Administration, DOT § 25.109 (ii) Allow the airplane to accelerate (1) The accelerate-stop distance on a from V to the highest speed reached dry runway determined in accordance EF during the rejected takeoff, assuming with paragraph (a) of this section; or the critical engine fails at V and the (2) The accelerate-stop distance de- EF pilot takes the first action to reject termined in accordance with paragraph the takeoff at the V for takeoff from a (a) of this section, except that the run- dry runway; and way is wet and the corresponding wet (iii) Come to a full stop on a dry run- runway values of V and V are used.

EF 1 way from the speed reached as pre- In determining the wet runway accel- scribed in paragraph (a)(1)(ii) of this erate-stop distance, the stopping force section; plus from the wheel brakes may never ex- (iv) A distance equivalent to 2 sec- ceed: onds at the V for takeoff from a dry (i) The wheel brakes stopping force runway. determined in meeting the require- (2) The sum of the distances nec- ments of § 25.101(i) and paragraph (a) of essary to— this section; and (i) Accelerate the airplane from a (ii) The force resulting from the wet standing start with all engines oper- runway braking coefficient of friction ating to the highest speed reached dur- determined in accordance with para- ing the rejected takeoff, assuming the graphs (c) or (d) of this section, as ap- pilot takes the first action to reject plicable, taking into account the dis- the takeoff at the V for takeoff from a tribution of the normal load between dry runway; and braked and unbraked wheels at the (ii) With all engines still operating, most adverse center-of-gravity position come to a full stop on dry runway from approved for takeoff.

the speed reached as prescribed in para- (c) The wet runway braking coeffi- graph (a)(2)(i) of this section; plus cient of friction for a smooth wet run- (iii) A distance equivalent to 2 sec- way is defined as a curve of friction co- onds at the V for takeoff from a dry efficient versus ground speed and must runway. be computed as follows: (b) The accelerate-stop distance on a (1) The maximum tire-to-ground wet wet runway is the greater of the fol- runway braking coefficient of friction lowing distances: is defined as: Where— (2) The maximum tire-to-ground wet Tire Pressure = maximum airplane operating runway braking coefficient of friction tire pressure (psi); must be adjusted to take into account = maximum tire-to-ground braking μ t/gMAX the efficiency of the anti-skid system coefficient; on a wet runway. Anti-skid system op- V = airplane true ground speed (knots); and eration must be demonstrated by flight Linear interpolation may be used for tire pressures other than those listed. testing on a smooth wet runway, and 14 CFR Ch. I (1–1–25 Edition) § 25.109 its efficiency must be determined. Un- of friction may be used for runway sur- less a specific anti-skid system effi- faces that have been grooved or treated ciency is determined from a quan- with a porous friction course material.

titative analysis of the flight testing For grooved and porous friction course on a smooth wet runway, the max- runways, the wet runway braking imum tire-to-ground wet runway brak- coefficent of friction is defined as ei- ing coefficient of friction determined ther: in paragraph (c)(1) of this section must (1) 70 percent of the dry runway brak- be multiplied by the efficiency value ing coefficient of friction used to deter- associated with the type of anti-skid mine the dry runway accelerate-stop system installed on the airplane: distance; or Effi- (2) The wet runway braking coeffi- Type of anti-skid system ciency cient defined in paragraph (c) of this value section, except that a specific anti-skid On-Off ........................................................................ 0.30 system efficiency, if determined, is ap- Quasi-Modulating ....................................................... 0.50 propriate for a grooved or porous fric- Fully Modulating ........................................................ 0.80 tion course wet runway, and the max- (d) At the option of the applicant, a imum tire-to-ground wet runway brak- higher wet runway braking coefficient ing coefficient of friction is defined as: Where— (1) Shall not be included as an addi- Tire Pressure = maximum airplane operating tional means of deceleration when de- tire pressure (psi); termining the accelerate-stop distance μ = maximum tire-to-ground braking t/gMAX on a dry runway; and coefficient; (2) May be included as an additional V = airplane true ground speed (knots); and means of deceleration using rec- Linear interpolation may be used for tire ommended reverse thrust procedures pressures other than those listed.

when determining the accelerate-stop (e) Except as provided in paragraph distance on a wet runway, provided the (f)(1) of this section, means other than requirements of paragraph (e) of this wheel brakes may be used to determine section are met.

the accelerate-stop distance if that (g) The landing gear must remain ex- means— tended throughout the accelerate-stop (1) Is safe and reliable; distance.

(2) Is used so that consistent results (h) If the accelerate-stop distance in- can be expected under normal oper- cludes a stopway with surface charac- ating conditions; and teristics substantially different from (3) Is such that exceptional skill is those of the runway, the takeoff data not required to control the airplane.

must include operational correction (f) The effects of available reverse thrust— Federal Aviation Administration, DOT § 25.111 factors for the accelerate-stop dis- (i) 1.2 percent for two-engine air- tance. The correction factors must ac- planes; count for the particular surface charac- (ii) 1.5 percent for three-engine air- teristics of the stopway and the vari- planes; and ations in these characteristics with (iii) 1.7 percent for four-engine air- seasonal weather conditions (such as planes.

temperature, rain, snow, and ice) with- (4) The airplane configuration may in the established operational limits.

not be changed, except for gear retrac- (i) A flight test demonstration of the tion and automatic propeller feath- maximum brake kinetic energy accel- ering, and no change in power or thrust erate-stop distance must be conducted that requires action by the pilot may with not more than 10 percent of the be made until the airplane is 400 feet allowable brake wear range remaining above the takeoff surface; and on each of the airplane wheel brakes.

(5) If § 25.105(a)(2) requires the takeoff [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as path to be determined for flight in amended by Amdt. 25–42, 43 FR 2321, Jan. 16, icing conditions, the airborne part of 1978; Amdt. 25–92, 63 FR 8318, Feb. 18, 1998] the takeoff must be based on the air- plane drag: § 25.111 Takeoff path.

(i) With the most critical of the take- (a) The takeoff path extends from a off ice accretion(s) defined in Appen- standing start to a point in the takeoff dices C and O of this part, as applica- at which the airplane is 1,500 feet above ble, in accordance with § 25.21(g), from the takeoff surface, or at which the a height of 35 feet above the takeoff transition from the takeoff to the en surface up to the point where the air- route configuration is completed and plane is 400 feet above the takeoff sur- V is reached, whichever point is FTO face; and higher. In addition— (ii) With the most critical of the final (1) The takeoff path must be based on takeoff ice accretion(s) defined in Ap- the procedures prescribed in § 25.101(f); pendices C and O of this part, as appli- (2) The airplane must be accelerated cable, in accordance with § 25.21(g), on the ground to V at which point EF, from the point where the airplane is 400 the critical engine must be made inop- feet above the takeoff surface to the erative and remain inoperative for the end of the takeoff path.

rest of the takeoff; and (d) The takeoff path must be deter- (3) After reaching V the airplane EF, mined by a continuous demonstrated must be accelerated to V .

takeoff or by synthesis from segments.

(b) During the acceleration to speed If the takeoff path is determined by the V , the nose gear may be raised off the segmental method— ground at a speed not less than V .

R (1) The segments must be clearly de- However, landing gear retraction may fined and must be related to the dis- not be begun until the airplane is air- tinct changes in the configuration, borne.

power or thrust, and speed; (c) During the takeoff path deter- (2) The weight of the airplane, the mination in accordance with para- configuration, and the power or thrust graphs (a) and (b) of this section— must be constant throughout each seg- (1) The slope of the airborne part of ment and must correspond to the most the takeoff path must be positive at critical condition prevailing in the seg- each point; ment; (2) The airplane must reach V before (3) The flight path must be based on it is 35 feet above the takeoff surface the airplane’s performance without and must continue at a speed as close ground effect; and as practical to, but not less than V , until it is 400 feet above the takeoff (4) The takeoff path data must be surface; checked by continuous demonstrated (3) At each point along the takeoff takeoffs up to the point at which the path, starting at the point at which the airplane is out of ground effect and its airplane reaches 400 feet above the speed is stabilized, to ensure that the takeoff surface, the available gradient path is conservative relative to the of climb may not be less than— continous path.

14 CFR Ch. I (1–1–25 Edition) § 25.113 The airplane is considered to be out of mined under § 25.111 for a dry runway; the ground effect when it reaches a or height equal to its wing span. (ii) 115 percent of the horizontal dis- (e) For airplanes equipped with tance along the takeoff path, with all engines operating, from the start of the standby power rocket engines, the takeoff to a point equidistant between takeoff path may be determined in ac- the point at which V is reached and cordance with section II of appendix E. LOF the point at which the airplane is 35 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as feet above the takeoff surface, deter- amended by Amdt. 25–6, 30 FR 8468, July 2, mined by a procedure consistent with 1965; Amdt. 25–42, 43 FR 2321, Jan. 16, 1978; § 25.111.

Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; Amdt.

(2) The takeoff run on a wet runway 25–72, 55 FR 29774, July 20, 1990; Amdt. 25–94, is the greater of— 63 FR 8848, Feb. 23, 1998; Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; Amdt. 25–115, 69 FR 40527, (i) The horizontal distance along the July 2, 2004; Amdt. 25–121, 72 FR 44666; Aug.

takeoff path from the start of the take- 8, 2007; Amdt. 25–140, 79 FR 65525, Nov. 4, 2014] off to the point at which the airplane is 15 feet above the takeoff surface, § 25.113 Takeoff distance and takeoff achieved in a manner consistent with run.

the achievement of V before reaching (a) Takeoff distance on a dry runway 35 feet above the takeoff surface, as de- is the greater of— termined under § 25.111 for a wet run- (1) The horizontal distance along the way; or takeoff path from the start of the take- (ii) 115 percent of the horizontal dis- off to the point at which the airplane is tance along the takeoff path, with all 35 feet above the takeoff surface, deter- engines operating, from the start of the mined under § 25.111 for a dry runway; takeoff to a point equidistant between or the point at which V is reached and LOF (2) 115 percent of the horizontal dis- the point at which the airplane is 35 tance along the takeoff path, with all feet above the takeoff surface, deter- engines operating, from the start of the mined by a procedure consistent with takeoff to the point at which the air- § 25.111.

plane is 35 feet above the takeoff sur- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as face, as determined by a procedure con- amended by Amdt. 25–23, 35 FR 5671, Apr. 8, sistent with § 25.111.

1970; Amdt. 25–92, 63 FR 8320, Feb. 18, 1998] (b) Takeoff distance on a wet runway is the greater of— § 25.115 Takeoff flight path.

(1) The takeoff distance on a dry run- (a) The takeoff flight path shall be way determined in accordance with considered to begin 35 feet above the paragraph (a) of this section; or takeoff surface at the end of the take- (2) The horizontal distance along the off distance determined in accordance takeoff path from the start of the take- with § 25.113(a) or (b), as appropriate for off to the point at which the airplane is the runway surface condition.

15 feet above the takeoff surface, (b) The net takeoff flight path data achieved in a manner consistent with must be determined so that they rep- the achievement of V before reaching resent the actual takeoff flight paths 35 feet above the takeoff surface, deter- (determined in accordance with § 25.111 mined under § 25.111 for a wet runway.

and with paragraph (a) of this section) (c) If the takeoff distance does not in- reduced at each point by a gradient of clude a clearway, the takeoff run is climb equal to— equal to the takeoff distance. If the (1) 0.8 percent for two-engine air- takeoff distance includes a clearway— planes; (1) The takeoff run on a dry runway (2) 0.9 percent for three-engine air- is the greater of— planes; and (i) The horizontal distance along the (3) 1.0 percent for four-engine air- takeoff path from the start of the take- planes.

off to a point equidistant between the (c) The prescribed reduction in climb point at which V is reached and the gradient may be applied as an equiva- LOF point at which the airplane is 35 feet lent reduction in acceleration along above the takeoff surface, as deter- that part of the takeoff flight path at Federal Aviation Administration, DOT § 25.121 which the airplane is accelerated in ical power operating condition existing level flight. later along the flight path but before the point at which the landing gear is [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as fully retracted; and amended by Amdt. 25–92, 63 FR 8320, Feb. 18, (2) The weight equal to the weight 1998] existing when retraction of the landing § 25.117 Climb: general.

gear is begun, determined under § 25.111.

Compliance with the requirements of §§ 25.119 and 25.121 must be shown at (b) Takeoff; landing gear retracted. In each weight, altitude, and ambient the takeoff configuration existing at temperature within the operational the point of the flight path at which limits established for the airplane and the landing gear is fully retracted, and with the most unfavorable center of in the configuration used in § 25.111 but gravity for each configuration.

without ground effect: (1) The steady gradient of climb may § 25.119 Landing climb: All-engines-op- not be less than 2.4 percent for two-en- erating.

gine airplanes, 2.7 percent for three-en- In the landing configuration, the gine airplanes, and 3.0 percent for four- steady gradient of climb may not be engine airplanes, at V with: less than 3.2 percent, with the engines (i) The critical engine inoperative, at the power or thrust that is available the remaining engines at the takeoff 8 seconds after initiation of movement power or thrust available at the time of the power or thrust controls from the landing gear is fully retracted, de- the minimum flight idle to the go- termined under § 25.111, unless there is around power or thrust setting— a more critical power operating condi- (a) In non-icing conditions, with a tion existing later along the flight path climb speed of V determined in ac- REF but before the point where the airplane cordance with § 25.125(b)(2)(i); and reaches a height of 400 feet above the (b) In icing conditions with the most takeoff surface; and critical of the landing ice accretion(s) (ii) The weight equal to the weight defined in Appendices C and O of this existing when the airplane’s landing part, as applicable, in accordance with gear is fully retracted, determined § 25.21(g), and with a climb speed of under § 25.111.

V determined in accordance with REF (2) The requirements of paragraph § 25.125(b)(2)(ii).

(b)(1) of this section must be met: [Amdt. 25–121, 72 FR 44666; Aug. 8, 2007, as (i) In non-icing conditions; and amended by Amdt. 25–,140, 79 FR 65525, Nov.

(ii) In icing conditions with the most 4, 2014] critical of the takeoff ice accretion(s) § 25.121 Climb: One-engine-inoper- defined in Appendices C and O of this ative.

part, as applicable, in accordance with § 25.21(g), if in the configuration used to (a) Takeoff; landing gear extended. In show compliance with § 25.121(b) with the critical takeoff configuration exist- this takeoff ice accretion: ing along the flight path (between the (A) The stall speed at maximum points at which the airplane reaches takeoff weight exceeds that in non- V and at which the landing gear is LOF icing conditions by more than the fully retracted) and in the configura- greater of 3 knots CAS or 3 percent of tion used in § 25.111 but without ground V ; or SR effect, the steady gradient of climb (B) The degradation of the gradient must be positive for two-engine air- of climb determined in accordance with planes, and not less than 0.3 percent for § 25.121(b) is greater than one-half of three-engine airplanes or 0.5 percent the applicable actual-to-net takeoff for four-engine airplanes, at V and LOF flight path gradient reduction defined with— in § 25.115(b).

(1) The critical engine inoperative and the remaining engines at the power (c) Final takeoff. In the en route con- or thrust available when retraction of figuration at the end of the takeoff the landing gear is begun in accordance path determined in accordance with with § 25.111 unless there is a more crit- § 25.111: 14 CFR Ch. I (1–1–25 Edition) § 25.123 (1) The steady gradient of climb may defined in Appendices C and O of this not be less than 1.2 percent for two-en- part, as applicable, in accordance with gine airplanes, 1.5 percent for three-en- § 25.21(g). The climb speed selected for gine airplanes, and 1.7 percent for four- non-icing conditions may be used if the engine airplanes, at V with— climb speed for icing conditions, com- FTO (i) The critical engine inoperative puted in accordance with paragraph and the remaining engines at the avail- (d)(1)(iii) of this section, does not ex- able maximum continuous power or ceed that for non-icing conditions by thrust; and more than the greater of 3 knots CAS (ii) The weight equal to the weight or 3 percent.

existing at the end of the takeoff path, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as determined under § 25.111.

amended by Amdt. 25–84, 60 FR 30749, June 9, (2) The requirements of paragraph 1995; Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; (c)(1) of this section must be met: Amdt. 25–121, 72 FR 44666; Aug. 8, 2007; Amdt.

(i) In non-icing conditions; and 25–140, 79 FR 65525, Nov. 4, 2014] (ii) In icing conditions with the most critical of the final takeoff ice accre- § 25.123 En route flight paths.

tion(s) defined in Appendices C and O of (a) For the en route configuration, this part, as applicable, in accordance the flight paths prescribed in para- with § 25.21(g), if in the configuration graph (b) and (c) of this section must used to show compliance with § 25.121(b) be determined at each weight, altitude, with the takeoff ice accretion used to and ambient temperature, within the show compliance with § 25.111(c)(5)(i): operating limits established for the (A) The stall speed at maximum airplane. The variation of weight along takeoff weight exceeds that in non- the flight path, accounting for the pro- icing conditions by more than the gressive consumption of fuel and oil by greater of 3 knots CAS or 3 percent of the operating engines, may be included V ; or SR in the computation. The flight paths (B) The degradation of the gradient must be determined at a speed not less of climb determined in accordance with than V , with— FTO § 25.121(b) is greater than one-half of (1) The most unfavorable center of the applicable actual-to-net takeoff gravity; flight path gradient reduction defined (2) The critical engines inoperative; in § 25.115(b).

(3) The remaining engines at the (d) Approach. In a configuration cor- available maximum continuous power responding to the normal all-engines- or thrust; and operating procedure in which V for SR (4) The means for controlling the en- this configuration does not exceed 110 gine-cooling air supply in the position percent of the V for the related all- SR that provides adequate cooling in the engines-operating landing configura- hot-day condition.

tion: (b) The one-engine-inoperative net (1) The steady gradient of climb may flight path data must represent the ac- not be less than 2.1 percent for two-en- tual climb performance diminished by gine airplanes, 2.4 percent for three-en- a gradient of climb of 1.1 percent for gine airplanes, and 2.7 percent for four- two-engine airplanes, 1.4 percent for engine airplanes, with— three-engine airplanes, and 1.6 percent (i) The critical engine inoperative, for four-engine airplanes— the remaining engines at the go-around (1) In non-icing conditions; and power or thrust setting; (2) In icing conditions with the most (ii) The maximum landing weight; critical of the en route ice accretion(s) (iii) A climb speed established in con- defined in Appendices C and O of this nection with normal landing proce- part, as applicable, in accordance with dures, but not exceeding 1.4 V ; and SR § 25.21(g), if: (iv) Landing gear retracted.

(2) The requirements of paragraph (i) A speed of 1.18 ‘‘V with the en SR0 (d)(1) of this section must be met: route ice accretion exceeds the en (i) In non-icing conditions; and route speed selected for non-icing con- (ii) In icing conditions with the most ditions by more than the greater of 3 critical of the approach ice accretion(s) knots CAS or 3 percent of V ; or SR Federal Aviation Administration, DOT § 25.125 (ii) The degradation of the gradient that speed exceeds V selected for REF of climb is greater than one-half of the non-icing conditions by more than 5 applicable actual-to-net flight path re- knots CAS; and duction defined in paragraph (b) of this (C) A speed that provides the maneu- section.

vering capability specified in § 25.143(h) (c) For three- or four-engine air- with the most critical of the landing planes, the two-engine-inoperative net ice accretion(s) defined in Appendices C flight path data must represent the ac- and O of this part, as applicable, in ac- tual climb performance diminished by cordance with § 25.21(g).

a gradient of climb of 0.3 percent for (3) Changes in configuration, power three-engine airplanes and 0.5 percent or thrust, and speed, must be made in for four-engine airplanes.

accordance with the established proce- dures for service operation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (4) The landing must be made with- amended by Amdt. 25–121, 72 FR 44666; Aug. 8, 2007; Amdt. 25–140, 79 FR 65525, Nov. 4, 2014] out excessive vertical acceleration, tendency to bounce, nose over, ground § 25.125 Landing.

loop, porpoise, or water loop.

(a) The horizontal distance necessary (5) The landings may not require ex- to land and to come to a complete stop ceptional piloting skill or alertness.

(or to a speed of approximately 3 knots (c) For landplanes and amphibians, for water landings) from a point 50 feet the landing distance on land must be above the landing surface must be de- determined on a level, smooth, dry, termined (for standard temperatures, hard-surfaced runway. In addition— at each weight, altitude, and wind (1) The pressures on the wheel brak- within the operational limits estab- ing systems may not exceed those spec- lished by the applicant for the air- ified by the brake manufacturer; plane): (2) The brakes may not be used so as (1) In non-icing conditions; and to cause excessive wear of brakes or (2) In icing conditions with the most tires; and critical of the landing ice accretion(s) (3) Means other than wheel brakes defined in Appendices C and O of this may be used if that means— part, as applicable, in accordance with (i) Is safe and reliable; § 25.21(g), if V for icing conditions ex- REF (ii) Is used so that consistent results ceeds V for non-icing conditions by REF can be expected in service; and more than 5 knots CAS at the max- (iii) Is such that exceptional skill is imum landing weight.

not required to control the airplane.

(b) In determining the distance in (d) For seaplanes and amphibians, paragraph (a) of this section: the landing distance on water must be (1) The airplane must be in the land- determined on smooth water.

ing configuration.

(e) For skiplanes, the landing dis- (2) A stabilized approach, with a cali- tance on snow must be determined on brated airspeed of not less than V , REF smooth, dry, snow.

must be maintained down to the 50-foot (f) The landing distance data must height.

include correction factors for not more (i) In non-icing conditions, V may REF than 50 percent of the nominal wind not be less than: components along the landing path op- (A) 1.23 V 0; SR posite to the direction of landing, and (B) V established under § 25.149(f); MCL not less than 150 percent of the nomi- and nal wind components along the landing (C) A speed that provides the maneu- path in the direction of landing.

vering capability specified in § 25.143(h).

(ii) In icing conditions, V may not (g) If any device is used that depends REF be less than: on the operation of any engine, and if (A) The speed determined in para- the landing distance would be notice- graph (b)(2)(i) of this section; ably increased when a landing is made (B) 1.23 V with the most critical of with that engine inoperative, the land- SR0 the landing ice accretion(s) defined in ing distance must be determined with Appendices C and O of this part, as ap- that engine inoperative unless the use plicable, in accordance with § 25.21(g), if of compensating means will result in a 14 CFR Ch. I (1–1–25 Edition) § 25.143 landing distance not more than that Force, in pounds, applied to the Pitch Roll Yaw control wheel or rudder pedals with each engine operating.

For short term application for [Amdt. 25–121, 72 FR 44666; Aug. 8, 2007; 72 FR pitch and roll control—two 50467, Aug. 31, 2007; Amdt. 25–140, 79 FR 65525, hands available for control .... 75 50 Nov. 4, 2014] For short term application for pitch and roll control—one C ONTROLLABILITY AND hand available for control ...... 50 25 M ANEUVERABILITY For short term application for § 25.143 General.

For long term application .......... 10 5 20 (a) The airplane must be safely con- (e) Approved operating procedures or trollable and maneuverable during— conventional operating practices must (1) Takeoff; be followed when demonstrating com- (2) Climb; (3) Level flight; pliance with the control force limita- (4) Descent; and tions for short term application that (5) Landing.

are prescribed in paragraph (d) of this (b) It must be possible to make a section. The airplane must be in trim, smooth transition from one flight con- or as near to being in trim as practical, dition to any other flight condition in the preceding steady flight condi- without exceptional piloting skill, tion. For the takeoff condition, the air- alertness, or strength, and without plane must be trimmed according to danger of exceeding the airplane limit- the approved operating procedures.

load factor under any probable oper- (f) When demonstrating compliance ating conditions, including— with the control force limitations for (1) The sudden failure of the critical long term application that are pre- engine; scribed in paragraph (d) of this section, (2) For airplanes with three or more the airplane must be in trim, or as near engines, the sudden failure of the sec- to being in trim as practical.

ond critical engine when the airplane is (g) When maneuvering at a constant in the en route, approach, or landing airspeed or Mach number (up to V / configuration and is trimmed with the FC M ), the stick forces and the gradient critical engine inoperative; and FC of the stick force versus maneuvering (3) Configuration changes, including deployment or retraction of decelera- load factor must lie within satisfactory tion devices.

limits. The stick forces must not be so (c) The airplane must be shown to be great as to make excessive demands on safely controllable and maneuverable the pilot’s strength when maneuvering with the most critical of the ice accre- the airplane, and must not be so low tion(s) appropriate to the phase of that the airplane can easily be over- flight as defined in Appendices C and O stressed inadvertently. Changes of gra- of this part, as applicable, in accord- dient that occur with changes of load ance with § 25.21(g), and with the crit- factor must not cause undue difficulty ical engine inoperative and its pro- in maintaining control of the airplane, peller (if applicable) in the minimum and local gradients must not be so low drag position: as to result in a danger of overcontrol- (1) At the minimum V for takeoff; ling.

(2) During an approach and go- (h) The maneuvering capabilities in a around; and constant speed coordinated turn at for- (3) During an approach and landing.

ward center of gravity, as specified in (d) The following table prescribes, for the following table, must be free of conventional wheel type controls, the stall warning or other characteristics maximum control forces permitted during the testing required by para- that might interfere with normal ma- graph (a) through (c) of this section: neuvering: Maneuvering Configuration Speed bank angle in a Thrust/power setting coordinated turn Takeoff .......... V 30 ° Asymmetric WAT-Limited.

Federal Aviation Administration, DOT § 25.145 Maneuvering Configuration Speed bank angle in a Thrust/power setting coordinated turn 2 3 Takeoff .......... V + XX 40 ° All-engines-operating climb.

En route ........ V 40 ° Asymmetric WAT-Limited.

FTO Landing ......... V 40 ° Symmetric for ¥ 3 ° flight path angle.

REF A combination of weight, altitude, and temperature (WAT) such that the thrust or power setting produces the minimum climb gradient specified in § 25.121 for the flight condition.

Airspeed approved for all-engines-operating initial climb.

That thrust or power setting which, in the event of failure of the critical engine and without any crew action to adjust the thrust or power of the remaining engines, would result in the thrust or power specified for the takeoff condition at V , or any lesser thrust or power setting that is used for all-engines-operating initial climb procedures.

(i) When demonstrating compliance (2) There is no pitch control force re- with § 25.143 in icing conditions— versal during a pushover maneuver down to 0.5 g load factor.

(1) Controllability must be dem- onstrated with the most critical of the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ice accretion(s) for the particular amended by Amdt. 25–42, 43 FR 2321, Jan. 16, flight phase as defined in Appendices C 1978; Amdt. 25–84, 60 FR 30749, June 9, 1995; Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; and O of this part, as applicable, in ac- Amdt. 25–121, 72 FR 44667, Aug. 8, 2007; Amdt.

cordance with § 25.21(g); 25–129, 74 FR 38339, Aug. 3, 2009; Amdt. 25–140, (2) It must be shown that a push force 79 FR 65525, Nov. 4, 2014] is required throughout a pushover ma- neuver down to a zero g load factor, or § 25.145 Longitudinal control.

the lowest load factor obtainable if (a) It must be possible, at any point limited by elevator power or other de- between the trim speed prescribed in sign characteristic of the flight control § 25.103(b)(6) and stall identification (as system. It must be possible to prompt- defined in § 25.201(d)), to pitch the nose ly recover from the maneuver without downward so that the acceleration to exceeding a pull control force of 50 this selected trim speed is prompt with pounds; and (1) The airplane trimmed at the trim (3) Any changes in force that the speed prescribed in § 25.103(b)(6); pilot must apply to the pitch control to (2) The landing gear extended; maintain speed with increasing sideslip (3) The wing flaps (i) retracted and angle must be steadily increasing with (ii) extended; and no force reversals, unless the change in (4) Power (i) off and (ii) at maximum control force is gradual and easily con- continuous power on the engines.

trollable by the pilot without using ex- (b) With the landing gear extended, ceptional piloting skill, alertness, or no change in trim control, or exertion strength.

of more than 50 pounds control force (j) For flight in icing conditions be- (representative of the maximum short fore the ice protection system has been term force that can be applied readily activated and is performing its in- by one hand) may be required for the tended function, it must be dem- following maneuvers: onstrated in flight with the most crit- (1) With power off, flaps retracted, ical of the ice accretion(s) defined in and the airplane trimmed at 1.3 V , SR1 Appendix C, part II, paragraph (e) of extend the flaps as rapidly as possible this part and Appendix O, part II, para- while maintaining the airspeed at ap- graph (d) of this part, as applicable, in proximately 30 percent above the ref- accordance with § 25.21(g), that: erence stall speed existing at each in- (1) The airplane is controllable in a stant throughout the maneuver.

pull-up maneuver up to 1.5 g load fac- (2) Repeat paragraph (b)(1) except ini- tor; and tially extend the flaps and then retract them as rapidly as possible.

(3) Repeat paragraph (b)(2), except at the go-around power or thrust setting.

(4) With power off, flaps retracted, and the airplane trimmed at 1.3 V , SR1 14 CFR Ch. I (1–1–25 Edition) § 25.147 rapidly set go-around power or thrust and distinct motion once the control while maintaining the same airspeed. has reached the gated position.

(5) Repeat paragraph (b)(4) except [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as with flaps extended.

amended by Amdt. 25–23, 35 FR 5671, Apr. 8, (6) With power off, flaps extended, 1970; Amdt. 25–72, 55 FR 29774, July 20, 1990; Amdt. 25–84, 60 FR 30749, June 9, 1995; Amdt.

and the airplane trimmed at 1.3 V , SR1 25–98, 64 FR 6164, Feb. 8, 1999; 64 FR 10740, obtain and maintain airspeeds between Mar. 5, 1999; Amdt. 25–108, 67 FR 70827, Nov.

V and either 1.6 V or V , which- SW SR1 FE 26, 2002] ever is lower.

§ 25.147 Directional and lateral con- (c) It must be possible, without ex- trol.

ceptional piloting skill, to prevent loss of altitude when complete retraction of (a) Directional control; general. It must be possible, with the wings level, to the high lift devices from any position yaw into the operative engine and to is begun during steady, straight, level safely make a reasonably sudden flight at 1.08 V for propeller powered SR1 change in heading of up to 15 degrees in airplanes, or 1.13 V for turbojet pow- SR1 the direction of the critical inoperative ered airplanes, with— engine. This must be shown at 1.3 V R1 S (1) Simultaneous movement of the for heading changes up to 15 degrees power or thrust controls to the go- (except that the heading change at around power or thrust setting; which the rudder pedal force is 150 (2) The landing gear extended; and pounds need not be exceeded), and (3) The critical combinations of land- with— ing weights and altitudes.

(1) The critical engine inoperative (d) If gated high-lift device control and its propeller in the minimum drag positions are provided, paragraph (c) of position; this section applies to retractions of (2) The power required for level flight the high-lift devices from any position at 1.3 V R1, but not more than max- S imum continuous power; from the maximum landing position to (3) The most unfavorable center of the first gated position, between gated gravity; positions, and from the last gated posi- (4) Landing gear retracted; tion to the fully retracted position.

(5) Flaps in the approach position; The requirements of paragraph (c) of and this section also apply to retractions (6) Maximum landing weight.

from each approved landing position to (b) Directional control; airplanes with the control position(s) associated with four or more engines. Airplanes with the high-lift device configuration(s) four or more engines must meet the re- used to establish the go-around proce- quirements of paragraph (a) of this sec- dure(s) from that landing position. In tion except that— addition, the first gated control posi- (1) The two critical engines must be tion from the maximum landing posi- inoperative with their propellers (if ap- tion must correspond with a configura- plicable) in the minimum drag posi- tion of the high-lift devices used to es- tion; (2) [Reserved] tablish a go-around procedure from a (3) The flaps must be in the most fa- landing configuration. Each gated con- vorable climb position.

trol position must require a separate (c) Lateral control; general. It must be and distinct motion of the control to possible to make 20 ° banked turns, with pass through the gated position and and against the inoperative engine, must have features to prevent inad- from steady flight at a speed equal to vertent movement of the control 1.3 V R1, with— S through the gated position. It must (1) The critical engine inoperative only be possible to make this separate and its propeller (if applicable) in the minimum drag position; (2) The remaining engines at max- imum continuous power; (3) The most unfavorable center of gravity; Federal Aviation Administration, DOT § 25.149 (4) Landing gear (i) retracted and (ii) (c) V may not exceed 1.13 V MC SR extended; with— (1) Maximum available takeoff power (5) Flaps in the most favorable climb or thrust on the engines; position; and (2) The most unfavorable center of (6) Maximum takeoff weight.

gravity; (d) Lateral control; roll capability. With (3) The airplane trimmed for takeoff; the critical engine inoperative, roll re- (4) The maximum sea level takeoff sponse must allow normal maneuvers.

weight (or any lesser weight necessary Lateral control must be sufficient, at to show V ); MC the speeds likely to be used with one (5) The airplane in the most critical engine inoperative, to provide a roll takeoff configuration existing along rate necessary for safety without ex- the flight path after the airplane be- cessive control forces or travel.

comes airborne, except with the land- (e) Lateral control; airplanes with four ing gear retracted; or more engines. Airplanes with four or (6) The airplane airborne and the more engines must be able to make 20 ° ground effect negligible; and banked turns, with and against the in- (7) If applicable, the propeller of the operative engines, from steady flight at inoperative engine— a speed equal to 1.3 V R1, with max- S (i) Windmilling; imum continuous power, and with the (ii) In the most probable position for airplane in the configuration pre- the specific design of the propeller con- scribed by paragraph (b) of this section.

trol; or (f) Lateral control; all engines oper- (iii) Feathered, if the airplane has an ating. With the engines operating, roll automatic feathering device acceptable response must allow normal maneuvers for showing compliance with the climb (such as recovery from upsets produced requirements of § 25.121.

by gusts and the initiation of evasive (d) The rudder forces required to maneuvers). There must be enough ex- maintain control at V may not ex- MC cess lateral control in sideslips (up to ceed 150 pounds nor may it be nec- sideslip angles that might be required essary to reduce power or thrust of the in normal operation), to allow a lim- operative engines. During recovery, the ited amount of maneuvering and to airplane may not assume any dan- correct for gusts. Lateral control must gerous attitude or require exceptional be enough at any speed up to V / M FC FC piloting skill, alertness, or strength to to provide a peak roll rate necessary prevent a heading change of more than for safety, without excessive control 20 degrees.

forces or travel.

(e) V , the minimum control speed MCG on the ground, is the calibrated air- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–42, 43 FR 2321, Jan. 16, speed during the takeoff run at which, 1978; Amdt. 25–72, 55 FR 29774, July 20, 1990; when the critical engine is suddenly Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; made inoperative, it is possible to Amdt. 25–115, 69 FR 40527, July 2, 2004] maintain control of the airplane using the rudder control alone (without the § 25.149 Minimum control speed.

use of nosewheel steering), as limited (a) In establishing the minimum con- by 150 pounds of force, and the lateral trol speeds required by this section, the control to the extent of keeping the method used to simulate critical en- wings level to enable the takeoff to be gine failure must represent the most safely continued using normal piloting critical mode of powerplant failure skill. In the determination of V , as- MCG with respect to controllability ex- suming that the path of the airplane pected in service.

accelerating with all engines operating (b) V is the calibrated airspeed at is along the centerline of the runway, MC which, when the critical engine is sud- its path from the point at which the denly made inoperative, it is possible critical engine is made inoperative to to maintain control of the airplane the point at which recovery to a direc- with that engine still inoperative and tion parallel to the centerline is com- maintain straight flight with an angle pleted may not deviate more than 30 of bank of not more than 5 degrees. feet laterally from the centerline at 14 CFR Ch. I (1–1–25 Edition) § 25.149 any point. V must be established (1) The airplane in the most critical MCG with— configuration (or, at the option of the (1) The airplane in each takeoff con- applicant, each configuration) for ap- figuration or, at the option of the ap- proach and landing with one critical plicant, in the most critical takeoff engine inoperative; configuration; (2) The most unfavorable center of (2) Maximum available takeoff power gravity; or thrust on the operating engines; (3) The airplane trimmed for ap- (3) The most unfavorable center of proach with one critical engine inoper- gravity; ative; (4) The airplane trimmed for takeoff; (4) The most unfavorable weight, or, and at the option of the applicant, as a (5) The most unfavorable weight in function of weight; the range of takeoff weights.

(5) For propeller airplanes, the pro- (f) V , the minimum control speed MCL peller of the more critical inoperative during approach and landing with all engine in the position it achieves with- engines operating, is the calibrated air- out pilot action, assuming the engine speed at which, when the critical en- fails while at the power or thrust nec- gine is suddenly made inoperative, it is essary to maintain a three degree ap- possible to maintain control of the air- proach path angle, and the propeller of plane with that engine still inoper- the other inoperative engine feathered; ative, and maintain straight flight (6) The power or thrust on the oper- with an angle of bank of not more than ating engine(s) necessary to maintain 5 degrees. V must be established MCL an approach path angle of three de- with— grees when one critical engine is inop- (1) The airplane in the most critical erative; and configuration (or, at the option of the (7) The power or thrust on the oper- applicant, each configuration) for ap- ating engine(s) rapidly changed, imme- proach and landing with all engines op- erating; diately after the second critical engine is made inoperative, from the power or (2) The most unfavorable center of thrust prescribed in paragraph (g)(6) of gravity; (3) The airplane trimmed for ap- this section to— proach with all engines operating; (i) Minimum power or thrust; and (4) The most favorable weight, or, at (ii) Go-around power or thrust set- the option of the applicant, as a func- ting.

tion of weight; (h) In demonstrations of V and MCL (5) For propeller airplanes, the pro- V — MCL-2 peller of the inoperative engine in the (1) The rudder force may not exceed position it achieves without pilot ac- 150 pounds; tion, assuming the engine fails while at (2) The airplane may not exhibit haz- the power or thrust necessary to main- ardous flight characteristics or require tain a three degree approach path exceptional piloting skill, alertness, or angle; and strength; (6) Go-around power or thrust setting (3) Lateral control must be sufficient on the operating engine(s).

to roll the airplane, from an initial (g) For airplanes with three or more condition of steady flight, through an engines, V , the minimum control MCL-2 angle of 20 degrees in the direction nec- speed during approach and landing essary to initiate a turn away from the with one critical engine inoperative, is inoperative engine(s), in not more than the calibrated airspeed at which, when 5 seconds; and a second critical engine is suddenly (4) For propeller airplanes, hazardous made inoperative, it is possible to flight characteristics must not be ex- maintain control of the airplane with hibited due to any propeller position both engines still inoperative, and maintain straight flight with an angle of bank of not more than 5 degrees.

V must be established with— MCL-2 Federal Aviation Administration, DOT § 25.173 achieved when the engine fails or dur- grees) at 1.3 V during climbing flight SR 1 ing any likely subsequent movements with— of the engine or propeller controls. (1) The critical engine inoperative; (2) The remaining engines at max- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as imum continuous power; and amended by Amdt. 25–42, 43 FR 2321, Jan. 16, (3) The landing gear and flaps re- 1978; Amdt. 25–72, 55 FR 29774, July 20, 1990; 55 tracted.

FR 37607, Sept. 12, 1990; Amdt. 25–84, 60 FR (e) Airplanes with four or more en- 30749, June 9, 1995; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002] gines. Each airplane with four or more engines must also maintain trim in T RIM rectilinear flight with the most unfa- vorable center of gravity and at the § 25.161 Trim.

climb speed, configuration, and power (a) General. Each airplane must meet required by § 25.123(a) for the purpose of the trim requirements of this section establishing the en route flight paths after being trimmed, and without fur- with two engines inoperative.

ther pressure upon, or movement of, ei- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ther the primary controls or their cor- amended by Amdt. 25–23, 35 FR 5671, Apr. 8, responding trim controls by the pilot 1970; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; or the automatic pilot.

Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; (b) Lateral and directional trim. The Amdt. 25–115, 69 FR 40527, July 2, 2004] airplane must maintain lateral and di- S TABILITY rectional trim with the most adverse lateral displacement of the center of § 25.171 General.

gravity within the relevant operating The airplane must be longitudinally, limitations, during normally expected directionally, and laterally stable in conditions of operation (including op- accordance with the provisions of eration at any speed from 1.3 V to SR 1 §§ 25.173 through 25.177. In addition, V /M ).

MO MO suitable stability and control feel (c) Longitudinal trim. The airplane (static stability) is required in any con- must maintain longitudinal trim dur- dition normally encountered in service, ing— if flight tests show it is necessary for (1) A climb with maximum contin- safe operation.

uous power at a speed not more than 1.3 V , with the landing gear re- SR 1 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tracted, and the flaps (i) retracted and amended by Amdt. 25–7, 30 FR 13117, Oct. 15, (ii) in the takeoff position; 1965] (2) Either a glide with power off at a § 25.173 Static longitudinal stability.

speed not more than 1.3 V , or an ap- SR1 proach within the normal range of ap- Under the conditions specified in proach speeds appropriate to the § 25.175, the characteristics of the eleva- weight and configuration with power tor control forces (including friction) settings corresponding to a 3 degree must be as follows: glidepath, whichever is the most se- (a) A pull must be required to obtain vere, with the landing gear extended, and maintain speeds below the speci- the wing flaps (i) retracted and (ii) ex- fied trim speed, and a push must be re- tended, and with the most unfavorable quired to obtain and maintain speeds combination of center of gravity posi- above the specified trim speed. This tion and weight approved for landing; must be shown at any speed that can be and obtained except speeds higher than the (3) Level flight at any speed from 1.3 landing gear or wing flap operating V , to V /M with the landing gear limit speeds or V /M whichever is SR 1 MO MO, FC FC, and flaps retracted, and from 1.3 V to appropriate, or lower than the min- SR 1 V with the landing gear extended. imum speed for steady unstalled flight.

LE (d) Longitudinal, directional, and lat- (b) The airspeed must return to with- eral trim. The airplane must maintain in 10 percent of the original trim speed longitudinal, directional, and lateral for the climb, approach, and landing trim (and for the lateral trim, the conditions specified in § 25.175 (a), (c), angle of bank may not exceed five de- and (d), and must return to within 7.5 14 CFR Ch. I (1–1–25 Edition) § 25.175 percent of the original trim speed for (ii) The center of gravity in the most the cruising condition specified in adverse position (see § 25.27); § 25.175(b), when the control force is (iii) The most critical weight be- slowly released from any speed within tween the maximum takeoff and max- the range specified in paragraph (a) of imum landing weights; this section.

(iv) 75 percent of maximum contin- (c) The average gradient of the stable uous power for reciprocating engines or slope of the stick force versus speed for turbine engines, the maximum curve may not be less than 1 pound for cruising power selected by the appli- each 6 knots.

cant as an operating limitation (see (d) Within the free return speed range § 25.1521), except that the power need specified in paragraph (b) of this sec- not exceed that required at V / M ; MO MO tion, it is permissible for the airplane, and without control forces, to stabilize on (v) The airplane trimmed for level speeds above or below the desired trim flight with the power required in para- speeds if exceptional attention on the graph (b)(1)(iv) of this section.

part of the pilot is not required to re- (2) With the landing gear retracted at turn to and maintain the desired trim low speed, the stick force curve must speed and altitude.

have a stable slope at all speeds within [Amdt. 25–7, 30 FR 13117, Oct. 15, 1965] a range which is the greater of 15 per- cent of the trim speed plus the result- § 25.175 Demonstration of static longi- ing free return speed range, or 50 knots tudinal stability.

plus the resulting free return speed Static longitudinal stability must be range, above and below the trim speed shown as follows: (except that the speed range need not (a) Climb. The stick force curve must include speeds less than 1.3 V , nor SR 1 have a stable slope at speeds between speeds greater than the minimum 85 and 115 percent of the speed at which speed of the applicable speed range pre- the airplane— scribed in paragraph (b)(1), nor speeds (1) Is trimmed, with— that require a stick force of more than (i) Wing flaps retracted; 50 pounds), with— (ii) Landing gear retracted; (i) Wing flaps, center of gravity posi- (iii) Maximum takeoff weight; and tion, and weight as specified in para- (iv) 75 percent of maximum contin- graph (b)(1) of this section; uous power for reciprocating engines or (ii) Power required for level flight at the maximum power or thrust selected a speed equal to ( V + 1.3 V )/2; and MO SR 1 by the applicant as an operating limi- (iii) The airplane trimmed for level tation for use during climb for turbine flight with the power required in para- engines; and graph (b)(2)(ii) of this section.

(2) Is trimmed at the speed for best (3) With the landing gear extended, rate-of-climb except that the speed the stick force curve must have a sta- need not be less than 1.3 V .

SR 1 ble slope at all speeds within a range (b) Cruise. Static longitudinal sta- which is the greater of 15 percent of the bility must be shown in the cruise con- trim speed plus the resulting free re- dition as follows: turn speed range, or 50 knots plus the (1) With the landing gear retracted at resulting free return speed range, high speed, the stick force curve must above and below the trim speed (except have a stable slope at all speeds within that the speed range need not include a range which is the greater of 15 per- speeds less than 1.3 V , nor speeds SR 1 cent of the trim speed plus the result- greater than V nor speeds that re- LE, ing free return speed range, or 50 knots quire a stick force of more than 50 plus the resulting free return speed pounds), with— range, above and below the trim speed (i) Wing flap, center of gravity posi- (except that the speed range need not tion, and weight as specified in para- include speeds less than 1.3 V , nor SR 1 graph (b)(1) of this section; speeds greater than V /M nor speeds FC FC, that require a stick force of more than (ii) 75 percent of maximum contin- 50 pounds), with— uous power for reciprocating engines (i) The wing flaps retracted; or, for turbine engines, the maximum Federal Aviation Administration, DOT § 25.181 cruising power selected by the appli- (2) From V /M to V /M , unless MO MO FC FC cant as an operating limitation, except the divergence is— that the power need not exceed that re- (i) Gradual; (ii) Easily recognizable by the pilot; quired for level flight at V ; and LE and (iii) The aircraft trimmed for level (iii) Easily controllable by the pilot.

flight with the power required in para- (c) The following requirement must graph (b)(3)(ii) of this section.

be met for the configurations and speed (c) Approach. The stick force curve specified in paragraph (a) of this sec- must have a stable slope at speeds be- tion. In straight, steady sideslips over tween V and 1.7 V , with— SW SR 1 the range of sideslip angles appropriate (1) Wing flaps in the approach posi- to the operation of the airplane, the ai- tion; leron and rudder control movements (2) Landing gear retracted; and forces must be substantially pro- (3) Maximum landing weight; and portional to the angle of sideslip in a (4) The airplane trimmed at 1.3 V SR 1 stable sense. This factor of proportion- with enough power to maintain level ality must lie between limits found flight at this speed.

necessary for safe operation. The range (d) Landing. The stick force curve of sideslip angles evaluated must in- must have a stable slope, and the stick clude those sideslip angles resulting force may not exceed 80 pounds, at from the lesser of: speeds between V and 1.7 V with— SW SR 0 (1) One-half of the available rudder (1) Wing flaps in the landing position; control input; and (2) Landing gear extended; (2) A rudder control force of 180 (3) Maximum landing weight; pounds.

(4) The airplane trimmed at 1.3 V SR0 (d) For sideslip angles greater than with— those prescribed by paragraph (c) of (i) Power or thrust off, and this section, up to the angle at which full rudder control is used or a rudder (ii) Power or thrust for level flight.

control force of 180 pounds is obtained, (5) The airplane trimmed at 1.3 V SR 0 the rudder control forces may not re- with power or thrust off.

verse, and increased rudder deflection [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as must be needed for increased angles of amended by Amdt. 25–7, 30 FR 13117, Oct. 15, sideslip. Compliance with this require- 1965; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; ment must be shown using straight, Amdt. 25–115, 69 FR 40527, July 2, 2004] steady sideslips, unless full lateral con- trol input is achieved before reaching § 25.177 Static lateral-directional sta- bility. either full rudder control input or a rudder control force of 180 pounds; a (a) The static directional stability straight, steady sideslip need not be (as shown by the tendency to recover maintained after achieving full lateral from a skid with the rudder free) must control input. This requirement must be positive for any landing gear and be met at all approved landing gear and flap position and symmetric power con- flap positions for the range of oper- dition, at speeds from 1.13 V , up to SR1 ating speeds and power conditions ap- V , V , or V /M (as appropriate for FE LE FC FC propriate to each landing gear and flap the airplane configuration).

position with all engines operating.

(b) The static lateral stability (as shown by the tendency to raise the low [Amdt. 25–135, 76 FR 74654, Dec. 1, 2011] wing in a sideslip with the aileron con- § 25.181 Dynamic stability.

trols free) for any landing gear and flap position and symmetric power condi- (a) Any short period oscillation, not tion, may not be negative at any air- including combined lateral-directional speed (except that speeds higher than oscillations, occurring between 1.13 V SR V need not be considered for flaps ex- and maximum allowable speed appro- FE tended configurations nor speeds high- priate to the configuration of the air- er than V for landing gear extended plane must be heavily damped with the LE configurations) in the following air- primary controls— speed ranges: (1) Free; and (1) From 1.13 V to V /M . (2) In a fixed position.

SR1 MO MO 14 CFR Ch. I (1–1–25 Edition) § 25.201 (b) Any combined lateral-directional the pilot a clear and distinctive indica- oscillations (‘‘Dutch roll’’) occurring tion of an acceptable nature that the between 1.13 V and maximum allow- SR airplane is stalled. Acceptable indica- able speed appropriate to the configu- tions of a stall, occurring either indi- ration of the airplane must be posi- vidually or in combination, are— tively damped with controls free, and (1) A nose-down pitch that cannot be must be controllable with normal use readily arrested; of the primary controls without requir- (2) Buffeting, of a magnitude and se- ing exceptional pilot skill.

verity that is a strong and effective de- [Amdt. 25–42, 43 FR 2322, Jan. 16, 1978, as terrent to further speed reduction; or amended by Amdt. 25–72, 55 FR 29775, July 20, (3) The pitch control reaches the aft 1990; 55 FR 37607, Sept. 12, 1990; Amdt. 25–108, stop and no further increase in pitch 67 FR 70827, Nov. 26, 2002] attitude occurs when the control is held full aft for a short time before re- S TALLS covery is initiated.

§ 25.201 Stall demonstration.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) Stalls must be shown in straight amended by Amdt. 25–84, 60 FR 30750, June 9, flight and in 30 degree banked turns 1995; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002] with— § 25.203 Stall characteristics.

(1) Power off; and (2) The power necessary to maintain (a) It must be possible to produce and level flight at 1.5 V (where V cor- SR1 SR1 to correct roll and yaw by unreversed responds to the reference stall speed at use of the aileron and rudder controls, maximum landing weight with flaps in up to the time the airplane is stalled.

the approach position and the landing No abnormal nose-up pitching may gear retracted).

occur. The longitudinal control force (b) In each condition required by must be positive up to and throughout paragraph (a) of this section, it must the stall. In addition, it must be pos- be possible to meet the applicable re- sible to promptly prevent stalling and quirements of § 25.203 with— to recover from a stall by normal use (1) Flaps, landing gear, and decelera- of the controls.

tion devices in any likely combination (b) For level wing stalls, the roll oc- of positions approved for operation; curring between the stall and the com- (2) Representative weights within the pletion of the recovery may not exceed range for which certification is re- approximately 20 degrees.

quested; (3) The most adverse center of grav- (c) For turning flight stalls, the ac- ity for recovery; and tion of the airplane after the stall may (4) The airplane trimmed for straight not be so violent or extreme as to flight at the speed prescribed in make it difficult, with normal piloting § 25.103(b)(6).

skill, to effect a prompt recovery and (c) The following procedures must be to regain control of the airplane. The used to show compliance with § 25.203; maximum bank angle that occurs dur- (1) Starting at a speed sufficiently ing the recovery may not exceed— above the stalling speed to ensure that (1) Approximately 60 degrees in the a steady rate of speed reduction can be original direction of the turn, or 30 de- established, apply the longitudinal grees in the opposite direction, for de- control so that the speed reduction celeration rates up to 1 knot per sec- does not exceed one knot per second ond; and until the airplane is stalled.

(2) Approximately 90 degrees in the (2) In addition, for turning flight original direction of the turn, or 60 de- stalls, apply the longitudinal control grees in the opposite direction, for de- to achieve airspeed deceleration rates celeration rates in excess of 1 knot per up to 3 knots per second.

second.

(3) As soon as the airplane is stalled, recover by normal recovery techniques.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (d) The airplane is considered stalled amended by Amdt. 25–84, 60 FR 30750, June 9, when the behavior of the airplane gives 1995] Federal Aviation Administration, DOT § 25.207 same way as for the airplane in non- § 25.207 Stall warning.

icing conditions. Compliance with this (a) Stall warning with sufficient mar- requirement must be demonstrated in gin to prevent inadvertent stalling flight with the speed reduced at rates with the flaps and landing gear in any not exceeding one knot per second, normal position must be clear and dis- with— tinctive to the pilot in straight and (1) The most critical of the takeoff turning flight.

ice and final takeoff ice accretions de- (b) The warning must be furnished ei- fined in Appendices C and O of this ther through the inherent aerodynamic part, as applicable, in accordance with qualities of the airplane or by a device § 25.21(g), for each configuration used in that will give clearly distinguishable the takeoff phase of flight; indications under expected conditions of flight. However, a visual stall warn- (2) The most critical of the en route ing device that requires the attention ice accretion(s) defined in Appendices C of the crew within the cockpit is not and O of this part, as applicable, in ac- acceptable by itself. If a warning de- cordance with § 25.21(g), for the en vice is used, it must provide a warning route configuration; in each of the airplane configurations (3) The most critical of the holding prescribed in paragraph (a) of this sec- ice accretion(s) defined in Appendices C tion at the speed prescribed in para- and O of this part, as applicable, in ac- graphs (c) and (d) of this section. Ex- cordance with § 25.21(g), for the holding cept for the stall warning prescribed in configuration(s); paragraph (h)(3)(ii) of this section, the (4) The most critical of the approach stall warning for flight in icing condi- ice accretion(s) defined in Appendices C tions must be provided by the same and O of this part, as applicable, in ac- means as the stall warning for flight in cordance with § 25.21(g), for the ap- non-icing conditions.

proach configuration(s); and (c) When the speed is reduced at rates (5) The most critical of the landing not exceeding one knot per second, ice accretion(s) defined in Appendices C stall warning must begin, in each nor- and O of this part, as applicable, in ac- mal configuration, at a speed, V , ex- SW cordance with § 25.21(g), for the landing ceeding the speed at which the stall is and go-around configuration(s).

identified in accordance with § 25.201(d) (f) The stall warning margin must be by not less than five knots or five per- sufficient in both non-icing and icing cent CAS, whichever is greater. Once conditions to allow the pilot to prevent initiated, stall warning must continue stalling when the pilot starts a recov- until the angle of attack is reduced to ery maneuver not less than one second approximately that at which stall after the onset of stall warning in slow- warning began.

down turns with at least 1.5 g load fac- (d) In addition to the requirement of tor normal to the flight path and air- paragraph (c) of this section, when the speed deceleration rates of at least 2 speed is reduced at rates not exceeding knots per second. When demonstrating one knot per second, in straight flight compliance with this paragraph for with engines idling and at the center- icing conditions, the pilot must per- of-gravity position specified in form the recovery maneuver in the § 25.103(b)(5), V , in each normal con- SW same way as for the airplane in non- figuration, must exceed V by not less SR icing conditions. Compliance with this than three knots or three percent CAS, requirement must be demonstrated in whichever is greater.

flight with— (e) In icing conditions, the stall (1) The flaps and landing gear in any warning margin in straight and turn- normal position; ing flight must be sufficient to allow (2) The airplane trimmed for straight the pilot to prevent stalling (as defined flight at a speed of 1.3 V ; and SR in § 25.201(d)) when the pilot starts a re- (3) The power or thrust necessary to covery maneuver not less than three maintain level flight at 1.3 V .

seconds after the onset of stall warn- SR ing. When demonstrating compliance (g) Stall warning must also be pro- with this paragraph, the pilot must vided in each abnormal configuration perform the recovery maneuver in the of the high lift devices that is likely to 14 CFR Ch. I (1–1–25 Edition) § 25.231 be used in flight following system fail- G ROUND AND W ATER H ANDLING ures (including all configurations cov- C HARACTERISTICS ered by Airplane Flight Manual proce- § 25.231 Longitudinal stability and dures).

control.

(h) The following stall warning mar- gin is required for flight in icing condi- (a) Landplanes may have no uncon- trollable tendency to nose over in any tions before the ice protection system reasonably expected operating condi- has been activated and is performing tion or when rebound occurs during its intended function. Compliance must landing or takeoff. In addition— be shown using the most critical of the ice accretion(s) defined in Appendix C, (1) Wheel brakes must operate smoothly and may not cause any undue part II, paragraph (e) of this part and tendency to nose over; and Appendix O, part II, paragraph (d) of (2) If a tail-wheel landing gear is this part, as applicable, in accordance used, it must be possible, during the with § 25.21(g). The stall warning mar- takeoff ground run on concrete, to gin in straight and turning flight must maintain any attitude up to thrust line be sufficient to allow the pilot to pre- level, at 75 percent of V .

vent stalling without encountering any SR 1 (b) For seaplanes and amphibians, adverse flight characteristics when: the most adverse water conditions safe (1) The speed is reduced at rates not for takeoff, taxiing, and landing, must exceeding one knot per second; be established.

(2) The pilot performs the recovery maneuver in the same way as for flight [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as in non-icing conditions; and amended by Amdt. 25–108, 67 FR 70828, Nov.

26, 2002] (3) The recovery maneuver is started no earlier than: § 25.233 Directional stability and con- (i) One second after the onset of stall trol.

warning if stall warning is provided by (a) There may be no uncontrollable the same means as for flight in non- ground-looping tendency in 90 ° cross icing conditions; or winds, up to a wind velocity of 20 knots (ii) Three seconds after the onset of or 0.2 V , whichever is greater, except SR 0 stall warning if stall warning is pro- that the wind velocity need not exceed vided by a different means than for 25 knots at any speed at which the air- flight in non-icing conditions.

plane may be expected to be operated (i) In showing compliance with para- on the ground. This may be shown graph (h) of this section, if stall warn- while establishing the 90 ° cross compo- ing is provided by a different means in nent of wind velocity required by icing conditions than for non-icing con- § 25.237.

ditions, compliance with § 25.203 must (b) Landplanes must be satisfactorily be shown using the accretion defined in controllable, without exceptional pilot- appendix C, part II(e) of this part. Com- ing skill or alertness, in power-off land- pliance with this requirement must be ings at normal landing speed, without shown using the demonstration pre- using brakes or engine power to main- scribed by § 25.201, except that the de- tain a straight path. This may be celeration rates of § 25.201(c)(2) need not shown during power-off landings made be demonstrated.

in conjunction with other tests.

(c) The airplane must have adequate [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–7, 30 FR 13118, Oct. 15, directional control during taxiing. This 1965; Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; may be shown during taxiing prior to Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; takeoffs made in conjunction with Amdt. 25–121, 72 FR 44668, Aug. 8, 2007; Amdt.

other tests.

25–129, 74 FR 38339, Aug. 3, 2009; Amdt. 25–140, 79 FR 65526, Nov. 4, 2014] [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5671, Apr. 8, 1970; Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998; Amdt.

25–108, 67 FR 70828, Nov. 26, 2002] Federal Aviation Administration, DOT § 25.251 or result in the taking in of an undue § 25.235 Taxiing condition.

quantity of water; The shock absorbing mechanism may (2) Dangerously uncontrollable not damage the structure of the air- porpoising, bounding, or swinging tend- plane when the airplane is taxied on ency; or the roughest ground that may reason- (3) Immersion of auxiliary floats or ably be expected in normal operation.

sponsons, wing tips, propeller blades, or other parts not designed to with- § 25.237 Wind velocities.

stand the resulting water loads.

(a) For land planes and amphibians, (b) Compliance with the require- the following applies: ments of paragraph (a) of this section (1) A 90-degree cross component of must be shown— wind velocity, demonstrated to be safe (1) In water conditions, from smooth for takeoff and landing, must be estab- to the most adverse condition estab- lished for dry runways and must be at lished in accordance with § 25.231; least 20 knots or 0.2 V , whichever is SR0 (2) In wind and cross-wind velocities, greater, except that it need not exceed water currents, and associated waves 25 knots.

and swells that may reasonably be ex- (2) The crosswind component for pected in operation on water; takeoff established without ice accre- (3) At speeds that may reasonably be tions is valid in icing conditions.

expected in operation on water; (3) The landing crosswind component (4) With sudden failure of the critical must be established for: engine at any time while on water; and (i) Non-icing conditions, and (5) At each weight and center of grav- (ii) Icing conditions with the most ity position, relevant to each operating critical of the landing ice accretion(s) condition, within the range of loading defined in Appendices C and O of this conditions for which certification is re- part, as applicable, in accordance with quested.

§ 25.21(g).

(b) For seaplanes and amphibians, (c) In the water conditions of para- the following applies: graph (b) of this section, and in the (1) A 90-degree cross component of corresponding wind conditions, the sea- wind velocity, up to which takeoff and plane or amphibian must be able to landing is safe under all water condi- drift for five minutes with engines in- tions that may reasonably be expected operative, aided, if necessary, by a sea in normal operation, must be estab- anchor.

lished and must be at least 20 knots or M ISCELLANEOUS F LIGHT R EQUIREMENTS 0.2 V , whichever is greater, except SR 0 that it need not exceed 25 knots.

§ 25.251 Vibration and buffeting.

(2) A wind velocity, for which taxiing is safe in any direction under all water (a) The airplane must be dem- conditions that may reasonably be ex- onstrated in flight to be free from any pected in normal operation, must be es- vibration and buffeting that would pre- tablished and must be at least 20 knots vent continued safe flight in any likely or 0.2 V , whichever is greater, except operating condition.

SR0 that it need not exceed 25 knots. (b) Each part of the airplane must be demonstrated in flight to be free from [Amdt. 25–42, 43 FR 2322, Jan. 16, 1978, as excessive vibration under any appro- amended by Amdt. 25–108, 67 FR 70827, Nov.

priate speed and power conditions up to 26, 2002; Amdt. 25–121, 72 FR 44668, Aug. 8, V /M . The maximum speeds shown 2007; Amdt. 25–140, 79 FR 65525, Nov. 4, 2014] DF DF must be used in establishing the oper- § 25.239 Spray characteristics, control, ating limitations of the airplane in ac- and stability on water.

cordance with § 25.1505.

(a) For seaplanes and amphibians, (c) Except as provided in paragraph during takeoff, taxiing, and landing, (d) of this section, there may be no buf- and in the conditions set forth in para- feting condition, in normal flight, in- graph (b) of this section, there may be cluding configuration changes during no— cruise, severe enough to interfere with (1) Spray characteristics that would the control of the airplane, to cause ex- impair the pilot’s view, cause damage, cessive fatigue to the crew, or to cause 14 CFR Ch. I (1–1–25 Edition) § 25.253 structural damage. Stall warning buf- (iii) Buffeting that would impair the feting within these limits is allowable. pilot’s ability to read the instruments or control the airplane for recovery.

(d) There may be no perceptible buf- (3) With the airplane trimmed at any feting condition in the cruise configu- speed up to V /M , there must be no ration in straight flight at any speed MO MO reversal of the response to control up to V / M except that stall warn- MO MO, input about any axis at any speed up to ing buffeting is allowable.

V /M . Any tendency to pitch, roll, or DF DF (e) For an airplane with M greater D yaw must be mild and readily control- than .6 or with a maximum operating lable, using normal piloting tech- altitude greater than 25,000 feet, the niques. When the airplane is trimmed positive maneuvering load factors at at V /M , the slope of the elevator MO MO which the onset of perceptible buf- control force versus speed curve need feting occurs must be determined with not be stable at speeds greater than the airplane in the cruise configuration V /M , but there must be a push force FC FC for the ranges of airspeed or Mach at all speeds up to V /M and there DF DF number, weight, and altitude for which must be no sudden or excessive reduc- the airplane is to be certificated. The tion of elevator control force as V / DF envelopes of load factor, speed, alti- M is reached.

DF tude, and weight must provide a suffi- (4) Adequate roll capability to assure cient range of speeds and load factors a prompt recovery from a lateral upset for normal operations. Probable inad- condition must be available at any vertent excursions beyond the bound- speed up to V /M .

DF DF aries of the buffet onset envelopes may (5) With the airplane trimmed at not result in unsafe conditions.

V /M , extension of the speedbrakes MO MO over the available range of movements [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as of the pilot’s control, at all speeds amended by Amdt. 25–23, 35 FR 5671, Apr. 8, 1970; Amdt. 25–72, 55 FR 29775, July 20, 1990; above V /M , but not so high that MO MO Amdt. 25–77, 57 FR 28949, June 29, 1992] V /M would be exceeded during the DF DF maneuver, must not result in: § 25.253 High-speed characteristics.

(i) An excessive positive load factor when the pilot does not take action to (a) Speed increase and recovery charac- counteract the effects of extension; teristics. The following speed increase (ii) Buffeting that would impair the and recovery characteristics must be pilot’s ability to read the instruments met: or control the airplane for recovery; or (1) Operating conditions and charac- (iii) A nose down pitching moment, teristics likely to cause inadvertent unless it is small.

speed increases (including upsets in (b) Maximum speed for stability charac- pitch and roll) must be simulated with teristics, V /M . V /M is the max- FC FC FC FC the airplane trimmed at any likely imum speed at which the requirements cruise speed up to V / M . These con- MO MO of §§ 25.143(g), 25.147(f), 25.175(b)(1), ditions and characteristics include gust 25.177(a) through (c), and 25.181 must be upsets, inadvertent control move- met with flaps and landing gear re- ments, low stick force gradient in rela- tracted. Except as noted in § 25.253(c), tion to control friction, passenger V /M may not be less than a speed FC FC movement, leveling off from climb, and midway between V /M and V /M , MO MO DF DF descent from Mach to airspeed limit al- except that, for altitudes where Mach titudes.

number is the limiting factor, M need FC (2) Allowing for pilot reaction time not exceed the Mach number at which after effective inherent or artificial effective speed warning occurs.

speed warning occurs, it must be shown (c) Maximum speed for stability charac- that the airplane can be recovered to a teristics in icing conditions. The max- normal attitude and its speed reduced imum speed for stability characteris- to V / M without— MO MO, tics with the most critical of the ice (i) Exceptional piloting strength or accretions defined in Appendices C and skill; O of this part, as applicable, in accord- (ii) Exceeding V / M V / M or the ance with § 25.21(g), at which the re- D D, DF DF, structural limitations; and quirements of §§ 25.143(g), 25.147(f), Federal Aviation Administration, DOT § 25.255 25.175(b)(1), 25.177(a) through (c), and (2) 0 g to 2.0 g, and extrapolating by 25.181 must be met, is the lower of: an acceptable method to ¥ 1 g and + 2.5 (1) 300 knots CAS; g.

(2) V ; or FC (d) If the procedure set forth in para- (3) A speed at which it is dem- graph (c)(2) of this section is used to onstrated that the airframe will be free demonstrate compliance and marginal of ice accretion due to the effects of in- conditions exist during flight test with creased dynamic pressure.

regard to reversal of primary longitu- dinal control force, flight tests must be [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as accomplished from the normal accel- amended by Amdt. 25–23, 35 FR 5671, Apr. 8, 1970; Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; eration at which a marginal condition Amdt. 25–72, 55 FR 29775, July 20, 1990; Amdt.

is found to exist to the applicable limit 25–84, 60 FR 30750, June 9, 1995; Amdt. 25–121, specified in paragraph (b)(1) of this sec- 72 FR 44668, Aug. 8, 2007; Amdt. 25–135, 76 FR tion.

74654, Dec. 1, 2011; Amdt. 25–140,79 FR 65525, (e) During flight tests required by Nov. 4, 2014] paragraph (a) of this section, the limit maneuvering load factors prescribed in § 25.255 Out-of-trim characteristics.

§§ 25.333(b) and 25.337, and the maneu- (a) From an initial condition with vering load factors associated with the airplane trimmed at cruise speeds probable inadvertent excursions be- up to V /M the airplane must have MO MO, yond the boundaries of the buffet onset satisfactory maneuvering stability and envelopes determined under § 25.251(e), controllability with the degree of out- need not be exceeded. In addition, the of-trim in both the airplane nose-up entry speeds for flight test demonstra- and nose-down directions, which re- tions at normal acceleration values sults from the greater of— less than 1 g must be limited to the ex- (1) A three-second movement of the tent necessary to accomplish a recov- longitudinal trim system at its normal ery without exceeding V /M .

rate for the particular flight condition DF DF (f) In the out-of-trim condition speci- with no aerodynamic load (or an equiv- alent degree of trim for airplanes that fied in paragraph (a) of this section, it do not have a power-operated trim sys- must be possible from an overspeed tem), except as limited by stops in the condition at V /M to produce at DF DF trim system, including those required least 1.5 g for recovery by applying not by § 25.655(b) for adjustable stabilizers; more than 125 pounds of longitudinal or control force using either the primary (2) The maximum mistrim that can longitudinal control alone or the pri- be sustained by the autopilot while mary longitudinal control and the lon- maintaining level flight in the high gitudinal trim system. If the longitu- speed cruising condition.

dinal trim is used to assist in pro- (b) In the out-of-trim condition speci- ducing the required load factor, it must fied in paragraph (a) of this section, be shown at V /M that the longitu- DF DF when the normal acceleration is varied dinal trim can be actuated in the air- from + 1 g to the positive and negative plane nose-up direction with the pri- values specified in paragraph (c) of this mary surface loaded to correspond to section— the least of the following airplane (1) The stick force vs. g curve must nose-up control forces: have a positive slope at any speed up to (1) The maximum control forces ex- and including V /M ; and FC FC pected in service as specified in §§ 25.301 (2) At speeds between V /M and FC FC and 25.397.

V /M the direction of the primary DF DF (2) The control force required to longitudinal control force may not re- produce 1.5 g.

verse.

(3) The control force corresponding to (c) Except as provided in paragraphs buffeting or other phenomena of such (d) and (e) of this section, compliance intensity that it is a strong deterrent with the provisions of paragraph (a) of to further application of primary longi- this section must be demonstrated in tudinal control force.

flight over the acceleration range— (1) ¥ 1 g to + 2.5 g; or [Amdt. 25–42, 43 FR 2322, Jan. 16, 1978]

14 CFR Ch. I (1–1–25 Edition) § 25.301 ure conditions on the airplane struc- Subpart C—Structure ture.

(b) System fully operative. With the G ENERAL system fully operative, the following § 25.301 Loads.

criteria apply: (1) The applicant must derive limit (a) Strength requirements are speci- loads for the limit conditions specified fied in terms of limit loads (the max- in subpart C of this part, taking into imum loads to be expected in service) account the behavior of the system up and ultimate loads (limit loads multi- to the limit loads. System nonlinear- plied by prescribed factors of safety).

ities must be taken into account.

Unless otherwise provided, prescribed (2) The applicant must show that the loads are limit loads.

airplane meets the strength require- (b) Unless otherwise provided, the ments of subparts C and D of this part, specified air, ground, and water loads using the appropriate factor of safety must be placed in equilibrium with in- to derive ultimate loads from the limit ertia forces, considering each item of loads defined in paragraph (b)(1) of this mass in the airplane. These loads must section. The effect of nonlinearities be distributed to conservatively ap- must be investigated sufficiently be- proximate or closely represent actual yond limit conditions to ensure the be- conditions. Methods used to determine havior of the system presents no detri- load intensities and distribution must mental effects compared to the behav- be validated by flight load measure- ior below limit conditions. However, ment unless the methods used for de- conditions beyond limit conditions termining those loading conditions are need not be considered when it can be shown to be reliable.

shown that the airplane has design fea- (c) If deflections under load would tures that will not allow it to exceed significantly change the distribution of those limit conditions.

external or internal loads, this redis- (3) [Reserved] tribution must be taken into account.

(c) System in the failure condition. For [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as any system failure condition not shown amended by Amdt. 25–23, 35 FR 5672, Apr. 8, to be extremely improbable or that re- 1970] sults from a single failure, the fol- lowing criteria apply: § 25.302 Interaction of systems and (1) At the time of occurrence. The ap- structures.

plicant must establish a realistic sce- For airplanes equipped with systems nario, starting from 1g level flight con- that affect structural performance, ei- ditions, and including pilot corrective ther directly or as a result of a failure actions, to determine the loads occur- or malfunction, the influence of these ring at the time of failure and imme- systems and their failure conditions diately after failure.

must be taken into account when (i) For static strength substan- showing compliance with the require- tiation, the airplane must be able to ments of subparts C and D of this part.

withstand the ultimate loads deter- These criteria are only applicable to mined by multiplying the loads in structure whose failure could prevent paragraph (c)(1) of this section by a continued safe flight and landing.

factor of safety that is related to the (a) General. The applicant must use probability of occurrence of the failure.

the following criteria in determining The factor of safety (F.S.) is defined in the influence of a system and its fail- Figure 1.

Federal Aviation Administration, DOT § 25.302 Figure 1 to paragraph (c)(1)(i) (ii) For residual strength substan- (B) the limit gust and turbulence tiation, the airplane must be able to conditions specified in §§ 25.341 and withstand two thirds of the ultimate 25.345, loads defined in paragraph (c)(1)(i) of (C) the limit rolling conditions speci- this section. For pressurized cabins, fied in § 25.349 and the limit unsymmet- these loads must be combined with the rical conditions specified in §§ 25.367 normal operating differential pressure.

and 25.427(b) and (c), (iii) [Reserved] (D) the limit yaw maneuvering condi- (iv) Failures of the system that re- tions specified in § 25.351, sult in forced structural vibrations (os- (E) the limit ground loading condi- cillatory failures) must not produce tions specified in §§ 25.473 and 25.491, loads that could result in detrimental and deformation of primary structure.

(F) any other subpart C of this part (2) For the continuation of the flight.

For the airplane, in the system failed load condition for which a system is state and considering any appropriate specifically installed or tailored to re- reconfiguration and flight limitations, duce the loads of that condition.

the following apply: (ii) For static strength substan- (i) The loads derived from the fol- tiation, each part of the structure lowing conditions at speeds up to V / C must be able to withstand the loads in M , or the speed limitation prescribed C paragraph (c)(2)(i) of this section mul- for the remainder of the flight must be tiplied by a factor of safety that de- determined: pends on the probability of being in (A) the limit symmetrical maneu- this failure condition. The factor of vering conditions specified in §§ 25.331 safety is defined in Figure 2.

and 25.345, 14 CFR Ch. I (1–1–25 Edition) § 25.302 Figure 2 to paragraph (c)(2)(ii) Qj = (Tj)(Pj) where: that reduce the reliability of the re- Tj = Average time spent in failure condition maining system. As far as practicable, j (in hours) these failures must be indicated to the Pj = Probability of occurrence of failure flightcrew before flight.

mode j (per hour) ¥ 3 (2) The existence of any failure condi- If Pj is greater than 10 per flight hour, tion evaluated under paragraph (c) of then a 1.5 factor of safety must be ap- plied in lieu of the factor of safety de- this section that results in a factor of fined in Figure 2.

safety between the airplane strength and the loads of subpart C of this part (iii) For residual strength substan- below 1.25 must be indicated to the tiation, the airplane must be able to withstand two thirds of the ultimate flightcrew.

loads defined in paragraph (c)(2)(ii) of (e) Dispatch with known failure con- this section. For pressurized cabins, ditions. If the airplane is to be dis- these loads must be combined with the patched in a known system failure con- normal operating differential pressure.

dition that affects structural perform- (iv) If the loads induced by the fail- ance or affects the reliability of the re- ure condition have a significant effect maining system to maintain structural on fatigue or damage tolerance then performance, then the Master Min- their effects must be taken into ac- imum Equipment List must ensure the count.

provisions of § 25.302 are met for the (v)–(vi) [Reserved] dispatched condition and for any subse- (3) [Reserved] quent failures. Flight limitations and (d) Failure indications. For system operational limitations may be taken failure detection and indication, the into account in establishing Qj as the following apply: combined probability of being in the (1) The system must be checked for dispatched failure condition and the failure conditions evaluated under subsequent failure condition for the paragraph (c) of this section that de- grade the structural capability below safety margins in Figure 2. No reduc- the level required by subparts C (ex- tion in these safety margins is allowed cluding § 25.302) and D of this part or Federal Aviation Administration, DOT § 25.321 if the subsequent system failure rate is by analysis, flight tests, or other tests ¥ 3 greater than 10 per flight hour. found necessary by the Administrator.

(f) Unless shown to be extremely im- [Doc. No. FAA–2022–1544, 89 FR 68732, Aug. 27, probable, the airplane must be designed 2024] to withstand any forced structural vi- § 25.303 Factor of safety. bration resulting from any failure, malfunction or adverse condition in Unless otherwise specified, a factor of the flight control system. These must safety of 1.5 must be applied to the pre- be considered limit loads and must be scribed limit load which are considered investigated at airspeeds up to V /M .

C C external loads on the structure. When a loading condition is prescribed in [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as terms of ultimate loads, a factor of amended by Amdt. 25–23, 35 FR 5672, Apr. 8, safety need not be applied unless other- 1970; Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; wise specified. Amdt. 25–77, 57 FR 28949, June 29, 1992; Amdt.

25–86, 61 FR 5220, Feb. 9, 1996] [Amdt. 25–23, 35 FR 5672, Apr. 8, 1970] § 25.307 Proof of structure.

§ 25.305 Strength and deformation.

(a) Compliance with the strength and (a) The structure must be able to deformation requirements of this sub- support limit loads without detri- part must be shown for each critical mental permanent deformation. At any loading condition. Structural analysis load up to limit loads, the deformation may be used only if the structure con- may not interfere with safe operation.

forms to that for which experience has (b) The structure must be able to shown this method to be reliable. In support ultimate loads without failure other cases, substantiating tests must for at least 3 seconds. However, when be made to load levels that are suffi- proof of strength is shown by dynamic cient to verify structural behavior up tests simulating actual load condi- to loads specified in § 25.305.

tions, the 3-second limit does not (b)–(c) [Reserved] apply. Static tests conducted to ulti- (d) When static or dynamic tests are mate load must include the ultimate used to show compliance with the re- deflections and ultimate deformation quirements of § 25.305(b) for flight induced by the loading. When analyt- structures, appropriate material cor- ical methods are used to show compli- rection factors must be applied to the ance with the ultimate load strength test results, unless the structure, or requirements, it must be shown that— part thereof, being tested has features (1) The effects of deformation are not such that a number of elements con- significant; tribute to the total strength of the (2) The deformations involved are structure and the failure of one ele- fully accounted for in the analysis; or ment results in the redistribution of (3) The methods and assumptions the load through alternate load paths.

used are sufficient to cover the effects of these deformations.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) Where structural flexibility is amended by Amdt. 25–23, 35 FR 5672, Apr. 8, such that any rate of load application 1970; Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; likely to occur in the operating condi- Amdt. 25–72, 55 FR 29775, July 20, 1990; Amdt.

25–139, 79 FR 59429, Oct. 2, 2014] tions might produce transient stresses appreciably higher than those cor- F LIGHT L OADS responding to static loads, the effects of this rate of application must be con- § 25.321 General.

sidered.

(d) [Reserved] (a) Flight load factors represent the (e) The airplane must be designed to ratio of the aerodynamic force compo- withstand any vibration and buffeting nent (acting normal to the assumed that might occur in any likely oper- longitudinal axis of the airplane) to the ating condition up to V /M , including weight of the airplane. A positive load D D stall and probable inadvertent excur- factor is one in which the aerodynamic sions beyond the boundaries of the buf- force acts upward with respect to the fet onset envelope. This must be shown airplane.

14 CFR Ch. I (1–1–25 Edition) § 25.331 (b) Considering compressibility ef- corresponding pitching velocities must fects at each speed, compliance with be taken into account. The in-trim and the flight load requirements of this out-of-trim flight conditions specified subpart must be shown— in § 25.255 must be considered.

(b) Maneuvering balanced conditions.

(1) At each critical altitude within Assuming the airplane to be in equi- the range of altitudes selected by the librium with zero pitching accelera- applicant; tion, the maneuvering conditions A (2) At each weight from the design through I on the maneuvering envelope minimum weight to the design max- in § 25.333(b) must be investigated.

imum weight appropriate to each par- (c) Maneuvering pitching conditions.

ticular flight load condition; and The following conditions must be in- (3) For each required altitude and vestigated: weight, for any practicable distribution (1) Maximum pitch control displacement of disposable load within the operating at V . The airplane is assumed to be A limitations recorded in the Airplane flying in steady level flight (point A , Flight Manual.

§ 25.333(b)) and the cockpit pitch con- (c) Enough points on and within the trol is suddenly moved to obtain ex- boundaries of the design envelope must treme nose up pitching acceleration. In be investigated to ensure that the max- defining the tail load, the response of imum load for each part of the airplane the airplane must be taken into ac- structure is obtained.

count. Airplane loads that occur subse- (d) The significant forces acting on quent to the time when normal accel- the airplane must be placed in equi- eration at the c.g. exceeds the positive librium in a rational or conservative limit maneuvering load factor (at point manner. The linear inertia forces must A in § 25.333(b)), or the resulting be considered in equilibrium with the tailplane normal load reaches its max- thrust and all aerodynamic loads, imum, whichever occurs first, need not while the angular (pitching) inertia be considered.

forces must be considered in equi- (2) Checked maneuver between V and A librium with thrust and all aero- V . Nose-up checked pitching maneu- D dynamic moments, including moments vers must be analyzed in which the due to loads on components such as positive limit load factor prescribed in tail surfaces and nacelles. Critical § 25.337 is achieved. As a separate condi- thrust values in the range from zero to tion, nose-down checked pitching ma- maximum continuous thrust must be neuvers must be analyzed in which a considered.

limit load factor of 0g is achieved. In [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as defining the airplane loads, the flight amended by Amdt. 25–23, 35 FR 5672, Apr. 8, deck pitch control motions described 1970; Amdt. 25–86, 61 FR 5220, Feb. 9, 1996] in paragraphs (c)(2)(i) through (iv) of this section must be used: F LIGHT M ANEUVER AND G UST (i) The airplane is assumed to be fly- C ONDITIONS ing in steady level flight at any speed between V and V and the flight deck A D § 25.331 Symmetric maneuvering con- pitch control is moved in accordance ditions.

with the following formula: (a) Procedure. For the analysis of the d (t) = d sin( w t) for 0 ≤ t ≤ t 1 max maneuvering flight conditions specified in paragraphs (b) and (c) of this sec- Where— tion, the following provisions apply: d = the maximum available displacement of (1) Where sudden displacement of a the flight deck pitch control in the ini- control is specified, the assumed rate tial direction, as limited by the control system stops, control surface stops, or by of control surface displacement may pilot effort in accordance with § 25.397(b); not be less than the rate that could be d (t) = the displacement of the flight deck applied by the pilot through the con- pitch control as a function of time. In trol system.

the initial direction, d (t) is limited to d .

(2) In determining elevator angles In the reverse direction, d (t) may be and chordwise load distribution in the truncated at the maximum available dis- maneuvering conditions of paragraphs placement of the flight deck pitch con- (b) and (c) of this section, the effect of trol as limited by the control system Federal Aviation Administration, DOT § 25.331 stops, control surface stops, or by pilot undamped natural frequency of the short effort in accordance with 25.397(b); period rigid mode of the airplane, with t = 3 π /2 w ; max active control system effects included w = the circular frequency (radians/second) of where appropriate; but not less than: the control deflection taken equal to the Where (iii) In addition, for cases where the airplane response to the specified flight V = the speed of the airplane at entry to the deck pitch control motion does not maneuver.

V = the design maneuvering speed pre- achieve the prescribed limit load fac- A scribed in § 25.335(c). tors, then the following flight deck pitch control motion must be used: (ii) For nose-up pitching maneuvers, d (t) = d sin( w t) for 0 ≤ t ≤ t 1 1 the complete flight deck pitch control d (t) = d for t ≤ t ≤ t 1 1 2 displacement history may be scaled d (t) = d sin( w [t + t ¥ t ]) for t ≤ t ≤ 1 1 2 2 down in amplitude to the extent nec- t max essary to ensure that the positive limit Where— load factor prescribed in § 25.337 is not exceeded. For nose-down pitching ma- = π /2 w t 1 = t + D t neuvers, the complete flight deck con- t 2 1 = t + π / w ; t max 2 trol displacement history may be D t = the minimum period of time necessary scaled down in amplitude to the extent to allow the prescribed limit load factor necessary to ensure that the normal to be achieved in the initial direction, acceleration at the center of gravity but it need not exceed five seconds (see does not go below 0g.

figure below).

(iv) In cases where the flight deck (A) For the nose-up pitching maneu- pitch control motion may be affected ver, the time at which the normal ac- by inputs from systems (for example, celeration at the center of gravity goes by a stick pusher that can operate at below 0g; high load factor as well as at 1g), then (B) For the nose-down pitching ma- the effects of those systems shall be neuver, the time at which the normal taken into account. acceleration at the center of gravity (v) Airplane loads that occur beyond goes above the positive limit load fac- the following times need not be consid- tor prescribed in § 25.337; ered: 14 CFR Ch. I (1–1–25 Edition) § 25.333 (C) t .

max. § 25.333 Flight maneuvering envelope.

(a) General. The strength require- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5672, Apr. 8, ments must be met at each combina- 1970; Amdt. 25–46, 43 FR 50594, Oct. 30, 1978; 43 tion of airspeed and load factor on and FR 52495, Nov. 13, 1978; 43 FR 54082, Nov. 20, within the boundaries of the represent- 1978; Amdt. 25–72, 55 FR 29775, July 20, 1990; 55 ative maneuvering envelope ( V-n dia- FR 37607, Sept. 12, 1990; Amdt. 25–86, 61 FR gram) of paragraph (b) of this section.

5220, Feb. 9, 1996; Amdt. 25–91, 62 FR 40704, This envelope must also be used in de- July 29, 1997; Amdt. 25–141, 79 FR 73466, Dec.

termining the airplane structural oper- 11, 2014] ating limitations as specified in § 25.1501.

(b) Maneuvering envelope.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–86, 61 FR 5220, Feb. 9, 1996] (2) Except as provided in § 25.335(d)(2), § 25.335 Design airspeeds.

V may not be less than V + 1.32 U C B REF The selected design airspeeds are (with U as specified in REF equivalent airspeeds (EAS). Estimated § 25.341(a)(5)(i)). However V need not C values of V and V must be conserv- S 0 S 1 exceed the maximum speed in level ative.

flight at maximum continuous power (a) Design cruising speed, V . For V C C, for the corresponding altitude.

the following apply: (3) At altitudes where V is limited D (1) The minimum value of V must be C by Mach number, V may be limited to C sufficiently greater than V to provide B a selected Mach number.

for inadvertent speed increases likely (b) Design dive speed, V . V must be D D to occur as a result of severe atmos- selected so that V / M is not greater C C pheric turbulence.

than 0.8 V / M or so that the minimum D D, Federal Aviation Administration, DOT § 25.335 U = the reference gust velocity (feet per speed margin between V / M and V / M ref C C D D second equivalent airspeed) from is the greater of the following values: § 25.341(a)(5)(i); (1) From an initial condition of sta- w = average wing loading (pounds per square bilized flight at V / M the airplane is C C, foot) at the particular weight under con- upset, flown for 20 seconds along a sideration.

flight path 7.5 ° below the initial path, and then pulled up at a load factor of . 88 μ 1.5 g (0.5 g acceleration increment). The K = g speed increase occurring in this maneu- + . 5 3 μ ver may be calculated if reliable or conservative aerodynamic data is used.

w 2 = μ Power as specified in § 25.175(b)(1)(iv) is cag ρ assumed until the pullup is initiated, at which time power reduction and the 3 r = density of air (slugs/ft ); use of pilot controlled drag devices c = mean geometric chord of the wing (feet); g = acceleration due to gravity (ft/sec ); may be assumed; a = slope of the airplane normal force coeffi- (2) The minimum speed margin must cient curve, C per radian; NA be enough to provide for atmospheric variations (such as horizontal gusts, (2) At altitudes where V is limited C and penetration of jet streams and cold by Mach number— fronts) and for instrument errors and (i) V may be chosen to provide an B airframe production variations. These optimum margin between low and high factors may be considered on a prob- speed buffet boundaries; and, ability basis. The margin at altitude (ii) V need not be greater than V .

B C where M is limited by compressibility C (e) Design flap speeds, V . For V , the F F effects must not less than 0.07M unless following apply: a lower margin is determined using a (1) The design flap speed for each flap rational analysis that includes the ef- position (established in accordance fects of any automatic systems. In any with § 25.697(a)) must be sufficiently case, the margin may not be reduced to greater than the operating speed rec- less than 0.05M.

ommended for the corresponding stage (c) Design maneuvering speed V . For A of flight (including balked landings) to V , the following apply: A allow for probable variations in control (1) V may not be less than V √ n A S 1 of airspeed and for transition from one where— flap position to another.

(i) n is the limit positive maneu- (2) If an automatic flap positioning or vering load factor at V ; and C load limiting device is used, the speeds (ii) V is the stalling speed with flaps S 1 and corresponding flap positions pro- retracted.

grammed or allowed by the device may (2) V and V must be evaluated at A S be used.

the design weight and altitude under (3) V may not be less than— F consideration.

(i) 1.6 V with the flaps in takeoff po- S 1 need not be more than V or (3) V A C sition at maximum takeoff weight; the speed at which the positive C N max (ii) 1.8 V with the flaps in approach S 1 curve intersects the positive maneuver position at maximum landing weight, load factor line, whichever is less.

and (d) Design speed for maximum gust in- (iii) 1.8 V with the flaps in landing S 0 tensity, V .

B position at maximum landing weight.

(1) V may not be less than B (f) Design drag device speeds, V . The DD selected design speed for each drag de- 1 2 vice must be sufficiently greater than K U V a ⎡ ⎤ g ref c V 1 + the speed recommended for the oper- S 1 ⎢ ⎥ ation of the device to allow for prob- w 498 ⎣ ⎦ able variations in speed control. For where— drag devices intended for use in high V = the 1-g stalling speed based on C S1 NAmax speed descents, V may not be less DD with the flaps retracted at the particular than V . When an automatic drag de- D weight under consideration; vice positioning or load limiting means V = design cruise speed (knots equivalent c airspeed); is used, the speeds and corresponding 14 CFR Ch. I (1–1–25 Edition) § 25.337 drag device positions programmed or ⎡ ⎤ allowed by the automatic means must U ⎛ ⎞ s π ds U = 1- Cos ⎜ ⎟ ⎢ ⎥ be used for design.

⎝ ⎠ 2 H ⎣ ⎦ [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as for 0 ≤ s ≤ 2H amended by Amdt. 25–23, 35 FR 5672, Apr. 8, 1970; Amdt. 25–86, 61 FR 5220, Feb. 9, 1996; where— s = distance penetrated into the gust (feet); Amdt. 25–91, 62 FR 40704, July 29, 1997] U = the design gust velocity in equivalent ds § 25.337 Limit maneuvering load fac- airspeed specified in paragraph (a)(4) of tors. this section; and H = the gust gradient which is the distance (a) Except where limited by max- (feet) parallel to the airplane’s flight imum (static) lift coefficients, the air- path for the gust to reach its peak veloc- plane is assumed to be subjected to ity.

symmetrical maneuvers resulting in (3) A sufficient number of gust gra- the limit maneuvering load factors pre- dient distances in the range 30 feet to scribed in this section. Pitching veloci- 350 feet must be investigated to find ties appropriate to the corresponding the critical response for each load pull-up and steady turn maneuvers quantity.

must be taken into account.

(4) The design gust velocity must be: (b) The positive limit maneuvering load factor n for any speed up to Vn 1 6 may not be less than 2.1 + 24,000/ ( W + H U U F =

( )

ds ref g 10,000) except that n may not be less than 2.5 and need not be greater than where— 3.8—where W is the design maximum U = the reference gust velocity in equiva- ref takeoff weight. lent airspeed defined in paragraph (a)(5) of this section.

(c) The negative limit maneuvering F = the flight profile alleviation factor de- g load factor— fined in paragraph (a)(6) of this section.

(1) May not be less than ¥ 1.0 at speeds up to V ; and C (5) The following reference gust ve- (2) Must vary linearly with speed locities apply: from the value at V to zero at V .

C D (i) At airplane speeds between V and B (d) Maneuvering load factors lower V : Positive and negative gusts with C than those specified in this section reference gust velocities of 56.0 ft/sec may be used if the airplane has design EAS must be considered at sea level.

features that make it impossible to ex- The reference gust velocity may be re- ceed these values in flight.

duced linearly from 56.0 ft/sec EAS at sea level to 44.0 ft/sec EAS at 15,000 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as feet. The reference gust velocity may amended by Amdt. 25–23, 35 FR 5672, Apr. 8, 1970] be further reduced linearly from 44.0 ft/ sec EAS at 15,000 feet to 20.86 ft/sec § 25.341 Gust and turbulence loads.

EAS at 60,000 feet.

(ii) At the airplane design speed V : (a) Discrete Gust Design Criteria. The D The reference gust velocity must be 0.5 airplane is assumed to be subjected to times the value obtained under symmetrical vertical and lateral gusts § 25.341(a)(5)(i).

in level flight. Limit gust loads must be determined in accordance with the (6) The flight profile alleviation fac- tor, F , must be increased linearly from provisions: g the sea level value to a value of 1.0 at (1) Loads on each part of the struc- the maximum operating altitude de- ture must be determined by dynamic fined in § 25.1527. At sea level, the flight analysis. The analysis must take into account unsteady aerodynamic charac- profile alleviation factor is determined by the following equation: teristics and all significant structural degrees of freedom including rigid body motions.

(2) The shape of the gust must be: Federal Aviation Administration, DOT § 25.341 (b) Continuous turbulence design cri- teria. The dynamic response of the air- F F F = + 0 5 .

( )

g gz gm plane to vertical and lateral contin- uous turbulence must be taken into ac- Where : count. The dynamic analysis must take into account unsteady aerodynamic Z mo characteristics and all significant F = − 1 ; gz structural degrees of freedom including 250000 rigid body motions. The limit loads must be determined for all critical alti- R π ⎛ ⎞ tudes, weights, and weight distribu- F R Tan = ; gm 2 ⎝ ⎠ tions as specified in § 25.321(b), and all critical speeds within the ranges indi- Maximum Landing Weight cated in § 25.341(b)(3).

R = ; (1) Except as provided in paragraphs Maximum Take off Weight - (b)(4) and (5) of this section, the fol- lowing equation must be used: Maximum Zero Fuel Weight ¯ P = P ± U A L L ¥ 1 g σ R = ; Where— Maximum Take off Weight - P = limit load; L Z = Maximum operating altitude defined in mo P = steady 1g load for the condition; L ¥ 1g § 25.1527 (feet).

A = ratio of root-mean-square incremental load for the condition to root-mean- (7) When a stability augmentation square turbulence velocity; and system is included in the analysis, the U = limit turbulence intensity in true air- σ effect of any significant system non- speed, specified in paragraph (b)(3) of this linearities should be accounted for section.

when deriving limit loads from limit (2) Values of A must be determined gust conditions.

according to the following formula: Where— lates the loads in the aircraft structure to the atmospheric turbulence; and H( W ) = the frequency response function, de- F ( W ) = normalized power spectral density of termined by dynamic analysis, that re- atmospheric turbulence given by— Where— (i) At airplane speeds between V and B V : W = reduced frequency, radians per foot; and C L = scale of turbulence = 2,500 ft.

U = U F σ σ ref g (3) The limit turbulence intensities, Where— U , in feet per second true airspeed re- σ U σ ref is the reference turbulence intensity quired for compliance with this para- that varies linearly with altitude from 90 graph are— fps (TAS) at sea level to 79 fps (TAS) at 14 CFR Ch. I (1–1–25 Edition) § 25.343 24,000 feet and is then constant at 79 fps field as A U of the same load quantity σ (TAS) up to the altitude of 60,000 feet.

in a linear approximated model.

F is the flight profile alleviation factor de- g (c) Supplementary gust conditions for fined in paragraph (a)(6) of this section; wing-mounted engines. For airplanes (ii) At speed V : U is equal to ⁄2 the equipped with wing-mounted engines, D σ values obtained under paragraph the engine mounts, pylons, and wing (b)(3)(i) of this section. supporting structure must be designed (iii) At speeds between V and V : U for the maximum response at the na- C D σ is equal to a value obtained by linear celle center of gravity derived from the interpolation. following dynamic gust conditions ap- (iv) At all speeds, both positive and plied to the airplane: negative incremental loads due to con- (1) A discrete gust determined in ac- tinuous turbulence must be considered. cordance with § 25.341(a) at each angle (4) When an automatic system affect- normal to the flight path, and sepa- ing the dynamic response of the air- rately, plane is included in the analysis, the (2) A pair of discrete gusts, one effects of system non-linearities on vertical and one lateral. The length of loads at the limit load level must be each of these gusts must be independ- taken into account in a realistic or ently tuned to the maximum response conservative manner. in accordance with § 25.341(a). The pene- (5) If necessary for the assessment of tration of the airplane in the combined loads on airplanes with significant non- gust field and the phasing of the linearities, it must be assumed that vertical and lateral component gusts the turbulence field has a root-mean- must be established to develop the square velocity equal to 40 percent of maximum response to the gust pair. In the U values specified in paragraph the absence of a more rational anal- σ (b)(3) of this section. The value of limit ysis, the following formula must be load is that load with the same prob- used for each of the maximum engine ability of exceedance in the turbulence loads in all six degrees of freedom: Where— minimum fuel weight condition for P = limit load; showing compliance with the flight L P L-1g = steady 1g load for the condition; load requirements as prescribed in this = peak incremental response load due to L V subpart. In addition— a vertical gust according to § 25.341(a); (1) The structure must be designed and for a condition of zero fuel and oil in L = peak incremental response load due to L a lateral gust according to § 25.341(a).

the wing at limit loads corresponding to— [Doc. No. 27902, 61 FR 5221, Feb. 9, 1996; 61 FR 9533, Mar. 8, 1996; Doc. No. FAA–2013–0142; 79 (i) A maneuvering load factor of + FR 73467, Dec. 11, 2014; Amdt. 25–141, 80 FR 2.25; and 4762, Jan. 29, 2015; 80 FR 6435, Feb. 5, 2015] (ii) The gust and turbulence condi- tions of § 25.341(a) and (b), but assuming § 25.343 Design fuel and oil loads.

85% of the gust velocities prescribed in (a) The disposable load combinations § 25.341(a)(4) and 85% of the turbulence must include each fuel and oil load in intensities prescribed in § 25.341(b)(3).

the range from zero fuel and oil to the (2) Fatigue evaluation of the struc- selected maximum fuel and oil load. A ture must account for any increase in structural reserve fuel condition, not operating stresses resulting from the exceeding 45 minutes of fuel under the design condition of paragraph (b)(1) of operating conditions in § 25.1001(e) and this section; and (f), as applicable, may be selected.

(b) If a structural reserve fuel condi- tion is selected, it must be used as the Federal Aviation Administration, DOT § 25.349 (3) The flutter, deformation, and vi- (1) Maneuvering to a positive limit bration requirements must also be met load factor as prescribed in § 25.337(b); with zero fuel. and (2) The vertical gust and turbulence [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as conditions prescribed in § 25.341(a) and amended by Amdt. 25–18, 33 FR 12226, Aug. 30, (b).

1968; Amdt. 25–72, 55 FR 37607, Sept. 12, 1990; (d) The airplane must be designed for Amdt. 25–86, 61 FR 5221, Feb. 9, 1996; Amdt.

25–141, 79 FR 73468, Dec. 11, 2014] a maneuvering load factor of 1.5 g at the maximum take-off weight with the § 25.345 High lift devices.

wing-flaps and similar high lift devices in the landing configurations.

(a) If wing flaps are to be used during takeoff, approach, or landing, at the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as design flap speeds established for these amended by Amdt. 25–46, 43 FR 50595, Oct. 30, stages of flight under § 25.335(e) and 1978; Amdt. 25–72, 55 FR 37607, Sept. 17, 1990; Amdt. 25–86, 61 FR 5221, Feb. 9, 1996; Amdt.

with the wing flaps in the cor- 25–91, 62 FR 40704, July 29, 1997; Amdt. 25–141, responding positions, the airplane is 79 FR 73468, Dec. 11, 2014] assumed to be subjected to symmet- rical maneuvers and gusts. The result- § 25.349 Rolling conditions.

ing limit loads must correspond to the The airplane must be designed for conditions determined as follows: loads resulting from the rolling condi- (1) Maneuvering to a positive limit tions specified in paragraphs (a) and (b) load factor of 2.0; and of this section. Unbalanced aero- (2) Positive and negative gusts of 25 dynamic moments about the center of ft/sec EAS acting normal to the flight gravity must be reacted in a rational path in level flight. Gust loads result- or conservative manner, considering ing on each part of the structure must the principal masses furnishing the re- be determined by rational analysis.

acting inertia forces.

The analysis must take into account (a) Maneuvering. The following condi- the unsteady aerodynamic characteris- tions, speeds, and aileron deflections tics and rigid body motions of the air- (except as the deflections may be lim- craft. The shape of the gust must be as ited by pilot effort) must be considered described in § 25.341(a)(2) except that— in combination with an airplane load U = 25 ft/sec EAS; ds factor of zero and of two-thirds of the H = 12.5 c; and positive maneuvering factor used in de- c = mean geometric chord of the wing (feet).

sign. In determining the required aile- (b) The airplane must be designed for ron deflections, the torsional flexi- the conditions prescribed in paragraph bility of the wing must be considered (a) of this section, except that the air- in accordance with § 25.301(b): plane load factor need not exceed 1.0, (1) Conditions corresponding to taking into account, as separate condi- steady rolling velocities must be inves- tions, the effects of— tigated. In addition, conditions cor- (1) Propeller slipstream cor- responding to maximum angular accel- responding to maximum continuous eration must be investigated for air- power at the design flap speeds V and planes with engines or other weight F, with takeoff power at not less than 1.4 concentrations outboard of the fuse- times the stalling speed for the par- lage. For the angular acceleration con- ticular flap position and associated ditions, zero rolling velocity may be maximum weight; and assumed in the absence of a rational (2) A head-on gust of 25 feet per sec- time history investigation of the ma- ond velocity (EAS). neuver.

(c) If flaps or other high lift devices (2) At V a sudden deflection of the A, are to be used in en route conditions, aileron to the stop is assumed.

and with flaps in the appropriate posi- (3) At V the aileron deflection must C, tion at speeds up to the flap design be that required to produce a rate of speed chosen for these conditions, the roll not less than that obtained in airplane is assumed to be subjected to paragraph (a)(2) of this section.

symmetrical maneuvers and gusts (4) At V the aileron deflection must D, within the range determined by— be that required to produce a rate of 14 CFR Ch. I (1–1–25 Edition) § 25.351 roll not less than one-third of that in (d) With the airplane yawed to the paragraph (a)(2) of this section. static equilibrium sideslip angle of paragraph (c) of this section, it is as- (b) Unsymmetrical gusts. The airplane sumed that the cockpit rudder control is assumed to be subjected to unsym- is suddenly returned to neutral.

metrical vertical gusts in level flight.

The resulting limit loads must be de- [Amdt. 25–91, 62 FR 40704, July 29, 1997] termined from either the wing max- imum airload derived directly from § 25.353 Rudder control reversal condi- tions.

§ 25.341(a), or the wing maximum air- load derived indirectly from the Airplanes with a powered rudder con- vertical load factor calculated from trol surface or surfaces must be de- § 25.341(a). It must be assumed that 100 signed for loads, considered to be ulti- percent of the wing air load acts on one mate, resulting from the yaw maneu- side of the airplane and 80 percent of ver conditions specified in paragraphs the wing air load acts on the other (a) through (e) of this section at speeds side.

from V to V /M . Any permanent de- MC C C formation resulting from these ulti- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as mate load conditions must not prevent amended by Amdt. 25–23, 35 FR 5672, Apr. 8, continued safe flight and landing. The 1970; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; applicant must evaluate these condi- Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] tions with the landing gear retracted § 25.351 Yaw maneuver conditions.

and speed brakes (and spoilers when used as speed brakes) retracted. The The airplane must be designed for applicant must evaluate the effects of loads resulting from the yaw maneuver flaps, flaperons, or any other aero- conditions specified in paragraphs (a) dynamic devices when used as flaps, through (d) of this section at speeds and slats-extended configurations, if from V to V . Unbalanced aero- MC D they are used in en route conditions.

dynamic moments about the center of Unbalanced aerodynamic moments gravity must be reacted in a rational about the center of gravity must be re- or conservative manner considering the acted in a rational or conservative airplane inertia forces. In computing manner considering the airplane iner- the tail loads the yawing velocity may tia forces. In computing the loads on be assumed to be zero.

the airplane, the yawing velocity may (a) With the airplane in unacceler- be assumed to be zero. The applicant ated flight at zero yaw, it is assumed must assume a pilot force of 200 pounds that the cockpit rudder control is sud- when evaluating each of the following denly displaced to achieve the result- conditions: ing rudder deflection, as limited by: (a) With the airplane in unacceler- (1) The control system on control ated flight at zero yaw, the flightdeck surface stops; or rudder control is suddenly and fully (2) A limit pilot force of 300 pounds displaced to achieve the resulting rud- from V to V and 200 pounds from V / MC A C der deflection, as limited by the con- M to V /M , with a linear variation C D D trol system or the control surface between V and V /M .

A C C stops.

(b) With the cockpit rudder control (b) With the airplane yawed to the deflected so as always to maintain the overswing sideslip angle, the flightdeck maximum rudder deflection available rudder control is suddenly and fully within the limitations specified in displaced in the opposite direction, as paragraph (a) of this section, it is as- limited by the control system or con- sumed that the airplane yaws to the trol surface stops.

overswing sideslip angle.

(c) With the airplane yawed to the (c) With the airplane yawed to the opposite overswing sideslip angle, the static equilibrium sideslip angle, it is flightdeck rudder control is suddenly assumed that the cockpit rudder con- and fully displaced in the opposite di- trol is held so as to achieve the max- rection, as limited by the control sys- imum rudder deflection available with- tem or control surface stops.

in the limitations specified in para- (d) With the airplane yawed to the graph (a) of this section. subsequent overswing sideslip angle, Federal Aviation Administration, DOT § 25.363 the flightdeck rudder control is sud- by each of the following conditions to denly and fully displaced in the oppo- be considered separately: site direction, as limited by the control (i) Sudden maximum engine decelera- system or control surface stops. tion due to malfunction or abnormal (e) With the airplane yawed to the condition; and opposite overswing sideslip angle, the (ii) The maximum acceleration of en- flightdeck rudder control is suddenly gine.

returned to neutral. (b) For auxiliary power unit installa- tions, the power unit mounts and adja- [Amdt. No. 25–147, 87 FR 71210, Nov. 22, 2022] cent supporting airframe structure must be designed to withstand 1g level S UPPLEMENTARY C ONDITIONS flight loads acting simultaneously with § 25.361 Engine and auxiliary power the limit torque loads imposed by each unit torque.

of the following conditions to be con- sidered separately: (a) For engine installations— (1) Sudden maximum auxiliary power (1) Each engine mount, pylon, and ad- unit deceleration due to malfunction, jacent supporting airframe structures abnormal condition, or structural fail- must be designed for the effects of— ure; and (i) A limit engine torque cor- (2) The maximum acceleration of the responding to takeoff power/thrust and, auxiliary power unit.

if applicable, corresponding propeller speed, acting simultaneously with 75% [Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] of the limit loads from flight condition A of § 25.333(b); § 25.362 Engine failure loads.

(ii) A limit engine torque cor- (a) For engine mounts, pylons, and responding to the maximum contin- adjacent supporting airframe struc- uous power/thrust and, if applicable, ture, an ultimate loading condition corresponding propeller speed, acting must be considered that combines 1g simultaneously with the limit loads flight loads with the most critical from flight condition A of § 25.333(b); transient dynamic loads and vibra- and tions, as determined by dynamic anal- (iii) For turbopropeller installations ysis, resulting from failure of a blade, only, in addition to the conditions shaft, bearing or bearing support, or specified in paragraphs (a)(1)(i) and (ii) bird strike event. Any permanent de- of this section, a limit engine torque formation from these ultimate load corresponding to takeoff power and conditions must not prevent continued propeller speed, multiplied by a factor safe flight and landing.

accounting for propeller control sys- (b) The ultimate loads developed tem malfunction, including quick from the conditions specified in para- feathering, acting simultaneously with graph (a) of this section are to be— 1g level flight loads. In the absence of (1) Multiplied by a factor of 1.0 when a rational analysis, a factor of 1.6 must applied to engine mounts and pylons; be used.

and (2) The limit engine torque to be con- (2) Multiplied by a factor of 1.25 when sidered under paragraph (a)(1) of this applied to adjacent supporting air- section must be obtained by— frame structure.

(i) For turbopropeller installations, multiplying mean engine torque for the [Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] specified power/thrust and speed by a § 25.363 Side load on engine and auxil- factor of 1.25; iary power unit mounts.

(ii) For other turbine engines, the limit engine torque must be equal to (a) Each engine and auxiliary power the maximum accelerating torque for unit mount and its supporting struc- the case considered. ture must be designed for a limit load (3) The engine mounts, pylons, and factor in lateral direction, for the side adjacent supporting airframe structure load on the engine and auxiliary power must be designed to withstand 1g level unit mount, at least equal to the max- flight loads acting simultaneously with imum load factor obtained in the yaw- the limit engine torque loads imposed ing conditions but not less than— 14 CFR Ch. I (1–1–25 Edition) § 25.365 (1) 1.33; or openings that cannot reasonably be ex- (2) One-third of the limit load factor pected to be confined to the small com- for flight condition A as prescribed in partment. The size H must be com- o § 25.333(b).

puted by the following formula: (b) The side load prescribed in para- H = PA o s graph (a) of this section may be as- sumed to be independent of other flight where, conditions.

H o = Maximum opening in square feet, need [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as not exceed 20 square feet.

amended by Amdt. 25–23, 35 FR 5672, Apr. 8, P = (A /6240) + .024 s 1970; Amdt. 25–91, 62 FR 40704, July 29, 1997] A s = Maximum cross-sectional area of the pressurized shell normal to the longitu- § 25.365 Pressurized compartment dinal axis, in square feet; and loads.

(3) The maximum opening caused by For airplanes with one or more pres- airplane or equipment failures not surized compartments the following shown to be extremely improbable.

apply: (f) In complying with paragraph (e) of (a) The airplane structure must be this section, the fail-safe features of strong enough to withstand the flight the design may be considered in deter- loads combined with pressure differen- mining the probability of failure or tial loads from zero up to the max- penetration and probable size of open- imum relief valve setting.

(b) The external pressure distribution ings, provided that possible improper in flight, and stress concentrations and operation of closure devices and inad- fatigue effects must be accounted for.

vertent door openings are also consid- (c) If landings may be made with the ered. Furthermore, the resulting dif- compartment pressurized, landing ferential pressure loads must be com- loads must be combined with pressure bined in a rational and conservative differential loads from zero up to the manner with 1–g level flight loads and maximum allowed during landing.

any loads arising from emergency de- (d) The airplane structure must be pressurization conditions. These loads designed to be able to withstand the may be considered as ultimate condi- pressure differential loads cor- tions; however, any deformations asso- responding to the maximum relief ciated with these conditions must not valve setting multiplied by a factor of interfere with continued safe flight and 1.33 for airplanes to be approved for op- landing. The pressure relief provided by eration to 45,000 feet or by a factor of intercompartment venting may also be 1.67 for airplanes to be approved for op- considered.

eration above 45,000 feet, omitting (g)(1) Except as provided in para- other loads.

graph (g)(2) of this section, bulkheads, (e) Any structure, component or part, floors, and partitions in pressurized inside or outside a pressurized com- partment, the failure of which could compartments for occupants must be interfere with continued safe flight and designed to withstand the conditions landing, must be designed to withstand specified in paragraph (e) of this sec- the effects of a sudden release of pres- tion. In addition, reasonable design sure through an opening in any com- precautions must be taken to minimize partment at any operating altitude re- the probability of parts becoming de- sulting from each of the following con- tached and injuring occupants while in ditions: their seats.

(1) The penetration of the compart- (2) Partitions adjacent to the opening ment by a portion of an engine fol- specified in paragraph (e)(2) of this sec- lowing an engine disintegration; tion need not be designed to withstand (2) Any opening in any pressurized that condition provided— compartment up to the size H in o (i) Failure of the partition would not square feet; however, small compart- interfere with continued safe flight and ments may be combined with an adja- landing; and cent pressurized compartment and both considered as a single compartment for Federal Aviation Administration, DOT § 25.391 (ii) Designing the partition to with- signed for the loads, including gyro- stand the condition specified in para- scopic loads, arising from the condi- graph (e)(2) of this section would be im- tions specified in §§ 25.331, 25.341, 25.349, practical. 25.351, 25.473, 25.479, and 25.481, with the engine or auxiliary power unit at the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as maximum rotating speed appropriate amended by Amdt. 25–54, 45 FR 60172, Sept.

to the condition. For the purposes of 11, 1980; Amdt. 25–71, 55 FR 13477, Apr. 10, compliance with this paragraph, the 1990; Amdt. 25–72, 55 FR 29776, July 20, 1990; Amdt. 25–87, 61 FR 28695, June 5, 1996; Amdt. pitch maneuver in § 25.331(c)(1) must be No. 25–149, 88 FR 38382, June 13, 2023] carried out until the positive limit ma- neuvering load factor (point A in § 25.367 Unsymmetrical loads due to § 25.333(b)) is reached.

engine failure.

[Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] (a) The airplane must be designed for the unsymmetrical loads resulting § 25.373 Speed control devices.

from the failure of the critical engine.

If speed control devices (such as Turbopropeller airplanes must be de- spoilers and drag flaps) are installed signed for the following conditions in for use in en route conditions— combination with a single malfunction (a) The airplane must be designed for of the propeller drag limiting system, the symmetrical maneuvers prescribed considering the probable pilot correc- in §§ 25.333 and 25.337, the yawing ma- tive action on the flight controls: neuvers in § 25.351, and the vertical and (1) At speeds between V and V the MC D, lateral gust and turbulence conditions loads resulting from power failure be- prescribed in § 25.341(a) and (b) at each cause of fuel flow interruption are con- setting and the maximum speed associ- sidered to be limit loads.

ated with that setting; and (2) At speeds between V and V the MC C, (b) If the device has automatic oper- loads resulting from the disconnection ating or load limiting features, the air- of the engine compressor from the tur- plane must be designed for the maneu- bine or from loss of the turbine blades ver and gust conditions prescribed in are considered to be ultimate loads.

paragraph (a) of this section, at the (3) The time history of the thrust speeds and corresponding device posi- decay and drag build-up occurring as a tions that the mechanism allows.

result of the prescribed engine failures must be substantiated by test or other [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as data applicable to the particular en- amended by Amdt. 25–72, 55 FR 29776, July 20, gine-propeller combination. 1990; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] (4) The timing and magnitude of the probable pilot corrective action must C ONTROL S URFACE AND S YSTEM L OADS be conservatively estimated, consid- ering the characteristics of the par- § 25.391 Control surface loads: Gen- ticular engine-propeller-airplane com- eral.

bination.

The control surfaces must be de- (b) Pilot corrective action may be as- signed for the limit loads resulting sumed to be initiated at the time max- from the flight conditions in §§ 25.331, imum yawing velocity is reached, but 25.341(a) and (b), 25.349, and 25.351, con- not earlier than two seconds after the sidering the requirements for— engine failure. The magnitude of the (a) Loads parallel to hinge line, in corrective action may be based on the § 25.393; control forces specified in § 25.397(b) ex- (b) Pilot effort effects, in § 25.397; cept that lower forces may be assumed (c) Trim tab effects, in § 25.407; where it is shown by anaylsis or test (d) Unsymmetrical loads, in § 25.427; that these forces can control the yaw and and roll resulting from the prescribed (e) Auxiliary aerodynamic surfaces, engine failure conditions.

in § 25.445.

§ 25.371 Gyroscopic loads.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as The structure supporting any engine amended by Amdt. 25–86, 61 FR 5222, Feb. 9, or auxiliary power unit must be de- 1996; Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] 14 CFR Ch. I (1–1–25 Edition) § 25.393 and automatic pilot systems, must be § 25.393 Loads parallel to hinge line.

considered.

(a) Control surfaces and supporting (c) Limit pilot forces and torques. The hinge brackets must be designed for in- limit pilot forces and torques are as ertia loads acting parallel to the hinge follows: line.

(b) In the absence of more rational Maximum Minimum Control forces or forces or data, the inertia loads may be assumed torques torques to be equal to KW, where— Aileron: (1) K = 24 for vertical surfaces; Stick .............................. 100 lbs ............ 40 lbs.

(2) K = 12 for horizontal surfaces; and 1 2 Wheel .......................... 80 D in.-lbs ... 40 D in.-lbs.

(3) W = weight of the movable sur- Elevator: Stick .............................. 250 lbs ............ 100 lbs.

faces.

Wheel (symmetrical) ..... 300 lbs ............ 100 lbs.

Wheel (unsymmetrical) ......................... 100 lbs.

§ 25.395 Control system.

Rudder .............................. 300 lbs ............ 130 lbs.

(a) Longitudinal, lateral, directional, 1 The critical parts of the aileron control system must be de- signed for a single tangential force with a limit value equal to and drag control system and their sup- 1.25 times the couple force determined from these criteria.

porting structures must be designed for 2 D = wheel diameter (inches).

The unsymmetrical forces must be applied at one of the loads corresponding to 125 percent of normal handgrip points on the periphery of the control wheel.

the computed hinge moments of the movable control surface in the condi- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tions prescribed in § 25.391.

amended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; Amdt. 25–72, 55 FR 29776, July 20, 1990] (b) The system limit loads of para- graph (a) of this section need not ex- § 25.399 Dual control system.

ceed the loads that can be produced by the pilot (or pilots) and by automatic (a) Each dual control system must be or power devices operating the con- designed for the pilots operating in op- trols. position, using individual pilot forces (c) The loads must not be less than not less than— those resulting from application of the (1) 0.75 times those obtained under minimum forces prescribed in § 25.395; or § 25.397(c). (2) The minimum forces specified in § 25.397(c).

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (b) The control system must be de- amended by Amdt. 25–23, 35 FR 5672, Apr. 8, signed for pilot forces applied in the 1970; Amdt. 25–72, 55 FR 29776, July 20, 1990; same direction, using individual pilot Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] forces not less than 0.75 times those ob- § 25.397 Control system loads.

tained under § 25.395.

(a) General. The maximum and min- § 25.405 Secondary control system.

imum pilot forces, specified in para- Secondary controls, such as wheel graph (c) of this section, are assumed brake, spoiler, and tab controls, must to act at the appropriate control grips be designed for the maximum forces or pads (in a manner simulating flight that a pilot is likely to apply to those conditions) and to be reacted at the at- controls. The following values may be tachment of the control system to the used: control surface horn.

(b) Pilot effort effects. In the control P ILOT C ONTROL F ORCE L IMITS (S ECONDARY surface flight loading condition, the air C ONTROLS ) loads on movable surfaces and the cor- responding deflections need not exceed Control Limit pilot forces those that would result in flight from Miscellaneous: the application of any pilot force with- *Crank, wheel, or lever .. ((1 + R) / 3) × 50 lbs., but not in the ranges specified in paragraph (c) less than 50 lbs. nor more of this section. Two-thirds of the max- than 150 lbs. (R = radius).

(Applicable to any angle with- imum values specified for the aileron in 20 ° of plane of control).

and elevator may be used if control Twist ............................... 133 in.–lbs.

surface hinge moments are based on re- Push-pull ........................ To be chosen by applicant.

liable data. In applying this criterion, *Limited to flap, tab, stabilizer, spoiler, and landing gear op- eration controls.

the effects of servo mechanisms, tabs, Federal Aviation Administration, DOT § 25.415 moments H must be computed from the § 25.407 Trim tab effects.

formula: The effects of trim tabs on the con- H = K (1/2) r V c S o trol surface design conditions must be accounted for only where the surface Where— loads are limited by maximum pilot ef- K = hinge moment factor for ground gusts derived in paragraph (c) of this section; fort. In these cases, the tabs are con- r = density of air at sea level; o sidered to be deflected in the direction V = 65 knots relative to the aircraft; that would assist the pilot, and the de- S = area of the control surface aft of the flections are— hinge line; (a) For elevator trim tabs, those re- c = mean aerodynamic chord of the control quired to trim the airplane at any surface aft of the hinge line.

point within the positive portion of the (c) The hinge moment factor K for pertinent flight envelope in § 25.333(b), ground gusts must be taken from the except as limited by the stops; and following table: (b) For aileron and rudder trim tabs, those required to trim the airplane in Position of Surface K controls the critical unsymmetrical power and loading conditions, with appropriate (1) Aileron .................... 0.75 Control column locked or lashed in mid-po- allowance for rigging tolerances.

sition.

(2) Aileron .................... * ± 0.50 Ailerons at full throw.

§ 25.409 Tabs.

(3) Elevator .................. * ± 0.75 Elevator full down.

(4) Elevator .................. * ± 0.75 Elevator full up.

(a) Trim tabs. Trim tabs must be de- (5) Rudder .................... 0.75 Rudder in neutral.

signed to withstand loads arising from (6) Rudder .................... 0.75 Rudder at full throw.

all likely combinations of tab setting, * A positive value of K indicates a moment tending to de- primary control position, and airplane press the surface, while a negative value of K indicates a mo- ment tending to raise the surface.

speed (obtainable without exceeding the flight load conditions prescribed (d) The computed hinge moment of for the airplane as a whole), when the paragraph (b) of this section must be effect of the tab is opposed by pilot ef- used to determine the limit loads due fort forces up to those specified in to ground gust conditions for the con- § 25.397(b).

trol surface. A 1.25 factor on the com- (b) Balancing tabs. Balancing tabs puted hinge moments must be used in must be designed for deflections con- calculating limit control system loads.

sistent with the primary control sur- (e) Where control system flexibility face loading conditions.

is such that the rate of load applica- (c) Servo tabs. Servo tabs must be de- tion in the ground gust conditions signed for deflections consistent with might produce transient stresses appre- the primary control surface loading ciably higher than those corresponding conditions obtainable within the pilot to static loads, in the absence of a ra- maneuvering effort, considering pos- tional analysis substantiating a dif- sible opposition from the trim tabs.

ferent dynamic factor, an additional factor of 1.6 must be applied to the con- § 25.415 Ground gust conditions.

trol system loads of paragraph (d) of (a) The flight control systems and this section to obtain limit loads. If a surfaces must be designed for the limit rational analysis is used, the addi- loads generated when the airplane is tional factor must not be less than 1.2.

subjected to a horizontal 65-knot (f) For the condition of the control ground gust from any direction while locks engaged, the control surfaces, the taxiing and while parked. For airplanes control system locks, and the parts of equipped with control system gust any control systems between the sur- locks, the taxiing condition must be faces and the locks must be designed to evaluated with the controls locked and the resultant limit loads. Where con- unlocked, and the parked condition trol locks are not provided, then the must be evaluated with the controls control surfaces, the control system locked. stops nearest the surfaces, and the (b) The control system and surface parts of any control systems between loads due to ground gust may be as- the surfaces and the stops must be de- sumed to be static loads, and the hinge signed to the resultant limit loads. If 14 CFR Ch. I (1–1–25 Edition) § 25.427 the control system design is such as to § 25.341(a) acting in any orientation at allow any part of the control system to right angles to the flight path.

impact with the stops due to flexi- (d) Unsymmetrical loading on the bility, then the resultant impact loads empennage arising from buffet condi- must be taken into account in deriving tions of § 25.305(e) must be taken into the limit loads due to ground gust.

account.

(g) For the condition of taxiing with [Doc. No. 27902, 61 FR 5222, Feb. 9, 1996] the control locks disengaged, or where control locks are not provided, the fol- § 25.445 Auxiliary aerodynamic sur- lowing apply: faces.

(1) The control surfaces, the control (a) When significant, the aero- system stops nearest the surfaces, and dynamic influence between auxiliary the parts of any control systems be- aerodynamic surfaces, such as out- tween the surfaces and the stops must board fins and winglets, and their sup- be designed to the resultant limit porting aerodynamic surfaces, must be loads.

taken into account for all loading con- (2) The parts of the control systems ditions including pitch, roll, and yaw between the stops nearest the surfaces maneuvers, and gusts as specified in and the flight deck controls must be § 25.341(a) acting at any orientation at designed to the resultant limit loads, right angles to the flight path.

except that the parts of the control (b) To provide for unsymmetrical system where loads are eventually re- loading when outboard fins extend acted by the pilot need not exceed: above and below the horizontal surface, (i) The loads corresponding to the the critical vertical surface loading maximum pilot loads in § 25.397(c) for (load per unit area) determined under each pilot alone; or § 25.391 must also be applied as follows: (ii) 0.75 times these maximum loads (1) 100 percent to the area of the for each pilot when the pilot forces are vertical surfaces above (or below) the applied in the same direction.

horizontal surface.

[Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] (2) 80 percent to the area below (or above) the horizontal surface.

§ 25.427 Unsymmetrical loads.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) In designing the airplane for lat- amended by Amdt. 25–86, 61 FR 5222, Feb. 9, eral gust, yaw maneuver and roll ma- 1996] neuver conditions, account must be taken of unsymmetrical loads on the § 25.457 Wing flaps.

empennage arising from effects such as Wing flaps, their operating mecha- slipstream and aerodynamic inter- nisms, and their supporting structures ference with the wing, vertical fin and must be designed for critical loads oc- other aerodynamic surfaces.

curring in the conditions prescribed in (b) The horizontal tail must be as- § 25.345, accounting for the loads occur- sumed to be subjected to unsymmet- ring during transition from one flap po- rical loading conditions determined as sition and airspeed to another.

follows: (1) 100 percent of the maximum load- § 25.459 Special devices.

ing from the symmetrical maneuver The loading for special devices using conditions of § 25.331 and the vertical aerodynamic surfaces (such as slots, gust conditions of § 25.341(a) acting sep- slats and spoilers) must be determined arately on the surface on one side of from test data.

the plane of symmetry; and (2) 80 percent of these loadings acting [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as on the other side.

amended by Amdt. 25–72, 55 FR 29776, July 20, (c) For empennage arrangements 1990] where the horizontal tail surfaces have G ROUND L OADS dihedral angles greater than plus or minus 10 degrees, or are supported by § 25.471 General.

the vertical tail surfaces, the surfaces and the supporting structure must be (a) Loads and equilibrium. For limit designed for gust velocities specified in ground loads— Federal Aviation Administration, DOT § 25.479 (1) Limit ground loads obtained (3) With a limit descent velocity of 6 under this subpart are considered to be fps at the design take-off weight (the external forces applied to the airplane maximum weight for landing condi- structure; and tions at a reduced descent velocity).

(4) The prescribed descent velocities (2) In each specified ground load con- may be modified if it is shown that the dition, the external loads must be airplane has design features that make placed in equilibrium with the linear it impossible to develop these veloci- and angular inertia loads in a rational ties.

or conservative manner.

(b) Airplane lift, not exceeding air- (b) Critical centers of gravity. The crit- plane weight, may be assumed unless ical centers of gravity within the range the presence of systems or procedures for which certification is requested significantly affects the lift.

must be selected so that the maximum (c) The method of analysis of air- design loads are obtained in each land- plane and landing gear loads must take ing gear element. Fore and aft, into account at least the following ele- vertical, and lateral airplane centers of ments: gravity must be considered. Lateral (1) Landing gear dynamic character- displacements of the c.g. from the air- istics.

plane centerline which would result in (2) Spin-up and springback.

main gear loads not greater than 103 (3) Rigid body response.

percent of the critical design load for (4) Structural dynamic response of symmetrical loading conditions may be the airframe, if significant.

selected without considering the ef- (d) The landing gear dynamic charac- fects of these lateral c.g. displacements teristics must be validated by tests as on the loading of the main gear ele- defined in § 25.723(a).

ments, or on the airplane structure (e) The coefficient of friction between provided— the tires and the ground may be estab- (1) The lateral displacement of the lished by considering the effects of c.g. results from random passenger or skidding velocity and tire pressure.

cargo disposition within the fuselage or However, this coefficient of friction from random unsymmetrical fuel load- need not be more than 0.8.

ing or fuel usage; and [Amdt. 25–91, 62 FR 40705, July 29, 1997; Amdt.

(2) Appropriate loading instructions 25–91, 62 FR 45481, Aug. 27, 1997; Amdt. 25–103, for random disposable loads are in- 66 FR 27394, May 16, 2001] cluded under the provisions of § 25.1583(c)(2) to ensure that the lateral § 25.477 Landing gear arrangement.

displacement of the center of gravity is Sections 25.479 through 25.485 apply maintained within these limits.

to airplanes with conventional ar- (c) Landing gear dimension data. Fig- rangements of main and nose gears, or ure 1 of appendix A contains the basic main and tail gears, when normal oper- landing gear dimension data.

ating techniques are used.

[Amdt. 25–23, 35 FR 5673, Apr. 8, 1970, as § 25.479 Level landing conditions.

amended by Doc. No. FAA–2022–1355, Amdt.

25–148, 87 FR 75710, Dec. 9, 2022; 88 FR 2813, (a) In the level attitude, the airplane Jan. 18, 2023] is assumed to contact the ground at forward velocity components, ranging § 25.473 Landing load conditions and from V to 1.25 V parallel to the L1 L2 assumptions.

ground under the conditions prescribed (a) For the landing conditions speci- in § 25.473 with— fied in § 25.479 to § 25.485 the airplane is (1) V equal to V (TAS) at the ap- L1 S0 assumed to contact the ground— propriate landing weight and in stand- (1) In the attitudes defined in § 25.479 ard sea level conditions; and and § 25.481; (2) V equal to V (TAS) at the ap- L2 S0 (2) With a limit descent velocity of 10 propriate landing weight and altitudes fps at the design landing weight (the in a hot day temperature of 41 degrees maximum weight for landing condi- F. above standard.

tions at maximum descent velocity); (3) The effects of increased contact and speed must be investigated if approval 14 CFR Ch. I (1–1–25 Edition) § 25.481 of downwind landings exceeding 10 § 25.481 Tail-down landing conditions.

knots is requested.

(a) In the tail-down attitude, the air- (b) For the level landing attitude for plane is assumed to contact the ground airplanes with tail wheels, the condi- at forward velocity components, rang- tions specified in this section must be ing from V to V parallel to the L1 L2 investigated with the airplane hori- ground under the conditions prescribed zontal reference line horizontal in ac- in § 25.473 with— cordance with Figure 2 of Appendix A (1) V equal to V (TAS) at the ap- of this part. L 1 S 0 propriate landing weight and in stand- (c) For the level landing attitude for ard sea level conditions; and airplanes with nose wheels, shown in Figure 2 of Appendix A of this part, the (2) V equal to V (TAS) at the ap- L 2 S 0 conditions specified in this section propriate landing weight and altitudes must be investigated assuming the fol- in a hot day temperature of 41 degrees lowing attitudes: F. above standard.

(1) An attitude in which the main (3) The combination of vertical and wheels are assumed to contact the drag components considered to be act- ground with the nose wheel just clear ing at the main wheel axle centerline.

of the ground; and (b) For the tail-down landing condi- (2) If reasonably attainable at the tion for airplanes with tail wheels, the specified descent and forward veloci- main and tail wheels are assumed to ties, an attitude in which the nose and contact the ground simultaneously, in main wheels are assumed to contact accordance with figure 3 of appendix A.

the ground simultaneously.

Ground reaction conditions on the tail (d) In addition to the loading condi- wheel are assumed to act— tions prescribed in paragraph (a) of this (1) Vertically; and section, but with maximum vertical (2) Up and aft through the axle at 45 ground reactions calculated from para- degrees to the ground line.

graph (a), the following apply: (c) For the tail-down landing condi- (1) The landing gear and directly af- tion for airplanes with nose wheels, the fected attaching structure must be de- airplane is assumed to be at an atti- signed for the maximum vertical tude corresponding to either the stall- ground reaction combined with an aft ing angle or the maximum angle allow- acting drag component of not less than ing clearance with the ground by each 25% of this maximum vertical ground part of the airplane other than the reaction.

(2) The most severe combination of main wheels, in accordance with figure loads that are likely to arise during a 3 of appendix A, whichever is less.

lateral drift landing must be taken [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as into account. In absence of a more ra- amended by Amdt. 25–91, 62 FR 40705, July 29, tional analysis of this condition, the 1997; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] following must be investigated: (i) A vertical load equal to 75% of the § 25.483 One-gear landing conditions.

maximum ground reaction of § 25.473 For the one-gear landing conditions, must be considered in combination the airplane is assumed to be in the with a drag and side load of 40% and level attitude and to contact the 25% respectively of that vertical load.

ground on one main landing gear, in (ii) The shock absorber and tire de- accordance with Figure 4 of Appendix flections must be assumed to be 75% of A of this part. In this attitude— the deflection corresponding to the (a) The ground reactions must be the maximum ground reaction of same as those obtained on that side § 25.473(a)(2). This load case need not be under § 25.479(d)(1), and considered in combination with flat tires. (b) Each unbalanced external load (3) The combination of vertical and must be reacted by airplane inertia in drag components is considered to be a rational or conservative manner.

acting at the wheel axle centerline.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as [Amdt. 25–91, 62 FR 40705, July 29, 1997; Amdt. amended by Amdt. 25–91, 62 FR 40705, July 29, 25–91, 62 FR 45481, Aug. 27, 1997] 1997] Federal Aviation Administration, DOT § 25.493 the airplane structure and landing gear § 25.485 Side load conditions.

are assumed to be subjected to loads In addition to § 25.479(d)(2) the fol- not less than those obtained when the lowing conditions must be considered: aircraft is operating over the roughest (a) For the side load condition, the ground that may reasonably be ex- airplane is assumed to be in the level pected in normal operation.

attitude with only the main wheels contacting the ground, in accordance [Amdt. 25–91, 62 FR 40705, July 29, 1997] with figure 5 of appendix A.

(b) Side loads of 0.8 of the vertical re- § 25.493 Braked roll conditions.

action (on one side) acting inward and (a) An airplane with a tail wheel is 0.6 of the vertical reaction (on the assumed to be in the level attitude other side) acting outward must be with the load on the main wheels, in combined with one-half of the max- accordance with figure 6 of appendix A.

imum vertical ground reactions ob- The limit vertical load factor is 1.2 at tained in the level landing conditions.

the design landing weight and 1.0 at These loads are assumed to be applied the design ramp weight. A drag reac- at the ground contact point and to be tion equal to the vertical reaction mul- resisted by the inertia of the airplane.

tiplied by a coefficient of friction of The drag loads may be assumed to be 0.8, must be combined with the vertical zero.

ground reaction and applied at the ground contact point.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–91, 62 FR 40705, July 29, (b) For an airplane with a nose wheel 1997] the limit vertical load factor is 1.2 at the design landing weight, and 1.0 at § 25.487 Rebound landing condition.

the design ramp weight. A drag reac- (a) The landing gear and its sup- tion equal to the vertical reaction, porting structure must be investigated multiplied by a coefficient of friction for the loads occurring during rebound of 0.8, must be combined with the of the airplane from the landing sur- vertical reaction and applied at the face.

ground contact point of each wheel (b) With the landing gear fully ex- with brakes. The following two atti- tended and not in contact with the tudes, in accordance with figure 6 of ground, a load factor of 20.0 must act appendix A, must be considered: on the unsprung weights of the landing (1) The level attitude with the wheels gear. This load factor must act in the contacting the ground and the loads direction of motion of the unsprung distributed between the main and nose weights as they reach their limiting gear. Zero pitching acceleration is as- positions in extending with relation to sumed.

the sprung parts of the landing gear.

(2) The level attitude with only the main gear contacting the ground and § 25.489 Ground handling conditions.

with the pitching moment resisted by Unless otherwise prescribed, the angular acceleration.

landing gear and airplane structure (c) A drag reaction lower than that must be investigated for the conditions prescribed in this section may be used in §§ 25.491 through 25.509 with the air- if it is substantiated that an effective plane at the design ramp weight (the drag force of 0.8 times the vertical re- maximum weight for ground handling action cannot be attained under any conditions). No wing lift may be con- likely loading condition.

sidered. The shock absorbers and tires (d) An airplane equipped with a nose may be assumed to be in their static gear must be designed to withstand the position.

loads arising from the dynamic pitch- ing motion of the airplane due to sud- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5673, Apr. 8, den application of maximum braking 1970] force. The airplane is considered to be at design takeoff weight with the nose § 25.491 Taxi, takeoff and landing roll.

and main gears in contact with the Within the range of appropriate ground, and with a steady-state ground speeds and approved weights, vertical load factor of 1.0. The steady- 14 CFR Ch. I (1–1–25 Edition) § 25.495 state nose gear reaction must be com- § 25.497 Tail-wheel yawing.

bined with the maximum incremental (a) A vertical ground reaction equal nose gear vertical reaction caused by to the static load on the tail wheel, in the sudden application of maximum combination with a side component of braking force as described in para- equal magnitude, is assumed.

graphs (b) and (c) of this section.

(b) If there is a swivel, the tail wheel (e) In the absence of a more rational is assumed to be swiveled 90 ° to the air- analysis, the nose gear vertical reac- plane longitudinal axis with the result- ant load passing through the axle.

tion prescribed in paragraph (d) of this (c) If there is a lock, steering device, section must be calculated according or shimmy damper the tail wheel is to the following formula: also assumed to be in the trailing posi- tion with the side load acting at the ⎡ ⎤ W f AE μ T ground contact point.

V B = + N ⎢ ⎥ A B A B E + + + μ ⎣ ⎦ § 25.499 Nose-wheel yaw and steering.

Where: (a) A vertical load factor of 1.0 at the V = Nose gear vertical reaction.

N airplane center of gravity, and a side W = Design takeoff weight.

T component at the nose wheel ground A = Horizontal distance between the c.g. of contact equal to 0.8 of the vertical the airplane and the nose wheel.

ground reaction at that point are as- B = Horizontal distance between the c.g. of sumed.

the airplane and the line joining the cen- (b) With the airplane assumed to be ters of the main wheels.

in static equilibrium with the loads re- E = Vertical height of the c.g. of the airplane sulting from the use of brakes on one above the ground in the 1.0 g static con- side of the main landing gear, the nose dition.

gear, its attaching structure, and the μ = Coefficient of friction of 0.80.

fuselage structure forward of the cen- f = Dynamic response factor; 2.0 is to be used ter of gravity must be designed for the unless a lower factor is substantiated. In following loads: the absence of other information, the dy- (1) A vertical load factor at the cen- namic response factor f may be defined ter of gravity of 1.0.

by the equation: (2) A forward acting load at the air- plane center of gravity of 0.8 times the ⎛ ⎞ − πξ vertical load on one main gear.

f = + 1 exp ⎜ ⎟ (3) Side and vertical loads at the ⎜ ⎟ 2 − 1 ξ ⎝ ⎠ ground contact point on the nose gear that are required for static equi- Where: librium.

x is the effective critical damping ratio of (4) A side load factor at the airplane the rigid body pitching mode about the center of gravity of zero.

main landing gear effective ground con- (c) If the loads prescribed in para- tact point.

graph (b) of this section result in a [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as nose gear side load higher than 0.8 amended by Amdt. 25–23, 35 FR 5673, Apr. 8, times the vertical nose gear load, the 1970; Amdt. 25–97, 63 FR 29072, May 27, 1998] design nose gear side load may be lim- ited to 0.8 times the vertical load, with § 25.495 Turning.

unbalanced yawing moments assumed In the static position, in accordance to be resisted by airplane inertia with figure 7 of appendix A, the air- forces.

plane is assumed to execute a steady (d) For other than the nose gear, its attaching structure, and the forward turn by nose gear steering, or by appli- fuselage structure, the loading condi- cation of sufficient differential power, tions are those prescribed in paragraph so that the limit load factors applied at (b) of this section, except that— the center of gravity are 1.0 vertically (1) A lower drag reaction may be used and 0.5 laterally. The side ground reac- if an effective drag force of 0.8 times tion of each wheel must be 0.5 of the the vertical reaction cannot be reached vertical reaction.

under any likely loading condition; and Federal Aviation Administration, DOT § 25.509 (2) The forward acting load at the § 25.509 Towing loads.

center of gravity need not exceed the (a) The towing loads specified in maximum drag reaction on one main paragraph (d) of this section must be gear, determined in accordance with considered separately. These loads § 25.493(b).

must be applied at the towing fittings (e) With the airplane at design ramp and must act parallel to the ground. In weight, and the nose gear in any steer- addition— able position, the combined application (1) A vertical load factor equal to 1.0 of full normal steering torque and must be considered acting at the center vertical force equal to 1.33 times the of gravity; maximum static reaction on the nose (2) The shock struts and tires must be in their static positions; and gear must be considered in designing (3) With W as the design ramp the nose gear, its attaching structure, T weight, the towing load, F is— TOW, and the forward fuselage structure.

(i) 0.3 W for W less than 30,000 T T [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as pounds; amended by Amdt. 25–23, 35 FR 5673, Apr. 8, (ii) ( 6W + 450,000)/70 for W between T T 1970; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; 30,000 and 100,000 pounds; and Amdt. 25–91, 62 FR 40705, July 29, 1997] (iii) 0.15 W for W over 100,000 T T pounds.

§ 25.503 Pivoting.

(b) For towing points not on the (a) The airplane is assumed to pivot landing gear but near the plane of sym- about one side of the main gear with metry of the airplane, the drag and the brakes on that side locked. The side tow load components specified for limit vertical load factor must be 1.0 the auxiliary gear apply. For towing and the coefficient of friction 0.8. points located outboard of the main gear, the drag and side tow load compo- (b) The airplane is assumed to be in nents specified for the main gear apply.

static equilibrium, with the loads being Where the specified angle of swivel applied at the ground contact points, cannot be reached, the maximum ob- in accordance with figure 8 of appendix tainable angle must be used.

A.

(c) The towing loads specified in paragraph (d) of this section must be § 25.507 Reversed braking.

reacted as follows: (a) The airplane must be in a three (1) The side component of the towing point static ground attitude. Hori- load at the main gear must be reacted zontal reactions parallel to the ground by a side force at the static ground line and directed forward must be applied of the wheel to which the load is ap- at the ground contact point of each plied.

wheel with brakes. The limit loads (2) The towing loads at the auxiliary must be equal to 0.55 times the vertical gear and the drag components of the load at each wheel or to the load devel- towing loads at the main gear must be oped by 1.2 times the nominal max- reacted as follows: imum static brake torque, whichever is (i) A reaction with a maximum value less. equal to the vertical reaction must be applied at the axle of the wheel to (b) For airplanes with nose wheels, which the load is applied. Enough air- the pitching moment must be balanced plane inertia to achieve equilibrium by rotational inertia.

must be applied.

(c) For airplanes with tail wheels, the (ii) The loads must be reacted by air- resultant of the ground reactions must plane inertia.

pass through the center of gravity of (d) The prescribed towing loads are as the airplane.

follows: 14 CFR Ch. I (1–1–25 Edition) § 25.511 Load Tow point Position Magnitude No. Direction Main gear ............................... ................................................ 0.75 F per main 1 Forward, parallel to drag axis.

TOW gear unit. 2 Forward, at 30 ° to drag axis.

3 Aft, parallel to drag axis.

4 Aft, at 30 ° to drag axis.

Auxiliary gear ......................... Swiveled forward .................... 1.0 F ................... 5 Forward.

TOW 6 Aft.

Swiveled aft ............................ ......do ....................... 7 Forward.

8 Aft.

Swiveled 45 ° from forward ..... 0.5 F ................... 9 Forward, in plane of wheel.

TOW 10 Aft, in plane of wheel.

Swiveled 45 ° from aft ............. ......do ....................... 11 Forward, in plane of wheel.

12 Aft, in plane of wheel.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5673, Apr. 8, 1970] (3) Any unequal tire inflation pres- § 25.511 Ground load: unsymmetrical loads on multiple-wheel units. sure, assuming the maximum variation to be ± 5 percent of the nominal tire in- (a) General. Multiple-wheel landing flation pressure.

gear units are assumed to be subjected (4) A runway crown of zero and a run- to the limit ground loads prescribed in way crown having a convex upward this subpart under paragraphs (b) shape that may be approximated by a through (f) of this section. In addi- slope of 1 ⁄2 percent with the hori- tion— zontal. Runway crown effects must be (1) A tandem strut gear arrangement considered with the nose gear unit on is a multiple-wheel unit; and either slope of the crown.

(2) In determining the total load on a (5) The airplane attitude.

gear unit with respect to the provisions (6) Any structural deflections.

of paragraphs (b) through (f) of this (c) Deflated tires. The effect of de- section, the transverse shift in the load flated tires on the structure must be centroid, due to unsymmetrical load considered with respect to the loading distribution on the wheels, may be ne- conditions specified in paragraphs (d) glected.

through (f) of this section, taking into (b) Distribution of limit loads to wheels; account the physical arrangement of tires inflated. The distribution of the the gear components. In addition— limit loads among the wheels of the (1) The deflation of any one tire for landing gear must be established for each multiple wheel landing gear unit, each landing, taxiing, and ground han- and the deflation of any two critical dling condition, taking into account tires for each landing gear unit using the effects of the following factors: four or more wheels per unit, must be (1) The number of wheels and their considered; and physical arrangements. For truck type (2) The ground reactions must be ap- landing gear units, the effects of any plied to the wheels with inflated tires seesaw motion of the truck during the except that, for multiple-wheel gear landing impact must be considered in units with more than one shock strut, determining the maximum design loads a rational distribution of the ground for the fore and aft wheel pairs.

reactions between the deflated and in- (2) Any differentials in tire diameters flated tires, accounting for the dif- resulting from a combination of manu- ferences in shock strut extensions re- facturing tolerances, tire growth, and sulting from a deflated tire, may be tire wear. A maximum tire-diameter used.

differential equal to ⁄3 of the most un- (d) Landing conditions. For one and favorable combination of diameter for two deflated tires, the applied load variations that is obtained when tak- ing into account manufacturing toler- to each gear unit is assumed to be 60 ances, tire growth, and tire wear, may percent and 50 percent, respectively, of be assumed. the limit load applied to each gear for Federal Aviation Administration, DOT § 25.523 each of the prescribed landing condi- (2) For jacking by other airplane tions. However, for the drift landing structure at maximum approved jack- condition of § 25.485, 100 percent of the ing weight: vertical load must be applied. (i) The airplane structure must be de- signed for a vertical load of 1.33 times (e) Taxiing and ground handling condi- the vertical reaction at each jacking tions. For one and for two deflated point acting singly and in combination tires— with a horizontal load of 0.33 times the (1) The applied side or drag load fac- vertical static reaction applied in any tor, or both factors, at the center of direction.

gravity must be the most critical value (ii) The jacking pads and local struc- up to 50 percent and 40 percent, respec- ture must be designed for a vertical tively, of the limit side or drag load load of 2.0 times the vertical static re- factors, or both factors, corresponding action at each jacking point, acting to the most severe condition resulting singly and in combination with a hori- from consideration of the prescribed zontal load of 0.33 times the vertical taxiing and ground handling condi- static reaction applied in any direc- tions; tion.

(2) For the braked roll conditions of (c) Tie-down. If tie-down points are § 25.493 (a) and (b)(2), the drag loads on provided, the main tie-down points and each inflated tire may not be less than local structure must withstand the those at each tire for the symmetrical limit loads resulting from a 65-knot load distribution with no deflated tires; horizontal wind from any direction.

(3) The vertical load factor at the center of gravity must be 60 percent [Doc. No. 26129, 59 FR 22102, Apr. 28, 1994] and 50 percent, respectively, of the fac- tor with no deflated tires, except that W ATER L OADS it may not be less than 1g; and § 25.521 General.

(4) Pivoting need not be considered.

(f) Towing conditions. For one and for (a) Seaplanes must be designed for two deflated tires, the towing load, the water loads developed during take- F must be 60 percent and 50 percent, off and landing, with the seaplane in TOW, respectively, of the load prescribed. any attitude likely to occur in normal operation, and at the appropriate for- § 25.519 Jacking and tie-down provi- ward and sinking velocities under the sions.

most severe sea conditions likely to be encountered.

(a) General. The airplane must be de- signed to withstand the limit load con- (b) Unless a more rational analysis of ditions resulting from the static the water loads is made, or the stand- ground load conditions of paragraph (b) ards in ANC–3 are used, §§ 25.523 of this section and, if applicable, para- through 25.537 apply.

graph (c) of this section at the most (c) The requirements of this section critical combinations of airplane and §§ 25.523 through 25.537 apply also to weight and center of gravity. The max- amphibians.

imum allowable load at each jack pad § 25.523 Design weights and center of must be specified.

gravity positions.

(b) Jacking. The airplane must have provisions for jacking and must with- (a) Design weights. The water load re- stand the following limit loads when quirements must be met at each oper- the airplane is supported on jacks— ating weight up to the design landing weight except that, for the takeoff con- (1) For jacking by the landing gear at dition prescribed in § 25.531, the design the maximum ramp weight of the air- water takeoff weight (the maximum plane, the airplane structure must be weight for water taxi and takeoff run) designed for a vertical load of 1.33 must be used.

times the vertical static reaction at each jacking point acting singly and in (b) Center of gravity positions. The combination with a horizontal load of critical centers of gravity within the 0.33 times the vertical static reaction limits for which certification is re- applied in any direction. quested must be considered to reach 14 CFR Ch. I (1–1–25 Edition) § 25.525 maximum design loads for each part of (2) C = empirical seaplane operations the seaplane structure. factor equal to 0.012 (except that this factor may not be less than that nec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as essary to obtain the minimum value of amended by Amdt. 25–23, 35 FR 5673, Apr. 8, step load factor of 2.33).

1970] (3) V = seaplane stalling speed in S 0 knots with flaps extended in the appro- § 25.525 Application of loads.

priate landing position and with no (a) Unless otherwise prescribed, the slipstream effect.

seaplane as a whole is assumed to be (4) b = angle of dead rise at the longi- subjected to the loads corresponding to tudinal station at which the load fac- the load factors specified in § 25.527.

tor is being determined in accordance (b) In applying the loads resulting with figure 1 of appendix B.

from the load factors prescribed in (5) W= seaplane design landing § 25.527, the loads may be distributed weight in pounds.

over the hull or main float bottom (in (6) K = empirical hull station weigh- order to avoid excessive local shear ing factor, in accordance with figure 2 loads and bending moments at the lo- of appendix B.

cation of water load application) using (7) r = ratio of distance, measured x pressures not less than those pre- parallel to hull reference axis, from the scribed in § 25.533(c).

center of gravity of the seaplane to the (c) For twin float seaplanes, each hull longitudinal station at which the float must be treated as an equivalent load factor is being computed to the ra- hull on a fictitious seaplane with a dius of gyration in pitch of the sea- weight equal to one-half the weight of plane, the hull reference axis being a the twin float seaplane.

straight line, in the plane of sym- (d) Except in the takeoff condition of metry, tangential to the keel at the § 25.531, the aerodynamic lift on the main step.

seaplane during the impact is assumed (c) For a twin float seaplane, because to be ⁄3 of the weight of the seaplane.

of the effect of flexibility of the attach- ment of the floats to the seaplane, the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Doc. No. FAA–2022–1355, Amdt. factor K may be reduced at the bow 25–148, 87 FR 75710, Dec. 9, 2022; 88 FR 2813, and stern to 0.8 of the value shown in Jan. 18, 2023] figure 2 of appendix B. This reduction applies only to the design of the carry- § 25.527 Hull and main float load fac- through and seaplane structure.

tors.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) Water reaction load factors n W amended by Amdt. 25–23, 35 FR 5673, Apr. 8, must be computed in the following 1970] manner: (1) For the step landing case § 25.529 Hull and main float landing conditions.

C V S 0 (a) Symmetrical step, bow, and stern n = w 2 1 landing. For symmetrical step, bow, ⎛ ⎞ 3 3 W Tan β and stern landings, the limit water re- ⎝ ⎠ action load factors are those computed under § 25.527. In addition— (2) For the bow and stern landing (1) For symmetrical step landings, cases the resultant water load must be ap- C V plied at the keel, through the center of K 1 S 0 1 n = × gravity, and must be directed per- w 1 2 3 3 pendicularly to the keel line; ⎛ ⎞ W Tan β r + 1 (2) For symmetrical bow landings,

( )

x ⎝ ⎠ the resultant water load must be ap- (b) The following values are used: plied at the keel, one-fifth of the longi- (1) n = water reaction load factor tudinal distance from the bow to the W (that is, the water reaction divided by step, and must be directed perpendicu- seaplane weight). larly to the keel line; and Federal Aviation Administration, DOT § 25.533 tended in the appropriate takeoff posi- (3) For symmetrical stern landings, tion; the resultant water load must be ap- b = angle of dead rise at the main step (de- plied at the keel, at a point 85 percent grees); and of the longitudinal distance from the W = design water takeoff weight in pounds.

step to the stern post, and must be di- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as rected perpendicularly to the keel line.

amended by Amdt. 25–23, 35 FR 5673, Apr. 8, (b) Unsymmetrical landing for hull and 1970] single float seaplanes. Unsymmetrical step, bow, and stern landing conditions § 25.533 Hull and main float bottom must be investigated. In addition— pressures.

(1) The loading for each condition (a) General. The hull and main float consists of an upward component and a structure, including frames and bulk- side component equal, respectively, to heads, stringers, and bottom plating, 0.75 and 0.25 tan b times the resultant must be designed under this section.

load in the corresponding symmetrical (b) Local pressures. For the design of landing condition; and the bottom plating and stringers and (2) The point of application and di- their attachments to the supporting rection of the upward component of the structure, the following pressure dis- load is the same as that in the sym- tributions must be applied: metrical condition, and the point of ap- (1) For an unflared bottom, the pres- plication of the side component is at sure at the chine is 0.75 times the pres- the same longitudinal station as the sure at the keel, and the pressures be- upward component but is directed in- tween the keel and chine vary linearly, ward perpendicularly to the plane of in accordance with figure 3 of appendix symmetry at a point midway between B. The pressure at the keel (psi) is the keel and chine lines.

computed as follows: (c) Unsymmetrical landing; twin float seaplanes. The unsymmetrical loading K V S 1 consists of an upward load at the step P C = × k 2 of each float of 0.75 and a side load of tan β k 0.25 tan b at one float times the step where— landing load reached under § 25.527. The P = pressure (p.s.i.) at the keel; k side load is directed inboard, per- C = 0.00213; pendicularly to the plane of symmetry K = hull station weighing factor, in accord- midway between the keel and chine ance with figure 2 of appendix B; lines of the float, at the same longitu- V S 1 = seaplane stalling speed (Knots) at the dinal station as the upward load. design water takeoff weight with flaps extended in the appropriate takeoff posi- § 25.531 Hull and main float takeoff tion; and b K = angle of dead rise at keel, in accordance condition.

with figure 1 of appendix B.

For the wing and its attachment to (2) For a flared bottom, the pressure the hull or main float— at the beginning of the flare is the (a) The aerodynamic wing lift is as- same as that for an unflared bottom, sumed to be zero; and and the pressure between the chine and (b) A downward inertia load, cor- the beginning of the flare varies lin- responding to a load factor computed early, in accordance with figure 3 of ap- from the following formula, must be pendix B. The pressure distribution is applied: the same as that prescribed in para- C V graph (b)(1) of this section for an TO S 1 n = unflared bottom except that the pres- 2 1 ⎛ ⎞ sure at the chine is computed as fol- 3 3 W tan β lows: ⎝ ⎠ where— K V S 1 n = inertia load factor; P C = × ch 3 C = empirical seaplane operations factor TO tan β equal to 0.004; where— V = seaplane stalling speed (knots) at the S 1 design takeoff weight with the flaps ex- P = pressure (p.s.i.) at the chine; ch 14 CFR Ch. I (1–1–25 Edition) § 25.535 C = 0.0016; § 25.535 Auxiliary float loads.

K 2 = hull station weighing factor, in accord- (a) General. Auxiliary floats and their ance with figure 2 of appendix B; attachments and supporting structures V = seaplane stalling speed at the design S 1 must be designed for the conditions water takeoff weight with flaps extended prescribed in this section. In the cases in the appropriate takeoff position; and specified in paragraphs (b) through (e) b = angle of dead rise at appropriate station.

of this section, the prescribed water The area over which these pressures loads may be distributed over the float are applied must simulate pressures oc- bottom to avoid excessive local loads, curring during high localized impacts using bottom pressures not less than on the hull or float, but need not ex- those prescribed in paragraph (g) of tend over an area that would induce this section.

critical stresses in the frames or in the (b) Step loading. The resultant water overall structure.

load must be applied in the plane of (c) Distributed pressures. For the de- symmetry of the float at a point three- sign of the frames, keel, and chine fourths of the distance from the bow to structure, the following pressure dis- the step and must be perpendicular to tributions apply: the keel. The resultant limit load is computed as follows, except that the (1) Symmetrical pressures are com- value of L need not exceed three times puted as follows: the weight of the displaced water when K V the float is completely submerged: S 0 P C = × tan β C V W 5 2 where— S L = P = pressure (p.s.i.); 2 2 3 C = 0.078 C (with C computed under 4 1 1 r + 1 tan β

s y ( )

§ 25.527); K = hull station weighing factor, deter- where— mined in accordance with figure 2 of ap- L = limit load (lbs.); pendix B; C = 0.0053; V S 0 = seaplane stalling speed (Knots) with V = seaplane stalling speed (knots) with S 0 landing flaps extended in the appropriate landing flaps extended in the appropriate position and with no slipstream effect; position and with no slipstream effect; and W = seaplane design landing weight in V = seaplane stalling speed with landing pounds; S 0 flaps extended in the appropriate posi- b = angle of dead rise at a station ⁄4 of the S tion and with no slipstream effect; and b distance from the bow to the step, but = angle of dead rise at appropriate sta- need not be less than 15 degrees; and tion. r = ratio of the lateral distance between the y center of gravity and the plane of sym- (2) The unsymmetrical pressure dis- metry of the float to the radius of gyra- tribution consists of the pressures pre- tion in roll.

scribed in paragraph (c)(1) of this sec- (c) Bow loading. The resultant limit tion on one side of the hull or main load must be applied in the plane of float centerline and one-half of that symmetry of the float at a point one- pressure on the other side of the hull or fourth of the distance from the bow to main float centerline, in accordance the step and must be perpendicular to with figure 3 of appendix B.

the tangent to the keel line at that These pressures are uniform and must point. The magnitude of the resultant be applied simultaneously over the en- load is that specified in paragraph (b) tire hull or main float bottom. The of this section.

loads obtained must be carried into the (d) Unsymmetrical step loading. The re- sidewall structure of the hull proper, sultant water load consists of a compo- but need not be transmitted in a fore nent equal to 0.75 times the load speci- and aft direction as shear and bending fied in paragraph (a) of this section and loads.

a side component equal to 0.25 tan b times the load specified in paragraph [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (b) of this section. The side load must amended by Amdt. 25–23, 35 FR 5673, Apr. 8, 1970] be applied perpendicularly to the plane Federal Aviation Administration, DOT § 25.561 of symmetry of the float at a point E MERGENCY L ANDING C ONDITIONS midway between the keel and the § 25.561 General.

chine.

(e) Unsymmetrical bow loading. The re- (a) The airplane, although it may be sultant water load consists of a compo- damaged in emergency landing condi- nent equal to 0.75 times the load speci- tions on land or water, must be de- fied in paragraph (b) of this section and signed as prescribed in this section to a side component equal to 0.25 tan b protect each occupant under those con- times the load specified in paragraph ditions.

(c) of this section. The side load must (b) The structure must be designed to be applied perpendicularly to the plane give each occupant every reasonable of symmetry at a point midway be- chance of escaping serious injury in a tween the keel and the chine. minor crash landing when— (1) Proper use is made of seats, belts, (f) Immersed float condition. The re- and all other safety design provisions; sultant load must be applied at the (2) The wheels are retracted (where centroid of the cross section of the applicable); and float at a point one-third of the dis- (3) The occupant experiences the fol- tance from the bow to the step. The lowing ultimate inertia forces acting limit load components are as follows: separately relative to the surrounding structure: vertical = V ρ g (i) Upward, 3.0g (ii) Forward, 9.0g ⎛ ⎞ ρ (iii) Sideward, 3.0g on the airframe; aft = C V KV x S ⎝ ⎠ and 4.0g on the seats and their attach- ments.

⎛ ⎞ ρ (iv) Downward, 6.0g side = C V KV y S 2 (v) Rearward, 1.5g ⎝ ⎠ (c) For equipment, cargo in the pas- where— senger compartments and any other r = mass density of water (slugs/ft. ); large masses, the following apply: V = volume of float (ft. ); (1) Except as provided in paragraph C = coefficient of drag force, equal to 0.133; x (c)(2) of this section, these items must C y = coefficient of side force, equal to 0.106; be positioned so that if they break K = 0.8, except that lower values may be used loose they will be unlikely to: if it is shown that the floats are incapa- (i) Cause direct injury to occupants; ble of submerging at a speed of 0.8 V in S 0 (ii) Penetrate fuel tanks or lines or normal operations; cause fire or explosion hazard by dam- V = seaplane stalling speed (knots) with S 0 age to adjacent systems; or landing flaps extended in the appropriate (iii) Nullify any of the escape facili- position and with no slipstream effect; and ties provided for use after an emer- g = acceleration due to gravity (ft./sec. ). gency landing.

(2) When such positioning is not prac- (g) Float bottom pressures. The float tical (e.g. fuselage mounted engines or bottom pressures must be established auxiliary power units) each such item under § 25.533, except that the value of of mass shall be restrained under all K in the formulae may be taken as 1.0.

loads up to those specified in paragraph The angle of dead rise to be used in de- (b)(3) of this section. The local attach- termining the float bottom pressures is ments for these items should be de- set forth in paragraph (b) of this sec- signed to withstand 1.33 times the spec- tion.

ified loads if these items are subject to severe wear and tear through frequent [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5673, Apr. 8, removal (e.g. quick change interior 1970; Amdt. 25–148, 87 FR 75710, Dec. 9, 2022; 88 items).

FR 2813, Jan. 18, 2023] (d) Seats and items of mass (and their supporting structure) must not § 25.537 Seawing loads.

deform under any loads up to those Seawing design loads must be based specified in paragraph (b)(3) of this sec- on applicable test data. tion in any manner that would impede 14 CFR Ch. I (1–1–25 Edition) § 25.562 subsequent rapid evacuation of occu- grees either right or left, whichever pants. would cause the greatest likelihood of the upper torso restraint system [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (where installed) moving off the occu- amended by Amdt. 25–23, 35 FR 5673, Apr. 8, pant’s shoulder, and with the wings 1970; Amdt. 25–64, 53 FR 17646, May 17, 1988; Amdt. 25–91, 62 FR 40706, July 29, 1997] level. Peak floor deceleration must occur in not more than 0.09 seconds § 25.562 Emergency landing dynamic after impact and must reach a min- conditions.

imum of 16g. Where floor rails or floor (a) The seat and restraint system in fittings are used to attach the seating the airplane must be designed as pre- devices to the test fixture, the rails or scribed in this section to protect each fittings must be misaligned with re- occupant during an emergency landing spect to the adjacent set of rails or fit- condition when— tings by at least 10 degrees vertically (1) Proper use is made of seats, safety ( i.e. , out of Parallel) with one rolled 10 belts, and shoulder harnesses provided degrees.

for in the design; and (c) The following performance meas- (2) The occupant is exposed to loads ures must not be exceeded during the resulting from the conditions pre- dynamic tests conducted in accordance scribed in this section.

with paragraph (b) of this section: (b) Each seat type design approved (1) Where upper torso straps are used for crew or passenger occupancy during for crewmembers, tension loads in indi- takeoff and landing must successfully vidual straps must not exceed 1,750 complete dynamic tests or be dem- pounds. If dual straps are used for re- onstrated by rational analysis based on straining the upper torso, the total dynamic tests of a similar type seat, in strap tension loads must not exceed accordance with each of the following 2,000 pounds.

emergency landing conditions. The (2) The maximum compressive load tests must be conducted with an occu- measured between the pelvis and the pant simulated by a 170-pound lumbar column of the anthropomorphic anthropomorphic test dummy, as de- dummy must not exceed 1,500 pounds.

fined by 49 CFR Part 572, Subpart B, or (3) The upper torso restraint straps its equivalent, sitting in the normal (where installed) must remain on the upright position.

occupant’s shoulder during the impact.

(1) A change in downward vertical ve- (4) The lap safety belt must remain locity ( D v) of not less than 35 feet per on the occupant’s pelvis during the im- second, with the airplane’s longitu- pact.

dinal axis canted downward 30 degrees (5) Each occupant must be protected with respect to the horizontal plane and with the wings level. Peak floor de- from serious head injury under the con- celeration must occur in not more than ditions prescribed in paragraph (b) of 0.08 seconds after impact and must this section. Where head contact with reach a minimum of 14g. seats or other structure can occur, pro- (2) A change in forward longitudinal tection must be provided so that the velocity ( D v) of not less than 44 feet head impact does not exceed a Head In- per second, with the airplane’s longitu- jury Criterion (HIC) of 1,000 units. The dinal axis horizontal and yawed 10 de- level of HIC is defined by the equation: 2 5 .

⎧ ⎫ ⎡ ⎤ t ⎪ ⎪ 1 2 HIC t t a t dt = − ( )

( ) ⎢ ⎥

⎨ ⎬ 2 1

∫

t t t − 1

( )

⎢ ⎥ 2 1 ⎪ ⎪ ⎣ ⎦ ⎩ ⎭ max Where: a(t) is the total acceleration vs. time curve for the head strike, and where t is the initial integration time, t 2 is the final integration time, and Federal Aviation Administration, DOT § 25.571 (t) is in seconds, and (a) is in units of gravity catastrophic failure of the airplane; (g).

and (iii) An analysis, supported by test (6) Where leg injuries may result evidence, of the principal structural from contact with seats or other struc- elements and detail design points iden- ture, protection must be provided to tified in paragraph (a)(1)(ii) of this sec- prevent axially compressive loads ex- tion.

ceeding 2,250 pounds in each femur.

(2) The service history of airplanes of (7) The seat must remain attached at similar structural design, taking due all points of attachment, although the account of differences in operating con- structure may have yielded.

ditions and procedures, may be used in (8) Seats must not yield under the the evaluations required by this sec- tests specified in paragraphs (b)(1) and tion.

(b)(2) of this section to the extent they (3) Based on the evaluations required would impede rapid evacuation of the by this section, inspections or other airplane occupants.

procedures must be established, as nec- [Amdt. 25–64, 53 FR 17646, May 17, 1988] essary, to prevent catastrophic failure, and must be included in the Airworthi- § 25.563 Structural ditching provi- ness Limitations section of the In- sions.

structions for Continued Airworthiness Structural strength considerations of required by § 25.1529. The limit of valid- ditching provisions must be in accord- ity of the engineering data that sup- ance with § 25.801(e). ports the structural maintenance pro- gram (hereafter referred to as LOV), F ATIGUE E VALUATION stated as a number of total accumu- lated flight cycles or flight hours or § 25.571 Damage-tolerance and fatigue both, established by this section must evaluation of structure.

also be included in the Airworthiness (a) General. An evaluation of the Limitations section of the Instructions strength, detail design, and fabrication for Continued Airworthiness required must show that catastrophic failure by § 25.1529. Inspection thresholds for due to fatigue, corrosion, manufac- the following types of structure must turing defects, or accidental damage, be established based on crack growth will be avoided throughout the oper- analyses and/or tests, assuming the ational life of the airplane. This eval- structure contains an initial flaw of uation must be conducted in accord- the maximum probable size that could ance with the provisions of paragraphs exist as a result of manufacturing or (b) and (e) of this section, except as service-induced damage: specified in paragraph (c) of this sec- (i) Single load path structure, and tion, for each part of the structure that (ii) Multiple load path ‘‘fail-safe’’ could contribute to a catastrophic fail- structure and crack arrest ‘‘fail-safe’’ ure (such as wing, empennage, control structure, where it cannot be dem- surfaces and their systems, the fuse- onstrated that load path failure, par- lage, engine mounting, landing gear, tial failure, or crack arrest will be de- and their related primary attach- tected and repaired during normal ments). For turbojet powered air- maintenance, inspection, or operation planes, those parts that could con- of an airplane prior to failure of the re- tribute to a catastrophic failure must maining structure.

also be evaluated under paragraph (d) (b) Damage-tolerance evaluation. The of this section. In addition, the fol- evaluation must include a determina- lowing apply: tion of the probable locations and (1) Each evaluation required by this modes of damage due to fatigue, corro- section must include— sion, or accidental damage. Repeated (i) The typical loading spectra, tem- load and static analyses supported by peratures, and humidities expected in test evidence and (if available) service service; experience must also be incorporated (ii) The identification of principal in the evaluation. Special consider- structural elements and detail design ation for widespread fatigue damage points, the failure of which could cause must be included where the design is 14 CFR Ch. I (1–1–25 Edition) § 25.571 such that this type of damage could pressures during 1 g level flight) multi- occur. An LOV must be established plied by a factor of 1.15, omitting other that corresponds to the period of time, loads.

stated as a number of total accumu- (6) For landing gear and directly-af- lated flight cycles or flight hours or fected airframe structure, the limit both, during which it is demonstrated ground loading conditions specified in that widespread fatigue damage will §§ 25.473, 25.491, and 25.493.

not occur in the airplane structure.

If significant changes in structural This demonstration must be by full- stiffness or geometry, or both, follow scale fatigue test evidence. The type from a structural failure, or partial certificate may be issued prior to com- failure, the effect on damage tolerance pletion of full-scale fatigue testing, must be further investigated.

provided the Administrator has ap- (c) Fatigue (safe-life) evaluation. Com- proved a plan for completing the re- pliance with the damage-tolerance re- quired tests. In that case, the Air- quirements of paragraph (b) of this sec- worthiness Limitations section of the tion is not required if the applicant es- Instructions for Continued Airworthi- tablishes that their application for par- ness required by § 25.1529 must specify ticular structure is impractical. This that no airplane may be operated be- structure must be shown by analysis, yond a number of cycles equal to ⁄2 the supported by test evidence, to be able number of cycles accumulated on the to withstand the repeated loads of vari- fatigue test article, until such testing able magnitude expected during its is completed. The extent of damage for service life without detectable cracks.

residual strength evaluation at any Appropriate safe-life scatter factors time within the operational life of the must be applied.

airplane must be consistent with the (d) Sonic fatigue strength. It must be initial detectability and subsequent shown by analysis, supported by test growth under repeated loads. The resid- evidence, or by the service history of ual strength evaluation must show airplanes of similar structural design that the remaining structure is able to and sonic excitation environment, withstand loads (considered as static that— ultimate loads) corresponding to the (1) Sonic fatigue cracks are not prob- following conditions: able in any part of the flight structure (1) The limit symmetrical maneu- subject to sonic excitation; or vering conditions specified in § 25.337 at (2) Catastrophic failure caused by all speeds up to V and in § 25.345.

c sonic cracks is not probable assuming (2) The limit gust conditions speci- that the loads prescribed in paragraph fied in § 25.341 at the specified speeds up (b) of this section are applied to all to V and in § 25.345.

C areas affected by those cracks.

(3) The limit rolling conditions speci- (e) Damage-tolerance (discrete source) fied in § 25.349 and the limit unsymmet- evaluation. The airplane must be capa- rical conditions specified in §§ 25.367 ble of successfully completing a flight and 25.427 (a) through (c), at speeds up during which likely structural damage to V .

C occurs as a result of— (4) The limit yaw maneuvering condi- (1) Impact with a 4-pound bird when tions specified in § 25.351(a) at the spec- the velocity of the airplane relative to ified speeds up to V .

C the bird along the airplane’s flight (5) For pressurized cabins, the fol- path is equal to V at sea level or 0.85V c c lowing conditions: at 8,000 feet, whichever is more critical; (i) The normal operating differential (2) Uncontained fan blade impact; pressure combined with the expected (3) Uncontained engine failure; or external aerodynamic pressures applied (4) Uncontained high energy rotating simultaneously with the flight loading machinery failure.

conditions specified in paragraphs (b)(1) through (4) of this section, if they The damaged structure must be able to have a significant effect. withstand the static loads (considered (ii) The maximum value of normal as ultimate loads) which are reason- operating differential pressure (includ- ably expected to occur on the flight.

ing the expected external aerodynamic Dynamic effects on these static loads Federal Aviation Administration, DOT § 25.607 need not be considered. Corrective ac- which could adversely affect safety, tion to be taken by the pilot following must— the incident, such as limiting maneu- (a) Be established on the basis of ex- vers, avoiding turbulence, and reducing perience or tests; speed, must be considered. If signifi- (b) Conform to approved specifica- cant changes in structural stiffness or tions (such as industry or military geometry, or both, follow from a struc- specifications, or Technical Standard tural failure or partial failure, the ef- Orders) that ensure their having the fect on damage tolerance must be fur- strength and other properties assumed ther investigated.

in the design data; and [Amdt. 25–45, 43 FR 46242, Oct. 5, 1978, as (c) Take into account the effects of amended by Amdt. 25–54, 45 FR 60173, Sept.

environmental conditions, such as tem- 11, 1980; Amdt. 25–72, 55 FR 29776, July 20, perature and humidity, expected in 1990; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; service.

Amdt. 25–96, 63 FR 15714, Mar. 31, 1998; 63 FR 23338, Apr. 28, 1998; Amdt. 25–132, 75 FR 69781, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as Nov. 15, 2010; Amdt. No. 25–148, 87 FR 75710, amended by Amdt. 25–38, 41 FR 55466, Dec. 20, Dec. 9, 2022; 88 FR 2813, Jan. 18, 2023] 1976; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978] L IGHTNING PROTECTION § 25.605 Fabrication methods.

§ 25.581 Lightning protection.

(a) The methods of fabrication used must produce a consistently sound (a) The airplane must be protected structure. If a fabrication process (such against catastrophic effects from light- as gluing, spot welding, or heat treat- ning.

ing) requires close control to reach this (b) For metallic components, compli- ance with paragraph (a) of this section objective, the process must be per- may be shown by— formed under an approved process spec- ification.

(1) Bonding the components properly to the airframe; or (b) Each new aircraft fabrication (2) Designing the components so that method must be substantiated by a a strike will not endanger the airplane.

test program.

(c) For nonmetallic components, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as compliance with paragraph (a) of this amended by Amdt. 25–46, 43 FR 50595, Oct. 30, section may be shown by— 1978] (1) Designing the components to min- imize the effect of a strike; or § 25.607 Fasteners.

(2) Incorporating acceptable means of (a) Each removable bolt, screw, nut, diverting the resulting electrical cur- pin, or other removable fastener must rent so as not to endanger the airplane.

incorporate two separate locking de- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] vices if— (1) Its loss could preclude continued Subpart D—Design and flight and landing within the design Construction limitations of the airplane using nor- mal pilot skill and strength; or G ENERAL (2) Its loss could result in reduction in pitch, yaw, or roll control capability § 25.601 General.

or response below that required by The airplane may not have design Subpart B of this chapter.

features or details that experience has (b) The fasteners specified in para- shown to be hazardous or unreliable.

graph (a) of this section and their lock- The suitability of each questionable ing devices may not be adversely af- design detail and part must be estab- fected by the environmental conditions lished by tests.

associated with the particular installa- tion.

§ 25.603 Materials.

The suitability and durability of ma- terials used for parts, the failure of 14 CFR Ch. I (1–1–25 Edition) § 25.609 (c) No self-locking nut may be used rial strength with the following prob- on any bolt subject to rotation in oper- ability: ation unless a nonfriction locking de- (1) Where applied loads are eventu- vice is used in addition to the self-lock- ally distributed through a single mem- ing device. ber within an assembly, the failure of which would result in loss of structural [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] integrity of the component, 99 percent probability with 95 percent confidence.

§ 25.609 Protection of structure.

(2) For redundant structure, in which Each part of the structure must— the failure of individual elements (a) Be suitably protected against de- would result in applied loads being terioration or loss of strength in serv- safely distributed to other load car- ice due to any cause, including— rying members, 90 percent probability (1) Weathering; with 95 percent confidence.

(2) Corrosion; and (c) The effects of environmental con- (3) Abrasion; and ditions, such as temperature and mois- ture, on material design values used in (b) Have provisions for ventilation an essential component or structure and drainage where necessary for pro- must be considered where these effects tection.

are significant within the airplane op- § 25.611 Accessibility provisions.

erating envelope.

(d) [Reserved] (a)Means must be provided to allow (e) Greater material design values inspection (including inspection of may be used if a ‘‘premium selection’’ principal structural elements and con- of the material is made in which a trol systems), replacement of parts specimen of each individual item is normally requiring replacement, ad- tested before use to determine that the justment, and lubrication as necessary actual strength properties of that par- for continued airworthiness. The in- ticular item will equal or exceed those spection means for each item must be used in design.

practicable for the inspection interval (f) Other material design values may for the item. Nondestructive inspection be used if approved by the Adminis- aids may be used to inspect structural trator.

elements where it is impracticable to provide means for direct visual inspec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tion if it is shown that the inspection amended by Amdt. 25–46, 43 FR 50595, Oct. 30, is effective and the inspection proce- 1978; Amdt. 25–72, 55 FR 29776, July 20, 1990; Amdt. 25–112, 68 FR 46431, Aug. 5, 2003] dures are specified in the maintenance manual required by § 25.1529.

§ 25.619 Special factors.

(b) EWIS must meet the accessibility requirements of § 25.1719. The factor of safety prescribed in § 25.303 must be multiplied by the high- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970, as est pertinent special factor of safety amended by Amdt. 25–123, 72 FR 63404, Nov. 8, prescribed in §§ 25.621 through 25.625 for 2007] each part of the structure whose strength is— § 25.613 Material strength properties and material design values. (a) Uncertain; (b) Likely to deteriorate in service (a) Material strength properties must before normal replacement; or be based on enough tests of material (c) Subject to appreciable variability meeting approved specifications to es- because of uncertainties in manufac- tablish design values on a statistical turing processes or inspection methods.

basis.

(b) Material design values must be [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as chosen to minimize the probability of amended by Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] structural failures due to material var- iability. Except as provided in para- § 25.621 Casting factors.

graphs (e) and (f) of this section, com- pliance must be shown by selecting ma- (a) General. For castings used in terial design values which assure mate- structural applications, the factors, Federal Aviation Administration, DOT § 25.621 tests, and inspections specified in para- (A) Inspection of 100 percent of its graphs (b) through (d) of this section surface, using visual inspection and liq- must be applied in addition to those uid penetrant or equivalent inspection necessary to establish foundry quality methods; and control. The inspections must meet ap- (B) Inspection of structurally signifi- cant internal areas and areas where de- proved specifications. Paragraphs (c) fects are likely to occur, using radio- and (d) of this section apply to any graphic or equivalent inspection meth- structural castings, except castings ods.

that are pressure tested as parts of hy- (iii) One casting undergoes a static draulic or other fluid systems and do test and is shown to meet the strength not support structural loads.

and deformation requirements of (b) Bearing stresses and surfaces. The § 25.305(a) and (b).

casting factors specified in paragraphs (2) A casting factor of 1.25 or greater (c) and (d) of this section— may be used, provided that— (1) Need not exceed 1.25 with respect (i) Each casting receives: to bearing stresses regardless of the (A) Inspection of 100 percent of its method of inspection used; and surface, using visual inspection and liq- (2) Need not be used with respect to uid penetrant or equivalent inspection the bearing surfaces of a part whose methods; and bearing factor is larger than the appli- (B) Inspection of structurally signifi- cable casting factor.

cant internal areas and areas where de- (c) Critical castings. Each casting fects are likely to occur, using radio- whose failure could preclude continued graphic or equivalent inspection meth- safe flight and landing of the airplane ods.

or could result in serious injury to oc- (ii) Three castings undergo static cupants is a critical casting. Each crit- tests and are shown to meet: ical casting must have a factor associ- (A) The strength requirements of ated with it for showing compliance § 25.305(b) at an ultimate load cor- with strength and deformation require- responding to a casting factor of 1.25; ments of § 25.305, and must comply with and the following criteria associated with (B) The deformation requirements of that factor: § 25.305(a) at a load of 1.15 times the (1) A casting factor of 1.0 or greater limit load.

may be used, provided that— (3) A casting factor of 1.50 or greater (i) It is demonstrated, in the form of may be used, provided that— process qualification, proof of product, (i) Each casting receives: and process monitoring that, for each (A) Inspection of 100 percent of its casting design and part number, the surface, using visual inspection and liq- castings produced by each foundry and uid penetrant or equivalent inspection process combination have coefficients methods; and of variation of the material properties (B) Inspection of structurally signifi- that are equivalent to those of wrought cant internal areas and areas where de- alloy products of similar composition.

fects are likely to occur, using radio- Process monitoring must include test- graphic or equivalent inspection meth- ing of coupons cut from the prolonga- ods.

tions of each casting (or each set of (ii) One casting undergoes a static castings, if produced from a single pour test and is shown to meet: into a single mold in a runner system) (A) The strength requirements of and, on a sampling basis, coupons cut § 25.305(b) at an ultimate load cor- from critical areas of production cast- responding to a casting factor of 1.50; ings. The acceptance criteria for the and process monitoring inspections and (B) The deformation requirements of tests must be established and included § 25.305(a) at a load of 1.15 times the in the process specifications to ensure limit load.

the properties of the production cast- (d) Non-critical castings. For each ings are controlled to within levels casting other than critical castings, as used in design.

specified in paragraph (c) of this sec- (ii) Each casting receives: tion, the following apply: 14 CFR Ch. I (1–1–25 Edition) § 25.623 (1) A casting factor of 1.0 or greater § 25.623 Bearing factors.

may be used, provided that the require- (a) Except as provided in paragraph ments of (c)(1) of this section are met, (b) of this section, each part that has or all of the following conditions are clearance (free fit), and that is subject met: to pounding or vibration, must have a (i) Castings are manufactured to ap- bearing factor large enough to provide proved specifications that specify the for the effects of normal relative mo- minimum mechanical properties of the tion.

material in the casting and provides (b) No bearing factor need be used for for demonstration of these properties a part for which any larger special fac- by testing of coupons cut from the tor is prescribed.

castings on a sampling basis.

(ii) Each casting receives: § 25.625 Fitting factors.

(A) Inspection of 100 percent of its For each fitting (a part or terminal surface, using visual inspection and liq- used to join one structural member to uid penetrant or equivalent inspection another), the following apply: methods; and (a) For each fitting whose strength is (B) Inspection of structurally signifi- not proven by limit and ultimate load cant internal areas and areas where de- tests in which actual stress conditions fects are likely to occur, using radio- are simulated in the fitting and sur- graphic or equivalent inspection meth- rounding structures, a fitting factor of ods. at least 1.15 must be applied to each part of— (iii) Three sample castings undergo (1) The fitting; static tests and are shown to meet the (2) The means of attachment; and strength and deformation requirements (3) The bearing on the joined mem- of § 25.305(a) and (b).

bers.

(2) A casting factor of 1.25 or greater (b) No fitting factor need be used— may be used, provided that each cast- (1) For joints made under approved ing receives: practices and based on comprehensive (i) Inspection of 100 percent of its sur- test data (such as continuous joints in face, using visual inspection and liquid metal plating, welded joints, and scarf penetrant or equivalent inspection joints in wood); or methods; and (2) With respect to any bearing sur- (ii) Inspection of structurally signifi- face for which a larger special factor is cant internal areas and areas where de- used.

fects are likely to occur, using radio- (c) For each integral fitting, the part graphic or equivalent inspection meth- must be treated as a fitting up to the ods.

point at which the section properties (3) A casting factor of 1.5 or greater become typical of the member.

may be used, provided that each cast- (d) For each seat, berth, safety belt, ing receives inspection of 100 percent of and harness, the fitting factor specified its surface using visual inspection and in § 25.785(f)(3) applies.

liquid penetrant or equivalent inspec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tion methods.

amended by Amdt. 25–23, 35 FR 5674, Apr. 8, (4) A casting factor of 2.0 or greater 1970; Amdt. 25–72, 55 FR 29776, July 20, 1990] may be used, provided that each cast- ing receives inspection of 100 percent of § 25.629 Aeroelastic stability require- its surface using visual inspection ments.

methods.

(a) General. The aeroelastic stability (5) The number of castings per pro- evaluation required under this section duction batch to be inspected by non- includes flutter, divergence, control re- visual methods in accordance with versal and any undue loss of stability paragraphs (d)(2) and (3) of this section and control as a result of structural de- may be reduced when an approved qual- formation. The aeroelastic evaluation ity control procedure is established.

must include whirl modes associated with any propeller or rotating device [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as that contributes significant dynamic amended by Amdt. 25–139, 79 FR 59429, Oct. 2, 2014] forces. Additionally, the evaluation Federal Aviation Administration, DOT § 25.629 must include any condition of oper- probable, which may result from mis- ation within the maneuvering enve- management of fuel.

lope. Compliance with this section (2) Any single failure in any flutter must be shown by analyses, wind tun- damper system.

nel tests, ground vibration tests, flight (3) For airplanes not approved for op- tests, or other means found necessary eration in icing conditions, the max- by the Administrator. imum likely ice accumulation expected (b) Aeroelastic stability envelopes. The as a result of an inadvertent encounter.

airplane must be designed to be free (4) Failure of any single element of from aeroelastic instability for all con- the structure supporting any engine, figurations and design conditions with- independently mounted propeller shaft, in the aeroelastic stability envelopes large auxiliary power unit, or large ex- as follows: ternally mounted aerodynamic body (1) For normal conditions without (such as an external fuel tank).

failures, malfunctions, or adverse con- (5) For airplanes with engines that ditions, all combinations of altitudes have propellers or large rotating de- and speeds encompassed by the V /M vices capable of significant dynamic D D versus altitude envelope enlarged at all forces, any single failure of the engine points by an increase of 15 percent in structure that would reduce the rigid- equivalent airspeed at both constant ity of the rotational axis.

Mach number and constant altitude. In (6) The absence of aerodynamic or gy- addition, a proper margin of stability roscopic forces resulting from the most must exist at all speeds up to V /M adverse combination of feathered pro- D D and, there must be no large and rapid pellers or other rotating devices capa- reduction in stability as V /M is ap- ble of significant dynamic forces. In D D proached. The enlarged envelope may addition, the effect of a single feath- be limited to Mach 1.0 when M is less ered propeller or rotating device must D than 1.0 at all design altitudes, and be coupled with the failures of para- (2) For the conditions described in graphs (d)(4) and (d)(5) of this section.

§ 25.629(d) below, for all approved alti- (7) Any single propeller or rotating tudes, any airspeed up to the greater device capable of significant dynamic airspeed defined by; forces rotating at the highest likely (i) The V /M envelope determined by overspeed.

D D § 25.335(b); or, (8) Any damage or failure condition, (ii) An altitude-airspeed envelope de- required or selected for investigation fined by a 15 percent increase in equiv- by § 25.571. The single structural fail- alent airspeed above V at constant al- ures described in paragraphs (d)(4) and C titude, from sea level to the altitude of (d)(5) of this section need not be consid- the intersection of 1.15 V with the ex- ered in showing compliance with this C tension of the constant cruise Mach section if; number line, M , then a linear vari- (i) The structural element could not C ation in equivalent airspeed to M + .05 fail due to discrete source damage re- C at the altitude of the lowest V /M sulting from the conditions described C C intersection; then, at higher altitudes, in § 25.571(e), and up to the maximum flight altitude, the (ii) A damage tolerance investigation boundary defined by a .05 Mach in- in accordance with § 25.571(b) shows crease in M at constant altitude. that the maximum extent of damage C (c) Balance weights. If concentrated assumed for the purpose of residual balance weights are used, their effec- strength evaluation does not involve tiveness and strength, including sup- complete failure of the structural ele- porting structure, must be substan- ment.

tiated. (9) The following flight control sys- (d) Failures, malfunctions, and adverse tem failure combinations in which conditions. The failures, malfunctions, aeroelastic stability relies on flight and adverse conditions that must be control system stiffness, damping or considered in showing compliance with both: this section are: (i) Any dual hydraulic system failure.

(1) Any critical fuel loading condi- (ii) Any dual electrical system fail- tions, not shown to be extremely im- ure.

14 CFR Ch. I (1–1–25 Edition) § 25.631 (iii) Any single failure in combina- C ONTROL SURFACES tion with any probable hydraulic or § 25.651 Proof of strength.

electrical system failure.

(10) Any damage, failure, or malfunc- (a) Limit load tests of control sur- faces are required. These tests must in- tion considered under §§ 25.631, 25.671, clude the horn or fitting to which the 25.672, and 25.1309.

control system is attached.

(11) Any other combination of fail- (b) Compliance with the special fac- ures, malfunctions, or adverse condi- tors requirements of §§ 25.619 through tions not shown to be extremely im- 25.625 and 25.657 for control surface probable.

hinges must be shown by analysis or (e) Flight flutter testing. Full scale individual load tests.

flight flutter tests at speeds up to V / DF M must be conducted for new type DF § 25.655 Installation.

designs and for modifications to a type (a) Movable tail surfaces must be in- design unless the modifications have stalled so that there is no interference been shown to have an insignificant ef- between any surfaces when one is held fect on the aeroelastic stability. These in its extreme position and the others tests must demonstrate that the air- are operated through their full angular plane has a proper margin of damping movement.

at all speeds up to V /M , and that DF DF (b) If an adjustable stabilizer is used, there is no large and rapid reduction in it must have stops that will limit its damping as V /M , is approached. If a DF DF range of travel to the maximum for failure, malfunction, or adverse condi- which the airplane is shown to meet tion is simulated during flight test in the trim requirements of § 25.161.

showing compliance with paragraph (d) of this section, the maximum speed in- § 25.657 Hinges.

vestigated need not exceed V /M if it FC FC (a) For control surface hinges, in- is shown, by correlation of the flight cluding ball, roller, and self-lubricated test data with other test data or anal- bearing hinges, the approved rating of yses, that the airplane is free from any the bearing may not be exceeded. For aeroelastic instability at all speeds nonstandard bearing hinge configura- within the altitude-airspeed envelope tions, the rating must be established described in paragraph (b)(2) of this on the basis of experience or tests and, section.

in the absence of a rational investiga- tion, a factor of safety of not less than [Doc. No. 26007, 57 FR 28949, June 29, 1992, as 6.67 must be used with respect to the amended by Doc. No. FAA–2022–1544, 89 FR ultimate bearing strength of the soft- 68731, Aug. 27, 2024] est material used as a bearing.

(b) Hinges must have enough § 25.631 Bird strike damage.

strength and rigidity for loads parallel The empennage structure must be de- to the hinge line.

signed to assure capability of contin- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] ued safe flight and landing of the air- plane after impact with an 8-pound bird C ONTROL S YSTEMS when the velocity of the airplane (rel- ative to the bird along the airplane’s § 25.671 General.

flight path) is equal to V at sea level, C (a) Each flight control system must selected under § 25.335(a). Compliance operate with the ease, smoothness, and with this section by provision of redun- positiveness appropriate to its func- dant structure and protected location tion. The flight control system must of control system elements or protec- continue to operate and respond appro- tive devices such as splitter plates or priately to commands, and must not energy absorbing material is accept- hinder airplane recovery, when the air- able. Where compliance is shown by plane is experiencing any pitch, roll, or analysis, tests, or both, use of data on yaw rate, or vertical load factor that airplanes having similar structural de- could occur due to operating or envi- sign is acceptable.

ronmental conditions, or when the air- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] plane is in any attitude.

Federal Aviation Administration, DOT § 25.672 (b) Each element of each flight con- vided whenever the airplane enters any trol system must be designed, or dis- mode that significantly changes or de- tinctively and permanently marked, to grades the normal handling or oper- minimize the probability of incorrect ational characteristics of the airplane.

assembly that could result in failure or [Doc. No. FAA–2022–1544, 89 FR 68734, Aug. 27, malfunctioning of the system. The ap- 2024] plicant may use distinctive and perma- nent marking only where design means § 25.672 Stability augmentation and automatic and power-operated sys- are impractical.

tems.

(c) The airplane must be shown by analysis, test, or both, to be capable of If the functioning of stability aug- continued safe flight and landing after mentation or other automatic or any of the following failures or jams in power-operated systems is necessary to the flight control system within the show compliance with the flight char- normal flight envelope. Probable mal- acteristics requirements of this part, functions must have only minor effects such systems must comply with § 25.671 on control system operation and must and the following: be capable of being readily counter- (a) A warning which is clearly distin- acted by the pilot.

guishable to the pilot under expected (1) Any single failure, excluding fail- flight conditions without requiring his ures of the type defined in § 25.671(c)(3); attention must be provided for any (2) Any combination of failures not failure in the stability augmentation shown to be extremely improbable, ex- system or in any other automatic or cluding failures of the type defined in power-operated system which could re- § 25.671(c)(3); and sult in an unsafe condition if the pilot (3) Any failure or event that results were not aware of the failure. Warning in a jam of a flight control surface or systems must not activate the control pilot control that is fixed in position systems.

due to a physical interference. The jam (b) The design of the stability aug- must be evaluated as follows: mentation system or of any other auto- (i) The jam must be considered at matic or power-operated system must any normally encountered position of permit initial counteraction of failures the control surface or pilot control.

of the type specified in § 25.671(c) with- (ii) The jam must be assumed to out requiring exceptional pilot skill or occur anywhere within the normal strength, by either the deactivation of flight envelope and during any flight the system, or a failed portion thereof, phase except during the time imme- or by overriding the failure by move- diately before touchdown if the risk of ment of the flight controls in the nor- a potential jam is minimized to the ex- mal sense.

tent practical.

(c) It must be shown that after any (iii) In the presence of the jam, any single failure of the stability aug- additional failure conditions that could mentation system or any other auto- prevent continued safe flight and land- matic or power-operated system— ing must have a combined probability (1) The airplane is safely controllable of 1/1000 or less. when the failure or malfunction occurs (d) If all engines fail at any point in at any speed or altitude within the ap- the flight, the airplane must be con- proved operating limitations that is trollable, and an approach and flare to critical for the type of failure being a landing and controlled stop, and flare considered; to a ditching, must be possible, with- (2) The controllability and maneuver- out requiring exceptional piloting skill ability requirements of this part are or strength. met within a practical operational (e) The airplane must be designed to flight envelope (for example, speed, al- indicate to the flightcrew whenever the titude, normal acceleration, and air- primary control means is near the plane configurations) which is de- limit of control authority. scribed in the Airplane Flight Manual; (f) If the flight control system has and multiple modes of operation, appro- (3) The trim, stability, and stall char- priate flightcrew alerting must be pro- acteristics are not impaired below a 14 CFR Ch. I (1–1–25 Edition) § 25.675 level needed to permit continued safe § 25.679 Control system gust locks.

flight and landing.

(a) There must be a device to prevent damage to the control surfaces (includ- [Amdt. 25–23, 35 FR 5675 Apr. 8, 1970] ing tabs), and to the control system, from gusts striking the airplane while § 25.675 Stops.

it is on the ground or water. If the de- (a) Each control system must have vice, when engaged, prevents normal stops that positively limit the range of operation of the control surfaces by the motion of each movable aerodynamic pilot, it must— surface controlled by the system.

(1) Automatically disengage when the (b) Each stop must be located so that pilot operates the primary flight con- wear, slackness, or take-up adjust- trols in a normal manner; or ments will not adversely affect the (2) Limit the operation of the air- control characteristics of the airplane plane so that the pilot receives unmis- because of a change in the range of sur- takable warning at the start of takeoff.

(b) The device must have means to face travel.

preclude the possibility of it becoming (c) Each stop must be able to with- inadvertently engaged in flight.

stand any loads corresponding to the design conditions for the control sys- § 25.681 Limit load static tests.

tem.

(a) Compliance with the limit load [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as requirements of this Part must be amended by Amdt. 25–38, 41 FR 55466, Dec. 20, shown by tests in which— 1976] (1) The direction of the test loads produces the most severe loading in the § 25.677 Trim systems.

control system; and (a) Trim controls must be designed to (2) Each fitting, pulley, and bracket prevent inadvertent or abrupt oper- used in attaching the system to the main structure is included.

ation and to operate in the plane, and (b) Compliance must be shown (by with the sense of motion, of the air- analyses or individual load tests) with plane.

the special factor requirements for (b) There must be means adjacent to control system joints subject to angu- the trim control to indicate the direc- lar motion.

tion of the control movement relative to the airplane motion. In addition, § 25.683 Operation tests.

there must be clearly visible means to (a) It must be shown by operation indicate the position of the trim device tests that when portions of the control with respect to the range of adjust- system subject to pilot effort loads are ment. The indicator must be clearly loaded to 80 percent of the limit load marked with the range within which it specified for the system and the pow- has been demonstrated that takeoff is ered portions of the control system are safe for all center of gravity positions loaded to the maximum load expected approved for takeoff.

in normal operation, the system is free (c) Trim control systems must be de- from— signed to prevent creeping in flight. (1) Jamming; Trim tab controls must be irreversible (2) Excessive friction; and (3) Excessive deflection.

unless the tab is appropriately bal- (b) It must be shown by analysis and, anced and shown to be free from flut- where necessary, by tests, that in the ter.

presence of deflections of the airplane (d) If an irreversible tab control sys- structure due to the separate applica- tem is used, the part from the tab to tion of pitch, roll, and yaw limit ma- the attachment of the irreversible unit neuver loads, the control system, when to the airplane structure must consist loaded to obtain these limit loads and of a rigid connection.

operated within its operational range of deflections, can be exercised about [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5675, Apr. 8, all control axes and remain free from— 1970; Amdt. 25–115, 69 FR 40527, July 2, 2004] (1) Jamming; Federal Aviation Administration, DOT § 25.697 (2) Excessive friction; (e) Turnbuckles must be attached to (3) Disconnection; and parts having angular motion in a man- (4) Any form of permanent damage.

ner that will positively prevent binding (c) It must be shown that under vi- throughout the range of travel.

bration loads in the normal flight and (f) There must be provisions for vis- ground operating conditions, no hazard ual inspection of fairleads, pulleys, ter- can result from interference or contact minals, and turnbuckles.

with adjacent elements.

§ 25.693 Joints.

[Amdt. 25–139, 79 FR 59430, Oct. 2, 2014] Control system joints (in push-pull § 25.685 Control system details.

systems) that are subject to angular (a) Each detail of each control sys- motion, except those in ball and roller tem must be designed and installed to bearing systems, must have a special prevent jamming, chafing, and inter- factor of safety of not less than 3.33 ference from cargo, passengers, loose with respect to the ultimate bearing objects, or the freezing of moisture.

strength of the softest material used as (b) There must be means in the cock- a bearing. This factor may be reduced pit to prevent the entry of foreign ob- to 2.0 for joints in cable control sys- jects into places where they would jam tems. For ball or roller bearings, the the system.

approved ratings may not be exceeded.

(c) There must be means to prevent [Amdt. 25–72, 55 FR 29777, July 20, 1990] the slapping of cables or tubes against other parts.

§ 25.697 Lift and drag devices, con- (d) Sections 25.689 and 25.693 apply to trols.

cable systems and joints.

(a) Each lift device control must be [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as designed so that the pilots can place amended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976] the device in any takeoff, en route, ap- proach, or landing position established § 25.689 Cable systems.

under § 25.101(d). Lift and drag devices (a) Each cable, cable fitting, turn- must maintain the selected positions, buckle, splice, and pulley must be ap- except for movement produced by an proved. In addition— automatic positioning or load limiting (1) No cable smaller than ⁄ 8 inch in device, without further attention by diameter may be used in the aileron, the pilots.

elevator, or rudder systems; and (b) Each lift and drag device control (2) Each cable system must be de- must be designed and located to make signed so that there will be no haz- inadvertent operation improbable. Lift ardous change in cable tension and drag devices intended for ground throughout the range of travel under operation only must have means to operating conditions and temperature prevent the inadvertant operation of variations.

their controls in flight if that oper- (b) Each kind and size of pulley must ation could be hazardous.

correspond to the cable with which it is (c) The rate of motion of the surfaces used. Pulleys and sprockets must have in response to the operation of the con- closely fitted guards to prevent the ca- trol and the characteristics of the bles and chains from being displaced or automatic positioning or load limiting fouled. Each pulley must lie in the device must give satisfactory flight plane passing through the cable so that and performance characteristics under the cable does not rub against the pul- ley flange. steady or changing conditions of air- (c) Fairleads must be installed so speed, engine power, and airplane atti- that they do not cause a change in tude.

cable direction of more than three de- (d) The lift device control must be grees.

designed to retract the surfaces from (d) Clevis pins subject to load or mo- the fully extended position, during tion and retained only by cotter pins steady flight at maximum continuous may not be used in the control system.

14 CFR Ch. I (1–1–25 Edition) § 25.699 engine power at any speed below V + (d) The interconnection must be de- F 9.0 (knots). signed for the loads resulting when interconnected flap or slat surfaces on [Amdt. 25–23, 35 FR 5675, Apr. 8, 1970, as one side of the plane of symmetry are amended by Amdt. 25–46, 43 FR 50595, Oct. 30, jammed and immovable while the sur- 1978; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] faces on the other side are free to move § 25.699 Lift and drag device indicator. and the full power of the surface actu- ating system is applied.

(a) There must be means to indicate to the pilots the position of each lift or [Amdt. 25–72, 55 FR 29777, July 20, 1990] drag device having a separate control § 25.703 Takeoff warning system.

in the cockpit to adjust its position. In addition, an indication of unsymmet- A takeoff warning system must be in- rical operation or other malfunction in stalled and must meet the following re- the lift or drag device systems must be quirements: provided when such indication is nec- (a) The system must provide to the essary to enable the pilots to prevent pilots an aural warning that is auto- or counteract an unsafe flight or matically activated during the initial ground condition, considering the ef- portion of the takeoff roll if the air- fects on flight characteristics and per- plane is in a configuration, including formance.

any of the following, that would not (b) There must be means to indicate allow a safe takeoff: to the pilots the takeoff, en route, ap- (1) The wing flaps or leading edge de- proach, and landing lift device posi- vices are not within the approved range tions.

of takeoff positions.

(c) If any extension of the lift and (2) Wing spoilers (except lateral con- drag devices beyond the landing posi- trol spoilers meeting the requirements tion is possible, the controls must be of § 25.671), speed brakes, or longitu- clearly marked to identify this range dinal trim devices are in a position of extension.

that would not allow a safe takeoff.

(b) The warning required by para- [Amdt. 25–23, 35 FR 5675, Apr. 8, 1970] graph (a) of this section must continue § 25.701 Flap and slat interconnection. until— (1) The configuration is changed to (a) Unless the airplane has safe flight allow a safe takeoff; characteristics with the flaps or slats (2) Action is taken by the pilot to retracted on one side and extended on terminate the takeoff roll; the other, the motion of flaps or slats (3) The airplane is rotated for take- on opposite sides of the plane of sym- off; or metry must be synchronized by a me- (4) The warning is manually deacti- chanical interconnection or approved vated by the pilot.

equivalent means.

(c) The means used to activate the (b) If a wing flap or slat interconnec- system must function properly tion or equivalent means is used, it throughout the ranges of takeoff must be designed to account for the ap- weights, altitudes, and temperatures plicable unsymmetrical loads, includ- for which certification is requested.

ing those resulting from flight with the engines on one side of the plane of sym- [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978] metry inoperative and the remaining L ANDING G EAR engines at takeoff power.

(c) For airplanes with flaps or slats § 25.721 General.

that are not subjected to slipstream conditions, the structure must be de- (a) The landing gear system must be signed for the loads imposed when the designed so that when it fails due to wing flaps or slats on one side are car- overloads during takeoff and landing, rying the most severe load occurring in the failure mode is not likely to cause the prescribed symmetrical conditions spillage of enough fuel to constitute a and those on the other side are car- fire hazard. The overloads must be as- rying not more than 80 percent of that sumed to act in the upward and aft di- load. rections in combination with side loads Federal Aviation Administration, DOT § 25.729 acting inboard and outboard. In the ab- tions in a manner consistent with the sence of a more rational analysis, the development of rational or conserv- side loads must be assumed to be up to ative limit loads.

20 percent of the vertical load or 20 per- (b) The landing gear may not fail in a test, demonstrating its reserve en- cent of the drag load, whichever is greater. ergy absorption capacity, simulating a (b) The airplane must be designed to descent velocity of 12 f.p.s. at design avoid any rupture leading to the spill- landing weight, assuming airplane lift age of enough fuel to constitute a fire not greater than airplane weight act- hazard as a result of a wheels-up land- ing during the landing impact.

ing on a paved runway, under the fol- (c) In lieu of the tests prescribed in lowing minor crash landing conditions: this section, changes in previously ap- (1) Impact at 5 feet-per-second proved design weights and minor vertical velocity, with the airplane changes in design may be substantiated under control, at Maximum Design by analyses based on previous tests Landing Weight— conducted on the same basic landing (i) With the landing gear fully re- gear system that has similar energy tracted; and absorption characteristics.

(ii) With any one or more landing [Doc. No. 1999–5835, 66 FR 27394, May 16, 2001] gear legs not extended.

(2) Sliding on the ground, with— §§ 25.725–25.727 [Reserved] (i) The landing gear fully retracted and with up to a 20 ° yaw angle; and § 25.729 Retracting mechanism.

(ii) Any one or more landing gear (a) General. For airplanes with re- legs not extended and with 0 ° yaw tractable landing gear, the following angle.

apply: (c) For configurations where the en- (1) The landing gear retracting mech- gine nacelle is likely to come into con- anism, wheel well doors, and sup- tact with the ground, the engine pylon porting structure, must be designed or engine mounting must be designed for— so that when it fails due to overloads (i) The loads occurring in the flight (assuming the overloads to act pre- conditions when the gear is in the re- dominantly in the upward direction tracted position, and separately, predominantly in the (ii) The combination of friction aft direction), the failure mode is not loads, inertia loads, brake torque loads, likely to cause the spillage of enough air loads, and gyroscopic loads result- fuel to constitute a fire hazard.

ing from the wheels rotating at a pe- [Amdt. 25–139, 79 FR 59430, Oct. 2, 2014] ripheral speed equal to 1.23V (with SR the wing-flaps in take-off position at § 25.723 Shock absorption tests.

design take-off weight), occurring dur- (a) The analytical representation of ing retraction and extension at any the landing gear dynamic characteris- airspeed up to 1.5 V (with the wing- SR1 tics that is used in determining the flaps in the approach position at design landing loads must be validated by en- landing weight), and ergy absorption tests. A range of tests (iii) Any load factor up to those spec- must be conducted to ensure that the ified in § 25.345(a) for the wing-flaps ex- analytical representation is valid for tended condition.

the design conditions specified in (2) Unless there are other means to § 25.473. decelerate the airplane in flight at this (1) The configurations subjected to speed, the landing gear, the retracting energy absorption tests at limit design mechanism, and the airplane structure conditions must include at least the (including wheel well doors) must be design landing weight or the design designed to withstand the flight loads takeoff weight, whichever produces the occurring with the landing gear in the greater value of landing impact energy. extended position at any speed up to (2) The test attitude of the landing 0.67 V .

C gear unit and the application of appro- (3) Landing gear doors, their oper- priate drag loads during the test must ating mechanism, and their supporting simulate the airplane landing condi- structures must be designed for the 14 CFR Ch. I (1–1–25 Edition) § 25.731 yawing maneuvers prescribed for the (4) There must not be a manual shut- airplane in addition to the conditions off means readily available to the of airspeed and load factor prescribed flightcrew for the warning required by in paragraphs (a)(1) and (2) of this sec- paragraph (e)(2) of this section such tion.

that it could be operated instinctively, (b) Landing gear lock. There must be inadvertently, or by habitual reflexive positive means to keep the landing action.

gear extended in flight and on the (5) The system used to generate the ground. There must be positive means aural warning must be designed to to keep the landing gear and doors in minimize false or inappropriate alerts.

the correct retracted position in flight, (6) Failures of systems used to in- unless it can be shown that lowering of hibit the landing gear aural warning, the landing gear or doors, or flight that would prevent the warning system with the landing gear or doors ex- from operating, must be improbable.

tended, at any speed, is not hazardous.

(7) A flightcrew alert must be pro- (c) Emergency operation. There must vided whenever the landing gear posi- be an emergency means for extending tion is not consistent with the landing the landing gear in the event of— gear selector lever position.

(1) Any reasonably probable failure in (f) Protection of equipment on landing the normal retraction system; or gear and in wheel wells. Equipment that (2) The failure of any single source of is essential to the safe operation of the hydraulic, electric, or equivalent en- airplane and that is located on the ergy supply.

landing gear and in wheel wells must (d) Operation test. The proper func- be protected from the damaging effects tioning of the retracting mechanism of— must be shown by operation tests.

(1) A bursting tire; (e) Position indicator and warning de- (2) A loose tire tread, unless it is vice. If a retractable landing gear is shown that a loose tire tread cannot used, there must be a landing gear po- cause damage.

sition indicator easily visible to the (3) Possible wheel brake tempera- pilot or to the appropriate crew mem- tures.

bers (as well as necessary devices to ac- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tuate the indicator) to indicate with- amended by Amdt. 25–23, 35 FR 5676, Apr. 8, out ambiguity that the retractable 1970; Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; units and their associated doors are se- Amdt. 25–72, 55 FR 29777, July 20, 1990; Amdt.

cured in the extended (or retracted) po- 25–75, 56 FR 63762, Dec. 5, 1991; Amdt. 25–136, sition. The means must be designed as 77 FR 1617, Jan. 11, 2012] follows: (1) If switches are used, they must be § 25.731 Wheels.

located and coupled to the landing gear (a) Each main and nose wheel must mechanical systems in a manner that be approved.

prevents an erroneous indication of (b) The maximum static load rating ‘‘down and locked’’ if the landing gear of each wheel may not be less than the is not in a fully extended position, or of corresponding static ground reaction ‘‘up and locked’’ if the landing gear is with— not in the fully retracted position. The (1) Design maximum weight; and switches may be located where they are operated by the actual landing gear (2) Critical center of gravity.

locking latch or device. (c) The maximum limit load rating of each wheel must equal or exceed the (2) The flightcrew must be given an aural warning that functions continu- maximum radial limit load determined ously, or is periodically repeated, if a under the applicable ground load re- landing is attempted when the landing quirements of this part.

gear is not locked down.

(d) Overpressure burst prevention.

(3) The warning must be given in suf- Means must be provided in each wheel ficient time to allow the landing gear to prevent wheel failure and tire burst to be locked down or a go-around to be that may result from excessive pressur- made. ization of the wheel and tire assembly.

Federal Aviation Administration, DOT § 25.735 (e) Braked wheels. Each braked wheel ing at the center of gravity. The reac- must meet the applicable requirements tions in this case must be distributed of § 25.735. to the nose and main wheels by the principles of statics with a drag reac- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tion equal to 0.20 times the vertical amended by Amdt. 25–72, 55 FR 29777, July 20, 1990; Amdt. 25–107, 67 FR 20420, Apr. 24, 2002] load at each wheel with brakes capable of producing this ground reaction. This § 25.733 Tires.

nose tire load may not exceed 1.5 times (a) When a landing gear axle is fitted the load rating of the tire.

with a single wheel and tire assembly, (c) When a landing gear axle is fitted the wheel must be fitted with a suit- with more than one wheel and tire as- able tire of proper fit with a speed rat- sembly, such as dual or dual-tandem, ing approved by the Administrator each wheel must be fitted with a suit- that is not exceeded under critical con- able tire of proper fit with a speed rat- ditions and with a load rating approved ing approved by the Administrator by the Administrator that is not ex- that is not exceeded under critical con- ceeded under— ditions, and with a load rating ap- (1) The loads on the main wheel tire, proved by the Administrator that is corresponding to the most critical not exceeded by— combination of airplane weight (up to (1) The loads on each main wheel maximum weight) and center of grav- tire, corresponding to the most critical ity position, and combination of airplane weight (up to (2) The loads corresponding to the maximum weight) and center of grav- ground reactions in paragraph (b) of ity position, when multiplied by a fac- this section, on the nose wheel tire, ex- tor of 1.07; and cept as provided in paragraphs (b)(2) (2) Loads specified in paragraphs and (b)(3) of this section.

(a)(2), (b)(1), (b)(2), and (b)(3) of this (b) The applicable ground reactions section on each nose wheel tire.

for nose wheel tires are as follows: (d) Each tire installed on a retract- (1) The static ground reaction for the able landing gear system must, at the tire corresponding to the most critical maximum size of the tire type expected combination of airplane weight (up to maximum ramp weight) and center of in service, have a clearance to sur- gravity position with a force of 1.0g rounding structure and systems that is acting downward at the center of grav- adequate to prevent unintended con- ity. This load may not exceed the load tact between the tire and any part of rating of the tire. the structure or systems.

(2) The ground reaction of the tire (e) For an airplane with a maximum corresponding to the most critical certificated takeoff weight of more combination of airplane weight (up to than 75,000 pounds, tires mounted on maximum landing weight) and center braked wheels must be inflated with of gravity position combined with dry nitrogen or other gases shown to be forces of 1.0g downward and 0.31g for- inert so that the gas mixture in the ward acting at the center of gravity.

tire does not contain oxygen in excess The reactions in this case must be dis- of 5 percent by volume, unless it can be tributed to the nose and main wheels shown that the tire liner material will by the principles of statics with a drag not produce a volatile gas when heated reaction equal to 0.31 times the or that means are provided to prevent vertical load at each wheel with brakes tire temperatures from reaching unsafe capable of producing this ground reac- levels.

tion. This nose tire load may not ex- [Amdt. 25–48, 44 FR 68752, Nov. 29, 1979; Amdt.

ceed 1.5 times the load rating of the 25–72, 55 FR 29777, July 20, 1990, as amended tire.

by Amdt. 25–78, 58 FR 11781, Feb. 26, 1993] (3) The ground reaction of the tire corresponding to the most critical § 25.735 Brakes and braking systems.

combination of airplane weight (up to maximum ramp weight) and center of (a) Approval. Each assembly con- gravity position combined with forces sisting of a wheel(s) and brake(s) must of 1.0g downward and 0.20g forward act- be approved.

14 CFR Ch. I (1–1–25 Edition) § 25.735 (b) Brake system capability. The brake eter testing that the wheel, brake and system, associated systems and compo- tire assembly is capable of absorbing nents must be designed and con- not less than this level of kinetic en- structed so that: ergy throughout the defined wear (1) If any electrical, pneumatic, hy- range of the brake. The energy absorp- draulic, or mechanical connecting or tion rate derived from the airplane transmitting element fails, or if any manufacturer’s braking requirements single source of hydraulic or other must be achieved. The mean decelera- brake operating energy supply is lost, tion must not be less than 10 fps .

it is possible to bring the airplane to (2) Maximum kinetic energy accelerate- rest with a braked roll stopping dis- stop. The maximum kinetic energy ac- tance of not more than two times that celerate-stop is a rejected takeoff for obtained in determining the landing the most critical combination of air- distance as prescribed in § 25.125. plane takeoff weight and speed. The ac- (2) Fluid lost from a brake hydraulic celerate-stop brake kinetic energy ab- system following a failure in, or in the sorption requirement of each wheel, vicinity of, the brakes is insufficient to brake, and tire assembly must be de- cause or support a hazardous fire on termined. It must be substantiated by the ground or in flight. dynamometer testing that the wheel, (c) Brake controls. The brake controls brake, and tire assembly is capable of must be designed and constructed so absorbing not less than this level of ki- that: netic energy throughout the defined (1) Excessive control force is not re- wear range of the brake. The energy quired for their operation. absorption rate derived from the air- (2) If an automatic braking system is plane manufacturer’s braking require- installed, means are provided to: ments must be achieved. The mean de- (i) Arm and disarm the system, and celeration must not be less than 6 fps .

(ii) Allow the pilot(s) to override the (3) Most severe landing stop. The most system by use of manual braking. severe landing stop is a stop at the (d) Parking brake. The airplane must most critical combination of airplane have a parking brake control that, landing weight and speed. The most se- when selected on, will, without further vere landing stop brake kinetic energy attention, prevent the airplane from absorption requirement of each wheel, rolling on a dry and level paved runway brake, and tire assembly must be de- when the most adverse combination of termined. It must be substantiated by maximum thrust on one engine and up dynamometer testing that, at the de- to maximum ground idle thrust on any, clared fully worn limit(s) of the brake or all, other engine(s) is applied. The heat sink, the wheel, brake and tire as- control must be suitably located or be sembly is capable of absorbing not less adequately protected to prevent inad- than this level of kinetic energy. The vertent operation. There must be indi- most severe landing stop need not be cation in the cockpit when the parking considered for extremely improbable brake is not fully released. failure conditions or if the maximum (e) Antiskid system. If an antiskid sys- kinetic energy accelerate-stop energy tem is installed: is more severe.

(1) It must operate satisfactorily over (g) Brake condition after high kinetic the range of expected runway condi- energy dynamometer stop(s). Following tions, without external adjustment. the high kinetic energy stop dem- (2) It must, at all times, have pri- onstration(s) required by paragraph (f) ority over the automatic braking sys- of this section, with the parking brake tem, if installed. promptly and fully applied for at least (f) Kinetic energy capacity —(1) Design 3 minutes, it must be demonstrated landing stop. The design landing stop is that for at least 5 minutes from appli- an operational landing stop at max- cation of the parking brake, no condi- imum landing weight. The design land- tion occurs (or has occurred during the ing stop brake kinetic energy absorp- stop), including fire associated with tion requirement of each wheel, brake, the tire or wheel and brake assembly, and tire assembly must be determined. that could prejudice the safe and com- It must be substantiated by dynamom- plete evacuation of the airplane.

Federal Aviation Administration, DOT § 25.772 (h) Stored energy systems. An indica- § 25.755 Hulls.

tion to the flightcrew of the usable (a) Each hull must have enough wa- stored energy must be provided if a tertight compartments so that, with stored energy system is used to show any two adjacent compartments flood- compliance with paragraph (b)(1) of ed, the buoyancy of the hull and auxil- this section. The available stored en- iary floats (and wheel tires, if used) ergy must be sufficient for: provides a margin of positive stability (1) At least 6 full applications of the great enough to minimize the prob- brakes when an antiskid system is not ability of capsizing in rough, fresh operating; and water.

(2) Bringing the airplane to a com- (b) Bulkheads with watertight doors plete stop when an antiskid system is may be used for communication be- operating, under all runway surface tween compartments.

conditions for which the airplane is certificated.

P ERSONNEL AND C ARGO (i) Brake wear indicators. Means must A CCOMMODATIONS be provided for each brake assembly to indicate when the heat sink is worn to § 25.771 Pilot compartment.

the permissible limit. The means must (a) Each pilot compartment and its be reliable and readily visible.

equipment must allow the minimum (j) Overtemperature burst prevention.

flight crew (established under § 25.1523) Means must be provided in each braked to perform their duties without unrea- wheel to prevent a wheel failure, a tire sonable concentration or fatigue.

burst, or both, that may result from (b) The primary controls listed in elevated brake temperatures. Addition- § 25.779(a), excluding cables and control ally, all wheels must meet the require- rods, must be located with respect to ments of § 25.731(d).

the propellers so that no member of the (k) Compatibility. Compatibility of minimum flight crew (established the wheel and brake assemblies with under § 25.1523), or part of the controls, the airplane and its systems must be lies in the region between the plane of substantiated.

rotation of any inboard propeller and [Doc. No. FAA–1999–6063, 67 FR 20420, Apr. 24, the surface generated by a line passing 2002, as amended by Amdt. 25–108, 67 FR through the center of the propeller hub 70827, Nov. 26, 2002; 68 FR 1955, Jan. 15, 2003] making an angle of five degrees for- ward or aft of the plane of rotation of § 25.737 Skis.

the propeller.

Each ski must be approved. The max- (c) If provision is made for a second imum limit load rating of each ski pilot, the airplane must be controllable must equal or exceed the maximum with equal safety from either pilot limit load determined under the appli- seat.

cable ground load requirements of this (d) The pilot compartment must be part.

constructed so that, when flying in rain or snow, it will not leak in a man- F LOATS AND H ULLS ner that will distract the crew or harm the structure.

§ 25.751 Main float buoyancy.

(e) Vibration and noise characteris- Each main float must have— tics of cockpit equipment may not (a) A buoyancy of 80 percent in excess interfere with safe operation of the air- of that required to support the max- plane.

imum weight of the seaplane or am- phibian in fresh water; and [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (b) Not less than five watertight com- amended by Amdt. 25–4, 30 FR 6113, Apr. 30, partments approximately equal in vol- 1965] ume.

§ 25.772 Pilot compartment doors.

§ 25.753 Main float design.

For an airplane that has a lockable Each main float must be approved door installed between the pilot com- and must meet the requirements of partment and the passenger compart- § 25.521. ment: 14 CFR Ch. I (1–1–25 Edition) § 25.773 (a) For airplanes with a maximum pendix O of this part, if certification passenger seating configuration of for flight in icing conditions is sought: more than 20 seats, the emergency exit (A) For airplanes certificated in ac- configuration must be designed so that cordance with § 25.1420(a)(1), the icing neither crewmembers nor passengers conditions that the airplane is certified require use of the flightdeck door in to safely exit following detection.

order to reach the emergency exits pro- (B) For airplanes certificated in ac- vided for them; and cordance with § 25.1420(a)(2), the icing (b) Means must be provided to enable conditions that the airplane is certified flight crewmembers to directly enter to safely operate in and the icing con- the passenger compartment from the ditions that the airplane is certified to pilot compartment if the cockpit door safely exit following detection.

becomes jammed.

(C) For airplanes certificated in ac- (c) There must be an emergency cordance with § 25.1420(a)(3) and for air- means to enable a flight attendant to planes not subject to § 25.1420, all icing enter the pilot compartment in the conditions.

event that the flightcrew becomes in- (2) No single failure of the systems capacitated.

used to provide the view required by paragraph (b)(1) of this section must [Doc. No. 24344, 55 FR 29777, July 20, 1990, as cause the loss of that view by both pi- amended by Amdt. 25–106, 67 FR 2127, Jan. 15, lots in the specified precipitation con- 2002] ditions.

§ 25.773 Pilot compartment view.

(3) The first pilot must have a win- dow that— (a) Nonprecipitation conditions. For (i) Is openable under the conditions nonprecipitation conditions, the fol- prescribed in paragraph (b)(1) of this lowing apply: section when the cabin is not pressur- (1) Each pilot compartment must be ized; arranged to give the pilots a suffi- (ii) Provides the view specified in ciently extensive, clear, and undis- paragraph (b)(1) of this section; and torted view, to enable them to safely (iii) Provides sufficient protection perform any maneuvers within the op- from the elements against impairment erating limitations of the airplane, in- of the pilot’s vision.

cluding taxiing takeoff, approach, and (4) The openable window specified in landing.

paragraph (b)(3) of this section need (2) Each pilot compartment must be not be provided if it is shown that an free of glare and reflection that could area of the transparent surface will re- interfere with the normal duties of the main clear sufficient for at least one minimum flight crew (established pilot to land the airplane safely in the under § 25.1523). This must be shown in event of— day and night flight tests under non- (i) Any system failure or combina- precipitation conditions.

tion of failures which is not extremely (b) Precipitation conditions. For pre- improbable, in accordance with cipitation conditions, the following § 25.1309, under the precipitation condi- apply: (1) The airplane must have a means tions specified in paragraph (b)(1) of to maintain a clear portion of the this section.

windshield, during precipitation condi- (ii) An encounter with severe hail, tions, sufficient for both pilots to have birds, or insects.

a sufficiently extensive view along the (c) Internal windshield and window flight path in normal flight attitudes fogging. The airplane must have a of the airplane. This means must be de- means to prevent fogging of the inter- signed to function, without continuous nal portions of the windshield and win- attention on the part of the crew, in— dow panels over an area which would (i) Heavy rain at speeds up to 1.5 V provide the visibility specified in para- SR1 with lift and drag devices retracted; graph (a) of this section under all in- and ternal and external ambient condi- (ii) The icing conditions specified in tions, including precipitation condi- Appendix C of this part and the fol- tions, in which the airplane is intended lowing icing conditions specified in Ap- to be operated.

Federal Aviation Administration, DOT § 25.775 (d) Fixed markers or other guides ing the provisions of paragraph (a)(2) of must be installed at each pilot station this section.

to enable the pilots to position them- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as selves in their seats for an optimum amended by Amdt. 25–23, 35 FR 5676, Apr. 8, combination of outside visibility and 1970; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; Amdt. 25–72, 55 FR 29778, July 20, 1990; Amdt.

instrument scan. If lighted markers or 25–108, 67 FR 70827, Nov. 26, 2002; Amdt. 25– guides are used they must comply with 121, 72 FR 44669, Aug. 8, 2007; Amdt. 25–136, 77 the requirements specified in § 25.1381.

FR 1618, Jan. 11, 2012; Amdt. 25–140, 79 FR (e) Vision systems with transparent dis- 65525, Nov. 4, 2014; Docket FAA–2013–0485, plays. A vision system with a trans- Amdt. 25–144, 81 FR 90169, Dec. 13, 2016] parent display surface located in the § 25.775 Windshields and windows.

pilot’s outside field of view, such as a (a) Internal panes must be made of head up-display, head mounted display, nonsplintering material.

or other equivalent display, must meet (b) Windshield panes directly in front the following requirements in non- of the pilots in the normal conduct of precipitation and precipitation condi- their duties, and the supporting struc- tions: tures for these panes, must withstand, (1) While the vision system display is without penetration, the impact of a in operation, it must compensate for four-pound bird when the velocity of interference with the pilot’s outside the airplane (relative to the bird along field of view such that the combination the airplane’s flight path) is equal to of what is visible in the display and the value of V at sea level, selected C, what remains visible through and under § 25.335(a).

around it, enables the pilot to perform (c) Unless it can be shown by analysis the maneuvers and normal duties of or tests that the probability of occur- paragraph (a) of this section.

rence of a critical windshield frag- (2) The pilot’s view of the external mentation condition is of a low order, scene may not be distorted by the the airplane must have a means to transparent display surface or by the minimize the danger to the pilots from flying windshield fragments due to bird vision system imagery. When the vi- impact. This must be shown for each sion system displays imagery or any transparent pane in the cockpit that— symbology that is referenced to the im- (1) Appears in the front view of the agery and outside scene topography, airplane; including attitude symbology, flight (2) Is inclined 15 degrees or more to path vector, and flight path angle ref- the longitudinal axis of the airplane; erence cue, that imagery and sym- and bology must be aligned with, and (3) Has any part of the pane located scaled to, the external scene.

where its fragmentation will constitute (3) The vision system must provide a a hazard to the pilots.

means to allow the pilot using the dis- (d) The design of windshields and play to immediately deactivate and re- windows in pressurized airplanes must activate the vision system imagery, on be based on factors peculiar to high al- demand, without removing the pilot’s titude operation, including the effects hands from the primary flight controls of continuous and cyclic pressurization loadings, the inherent characteristics or thrust controls.

of the material used, and the effects of (4) When the vision system is not in temperatures and temperature dif- operation it may not restrict the pilot ferentials. The windshield and window from performing the maneuvers speci- panels must be capable of withstanding fied in paragraph (a)(1) of this section the maximum cabin pressure differen- or the pilot compartment from meet- tial loads combined with critical aero- dynamic pressure and temperature ef- fects after any single failure in the in- stallation or associated systems. It may be assumed that, after a single failure that is obvious to the flight crew (established under § 25.1523), the 14 CFR Ch. I (1–1–25 Edition) § 25.777 cabin pressure differential is reduced seated with seat belt and shoulder har- from the maximum, in accordance with ness (if provided) fastened.

appropriate operating limitations, to (g) Control knobs must be shaped in allow continued safe flight of the air- accordance with § 25.781. In addition, plane with a cabin pressure altitude of the knobs must be of the same color, not more than 15,000 feet.

and this color must contrast with the (e) The windshield panels in front of color of control knobs for other pur- the pilots must be arranged so that, as- poses and the surrounding cockpit.

suming the loss of vision through any (h) If a flight engineer is required as one panel, one or more panels remain part of the minimum flight crew (es- available for use by a pilot seated at a tablished under § 25.1523), the airplane pilot station to permit continued safe must have a flight engineer station lo- flight and landing.

cated and arranged so that the flight [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as crewmembers can perform their func- amended by Amdt. 25–23, 35 FR 5676, Apr. 8, tions efficiently and without inter- 1970; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976] fering with each other.

§ 25.777 Cockpit controls.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–46, 43 FR 50596, Oct. 30, (a) Each cockpit control must be lo- 1978] cated to provide convenient operation and to prevent confusion and inad- § 25.779 Motion and effect of cockpit vertent operation.

controls.

(b) The direction of movement of Cockpit controls must be designed so cockpit controls must meet the re- that they operate in accordance with quirements of § 25.779. Wherever prac- the following movement and actuation: ticable, the sense of motion involved in (a) Aerodynamic controls: the operation of other controls must (1) Primary.

correspond to the sense of the effect of the operation upon the airplane or Controls Motion and effect upon the part operated. Controls of a variable nature using a rotary motion Aileron ......................... Right (clockwise) for right wing down.

must move clockwise from the off posi- Elevator ....................... Rearward for nose up.

tion, through an increasing range, to Rudder ......................... Right pedal forward for nose right.

the full on position.

(c) The controls must be located and (2) Secondary.

arranged, with respect to the pilots’ seats, so that there is full and unre- Controls Motion and effect stricted movement of each control Flaps (or auxiliary lift Forward for flaps up; rearward for without interference from the cockpit devices). flaps down.

structure or the clothing of the min- Trim tabs (or equiva- Rotate to produce similar rotation of imum flight crew (established under lent). the airplane about an axis parallel § 25.1523) when any member of this to the axis of the control.

flight crew, from 5 ′ 2 ″ to 6 ′ 3 ″ in height, is seated with the seat belt and shoul- (b) Powerplant and auxiliary con- der harness (if provided) fastened.

trols: (d) Identical powerplant controls for (1) Powerplant.

each engine must be located to prevent confusion as to the engines they con- Controls Motion and effect trol.

Power or thrust ............ Forward to increase forward thrust (e) Wing flap controls and other aux- and rearward to increase rear- iliary lift device controls must be lo- ward thrust.

cated on top of the pedestal, aft of the Propellers .................... Forward to increase rpm.

Mixture ......................... Forward or upward for rich.

throttles, centrally or to the right of Carburetor air heat ...... Forward or upward for cold.

the pedestal centerline, and not less Supercharger ............... Forward or upward for low blower.

than 10 inches aft of the landing gear For turbosuperchargers, forward, control.

upward, or clockwise, to increase pressure.

(f) The landing gear control must be located forward of the throttles and (2) Auxiliary.

must be operable by each pilot when Federal Aviation Administration, DOT § 25.781 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as Controls Motion and effect amended by Amdt. 25–72, 55 FR 29778, July 20, Landing gear ............... Down to extend.

1990] § 25.781 Cockpit control knob shape.

Cockpit control knobs must conform to the general shapes (but not necessarily the exact sizes or specific proportions) in the following figure: [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–72, 55 FR 29779, July 20, 1990] 14 CFR Ch. I (1–1–25 Edition) § 25.783 their controlling systems must be de- § 25.783 Fuselage doors.

signed so that— (a) General. This section applies to fu- (1) No single failure will prevent selage doors, which includes all doors, more than one exit from being opened; hatches, openable windows, access pan- and els, covers, etc., on the exterior of the (2) Failures that would prevent open- fuselage that do not require the use of ing of the exit after landing are im- tools to open or close. This also applies probable.

to each door or hatch through a pres- (c) Pressurization prevention means.

sure bulkhead, including any bulkhead There must be a provision to prevent that is specifically designed to func- pressurization of the airplane to an un- tion as a secondary bulkhead under the safe level if any door subject to pres- prescribed failure conditions of part 25.

surization is not fully closed, latched, These doors must meet the require- and locked.

ments of this section, taking into ac- (1) The provision must be designed to count both pressurized and unpres- function after any single failure, or surized flight, and must be designed as after any combination of failures not follows: shown to be extremely improbable.

(1) Each door must have means to (2) Doors that meet the conditions safeguard against opening in flight as a described in paragraph (h) of this sec- result of mechanical failure, or failure tion are not required to have a dedi- of any single structural element.

cated pressurization prevention means (2) Each door that could be a hazard if, from every possible position of the if it unlatches must be designed so that door, it will remain open to the extent unlatching during pressurized and un- that it prevents pressurization or safe- pressurized flight from the fully closed, ly close and latch as pressurization latched, and locked condition is ex- takes place. This must also be shown tremely improbable. This must be with any single failure and malfunc- shown by safety analysis.

tion, except that— (3) Each element of each door oper- (i) With failures or malfunctions in ating system must be designed or, the latching mechanism, it need not where impracticable, distinctively and latch after closing; and permanently marked, to minimize the (ii) With jamming as a result of me- probability of incorrect assembly and chanical failure or blocking debris, the adjustment that could result in a mal- door need not close and latch if it can function.

be shown that the pressurization loads (4) All sources of power that could on the jammed door or mechanism initiate unlocking or unlatching of any would not result in an unsafe condi- door must be automatically isolated tion.

from the latching and locking systems (d) Latching and locking. The latching prior to flight and it must not be pos- and locking mechanisms must be de- sible to restore power to the door dur- signed as follows: ing flight.

(5) Each removable bolt, screw, nut, (1) There must be a provision to latch pin, or other removable fastener must each door.

meet the locking requirements of (2) The latches and their operating § 25.607. mechanism must be designed so that, (6) Certain doors, as specified by under all airplane flight and ground § 25.807(h), must also meet the applica- loading conditions, with the door ble requirements of §§ 25.809 through latched, there is no force or torque 25.812 for emergency exits. tending to unlatch the latches. In addi- (b) Opening by persons. There must be tion, the latching system must include a means to safeguard each door against a means to secure the latches in the opening during flight due to inad- latched position. This means must be vertent action by persons. In addition, independent of the locking system.

design precautions must be taken to (3) Each door subject to pressuriza- minimize the possibility for a person to tion, and for which the initial opening open a door intentionally during flight. movement is not inward, must— If these precautions include the use of (i) Have an individual lock for each auxiliary devices, those devices and latch; Federal Aviation Administration, DOT § 25.783 (ii) Have the lock located as close as (i) Each door that is subject to pres- practicable to the latch; and surization and for which the initial (iii) Be designed so that, during pres- opening movement is not inward; or surized flight, no single failure in the (ii) Each door that could be a hazard locking system would prevent the if unlatched.

locks from restraining the latches nec- (4) There must be an aural warning essary to secure the door. to the pilots prior to or during the ini- (4) Each door for which the initial tial portion of takeoff roll if any door opening movement is inward, and is not fully closed, latched, and locked, unlatching of the door could result in a and its opening would prevent a safe hazard, must have a locking means to takeoff and return to landing.

prevent the latches from becoming dis- (f) Visual inspection provision. Each engaged. The locking means must en- door for which unlatching of the door sure sufficient latching to prevent could be a hazard must have a provi- opening of the door even with a single sion for direct visual inspection to de- failure of the latching mechanism. termine, without ambiguity, if the (5) It must not be possible to position door is fully closed, latched, and the lock in the locked position if the locked. The provision must be perma- latch and the latching mechanism are nent and discernible under operational not in the latched position. lighting conditions, or by means of a (6) It must not be possible to unlatch flashlight or equivalent light source.

the latches with the locks in the (g) Certain maintenance doors, remov- locked position. Locks must be de- able emergency exits, and access panels.

signed to withstand the limit loads re- Some doors not normally opened ex- sulting from— cept for maintenance purposes or emer- (i) The maximum operator effort gency evacuation and some access pan- when the latches are operated manu- els need not comply with certain para- ally; graphs of this section as follows: (ii) The powered latch actuators, if (1) Access panels that are not subject installed; and to cabin pressurization and would not (iii) The relative motion between the be a hazard if open during flight need latch and the structural counterpart. not comply with paragraphs (a) (7) Each door for which unlatching through (f) of this section, but must would not result in a hazard is not re- have a means to prevent inadvertent quired to have a locking mechanism opening during flight.

meeting the requirements of para- (2) Inward-opening removable emer- graphs (d)(3) through (d)(6) of this sec- gency exits that are not normally re- tion. moved, except for maintenance pur- (e) Warning, caution, and advisory in- poses or emergency evacuation, and dications. Doors must be provided with flight deck-openable windows need not the following indications: comply with paragraphs (c) and (f) of (1) There must be a positive means to this section.

indicate at each door operator’s station (3) Maintenance doors that meet the that all required operations to close, conditions of paragraph (h) of this sec- latch, and lock the door(s) have been tion, and for which a placard is pro- completed. vided limiting use to maintenance ac- (2) There must be a positive means cess, need not comply with paragraphs clearly visible from each operator sta- (c) and (f) of this section.

tion for any door that could be a haz- (h) Doors that are not a hazard. For ard if unlatched to indicate if the door the purposes of this section, a door is is not fully closed, latched, and locked. considered not to be a hazard in the un- (3) There must be a visual means on latched condition during flight, pro- the flight deck to signal the pilots if vided it can be shown to meet all of the any door is not fully closed, latched, following conditions: and locked. The means must be de- (1) Doors in pressurized compart- signed such that any failure or com- ments would remain in the fully closed bination of failures that would result position if not restrained by the in an erroneous closed, latched, and latches when subject to a pressure locked indication is improbable for— greater than ⁄2 psi. Opening by persons, 14 CFR Ch. I (1–1–25 Edition) § 25.785 either inadvertently or intentionally, spine, or by a safety belt and shoulder need not be considered in making this harness that will prevent the head determination. from contacting any injurious object.

(2) The door would remain inside the Each occupant of any other seat must airplane or remain attached to the air- be protected from head injury by a plane if it opens either in pressurized safety belt and, as appropriate to the or unpressurized portions of the flight.

type, location, and angle of facing of This determination must include the each seat, by one or more of the fol- consideration of inadvertent and inten- lowing: tional opening by persons during either (1) A shoulder harness that will pre- pressurized or unpressurized portions vent the head from contacting any in- of the flight.

jurious object.

(3) The disengagement of the latches (2) The elimination of any injurious during flight would not allow depres- object within striking radius of the surization of the cabin to an unsafe head.

level. This safety assessment must in- (3) An energy absorbing rest that will clude the physiological effects on the support the arms, shoulders, head, and occupants.

spine.

(4) The open door during flight would (e) Each berth must be designed so not create aerodynamic interference that the forward part has a padded end that could preclude safe flight and board, canvas diaphragm, or equivalent landing.

means, that can withstand the static (5) The airplane would meet the load reaction of the occupant when structural design requirements with subjected to the forward inertia force the door open. This assessment must specified in § 25.561. Berths must be free include the aeroelastic stability re- from corners and protuberances likely quirements of § 25.629, as well as the to cause injury to a person occupying strength requirements of subpart C of the berth during emergency conditions.

this part.

(f) Each seat or berth, and its sup- (6) The unlatching or opening of the porting structure, and each safety belt door must not preclude safe flight and or harness and its anchorage must be landing as a result of interaction with designed for an occupant weight of 170 other systems or structures.

pounds, considering the maximum load factors, inertia forces, and reactions [Doc. No. 2003–14193, 69 FR 24501, May 3, 2004] among the occupant, seat, safety belt, § 25.785 Seats, berths, safety belts, and and harness for each relevant flight harnesses.

and ground load condition (including the emergency landing conditions pre- (a) A seat (or berth for a nonambu- scribed in § 25.561). In addition— lant person) must be provided for each (1) The structural analysis and test- occupant who has reached his or her ing of the seats, berths, and their sup- second birthday.

(b) Each seat, berth, safety belt, har- porting structures may be determined ness, and adjacent part of the airplane by assuming that the critical load in at each station designated as occupi- the forward, sideward, downward, up- able during takeoff and landing must ward, and rearward directions (as de- be designed so that a person making termined from the prescribed flight, proper use of these facilities will not ground, and emergency landing condi- suffer serious injury in an emergency tions) acts separately or using selected landing as a result of the inertia forces combinations of loads if the required specified in §§ 25.561 and 25.562. strength in each specified direction is (c) Each seat or berth must be ap- substantiated. The forward load factor proved. need not be applied to safety belts for (d) Each occupant of a seat that berths.

makes more than an 18-degree angle (2) Each pilot seat must be designed with the vertical plane containing the for the reactions resulting from the ap- airplane centerline must be protected plication of the pilot forces prescribed from head injury by a safety belt and in § 25.395.

an energy absorbing rest that will sup- (3) The inertia forces specified in port the arms, shoulders, head, and § 25.561 must be multiplied by a factor Federal Aviation Administration, DOT § 25.787 of 1.33 (instead of the fitting factor pre- (i) Each safety belt must be equipped scribed in § 25.625) in determining the with a metal to metal latching device.

strength of the attachment of each (j) If the seat backs do not provide a seat to the structure and each belt or firm handhold, there must be a hand- harness to the seat or structure. grip or rail along each aisle to enable (g) Each seat at a flight deck station persons to steady themselves while must have a restraint system con- using the aisles in moderately rough sisting of a combined safety belt and air.

shoulder harness with a single-point re- (k) Each projecting object that would lease that permits the flight deck occu- injure persons seated or moving about pant, when seated with the restraint the airplane in normal flight must be system fastened, to perform all of the padded.

occupant’s necessary flight deck func- (l) Each forward observer’s seat re- tions. There must be a means to secure quired by the operating rules must be each combined restraint system when shown to be suitable for use in con- not in use to prevent interference with ducting the necessary enroute inspec- the operation of the airplane and with tion.

rapid egress in an emergency.

[Amdt. 25–72, 55 FR 29780, July 20, 1990, as (h) Each seat located in the pas- amended by Amdt. 25–88, 61 FR 57956, Nov. 8, senger compartment and designated for 1996] use during takeoff and landing by a flight attendant required by the oper- § 25.787 Stowage compartments.

ating rules of this chapter must be: (a) Each compartment for the stow- (1) Near a required floor level emer- age of cargo, baggage, carry-on arti- gency exit, except that another loca- cles, and equipment (such as life rafts), tion is acceptable if the emergency and any other stowage compartment, egress of passengers would be enhanced must be designed for its placarded max- with that location. A flight attendant imum weight of contents and for the seat must be located adjacent to each critical load distribution at the appro- Type A or B emergency exit. Other priate maximum load factors cor- flight attendant seats must be evenly responding to the specified flight and distributed among the required floor- ground load conditions, and to those level emergency exits to the extent emergency landing conditions of feasible.

§ 25.561(b)(3) for which the breaking (2) To the extent possible, without loose of the contents of such compart- compromising proximity to a required ments in the specified direction could— floor level emergency exit, located to (1) Cause direct injury to occupants; provide a direct view of the cabin area (2) Penetrate fuel tanks or lines or for which the flight attendant is re- cause fire or explosion hazard by dam- sponsible.

age to adjacent systems; or (3) Positioned so that the seat will (3) Nullify any of the escape facilities not interfere with the use of a passage- provided for use after an emergency way or exit when the seat is not in use.

landing.

(4) Located to minimize the prob- ability that occupants would suffer in- If the airplane has a passenger-seating jury by being struck by items dislodged configuration, excluding pilot seats, of from service areas, stowage compart- 10 seats or more, each stowage com- ments, or service equipment. partment in the passenger cabin, ex- (5) Either forward or rearward facing cept for under seat and overhead com- with an energy absorbing rest that is partments for passenger convenience, designed to support the arms, shoul- must be completely enclosed.

ders, head, and spine. (b) There must be a means to prevent (6) Equipped with a restraint system the contents in the compartments from consisting of a combined safety belt becoming a hazard by shifting, under and shoulder harness unit with a single the loads specified in paragraph (a) of point release. There must be means to this section. For stowage compart- secure each restraint system when not ments in the passenger and crew cabin, in use to prevent interference with if the means used is a latched door, the rapid egress in an emergency. design must take into consideration 14 CFR Ch. I (1–1–25 Edition) § 25.789 the wear and deterioration expected in (c) A placard must be located on or service. adjacent to the door of each receptacle (c) If cargo compartment lamps are used for the disposal of flammable installed, each lamp must be installed waste materials to indicate that use of so as to prevent contact between lamp the receptacle for disposal of ciga- bulb and cargo.

rettes, etc., is prohibited.

(d) Lavatories must have ‘‘No Smok- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–32, 37 FR 3969, Feb. 24, ing’’ or ‘‘No Smoking in Lavatory’’ 1972; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; placards conspicuously located on or Amdt. 25–51, 45 FR 7755, Feb. 4, 1980; Amdt.

adjacent to each side of the entry door.

25–139, 79 FR 59430, Oct. 2, 2014] (e) Symbols that clearly express the intent of the sign or placard may be § 25.789 Retention of items of mass in used in lieu of letters.

passenger and crew compartments and galleys.

[Amdt. 25–72, 55 FR 29780, July 20, 1990, as (a) Means must be provided to pre- amended by Amdt. No. 25–153, 89 FR 68099, vent each item of mass (that is part of Aug. 23, 2024] the airplane type design) in a passenger § 25.793 Floor surfaces.

or crew compartment or galley from becoming a hazard by shifting under The floor surface of all areas which the appropriate maximum load factors are likely to become wet in service corresponding to the specified flight must have slip resistant properties.

and ground load conditions, and to the [Amdt. 25–51, 45 FR 7755, Feb. 4, 1980] emergency landing conditions of § 25.561(b).

§ 25.795 Security considerations.

(b) Each interphone restraint system must be designed so that when sub- (a) Protection of flightcrew compart- jected to the load factors specified in ment. If a flightdeck door is required by § 25.561(b)(3), the interphone will re- operating rules: main in its stowed position.

(1) The bulkhead, door, and any other [Amdt. 25–32, 37 FR 3969, Feb. 24, 1972, as accessible boundary separating the amended by Amdt. 25–46, 43 FR 50596, Oct. 30, flightcrew compartment from occupied 1978] areas must be designed to resist forc- ible intrusion by unauthorized persons § 25.791 Passenger information signs and be capable of withstanding impacts and placards.

of 300 joules (221.3 foot pounds).

(a) Regarding ‘‘No Smoking’’ signs (2) The bulkhead, door, and any other and placards: accessible boundary separating the (1) There must be at least one flightcrew compartment from occupied placard, or lighted sign, stating if areas must be designed to resist a con- smoking is prohibited. The placard or stant 250 pound (1,113 Newtons) tensile lighted sign must be legible to each load on accessible handholds, including person seated in the cabin.

the doorknob or handle.

(2) Lighted ‘‘No Smoking’’ signs must (3) The bulkhead, door, and any other either be operable by a member of the boundary separating the flightcrew flightcrew or be illuminated continu- compartment from any occupied areas ously during airplane operations. Illu- must be designed to resist penetration minated signs must be legible under all by small arms fire and fragmentation probable conditions of cabin illumina- devices to a level equivalent to level tion to each person seated in the cabin.

IIIa of the National Institute of Justice (b) Signs that notify when seat belts (NIJ) Standard 0101.04.

should be fastened and that are in- (4) If required by the operating rules stalled to comply with the operating of this chapter, an installed physical rules of this chapter must be operable secondary barrier (IPSB) must be in- by a member of the flightcrew and, stalled to resist intrusion into the when illuminated, must be legible under all probable conditions of cabin flightdeck whenever the flightdeck illumination to each person seated in door is opened. When deployed, the the cabin. IPSB must: Federal Aviation Administration, DOT § 25.795 (i) Resist a 250 pound (1113 Newtons) may be assumed to be 4 feet (1.22 me- static load in the direction of the pas- ters) in length; and senger cabin applied at the most crit- (iii) A 6-inch (0.152 meters) displace- ical locations on the IPSB; ment, except where limited by the fu- selage contour, from a single point (ii) Resist a 600 pound (2669 Newtons) force applied anywhere along the dis- static load in the direction of the tribution system where relative move- flightdeck applied at the most critical ment between the system and its at- locations on the IPSB; tachment can occur.

(iii) Delay a person attempting to ac- (iv) Paragraphs (b)(3)(i) through (iii) cess the flightdeck by at least the time of this section do not apply to compo- required for a crewmember to open and nents that are redundant and separated reclose the flightdeck door, but no less in accordance with paragraph (c)(2) of than 5 seconds; this section or are installed remotely (iv) Prevent a person from reaching from the cargo compartment.

through and touching the flightdeck (c) An airplane with a maximum cer- door; tificated passenger seating capacity of (v) Allow for necessary crewmember more than 60 persons or a maximum activities; and certificated takeoff gross weight of (vi) Provide line-of-sight visibility over 100,000 pounds (45,359 Kilograms) between the flightdeck door and the must comply with the following: cabin.

(1) Least risk bomb location. An air- (b) Airplanes with a maximum cer- plane must be designed with a des- tificated passenger seating capacity of ignated location where a bomb or other more than 60 persons or a maximum explosive device could be placed to best certificated takeoff gross weight of protect flight-critical structures and over 100,000 pounds (45,359 Kilograms) systems from damage in the case of must be designed to limit the effects of detonation.

an explosive or incendiary device as (2) Survivability of systems. (i) Except follows: where impracticable, redundant air- (1) Flightdeck smoke protection. Means plane systems necessary for continued must be provided to limit entry of safe flight and landing must be phys- smoke, fumes, and noxious gases into ically separated, at a minimum, by an the flightdeck.

amount equal to a sphere of diameter (2) Passenger cabin smoke protection.

Means must be provided to prevent pas- senger incapacitation in the cabin re-

D H = ( ) 2 / π

sulting from smoke, fumes, and nox- ious gases as represented by the initial (where H is defined under § 25.365(e)(2) combined volumetric concentrations of of this part and D need not exceed 5.05 0.59% carbon monoxide and 1.23% car- feet (1.54 meters)). The sphere is ap- bon dioxide.

plied everywhere within the fuselage— (3) Cargo compartment fire suppression.

limited by the forward bulkhead and An extinguishing agent must be capa- the aft bulkhead of the passenger cabin ble of suppressing a fire. All cargo- and cargo compartment beyond which compartment fire suppression systems only one-half the sphere is applied.

must be designed to withstand the fol- (ii) Where compliance with paragraph lowing effects, including support struc- (c)(2)(i) of this section is impracticable, ture displacements or adjacent mate- other design precautions must be taken rials displacing against the distribu- to maximize the survivability of those tion system: systems.

(i) Impact or damage from a 0.5-inch (3) Interior design to facilitate searches.

diameter aluminum sphere traveling at Design features must be incorporated 430 feet per second (131.1 meters per that will deter concealment or promote second); discovery of weapons, explosives, or (ii) A 15-pound per square-inch (103.4 other objects from a simple inspection kPa) pressure load if the projected sur- in the following areas of the airplane face area of the component is greater cabin: than 4 square feet. Any single dimen- (i) Areas above the overhead bins sion greater than 4 feet (1.22 meters) must be designed to prevent objects 14 CFR Ch. I (1–1–25 Edition) § 25.801 from being hidden from view in a sim- P.O. Box 6000, Rockville, MD 20849–6000, ple search from the aisle. Designs that telephone (800) 851–3420.

prevent concealment of objects with [Amdt. 25–127; 121–341, 73 FR 63879, Oct. 28, volumes 20 cubic inches and greater 2008, as amended at 74 FR 22819, May 15, 2009; satisfy this requirement.

Amdt. 25–138, 79 FR 13519, Mar. 11, 2014; Doc.

(ii) Toilets must be designed to pre- No. FAA–2018–0119, Amdt. 25–145, 83 FR 9169, vent the passage of solid objects great- Mar. 5, 2018; Amdt. 25–150, 88 FR 41308, June er than 2.0 inches in diameter.

26, 2023] (iii) Life preservers or their storage E MERGENCY P ROVISIONS locations must be designed so that tampering is evident.

§ 25.801 Ditching.

(d) Each chemical oxygen generator or its installation must be designed to (a) If certification with ditching pro- be secure from deliberate manipulation visions is requested, the airplane must by one of the following: meet the requirements of this section (1) By providing effective resistance and §§ 25.807(e), 25.1411, and 25.1415(a).

to tampering, (b) Each practicable design measure, (2) By providing an effective com- compatible with the general character- bination of resistance to tampering and istics of the airplane, must be taken to active tamper-evident features, minimize the probability that in an (3) By installation in a location or emergency landing on water, the be- manner whereby any attempt to access havior of the airplane would cause im- the generator would be immediately mediate injury to the occupants or obvious, or would make it impossible for them to (4) By a combination of approaches escape.

specified in paragraphs (d)(1), (d)(2) and (c) The probable behavior of the air- (d)(3) of this section that the Adminis- plane in a water landing must be inves- trator finds provides a secure installa- tigated by model tests or by compari- tion.

son with airplanes of similar configura- (e) Exceptions. Airplanes used solely tion for which the ditching characteris- to transport cargo only need to meet tics are known. Scoops, flaps, projec- the requirements of paragraphs (b)(1), tions, and any other factor likely to af- (b)(3), and (c)(2) of this section.

fect the hydrodynamic characteristics (f) Material Incorporated by Reference.

of the airplane, must be considered.

You must use National Institute of (d) It must be shown that, under rea- Justice (NIJ) Standard 0101.04, Ballistic sonably probable water conditions, the Resistance of Personal Body Armor, flotation time and trim of the airplane June 2001, Revision A, to establish bal- will allow the occupants to leave the listic resistance as required by para- airplane and enter the liferafts re- graph (a)(3) of this section.

quired by § 25.1415. If compliance with (1) The Director of the Federal Reg- this provision is shown by buoyancy ister approved the incorporation by ref- and trim computations, appropriate al- erence of this document under 5 U.S.C.

lowances must be made for probable 552(a) and 1 CFR part 51.

structural damage and leakage. If the (2) You may review copies of NIJ airplane has fuel tanks (with fuel jetti- Standard 0101.04 at the: soning provisions) that can reasonably (i) National Institute of Justice be expected to withstand a ditching (NIJ), http://www.ojp.usdoj.gov/nij, tele- without leakage, the jettisonable vol- phone (202) 307–2942; or ume of fuel may be considered as buoy- (ii) National Archives and Records ancy volume.

Administration (NARA). For informa- (e) Unless the effects of the collapse tion on the availability of this mate- rial at NARA, call (202) 741–6030, or go of external doors and windows are ac- to http://www.archives.gov/federal-reg- counted for in the investigation of the ister/cfr/ibr-locations.html. probable behavior of the airplane in a (3) You may obtain copies of NIJ water landing (as prescribed in para- Standard 0101.04 from the National graphs (c) and (d) of this section), the Criminal Justice Reference Service, external doors and windows must be Federal Aviation Administration, DOT § 25.807 designed to withstand the probable radii not greater than seven inches, maximum local pressures. and with a step-up inside the airplane of not more than 20 inches. If the exit [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as is located over the wing, the step-down amended by Amdt. 25–72, 55 FR 29781, July 20, outside the airplane may not exceed 27 1990] inches.

§ 25.803 Emergency evacuation. (4) Type IV. This type is a rectan- gular opening of not less than 19 inches (a) Each crew and passenger area wide by 26 inches high, with corner must have emergency means to allow radii not greater than 6.3 inches, lo- rapid evacuation in crash landings, cated over the wing, with a step-up in- with the landing gear extended as well side the airplane of not more than 29 as with the landing gear retracted, con- inches and a step-down outside the air- sidering the possibility of the airplane plane of not more than 36 inches.

being on fire.

(5) Ventral. This type is an exit from (b) [Reserved] the passenger compartment through (c) For airplanes having a seating ca- the pressure shell and the bottom fuse- pacity of more than 44 passengers, it lage skin. The dimensions and physical must be shown that the maximum configuration of this type of exit must seating capacity, including the number allow at least the same rate of egress of crewmembers required by the oper- as a Type I exit with the airplane in ating rules for which certification is the normal ground attitude, with land- requested, can be evacuated from the ing gear extended.

airplane to the ground under simulated (6) Tailcone. This type is an aft exit emergency conditions within 90 sec- from the passenger compartment onds. Compliance with this require- through the pressure shell and through ment must be shown by actual dem- an openable cone of the fuselage aft of onstration using the test criteria out- the pressure shell. The means of open- lined in appendix J of this part unless ing the tailcone must be simple and ob- the Administrator finds that a com- vious and must employ a single oper- bination of analysis and testing will ation.

provide data equivalent to that which (7) Type A. This type is a floor-level would be obtained by actual dem- exit with a rectangular opening of not onstration.

less than 42 inches wide by 72 inches (d)–(e) [Reserved] high, with corner radii not greater [Doc. No. 24344, 55 FR 29781, July 20, 1990] than seven inches.

(8) Type B. This type is a floor-level § 25.807 Emergency exits.

exit with a rectangular opening of not (a) Type. For the purpose of this part, less than 32 inches wide by 72 inches the types of exits are defined as fol- high, with corner radii not greater lows: than six inches.

(1) Type I. This type is a floor-level (9) Type C. This type is a floor-level exit with a rectangular opening of not exit with a rectangular opening of not less than 24 inches wide by 48 inches less than 30 inches wide by 48 inches high, with corner radii not greater high, with corner radii not greater than eight inches. than 10 inches.

(2) Type II. This type is a rectangular (b) Step down distance. Step down dis- opening of not less than 20 inches wide tance, as used in this section, means by 44 inches high, with corner radii not the actual distance between the bot- greater than seven inches. Type II exits tom of the required opening and a usa- must be floor-level exits unless located ble foot hold, extending out from the over the wing, in which case they must fuselage, that is large enough to be ef- not have a step-up inside the airplane fective without searching by sight or of more than 10 inches nor a step-down feel.

outside the airplane of more than 17 (c) Over-sized exits. Openings larger inches. than those specified in this section, (3) Type III. This type is a rectan- whether or not of rectangular shape, gular opening of not less than 20 inches may be used if the specified rectan- wide by 36 inches high with corner gular opening can be inscribed within 14 CFR Ch. I (1–1–25 Edition) § 25.807 the opening and the base of the in- Type III 35 Type IV 9 scribed rectangular opening meets the specified step-up and step-down (1) For a passenger seating configura- heights.

tion of 1 to 9 seats, there must be at (d) Asymmetry. Exits of an exit pair least one Type IV or larger overwing need not be diametrically opposite exit in each side of the fuselage or, if each other nor of the same size; how- overwing exits are not provided, at ever, the number of passenger seats least one exit in each side that meets permitted under paragraph (g) of this the minimum dimensions of a Type III section is based on the smaller of the exit.

two exits.

(2) For a passenger seating configura- (e) Uniformity. Exits must be distrib- tion of more than 9 seats, each exit uted as uniformly as practical, taking must be a Type III or larger exit.

into account passenger seat distribu- (3) For a passenger seating configura- tion.

tion of 10 to 19 seats, there must be at (f) Location. (1) Each required pas- least one Type III or larger exit in each senger emergency exit must be acces- side of the fuselage.

sible to the passengers and located (4) For a passenger seating configura- where it will afford the most effective tion of 20 to 40 seats, there must be at means of passenger evacuation.

least two exits, one of which must be a (2) If only one floor-level exit per side Type II or larger exit, in each side of is prescribed, and the airplane does not the fuselage.

have a tailcone or ventral emergency (5) For a passenger seating configura- exit, the floor-level exits must be in tion of 41 to 110 seats, there must be at the rearward part of the passenger least two exits, one of which must be a compartment unless another location Type I or larger exit, in each side of affords a more effective means of pas- the fuselage.

senger evacuation.

(6) For a passenger seating configura- (3) If more than one floor-level exit tion of more than 110 seats, the emer- per side is prescribed, and the airplane gency exits in each side of the fuselage does not have a combination cargo and must include at least two Type I or passenger configuration, at least one larger exits.

floor-level exit must be located in each (7) The combined maximum number side near each end of the cabin.

of passenger seats permitted for all (4) For an airplane that is required to Type III exits is 70, and the combined have more than one passenger emer- maximum number of passenger seats gency exit for each side of the fuselage, permitted for two Type III exits in no passenger emergency exit shall be each side of the fuselage that are sepa- more than 60 feet from any adjacent rated by fewer than three passenger passenger emergency exit on the same seat rows is 65.

side of the same deck of the fuselage, (8) If a Type A, Type B, or Type C as measured parallel to the airplane’s exit is installed, there must be at least longitudinal axis between the nearest two Type C or larger exits in each side exit edges.

of the fuselage.

(g) Type and number required. The (9) If a passenger ventral or tailcone maximum number of passenger seats exit is installed and that exit provides permitted depends on the type and at least the same rate of egress as a number of exits installed in each side Type III exit with the airplane in the of the fuselage. Except as further re- most adverse exit opening condition stricted in paragraphs (g)(1) through that would result from the collapse of (g)(9) of this section, the maximum one or more legs of the landing gear, an number of passenger seats permitted increase in the passenger seating con- for each exit of a specific type installed figuration is permitted as follows: in each side of the fuselage is as fol- (i) For a ventral exit, 12 additional lows: passenger seats.

Type A 110 (ii) For a tailcone exit incorporating Type B 75 a floor level opening of not less than 20 Type C 55 Type I 45 inches wide by 60 inches high, with cor- Type II 40 ner radii not greater than seven inches, Federal Aviation Administration, DOT § 25.809 in the pressure shell and incorporating (3) If it is impractical to locate side an approved assist means in accordance exits above the waterline, the side with § 25.810(a), 25 additional passenger exits must be replaced by an equal seats. number of readily accessible overhead hatches of not less than the dimensions (iii) For a tailcone exit incorporating an opening in the pressure shell which of a Type III exit, except that for air- is at least equivalent to a Type III planes with a passenger configuration emergency exit with respect to dimen- of 35 or fewer seats, excluding pilot seats, the two required Type III side sions, step-up and step-down distance, exits need be replaced by only one and with the top of the opening not overhead hatch.

less than 56 inches from the passenger compartment floor, 15 additional pas- (j) Flightcrew emergency exits. For air- senger seats. planes in which the proximity of pas- senger emergency exits to the (h) Other exits. The following exits flightcrew area does not offer a conven- also must meet the applicable emer- ient and readily accessible means of gency exit requirements of §§ 25.809 evacuation of the flightcrew, and for through 25.812, and must be readily ac- all airplanes having a passenger seat- cessible: ing capacity greater than 20, flightcrew (1) Each emergency exit in the pas- exits shall be located in the flightcrew senger compartment in excess of the area. Such exits shall be of sufficient minimum number of required emer- size and so located as to permit rapid gency exits.

evacuation by the crew. One exit shall (2) Any other floor-level door or exit be provided on each side of the air- that is accessible from the passenger plane; or, alternatively, a top hatch compartment and is as large or larger shall be provided. Each exit must en- than a Type II exit, but less than 46 compass an unobstructed rectangular inches wide.

opening of at least 19 by 20 inches un- (3) Any other ventral or tail cone less satisfactory exit utility can be passenger exit.

demonstrated by a typical crew- (i) Ditching emergency exits for pas- member.

sengers. Whether or not ditching cer- tification is requested, ditching emer- [Amdt. 25–72, 55 FR 29781, July 20, 1990, as gency exits must be provided in accord- amended by Amdt. 25–88, 61 FR 57956, Nov. 8, ance with the following requirements, 1996; 62 FR 1817, Jan. 13, 1997; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998; 63 FR 12862, Mar. 16, unless the emergency exits required by 1998; Amdt. 25–114, 69 FR 24502, May 3, 2004] paragraph (g) of this section already meet them: § 25.809 Emergency exit arrangement.

(1) For airplanes that have a pas- (a) Each emergency exit, including senger seating configuration of nine or fewer seats, excluding pilot seats, one each flightcrew emergency exit, must exit above the waterline in each side of be a moveable door or hatch in the ex- the airplane, meeting at least the di- ternal walls of the fuselage, allowing mensions of a Type IV exit. an unobstructed opening to the out- side. In addition, each emergency exit (2) For airplanes that have a pas- must have means to permit viewing of senger seating configuration of 10 of the conditions outside the exit when more seats, excluding pilot seats, one the exit is closed. The viewing means exit above the waterline in a side of the may be on or adjacent to the exit pro- airplane, meeting at least the dimen- vided no obstructions exist between the sions of a Type III exit for each unit (or part of a unit) of 35 passenger seats, exit and the viewing means. Means but no less than two such exits in the must also be provided to permit view- passenger cabin, with one on each side ing of the likely areas of evacuee of the airplane. The passenger seat/ ground contact. The likely areas of exit ratio may be increased through evacuee ground contact must be the use of larger exits, or other means, viewable during all lighting conditions provided it is shown that the evacu- with the landing gear extended as well ation capability during ditching has as in all conditions of landing gear col- been improved accordingly. lapse.

14 CFR Ch. I (1–1–25 Edition) § 25.810 (b) Each emergency exit must be (h) When required by the operating openable from the inside and the out- rules for any large passenger-carrying side except that sliding window emer- turbojet-powered airplane, each ven- gency exits in the flight crew area need tral exit and tailcone exit must be— not be openable from the outside if (1) Designed and constructed so that other approved exits are convenient it cannot be opened during flight; and and readily accessible to the flight (2) Marked with a placard readable crew area. Each emergency exit must from a distance of 30 inches and in- be capable of being opened, when there stalled at a conspicuous location near is no fuselage deformation— the means of opening the exit, stating (1) With the airplane in the normal that the exit has been designed and ground attitude and in each of the atti- constructed so that it cannot be opened tudes corresponding to collapse of one during flight.

or more legs of the landing gear; and (i) Each emergency exit must have a (2) Within 10 seconds measured from means to retain the exit in the open the time when the opening means is ac- position, once the exit is opened in an tuated to the time when the exit is emergency. The means must not re- fully opened.

quire separate action to engage when (3) Even though persons may be the exit is opened, and must require crowded against the door on the inside positive action to disengage.

of the airplane.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) The means of opening emergency amended by Amdt. 25–15, 32 FR 13264, Sept.

exits must be simple and obvious; may 20, 1967; Amdt. 25–32, 37 FR 3970, Feb. 24, 1972; not require exceptional effort; and Amdt. 25–34, 37 FR 25355, Nov. 30, 1972; Amdt.

must be arranged and marked so that 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. 25–47, it can be readily located and operated, 44 FR 61325, Oct. 25, 1979; Amdt. 25–72, 55 FR even in darkness. Internal exit-opening 29782, July 20, 1990; Amdt. 25–114, 69 FR 24502, May 3, 2004; Amdt. 25–116, 69 FR 62788, Oct. 27, means involving sequence operations 2004] (such as operation of two handles or latches, or the release of safety § 25.810 Emergency egress assist catches) may be used for flightcrew means and escape routes.

emergency exits if it can be reasonably (a) Each non over-wing Type A, Type established that these means are sim- B or Type C exit, and any other non ple and obvious to crewmembers over-wing landplane emergency exit trained in their use.

more than 6 feet from the ground with (d) If a single power-boost or single the airplane on the ground and the power-operated system is the primary landing gear extended, must have an system for operating more than one approved means to assist the occupants exit in an emergency, each exit must in descending to the ground.

be capable of meeting the requirements (1) The assisting means for each pas- of paragraph (b) of this section in the senger emergency exit must be a self- event of failure of the primary system.

supporting slide or equivalent; and, in Manual operation of the exit (after the case of Type A or Type B exits, it failure of the primary system) is ac- must be capable of carrying simulta- ceptable.

neously two parallel lines of evacuees.

(e) Each emergency exit must be In addition, the assisting means must shown by tests, or by a combination of be designed to meet the following re- analysis and tests, to meet the require- quirements— ments of paragraphs (b) and (c) of this section. (i) It must be automatically deployed (f) Each door must be located where and deployment must begin during the persons using them will not be endan- interval between the time the exit gered by the propellers when appro- opening means is actuated from inside priate operating procedures are used. the airplane and the time the exit is (g) There must be provisions to mini- fully opened. However, each passenger mize the probability of jamming of the emergency exit which is also a pas- emergency exits resulting from fuse- senger entrance door or a service door lage deformation in a minor crash must be provided with means to pre- landing. vent deployment of the assisting means Federal Aviation Administration, DOT § 25.810 when it is opened from either the in- (b) Assist means from the cabin to side or the outside under non- the wing are required for each type A emergency conditions for normal use. or Type B exit located above the wing (ii) Except for assisting means in- and having a stepdown unless the exit stalled at Type C exits, it must be without an assist-means can be shown automatically erected within 6 seconds to have a rate of passenger egress at after deployment is begun. Assisting least equal to that of the same type of means installed at Type C exits must non over-wing exit. If an assist means be automatically erected within 10 sec- is required, it must be automatically onds from the time the opening means deployed and automatically erected of the exit is actuated.

concurrent with the opening of the (iii) It must be of such length after exit. In the case of assist means in- full deployment that the lower end is stalled at Type C exits, it must be self- self-supporting on the ground and pro- supporting within 10 seconds from the vides safe evacuation of occupants to time the opening means of the exits is the ground after collapse of one or actuated. For all other exit types, it more legs of the landing gear.

must be self-supporting 6 seconds after (iv) It must have the capability, in deployment is begun.

25-knot winds directed from the most (c) An escape route must be estab- critical angle, to deploy and, with the lished from each overwing emergency assistance of only one person, to re- exit, and (except for flap surfaces suit- main usable after full deployment to able as slides) covered with a slip re- evacuate occupants safely to the sistant surface. Except where a means ground.

for channeling the flow of evacuees is (v) For each system installation provided— (mockup or airplane installed), five (1) The escape route from each Type consecutive deployment and inflation A or Type B passenger emergency exit, tests must be conducted (per exit) or any common escape route from two without failure, and at least three tests Type III passenger emergency exits, of each such five-test series must be must be at least 42 inches wide; that conducted using a single representative from any other passenger emergency sample of the device. The sample de- exit must be at least 24 inches wide; vices must be deployed and inflated by and the system’s primary means after (2) The escape route surface must being subjected to the inertia forces have a reflectance of at least 80 per- specified in § 25.561(b). If any part of the cent, and must be defined by markings system fails or does not function prop- with a surface-to-marking contrast erly during the required tests, the ratio of at least 5:1.

cause of the failure or malfunction (d) Means must be provided to assist must be corrected by positive means evacuees to reach the ground for all and after that, the full series of five Type C exits located over the wing and, consecutive deployment and inflation if the place on the airplane structure tests must be conducted without fail- at which the escape route required in ure.

paragraph (c) of this section termi- (2) The assisting means for flightcrew nates is more than 6 feet from the emergency exits may be a rope or any ground with the airplane on the ground other means demonstrated to be suit- and the landing gear extended, for all able for the purpose. If the assisting other exit types.

means is a rope, or an approved device (1) If the escape route is over the equivalent to a rope, it must be— flap, the height of the terminal edge (i) Attached to the fuselage structure must be measured with the flap in the at or above the top of the emergency takeoff or landing position, whichever exit opening, or, for a device at a pi- is higher from the ground.

lot’s emergency exit window, at an- other approved location if the stowed (2) The assisting means must be usa- device, or its attachment, would reduce ble and self-supporting with one or the pilot’s view in flight; more landing gear legs collapsed and (ii) Able (with its attachment) to under a 25-knot wind directed from the withstand a 400-pound static load. most critical angle.

14 CFR Ch. I (1–1–25 Edition) § 25.811 (3) The assisting means provided for each passenger emergency exit, or at each escape route leading from a Type another overhead location if it is more A or B emergency exit must be capable practical because of low headroom, ex- of carrying simultaneously two par- cept that one sign may serve more allel lines of evacuees; and, the assist- than one exit if each exit can be seen ing means leading from any other exit readily from the sign; type must be capable of carrying as (2) A passenger emergency exit mark- many parallel lines of evacuees as ing sign next to each passenger emer- there are required escape routes. gency exit, except that one sign may (4) The assisting means provided for serve two such exits if they both can be each escape route leading from a Type seen readily from the sign; and C exit must be automatically erected (3) A sign on each bulkhead or divider within 10 seconds from the time the that prevents fore and aft vision along opening means of the exit is actuated, the passenger cabin to indicate emer- and that provided for the escape route gency exits beyond and obscured by the leading from any other exit type must bulkhead or divider, except that if this be automatically erected within 10 sec- is not possible the sign may be placed onds after actuation of the erection at another appropriate location.

system. (e) The location of the operating han- (e) If an integral stair is installed in dle and instructions for opening exits a passenger entry door that is qualified from the inside of the airplane must be as a passenger emergency exit, the shown in the following manner: stair must be designed so that, under (1) Each passenger emergency exit the following conditions, the effective- must have, on or near the exit, a mark- ness of passenger emergency egress will ing that is readable from a distance of not be impaired: 30 inches.

(1) The door, integral stair, and oper- (2) Each Type A, Type B, Type C or ating mechanism have been subjected Type I passenger emergency exit oper- to the inertia forces specified in ating handle must— § 25.561(b)(3), acting separately relative (i) Be self-illuminated with an initial to the surrounding structure.

brightness of at least 160 micro- (2) The airplane is in the normal lamberts; or ground attitude and in each of the atti- (ii) Be conspicuously located and well tudes corresponding to collapse of one illuminated by the emergency lighting or more legs of the landing gear.

even in conditions of occupant crowd- ing at the exit.

[Amdt. 25–72, 55 FR 29782, July 20, 1990, as (3) [Reserved] amended by Amdt. 25–88, 61 FR 57958, Nov. 8, (4) Each Type A, Type B, Type C, 1996; 62 FR 1817, Jan. 13, 1997; Amdt. 25–114, 69 FR 24502, May 3, 2004] Type I, or Type II passenger emergency exit with a locking mechanism re- § 25.811 Emergency exit marking.

leased by rotary motion of the handle must be marked— (a) Each passenger emergency exit, (i) With a red arrow, with a shaft at its means of access, and its means of least three-fourths of an inch wide and opening must be conspicuously a head twice the width of the shaft, ex- marked.

(b) The identity and location of each tending along at least 70 degrees of arc passenger emergency exit must be rec- at a radius approximately equal to ognizable from a distance equal to the three-fourths of the handle length.

width of the cabin. (ii) So that the centerline of the exit (c) Means must be provided to assist handle is within ± 1 inch of the pro- the occupants in locating the exits in jected point of the arrow when the han- conditions of dense smoke. dle has reached full travel and has re- (d) The location of each passenger leased the locking mechanism, and emergency exit must be indicated by a (iii) With the word ‘‘open’’ in red let- sign visible to occupants approaching ters 1 inch high, placed horizontally along the main passenger aisle (or near the head of the arrow.

aisles). There must be— (f) Each emergency exit that is re- (1) A passenger emergency exit loca- quired to be openable from the outside, tor sign above the aisle (or aisles) near and its means of opening, must be Federal Aviation Administration, DOT § 25.812 marked on the outside of the airplane. eral cabin illumination, interior light- In addition, the following apply: ing in emergency exit areas, and floor (1) The outside marking for each pas- proximity escape path marking.

senger emergency exit in the side of (2) Exterior emergency lighting.

the fuselage must include a 2-inch col- (b) Emergency exit signs— ored band outlining the exit.

(1) For airplanes that have a pas- (2) Each outside marking including senger seating configuration, excluding the band, must have color contrast to pilot seats, of 10 seats or more must be readily distinguishable from the sur- meet the following requirements: rounding fuselage surface. The contrast (i) Each passenger emergency exit lo- must be such that if the reflectance of cator sign required by § 25.811(d)(1) and the darker color is 15 percent or less, each passenger emergency exit mark- the reflectance of the lighter color ing sign required by § 25.811(d)(2) must must be at least 45 percent. ‘‘Reflec- have red letters at least 1 ⁄2 inches high tance’’ is the ratio of the luminous flux on an illuminated white background, reflected by a body to the luminous and must have an area of at least 21 flux it receives. When the reflectance square inches excluding the letters.

of the darker color is greater than 15 The lighted background-to-letter con- percent, at least a 30-percent difference trast must be at least 10:1. The letter between its reflectance and the reflec- height to stroke-width ratio may not tance of the lighter color must be pro- be more than 7:1 nor less than 6:1.

vided.

These signs must be internally elec- (3) In the case of exists other than trically illuminated with a background those in the side of the fuselage, such brightness of at least 25 foot-lamberts as ventral or tailcone exists, the exter- and a high-to-low background contrast nal means of opening, including in- no greater than 3:1.

structions if applicable, must be con- (ii) Each passenger emergency exit spicuously marked in red, or bright sign required by § 25.811(d)(3) must have chrome yellow if the background color red letters at least 1 ⁄2 inches high on a is such that red is inconspicuous. When white background having an area of at the opening means is located on only least 21 square inches excluding the one side of the fuselage, a conspicuous letters. These signs must be internally marking to that effect must be pro- electrically illuminated or self-illumi- vided on the other side.

nated by other than electrical means (g) Each sign required by paragraph and must have an initial brightness of (d) of this section may use the word at least 400 microlamberts. The colors ‘‘exit’’ in its legend in place of the may be reversed in the case of a sign term ‘‘emergency exit’’.

that is self-illuminated by other than [Amdt. 25–15, 32 FR 13264, Sept. 20, 1967, as electrical means.

amended by Amdt. 25–32, 37 FR 3970, Feb. 24, (2) For airplanes that have a pas- 1972; Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; 43 senger seating configuration, excluding FR 52495, Nov. 13, 1978; Amdt. 25–79, 58 FR pilot seats, of nine seats or less, that 45229, Aug. 26, 1993; Amdt. 25–88, 61 FR 57958, Nov. 8, 1996] are required by § 25.811(d)(1), (2), and (3) must have red letters at least 1 inch § 25.812 Emergency lighting.

high on a white background at least 2 inches high. These signs may be inter- (a) An emergency lighting system, nally electrically illuminated, or self- independent of the main lighting sys- illuminated by other than electrical tem, must be installed. However, the means, with an initial brightness of at sources of general cabin illumination least 160 microlamberts. The colors may be common to both the emergency may be reversed in the case of a sign and the main lighting systems if the that is self-illuminated by other than power supply to the emergency light- electrical means.

ing system is independent of the power supply to the main lighting system. (c) General illumination in the pas- The emergency lighting system must senger cabin must be provided so that include: when measured along the centerline of (1) Illuminated emergency exit mark- main passenger aisle(s), and cross ing and locating signs, sources of gen- aisle(s) between main aisles, at seat 14 CFR Ch. I (1–1–25 Edition) § 25.812 arm-rest height and at 40-inch inter- (3) The cockpit control device must vals, the average illumination is not have an ‘‘on,’’ ‘‘off,’’ and ‘‘armed’’ posi- less than 0.05 foot-candle and the illu- tion so that when armed in the cockpit mination at each 40-inch interval is not or turned on at either the cockpit or less than 0.01 foot-candle. A main pas- flight attendant station the lights will either light or remain lighted upon senger aisle(s) is considered to extend interruption (except an interruption along the fuselage from the most for- caused by a transverse vertical separa- ward passenger emergency exit or tion of the fuselage during crash land- cabin occupant seat, whichever is far- ing) of the airplane’s normal electric ther forward, to the most rearward pas- power. There must be a means to safe- senger emergency exit or cabin occu- guard against inadvertent operation of pant seat, whichever is farther aft.

the control device from the ‘‘armed’’ or (d) The floor of the passageway lead- ‘‘on’’ positions.

ing to each floor-level passenger emer- (g) Exterior emergency lighting must gency exit, between the main aisles be provided as follows: and the exit openings, must be pro- (1) At each overwing emergency exit vided with illumination that is not less the illumination must be— than 0.02 foot-candle measured along a (i) Not less than 0.03 foot-candle line that is within 6 inches of and par- (measured normal to the direction of allel to the floor and is centered on the the incident light) on a 2-square-foot passenger evacuation path.

area where an evacuee is likely to (e) Floor proximity emergency es- make his first step outside the cabin; cape path marking must provide emer- (ii) Not less than 0.05 foot-candle gency evacuation guidance for pas- (measured normal to the direction of sengers when all sources of illumina- the incident light) for a minimum tion more than 4 feet above the cabin width of 42 inches for a Type A aisle floor are totally obscured. In the overwing emergency exit and two feet dark of the night, the floor proximity for all other overwing emergency exits emergency escape path marking must along the 30 percent of the slip-resist- enable each passenger to— ant portion of the escape route re- (1) After leaving the passenger seat, quired in § 25.810(c) that is farthest visually identify the emergency escape from the exit; and path along the cabin aisle floor to the (iii) Not less than 0.03 foot-candle on first exits or pair of exits forward and the ground surface with the landing aft of the seat; and gear extended (measured normal to the (2) Readily identify each exit from direction of the incident light) where the emergency escape path by ref- an evacuee using the established escape erence only to markings and visual fea- route would normally make first con- tures not more than 4 feet above the tact with the ground.

cabin floor.

(2) At each non-overwing emergency (f) Except for subsystems provided in exit not required by § 25.810(a) to have accordance with paragraph (h) of this descent assist means the illumination section that serve no more than one as- must be not less than 0.03 foot-candle sist means, are independent of the air- (measured normal to the direction of plane’s main emergency lighting sys- the incident light) on the ground sur- tem, and are automatically activated face with the landing gear extended when the assist means is erected, the where an evacuee is likely to make emergency lighting system must be de- first contact with the ground outside signed as follows.

the cabin.

(1) The lights must be operable (h) The means required in manually from the flight crew station §§ 25.810(a)(1) and (d) to assist the occu- and from a point in the passenger com- pants in descending to the ground must partment that is readily accessible to a be illuminated so that the erected as- normal flight attendant seat.

sist means is visible from the airplane.

(2) There must be a flight crew warn- (1) If the assist means is illuminated ing light which illuminates when power by exterior emergency lighting, it is on in the airplane and the emergency must provide illumination of not less lighting control device is not armed. than 0.03 foot-candle (measured normal Federal Aviation Administration, DOT § 25.813 to the direction of the incident light) mains operative exclusive of those that at the ground end of the erected assist are directly damaged by the separa- means where an evacuee using the es- tion; and tablished escape route would normally (3) At least one required exterior make first contact with the ground, emergency light for each side of the with the airplane in each of the atti- airplane remains operative exclusive of tudes corresponding to the collapse of those that are directly damaged by the one or more legs of the landing gear.

separation.

(2) If the emergency lighting sub- [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967, as system illuminating the assist means amended by Amdt. 25–28, 36 FR 16899, Aug. 26, serves no other assist means, is inde- 1971; Amdt. 25–32, 37 FR 3971, Feb. 24, 1972; pendent of the airplane’s main emer- Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt.

gency lighting system, and is auto- 25–58, 49 FR 43186, Oct. 26, 1984; Amdt. 25–88, matically activated when the assist 61 FR 57958, Nov. 8, 1996; Amdt. 25–116, 69 FR means is erected, the lighting provi- 62788, Oct. 27, 2004; Amdt. 25–128, 74 FR 25645, sions— May 29, 2009] (i) May not be adversely affected by stowage; and § 25.813 Emergency exit access.

(ii) Must provide illumination of not Each required emergency exit must less than 0.03 foot-candle (measured be accessible to the passengers and lo- normal to the direction of incident cated where it will afford an effective light) at the ground and of the erected means of evacuation. Emergency exit assist means where an evacuee would distribution must be as uniform as normally make first contact with the practical, taking passenger distribu- ground, with the airplane in each of tion into account; however, the size the attitudes corresponding to the col- and location of exits on both sides of lapse of one or more legs of the landing the cabin need not be symmetrical. If gear.

only one floor level exit per side is pre- (i) The energy supply to each emer- scribed, and the airplane does not have gency lighting unit must provide the a tailcone or ventral emergency exit, required level of illumination for at the floor level exit must be in the rear- least 10 minutes at the critical ambient ward part of the passenger compart- conditions after emergency landing.

ment, unless another location affords a (j) If storage batteries are used as the more effective means of passenger energy supply for the emergency light- evacuation. Where more than one floor ing system, they may be recharged level exit per side is prescribed, at from the airplane’s main electric power least one floor level exit per side must system: Provided, That, the charging be located near each end of the cabin, circuit is designed to preclude inad- except that this provision does not vertent battery discharge into charg- apply to combination cargo/passenger ing circuit faults.

configurations. In addition— (k) Components of the emergency (a) There must be a passageway lead- lighting system, including batteries, ing from the nearest main aisle to each wiring relays, lamps, and switches Type A, Type B, Type C, Type I, or must be capable of normal operation Type II emergency exit and between in- after having been subjected to the iner- dividual passenger areas. Each passage- tia forces listed in § 25.561(b).

way leading to a Type A or Type B exit (l) The emergency lighting system must be unobstructed and at least 36 must be designed so that after any sin- inches wide. Passageways between indi- gle transverse vertical separation of vidual passenger areas and those lead- the fuselage during crash landing— ing to Type I, Type II, or Type C emer- (1) Not more than 25 percent of all gency exits must be unobstructed and electrically illuminated emergency at least 20 inches wide. Unless there lights required by this section are ren- dered inoperative, in addition to the are two or more main aisles, each Type lights that are directly damaged by the A or B exit must be located so that separation; there is passenger flow along the main (2) Each electrically illuminated exit aisle to that exit from both the forward sign required under § 25.811(d)(2) re- and aft directions. If two or more main 14 CFR Ch. I (1–1–25 Edition) § 25.813 aisles are provided, there must be un- There must be access from the nearest obstructed cross-aisles at least 20 aisle to each exit. In addition, for each inches wide between main aisles. There Type III exit in an airplane that has a must be— passenger seating configuration of 60 or (1) A cross-aisle which leads directly more— to each passageway between the near- (i) Except as provided in paragraph est main aisle and a Type A or B exit; (c)(1)(ii), the access must be provided and by an unobstructed passageway that is (2) A cross-aisle which leads to the at least 10 inches in width for interior immediate vicinity of each passageway arrangements in which the adjacent between the nearest main aisle and a seat rows on the exit side of the aisle Type 1, Type II, or Type III exit; except contain no more than two seats, or 20 that when two Type III exits are lo- inches in width for interior arrange- cated within three passenger rows of ments in which those rows contain each other, a single cross-aisle may be three seats. The width of the passage- used if it leads to the vicinity between way must be measured with adjacent the passageways from the nearest main seats adjusted to their most adverse aisle to each exit.

position. The centerline of the required (b) Adequate space to allow crew- passageway width must not be dis- member(s) to assist in the evacuation placed more than 5 inches horizontally of passengers must be provided as fol- from that of the exit.

lows: (ii) In lieu of one 10- or 20-inch pas- (1) Each assist space must be a rec- sageway, there may be two passage- tangle on the floor, of sufficient size to ways, between seat rows only, that enable a crewmember, standing erect, must be at least 6 inches in width and to effectively assist evacuees. The as- lead to an unobstructed space adjacent sist space must not reduce the unob- to each exit. (Adjacent exits must not structed width of the passageway below share a common passageway.) The that required for the exit.

width of the passageways must be (2) For each Type A or B exit, assist measured with adjacent seats adjusted space must be provided at each side of to their most adverse position. The un- the exit regardless of whether an assist obstructed space adjacent to the exit means is required by § 25.810(a).

must extend vertically from the floor (3) For each Type C, I or II exit in- to the ceiling (or bottom of sidewall stalled in an airplane with seating for stowage bins), inboard from the exit for more than 80 passengers, an assist a distance not less than the width of space must be provided at one side of the narrowest passenger seat installed the passageway regardless of whether on the airplane, and from the forward an assist means is required by edge of the forward passageway to the § 25.810(a).

aft edge of the aft passageway. The exit (4) For each Type C, I or II exit, an opening must be totally within the fore assist space must be provided at one and aft bounds of the unobstructed side of the passageway if an assist space.

means is required by § 25.810(a).

(2) In addition to the access— (5) For any tailcone exit that quali- (i) For airplanes that have a pas- fies for 25 additional passenger seats senger seating configuration of 20 or under the provisions of § 25.807(g)(9)(ii), more, the projected opening of the exit an assist space must be provided, if an provided must not be obstructed and assist means is required by § 25.810(a).

there must be no interference in open- (6) There must be a handle, or han- ing the exit by seats, berths, or other dles, at each assist space, located to protrusions (including any seatback in enable the crewmember to steady him- the most adverse position) for a dis- self or herself: tance from that exit not less than the (i) While manually activating the as- width of the narrowest passenger seat sist means (where applicable) and, installed on the airplane.

(ii) While assisting passengers during an evacuation. (ii) For airplanes that have a pas- (c) The following must be provided senger seating configuration of 19 or for each Type III or Type IV exit—(1) fewer, there may be minor obstructions Federal Aviation Administration, DOT § 25.819 in this region, if there are compen- Minimum passenger aisle width (inches) sating factors to maintain the effec- Passenger seating capacity tiveness of the exit. Less than 25 in. and 25 in. from more from (3) For each Type III exit, regardless floor floor of the passenger capacity of the air- 10 or less ....................................... 12 15 plane in which it is installed, there 11 through 19 ................................ 12 20 must be placards that— 20 or more ..................................... 15 20 (i) Are readable by all persons seated A narrower width not less than 9 inches may be approved adjacent to and facing a passageway to when substantiated by tests found necessary by the the exit; Administrator.

(ii) Accurately state or illustrate the [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967, as proper method of opening the exit, in- amended by Amdt. 25–38, 41 FR 55466, Dec. 20, cluding the use of handholds; and 1976] (iii) If the exit is a removable hatch, state the weight of the hatch and indi- § 25.817 Maximum number of seats cate an appropriate location to place abreast.

the hatch after removal.

On airplanes having only one pas- (d) If it is necessary to pass through senger aisle, no more than three seats a passageway between passenger com- abreast may be placed on each side of partments to reach any required emer- the aisle in any one row.

gency exit from any seat in the pas- senger cabin, the passageway must be [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967] unobstructed. However, curtains may § 25.819 Lower deck service compart- be used if they allow free entry ments (including galleys).

through the passageway.

(e) No door may be installed between For airplanes with a service compart- any passenger seat that is occupiable ment located below the main deck, for takeoff and landing and any pas- which may be occupied during taxi or senger emergency exit, such that the flight but not during takeoff or land- ing, the following apply: door crosses any egress path (including aisles, crossaisles and passageways). (a) There must be at least two emer- (f) If it is necessary to pass through a gency evacuation routes, one at each end of each lower deck service com- doorway separating any crewmember partment or two having sufficient sepa- seat (except those seats on the ration within each compartment, flightdeck), occupiable for takeoff and which could be used by each occupant landing, from any emergency exit, the of the lower deck service compartment door must have a means to latch it in to rapidly evacuate to the main deck the open position. The latching means under normal and emergency lighting must be able to withstand the loads conditions. The routes must provide for imposed upon it when the door is sub- the evacuation of incapacitated per- jected to the ultimate inertia forces, sons, with assistance. The use of the relative to the surrounding structure, evacuation routes may not be depend- listed in § 25.561(b).

ent on any powered device. The routes [Amdt. 25–1, 30 FR 3204, Mar. 9, 1965, as must be designed to minimize the pos- amended by Amdt. 25–15, 32 FR 13265, Sept.

sibility of blockage which might result 20, 1967; Amdt. 25–32, 37 FR 3971, Feb. 24, 1972; from fire, mechanical or structural Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt.

failure, or persons standing on top of or 25–72, 55 FR 29783, July 20, 1990; Amdt. 25–76, against the escape routes. In the event 57 FR 19244, May 4, 1992; Amdt. 25–76, 57 FR the airplane’s main power system or 29120, June 30, 1992; Amdt. 25–88, 61 FR 57958, Nov. 8, 1996; Amdt. 25–116, 69 FR 62788, Oct. compartment main lighting system 27, 2004; Amdt. 25–128, 74 FR 25645, May 29, should fail, emergency illumination for 2009] each lower deck service compartment must be automatically provided.

§ 25.815 Width of aisle.

(b) There must be a means for two- The passenger aisle width at any way voice communication between the point between seats must equal or ex- flight deck and each lower deck service ceed the values in the following table: compartment, which remains available 14 CFR Ch. I (1–1–25 Edition) § 25.820 following loss of normal electrical capable of being unlocked from the power generating system. outside without the aid of special tools.

(c) There must be an aural emer- [Doc. No. 2003–14193, 69 FR 24502, May 3, 2004] gency alarm system, audible during V ENTILATION AND H EATING normal and emergency conditions, to enable crewmembers on the flight deck § 25.831 Ventilation.

and at each required floor level emer- (a) Under normal operating condi- gency exit to alert occupants of each tions and in the event of any probable lower deck service compartment of an failure conditions of any system which emergency situation.

would adversely affect the ventilating (d) There must be a means, readily air, the ventilation system must be de- detectable by occupants of each lower signed to provide a sufficient amount deck service compartment, that indi- of uncontaminated air to enable the cates when seat belts should be fas- crewmembers to perform their duties tened.

without undue discomfort or fatigue (e) If a public address system is in- and to provide reasonable passenger stalled in the airplane, speakers must comfort. For normal operating condi- be provided in each lower deck service tions, the ventilation system must be compartment.

designed to provide each occupant with (f) For each occupant permitted in a an airflow containing at least 0.55 lower deck service compartment, there pounds of fresh air per minute.

must be a forward or aft facing seat (b) Crew and passenger compartment which meets the requirements of air must be free from harmful or haz- ardous concentrations of gases or va- § 25.785(d), and must be able to with- pors. In meeting this requirement, the stand maximum flight loads when oc- following apply: cupied.

(1) Carbon monoxide concentrations (g) For each powered lift system in- in excess of 1 part in 20,000 parts of air stalled between a lower deck service are considered hazardous. For test pur- compartment and the main deck for poses, any acceptable carbon monoxide the carriage of persons or equipment, detection method may be used.

or both, the system must meet the fol- (2) Carbon dioxide concentration dur- lowing requirements: ing flight must be shown not to exceed (1) Each lift control switch outside 0.5 percent by volume (sea level equiva- the lift, except emergency stop but- lent) in compartments normally occu- tons, must be designed to prevent the pied by passengers or crewmembers.

activation of the life if the lift door, or (c) There must be provisions made to the hatch required by paragraph (g)(3) ensure that the conditions prescribed of this section, or both are open.

in paragraph (b) of this section are met (2) An emergency stop button, that after reasonably probable failures or when activated will immediately stop malfunctioning of the ventilating, the lift, must be installed within the heating, pressurization, or other sys- lift and at each entrance to the lift. tems and equipment.

(d) If accumulation of hazardous (3) There must be a hatch capable of quantities of smoke in the cockpit area being used for evacuating persons from is reasonably probable, smoke evacu- the lift that is openable from inside ation must be readily accomplished, and outside the lift without tools, with starting with full pressurization and the lift in any position.

without depressurizing beyond safe [Amdt. 25–53, 45 FR 41593, June 19, 1980; 45 FR limits.

43154, June 26, 1980; Amdt. 25–110; 68 FR 36883, (e) Except as provided in paragraph June 19, 2003] (f) of this section, means must be pro- vided to enable the occupants of the § 25.820 Lavatory doors.

following compartments and areas to All lavatory doors must be designed control the temperature and quantity to preclude anyone from becoming of ventilating air supplied to their trapped inside the lavatory. If a lock- compartment or area independently of ing mechanism is installed, it must be the temperature and quantity of air Federal Aviation Administration, DOT § 25.832 supplied to other compartments and (1) The total volume of the flight areas: crew and passenger compartments is (1) The flight crew compartment. 800 cubic feet or less.

(2) Crewmember compartments and (2) The air inlets and passages for air areas other than the flight crew com- to flow between flight crew and pas- partment unless the crewmember com- senger compartments are arranged to partment or area is ventilated by air provide compartment temperatures interchange with other compartments within 5 degrees F. of each other and or areas under all operating conditions. adequate ventilation to occupants in (f) Means to enable the flight crew to both compartments.

control the temperature and quantity (3) The temperature and ventilation of ventilating air supplied to the flight controls are accessible to the flight crew compartment independently of crew.

the temperature and quantity of ven- (g) The exposure time at any given tilating air supplied to other compart- temperature must not exceed the val- ments are not required if all of the fol- ues shown in the following graph after lowing conditions are met: any improbable failure condition.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. 25–87, 61 FR 28695, June 5, 1996; Amdt. 25–89, 61 FR 63956, Dec. 2, 1996] (c) Compliance with this section § 25.832 Cabin ozone concentration.

must be shown by analysis or tests (a) The airplane cabin ozone con- based on airplane operational proce- centration during flight must be shown dures and performance limitations, not to exceed— that demonstrate that either— (1) 0.25 parts per million by volume, (1) The airplane cannot be operated sea level equivalent, at any time above flight level 320; and at an altitude which would result in (2) 0.1 parts per million by volume, cabin ozone concentrations exceeding sea level equivalent, time-weighted av- the limits prescribed by paragraph (a) erage during any 3-hour interval above of this section; or flight level 270.

(2) The airplane ventilation system, (b) For the purpose of this section, including any ozone control equipment, ‘‘sea level equivalent’’ refers to condi- tions of 25 ° C and 760 millimeters of mercury pressure.

14 CFR Ch. I (1–1–25 Edition) § 25.833 will maintain cabin ozone concentra- (2) Two reverse pressure differential tions at or below the limits prescribed relief valves (or their equivalents) to by paragraph (a) of this section. automatically prevent a negative pres- sure differential that would damage [Amdt. 25–50, 45 FR 3883, Jan. 1, 1980, as the structure. One valve is enough, amended by Amdt. 25–56, 47 FR 58489, Dec. 30, however, if it is of a design that rea- 1982; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] sonably precludes its malfunctioning.

§ 25.833 Combustion heating systems.

(3) A means by which the pressure differential can be rapidly equalized.

Combustion heaters must be ap- proved. (4) An automatic or manual regulator for controlling the intake or exhaust [Amdt. 25–72, 55 FR 29783, July 20, 1990] airflow, or both, for maintaining the P RESSURIZATION required internal pressures and airflow rates.

§ 25.841 Pressurized cabins.

(5) Instruments at the pilot or flight (a) Except as provided in paragraph engineer station to show the pressure (c) of this section, pressurized cabins differential, the cabin pressure alti- and compartments to be occupied must tude, and the rate of change of the be equipped to provide a cabin pressure cabin pressure altitude.

altitude of not more than 8,000 feet (6) Warning indication to the under normal operating conditions.

flightcrew when the safe or preset pres- (1) If certification for operation sure differential or cabin pressure alti- above 25,000 feet is requested, the air- tude limit is exceeded. Appropriate plane must be designed so that occu- warning markings on the cabin pres- pants will not be exposed to cabin pres- sure differential indicator meet the sure altitudes in excess of 15,000 feet warning requirement for pressure dif- after any probable failure condition in ferential limits. An alert meets the the pressurization system except as warning requirement for cabin pressure provided in paragraph (c) of this sec- altitude limits if it warns the tion.

flightcrew when the cabin pressure al- (2) The airplane must be designed so titude exceeds 10,000 feet, except as that occupants will not be exposed to a provided in paragraph (d) of this sec- cabin pressure altitude that exceeds tion.

the following after decompression from (7) A warning placard at the pilot or any failure condition not shown to be flight engineer station if the structure extremely improbable: is not designed for pressure differen- (i) Twenty-five thousand (25,000) feet tials up to the maximum relief valve for more than 2 minutes; or setting in combination with landing (ii) Forty thousand (40,000) feet for loads.

any duration.

(8) The pressure sensors necessary to (3) Fuselage structure, engine and meet the requirements of paragraphs system failures are to be considered in (b)(5) and (b)(6) of this section and evaluating the cabin decompression.

§ 25.1447(c), must be located and the (b) Pressurized cabins must have at sensing system designed so that, in the least the following valves, controls, event of loss of cabin pressure in any and indicators for controlling cabin passenger or crew compartment (in- pressure: cluding upper and lower lobe galleys), (1) Two pressure relief valves to auto- the warning and automatic presen- matically limit the positive pressure tation devices, required by those provi- differential to a predetermined value sions, will be actuated without any at the maximum rate of flow delivered delay that would significantly increase by the pressure source. The combined the hazards resulting from decompres- capacity of the relief valves must be sion.

large enough so that the failure of any one valve would not cause an appre- (c) When operating into or out of air- ciable rise in the pressure differential. ports with elevations at or above 8,000 The pressure differential is positive feet, the cabin pressure altitude in when the internal pressure is greater pressurized cabins and occupied com- than the external. partments may be up to, or greater Federal Aviation Administration, DOT § 25.851 than, the airport elevation by 2,000 pressure vessel for the pressure dif- feet, provided— ferential specified in § 25.365(d).

(1) In the event of probable failure (b) Functional tests. The following conditions of the cabin pressurization functional tests must be performed: system, the cabin pressure altitude (1) Tests of the functioning and ca- must not exceed 15,000 feet, or 2,000 feet pacity of the positive and negative above the airport elevation, whichever pressure differential valves, and of the is higher; and emergency release valve, to stimulate (2) The cabin pressurization system is the effects of closed regulator valves.

(2) Tests of the pressurization system designed to minimize the time in flight to show proper functioning under each that occupants may be exposed to possible condition of pressure, tem- cabin pressure altitudes exceeding 8,000 perature, and moisture, up to the max- feet.

imum altitude for which certification (d) When operating into or out of air- is requested.

ports with elevations at or above 8,000 (3) Flight tests, to show the perform- feet, the cabin pressure high altitude ance of the pressure supply, pressure warning alert may be provided at up to and flow regulators, indicators, and 15,000 feet, or 2,000 feet above the air- warning signals, in steady and stepped plane’s maximum takeoff and landing climbs and descents at rates cor- altitude, whichever is greater, pro- responding to the maximum attainable vided: (1) During landing, the change in within the operating limitations of the cabin pressure high altitude warning airplane, up to the maximum altitude alert may not occur before the start of for which certification is requested.

descent into the high elevation airport (4) Tests of each door and emergency and, following takeoff, the cabin pres- exit, to show that they operate prop- sure high altitude warning alert must erly after being subjected to the flight be reset to 10,000 feet before beginning tests prescribed in paragraph (b)(3) of cruise operation; this section.

(2) Indication is provided to the F IRE P ROTECTION flightcrew that the cabin pressure high altitude warning alert has shifted § 25.851 Fire extinguishers.

above 10,000 feet cabin pressure alti- (a) Hand fire extinguishers. (1) The fol- tude; and lowing minimum number of hand fire (3) Either an alerting system is in- extinguishers must be conveniently lo- stalled that notifies the flightcrew cated and evenly distributed in pas- members on flight deck duty when to senger compartments: don oxygen in accordance with the ap- plicable operating regulations, or a Passenger capacity No. of extinguishers limitation is provided in the airplane flight manual that requires the pilot flying the airplane to don oxygen when the cabin pressure altitude warning has shifted above 10,000 feet, and requires other flightcrew members on flight deck duty to monitor the cabin pres- sure and utilize oxygen in accordance with the applicable operating regula- (2) At least one hand fire extin- tions.

guisher must be conveniently located in the pilot compartment.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (3) At least one readily accessible amended by Amdt. 25–38, 41 FR 55466, Dec. 20, hand fire extinguisher must be avail- 1976; Amdt. 25–87, 61 FR 28696, June 5, 1996; Amdt. No. 25–151, 88 FR 39160, June 15, 2023; able for use in each Class A or Class B 88 FR 44032, July 11, 2023] cargo or baggage compartment and in each Class E or Class F cargo or bag- § 25.843 Tests for pressurized cabins.

gage compartment that is accessible to (a) Strength test. The complete pres- crewmembers in flight.

surized cabin, including doors, win- (4) At least one hand fire extin- dows, and valves, must be tested as a guisher must be located in, or readily 14 CFR Ch. I (1–1–25 Edition) § 25.853 accessible for use in, each galley lo- § 25.853 Compartment interiors.

cated above or below the passenger For each compartment occupied by compartment.

the crew or passengers, the following (5) Each hand fire extinguisher must apply: be approved.

(a) Materials (including finishes or (6) At least one of the required fire decorative surfaces applied to the ma- extinguishers located in the passenger terials) must meet the applicable test criteria prescribed in part I of appendix compartment of an airplane with a pas- F of this part, or other approved equiv- senger capacity of at least 31 and not alent methods, regardless of the pas- more than 60, and at least two of the senger capacity of the airplane.

fire extinguishers located in the pas- (b) [Reserved] senger compartment of an airplane (c) In addition to meeting the re- with a passenger capacity of 61 or more quirements of paragraph (a) of this sec- must contain Halon 1211 tion, seat cushions, except those on (bromochlorodifluoromethane CBrC flight crewmember seats, must meet F ), or equivalent, as the extinguishing the test requirements of part II of ap- agent. The type of extinguishing agent pendix F of this part, or other equiva- used in any other extinguisher required lent methods, regardless of the pas- by this section must be appropriate for senger capacity of the airplane.

the kinds of fires likely to occur where (d) Except as provided in paragraph used.

(e) of this section, the following inte- (7) The quantity of extinguishing rior components of airplanes with pas- agent used in each extinguisher re- senger capacities of 20 or more must quired by this section must be appro- also meet the test requirements of parts IV and V of appendix F of this priate for the kinds of fires likely to part, or other approved equivalent occur where used.

method, in addition to the flamma- (8) Each extinguisher intended for bility requirements prescribed in para- use in a personnel compartment must graph (a) of this section: be designed to minimize the hazard of (1) Interior ceiling and wall panels, toxic gas concentration.

other than lighting lenses and win- (b) Built-in fire extinguishers. If a dows; built-in fire extinguisher is provided— (2) Partitions, other than transparent (1) Each built-in fire extinguishing panels needed to enhance cabin safety; system must be installed so that— (3) Galley structure, including ex- (i) No extinguishing agent likely to posed surfaces of stowed carts and enter personnel compartments will be standard containers and the cavity hazardous to the occupants; and walls that are exposed when a full com- (ii) No discharge of the extinguisher plement of such carts or containers is can cause structural damage.

not carried; and (4) Large cabinets and cabin stowage (2) The capacity of each required compartments, other than underseat built-in fire extinguishing system must stowage compartments for stowing be adequate for any fire likely to occur small items such as magazines and in the compartment where used, con- maps.

sidering the volume of the compart- (e) The interiors of compartments, ment and the ventilation rate. The ca- such as pilot compartments, galleys, pacity of each system is adequate if lavatories, crew rest quarters, cabinets there is sufficient quantity of agent to and stowage compartments, need not extinguish the fire or suppress the fire meet the standards of paragraph (d) of anywhere baggage or cargo is placed this section, provided the interiors of within the cargo compartment for the such compartments are isolated from duration required to land and evacuate the main passenger cabin by doors or the airplane.

equivalent means that would normally be closed during an emergency landing [Amdt. 25–74, 56 FR 15456, Apr. 16, 1991, as condition.

amended by Doc. No. Docket FAA–2014–0001, (f) Smoking is not allowed in lava- Amdt. 25–142, 81 FR 7703, Feb. 16, 2016] tories. If smoking is allowed in any Federal Aviation Administration, DOT § 25.855 area occupied by the crew or pas- rate from, but may be attached to, the sengers, an adequate number of self- airplane structure: contained, removable ashtrays must be (1) Any Class B through Class E cargo provided in designated smoking sec- or baggage compartment, and tions for all seated occupants. (2) Any Class F cargo or baggage (g) Regardless of whether smoking is compartment, unless other means of allowed in any other part of the air- containing a fire and protecting crit- plane, lavatories must have self-con- ical systems and structure are pro- tained, removable ashtrays located vided.

conspicuously on or near the entry side (c) Ceiling and sidewall liner panels of each lavatory door, except that one of Class C cargo or baggage compart- ashtray may serve more than one lava- ments, and ceiling and sidewall liner tory door if the ashtray can be seen panels in Class F cargo or baggage readily from the cabin side of each lav- compartments, if installed to meet the atory served. requirements of paragraph (b)(2) of this (h) Each receptacle used for the dis- section, must meet the test require- posal of flammable waste material ments of part III of appendix F of this must be fully enclosed, constructed of part or other approved equivalent at least fire resistant materials, and methods.

must contain fires likely to occur in it (d) All other materials used in the under normal use. The capability of the construction of the cargo or baggage receptacle to contain those fires under compartment must meet the applicable all probable conditions of wear, mis- test criteria prescribed in part I of ap- alignment, and ventilation expected in pendix F of this part or other approved service must be demonstrated by test. equivalent methods.

(e) No compartment may contain any [Amdt. 25–83, 60 FR 6623, Feb. 2, 1995, as controls, lines, equipment, or acces- amended by Amdt. 25–116, 69 FR 62788, Oct.

sories whose damage or failure would 27, 2004] affect safe operation, unless those items are protected so that— § 25.854 Lavatory fire protection.

(1) They cannot be damaged by the For airplanes with a passenger capac- movement of cargo in the compart- ity of 20 or more: ment, and (a) Each lavatory must be equipped (2) Their breakage or failure will not with a smoke detector system or equiv- create a fire hazard.

alent that provides a warning light in (f) There must be means to prevent the cockpit, or provides a warning cargo or baggage from interfering with light or audible warning in the pas- the functioning of the fire protective senger cabin that would be readily de- features of the compartment.

tected by a flight attendant; and (g) Sources of heat within the com- (b) Each lavatory must be equipped partment must be shielded and insu- with a built-in fire extinguisher for lated to prevent igniting the cargo or each disposal receptacle for towels, baggage.

paper, or waste, located within the lav- (h) Flight tests must be conducted to atory. The extinguisher must be de- show compliance with the provisions of signed to discharge automatically into § 25.857 concerning— each disposal receptacle upon occur- (1) Compartment accessibility, rence of a fire in that receptacle.

(2) The entries of hazardous quan- [Amdt. 25–74, 56 FR 15456, Apr. 16, 1991] tities of smoke or extinguishing agent into compartments occupied by the § 25.855 Cargo or baggage compart- crew or passengers, and ments.

(3) The dissipation of the extin- For each cargo or baggage compart- guishing agent in all Class C compart- ment, the following apply: ments and, if applicable, in any Class F (a) The compartment must meet one compartments.

of the class requirements of § 25.857. (i) During the above tests, it must be (b) Each of the following cargo or shown that no inadvertent operation of baggage compartments, as defined in smoke or fire detectors in any com- § 25.857, must have a liner that is sepa- partment would occur as a result of 14 CFR Ch. I (1–1–25 Edition) § 25.856 fire contained in any other compart- the compartment using a hand fire ex- ment, either during or after extin- tinguisher; guishment, unless the extinguishing (2) When the access provisions are system floods each such compartment being used, no hazardous quantity of simultaneously.

smoke, flames, or extinguishing agent, (j) Cargo or baggage compartment will enter any compartment occupied electrical wiring interconnection sys- by the crew or passengers; tem components must meet the re- (3) There is a separate approved quirements of § 25.1721.

smoke detector or fire detector system to give warning at the pilot or flight [Amdt. 25–72, 55 FR 29784, July 20, 1990, as engineer station.

amended by Amdt. 25–93, 63 FR 8048, Feb. 17, 1998; Amdt. 25–116, 69 FR 62788, Oct. 27, 2004; (c) Class C. A Class C cargo or bag- Amdt. 25–123, 72 FR 63405, Nov. 8, 2007; Doc.

gage compartment is one not meeting No. Docket FAA–2014–0001, Amdt. 25–142, 81 the requirements for either a Class A FR 7704, Feb. 16, 2016] or B compartment but in which— (1) There is a separate approved § 25.856 Thermal/Acoustic insulation smoke detector or fire detector system materials.

to give warning at the pilot or flight (a) Thermal/acoustic insulation ma- engineer station; terial installed in the fuselage must (2) There is an approved built-in fire meet the flame propagation test re- extinguishing or suppression system quirements of part VI of Appendix F to controllable from the cockpit.

this part, or other approved equivalent (3) There are means to exclude haz- test requirements. This requirement ardous quantities of smoke, flames, or does not apply to ‘‘small parts,’’ as de- extinguishing agent, from any com- fined in part I of Appendix F of this partment occupied by the crew or pas- part.

sengers; (b) For airplanes with a passenger ca- (4) There are means to control ven- pacity of 20 or greater, thermal/acous- tilation and drafts within the compart- tic insulation materials (including the ment so that the extinguishing agent means of fastening the materials to the used can control any fire that may fuselage) installed in the lower half of start within the compartment.

the airplane fuselage must meet the (d) [Reserved] flame penetration resistance test re- (e) Class E. A Class E cargo compart- quirements of part VII of Appendix F ment is one on airplanes used only for to this part, or other approved equiva- the carriage of cargo and in which— lent test requirements. This require- ment does not apply to thermal/acous- (1) [Reserved] tic insulation installations that the (2) There is a separate approved FAA finds would not contribute to fire smoke or fire detector system to give penetration resistance. warning at the pilot or flight engineer station; [Amdt. 25–111, 68 FR 45059, July 31, 2003] (3) There are means to shut off the ventilating airflow to, or within, the § 25.857 Cargo compartment classifica- tion. compartment, and the controls for these means are accessible to the flight (a) Class A; A Class A cargo or bag- crew in the crew compartment; gage compartment is one in which— (4) There are means to exclude haz- (1) The presence of a fire would be ardous quantities of smoke, flames, or easily discovered by a crewmember noxious gases, from the flight crew while at his station; and compartment; and (2) Each part of the compartment is (5) The required crew emergency easily accessible in flight.

exits are accessible under any cargo (b) Class B. A Class B cargo or bag- loading condition.

gage compartment is one in which— (f) Class F. A Class F cargo or bag- (1) There is sufficient access in flight gage compartment must be located on to enable a crewmember, standing at the main deck and is one in which— any one access point and without step- ping into the compartment, to extin- (1) There is a separate approved guish a fire occurring in any part of smoke detector or fire detector system Federal Aviation Administration, DOT § 25.859 to give warning at the pilot or flight (i) Be damaged by heater malfunc- engineer station; tioning; or (2) There are means to extinguish or (ii) Allow flammable fluids or vapors control a fire without requiring a crew- to reach the heater in case of leakage.

member to enter the compartment; and (2) The region surrounding the heat- (3) There are means to exclude haz- er, if the heater fuel system has fit- ardous quantities of smoke, flames, or tings that, if they leaked, would allow extinguishing agent from any compart- fuel or vapors to enter this region.

ment occupied by the crew or pas- (3) The part of the ventilating air sengers.

passage that surrounds the combustion chamber. However, no fire extinguish- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ment is required in cabin ventilating amended by Amdt. 25–32, 37 FR 3972, Feb. 24, 1972; Amdt. 25–60, 51 FR 18243, May 16, 1986; air passages.

Amdt. 25–93, 63 FR 8048, Feb. 17, 1998; Doc.

(b) Ventilating air ducts. Each ven- No. Docket FAA–2014–0001, Amdt. 25–142, 81 tilating air duct passing through any FR 7704, Feb. 16, 2016] fire zone must be fireproof. In addi- tion— § 25.858 Cargo or baggage compart- (1) Unless isolation is provided by ment smoke or fire detection sys- tems. fireproof valves or by equally effective means, the ventilating air duct down- If certification with cargo or baggage stream of each heater must be fireproof compartment smoke or fire detection for a distance great enough to ensure provisions is requested, the following that any fire originating in the heater must be met for each cargo or baggage can be contained in the duct; and compartment with those provisions: (2) Each part of any ventilating duct (a) The detection system must pro- passing through any region having a vide a visual indication to the flight flammable fluid system must be con- crew within one minute after the start structed or isolated from that system of a fire.

so that the malfunctioning of any com- (b) The system must be capable of de- ponent of that system cannot intro- tecting a fire at a temperature signifi- duce flammable fluids or vapors into cantly below that at which the struc- the ventilating airstream.

tural integrity of the airplane is sub- (c) Combustion air ducts. Each com- stantially decreased.

bustion air duct must be fireproof for a (c) There must be means to allow the distance great enough to prevent dam- crew to check in flight, the functioning age from backfiring or reverse flame of each fire detector circuit.

propagation. In addition— (d) The effectiveness of the detection (1) No combustion air duct may have system must be shown for all approved operating configurations and condi- a common opening with the ventilating tions. airstream unless flames from backfires or reverse burning cannot enter the [Amdt. 25–54, 45 FR 60173, Sept. 11, 1980, as ventilating airstream under any oper- amended by Amdt. 25–93, 63 FR 8048, Feb. 17, ating condition, including reverse flow 1998] or malfunctioning of the heater or its associated components; and § 25.859 Combustion heater fire pro- tection.

(2) No combustion air duct may re- strict the prompt relief of any backfire (a) Combustion heater fire zones. The that, if so restricted, could cause heat- following combustion heater fire zones er failure.

must be protected from fire in accord- (d) Heater controls; general. Provision ance with the applicable provisions of must be made to prevent the hazardous §§ 25.1181 through 25.1191 and §§ 25.1195 accumulation of water or ice on or in through 25.1203; any heater control component, control (1) The region surrounding the heat- system tubing, or safety control.

er, if this region contains any flam- mable fluid system components (ex- (e) Heater safety controls. For each cluding the heater fuel system), that combustion heater there must be the could— following safety control means: 14 CFR Ch. I (1–1–25 Edition) § 25.863 (1) Means independent of the compo- (h) Heater fuel systems. Each heater nents provided for the normal contin- fuel system must meet each power- uous control of air temperature, air- plant fuel system requirement affect- flow, and fuel flow must be provided, ing safe heater operation. Each heater for each heater, to automatically shut fuel system component within the ven- off the ignition and fuel supply to that tilating airstream must be protected heater at a point remote from that by shrouds so that no leakage from heater when any of the following oc- those components can enter the ven- curs: tilating airstream.

(i) The heat exchanger temperature (i) Drains. There must be means to exceeds safe limits.

safely drain fuel that might accumu- (ii) The ventilating air temperature late within the combustion chamber or exceeds safe limits.

the heat exchanger. In addition— (iii) The combustion airflow becomes (1) Each part of any drain that oper- inadequate for safe operation.

ates at high temperatures must be pro- (iv) The ventilating airflow becomes tected in the same manner as heater inadequate for safe operation.

exhausts; and (2) The means of complying with (2) Each drain must be protected paragraph (e)(1) of this section for any from hazardous ice accumulation under individual heater must— any operating condition.

(i) Be independent of components [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as serving any other heater whose heat amended by Amdt. 25–11, 32 FR 6912, May 5, output is essential for safe operation; 1967; Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] and (ii) Keep the heater off until re- § 25.863 Flammable fluid fire protec- started by the crew.

tion.

(3) There must be means to warn the (a) In each area where flammable crew when any heater whose heat out- fluids or vapors might escape by leak- put is essential for safe operation has age of a fluid system, there must be been shut off by the automatic means means to minimize the probability of prescribed in paragraph (e)(1) of this ignition of the fluids and vapors, and section.

the resultant hazards if ignition does (f) Air intakes. Each combustion and occur.

ventilating air intake must be located (b) Compliance with paragraph (a) of so that no flammable fluids or vapors this section must be shown by analysis can enter the heater system under any or tests, and the following factors must operating condition— be considered: (1) During normal operation; or (1) Possible sources and paths of fluid (2) As a result of the malfunctioning leakage, and means of detecting leak- of any other component.

age.

(g) Heater exhaust. Heater exhaust (2) Flammability characteristics of systems must meet the provisions of fluids, including effects of any combus- §§ 25.1121 and 25.1123. In addition, there tible or absorbing materials.

must be provisions in the design of the (3) Possible ignition sources, includ- heater exhaust system to safely expel ing electrical faults, overheating of the products of combustion to prevent equipment, and malfunctioning of pro- the occurrence of— tective devices.

(1) Fuel leakage from the exhaust to (4) Means available for controlling or surrounding compartments; extinguishing a fire, such as stopping (2) Exhaust gas impingement on sur- flow of fluids, shutting down equip- rounding equipment or structure; ment, fireproof containment, or use of (3) Ignition of flammable fluids by extinguishing agents.

the exhaust, if the exhaust is in a com- (5) Ability of airplane components partment containing flammable fluid that are critical to safety of flight to lines; and withstand fire and heat.

(4) Restriction by the exhaust of the prompt relief of backfires that, if so re- (c) If action by the flight crew is re- stricted, could cause heater failure. quired to prevent or counteract a fluid Federal Aviation Administration, DOT § 25.899 fire (e.g., equipment shutdown or actu- fire zones must be at least fire resist- ation of a fire extinguisher) quick act- ant.

ing means must be provided to alert (c) Oxygen equipment and lines the crew.

must— (d) Each area where flammable fluids (1) Not be located in any designated or vapors might escape by leakage of a fire zone, fluid system must be identified and de- (2) Be protected from heat that may fined.

be generated in, or escape from, any [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970, as designated fire zone, and amended by Amdt. 25–46, 43 FR 50597, Oct. 30, (3) Be installed so that escaping oxy- 1978] gen cannot cause ignition of grease, fluid, or vapor accumulations that are § 25.865 Fire protection of flight con- present in normal operation or as a re- trols, engine mounts, and other sult of failure or malfunction of any flight structure.

system.

Essential flight controls, engine mounts, and other flight structures lo- [Amdt. 25–72, 55 FR 29784, July 20, 1990, as cated in designated fire zones or in ad- amended by Amdt. 25–113, 69 FR 12530, Mar.

16, 2004; Amdt. 25–123, 72 FR 63405, Nov. 8, jacent areas which would be subjected 2007] to the effects of fire in the fire zone must be constructed of fireproof mate- M ISCELLANEOUS rial or shielded so that they are capa- ble of withstanding the effects of fire.

§ 25.871 Leveling means.

[Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] There must be means for determining when the airplane is in a level position § 25.867 Fire protection: other compo- nents. on the ground.

(a) Surfaces to the rear of the na- [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] celles, within one nacelle diameter of the nacelle centerline, must be at least § 25.875 Reinforcement near propel- fire-resistant. lers.

(b) Paragraph (a) of this section does (a) Each part of the airplane near the not apply to tail surfaces to the rear of propeller tips must be strong and stiff the nacelles that could not be readily enough to withstand the effects of the affected by heat, flames, or sparks induced vibration and of ice thrown coming from a designated fire zone or from the propeller.

engine compartment of any nacelle.

(b) No window may be near the pro- [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] peller tips unless it can withstand the most severe ice impact likely to occur.

§ 25.869 Fire protection: systems.

(a) Electrical system components: § 25.899 Electrical bonding and protec- tion against static electricity.

(1) Components of the electrical sys- tem must meet the applicable fire and (a) Electrical bonding and protection smoke protection requirements of against static electricity must be de- §§ 25.831(c) and 25.863.

signed to minimize accumulation of (2) Equipment that is located in des- electrostatic charge that would cause— ignated fire zones and is used during (1) Human injury from electrical emergency procedures must be at least shock, fire resistant.

(2) Ignition of flammable vapors, or (3) EWIS components must meet the (3) Interference with installed elec- requirements of § 25.1713.

trical/electronic equipment.

(b) Each vacuum air system line and (b) Compliance with paragraph (a) of fitting on the discharge side of the this section may be shown by— pump that might contain flammable (1) Bonding the components properly vapors or fluids must meet the require- to the airframe; or ments of § 25.1183 if the line or fitting is in a designated fire zone. Other vacuum (2) Incorporating other acceptable air systems components in designated means to dissipate the static charge so 14 CFR Ch. I (1–1–25 Edition) § 25.901 as not to endanger the airplane, per- § 25.903 Engines.

sonnel, or operation of the installed (a) Engine type certificate. (1) Each en- electrical/electronic systems.

gine must have a type certificate and [Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] must meet the applicable requirements of part 34 of this chapter.

(2) Each turbine engine must comply Subpart E—Powerplant with one of the following: (i) Sections 33.76, 33.77 and 33.78 of G ENERAL this chapter in effect on December 13, § 25.901 Installation.

2000, or as subsequently amended; or (ii) Sections 33.77 and 33.78 of this (a) For the purpose of this part, the chapter in effect on April 30, 1998, or as airplane powerplant installation in- subsequently amended before Decem- cludes each component that— ber 13, 2000; or (1) Is necessary for propulsion; (iii) Comply with § 33.77 of this chap- (2) Affects the control of the major ter in effect on October 31, 1974, or as propulsive units; or subsequently amended prior to April (3) Affects the safety of the major 30, 1998, unless that engine’s foreign ob- propulsive units between normal in- ject ingestion service history has re- spections or overhauls.

sulted in an unsafe condition; or (b) For each powerplant— (iv) Be shown to have a foreign object (1) The installation must comply ingestion service history in similar in- with— stallation locations which has not re- (i) The installation instructions pro- sulted in any unsafe condition.

vided under §§ 33.5 and 35.3 of this chap- ter; and N OTE : § 33.77 of this chapter in effect on Oc- tober 31, 1974, was published in 14 CFR parts (ii) The applicable provisions of this 1 to 59, Revised as of January 1, 1975. See 39 subpart; FR 35467, October 1, 1974.

(2) The components of the installa- tion must be constructed, arranged, (3) Each turbine engine must comply and installed so as to ensure their con- with one of the following paragraphs: tinued safe operation between normal (i) Section 33.68 of this chapter in ef- inspections or overhauls; fect on January 5, 2015, or as subse- (3) The installation must be acces- quently amended; or sible for necessary inspections and (ii) Section 33.68 of this chapter in ef- maintenance; and fect on March 26, 1984, or as subse- (4) The major components of the in- quently amended before January 5, stallation must be electrically bonded 2015, unless that engine’s ice accumula- to the other parts of the airplane. tion service history has resulted in an (c) For each powerplant and auxiliary unsafe condition; or power unit installation, the applicant (iii) Section 33.68 of this chapter in must comply with the requirements of effect on October 31, 1974, or as subse- § 25.1309, except that the effects of the quently amended prior to February 23, following failures need not comply 1984, unless that engine’s ice accumula- with § 25.1309(b)— tion service history has resulted in an (1) Engine case burn-through or rup- unsafe condition; or ture, (iv) Be shown to have an ice accumu- (2) Uncontained engine rotor failure, lation service history in similar instal- and lation locations which has not resulted in any unsafe conditions.

(3) Propeller debris release.

(b) Engine isolation. The powerplants (d) Each auxiliary power unit instal- must be arranged and isolated from lation must meet the applicable provi- each other to allow operation, in at sions of this subpart.

least one configuration, so that the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as failure or malfunction of any engine, or amended by Amdt. 25–23, 35 FR 5676, Apr. 8, of any system that can affect the en- 1970; Amdt. 25–40, 42 FR 15042, Mar. 17, 1977; gine, will not— Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt.

(1) Prevent the continued safe oper- 25–126, 73 FR 63345, Oct. 24, 2008; Doc. No.

FAA–2022–1544, 89 FR 68734, Aug. 27, 2024] ation of the remaining engines; or Federal Aviation Administration, DOT § 25.907 (2) Require immediate action by any meet the requirements of the category crewmember for continued safe oper- for its intended use.

ation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) Control of engine rotation. There amended by Amdt. 25–23, 35 FR 5676, Apr. 8, must be means for stopping the rota- 1970; Amdt. 25–40, 42 FR 15042, Mar. 17, 1977; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; Amdt.

tion of any engine individually in 25–72, 55 FR 29784, July 20, 1990; Amdt. 25–73, flight, except that, for turbine engine 55 FR 32861, Aug. 10, 1990; Amdt. 25–94, 63 FR installations, the means for stopping 8848, Feb. 23, 1998; Amdt. 25–95, 63 FR 14798, the rotation of any engine need be pro- Mar. 26, 1998; Amdt. 25–100, 65 FR 55854, Sept.

vided only where continued rotation 14, 2000; Amdt. 25–140, 79 FR 65525, Nov. 4, could jeopardize the safety of the air- 2014; Amdt. No. 25–148, 87 FR 75710, Dec. 9, 2022; 88 FR 2813, Jan. 18, 2023] plane. Each component of the stopping system on the engine side of the fire- § 25.904 Automatic takeoff thrust con- wall that might be exposed to fire must trol system (ATTCS).

be at least fire-resistant. If hydraulic Each applicant seeking approval for propeller feathering systems are used installation of an engine power control for this purpose, the feathering lines system that automatically resets the must be at least fire resistant under power or thrust on the operating en- the operating conditions that may be gine(s) when any engine fails during expected to exist during feathering.

the takeoff must comply with the re- (d) Turbine engine installations. For quirements of appendix I of this part.

turbine engine installations— [Amdt. 25–62, 52 FR 43156, Nov. 9, 1987] (1) Design precautions must be taken to minimize the hazards to the airplane § 25.905 Propellers.

in the event of an engine rotor failure (a) Each propeller must have a type or of a fire originating within the en- certificate.

gine which burns through the engine (b) Engine power and propeller shaft case.

rotational speed may not exceed the (2) The powerplant systems associ- limits for which the propeller is certifi- ated with engine control devices, sys- cated.

tems, and instrumentation, must be de- (c) The propeller blade pitch control signed to give reasonable assurance system must meet the requirements of that those engine operating limitations §§ 35.21, 35.23, 35.42 and 35.43 of this that adversely affect turbine rotor chapter.

structural integrity will not be exceed- (d) Design precautions must be taken ed in service.

to minimize the hazards to the airplane (e) Restart capability. (1) Means to re- in the event a propeller blade fails or is start any engine in flight must be pro- released by a hub failure. The hazards vided.

which must be considered include dam- (2) An altitude and airspeed envelope age to structure and vital systems due must be established for in-flight engine to impact of a failed or released blade and the unbalance created by such fail- restarting, and each engine must have ure or release.

a restart capability within that enve- lope.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (3) For turbine engine powered air- amended by Amdt. 25–54, 45 FR 60173, Sept.

11, 1980; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; planes, if the minimum windmilling Amdt. 25–72, 55 FR 29784, July 20, 1990; Amdt.

speed of the engines, following the 25–126, 73 FR 63345, Oct. 24, 2008] inflight shutdown of all engines, is in- sufficient to provide the necessary § 25.907 Propeller vibration and fa- electrical power for engine ignition, a tigue.

power source independent of the en- This section does not apply to fixed- gine-driven electrical power generating pitch wood propellers of conventional system must be provided to permit in- design.

flight engine ignition for restarting.

(a) The applicant must determine the (f) Auxiliary Power Unit. Each auxil- magnitude of the propeller vibration iary power unit must be approved or stresses or loads, including any stress 14 CFR Ch. I (1–1–25 Edition) § 25.925 peaks and resonant conditions, each propeller and the ground with the throughout the operational envelope of landing gear statically deflected and in the airplane by either: the level takeoff, or taxiing attitude, (1) Measurement of stresses or loads whichever is most critical. In addition, through direct testing or analysis there must be positive clearance be- based on direct testing of the propeller tween the propeller and the ground on the airplane and engine installation when in the level takeoff attitude with for which approval is sought; or the critical tire(s) completely deflated (2) Comparison of the propeller to and the corresponding landing gear similar propellers installed on similar strut bottomed.

airplane installations for which these (b) Water clearance. There must be a measurements have been made.

clearance of at least 18 inches between (b) The applicant must demonstrate each propeller and the water, unless by tests, analysis based on tests, or compliance with § 25.239(a) can be previous experience on similar designs shown with a lesser clearance.

that the propeller does not experience (c) Structural clearance. There must harmful effects of flutter throughout be— the operational envelope of the air- (1) At least one inch radial clearance plane.

between the blade tips and the airplane (c) The applicant must perform an structure, plus any additional radial evaluation of the propeller to show clearance necessary to prevent harmful that failure due to fatigue will be vibration; avoided throughout the operational life (2) At least one-half inch longitudinal of the propeller using the fatigue and clearance between the propeller blades structural data obtained in accordance or cuffs and stationary parts of the air- with part 35 of this chapter and the vi- plane; and bration data obtained from compliance (3) Positive clearance between other with paragraph (a) of this section. For rotating parts of the propeller or spin- the purpose of this paragraph, the pro- ner and stationary parts of the air- peller includes the hub, blades, blade plane.

retention component and any other propeller component whose failure due [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as to fatigue could be catastrophic to the amended by Amdt. 25–72, 55 FR 29784, July 20, 1990] airplane. This evaluation must include: (1) The intended loading spectra in- § 25.929 Propeller deicing.

cluding all reasonably foreseeable pro- peller vibration and cyclic load pat- (a) If certification for flight in icing terns, identified emergency conditions, is sought there must be a means to pre- allowable overspeeds and overtorques, vent or remove hazardous ice accumu- and the effects of temperatures and hu- lations that could form in the icing midity expected in service. conditions defined in Appendix C of (2) The effects of airplane and pro- this part and in the portions of Appen- peller operating and airworthiness lim- dix O of this part for which the air- itations. plane is approved for flight on propel- lers or on accessories where ice accu- [Amdt. 25–126, 73 FR 63345, Oct. 24, 2008] mulation would jeopardize engine per- formance.

§ 25.925 Propeller clearance.

(b) If combustible fluid is used for Unless smaller clearances are sub- propeller deicing, §§ 25.1181 through stantiated, propeller clearances with 25.1185 and 25.1189 apply.

the airplane at maximum weight, with the most adverse center of gravity, and [ Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–140, 79 FR 65525, Nov. 4, with the propeller in the most adverse 2014] pitch position, may not be less than the following: § 25.933 Reversing systems.

(a) Ground clearance. There must be a clearance of at least seven inches (for (a) For turbojet reversing systems— each airplane with nose wheel landing (1) For each system intended for gear) or nine inches (for each airplane ground operation only, the applicant with tail wheel landing gear) between must show— Federal Aviation Administration, DOT § 25.941 (i) The airplane is capable of contin- § 25.934 Turbojet engine thrust re- ued safe flight and landing during and verser system tests.

after any thrust reversal in flight; or Thrust reversers installed on tur- (ii) The system complies with bojet engines must meet the require- § 25.1309(b) using the assumption the ments of § 33.97 of this chapter.

airplane would not be capable of con- [Amdt. 25–23, 35 FR 5677, Apr. 8, 1970] tinued safe flight and landing during and after an in-flight thrust reversal.

§ 25.937 Turbopropeller-drag limiting (2) Each system intended for inflight systems.

use must be designed so that no unsafe Turbopropeller power airplane pro- condition will result during normal op- peller-drag limiting systems must be eration of the system, or from any fail- designed so that no single failure or ure (or reasonably likely combination malfunction of any of the systems dur- of failures) of the reversing system, ing normal or emergency operation re- under any anticipated condition of op- sults in propeller drag in excess of that eration of the airplane including for which the airplane was designed ground operation. Failure of structural under § 25.367. Failure of structural ele- elements need not be considered if the ments of the drag limiting systems probability of this kind of failure is ex- need not be considered if the prob- tremely remote.

ability of this kind of failure is ex- (3) Each system must have means to tremely remote.

prevent the engine from producing more than idle thrust when the revers- § 25.939 Turbine engine operating ing system malfunctions, except that it characteristics.

may produce any greater forward (a) Turbine engine operating charac- thrust that is shown to allow direc- teristics must be investigated in flight tional control to be maintained, with to determine that no adverse charac- aerodynamic means alone, under the teristics (such as stall, surge, or flame- most critical reversing condition ex- out) are present, to a hazardous degree, pected in operation.

during normal and emergency oper- (b) For propeller reversing systems— ation within the range of operating (1) Each system intended for ground limitations of the airplane and of the operation only must be designed so engine.

that no single failure (or reasonably (b) [Reserved] likely combination of failures) or mal- (c) The turbine engine air inlet sys- function of the system will result in tem may not, as a result of air flow dis- unwanted reverse thrust under any ex- tortion during normal operation, cause pected operating condition. Failure of vibration harmful to the engine.

structural elements need not be consid- [Amdt. 25–11, 32 FR 6912, May 5, 1967, as ered if this kind of failure is extremely amended by Amdt. 25–40, 42 FR 15043, Mar. 17, remote.

1977] (2) Compliance with this section may be shown by failure analysis or testing, § 25.941 Inlet, engine, and exhaust or both, for propeller systems that compatibility.

allow propeller blades to move from For airplanes using variable inlet or the flight low-pitch position to a posi- exhaust system geometry, or both— tion that is substantially less than (a) The system comprised of the that at the normal flight low-pitch po- inlet, engine (including thrust aug- sition. The analysis may include or be mentation systems, if incorporated), supported by the analysis made to and exhaust must be shown to function show compliance with the require- properly under all operating conditions ments of § 35.21 of this chapter for the for which approval is sought, including propeller and associated installation all engine rotating speeds and power components.

settings, and engine inlet and exhaust configurations; [Amdt. 25–72, 55 FR 29784, July 20, 1990, as (b) The dynamic effects of the oper- amended by Doc. No. FAA–2022–1544, 89 FR 68734, Aug. 27, 2024] ation of these (including consideration 14 CFR Ch. I (1–1–25 Edition) § 25.943 of probable malfunctions) upon the aer- (5) Each tank must have an expan- odynamic control of the airplane may sion space of not less than 2 percent of not result in any condition that would the tank capacity. It must be impos- require exceptional skill, alertness, or sible to fill the expansion space inad- strength on the part of the pilot to vertently with the airplane in the nor- mal ground attitude.

avoid exceeding an operational or structural limitation of the airplane; (c) Augmentation system drains and must be designed and located in ac- cordance with § 25.1455 if— (c) In showing compliance with para- (1) The augmentation system fluid is graph (b) of this section, the pilot subject to freezing; and strength required may not exceed the limits set forth in § 25.143(d), subject to (2) The fluid may be drained in flight the conditions set forth in paragraphs or during ground operation.

(e) and (f) of § 25.143.

(d) The augmentation liquid tank ca- pacity available for the use of each en- [Amdt. 25–38, 41 FR 55467, Dec. 20, 1976, as gine must be large enough to allow op- amended by Amdt. 25–121, 72 FR 44669, Aug. 8, eration of the airplane under the ap- 2007] proved procedures for the use of liquid- augmented power. The computation of § 25.943 Negative acceleration.

liquid consumption must be based on No hazardous malfunction of an en- the maximum approved rate appro- gine, an auxiliary power unit approved priate for the desired engine output for use in flight, or any component or and must include the effect of tempera- system associated with the powerplant ture on engine performance as well as or auxiliary power unit may occur any other factors that might vary the when the airplane is operated at the amount of liquid required.

negative accelerations within the (e) This section does not apply to fuel flight envelopes prescribed in § 25.333.

injection systems.

This must be shown for the greatest duration expected for the acceleration. [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977, as amended by Amdt. 25–72, 55 FR 29785, July 20, [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] 1990; Amdt. 25–115, 69 FR 40527, July 2, 2004] § 25.945 Thrust or power augmentation F UEL S YSTEM system.

§ 25.951 General.

(a) General. Each fluid injection sys- tem must provide a flow of fluid at the (a) Each fuel system must be con- rate and pressure established for proper structed and arranged to ensure a flow engine functioning under each intended of fuel at a rate and pressure estab- operating condition. If the fluid can lished for proper engine and auxiliary freeze, fluid freezing may not damage power unit functioning under each the airplane or adversely affect air- likely operating condition, including plane performance.

any maneuver for which certification is (b) Fluid tanks. Each augmentation requested and during which the engine system fluid tank must meet the fol- or auxiliary power unit is permitted to lowing requirements: be in operation.

(1) Each tank must be able to with- (b) Each fuel system must be ar- stand without failure the vibration, in- ranged so that any air which is intro- ertia, fluid, and structural loads that it duced into the system will not result may be subject to in operation. in— (2) The tanks as mounted in the air- (1) Power interruption for more than plane must be able to withstand with- 20 seconds for reciprocating engines; or out failure or leakage an internal pres- (2) Flameout for turbine engines.

sure 1.5 times the maximum operating (c) Each fuel system for a turbine en- pressure.

gine must be capable of sustained oper- (3) If a vent is provided, the venting ation throughout its flow and pressure must be effective under all normal range with fuel initially saturated with flight conditions.

water at 80 ° F and having 0.75cc of free (4) [Reserved] water per gallon added and cooled to Federal Aviation Administration, DOT § 25.955 the most critical condition for icing (2) Swept lightning strokes to areas likely to be encountered in operation. where swept strokes are highly prob- (d) Each fuel system for a turbine en- able; and gine powered airplane must meet the (3) Lightning-induced or conducted applicable fuel venting requirements of electrical transients.

part 34 of this chapter.

(c) To comply with paragraph (b) of this section, catastrophic fuel vapor ig- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5677, Apr. 8, nition must be extremely improbable, 1970; Amdt. 25–36, 39 FR 35460, Oct. 1, 1974; taking into account flammability, crit- Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt.

ical lightning strikes, and failures 25–73, 55 FR 32861, Aug. 10, 1990] within the fuel system.

(d) To protect design features that § 25.952 Fuel system analysis and test.

prevent catastrophic fuel vapor igni- (a) Proper fuel system functioning tion caused by lightning, the type de- under all probable operating conditions sign must include critical design con- must be shown by analysis and those figuration control limitations tests found necessary by the Adminis- (CDCCLs) identifying those features trator. Tests, if required, must be made and providing information to protect using the airplane fuel system or a test them. To ensure the continued effec- article that reproduces the operating tiveness of those design features, the characteristics of the portion of the type design must also include inspec- fuel system to be tested.

tion and test procedures, intervals be- (b) The likely failure of any heat ex- tween repetitive inspections and tests, changer using fuel as one of its fluids and mandatory replacement times for may not result in a hazardous condi- tion. those design features used in dem- onstrating compliance to paragraph (b) [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] of this section. The applicant must in- clude the information required by this § 25.953 Fuel system independence.

paragraph in the Airworthiness Limi- Each fuel system must meet the re- tations section of the Instructions for quirements of § 25.903(b) by— Continued Airworthiness required by (a) Allowing the supply of fuel to § 25.1529.

each engine through a system inde- pendent of each part of the system sup- [Doc. No. FAA–2014–1027, Amdt. 25–146, 83 FR plying fuel to any other engine; or 47556, Sept. 20, 2018] (b) Any other acceptable method.

§ 25.955 Fuel flow.

§ 25.954 Fuel system lightning protec- (a) Each fuel system must provide at tion.

least 100 percent of the fuel flow re- (a) For purposes of this section— quired under each intended operating (1) A critical lightning strike is a condition and maneuver. Compliance lightning strike that attaches to the must be shown as follows: airplane in a location that, when com- (1) Fuel must be delivered to each en- bined with the failure of any design gine at a pressure within the limits feature or structure, could create an specified in the engine type certificate.

ignition source.

(2) The quantity of fuel in the tank (2) A fuel system includes any compo- may not exceed the amount established nent within either the fuel tank struc- as the unusable fuel supply for that ture or the fuel tank systems, and any tank under the requirements of § 25.959 airplane structure or system compo- plus that necessary to show compliance nents that penetrate, connect to, or are with this section.

located within a fuel tank.

(3) Each main pump must be used (b) The design and installation of a that is necessary for each operating fuel system must prevent catastrophic fuel vapor ignition due to lightning and condition and attitude for which com- its effects, including: pliance with this section is shown, and (1) Direct lightning strikes to areas the appropriate emergency pump must having a high probability of stroke at- be substituted for each main pump so tachment; used.

14 CFR Ch. I (1–1–25 Edition) § 25.957 (4) If there is a fuel flowmeter, it the fuel system from the tank outlets must be blocked and the fuel must flow to each engine is pressurized, under all through the meter or its bypass. intended operations, so as to prevent (b) If an engine can be supplied with vapor formation, or must be shown by fuel from more than one tank, the fuel climbing from the altitude of the air- system must— port elected by the applicant to the (1) For each reciprocating engine, maximum altitude established as an supply the full fuel pressure to that en- operating limitation under § 25.1527. If a gine in not more than 20 seconds after climb test is elected, there may be no switching to any other fuel tank con- evidence of vapor lock or other mal- taining usable fuel when engine mal- functioning during the climb test con- functioning becomes apparent due to ducted under the following conditions: the depletion of the fuel supply in any (1) For reciprocating engine powered tank from which the engine can be fed; airplanes, the engines must operate at and maximum continuous power, except (2) For each turbine engine, in addi- that takeoff power must be used for the tion to having appropriate manual altitudes from 1,000 feet below the crit- switching capability, be designed to ical altitude through the critical alti- prevent interruption of fuel flow to tude. The time interval during which that engine, without attention by the takeoff power is used may not be less flight crew, when any tank supplying than the takeoff time limitation.

fuel to that engine is depleted of usable (2) For turbine engine powered air- fuel during normal operation, and any planes, the engines must operate at other tank, that normally supplies fuel takeoff power for the time interval se- to that engine alone, contains usable lected for showing the takeoff flight fuel.

path, and at maximum continuous power for the rest of the climb.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–11, 32 FR 6912, May 5, (3) The weight of the airplane must 1967] be the weight with full fuel tanks, min- imum crew, and the ballast necessary § 25.957 Flow between interconnected to maintain the center of gravity with- tanks.

in allowable limits.

If fuel can be pumped from one tank (4) The climb airspeed may not ex- to another in flight, the fuel tank ceed— vents and the fuel transfer system (i) For reciprocating engine powered must be designed so that no structural airplanes, the maximum airspeed es- damage to the tanks can occur because tablished for climbing from takeoff to of overfilling.

the maximum operating altitude with the airplane in the following configura- § 25.959 Unusable fuel supply.

tion: The unusable fuel quantity for each (A) Landing gear retracted.

fuel tank and its fuel system compo- (B) Wing flaps in the most favorable nents must be established at not less position.

than the quantity at which the first (C) Cowl flaps (or other means of con- evidence of engine malfunction occurs trolling the engine cooling supply) in under the most adverse fuel feed condi- the position that provides adequate tion for all intended operations and cooling in the hot-day condition.

flight maneuvers involving fuel feeding (D) Engine operating within the max- from that tank. Fuel system compo- imum continuous power limitations.

nent failures need not be considered.

(E) Maximum takeoff weight; and (ii) For turbine engine powered air- [Amdt. 25–23, 35 FR 5677, Apr. 8, 1970, as planes, the maximum airspeed estab- amended by Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] lished for climbing from takeoff to the maximum operating altitude.

§ 25.961 Fuel system hot weather oper- (5) The fuel temperature must be at ation.

least 110 ° F.

(a) The fuel system must perform (b) The test prescribed in paragraph satisfactorily in hot weather operation. (a) of this section may be performed in This must be shown by showing that flight or on the ground under closely Federal Aviation Administration, DOT § 25.963 K = 1.5 for the inboard and outboard loading simulated flight conditions. If a flight conditions for those parts of fuel tanks test is performed in weather cold outside the fuselage pressure boundary enough to interfere with the proper K = 6 for the downward loading condition conduct of the test, the fuel tank sur- K = 3 for the upward loading condition faces, fuel lines, and other fuel system (2) For those parts of wing fuel tanks parts subject to cold air must be insu- near the fuselage or near the engines, lated to simulate, insofar as prac- the greater of the fuel pressures result- ticable, flight in hot weather.

ing from paragraphs (d)(2)(i) or (d)(2)(ii) [Amdt. 25–11, 32 FR 6912, May 5, 1967, as of this section must be used: amended by Amdt. 25–57, 49 FR 6848, Feb. 23, (i) The fuel pressures resulting from 1984] paragraph (d)(1) of this section, and (ii) The lesser of the two following § 25.963 Fuel tanks: general.

conditions: (a) Each fuel tank must be able to (A) Fuel pressures resulting from the withstand, without failure, the vibra- accelerations specified in § 25.561(b)(3) tion, inertia, fluid, and structural loads considering the fuel tank full of fuel at that it may be subjected to in oper- maximum fuel density. Fuel pressures ation.

based on the 9.0g forward acceleration (b) Flexible fuel tank liners must be may be calculated using the fuel static approved or must be shown to be suit- head equal to the streamwise local able for the particular application.

chord of the tank. For inboard and out- (c) Integral fuel tanks must have fa- board conditions, an acceleration of cilities for interior inspection and re- 1.5g may be used in lieu of 3.0g as speci- pair.

fied in § 25.561(b)(3).

(d) Fuel tanks must, so far as it is (B) Fuel pressures resulting from the practicable, be designed, located, and accelerations as specified in installed so that no fuel is released in § 25.561(b)(3) considering a fuel volume or near the fuselage, or near the en- beyond 85 percent of the maximum per- gines, in quantities that would con- missible volume in each tank using the stitute a fire hazard in otherwise sur- static head associated with the 85 per- vivable emergency landing conditions, cent fuel level. A typical density of the and— appropriate fuel may be used. For in- (1) Fuel tanks must be able to resist board and outboard conditions, an ac- rupture and retain fuel under ultimate celeration of 1.5g may be used in lieu of hydrostatic design conditions in which 3.0g as specified in § 25.561(b)(3).

the pressure P within the tank varies (3) Fuel tank internal barriers and in accordance with the formula: baffles may be considered as solid boundaries if shown to be effective in P = K r gL limiting fuel flow.

Where— (4) For each fuel tank and sur- P = fuel pressure at each point within the rounding airframe structure, the ef- tank fects of crushing and scraping actions r = typical fuel density with the ground must not cause the g = acceleration due to gravity spillage of enough fuel, or generate L = a reference distance between the point of temperatures that would constitute a pressure and the tank farthest boundary fire hazard under the conditions speci- in the direction of loading fied in § 25.721(b).

K = 4.5 for the forward loading condition for (5) Fuel tank installations must be those parts of fuel tanks outside the fu- selage pressure boundary such that the tanks will not rupture as K = 9 for the forward loading condition for a result of the landing gear or an en- those parts of fuel tanks within the fuse- gine pylon or engine mount tearing lage pressure boundary, or that form part away as specified in § 25.721(a) and (c).

of the fuselage pressure boundary (e) Fuel tank access covers must K = 1.5 for the aft loading condition comply with the following criteria in K = 3.0 for the inboard and outboard loading order to avoid loss of hazardous quan- conditions for those parts of fuel tanks tities of fuel: within the fuselage pressure boundary, or (1) All covers located in an area that form part of the fuselage pressure boundary where experience or analysis indicates 14 CFR Ch. I (1–1–25 Edition) § 25.965 a strike is likely must be shown by tiated) while ⁄3 filled with water or analysis or tests to minimize penetra- other suitable test fluid.

tion and deformation by tire frag- (3) The test frequency of vibration ments, low energy engine debris, or must be as follows: (i) If no frequency of vibration result- other likely debris.

ing from any r.p.m. within the normal (2) All covers must be fire resistant operating range of engine speeds is as defined in part 1 of this chapter.

critical, the test frequency of vibration (f) For pressurized fuel tanks, a must be 2,000 cycles per minute.

means with fail-safe features must be (ii) If only one frequency of vibration provided to prevent the buildup of an resulting from any r.p.m. within the excessive pressure difference between normal operating range of engine the inside and the outside of the tank.

speeds is critical, that frequency of vi- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as bration must be the test frequency.

amended by Amdt. 25–40, 42 FR 15043, Mar. 17, (iii) If more than one frequency of vi- 1977; Amdt. 25–69, 54 FR 40354, Sept. 29, 1989; bration resulting from any r.p.m. with- Amdt. 25–139, 79 FR 59430, Oct. 2, 2014] in the normal operating range of en- gine speeds is critical, the most crit- § 25.965 Fuel tank tests.

ical of these frequencies must be the (a) It must be shown by tests that the test frequency.

fuel tanks, as mounted in the airplane, (4) Under paragraphs (b)(3)(ii) and can withstand, without failure or leak- (iii) of this section, the time of test age, the more critical of the pressures must be adjusted to accomplish the resulting from the conditions specified same number of vibration cycles that in paragraphs (a)(1) and (2) of this sec- would be accomplished in 25 hours at tion. In addition, it must be shown by the frequency specified in paragraph either analysis or tests, that tank sur- (b)(3)(i) of this section.

faces subjected to more critical pres- (5) During the test, the tank assem- sures resulting from the condition of bly must be rocked at the rate of 16 to paragraphs (a)(3) and (4) of this section, 20 complete cycles per minute, through are able to withstand the following an angle of 15 ° on both sides of the hor- pressures: izontal (30 ° total), about the most crit- (1) An internal pressure of 3.5 psi.

ical axis, for 25 hours. If motion about (2) 125 percent of the maximum air more than one axis is likely to be crit- pressure developed in the tank from ical, the tank must be rocked about ram effect.

each critical axis for 12 ⁄ 2 hours.

(3) Fluid pressures developed during (c) Except where satisfactory oper- maximum limit accelerations, and de- ating experience with a similar tank in flections, of the airplane with a full a similar installation is shown, non- tank. metallic tanks must withstand the test (4) Fluid pressures developed during specified in paragraph (b)(5) of this sec- the most adverse combination of air- tion, with fuel at a temperature of 110 plane roll and fuel load. ° F. During this test, a representative specimen of the tank must be installed (b) Each metallic tank with large un- in a supporting structure simulating supported or unstiffened flat surfaces, the installation in the airplane.

whose failure or deformation could (d) For pressurized fuel tanks, it cause fuel leakage, must be able to must be shown by analysis or tests withstand the following test, or its that the fuel tanks can withstand the equivalent, without leakage or exces- maximum pressure likely to occur on sive deformation of the tank walls: the ground or in flight.

(1) Each complete tank assembly and its supports must be vibration tested [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as while mounted to simulate the actual amended by Amdt. 25–11, 32 FR 6913, May 5, 1967; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] installation.

(2) Except as specified in paragraph § 25.967 Fuel tank installations.

(b)(4) of this section, the tank assembly must be vibrated for 25 hours at an am- (a) Each fuel tank must be supported plitude of not less than ⁄32 of an inch so that tank loads (resulting from the (unless another amplitude is substan- weight of the fuel in the tanks) are not Federal Aviation Administration, DOT § 25.975 concentrated on unsupported tank sur- (b) Each fuel tank must allow drain- faces. In addition— age of any hazardous quantity of water (1) There must be pads, if necessary, from any part of the tank to its sump to prevent chafing between the tank with the airplane in the ground atti- and its supports; tude.

(2) Padding must be nonabsorbent or (c) Each fuel tank sump must have treated to prevent the absorption of an accessible drain that— fluids; (1) Allows complete drainage of the sump on the ground; (3) If a flexible tank liner is used, it (2) Discharges clear of each part of must be supported so that it is not re- the airplane; and quired to withstand fluid loads; and (3) Has manual or automatic means (4) Each interior surface of the tank for positive locking in the closed posi- compartment must be smooth and free tion.

of projections that could cause wear of the liner unless— § 25.973 Fuel tank filler connection.

(i) Provisions are made for protection Each fuel tank filler connection must of the liner at these points; or prevent the entrance of fuel into any (ii) The construction of the liner part of the airplane other than the itself provides that protection.

tank itself. In addition— (b) Spaces adjacent to tank surfaces (a) [Reserved] must be ventilated to avoid fume accu- (b) Each recessed filler connection mulation due to minor leakage. If the that can retain any appreciable quan- tank is in a sealed compartment, ven- tity of fuel must have a drain that dis- tilation may be limited to drain holes charges clear of each part of the air- large enough to prevent excessive pres- plane; sure resulting from altitude changes.

(c) Each filler cap must provide a (c) The location of each tank must fuel-tight seal; and meet the requirements of § 25.1185(a).

(d) Each fuel filling point must have (d) No engine nacelle skin imme- a provision for electrically bonding the diately behind a major air outlet from airplane to ground fueling equipment.

the engine compartment may act as the wall of an integral tank.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (e) Each fuel tank must be isolated amended by Amdt. 25–40, 42 FR 15043, Mar. 17, from personnel compartments by a 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–115, 69 FR 40527, July 2, 2004] fumeproof and fuelproof enclosure.

§ 25.975 Fuel tank vents and carbu- § 25.969 Fuel tank expansion space.

retor vapor vents.

Each fuel tank must have an expan- (a) Fuel tank vents. Each fuel tank sion space of not less than 2 percent of must be vented from the top part of the the tank capacity. It must be impos- expansion space so that venting is ef- sible to fill the expansion space inad- fective under any normal flight condi- vertently with the airplane in the nor- tion. In addition— mal ground attitude. For pressure fuel- (1) Each vent must be arranged to ing systems, compliance with this sec- avoid stoppage by dirt or ice forma- tion may be shown with the means pro- tion; vided to comply with § 25.979(b).

(2) The vent arrangement must pre- [Amdt. 25–11, 32 FR 6913, May 5, 1967] vent siphoning of fuel during normal operation; § 25.971 Fuel tank sump.

(3) The venting capacity and vent (a) Each fuel tank must have a sump pressure levels must maintain accept- with an effective capacity, in the nor- able differences of pressure between mal ground attitude, of not less than the interior and exterior of the tank, the greater of 0.10 percent of the tank during— capacity or one-sixteenth of a gallon (i) Normal flight operation; unless operating limitations are estab- (ii) Maximum rate of ascent and de- lished to ensure that the accumulation scent; and of water in service will not exceed the (iii) Refueling and defueling (where sump capacity. applicable); 14 CFR Ch. I (1–1–25 Edition) § 25.977 (4) Airspaces of tanks with inter- (e) Each finger strainer must be ac- connected outlets must be inter- cessible for inspection and cleaning.

connected; [Amdt. 25–11, 32 FR 6913, May 5, 1967, as (5) There may be no point in any vent amended by Amdt. 25–36, 39 FR 35460, Oct. 1, line where moisture can accumulate 1974] with the airplane in the ground atti- § 25.979 Pressure fueling system.

tude or the level flight attitude, unless drainage is provided; For pressure fueling systems, the fol- (6) No vent or drainage provision may lowing apply: end at any point— (a) Each pressure fueling system fuel (i) Where the discharge of fuel from manifold connection must have means the vent outlet would constitute a fire to prevent the escape of hazardous hazard; or quantities of fuel from the system if the fuel entry valve fails.

(ii) From which fumes could enter (b) An automatic shutoff means must personnel compartments; and be provided to prevent the quantity of (7) Each fuel tank vent system must fuel in each tank from exceeding the prevent explosions, for a minimum of 2 maximum quantity approved for that minutes and 30 seconds, caused by tank. This means must— propagation of flames from outside the (1) Allow checking for proper shutoff tank through the fuel tank vents into operation before each fueling of the fuel tank vapor spaces when any fuel tank; and tank vent is continuously exposed to (2) Provide indication at each fueling flame.

station of failure of the shutoff means (b) Carburetor vapor vents. Each car- to stop the fuel flow at the maximum buretor with vapor elimination connec- quantity approved for that tank.

tions must have a vent line to lead va- (c) A means must be provided to pre- pors back to one of the fuel tanks. In vent damage to the fuel system in the addition— event of failure of the automatic shut- (1) Each vent system must have off means prescribed in paragraph (b) means to avoid stoppage by ice; and of this section.

(2) If there is more than one fuel (d) The airplane pressure fueling sys- tank, and it is necessary to use the tem (not including fuel tanks and fuel tanks in a definite sequence, each tank vents) must withstand an ulti- vapor vent return line must lead back mate load that is 2.0 times the load to the fuel tank used for takeoff and arising from the maximum pressures, landing.

including surge, that is likely to occur [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as during fueling. The maximum surge amended by Docket No. FAA–2014–0500, pressure must be established with any Amdt. No. 25–143, 81 FR 41207, June 24, 2016] combination of tank valves being ei- ther intentionally or inadvertently § 25.977 Fuel tank outlet.

closed.

(a) There must be a fuel strainer for (e) The airplane defueling system the fuel tank outlet or for the booster (not including fuel tanks and fuel tank pump. This strainer must— vents) must withstand an ultimate (1) For reciprocating engine powered load that is 2.0 times the load arising airplanes, have 8 to 16 meshes per inch; from the maximum permissible and defueling pressure (positive or nega- (2) For turbine engine powered air- tive) at the airplane fueling connec- planes, prevent the passage of any ob- tion.

ject that could restrict fuel flow or [Amdt. 25–11, 32 FR 6913, May 5, 1967, as damage any fuel system component.

amended by Amdt. 25–38, 41 FR 55467, Dec. 20, (b) [Reserved] 1976; Amdt. 25–72, 55 FR 29785, July 20, 1990] (c) The clear area of each fuel tank § 25.981 Fuel tank explosion preven- outlet strainer must be at least five tion.

times the area of the outlet line.

(d) The diameter of each strainer (a) No ignition source may be present must be at least that of the fuel tank at each point in the fuel tank or fuel outlet. tank system where catastrophic failure Federal Aviation Administration, DOT § 25.981 could occur due to ignition of fuel or (i) Equivalent Conventional Unheated vapors. This must be shown by: Aluminum Wing Tank is an integral (1) Determining the highest tempera- tank in an unheated semi-monocoque ture allowing a safe margin below the aluminum wing of a subsonic airplane lowest expected autoignition tempera- that is equivalent in aerodynamic per- ture of the fuel in the fuel tanks.

formance, structural capability, fuel (2) Demonstrating that no tempera- tank capacity and tank configuration ture at each place inside each fuel tank to the designed wing.

where fuel ignition is possible will ex- (ii) Fleet Average Flammability Expo- ceed the temperature determined under sure is defined in Appendix N to this paragraph (a)(1) of this section. This part and means the percentage of time must be verified under all probable op- each fuel tank ullage is flammable for erating, failure, and malfunction con- a fleet of an airplane type operating ditions of each component whose oper- over the range of flight lengths.

ation, failure, or malfunction could in- (iii) Main Fuel Tank means a fuel crease the temperature inside the tank.

tank that feeds fuel directly into one (3) Except for ignition sources due to or more engines and holds required fuel lightning addressed by § 25.954, dem- reserves continually throughout each onstrating that an ignition source flight.

could not result from each single fail- ure, from each single failure in com- (c) Paragraph (b) of this section does bination with each latent failure condi- not apply to a fuel tank if means are tion not shown to be extremely remote, provided to mitigate the effects of an and from all combinations of failures ignition of fuel vapors within that fuel not shown to be extremely improbable, tank such that no damage caused by an taking into account the effects of man- ignition will prevent continued safe ufacturing variability, aging, wear, flight and landing.

corrosion, and likely damage.

(d) To protect design features that (b) Except as provided in paragraphs prevent catastrophic ignition sources (b)(2) and (c) of this section, no fuel within the fuel tank or fuel tank sys- tank Fleet Average Flammability Ex- tem according to paragraph (a) of this posure on an airplane may exceed three section, and to prevent increasing the percent of the Flammability Exposure flammability exposure of the tanks Evaluation Time (FEET) as defined in above that permitted in paragraph (b) Appendix N of this part, or that of a of this section, the type design must fuel tank within the wing of the air- include critical design configuration plane model being evaluated, which- control limitations (CDCCLs) identi- ever is greater. If the wing is not a con- fying those features and providing in- ventional unheated aluminum wing, structions on how to protect them. To the analysis must be based on an as- ensure the continued effectiveness of sumed Equivalent Conventional those features, and prevent degrada- Unheated Aluminum Wing Tank.

(1) Fleet Average Flammability Ex- tion of the performance and reliability posure is determined in accordance of any means provided according to with Appendix N of this part. The as- paragraphs (a), (b), or (c) of this sec- sessment must be done in accordance tion, the type design must also include with the methods and procedures set necessary inspection and test proce- forth in the Fuel Tank Flammability dures, intervals between repetitive in- Assessment Method User’s Manual, spections and tests, and mandatory re- dated May 2008, document number placement times for those features.

DOT/FAA/AR–05/8 (incorporated by ref- The applicant must include informa- erence, see § 25.5).

tion required by this paragraph in the (2) Any fuel tank other than a main Airworthiness Limitations section of fuel tank on an airplane must meet the the Instructions for Continued Air- flammability exposure criteria of Ap- worthiness required by § 25.1529. The pendix M to this part if any portion of type design must also include visible the tank is located within the fuselage means of identifying critical features contour.

of the design in areas of the airplane (3) As used in this paragraph, 14 CFR Ch. I (1–1–25 Edition) § 25.991 where foreseeable maintenance ac- a reasonable degree of deformation and tions, repairs, or alterations may com- stretching without leakage.

promise the CDCCLs.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–15, 32 FR 13266, Sept.

[Doc. No. 1999–6411, 66 FR 23129, May 7, 2001, 20, 1967] as amended by Doc. No. FAA–2005–22997, 73 FR 42494, July 21, 2008; Doc. No. FAA– 2014– § 25.994 Fuel system components.

1027, Amdt. No. 25–146, 83 FR 47556, Sept. 20, 2018] Fuel system components in an engine nacelle or in the fuselage must be pro- F UEL S YSTEM C OMPONENTS tected from damage that could result in spillage of enough fuel to constitute § 25.991 Fuel pumps.

a fire hazard as a result of a wheels-up landing on a paved runway under each (a) Main pumps. Each fuel pump re- of the conditions prescribed in quired for proper engine operation, or § 25.721(b).

required to meet the fuel system re- quirements of this subpart (other than [Amdt. 25–139, 79 FR 59430, Oct. 2, 2014] those in paragraph (b) of this section, § 25.995 Fuel valves.

is a main pump. For each main pump, provision must be made to allow the In addition to the requirements of bypass of each positive displacement § 25.1189 for shutoff means, each fuel fuel pump other than a fuel injection valve must— pump (a pump that supplies the proper (a) [Reserved] flow and pressure for fuel injection (b) Be supported so that no loads re- when the injection is not accomplished sulting from their operation or from in a carburetor) approved as part of the accelerated flight conditions are trans- engine.

mitted to the lines attached to the (b) Emergency pumps. There must be valve.

emergency pumps or another main [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as pump to feed each engine immediately amended by Amdt. 25–40, 42 FR 15043, Mar. 17, after failure of any main pump (other 1977] than a fuel injection pump approved as § 25.997 Fuel strainer or filter.

part of the engine).

There must be a fuel strainer or filter § 25.993 Fuel system lines and fittings.

between the fuel tank outlet and the inlet of either the fuel metering device (a) Each fuel line must be installed or an engine driven positive displace- and supported to prevent excessive vi- ment pump, whichever is nearer the bration and to withstand loads due to fuel tank outlet. This fuel strainer or fuel pressure and accelerated flight filter must— conditions.

(a) Be accessible for draining and (b) Each fuel line connected to com- cleaning and must incorporate a screen ponents of the airplane between which or element which is easily removable; relative motion could exist must have (b) Have a sediment trap and drain provisions for flexibility.

except that it need not have a drain if (c) Each flexible connection in fuel the strainer or filter is easily remov- lines that may be under pressure and able for drain purposes; subjected to axial loading must use (c) Be mounted so that its weight is flexible hose assemblies.

not supported by the connecting lines (d) Flexible hose must be approved or or by the inlet or outlet connections of must be shown to be suitable for the the strainer or filter itself, unless ade- particular application.

quate strength margins under all load- (e) No flexible hose that might be ad- ing conditions are provided in the lines versely affected by exposure to high and connections; and temperatures may be used where exces- (d) Have the capacity (with respect to sive temperatures will exist during op- operating limitations established for eration or after engine shut-down.

the engine) to ensure that engine fuel (f) Each fuel line within the fuselage system functioning is not impaired, must be designed and installed to allow with the fuel contaminated to a degree Federal Aviation Administration, DOT § 25.1001 (with respect to particle size and den- off weight with flaps and landing gear sity) that is greater than that estab- up and in— lished for the engine in Part 33 of this (1) A power-off glide at 1.3 V ; SR1 chapter. (2) A climb at the one-engine inoper- ative best rate-of-climb speed, with the [Amdt. 25–36, 39 FR 35460, Oct. 1, 1974, as critical engine inoperative and the re- amended by Amdt. 25–57, 49 FR 6848, Feb. 23, maining engines at maximum contin- 1984] uous power; and § 25.999 Fuel system drains.

(3) Level flight at 1.3 V ; if the re- SR1 sults of the tests in the conditions (a) Drainage of the fuel system must specified in paragraphs (c)(1) and (2) of be accomplished by the use of fuel this section show that this condition strainer and fuel tank sump drains.

could be critical.

(b) Each drain required by paragraph (d) During the flight tests prescribed (a) of this section must— in paragraph (c) of this section, it must (1) Discharge clear of all parts of the be shown that— airplane; (1) The fuel jettisoning system and (2) Have manual or automatic means its operation are free from fire hazard; for positive locking in the closed posi- (2) The fuel discharges clear of any tion; and (3) Have a drain valve— part of the airplane; (i) That is readily accessible and (3) Fuel or fumes do not enter any which can be easily opened and closed; parts of the airplane; and and (4) The jettisoning operation does not (ii) That is either located or pro- adversely affect the controllability of tected to prevent fuel spillage in the the airplane.

event of a landing with landing gear re- (e) For reciprocating engine powered tracted. airplanes, means must be provided to prevent jettisoning the fuel in the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tanks used for takeoff and landing amended by Amdt. 25–38, 41 FR 55467, Dec. 20, below the level allowing 45 minutes 1976] flight at 75 percent maximum contin- § 25.1001 Fuel jettisoning system.

uous power. However, if there is an auxiliary control independent of the (a) A fuel jettisoning system must be main jettisoning control, the system installed on each airplane unless it is may be designed to jettison the re- shown that the airplane meets the maining fuel by means of the auxiliary climb requirements of §§ 25.119 and jettisoning control.

25.121(d) at maximum takeoff weight, (f) For turbine engine powered air- less the actual or computed weight of planes, means must be provided to pre- fuel necessary for a 15-minute flight vent jettisoning the fuel in the tanks comprised of a takeoff, go-around, and used for takeoff and landing below the landing at the airport of departure level allowing climb from sea level to with the airplane configuration, speed, 10,000 feet and thereafter allowing 45 power, and thrust the same as that minutes cruise at a speed for maximum used in meeting the applicable takeoff, range. However, if there is an auxiliary approach, and landing climb perform- ance requirements of this part. control independent of the main jetti- (b) If a fuel jettisoning system is re- soning control, the system may be de- quired it must be capable of jettisoning signed to jettison the remaining fuel enough fuel within 15 minutes, starting by means of the auxiliary jettisoning with the weight given in paragraph (a) control.

of this section, to enable the airplane (g) The fuel jettisoning valve must be to meet the climb requirements of designed to allow flight personnel to §§ 25.119 and 25.121(d), assuming that the close the valve during any part of the fuel is jettisoned under the conditions, jettisoning operation.

except weight, found least favorable (h) Unless it is shown that using any during the flight tests prescribed in means (including flaps, slots, and slats) paragraph (c) of this section. for changing the airflow across or (c) Fuel jettisoning must be dem- around the wings does not adversely af- onstrated beginning at maximum take- fect fuel jettisoning, there must be a 14 CFR Ch. I (1–1–25 Edition) § 25.1011 placard, adjacent to the jettisoning lon, and each oil tank used with a tur- control, to warn flight crewmembers bine engine must have an expansion against jettisoning fuel while the space of not less than 10 percent of the means that change the airflow are tank capacity.

being used. (2) Each reserve oil tank not directly (i) The fuel jettisoning system must connected to any engine may have an be designed so that any reasonably expansion space of not less than two probable single malfunction in the sys- percent of the tank capacity.

tem will not result in a hazardous con- (3) It must be impossible to fill the dition due to unsymmetrical jetti- expansion space inadvertently with the soning of, or inability to jettison, fuel. airplane in the normal ground attitude.

(c) Filler connection. Each recessed oil [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tank filler connection that can retain amended by Amdt. 25–18, 33 FR 12226, Aug. 30, any appreciable quantity of oil must 1968; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; have a drain that discharges clear of Amdt. 25–108, 67 FR 70827, Nov. 26, 2002] each part of the airplane. In addition, O IL S YSTEM each oil tank filler cap must provide an oil-tight seal.

§ 25.1011 General.

(d) Vent. Oil tanks must be vented as (a) Each engine must have an inde- follows: pendent oil system that can supply it (1) Each oil tank must be vented with an appropriate quantity of oil at a from the top part of the expansion temperature not above that safe for space so that venting is effective under continuous operation.

any normal flight condition.

(b) The usable oil capacity may not (2) Oil tank vents must be arranged be less than the product of the endur- so that condensed water vapor that ance of the airplane under critical op- might freeze and obstruct the line can- erating conditions and the approved not accumulate at any point.

maximum allowable oil consumption of (e) Outlet. There must be means to the engine under the same conditions, prevent entrance into the tank itself, plus a suitable margin to ensure sys- or into the tank outlet, of any object tem circulation. Instead of a rational that might obstruct the flow of oil analysis of airplane range for the pur- through the system. No oil tank outlet pose of computing oil requirements for may be enclosed by any screen or guard reciprocating engine powered air- that would reduce the flow of oil below planes, the following fuel/oil ratios a safe value at any operating tempera- may be used: ture. There must be a shutoff valve at (1) For airplanes without a reserve the outlet of each oil tank used with a oil or oil transfer system, a fuel/oil turbine engine, unless the external por- ratio of 30:1 by volume.

tion of the oil system (including the oil (2) For airplanes with either a re- tank supports) is fireproof.

serve oil or oil transfer system, a fuel/ (f) Flexible oil tank liners. Each flexi- oil ratio of 40:1 by volume.

ble oil tank liner must be approved or (c) Fuel/oil ratios higher than those must be shown to be suitable for the prescribed in paragraphs (b)(1) and (2) particular application.

of this section may be used if substan- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tiated by data on actual engine oil con- amended by Amdt. 25–19, 33 FR 15410, Oct. 17, sumption.

1968; Amdt. 25–23, 35 FR 5677, Apr. 8, 1970; Amdt. 25–36, 39 FR 35460, Oct. 1, 1974; Amdt.

§ 25.1013 Oil tanks.

25–57, 49 FR 6848, Feb. 23, 1984; Amdt. 25–72, 55 FR 29785, July 20, 1990] (a) Installation. Each oil tank instal- lation must meet the requirements of § 25.1015 Oil tank tests.

§ 25.967.

(b) Expansion space. Oil tank expan- Each oil tank must be designed and sion space must be provided as follows: installed so that— (1) Each oil tank used with a recipro- (a) It can withstand, without failure, cating engine must have an expansion each vibration, inertia, and fluid load space of not less than the greater of 10 that it may be subjected to in oper- percent of the tank capacity or 0.5 gal- ation; and Federal Aviation Administration, DOT § 25.1025 (b) It meets the provisions of § 25.965, (3) The oil strainer or filter, unless it except— is installed at an oil tank outlet, must (1) The test pressure— incorporate an indicator that will indi- cate contamination before it reaches (i) For pressurized tanks used with a the capacity established in accordance turbine engine, may not be less than 5 with paragraph (a)(2) of this section.

p.s.i. plus the maximum operating (4) The bypass of a strainer or filter pressure of the tank instead of the must be constructed and installed so pressure specified in § 25.965(a); and that the release of collected contami- (ii) For all other tanks may not be nants is minimized by appropriate lo- less than 5 p.s.i. instead of the pressure cation of the bypass to ensure that col- specified in § 25.965(a); and lected contaminants are not in the by- (2) The test fluid must be oil at 250 pass flow path.

° F. instead of the fluid specified in (5) An oil strainer or filter that has § 25.965(c).

no bypass, except one that is installed [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as at an oil tank outlet, must have a amended by Amdt. 25–36, 39 FR 35461, Oct. 1, means to connect it to the warning 1974] system required in § 25.1305(c)(7).

(b) Each oil strainer or filter in a § 25.1017 Oil lines and fittings.

powerplant installation using recipro- (a) Each oil line must meet the re- cating engines must be constructed and quirements of § 25.993 and each oil line installed so that oil will flow at the and fitting in any designated fire zone normal rate through the rest of the must meet the requirements of system with the strainer or filter ele- § 25.1183.

ment completely blocked.

(b) Breather lines must be arranged [Amdt. 25–36, 39 FR 35461, Oct. 1, 1974, as so that— amended by Amdt. 25–57, 49 FR 6848, Feb. 23, (1) Condensed water vapor that might 1984] freeze and obstruct the line cannot ac- cumulate at any point; § 25.1021 Oil system drains.

(2) The breather discharge does not A drain (or drains) must be provided constitute a fire hazard if foaming oc- to allow safe drainage of the oil sys- curs or causes emitted oil to strike the tem. Each drain must— pilot’s windshield; and (a) Be accessible; and (3) The breather does not discharge (b) Have manual or automatic means into the engine air induction system.

for positive locking in the closed posi- tion.

§ 25.1019 Oil strainer or filter.

[Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] (a) Each turbine engine installation must incorporate an oil strainer or fil- § 25.1023 Oil radiators.

ter through which all of the engine oil flows and which meets the following re- (a) Each oil radiator must be able to quirements: withstand, without failure, any vibra- tion, inertia, and oil pressure load to (1) Each oil strainer or filter that has which it would be subjected in oper- a bypass must be constructed and in- ation.

stalled so that oil will flow at the nor- (b) Each oil radiator air duct must be mal rate through the rest of the sys- located so that, in case of fire, flames tem with the strainer or filter com- coming from normal openings of the pletely blocked.

engine nacelle cannot impinge directly (2) The oil strainer or filter must upon the radiator.

have the capacity (with respect to op- erating limitations established for the § 25.1025 Oil valves.

engine) to ensure that engine oil sys- tem functioning is not impaired when (a) Each oil shutoff must meet the re- the oil is contaminated to a degree quirements of § 25.1189.

(with respect to particle size and den- (b) The closing of oil shutoff means sity) that is greater than that estab- may not prevent propeller feathering.

lished for the engine under Part 33 of (c) Each oil valve must have positive this chapter. stops or suitable index provisions in 14 CFR Ch. I (1–1–25 Edition) § 25.1027 the ‘‘on’’ and ‘‘off’’ positions and must tures must be corrected under para- be supported so that no loads resulting graphs (c) and (d) of this section.

from its operation or from accelerated (2) No corrected temperatures deter- flight conditions are transmitted to mined under paragraph (a)(1) of this the lines attached to the valve.

section may exceed established limits.

(3) For reciprocating engines, the fuel § 25.1027 Propeller feathering system.

used during the cooling tests must be (a) If the propeller feathering system the minimum grade approved for the depends on engine oil, there must be engines, and the mixture settings must means to trap an amount of oil in the be those normally used in the flight tank if the supply becomes depleted stages for which the cooling tests are due to failure of any part of the lubri- conducted. The test procedures must be cating system other than the tank as prescribed in § 25.1045.

itself.

(b) Maximum ambient atmospheric tem- (b) The amount of trapped oil must perature. A maximum ambient atmos- be enough to accomplish the feathering pheric temperature corresponding to operation and must be available only sea level conditions of at least 100 de- to the feathering pump.

grees F must be established. The as- (c) The ability of the system to ac- sumed temperature lapse rate is 3.6 de- complish feathering with the trapped grees F per thousand feet of altitude oil must be shown. This may be done above sea level until a temperature of on the ground using an auxiliary ¥ 69.7 degrees F is reached, above which source of oil for lubricating the engine altitude the temperature is considered during operation.

constant at ¥ 69.7 degrees F. However, (d) Provision must be made to pre- vent sludge or other foreign matter for winterization installations, the ap- from affecting the safe operation of the plicant may select a maximum ambi- propeller feathering system.

ent atmospheric temperature cor- responding to sea level conditions of [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as less than 100 degrees F.

amended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976] (c) Correction factor (except cylinder barrels). Unless a more rational correc- C OOLING tion applies, temperatures of engine fluids and powerplant components (ex- § 25.1041 General.

cept cylinder barrels) for which tem- The powerplant and auxiliary power perature limits are established, must unit cooling provisions must be able to be corrected by adding to them the dif- maintain the temperatures of power- ference between the maximum ambient plant components, engine fluids, and atmospheric temperature and the tem- auxiliary power unit components and perature of the ambient air at the time fluids within the temperature limits of the first occurrence of the maximum established for these components and component or fluid temperature re- fluids, under ground, water, and flight corded during the cooling test.

operating conditions, and after normal (d) Correction factor for cylinder barrel engine or auxiliary power unit shut- temperatures. Unless a more rational down, or both.

correction applies, cylinder barrel tem- [Amdt. 25–38, 41 FR 55467, Dec. 20, 1976] peratures must be corrected by adding to them 0.7 times the difference be- § 25.1043 Cooling tests.

tween the maximum ambient atmos- (a) General. Compliance with § 25.1041 pheric temperature and the tempera- must be shown by tests, under critical ture of the ambient air at the time of ground, water, and flight operating the first occurrence of the maximum conditions. For these tests, the fol- cylinder barrel temperature recorded lowing apply: during the cooling test.

(1) If the tests are conducted under conditions deviating from the max- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as imum ambient atmospheric tempera- amended by Amdt. 25–42, 43 FR 2323, Jan. 16, 1978] ture, the recorded powerplant tempera- Federal Aviation Administration, DOT § 25.1091 (4) Critical engine inoperative and its § 25.1045 Cooling test procedures.

propeller stopped.

(a) Compliance with § 25.1041 must be (5) Remaining engines at the max- shown for the takeoff, climb, en route, imum continuous power available for and landing stages of flight that cor- the altitude.

respond to the applicable performance (e) For hull seaplanes and amphib- requirements. The cooling tests must ians, cooling must be shown during be conducted with the airplane in the taxiing downwind for 10 minutes, at configuration, and operating under the five knots above step speed.

conditions, that are critical relative to cooling during each stage of flight. For [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the cooling tests, a temperature is amended by Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] ‘‘stabilized’’ when its rate of change is less than two degrees F. per minute.

I NDUCTION S YSTEM (b) Temperatures must be stabilized under the conditions from which entry § 25.1091 Air induction.

is made into each stage of flight being (a) The air induction system for each investigated, unless the entry condi- engine and auxiliary power unit must tion normally is not one during which supply— component and the engine fluid tem- (1) The air required by that engine peratures would stabilize (in which and auxiliary power unit under each case, operation through the full entry operating condition for which certifi- condition must be conducted before cation is requested; and entry into the stage of flight being in- (2) The air for proper fuel metering vestigated in order to allow tempera- and mixture distribution with the in- tures to reach their natural levels at duction system valves in any position.

the time of entry). The takeoff cooling (b) Each reciprocating engine must test must be preceded by a period dur- have an alternate air source that pre- ing which the powerplant component vents the entry of rain, ice, or any and engine fluid temperatures are sta- other foreign matter.

bilized with the engines at ground idle.

(c) Air intakes may not open within (c) Cooling tests for each stage of the cowling, unless— flight must be continued until— (1) That part of the cowling is iso- (1) The component and engine fluid lated from the engine accessory section temperatures stabilize; by means of a fireproof diaphragm; or (2) The stage of flight is completed; (2) For reciprocating engines, there or are means to prevent the emergence of (3) An operating limitation is backfire flames.

reached.

(d) For turbine engine powered air- (d) For reciprocating engine powered planes and airplanes incorporating aux- airplanes, it may be assumed, for cool- iliary power units— ing test purposes, that the takeoff (1) There must be means to prevent stage of flight is complete when the hazardous quantities of fuel leakage or airplane reaches an altitude of 1,500 overflow from drains, vents, or other feet above the takeoff surface or components of flammable fluid systems reaches a point in the takeoff where from entering the engine or auxiliary the transition from the takeoff to the power unit intake system; and en route configuration is completed (2) The airplane must be designed to and a speed is reached at which compli- prevent water or slush on the runway, ance with § 25.121(c) is shown, which- taxiway, or other airport operating ever point is at a higher altitude. The surfaces from being directed into the airplane must be in the following con- engine or auxiliary power unit air inlet figuration: (1) Landing gear retracted. ducts in hazardous quantities, and the (2) Wing flaps in the most favorable air inlet ducts must be located or pro- position. tected so as to minimize the ingestion (3) Cowl flaps (or other means of con- of foreign matter during takeoff, land- trolling the engine cooling supply) in ing, and taxiing.

the position that provides adequate (e) If the engine induction system cooling in the hot-day condition. contains parts or components that 14 CFR Ch. I (1–1–25 Edition) § 25.1093 could be damaged by foreign objects nents, or airframe components that entering the air inlet, it must be shown would do any of the following: by tests or, if appropriate, by analysis (i) Adversely affect installed engine that the induction system design can operation or cause a sustained loss of withstand the foreign object ingestion power or thrust; or an unacceptable in- crease in gas path operating tempera- test conditions of §§ 33.76, 33.77 and ture; or an airframe/engine incompati- 33.78(a)(1) of this chapter without fail- bility; or ure of parts or components that could (ii) Result in unacceptable temporary create a hazard.

power loss or engine damage; or [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (iii) Cause a stall, surge, or flameout amended by Amdt. 25–38, 41 FR 55467, Dec. 20, or loss of engine controllability (for ex- 1976; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977; ample, rollback).

Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt.

(2) Operate at ground idle speed for a 25–100, 65 FR 55854, Sept. 14, 2000] minimum of 30 minutes on the ground in the following icing conditions shown § 25.1093 Induction system icing pro- tection. in Table 1 of this section, unless re- placed by similar test conditions that (a) Reciprocating engines. Each recip- are more critical. These conditions rocating engine air induction system must be demonstrated with the avail- must have means to prevent and elimi- able air bleed for icing protection at its nate icing. Unless this is done by other critical condition, without adverse ef- means, it must be shown that, in air fect, followed by an acceleration to free of visible moisture at a tempera- takeoff power or thrust in accordance ture of 30 F., each airplane with alti- with the procedures defined in the air- tude engines using— plane flight manual. During the idle (1) Conventional venturi carburetors operation, the engine may be run up have a preheater that can provide a periodically to a moderate power or heat rise of 120 F. with the engine at 60 thrust setting in a manner acceptable percent of maximum continuous power; to the Administrator. Analysis may be or used to show ambient temperatures (2) Carburetors tending to reduce the below the tested temperature are less probability of ice formation has a pre- critical. The applicant must document heater that can provide a heat rise of the engine run-up procedure (including 100 ° F. with the engine at 60 percent of the maximum time interval between maximum continuous power.

run-ups from idle, run-up power set- (b) Turbine engines. Except as pro- ting, and duration at power), the asso- vided in paragraph (b)(3) of this sec- ciated minimum ambient temperature, tion, each engine, with all icing protec- and the maximum time interval. These tion systems operating, must: conditions must be used in the analysis (1) Operate throughout its flight that establishes the airplane operating power range, including the minimum limitations in accordance with § 25.1521.

descent idling speeds, in the icing con- (3) For the purposes of this section, ditions defined in Appendices C and O the icing conditions defined in appen- of this part, and Appendix D of part 33 dix O of this part, including the condi- of this chapter, and in falling and blow- tions specified in Condition 3 of Table 1 ing snow within the limitations estab- of this section, are not applicable to lished for the airplane for such oper- airplanes with a maximum takeoff ation, without the accumulation of ice weight equal to or greater than 60,000 on the engine, inlet system compo- pounds.

T ABLE 1—I CING CONDITIONS FOR G ROUND T ESTS Water concentration Mean effective par- Condition Total air temperature Demonstration (minimum) ticle diameter 1. Rime ice condition 0 to 15 ° F (18 to ¥ 9 ° C) Liquid—0.3 g/m ........ 15–25 microns ........... By test, analysis or com- bination of the two.

2. Glaze ice condition 20 to 30 ° F ( ¥ 7 to ¥ 1 Liquid—0.3 g/m ........ 15–25 microns ........... By test, analysis or com- ° C). bination of the two.

Federal Aviation Administration, DOT § 25.1105 T ABLE 1—I CING C ONDITIONS FOR G ROUND T ESTS —Continued Water concentration Mean effective par- Condition Total air temperature Demonstration (minimum) ticle diameter 3. Large drop condi- 15 to 30 ° F ( ¥ 9 to ¥ 1 Liquid—0.3 g/m ........ 100 microns (min- By test, analysis or com- tion. ° C). imum). bination of the two.

(c) Supercharged reciprocating engines. within the auxiliary power unit fire For each engine having a supercharger zone.

to pressurize the air before it enters (c) Each duct connected to compo- the carburetor, the heat rise in the air nents between which relative motion caused by that supercharging at any could exist must have means for flexi- altitude may be utilized in determining bility.

compliance with paragraph (a) of this (d) For turbine engine and auxiliary section if the heat rise utilized is that power unit bleed air duct systems, no which will be available, automatically, hazard may result if a duct failure oc- for the applicable altitude and oper- curs at any point between the air duct ating condition because of super- source and the airplane unit served by charging.

the air.

(e) Each auxiliary power unit induc- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tion system duct must be fireproof for amended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977; a sufficient distance upstream of the Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt.

auxiliary power unit compartment to 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–140, prevent hot gas reverse flow from burn- 79 FR 65526, Nov. 4, 2014] ing through auxiliary power unit ducts and entering any other compartment § 25.1101 Carburetor air preheater de- or area of the airplane in which a haz- sign.

ard would be created resulting from the Each carburetor air preheater must entry of hot gases. The materials used be designed and constructed to— to form the remainder of the induction (a) Ensure ventilation of the pre- system duct and plenum chamber of heater when the engine is operated in the auxiliary power unit must be capa- cold air; ble of resisting the maximum heat con- (b) Allow inspection of the exhaust ditions likely to occur.

manifold parts that it surrounds; and (f) Each auxiliary power unit induc- (c) Allow inspection of critical parts tion system duct must be constructed of the preheater itself.

of materials that will not absorb or trap hazardous quantities of flammable § 25.1103 Induction system ducts and fluids that could be ignited in the air duct systems.

event of a surge or reverse flow condi- (a) Each induction system duct up- tion.

stream of the first stage of the engine [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as supercharger and of the auxiliary amended by Amdt. 25–46, 43 FR 50597, Oct. 30, power unit compressor must have a 1978] drain to prevent the hazardous accu- mulation of fuel and moisture in the § 25.1105 Induction system screens.

ground attitude. No drain may dis- If induction system screens are charge where it might cause a fire haz- used— ard.

(a) Each screen must be upstream of (b) Each induction system duct must the carburetor; be— (b) No screen may be in any part of (1) Strong enough to prevent induc- the induction system that is the only tion system failures resulting from passage through which air can reach normal backfire conditions; and the engine, unless it can be deiced by (2) Fire-resistant if it is in any fire heated air; zone for which a fire-extinguishing sys- tem is required, except that ducts for (c) No screen may be deiced by alco- auxiliary power units must be fireproof hol alone; and 14 CFR Ch. I (1–1–25 Edition) § 25.1107 (d) It must be impossible for fuel to enough to ignite any flammable fluids strike any screen. or vapors external to the shroud.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.1107 Inter-coolers and after-cool- amended by Amdt. 25–40, 42 FR 15043, Mar. 17, ers.

1977] Each inter-cooler and after-cooler must be able to withstand any vibra- § 25.1123 Exhaust piping.

tion, inertia, and air pressure load to For powerplant and auxiliary power which it would be subjected in oper- unit installations, the following apply: ation.

(a) Exhaust piping must be heat and corrosion resistant, and must have pro- E XHAUST S YSTEM visions to prevent failure due to expan- sion by operating temperatures.

§ 25.1121 General.

(b) Piping must be supported to with- For powerplant and auxiliary power stand any vibration and inertia loads unit installations the following apply: to which it would be subjected in oper- (a) Each exhaust system must ensure ation; and safe disposal of exhaust gases without (c) Piping connected to components fire hazard or carbon monoxide con- between which relative motion could tamination in any personnel compart- exist must have means for flexibility.

ment. For test purposes, any accept- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as able carbon monoxide detection meth- amended by Amdt. 25–40, 42 FR 15044, Mar. 17, od may be used to show the absence of 1977] carbon monoxide.

(b) Each exhaust system part with a § 25.1125 Exhaust heat exchangers.

surface hot enough to ignite flammable For reciprocating engine powered fluids or vapors must be located or airplanes, the following apply: shielded so that leakage from any sys- (a) Each exhaust heat exchanger tem carrying flammable fluids or va- must be constructed and installed to pors will not result in a fire caused by withstand each vibration, inertia, and impingement of the fluids or vapors on other load to which it would be sub- any part of the exhaust system includ- ing shields for the exhaust system. jected in operation. In addition— (1) Each exchanger must be suitable (c) Each component that hot exhaust for continued operation at high tem- gases could strike, or that could be peratures and resistant to corrosion subjected to high temperatures from exhaust system parts, must be fire- from exhaust gases; (2) There must be means for the in- proof. All exhaust system components must be separated by fireproof shields spection of the critical parts of each exchanger; from adjacent parts of the airplane that are outside the engine and auxil- (3) Each exchanger must have cooling iary power unit compartments. provisions wherever it is subject to (d) No exhaust gases may discharge contact with exhaust gases; and so as to cause a fire hazard with re- (4) No exhaust heat exchanger or spect to any flammable fluid vent or muff may have any stagnant areas or drain. liquid traps that would increase the (e) No exhaust gases may discharge probability of ignition of flammable fluids or vapors that might be present where they will cause a glare seriously affecting pilot vision at night. in case of the failure or malfunction of components carrying flammable fluids.

(f) Each exhaust system component must be ventilated to prevent points of (b) If an exhaust heat exchanger is used for heating ventilating air— excessively high temperature.

(g) Each exhaust shroud must be ven- (1) There must be a secondary heat exchanger between the primary ex- tilated or insulated to avoid, during normal operation, a temperature high haust gas heat exchanger and the ven- tilating air system; or Federal Aviation Administration, DOT § 25.1143 (2) Other means must be used to pre- (e) The portion of each powerplant clude the harmful contamination of the control located in a designated fire ventilating air. zone that is required to be operated in the event of fire must be at least fire [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as resistant.

amended by Amdt. 25–38, 41 FR 55467, Dec. 20, (f) For powerplant valve controls lo- 1976] cated in the flight deck there must be § 25.1127 Exhaust driven turbo-super- a means: chargers.

(1) For the flightcrew to select each intended position or function of the (a) Each exhaust driven turbo-super- valve; and charger must be approved or shown to (2) To indicate to the flightcrew: be suitable for the particular applica- (i) The selected position or function tion. It must be installed and sup- of the valve; and ported to ensure safe operation be- tween normal inspections and over- (ii) When the valve has not responded hauls. In addition, there must be provi- as intended to the selected position or sions for expansion and flexibility be- function.

tween exhaust conduits and the tur- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as bine.

amended by Amdt. 25–40, 42 FR 15044, Mar. 17, (b) There must be provisions for lu- 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990; bricating the turbine and for cooling Amdt. 25–115, 69 FR 40527, July 2, 2004] turbine parts where temperatures are § 25.1142 Auxiliary power unit con- critical.

trols.

(c) If the normal turbo-supercharger control system malfunctions, the tur- Means must be provided on the flight bine speed may not exceed its max- deck for starting, stopping, and emer- imum allowable value. Except for the gency shutdown of each installed auxil- waste gate operating components, the iary power unit.

components provided for meeting this [Amdt. 25–46, 43 FR 50598, Oct. 30, 1978] requirement must be independent of the normal turbo-supercharger con- § 25.1143 Engine controls.

trols.

(a) There must be a separate power or P OWERPLANT CONTROLS AND thrust control for each engine.

A CCESSORIES (b) Power and thrust controls must be arranged to allow— § 25.1141 Powerplant controls: general.

(1) Separate control of each engine; and Each powerplant control must be lo- cated, arranged, and designed under (2) Simultaneous control of all en- §§ 25.777 through 25.781 and marked gines.

under § 25.1555. In addition, it must (c) Each power and thrust control meet the following requirements: must provide a positive and imme- (a) Each control must be located so diately responsive means of controlling that it cannot be inadvertently oper- its engine.

ated by persons entering, leaving, or (d) For each fluid injection (other moving normally in, the cockpit.

than fuel) system and its controls not (b) Each flexible control must be ap- provided and approved as part of the proved or must be shown to be suitable engine, the applicant must show that for the particular application. the flow of the injection fluid is ade- (c) Each control must have sufficient quately controlled.

strength and rigidity to withstand op- (e) If a power or thrust control incor- erating loads without failure and with- porates a fuel shutoff feature, the con- out excessive deflection. trol must have a means to prevent the inadvertent movement of the control (d) Each control must be able to into the shutoff position. The means maintain any set position without con- must— stant attention by flight crewmembers and without creep due to control loads (1) Have a positive lock or stop at the or vibration. idle position; and 14 CFR Ch. I (1–1–25 Edition) § 25.1145 (2) Require a separate and distinct (d) The propeller speed and pitch con- operation to place the control in the trols must be to the right of, and at shutoff position.

least one inch below, the pilot’s throt- tle controls.

[Amdt. 25–23, 35 FR 5677, Apr. 8, 1970, as amended by Amdt. 25–38, 41 FR 55467, Dec. 20, § 25.1153 Propeller feathering controls.

1976; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984] (a) There must be a separate pro- § 25.1145 Ignition switches.

peller feathering control for each pro- (a) Ignition switches must control peller. The control must have means to each engine ignition circuit on each prevent its inadvertent operation.

engine.

(b) If feathering is accomplished by (b) There must be means to quickly movement of the propeller pitch or shut off all ignition by the grouping of speed control lever, there must be switches or by a master ignition con- means to prevent the inadvertent trol.

movement of this lever to the feath- (c) Each group of ignition switches, ering position during normal oper- except ignition switches for turbine en- ation.

gines for which continuous ignition is [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as not required, and each master ignition amended by Amdt. 25–11, 32 FR 6913, May 5, control must have a means to prevent 1967] its inadvertent operation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.1155 Reverse thrust and propeller amended by Amdt. 25–40, 42 FR 15044 Mar. 17, pitch settings below the flight re- 1977] gime.

Each control for reverse thrust and § 25.1147 Mixture controls.

for propeller pitch settings below the (a) If there are mixture controls, flight regime must have means to pre- each engine must have a separate con- vent its inadvertent operation. The trol. The controls must be grouped and means must have a positive lock or arranged to allow— stop at the flight idle position and (1) Separate control of each engine; must require a separate and distinct and operation by the crew to displace the (2) Simultaneous control of all en- control from the flight regime (forward gines.

thrust regime for turbojet powered air- (b) Each intermediate position of the planes).

mixture controls that corresponds to a normal operating setting must be iden- [Amdt. 25–11, 32 FR 6913, May 5, 1967] tifiable by feel and sight.

(c) The mixture controls must be ac- § 25.1157 Carburetor air temperature cessible to both pilots. However, if controls.

there is a separate flight engineer sta- There must be a separate carburetor tion with a control panel, the controls air temperature control for each en- need be accessible only to the flight en- gine.

gineer.

§ 25.1159 Supercharger controls.

§ 25.1149 Propeller speed and pitch controls.

Each supercharger control must be accessible to the pilots or, if there is a (a) There must be a separate pro- separate flight engineer station with a peller speed and pitch control for each control panel, to the flight engineer.

propeller.

(b) The controls must be grouped and § 25.1161 Fuel jettisoning system con- arranged to allow— trols.

(1) Separate control of each pro- Each fuel jettisoning system control peller; and must have guards to prevent inad- (2) Simultaneous control of all pro- pellers. vertent operation. No control may be (c) The controls must allow synchro- near any fire extinguisher control or nization of all propellers. other control used to combat fire.

Federal Aviation Administration, DOT § 25.1181 (e) No ground wire for any engine § 25.1163 Powerplant accessories.

may be routed through a fire zone of (a) Each engine mounted accessory another engine unless each part of that must— wire within that zone is fireproof.

(1) Be approved for mounting on the (f) Each ignition system must be engine involved; independent of any electrical circuit, (2) Use the provisions on the engine not used for assisting, controlling, or for mounting; and analyzing the operation of that system.

(3) Be sealed to prevent contamina- (g) There must be means to warn ap- tion of the engine oil system and the propriate flight crewmembers if the accessory system.

malfunctioning of any part of the elec- (b) Electrical equipment subject to trical system is causing the continuous arcing or sparking must be installed to discharge of any battery necessary for minimize the probability of contact engine ignition.

with any flammable fluids or vapors (h) Each engine ignition system of a that might be present in a free state.

turbine powered airplane must be con- (c) If continued rotation of an engine- sidered an essential electrical load.

driven cabin supercharger or of any re- mote accessory driven by the engine is [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5677, Apr. 8, hazardous if malfunctioning occurs, 1970; Amdt. 25–72, 55 FR 29785, July 20, 1990] there must be means to prevent rota- tion without interfering with the con- § 25.1167 Accessory gearboxes.

tinued operation of the engine.

For airplanes equipped with an acces- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as sory gearbox that is not certificated as amended by Amdt. 25–57, 49 FR 6849, Feb. 23, part of an engine— 1984] (a) The engine with gearbox and con- necting transmissions and shafts at- § 25.1165 Engine ignition systems.

tached must be subjected to the tests (a) Each battery ignition system specified in § 33.49 or § 33.87 of this chap- must be supplemented by a generator ter, as applicable; that is automatically available as an (b) The accessory gearbox must meet alternate source of electrical energy to the requirements of §§ 33.25 and 33.53 or allow continued engine operation if 33.91 of this chapter, as applicable; and any battery becomes depleted.

(c) Possible misalignments and tor- (b) The capacity of batteries and gen- sional loadings of the gearbox, trans- erators must be large enough to meet mission, and shaft system, expected to the simultaneous demands of the en- result under normal operating condi- gine ignition system and the greatest tions must be evaluated.

demands of any electrical system com- ponents that draw electrical energy [Amdt. 25–38, 41 FR 55467, Dec. 20, 1976] from the same source.

(c) The design of the engine ignition P OWERPLANT F IRE P ROTECTION system must account for— § 25.1181 Designated fire zones; re- (1) The condition of an inoperative gions included.

generator; (2) The condition of a completely de- (a) Designated fire zones are— pleted battery with the generator run- (1) The engine power section; ning at its normal operating speed; and (2) The engine accessory section; (3) The condition of a completely de- (3) Except for reciprocating engines, pleted battery with the generator oper- any complete powerplant compartment ating at idling speed, if there is only in which no isolation is provided be- one battery. tween the engine power section and the (d) Magneto ground wiring (for sepa- engine accessory section; rate ignition circuits) that lies on the (4) Any auxiliary power unit com- engine side of the fire wall, must be in- partment; stalled, located, or protected, to mini- (5) Any fuel-burning heater and other mize the probability of simultaneous combustion equipment installation de- failure of two or more wires as a result scribed in § 25.859; of mechanical damage, electrical (6) The compressor and accessory sec- faults, or other cause. tions of turbine engines; and 14 CFR Ch. I (1–1–25 Edition) § 25.1182 (7) Combustor, turbine, and tailpipe (b) Paragraph (a) of this section does sections of turbine engine installations not apply to— that contain lines or components car- (1) Lines, fittings, and components rying flammable fluids or gases.

which are already approved as part of a (b) Each designated fire zone must type certificated engine; and meet the requirements of §§ 25.863, (2) Vent and drain lines, and their fit- 25.865, 25.867, 25.869, and 25.1185 through tings, whose failure will not result in, 25.1203.

or add to, a fire hazard.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) All components, including ducts, amended by Amdt. 25–11, 32 FR 6913, May 5, within a designated fire zone must be 1967; Amdt. 25–23, 35 FR 5677, Apr. 8, 1970; fireproof if, when exposed to or dam- Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt.

aged by fire, they could— 25–115, 69 FR 40527, July 2, 2004] (1) Result in fire spreading to other regions of the airplane; or § 25.1182 Nacelle areas behind fire- walls, and engine pod attaching (2) Cause unintentional operation of, structures containing flammable or inability to operate, essential serv- fluid lines.

ices or equipment.

(a) Each nacelle area immediately [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as behind the firewall, and each portion of amended by Amdt. 25–11, 32 FR 6913, May 5, any engine pod attaching structure 1967; Amdt. 25–36, 39 FR 35461, Oct. 1, 1974; containing flammable fluid lines, must Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt.

meet each requirement of §§ 25.1103(b), 25–101, 65 FR 79710, Dec. 19, 2000] 25.1165 (d) and (e), 25.1183, 25.1185(c), 25.1187, 25.1189, and 25.1195 through § 25.1185 Flammable fluids.

25.1203, including those concerning des- (a) Except for the integral oil sumps ignated fire zones. However, engine pod specified in § 25.1183(a), no tank or res- attaching structures need not contain ervoir that is a part of a system con- fire detection or extinguishing means.

taining flammable fluids or gases may (b) For each area covered by para- be in a designated fire zone unless the graph (a) of this section that contains fluid contained, the design of the sys- a retractable landing gear, compliance tem, the materials used in the tank, with that paragraph need only be the shut-off means, and all connec- shown with the landing gear retracted.

tions, lines, and control provide a de- [Amdt. 25–11, 32 FR 6913, May 5, 1967] gree of safety equal to that which would exist if the tank or reservoir § 25.1183 Flammable fluid-carrying were outside such a zone.

components.

(b) There must be at least one-half (a) Except as provided in paragraph inch of clear airspace between each (b) of this section, each line, fitting, tank or reservoir and each firewall or and other component carrying flam- shroud isolating a designated fire zone.

mable fluid in any area subject to en- (c) Absorbent materials close to gine fire conditions, and each compo- flammable fluid system components nent which conveys or contains flam- that might leak must be covered or mable fluid in a designated fire zone treated to prevent the absorption of must be fire resistant, except that hazardous quantities of fluids.

flammable fluid tanks and supports in a designated fire zone must be fireproof [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–19, 33 FR 15410, Oct. 17, or be enclosed by a fireproof shield un- 1968; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] less damage by fire to any non-fire- proof part will not cause leakage or § 25.1187 Drainage and ventilation of spillage of flammable fluid. Compo- fire zones.

nents must be shielded or located to (a) There must be complete drainage safeguard against the ignition of leak- ing flammable fluid. An integral oil of each part of each designated fire sump of less than 25-quart capacity on zone to minimize the hazards resulting a reciprocating engine need not be fire- from failure or malfunctioning of any proof nor be enclosed by a fireproof component containing flammable shield. fluids. The drainage means must be— Federal Aviation Administration, DOT § 25.1193 (1) Effective under conditions ex- (e) No hazardous quantity of flam- pected to prevail when drainage is mable fluid may drain into any des- needed; and ignated fire zone after shutoff.

(2) Arranged so that no discharged (f) There must be means to guard fluid will cause an additional fire haz- against inadvertent operation of the ard. shutoff means and to make it possible (b) Each designated fire zone must be for the crew to reopen the shutoff ventilated to prevent the accumulation means in flight after it has been closed.

of flammable vapors. (g) Each tank-to-engine shutoff valve must be located so that the operation (c) No ventilation opening may be of the valve will not be affected by where it would allow the entry of flam- powerplant or engine mount structural mable fluids, vapors, or flame from failure.

other zones.

(h) Each shutoff valve must have a (d) Each ventilation means must be means to relieve excessive pressure ac- arranged so that no discharged vapors cumulation unless a means for pressure will cause an additional fire hazard.

relief is otherwise provided in the sys- (e) Unless the extinguishing agent ca- tem.

pacity and rate of discharge are based on maximum air flow through a zone, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as there must be means to allow the crew amended by Amdt. 25–23, 35 FR 5677, Apr. 8, to shut off sources of forced ventilation 1970; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984] to any fire zone except the engine § 25.1191 Firewalls.

power section of the nacelle and the combustion heater ventilating air (a) Each engine, auxiliary power ducts.

unit, fuel-burning heater, other com- bustion equipment intended for oper- § 25.1189 Shutoff means.

ation in flight, and the combustion, (a) Each engine installation and each turbine, and tailpipe sections of tur- fire zone specified in § 25.1181(a)(4) and bine engines, must be isolated from the (5) must have a means to shut off or rest of the airplane by firewalls, otherwise prevent hazardous quantities shrouds, or equivalent means.

of fuel, oil, deicer, and other flammable (b) Each firewall and shroud must fluids, from flowing into, within, or be— through any designated fire zone, ex- (1) Fireproof; cept that shutoff means are not re- (2) Constructed so that no hazardous quired for— quantity of air, fluid, or flame can pass (1) Lines, fittings, and components from the compartment to other parts forming an integral part of an engine; of the airplane; and (3) Constructed so that each opening (2) Oil systems for turbine engine in- is sealed with close fitting fireproof stallations in which all components of grommets, bushings, or firewall fit- the system in a designated fire zone, tings; and including oil tanks, are fireproof or lo- (4) Protected against corrosion.

cated in areas not subject to engine fire conditions.

§ 25.1192 Engine accessory section dia- (b) The closing of any fuel shutoff phragm.

valve for any engine may not make For reciprocating engines, the engine fuel unavailable to the remaining en- power section and all portions of the gines.

exhaust system must be isolated from (c) Operation of any shutoff may not the engine accessory compartment by a interfere with the later emergency op- diaphragm that complies with the fire- eration of other equipment, such as the wall requirements of § 25.1191.

means for feathering the propeller.

[Amdt. 25–23, 35 FR 5678, Apr. 8, 1970] (d) Each flammable fluid shutoff means and control must be fireproof or § 25.1193 Cowling and nacelle skin.

must be located and protected so that any fire in a fire zone will not affect its (a) Each cowling must be constructed operation. and supported so that it can resist any 14 CFR Ch. I (1–1–25 Edition) § 25.1195 vibration, inertia, and air load to heaters, and other combustion equip- which it may be subjected in operation. ment. For each other designated fire (b) Cowling must meet the drainage zone, two discharges must be provided and ventilation requirements of each of which produces adequate agent § 25.1187. concentration.

(c) The fire extinguishing system for (c) On airplanes with a diaphragm a nacelle must be able to simulta- isolating the engine power section from neously protect each zone of the na- the engine accessory section, each part celle for which protection is provided.

of the accessory section cowling sub- ject to flame in case of fire in the en- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as gine power section of the powerplant amended by Amdt. 25–46, 43 FR 50598, Oct. 30, must— 1978] (1) Be fireproof; and § 25.1197 Fire extinguishing agents.

(2) Meet the requirements of § 25.1191.

(d) Each part of the cowling subject (a) Fire extinguishing agents must— to high temperatures due to its near- (1) Be capable of extinguishing ness to exhaust system parts or ex- flames emanating from any burning of haust gas impingement must be fire- fluids or other combustible materials proof.

in the area protected by the fire extin- (e) Each airplane must— guishing system; and (1) Be designed and constructed so (2) Have thermal stability over the that no fire originating in any fire zone temperature range likely to be experi- can enter, either through openings or enced in the compartment in which by burning through external skin, any they are stored.

other zone or region where it would (b) If any toxic extinguishing agent is create additional hazards; used, provisions must be made to pre- (2) Meet paragraph (e)(1) of this sec- vent harmful concentrations of fluid or tion with the landing gear retracted (if fluid vapors (from leakage during nor- applicable); and mal operation of the airplane or as a (3) Have fireproof skin in areas sub- result of discharging the fire extin- ject to flame if a fire starts in the en- guisher on the ground or in flight) from gine power or accessory sections. entering any personnel compartment, even though a defect may exist in the § 25.1195 Fire extinguishing systems.

extinguishing system. This must be shown by test except for built-in car- (a) Except for combustor, turbine, bon dioxide fuselage compartment fire and tail pipe sections of turbine engine extinguishing systems for which— installations that contain lines or com- (1) Five pounds or less of carbon diox- ponents carrying flammable fluids or ide will be discharged, under estab- gases for which it is shown that a fire lished fire control procedures, into any originating in these sections can be fuselage compartment; or controlled, there must be a fire extin- (2) There is protective breathing guisher system serving each designated equipment for each flight crewmember fire zone.

on flight deck duty.

(b) The fire extinguishing system, the quantity of the extinguishing agent, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the rate of discharge, and the discharge amended by Amdt. 25–38, 41 FR 55467, Dec. 20, distribution must be adequate to extin- 1976; Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] guish fires. It must be shown by either § 25.1199 Extinguishing agent con- actual or simulated flights tests that tainers.

under critical airflow conditions in flight the discharge of the extin- (a) Each extinguishing agent con- guishing agent in each designated fire tainer must have a pressure relief to zone specified in paragraph (a) of this prevent bursting of the container by section will provide an agent con- excessive internal pressures.

centration capable of extinguishing (b) The discharge end of each dis- fires in that zone and of minimizing charge line from a pressure relief con- the probability of reignition. An indi- nection must be located so that dis- vidual ‘‘one-shot’’ system may be used charge of the fire extinguishing agent for auxiliary power units, fuel burning would not damage the airplane. The Federal Aviation Administration, DOT § 25.1207 line must also be located or protected sensor or associated wiring within a to prevent clogging caused by ice or designated fire zone, unless the system other foreign matter.

continues to function as a satisfactory (c) There must be a means for each detection system after the short cir- fire extinguishing agent container to cuit.

indicate that the container has dis- (c) No fire or overheat detector may charged or that the charging pressure be affected by any oil, water, other is below the established minimum nec- fluids or fumes that might be present.

essary for proper functioning.

(d) There must be means to allow the (d) The temperature of each con- crew to check, in flight, the func- tainer must be maintained, under in- tioning of each fire or overheat detec- tended operating conditions, to prevent tor electric circuit.

the pressure in the container from— (e) Components of each fire or over- (1) Falling below that necessary to heat detector system in a fire zone provide an adequate rate of discharge; must be fire-resistant.

or (f) No fire or overheat detector sys- (2) Rising high enough to cause pre- tem component for any fire zone may mature discharge.

pass through another fire zone, un- (e) If a pyrotechnic capsule is used to less— discharge the extinguishing agent, (1) It is protected against the possi- each container must be installed so that temperature conditions will not bility of false warnings resulting from cause hazardous deterioration of the fires in zones through which it passes; pyrotechnic capsule. or (2) Each zone involved is simulta- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as neously protected by the same detector amended by Amdt. 25–23, 35 FR 5678, Apr. 8, and extinguishing system.

1970; Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] (g) Each fire detector system must be § 25.1201 Fire extinguishing system constructed so that when it is in the materials.

configuration for installation it will (a) No material in any fire extin- not exceed the alarm activation time guishing system may react chemically approved for the detectors using the re- with any extinguishing agent so as to sponse time criteria specified in the ap- create a hazard.

propriate Technical Standard Order for (b) Each system component in an en- the detector.

gine compartment must be fireproof.

(h) EWIS for each fire or overheat de- tector system in a fire zone must meet § 25.1203 Fire detector system.

the requirements of § 25.1731.

(a) There must be approved, quick [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as acting fire or overheat detectors in amended by Amdt. 25–23, 35 FR 5678, Apr. 8, each designated fire zone, and in the 1970; Amdt. 25–26, 36 FR 5493, Mar. 24, 1971; combustion, turbine, and tailpipe sec- Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] tions of turbine engine installations, in numbers and locations ensuring § 25.1207 Compliance.

prompt detection of fire in those zones.

Unless otherwise specified, compli- (b) Each fire detector system must be ance with the requirements of §§ 25.1181 constructed and installed so that— (1) It will withstand the vibration, in- through 25.1203 must be shown by a full ertia, and other loads to which it may scale fire test or by one or more of the be subjected in operation; following methods: (2) There is a means to warn the crew (a) Tests of similar powerplant con- in the event that the sensor or associ- figurations; ated wiring within a designated fire (b) Tests of components; zone is severed at one point, unless the (c) Service experience of aircraft system continues to function as a sat- with similar powerplant configura- isfactory detection system after the tions; severing; and (d) Analysis.

(3) There is a means to warn the crew in the event of a short circuit in the [Amdt. 25–46, 43 FR 50598, Oct. 30, 1978]

14 CFR Ch. I (1–1–25 Edition) § 25.1301 (c) Operationally-relevant behavior Subpart F—Equipment of the installed equipment must be: G ENERAL (1) Predictable and unambiguous; and (2) Designed to enable the flightcrew § 25.1301 Function and installation.

to intervene in a manner appropriate Each item of installed equipment to the task.

must— (d) To the extent practicable, in- (a) Be of a kind and design appro- stalled equipment must incorporate priate to its intended function; means to enable the flightcrew to man- (b) Be labeled as to its identification, age errors resulting from the kinds of function, or operating limitations, or flightcrew interactions with the equip- any applicable combination of these ment that can be reasonably expected factors; and in service. This paragraph does not (c) Be installed according to limita- apply to any of the following: tions specified for that equipment.

(1) Skill-related errors associated with manual control of the airplane; [Doc. No. FAA–2022–1544, 89 FR 68735, Aug. 27, (2) Errors that result from decisions, 2024] actions, or omissions committed with § 25.1302 Installed systems and equip- malicious intent; ment for use by the flightcrew.

(3) Errors arising from a crew- member’s reckless decisions, actions, This section applies to installed sys- or omissions reflecting a substantial tems and equipment intended for disregard for safety; and flightcrew members’ use in operating (4) Errors resulting from acts or the airplane from their normally seat- threats of violence, including actions ed positions on the flight deck. The ap- taken under duress.

plicant must show that these systems and installed equipment, individually [Doc. No. FAA–2010–1175, 78 FR 25846, May 3, and in combination with other such 2013] systems and equipment, are designed so that qualified flightcrew members § 25.1303 Flight and navigation instru- trained in their use can safely perform ments.

all of the tasks associated with the sys- (a) The following flight and naviga- tems’ and equipment’s intended func- tion instruments must be installed so tions. Such installed equipment and that the instrument is visible from systems must meet the following re- each pilot station: quirements: (1) A free air temperature indicator (a) Flight deck controls must be in- or an air-temperature indicator which stalled to allow accomplishment of all provides indications that are convert- the tasks required to safely perform ible to free-air temperature.

the equipment’s intended function, and (2) A clock displaying hours, min- information must be provided to the utes, and seconds with a sweep-second flightcrew that is necessary to accom- pointer or digital presentation.

plish the defined tasks.

(3) A direction indicator (non- (b) Flight deck controls and informa- stabilized magnetic compass).

tion intended for the flightcrew’s use (b) The following flight and naviga- must: tion instruments must be installed at (1) Be provided in a clear and unam- each pilot station: biguous manner at a resolution and (1) An airspeed indicator. If airspeed precision appropriate to the task; limitations vary with altitude, the in- (2) Be accessible and usable by the dicator must have a maximum allow- flightcrew in a manner consistent with able airspeed indicator showing the the urgency, frequency, and duration of variation of V with altitude.

MO their tasks; and (2) An altimeter (sensitive).

(3) Enable flightcrew awareness, if (3) A rate-of-climb indicator (vertical awareness is required for safe oper- speed).

ation, of the effects on the airplane or systems resulting from flightcrew ac- (4) A gyroscopic rate-of-turn indi- tions. cator combined with an integral slip- Federal Aviation Administration, DOT § 25.1305 skid indicator (turn-and-bank indi- for isolating the individual warning cator) except that only a slip-skid indi- means from the master warning means.

cator is required on large airplanes (6) An oil temperature indicator for with a third attitude instrument sys- each engine.

tem useable through flight attitudes of (7) Fire-warning devices that provide 360 ° of pitch and roll and installed in visual and audible warning.

accordance with § 121.305(k) of this (8) An augmentation liquid quantity title.

indicator (appropriate for the manner (5) A bank and pitch indicator (gyro- in which the liquid is to be used in op- scopically stabilized).

eration) for each tank.

(6) A direction indicator (gyroscop- (b) For reciprocating engine-powered ically stabilized, magnetic or non- airplanes. In addition to the powerplant magnetic).

instruments required by paragraph (a) (c) The following flight and naviga- of this section, the following power- tion instruments are required as pre- plant instruments are required: scribed in this paragraph: (1) A carburetor air temperature indi- (1) A speed warning device is required cator for each engine.

for turbine engine powered airplanes (2) A cylinder head temperature indi- and for airplanes with V /M greater MO MO cator for each air-cooled engine.

than 0.8 V /M or 0.8 V /M . The DF DF D D (3) A manifold pressure indicator for speed warning device must give effec- each engine.

tive aural warning (differing distinc- (4) A fuel pressure indicator (to indi- tively from aural warnings used for cate the pressure at which the fuel is other purposes) to the pilots, whenever supplied) for each engine.

the speed exceeds V plus 6 knots or MO (5) A fuel flowmeter, or fuel mixture M + 0.01. The upper limit of the pro- MO indicator, for each engine without an duction tolerance for the warning de- automatic altitude mixture control.

vice may not exceed the prescribed (6) A tachometer for each engine.

warning speed.

(7) A device that indicates, to the (2) A machmeter is required at each flight crew (during flight), any change pilot station for airplanes with com- in the power output, for each engine pressibility limitations not otherwise with— indicated to the pilot by the airspeed (i) An automatic propeller feathering indicating system required under para- system, whose operation is initiated by graph (b)(1) of this section.

a power output measuring system; or [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as (ii) A total engine piston displace- amended by Amdt. 25–24, 35 FR 7108, May 6, ment of 2,000 cubic inches or more.

1970; Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; (8) A means to indicate to the pilot Amdt. 25–90, 62 FR 13253, Mar. 19, 1997] when the propeller is in reverse pitch, § 25.1305 Powerplant instruments. for each reversing propeller.

(c) For turbine engine-powered air- The following are required power- planes. In addition to the powerplant plant instruments: instruments required by paragraph (a) (a) For all airplanes. (1) A fuel pres- of this section, the following power- sure warning means for each engine, or plant instruments are required: a master warning means for all engines (1) A gas temperature indicator for with provision for isolating the indi- each engine.

vidual warning means from the master (2) A fuel flowmeter indicator for warning means.

each engine.

(2) A fuel quantity indicator for each fuel tank. (3) A tachometer (to indicate the (3) An oil quantity indicator for each speed of the rotors with established oil tank. limiting speeds) for each engine.

(4) An oil pressure indicator for each (4) A means to indicate, to the flight independent pressure oil system of crew, the operation of each engine each engine. starter that can be operated continu- (5) An oil pressure warning means for ously but that is neither designed for each engine, or a master warning continuous operation nor designed to means for all engines with provision prevent hazard if it failed.

14 CFR Ch. I (1–1–25 Edition) § 25.1307 (5) An indicator to indicate the func- proper functioning of that system to tioning of the powerplant ice protec- the flight crew.

tion system for each engine.

[Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as (6) An indicator for the fuel strainer amended by Amdt. 25–35, 39 FR 1831, Jan. 15, or filter required by § 25.997 to indicate 1974; Amdt. 25–36, 39 FR 35461, Oct. 1, 1974; the occurrence of contamination of the Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt.

strainer or filter before it reaches the 25–54, 45 FR 60173, Sept. 11, 1980; Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–115, 69 FR capacity established in accordance 40527, July 2, 2004] with § 25.997(d).

(7) A warning means for the oil § 25.1307 Miscellaneous equipment.

strainer or filter required by § 25.1019, if it has no bypass, to warn the pilot of The following is required miscella- the occurrence of contamination of the neous equipment: strainer or filter screen before it (a) [Reserved] reaches the capacity established in ac- (b) Two or more independent sources cordance with § 25.1019(a)(2).

of electrical energy.

(8) An indicator to indicate the prop- (c) Electrical protective devices, as er functioning of any heater used to prescribed in this part.

prevent ice clogging of fuel system (d) Two systems for two-way radio components.

communications, with controls for (d) For turbojet engine powered air- each accessible from each pilot station, planes. In addition to the powerplant designed and installed so that failure of instruments required by paragraphs (a) one system will not preclude operation and (c) of this section, the following of the other system. The use of a com- powerplant instruments are required: mon antenna system is acceptable if (1) An indicator to indicate thrust, or adequate reliability is shown.

a parameter that is directly related to (e) Two systems for radio navigation, thrust, to the pilot. The indication with controls for each accessible from must be based on the direct measure- each pilot station, designed and in- ment of thrust or of parameters that stalled so that failure of one system are directly related to thrust. The indi- will not preclude operation of the other cator must indicate a change in thrust system. The use of a common antenna resulting from any engine malfunction, system is acceptable if adequate reli- damage, or deterioration.

ability is shown.

(2) A position indicating means to in- [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as dicate to the flightcrew when the amended by Amdt. 25–46, 43 FR 50598, Oct. 30, thrust reversing device— 1978; Amdt. 25–54, 45 FR 60173, Sept. 11, 1980; (i) Is not in the selected position, and Amdt. 25–72, 55 FR 29785, July 20, 1990] (ii) Is in the reverse thrust position, for each engine using a thrust revers- § 25.1309 Equipment, systems, and in- stallations.

ing device.

(3) An indicator to indicate rotor sys- The requirements of this section, ex- tem unbalance.

cept as identified below, apply to any (e) For turbopropeller-powered air- equipment or system as installed on planes. In addition to the powerplant the airplane. Although this section instruments required by paragraphs (a) does not apply to the performance and and (c) of this section, the following flight characteristic requirements of powerplant instruments are required: subpart B of this part, or to the struc- (1) A torque indicator for each en- tural requirements of subparts C and D gine.

of this part, it does apply to any sys- (2) Position indicating means to indi- tem on which compliance with any of cate to the flight crew when the pro- those requirements is dependent. Sec- peller blade angle is below the flight tion 25.1309(b) does not apply to the low pitch position, for each propeller.

flight control jam conditions addressed (f) For airplanes equipped with fluid by § 25.671(c)(3); single failures in the systems (other than fuel) for thrust or brake system addressed by § 25.735(b)(1); power augmentation, an approved the failure conditions addressed by means must be provided to indicate the §§ 25.810(a)(1)(v) and 25.812; uncontained Federal Aviation Administration, DOT § 25.1310 engine rotor failure, engine case rup- condition due to all subsequent active ture, or engine case burn-through fail- failures is remote; and ures addressed by §§ 25.903(d)(1) and (iii) The sum of the probabilities of 25.1193 and part 33 of this chapter; and the latent failures that are combined propeller debris release failures ad- with each active failure does not ex- dressed by § 25.905(d) and part 35 of this ceed 1/1000.

(c) The airplane and systems must chapter.

provide information concerning unsafe (a) The airplane’s equipment and sys- system operating conditions to the tems must be designed and installed so flightcrew to enable them to take ap- that: propriate corrective action in a timely (1) The equipment and systems re- manner. Systems and controls, includ- quired for type certification or by oper- ing information, indications, and ating rules, or whose improper func- annunciations, must be designed to tioning would reduce safety, perform as minimize flightcrew errors that could intended under the airplane operating create additional hazards.

and environmental conditions; and (d) [Reserved] (2) Other equipment and systems, (e) The applicant must establish cer- functioning normally or abnormally, tification maintenance requirements do not adversely affect the safety of as necessary to prevent the develop- the airplane or its occupants or the ment of the failure conditions de- proper functioning of the equipment scribed in paragraph (b) of this section.

and systems addressed by paragraph These requirements must be included (a)(1) of this section.

in the Airworthiness Limitations sec- (b) The airplane systems and associ- tion of the Instructions for Continued ated components, evaluated separately Airworthiness required by § 25.1529.

and in relation to other systems, must be designed and installed so that they [Doc. No. FAA–2022–1544, 89 FR 68735, Aug. 27, meet all of the following requirements: 2024] (1) Each catastrophic failure condi- tion— § 25.1310 Power source capacity and distribution.

(i) Must be extremely improbable; and (a) Each installation whose func- (ii) Must not result from a single fail- tioning is required for type certifi- ure.

cation or under operating rules and (2) Each hazardous failure condition that requires a power supply is an ‘‘es- must be extremely remote.

sential load’’ on the power supply. The (3) Each major failure condition must power sources and the system must be be remote.

able to supply the following power (4) Each significant latent failure loads in probable operating combina- must be eliminated as far as practical, tions and for probable durations: or, if not practical to eliminate, the la- (1) Loads connected to the system tency of the significant latent failure with the system functioning normally.

must be minimized. However, the re- (2) Essential loads, after failure of quirements of the previous sentence do any one prime mover, power converter, not apply if the associated system or energy storage device.

meets the requirements of paragraphs (3) Essential loads after failure of— (b)(1) and (b)(2) of this section, assum- (i) Any one engine on two-engine air- ing the significant latent failure has planes; and occurred. (ii) Any two engines on airplanes (5) For each catastrophic failure con- with three or more engines.

dition that results from two failures, (4) Essential loads for which an alter- either of which could be latent for nate source of power is required, after more than one flight, the applicant any failure or malfunction in any one must show that— power supply system, distribution sys- (i) It is impractical to provide addi- tem, or other utilization system.

tional fault tolerance; and (b) In determining compliance with (ii) Given the occurrence of any sin- paragraphs (a)(2) and (3) of this section, gle latent failure, the residual average the power loads may be assumed to be probability of the catastrophic failure reduced under a monitoring procedure 14 CFR Ch. I (1–1–25 Edition) § 25.1316 consistent with safety in the kinds of (3) The system is not adversely af- operation authorized. Loads not re- fected during and after the time the quired in controlled flight need not be airplane is exposed to HIRF environ- considered for the two-engine-inoper- ment II, as described in appendix L to ative condition on airplanes with three this part.

or more engines. (b) Each electrical and electronic system that performs a function whose [Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] failure would significantly reduce the capability of the airplane or the ability § 25.1316 Electrical and electronic sys- of the flightcrew to respond to an ad- tem lightning protection.

verse operating condition must be de- (a) Each electrical and electronic signed and installed so the system is system that performs a function, for not adversely affected when the equip- which failure would prevent the contin- ment providing these functions is ex- ued safe flight and landing of the air- posed to equipment HIRF test level 1 plane, must be designed and installed or 2, as described in appendix L to this so that— part.

(1) The function is not adversely af- (c) Each electrical and electronic sys- fected during and after the time the tem that performs a function whose airplane is exposed to lightning; and failure would reduce the capability of (2) The system automatically recov- the airplane or the ability of the ers normal operation of that function flightcrew to respond to an adverse op- in a timely manner after the airplane erating condition must be designed and is exposed to lightning.

installed so the system is not adversely (b) Each electrical and electronic affected when the equipment providing system that performs a function, for the function is exposed to equipment which failure would reduce the capa- HIRF test level 3, as described in ap- bility of the airplane or the ability of pendix L to this part.

the flightcrew to respond to an adverse (d) Before December 1, 2012, an elec- operating condition, must be designed trical or electronic system that per- and installed so that the function re- forms a function whose failure would covers normal operation in a timely prevent the continued safe flight and manner after the airplane is exposed to landing of an airplane may be designed lightning.

and installed without meeting the pro- visions of paragraph (a) provided— [Doc. No. FAA–2010–0224, Amdt. 25–134, 76 FR 33135, June 8, 2011] (1) The system has previously been shown to comply with special condi- § 25.1317 High-intensity Radiated tions for HIRF, prescribed under § 21.16, Fields (HIRF) Protection.

issued before December 1, 2007; (2) The HIRF immunity characteris- (a) Except as provided in paragraph tics of the system have not changed (d) of this section, each electrical and since compliance with the special con- electronic system that performs a func- ditions was demonstrated; and tion whose failure would prevent the (3) The data used to demonstrate continued safe flight and landing of the compliance with the special conditions airplane must be designed and installed is provided.

so that— (1) The function is not adversely af- [Doc. No. FAA–2006–23657, 72 FR 44025, Aug. 6, fected during and after the time the 2007] airplane is exposed to HIRF environ- I NSTRUMENTS : I NSTALLATION ment I, as described in appendix L to this part; § 25.1321 Arrangement and visibility.

(2) The system automatically recov- ers normal operation of that function, (a) Each flight, navigation, and pow- in a timely manner, after the airplane erplant instrument for use by any pilot is exposed to HIRF environment I, as must be plainly visible to him from his described in appendix L to this part, station with the minimum practicable unless the system’s recovery conflicts deviation from his normal position and with other operational or functional line of vision when he is looking for- requirements of the system; and ward along the flight path.

Federal Aviation Administration, DOT § 25.1322 (b) The flight instruments required (3) Be removed when the alerting by § 25.1303 must be grouped on the in- condition no longer exists.

strument panel and centered as nearly (b) Alerts must conform to the fol- as practicable about the vertical plane lowing prioritization hierarchy based of the pilot’s forward vision. In addi- on the urgency of flightcrew awareness tion— and response.

(1) The instrument that most effec- (1) Warning: For conditions that re- tively indicates attitude must be on quire immediate flightcrew awareness the panel in the top center position; and immediate flightcrew response.

(2) Caution: For conditions that re- (2) The instrument that most effec- quire immediate flightcrew awareness tively indicates airspeed must be adja- and subsequent flightcrew response.

cent to and directly to the left of the (3) Advisory: For conditions that re- instrument in the top center position: quire flightcrew awareness and may re- (3) The instrument that most effec- quire subsequent flightcrew response.

tively indicates altitude must be adja- (c) Warning and caution alerts must: cent to and directly to the right of the (1) Be prioritized within each cat- instrument in the top center position; egory, when necessary.

and (2) Provide timely attention-getting (4) The instrument that most effec- cues through at least two different tively indicates direction of flight senses by a combination of aural, vis- must be adjacent to and directly below ual, or tactile indications.

the instrument in the top center posi- (3) Permit each occurrence of the at- tion.

tention-getting cues required by para- (c) Required powerplant instruments graph (c)(2) of this section to be ac- must be closely grouped on the instru- knowledged and suppressed, unless ment panel. In addition— they are required to be continuous.

(1) The location of identical power- (d) The alert function must be de- plant instruments for the engines must signed to minimize the effects of false prevent confusion as to which engine and nuisance alerts. In particular, it each instrument relates; and must be designed to: (2) Powerplant instruments vital to (1) Prevent the presentation of an the safe operation of the airplane must alert that is inappropriate or unneces- be plainly visible to the appropriate sary.

crewmembers.

(2) Provide a means to suppress an (d) Instrument panel vibration may attention-getting component of an not damage or impair the accuracy of alert caused by a failure of the alerting any instrument.

function that interferes with the (e) If a visual indicator is provided to flightcrew’s ability to safely operate indicate malfunction of an instrument, the airplane. This means must not be it must be effective under all probable readily available to the flightcrew so cockpit lighting conditions.

that it could be operated inadvertently [Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as or by habitual reflexive action. When amended by Amdt. 25–41, 42 FR 36970, July 18, an alert is suppressed, there must be a 1977] clear and unmistakable annunciation to the flightcrew that the alert has § 25.1322 Flightcrew alerting.

been suppressed.

(a) Flightcrew alerts must: (e) Visual alert indications must: (1) Provide the flightcrew with the (1) Conform to the following color information needed to: convention: (i) Identify non-normal operation or (i) Red for warning alert indications.

airplane system conditions, and (ii) Amber or yellow for caution alert (ii) Determine the appropriate ac- indications.

tions, if any. (iii) Any color except red or green for (2) Be readily and easily detectable advisory alert indications.

and intelligible by the flightcrew under (2) Use visual coding techniques, to- all foreseeable operating conditions, gether with other alerting function ele- including conditions where multiple ments on the flight deck, to distin- alerts are provided. guish between warning, caution, and 14 CFR Ch. I (1–1–25 Edition) § 25.1323 advisory alert indications, if they are (d) From 1.23 V to the speed at SR presented on monochromatic displays which stall warning begins, the IAS that are not capable of conforming to must change perceptibly with CAS and the color convention in paragraph (e)(1) in the same sense, and at speeds below of this section. stall warning speed the IAS must not (f) Use of the colors red, amber, and change in an incorrect sense.

yellow on the flight deck for functions (e) From V to V + ⁄3 (V ¥ MO MO DF other than flightcrew alerting must be V ), the IAS must change perceptibly MO limited and must not adversely affect with CAS and in the same sense, and at flightcrew alerting. higher speeds up to V the IAS must DF not change in an incorrect sense.

[Amdt. 25–131, 75 FR 67209, Nov. 2, 2010] (f) There must be no indication of airspeed that would cause undue dif- § 25.1323 Airspeed indicating system.

ficulty to the pilot during the takeoff For each airspeed indicating system, between the initiation of rotation and the following apply: the achievement of a steady climbing (a) Each airspeed indicating instru- condition.

ment must be approved and must be (g) The effects of airspeed indicating calibrated to indicate true airspeed (at system lag may not introduce signifi- sea level with a standard atmosphere) cant takeoff indicated airspeed bias, or with a minimum practicable instru- significant errors in takeoff or accel- ment calibration error when the cor- erate-stop distances.

responding pitot and static pressures (h) Each system must be arranged, so are applied.

far as practicable, to prevent malfunc- (b) Each system must be calibrated tion or serious error due to the entry of to determine the system error (that is, moisture, dirt, or other substances.

the relation between IAS and CAS) in (i) Each system must have a heated flight and during the accelerated take- pitot tube or an equivalent means of off ground run. The ground run calibra- preventing malfunction in the heavy tion must be determined— rain conditions defined in Table 1 of (1) From 0.8 of the minimum value of this section; mixed phase and ice crys- V to the maximum value of V , consid- 1 2 tal conditions as defined in part 33, Ap- ering the approved ranges of altitude pendix D, of this chapter; the icing con- and weight; and ditions defined in Appendix C of this (2) With the flaps and power settings part; and the following icing conditions corresponding to the values determined specified in Appendix O of this part: in the establishment of the takeoff (1) For airplanes certificated in ac- path under § 25.111 assuming that the cordance with § 25.1420(a)(1), the icing critical engine fails at the minimum conditions that the airplane is certified value of V .

1 to safely exit following detection.

(c) The airspeed error of the installa- (2) For airplanes certificated in ac- tion, excluding the airspeed indicator cordance with § 25.1420(a)(2), the icing instrument calibration error, may not conditions that the airplane is certified exceed three percent or five knots, to safely operate in and the icing con- whichever is greater, throughout the ditions that the airplane is certified to speed range, from— safely exit following detection.

(1) V to 1.23 V , with flaps re- (3) For airplanes certificated in ac- MO SR 1 tracted; and cordance with § 25.1420(a)(3) and for air- (2) 1.23 V to V with flaps in the planes not subject to § 25.1420, all icing SR 0 FE landing position. conditions.

T ABLE 1—H EAVY R AIN C ONDITIONS FOR A IRSPEED I NDICATING S YSTEM T ESTS Altitude range Liquid water Horizontal extent Droplet MVD content (ft) (m) (km) (nmiles) ( μ m) (g/m3) 0 to 10 000 ......................... 0 to 3000 ............................ 1 100 50 1000 6 5 3 2000 15 1 0.5 2000 Federal Aviation Administration, DOT § 25.1325 (j) Where duplicate airspeed indica- (i) For airplanes certificated in ac- tors are required, their respective pitot cordance with § 25.1420(a)(1), the icing tubes must be far enough apart to conditions that the airplane is certified avoid damage to both tubes in a colli- to safely exit following detection.

sion with a bird. (ii) For airplanes certificated in ac- cordance with § 25.1420(a)(2), the icing [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as conditions that the airplane is certified amended by Amdt. 25–57, 49 FR 6849, Feb. 23, to safely operate in and the icing con- 1984; Amdt. 25–108, 67 FR 70828, Nov. 26, 2002; ditions that the airplane is certified to Amdt. 25–109, 67 FR 76656, Dec. 12, 2002; Amdt.

25–140, 79 FR 65526, Nov. 4, 2014] safely exit following detection.

(iii) For airplanes certificated in ac- § 25.1324 Angle of attack system.

cordance with § 25.1420(a)(3) and for air- planes not subject to § 25.1420, all icing Each angle of attack system sensor conditions.

must be heated or have an equivalent (c) The design and installation of the means of preventing malfunction in the static pressure system must be such heavy rain conditions defined in Table that— 1 of § 25.1323, the mixed phase and ice (1) Positive drainage of moisture is crystal conditions as defined in part 33, provided; chafing of the tubing and ex- Appendix D, of this chapter, the icing cessive distortion or restriction at conditions defined in Appendix C of bends in the tubing is avoided; and the this part, and the following icing con- materials used are durable, suitable for ditions specified in Appendix O of this the purpose intended, and protected part: against corrosion; and (a) For airplanes certificated in ac- (2) It is airtight except for the port cordance with § 25.1420(a)(1), the icing into the atmosphere. A proof test must conditions that the airplane is certified be conducted to demonstrate the integ- to safely exit following detection.

rity of the static pressure system in (b) For airplanes certificated in ac- the following manner: cordance with § 25.1420(a)(2), the icing (i) Unpressurized airplanes. Evacuate conditions that the airplane is certified the static pressure system to a pres- to safely operate in and the icing con- sure differential of approximately 1 ditions that the airplane is certified to inch of mercury or to a reading on the safely exit following detection.

altimeter, 1,000 feet above the airplane (c) For airplanes certificated in ac- elevation at the time of the test. With- cordance with § 25.1420(a)(3) and for air- out additional pumping for a period of planes not subject to § 25.1420, all icing 1 minute, the loss of indicated altitude conditions.

must not exceed 100 feet on the altim- [Amdt. 25–140, 79 FR 65527, Nov. 4, 2014] eter.

(ii) Pressurized airplanes. Evacuate § 25.1325 Static pressure systems.

the static pressure system until a pres- (a) Each instrument with static air sure differential equivalent to the max- case connections must be vented to the imum cabin pressure differential for outside atmosphere through an appro- which the airplane is type certificated priate piping system.

is achieved. Without additional pump- (b) Each static port must be designed ing for a period of 1 minute, the loss of and located so that: indicated altitude must not exceed 2 (1) The static pressure system per- percent of the equivalent altitude of formance is least affected by airflow the maximum cabin differential pres- variation, or by moisture or other for- sure or 100 feet, whichever is greater.

eign matter; and (d) Each pressure altimeter must be (2) The correlation between air pres- approved and must be calibrated to in- sure in the static pressure system and dicate pressure altitude in a standard true ambient atmospheric static pres- atmosphere, with a minimum prac- sure is not changed when the airplane ticable calibration error when the cor- is exposed to the icing conditions de- responding static pressures are applied.

fined in Appendix C of this part, and (e) Each system must be designed and the following icing conditions specified installed so that the error in indicated in Appendix O of this part: pressure altitude, at sea level, with a 14 CFR Ch. I (1–1–25 Edition) § 25.1326 standard atmosphere, excluding instru- (b) The indication provided must be ment calibration error, does not result designed to alert the flight crew if ei- in an error of more than ± 30 feet per 100 ther of the following conditions exist: knots speed for the appropriate con- (1) The pitot heating system is figuration in the speed range between switched ‘‘off’’.

1.23 V with flaps extended and 1.7 SR 0 (2) The pitot heating system is V with flaps retracted. However, the SR 1 switched ‘‘on’’ and any pitot tube heat- error need not be less than ± 30 feet.

ing element is inoperative.

(f) If an altimeter system is fitted [Amdt. 25–43, 43 FR 10339, Mar. 13, 1978] with a device that provides corrections to the altimeter indication, the device § 25.1327 Magnetic direction indicator.

must be designed and installed in such (a) Each magnetic direction indicator manner that it can be bypassed when it must be installed so that its accuracy malfunctions, unless an alternate al- is not excessively affected by the air- timeter system is provided. Each cor- plane’s vibration or magnetic fields.

rection device must be fitted with a (b) The compensated installation means for indicating the occurrence of may not have a deviation, in level reasonably probable malfunctions, in- flight, greater than 10 degrees on any cluding power failure, to the flight heading.

crew. The indicating means must be ef- fective for any cockpit lighting condi- § 25.1329 Flight guidance system.

tion likely to occur.

(a) Quick disengagement controls for (g) Except as provided in paragraph the autopilot and autothrust functions (h) of this section, if the static pressure must be provided for each pilot. The system incorporates both a primary autopilot quick disengagement con- and an alternate static pressure source, trols must be located on both control the means for selecting one or the wheels (or equivalent). The autothrust other source must be designed so quick disengagement controls must be that— located on the thrust control levers.

(1) When either source is selected, the Quick disengagement controls must be other is blocked off; and readily accessible to each pilot while (2) Both sources cannot be blocked operating the control wheel (or equiva- off simultaneously.

lent) and thrust control levers.

(h) For unpressurized airplanes, para- (b) The effects of a failure of the sys- graph (g)(1) of this section does not tem to disengage the autopilot or apply if it can be demonstrated that autothrust functions when manually the static pressure system calibration, commanded by the pilot must be as- when either static pressure source is sessed in accordance with the require- selected, is not changed by the other ments of § 25.1309.

static pressure source being open or blocked. (c) Engagement or switching of the flight guidance system, a mode, or a [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as sensor may not cause a transient re- amended by Amdt. 25–5, 30 FR 8261, June 29, sponse of the airplane’s flight path any 1965; Amdt. 25–12, 32 FR 7587, May 24, 1967; greater than a minor transient, as de- Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt.

25–108, 67 FR 70828, Nov. 26, 2002; Amdt. 25– fined in paragraph (n)(1) of this section.

140, 79 FR 65527, Nov. 4, 2014] (d) Under normal conditions, the dis- engagement of any automatic control § 25.1326 Pitot heat indication systems.

function of a flight guidance system If a flight instrument pitot heating may not cause a transient response of system is installed, an indication sys- the airplane’s flight path any greater tem must be provided to indicate to than a minor transient.

the flight crew when that pitot heating (e) Under rare normal and non-nor- system is not operating. The indication mal conditions, disengagement of any system must comply with the following automatic control function of a flight requirements: guidance system may not result in a (a) The indication provided must in- transient any greater than a signifi- corporate an amber light that is in cant transient, as defined in paragraph clear view of a flight crewmember. (n)(2) of this section.

Federal Aviation Administration, DOT § 25.1331 (f) The function and direction of mo- may not create a potential hazard tion of each command reference con- when the flightcrew applies an override trol, such as heading select or vertical force to the thrust levers.

speed, must be plainly indicated on, or (n) For purposes of this section, a transient is a disturbance in the con- adjacent to, each control if necessary trol or flight path of the airplane that to prevent inappropriate use or confu- is not consistent with response to sion.

flightcrew inputs or environmental (g) Under any condition of flight ap- conditions.

propriate to its use, the flight guidance (1) A minor transient would not sig- system may not produce hazardous nificantly reduce safety margins and loads on the airplane, nor create haz- would involve flightcrew actions that ardous deviations in the flight path.

are well within their capabilities. A This applies to both fault-free oper- minor transient may involve a slight ation and in the event of a malfunc- increase in flightcrew workload or tion, and assumes that the pilot begins some physical discomfort to passengers corrective action within a reasonable or cabin crew.

period of time.

(2) A significant transient may lead (h) When the flight guidance system to a significant reduction in safety is in use, a means must be provided to margins, an increase in flightcrew avoid excursions beyond an acceptable workload, discomfort to the flightcrew, margin from the speed range of the or physical distress to the passengers normal flight envelope. If the airplane or cabin crew, possibly including non- experiences an excursion outside this fatal injuries. Significant transients do range, a means must be provided to not require, in order to remain within prevent the flight guidance system or recover to the normal flight enve- from providing guidance or control to lope, any of the following: an unsafe speed.

(i) Exceptional piloting skill, alert- (i) The flight guidance system func- ness, or strength.

tions, controls, indications, and alerts (ii) Forces applied by the pilot which must be designed to minimize are greater than those specified in flightcrew errors and confusion con- § 25.143(c).

cerning the behavior and operation of (iii) Accelerations or attitudes in the the flight guidance system. Means airplane that might result in further must be provided to indicate the cur- hazard to secured or non-secured occu- rent mode of operation, including any pants.

armed modes, transitions, and rever- sions. Selector switch position is not [Doc. No. FAA–2004–18775, 71 FR 18191, Apr.

an acceptable means of indication. The 11, 2006] controls and indications must be § 25.1331 Instruments using a power grouped and presented in a logical and supply.

consistent manner. The indications must be visible to each pilot under all (a) For each instrument required by expected lighting conditions. § 25.1303(b) that uses a power supply, (j) Following disengagement of the the following apply: autopilot, a warning (visual and audi- (1) Each instrument must have a vis- tory) must be provided to each pilot ual means integral with, the instru- and be timely and distinct from all ment, to indicate when power adequate other cockpit warnings. to sustain proper instrument perform- (k) Following disengagement of the ance is not being supplied. The power autothrust function, a caution must be must be measured at or near the point provided to each pilot. where it enters the instruments. For (l) The autopilot may not create a electric instruments, the power is con- potential hazard when the flightcrew sidered to be adequate when the volt- applies an override force to the flight age is within approved limits.

controls. (2) Each instrument must, in the (m) During autothrust operation, it event of the failure of one power must be possible for the flightcrew to source, be supplied by another power move the thrust levers without requir- source. This may be accomplished ing excessive force. The autothrust automatically or by manual means.

14 CFR Ch. I (1–1–25 Edition) § 25.1333 (3) If an instrument presenting navi- § 25.1337 Powerplant instruments.

gation data receives information from (a) Instruments and instrument lines.

sources external to that instrument (1) Each powerplant and auxiliary and loss of that information would power unit instrument line must meet render the presented data unreliable, the requirements of §§ 25.993 and 25.1183.

the instrument must incorporate a vis- (2) Each line carrying flammable ual means to warn the crew, when such fluids under pressure must— loss of information occurs, that the (i) Have restricting orifices or other presented data should not be relied safety devices at the source of pressure upon.

to prevent the escape of excessive fluid (b) As used in this section, ‘‘instru- if the line fails; and ment’’ includes devices that are phys- (ii) Be installed and located so that ically contained in one unit, and de- the escape of fluids would not create a vices that are composed of two or more hazard.

physically separate units or compo- (3) Each powerplant and auxiliary nents connected together (such as a re- power unit instrument that utilizes mote indicating gyroscopic direction flammable fluids must be installed and indicator that includes a magnetic located so that the escape of fluid sensing element, a gyroscopic unit, an would not create a hazard.

amplifier and an indicator connected together). (b) Fuel quantity indicator. There must be means to indicate to the flight [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as crewmembers, the quantity, in gallons amended by Amdt. 25–41, 42 FR 36970, July 18, or equivalent units, of usable fuel in 1977] each tank during flight. In addition— § 25.1333 Instrument systems.

(1) Each fuel quantity indicator must be calibrated to read ‘‘zero’’ during For systems that operate the instru- level flight when the quantity of fuel ments required by § 25.1303(b) which are remaining in the tank is equal to the located at each pilot’s station— unusable fuel supply determined under (a) Means must be provided to con- § 25.959; nect the required instruments at the (2) Tanks with interconnected outlets first pilot’s station to operating sys- and airspaces may be treated as one tems which are independent of the op- tank and need not have separate indi- erating systems at other flight crew cators; and stations, or other equipment; (b) The equipment, systems, and in- (3) Each exposed sight gauge, used as stallations must be designed so that a fuel quantity indicator, must be pro- one display of the information essen- tected against damage.

tial to the safety of flight which is pro- (c) Fuel flowmeter system. If a fuel vided by the instruments, including at- flowmeter system is installed, each titude, direction, airspeed, and altitude metering component must have a will remain available to the pilots, means for bypassing the fuel supply if without additional crewmember ac- malfunction of that component se- tion, after any single failure or com- verely restricts fuel flow.

bination of failures that is not shown (d) Oil quantity indicator. There must to be extremely improbable; and be a stick gauge or equivalent means (c) Additional instruments, systems, to indicate the quantity of oil in each or equipment may not be connected to tank. If an oil transfer or reserve oil the operating systems for the required supply system is installed, there must instruments, unless provisions are be a means to indicate to the flight made to ensure the continued normal crew, in flight, the quantity of oil in functioning of the required instru- each tank.

ments in the event of any malfunction (e) Turbopropeller blade position indi- of the additional instruments, systems, cator. Required turbopropeller blade or equipment which is not shown to be position indicators must begin indi- extremely improbable.

cating before the blade moves more than eight degrees below the flight low [Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as pitch stop. The source of indication amended by Amdt. 25–41, 42 FR 36970, July 18, 1977] must directly sense the blade position.

Federal Aviation Administration, DOT § 25.1353 (f) Fuel pressure indicator. There must (6) There are means to indicate to ap- be means to measure fuel pressure, in propriate crewmembers the generating each system supplying reciprocating system quantities essential for the safe engines, at a point downstream of any operation of the system, such as the fuel pump except fuel injection pumps. voltage and current supplied by each In addition— generator.

(1) If necessary for the maintenance (c) External power. If provisions are of proper fuel delivery pressure, there made for connecting external power to must be a connection to transmit the the airplane, and that external power carburetor air intake static pressure to can be electrically connected to equip- the proper pump relief valve connec- ment other than that used for engine tion; and starting, means must be provided to (2) If a connection is required under ensure that no external power supply paragraph (f)(1) of this section, the having a reverse polarity, or a reverse gauge balance lines must be independ- phase sequence, can supply power to ently connected to the carburetor inlet the airplane’s electrical system.

pressure to avoid erroneous readings.

(d) Operation without normal electrical power. It must be shown by analysis, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tests, or both, that the airplane can be amended by Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] operated safely in VFR conditions, for a period of not less than five minutes, E LECTRICAL S YSTEMS AND E QUIPMENT with the normal electrical power (elec- trical power sources excluding the bat- § 25.1351 General.

tery) inoperative, with critical type (a) Electrical system capacity. The re- fuel (from the standpoint of flameout quired generating capacity, and num- and restart capability), and with the ber and kinds of power sources must— airplane initially at the maximum cer- (1) Be determined by an electrical tificated altitude. Parts of the elec- load analysis; and trical system may remain on if— (2) Meet the requirements of § 25.1309.

(1) A single malfunction, including a (b) Generating system. The generating wire bundle or junction box fire, can- system includes electrical power not result in loss of both the part sources, main power busses, trans- turned off and the part turned on; and mission cables, and associated control, (2) The parts turned on are elec- regulation, and protective devices. It trically and mechanically isolated must be designed so that— from the parts turned off.

(1) Power sources function properly [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as when independent and when connected amended by Amdt. 25–41, 42 FR 36970, July 18, in combination; 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990] (2) No failure or malfunction of any power source can create a hazard or § 25.1353 Electrical equipment and in- impair the ability of remaining sources stallations.

to supply essential loads; (a) Electrical equipment and controls (3) The system voltage and frequency must be installed so that operation of (as applicable) at the terminals of all any one unit or system of units will essential load equipment can be main- not adversely affect the simultaneous tained within the limits for which the operation of any other electrical unit equipment is designed, during any or system essential to safe operation.

probable operating condition; and Any electrical interference likely to be (4) System transients due to switch- present in the airplane must not result ing, fault clearing, or other causes do in hazardous effects on the airplane or not make essential loads inoperative, its systems.

and do not cause a smoke or fire haz- (b) Storage batteries must be de- ard.

signed and installed as follows: (5) There are means accessible, in flight, to appropriate crewmembers for (1) Safe cell temperatures and pres- the individual and collective dis- sures must be maintained during any connection of the electrical power probable charging or discharging con- sources from the system. dition. No uncontrolled increase in cell 14 CFR Ch. I (1–1–25 Edition) § 25.1355 temperature may result when the bat- § 25.1355 Distribution system.

tery is recharged (after previous com- (a) The distribution system includes plete discharge)— the distribution busses, their associ- (i) At maximum regulated voltage or ated feeders, and each control and pro- power; tective device.

(ii) During a flight of maximum dura- (b) [Reserved] tion; and (c) If two independent sources of elec- (iii) Under the most adverse cooling trical power for particular equipment condition likely to occur in service.

or systems are required by this chap- (2) Compliance with paragraph (b)(1) ter, in the event of the failure of one of this section must be shown by test power source for such equipment or unless experience with similar bat- system, another power source (includ- ing its separate feeder) must be auto- teries and installations has shown that matically provided or be manually se- maintaining safe cell temperatures and lectable to maintain equipment or sys- pressures presents no problem.

tem operation.

(3) No explosive or toxic gases emit- ted by any battery in normal oper- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ation, or as the result of any probable amended by Amdt. 25–23, 35 FR 5679, Apr. 8, 1970; Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] malfunction in the charging system or battery installation, may accumulate § 25.1357 Circuit protective devices.

in hazardous quantities within the air- (a) Automatic protective devices plane.

must be used to minimize distress to (4) No corrosive fluids or gases that the electrical system and hazard to the may escape from the battery may dam- airplane in the event of wiring faults or age surrounding airplane structures or serious malfunction of the system or adjacent essential equipment.

connected equipment.

(5) Each nickel cadmium battery in- (b) The protective and control de- stallation must have provisions to pre- vices in the generating system must be vent any hazardous effect on structure designed to de-energize and disconnect or essential systems that may be faulty power sources and power trans- caused by the maximum amount of mission equipment from their associ- heat the battery can generate during a ated busses with sufficient rapidity to short circuit of the battery or of indi- provide protection from hazardous vidual cells.

over-voltage and other malfunctioning.

(6) Nickel cadmium battery installa- (c) Each resettable circuit protective tions must have— device must be designed so that, when (i) A system to control the charging an overload or circuit fault exists, it rate of the battery automatically so as will open the circuit irrespective of the to prevent battery overheating; position of the operating control.

(ii) A battery temperature sensing (d) If the ability to reset a circuit and over-temperature warning system breaker or replace a fuse is essential to with a means for disconnecting the safety in flight, that circuit breaker or battery from its charging source in the fuse must be located and identified so event of an over-temperature condi- that it can be readily reset or replaced tion; or in flight. Where fuses are used, there must be spare fuses for use in flight (iii) A battery failure sensing and equal to at least 50% of the number of warning system with a means for dis- fuses of each rating required for com- connecting the battery from its charg- plete circuit protection.

ing source in the event of battery fail- (e) Each circuit for essential loads ure.

must have individual circuit protec- (c) Electrical bonding must provide tion. However, individual protection an adequate electrical return path for each circuit in an essential load under both normal and fault condi- system (such as each position light cir- tions, on airplanes having grounded cuit in a system) is not required.

electrical systems.

(f) For airplane systems for which [Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] the ability to remove or reset power Federal Aviation Administration, DOT § 25.1381 during normal operations is necessary, tion to generator loading, including the system must be designed so that loading due to faults.

circuit breakers are not the primary (b) For each flight condition that means to remove or reset system power cannot be simulated adequately in the unless specifically designed for use as a laboratory or by ground tests on the switch. airplane, flight tests must be made.

(g) Automatic reset circuit breakers § 25.1365 Electrical appliances, motors, may be used as integral protectors for and transformers.

electrical equipment (such as thermal cut-outs) if there is circuit protection (a) An applicant must show that, in to protect the cable to the equipment. the event of a failure of the electrical supply or control system, the design [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as and installation of domestic appliances amended by Amdt. 25–123, 72 FR 63405, Nov. 8, meet the requirements of § 25.1309(b) 2007] and (c). Domestic appliances are items such as cooktops, ovens, coffee makers, § 25.1360 Precautions against injury.

water heaters, refrigerators, and toilet (a) Shock. The electrical system flush systems that are placed on the must be designed to minimize risk of airplane to provide service amenities electric shock to crew, passengers, and to passengers.

servicing personnel and to mainte- (b) Galleys and cooking appliances nance personnel using normal pre- must be installed in a way that mini- cautions.

mizes risk of overheat or fire.

(b) Burns. The temperature of any (c) Domestic appliances, particularly part that may be handled by a crew- those in galley areas, must be installed member during normal operations or protected so as to prevent damage or must not cause dangerous inadvertent contamination of other equipment or movement by the crewmember or in- systems from fluids or vapors which jury to the crewmember.

may be present during normal oper- ation or as a result of spillage, if such [Amdt. 25–123, 72 FR 63406, Nov. 8, 2007] damage or contamination could create § 25.1362 Electrical supplies for emer- a hazardous condition.

gency conditions.

(d) Unless compliance with § 25.1309(b) is provided by the circuit protective A suitable electrical supply must be device required by § 25.1357(a), electric provided to those services required for motors and transformers, including emergency procedures after an emer- those installed in domestic systems, gency landing or ditching. The circuits must have a suitable thermal protec- for these services must be designed, tion device to prevent overheating protected, and installed so that the under normal operation and failure risk of the services being rendered inef- conditions, if overheating could create fective under these emergency condi- a smoke or fire hazard.

tions is minimized.

[Amdt. 25–123, 72 FR 63406, Nov. 8, 2007, as [Amdt. 25–123, 72 FR 63406, Nov. 8, 2007] amended by Doc. No. FAA–2022–1544, 89 FR 68735, Aug. 27, 2024] § 25.1363 Electrical system tests.

(a) When laboratory tests of the elec- L IGHTS trical system are conducted— § 25.1381 Instrument lights.

(1) The tests must be performed on a mock-up using the same generating (a) The instrument lights must— equipment used in the airplane; (1) Provide sufficient illumination to (2) The equipment must simulate the make each instrument, switch and electrical characteristics of the dis- other device necessary for safe oper- tribution wiring and connected loads to ation easily readable unless sufficient the extent necessary for valid test re- illumination is available from another sults; and source; and (3) Laboratory generator drives must (2) Be installed so that— simulate the actual prime movers on (i) Their direct rays are shielded from the airplane with respect to their reac- the pilot’s eyes; and 14 CFR Ch. I (1–1–25 Edition) § 25.1383 (ii) No objectionable reflections are § 25.1387 Position light system dihe- visible to the pilot. dral angles.

(b) Unless undimmed instrument (a) Except as provided in paragraph lights are satisfactory under each ex- (e) of this section, each forward and pected flight condition, there must be a rear position light must, as installed, means to control the intensity of illu- show unbroken light within the dihe- mination.

dral angles described in this section.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (b) Dihedral angle L (left) is formed amended by Amdt. 25–72, 55 FR 29785, July 20, by two intersecting vertical planes, the 1990] first parallel to the longitudinal axis of the airplane, and the other at 110 de- § 25.1383 Landing lights.

grees to the left of the first, as viewed (a) Each landing light must be ap- when looking forward along the longi- proved, and must be installed so that— tudinal axis.

(1) No objectionable glare is visible (c) Dihedral angle R (right) is formed to the pilot; by two intersecting vertical planes, the (2) The pilot is not adversely affected first parallel to the longitudinal axis of by halation; and the airplane, and the other at 110 de- (3) It provides enough light for night grees to the right of the first, as viewed landing.

when looking forward along the longi- (b) Except when one switch is used tudinal axis.

for the lights of a multiple light instal- (d) Dihedral angle A (aft) is formed lation at one location, there must be a by two intersecting vertical planes separate switch for each light.

making angles of 70 degrees to the (c) There must be a means to indicate right and to the left, respectively, to a to the pilots when the landing lights vertical plane passing through the lon- are extended.

gitudinal axis, as viewed when looking aft along the longitudinal axis.

§ 25.1385 Position light system installa- tion. (e) If the rear position light, when mounted as far aft as practicable in ac- (a) General. Each part of each posi- cordance with § 25.1385(c), cannot show tion light system must meet the appli- unbroken light within dihedral angle A cable requirements of this section and (as defined in paragraph (d) of this sec- each system as a whole must meet the tion), a solid angle or angles of ob- requirements of §§ 25.1387 through structed visibility totaling not more 25.1397.

than 0.04 steradians is allowable within (b) Forward position lights. Forward that dihedral angle, if such solid angle position lights must consist of a red is within a cone whose apex is at the and a green light spaced laterally as rear position light and whose elements far apart as practicable and installed make an angle of 30 ° with a vertical forward on the airplane so that, with line passing through the rear position the airplane in the normal flying posi- light.

tion, the red light is on the left side and the green light is on the right side.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as Each light must be approved.

amended by Amdt. 25–30, 36 FR 21278, Nov. 5, (c) Rear position light. The rear posi- 1971] tion light must be a white light mount- § 25.1389 Position light distribution ed as far aft as practicable on the tail and intensities.

or on each wing tip, and must be ap- proved.

(a) General. The intensities prescribed (d) Light covers and color filters. Each in this section must be provided by new light cover or color filter must be at equipment with light covers and color least flame resistant and may not filters in place. Intensities must be de- change color or shape or lose any ap- termined with the light source oper- preciable light transmission during ating at a steady value equal to the av- normal use.

erage luminous output of the source at the normal operating voltage of the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as airplane. The light distribution and in- amended by Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] tensity of each position light must Federal Aviation Administration, DOT § 25.1395 meet the requirements of paragraph (b) Angle from right or left of longitu- of this section. Dihedral angle (light in- Intensity dinal axis, meas- cluded) (candles) ured from dead (b) Forward and rear position lights.

ahead The light distribution and intensities must be expressed in terms of min- imum intensities in the horizontal plane, minimum intensities in any vertical plane, and maximum inten- § 25.1393 Minimum intensities in any sities in overlapping beams, within di- vertical plane of forward and rear hedral angles L, R, and A, and must position lights.

meet the following requirements: Each position light intensity must (1) Intensities in the horizontal plane.

equal or exceed the applicable values in Each intensity in the horizontal plane the following table: (the plane containing the longitudinal axis of the airplane and perpendicular Angle above or below the horizontal plane Intensity, l to the plane of symmetry of the air- 0 ° ......................................................................... 1.00 plane) must equal or exceed the values 0 ° to 5 ° ................................................................ 0.90 in § 25.1391.

5 ° to 10 ° .............................................................. 0.80 (2) Intensities in any vertical plane.

10 ° to 15 ° ............................................................ 0.70 Each intensity in any vertical plane 15 ° to 20 ° ............................................................ 0.50 20 ° to 30 ° ............................................................ 0.30 (the plane perpendicular to the hori- 30 ° to 40 ° ............................................................ 0.10 zontal plane) must equal or exceed the 40 ° to 90 ° ............................................................ 0.05 appropriate value in § 25.1393, where I is the minimum intensity prescribed in § 25.1395 Maximum intensities in over- § 25.1391 for the corresponding angles in lapping beams of forward and rear the horizontal plane.

position lights.

(3) Intensities in overlaps between adja- No position light intensity may ex- cent signals. No intensity in any over- ceed the applicable values in the fol- lap between adjacent signals may ex- lowing table, except as provided in ceed the values given in § 25.1395, except § 25.1389(b)(3).

that higher intensities in overlaps may be used with main beam intensities Maximum intensity substantially greater than the minima Overlaps Area A Area B specified in §§ 25.1391 and 25.1393 if the (candles) (candles) overlap intensities in relation to the Green in dihedral angle L ............. 10 1 main beam intensities do not adversely Red in dihedral angle R ................ 10 1 affect signal clarity. When the peak in- Green in dihedral angle A ............. 5 1 tensity of the forward position lights is Red in dihedral angle A ................ 5 1 more than 100 candles, the maximum Rear white in dihedral angle L ...... 5 1 overlap intensities between them may Rear white in dihedral angle R ..... 5 1 exceed the values given in § 25.1395 if the overlap intensity in Area A is not Where— more than 10 percent of peak position (a) Area A includes all directions in light intensity and the overlap inten- the adjacent dihedral angle that pass sity in Area B is not greater than 2.5 through the light source and intersect percent of peak position light inten- the common boundary plane at more sity.

than 10 degrees but less than 20 de- grees; and § 25.1391 Minimum intensities in the (b) Area B includes all directions in horizontal plane of forward and the adjacent dihedral angle that pass rear position lights.

through the light source and intersect Each position light intensity must the common boundary plane at more equal or exceed the applicable values in than 20 degrees.

the following table: 14 CFR Ch. I (1–1–25 Edition) § 25.1397 except that a solid angle or angles of § 25.1397 Color specifications.

obstructed visibility totaling not more Each position light color must have than 0.03 steradians is allowable within the applicable International Commis- a solid angle equal to 0.15 steradians sion on Illumination chromaticity co- centered about the longitudinal axis in ordinates as follows: the rearward direction.

(a) Aviation red — (c) Flashing characteristics. The ar- y is not greater than 0.335; and rangement of the system, that is, the z is not greater than 0.002.

number of light sources, beam width, speed of rotation, and other character- (b) Aviation green — istics, must give an effective flash fre- x is not greater than 0.440 ¥ 0.320 y ; quency of not less than 40, nor more x is not greater than y ¥ 0.170; and than 100 cycles per minute. The effec- y is not less than 0.390 ¥ 0.170 x .

tive flash frequency is the frequency at (c) Aviation white — which the airplane’s complete anti- collision light system is observed from x is not less than 0.300 and not greater than a distance, and applies to each sector 0.540; of light including any overlaps that y is not less than x ¥ 0.040; or y ¥ 0.010, which- exist when the system consists of more ever is the smaller; and y is not greater than x + 0.020 nor than one light source. In overlaps, 0.636 ¥ 0.400 x ; flash frequencies may exceed 100, but Where y is the y coordinate of the Planckian not 180 cycles per minute.

radiator for the value of x considered.

(d) Color. Each anticollision light [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as must be either aviation red or aviation amended by Amdt. 25–27, 36 FR 12972, July 10, white and must meet the applicable re- 1971] quirements of § 25.1397.

(e) Light intensity. The minimum § 25.1399 Riding light.

light intensities in all vertical planes, (a) Each riding (anchor) light re- measured with the red filter (if used) quired for a seaplane or amphibian and expressed in terms of ‘‘effective’’ must be installed so that it can— intensities, must meet the require- (1) Show a white light for at least 2 ments of paragraph (f) of this section.

nautical miles at night under clear at- The following relation must be as- mospheric conditions; and sumed: (2) Show the maximum unbroken t light practicable when the airplane is 2 I(t)dt

moored or drifting on the water. ∫

t I = (b) Externally hung lights may be e t t + − 0 2 .

( )

used.

2 1 where: § 25.1401 Anticollision light system.

I e = effective intensity (candles).

(a) General. The airplane must have I(t) = instantaneous intensity as a function an anticollision light system that— of time.

t —t = flash time interval (seconds).

(1) Consists of one or more approved 2 1 anticollision lights located so that Normally, the maximum value of effec- their light will not impair the crew’s tive intensity is obtained when t and t 2 1 vision or detract from the conspicuity are chosen so that the effective inten- of the position lights; and sity is equal to the instantaneous in- (2) Meets the requirements of para- tensity at t and t .

2 1 graphs (b) through (f) of this section.

(f) Minimum effective intensities for (b) Field of coverage. The system must anticollision lights. Each anticollision consist of enough lights to illuminate light effective intensity must equal or the vital areas around the airplane exceed the applicable values in the fol- considering the physical configuration lowing table.

and flight characteristics of the air- plane. The field of coverage must ex- Effective Angle above or below the horizontal plane intensity tend in each direction within at least (candles) 75 degrees above and 75 degrees below Federal Aviation Administration, DOT § 25.1415 Effective (4) The stowage provisions for each Angle above or below the horizontal plane intensity portable liferaft must allow rapid de- (candles) tachment and removal of the raft for use at other than the intended exits.

(e) Long-range signaling device. The stowage provisions for the long-range signaling device required by § 25.1415 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as must be near an exit available during amended by Amdt. 25–27, 36 FR 12972, July 10, an unplanned ditching.

1971; Amdt. 25–41, 42 FR 36970, July 18, 1977] (f) Life preserver stowage provisions.

The stowage provisions for life pre- § 25.1403 Wing icing detection lights.

servers described in § 25.1415 must ac- Unless operations at night in known commodate one life preserver for each or forecast icing conditions are prohib- occupant for which certification for ited by an operating limitation, a ditching is requested. Each life pre- means must be provided for illu- server must be within easy reach of minating or otherwise determining the each seated occupant.

formation of ice on the parts of the (g) Life line stowage provisions. If cer- wings that are critical from the stand- tification for ditching under § 25.801 is point of ice accumulation. Any illu- requested, there must be provisions to mination that is used must be of a type store life lines. These provisions that will not cause glare or reflection must— that would handicap crewmembers in (1) Allow one life line to be attached the performance of their duties.

to each side of the fuselage; and [Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] (2) Be arranged to allow the life lines to be used to enable the occupants to S AFETY E QUIPMENT stay on the wing after ditching.

§ 25.1411 General.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) Accessibility. Required safety amended by Amdt. 25–32, 37 FR 3972, Feb. 24, equipment to be used by the crew in an 1972; Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; emergency must be readily accessible. Amdt. 25–53, 45 FR 41593, June 19, 1980; Amdt.

(b) Stowage provisions. Stowage provi- 25–70, 54 FR 43925, Oct. 27, 1989; Amdt. 25–79, 58 FR 45229, Aug. 26, 1993; Amdt. 25–116, 69 FR sions for required emergency equip- 62789, Oct. 27, 2004] ment must be furnished and must— (1) Be arranged so that the equip- § 25.1415 Ditching equipment.

ment is directly accessible and its loca- tion is obvious; and (a) Ditching equipment used in air- (2) Protect the safety equipment planes to be certificated for ditching from inadvertent damage.

under § 25.801, and required by the oper- (c) Emergency exit descent device. The ating rules of this chapter, must meet stowage provisions for the emergency the requirements of this section.

exit descent devices required by (b) Each liferaft and each life pre- § 25.810(a) must be at each exit for server must be approved. In addition— which they are intended.

(1) Unless excess rafts of enough ca- (d) Liferafts. (1) The stowage provi- pacity are provided, the buoyancy and sions for the liferafts described in seating capacity beyond the rated ca- § 25.1415 must accommodate enough pacity of the rafts must accommodate rafts for the maximum number of occu- all occupants of the airplane in the pants for which certification for ditch- event of a loss of one raft of the largest ing is requested.

rated capacity; and (2) Liferafts must be stowed near (2) Each raft must have a trailing exits through which the rafts can be line, and must have a static line de- launched during an unplanned ditch- signed to hold the raft near the air- ing.

plane but to release it if the airplane (3) Rafts automatically or remotely becomes totally submerged.

released outside the airplane must be attached to the airplane by means of (c) Approved survival equipment the static line prescribed in § 25.1415. must be attached to each liferaft.

14 CFR Ch. I (1–1–25 Edition) § 25.1419 (d) There must be an approved sur- the powerplant installation, certain ad- vival type emergency locator trans- ditional provisions of subpart E of this mitter for use in one life raft. part may be found applicable.

(e) For airplanes not certificated for (e) One of the following methods of ditching under § 25.801 and not having icing detection and activation of the approved life preservers, there must be airframe ice protection system must be an approved flotation means for each provided: occupant. This means must be within (1) A primary ice detection system easy reach of each seated occupant and that automatically activates or alerts must be readily removable from the the flightcrew to activate the airframe airplane. ice protection system; (2) A definition of visual cues for rec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ognition of the first sign of ice accre- amended by Amdt. 25–29, 36 FR 18722, Sept.

tion on a specified surface combined 21, 1971; Amdt. 25–50, 45 FR 38348, June 9, 1980; with an advisory ice detection system Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt.

that alerts the flightcrew to activate 25–82, 59 FR 32057, June 21, 1994] the airframe ice protection system; or § 25.1419 Ice protection.

(3) Identification of conditions con- ducive to airframe icing as defined by If the applicant seeks certification an appropriate static or total air tem- for flight in icing conditions, the air- perature and visible moisture for use plane must be able to safely operate in by the flightcrew to activate the air- the continuous maximum and inter- frame ice protection system.

mittent maximum icing conditions of (f) Unless the applicant shows that appendix C. To establish this— the airframe ice protection system (a) An analysis must be performed to need not be operated during specific establish that the ice protection for phases of flight, the requirements of the various components of the airplane paragraph (e) of this section are appli- is adequate, taking into account the cable to all phases of flight.

various airplane operational configura- (g) After the initial activation of the tions; and airframe ice protection system— (b) To verify the ice protection anal- (1) The ice protection system must be ysis, to check for icing anomalies, and designed to operate continuously; to demonstrate that the ice protection (2) The airplane must be equipped system and its components are effec- with a system that automatically cy- tive, the airplane or its components cles the ice protection system; or must be flight tested in the various (3) An ice detection system must be operational configurations, in meas- provided to alert the flightcrew each ured natural atmospheric icing condi- time the ice protection system must be tions and, as found necessary, by one cycled.

or more of the following means: (h) Procedures for operation of the (1) Laboratory dry air or simulated ice protection system, including acti- icing tests, or a combination of both, of vation and deactivation, must be estab- the components or models of the com- lished and documented in the Airplane ponents.

Flight Manual.

(2) Flight dry air tests of the ice pro- tection system as a whole, or of its in- [Amdt. 25–72, 55 FR 29785, July 20, 1990, as dividual components.

amended by Amdt. 25–121, 72 FR 44669, Aug. 8, (3) Flight tests of the airplane or its 2007; Amdt. 25–129, 74 FR 38339, Aug. 3, 2009] components in measured simulated § 25.1420 Supercooled large drop icing icing conditions.

conditions.

(c) Caution information, such as an amber caution light or equivalent, (a) If certification for flight in icing must be provided to alert the conditions is sought, in addition to the flightcrew when the anti-ice or de-ice requirements of § 25.1419, an airplane system is not functioning normally. with a maximum takeoff weight less (d) For turbine engine powered air- than 60,000 pounds or with reversible planes, the ice protection provisions of flight controls must be capable of oper- this section are considered to be appli- ating in accordance with paragraphs cable primarily to the airframe. For (a)(1), (2), or (3), of this section.

Federal Aviation Administration, DOT § 25.1423 (1) Operating safely after encoun- (d) For the purposes of this section, tering the icing conditions defined in the following definitions apply: Appendix O of this part: (1) Reversible Flight Controls. Flight controls in the normal operating con- (i) The airplane must have a means figuration that have force or motion to detect that it is operating in Appen- originating at the airplane’s control dix O icing conditions; and surface (for example, through aero- (ii) Following detection of Appendix dynamic loads, static imbalance, or O icing conditions, the airplane must trim or servo tab inputs) that is trans- be capable of operating safely while mitted back to flight deck controls.

exiting all icing conditions.

This term refers to flight deck controls (2) Operating safely in a portion of connected to the pitch, roll, or yaw the icing conditions defined in Appen- control surfaces by direct mechanical dix O of this part as selected by the ap- linkages, cables, or push-pull rods in plicant: such a way that pilot effort produces (i) The airplane must have a means motion or force about the hinge line.

to detect that it is operating in condi- (2) Simulated Icing Test. Testing con- tions that exceed the selected portion ducted in simulated icing conditions, of Appendix O icing conditions; and such as in an icing tunnel or behind an (ii) Following detection, the airplane icing tanker.

must be capable of operating safely (3) Simulated Ice Shape. Ice shape fab- while exiting all icing conditions.

ricated from wood, epoxy, or other ma- (3) Operating safely in the icing con- terials by any construction technique.

ditions defined in Appendix O of this part. [Amdt. 25–140, 79 FR 65528, Nov. 4, 2014] (b) To establish that the airplane can § 25.1421 Megaphones.

operate safely as required in paragraph If a megaphone is installed, a re- (a) of this section, an applicant must straining means must be provided that show through analysis that the ice pro- is capable of restraining the mega- tection for the various components of phone when it is subjected to the ulti- the airplane is adequate, taking into mate inertia forces specified in account the various airplane oper- § 25.561(b)(3).

ational configurations. To verify the analysis, one, or more as found nec- [Amdt. 25–41, 42 FR 36970, July 18, 1977] essary, of the following methods must be used: § 25.1423 Public address system.

(1) Laboratory dry air or simulated A public address system required by icing tests, or a combination of both, of this chapter must— the components or models of the com- (a) Be powerable when the aircraft is ponents.

in flight or stopped on the ground, (2) Laboratory dry air or simulated after the shutdown or failure of all en- icing tests, or a combination of both, of gines and auxiliary power units, or the models of the airplane.

disconnection or failure of all power (3) Flight tests of the airplane or its sources dependent on their continued components in simulated icing condi- operation, for— tions, measured as necessary to sup- (1) A time duration of at least 10 min- port the analysis.

utes, including an aggregate time dura- (4) Flight tests of the airplane with tion of at least 5 minutes of announce- simulated ice shapes.

ments made by flight and cabin crew- (5) Flight tests of the airplane in nat- members, considering all other loads ural icing conditions, measured as nec- which may remain powered by the essary to support the analysis.

same source when all other power (c) For an airplane certified in ac- sources are inoperative; and cordance with paragraph (a)(2) or (3) of (2) An additional time duration in its this section, the requirements of standby state appropriate or required § 25.1419(e), (f), (g), and (h) must be met for any other loads that are powered by for the icing conditions defined in Ap- the same source and that are essential pendix O of this part in which the air- to safety of flight or required during plane is certified to operate. emergency conditions.

14 CFR Ch. I (1–1–25 Edition) § 25.1431 (b) Be capable of operation within 3 § 25.1433 Vacuum systems.

seconds from the time a microphone is There must be means, in addition to removed from its stowage.

the normal pressure relief, to auto- (c) Be intelligible at all passenger matically relieve the pressure in the seats, lavatories, and flight attendant discharge lines from the vacuum air seats and work stations.

pump when the delivery temperature of (d) Be designed so that no unused, the air becomes unsafe.

unstowed microphone will render the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as system inoperative.

amended by Amdt. 25–72, 55 FR 29785, July 20, (e) Be capable of functioning inde- 1990] pendently of any required crewmember interphone system.

§ 25.1435 Hydraulic systems.

(f) Be accessible for immediate use (a) Element design. Each element of from each of two flight crewmember the hydraulic system must be designed stations in the pilot compartment.

to: (g) For each required floor-level pas- (1) Withstand the proof pressure senger emergency exit which has an ad- without permanent deformation that jacent flight attendant seat, have a would prevent it from performing its microphone which is readily accessible intended functions, and the ultimate to the seated flight attendant, except pressure without rupture. The proof that one microphone may serve more and ultimate pressures are defined in than one exit, provided the proximity terms of the design operating pressure of the exits allows unassisted verbal (DOP) as follows: communication between seated flight Proof Ultimate attendants.

Element (xDOP) (xDOP) [Doc. No. 26003, 58 FR 45229, Aug. 26, 1993, as 1. Tubes and fittings. ......................... 1.5 3.0 amended by Amdt. 25–115, 69 FR 40527, July 2, 2. Pressure vessels containing gas: 2004] High pressure (e.g., accumula- tors) ......................................... 3.0 4.0 M ISCELLANEOUS E QUIPMENT Low pressure (e.g., reservoirs) .. 1.5 3.0 3. Hoses ............................................ 2.0 4.0 § 25.1431 Electronic equipment.

4. All other elements ......................... 1.5 2.0 (a) In showing compliance with (2) Withstand, without deformation § 25.1309 (a) and (b) with respect to that would prevent it from performing radio and electronic equipment and its intended function, the design oper- their installations, critical environ- ating pressure in combination with mental conditions must be considered.

limit structural loads that may be im- (b) Radio and electronic equipment posed; must be supplied with power under the (3) Withstand, without rupture, the requirements of § 25.1355(c).

design operating pressure multiplied by (c) Radio and electronic equipment, a factor of 1.5 in combination with ulti- controls, and wiring must be installed mate structural load that can reason- so that operation of any one unit or ably occur simultaneously; system of units will not adversely af- (4) Withstand the fatigue effects of fect the simultaneous operation of any all cyclic pressures, including tran- other radio or electronic unit, or sys- sients, and associated externally in- tem of units, required by this chapter.

duced loads, taking into account the (d) Electronic equipment must be de- consequences of element failure; and signed and installed such that it does (5) Perform as intended under all en- not cause essential loads to become in- vironmental conditions for which the operative as a result of electrical airplane is certificated.

power supply transients or transients (b) System design. Each hydraulic sys- from other causes.

tem must: (1) Have means located at a [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as flightcrew station to indicate appro- amended by Amdt. 25–113, 69 FR 12530, Mar.

16, 2004] priate system parameters, if Federal Aviation Administration, DOT § 25.1439 (i) It performs a function necessary range of motion of all associated user for continued safe flight and landing; systems. The test must be conducted at or the system relief pressure or 1.25 times (ii) In the event of hydraulic system the DOP if a system pressure relief de- malfunction, corrective action by the vice is not part of the system design.

crew to ensure continued safe flight Clearances between hydraulic system and landing is necessary; elements and other systems or struc- (2) Have means to ensure that system tural elements must remain adequate pressures, including transient pres- and there must be no detrimental ef- sures and pressures from fluid volu- fects.

metric changes in elements that are [Doc. No. 28617, 66 FR 27402, May 16, 2001] likely to remain closed long enough for such changes to occur, are within the § 25.1438 Pressurization and pneu- design capabilities of each element, matic systems.

such that they meet the requirements (a) Pressurization system elements defined in § 25.1435(a)(1) through (a)(5); must be burst pressure tested to 2.0 (3) Have means to minimize the re- times, and proof pressure tested to 1.5 lease of harmful or hazardous con- times, the maximum normal operating centrations of hydraulic fluid or vapors pressure.

into the crew and passenger compart- ments during flight; (b) Pneumatic system elements must (4) Meet the applicable requirements be burst pressure tested to 3.0 times, of §§ 25.863, 25.1183, 25.1185, and 25.1189 if and proof pressure tested to 1.5 times, a flammable hydraulic fluid is used; the maximum normal operating pres- and sure.

(5) Be designed to use any suitable (c) An analysis, or a combination of hydraulic fluid specified by the air- analysis and test, may be substituted plane manufacturer, which must be for any test required by paragraph (a) identified by appropriate markings as or (b) of this section if the Adminis- required by § 25.1541.

trator finds it equivalent to the re- (c) Tests. Tests must be conducted on quired test.

the hydraulic system(s), and/or sub- [Amdt. 25–41, 42 FR 36971, July 18, 1977] system(s) and elements, except that analysis may be used in place of or to § 25.1439 Protective breathing equip- supplement testing, where the analysis ment.

is shown to be reliable and appropriate.

(a) Fixed (stationary, or built in) pro- All internal and external influences tective breathing equipment must be must be taken into account to an ex- installed for the use of the flightcrew, tent necessary to evaluate their ef- and at least one portable protective fects, and to assure reliable system and breathing equipment shall be located element functioning and integration.

at or near the flight deck for use by a Failure or unacceptable deficiency of flight crewmember. In addition, port- an element or system must be cor- able protective breathing equipment rected and be sufficiently retested, must be installed for the use of appro- where necessary.

priate crewmembers for fighting fires (1) The system(s), subsystem(s), or in compartments accessible in flight element(s) must be subjected to per- other than the flight deck. This in- formance, fatigue, and endurance tests cludes isolated compartments and representative of airplane ground and upper and lower lobe galleys, in which flight operations.

crewmember occupancy is permitted (2) The complete system must be during flight. Equipment must be in- tested to determine proper functional stalled for the maximum number of performance and relation to the other crewmembers expected to be in the systems, including simulation of rel- area during any operation.

evant failure conditions, and to sup- port or validate element design. (b) For protective breathing equip- (3) The complete hydraulic system(s) ment required by paragraph (a) of this must be functionally tested on the air- section or by the applicable Operating plane in normal operation over the Regulations: 14 CFR Ch. I (1–1–25 Edition) § 25.1441 (1) The equipment must be designed § 25.1441 Oxygen equipment and sup- to protect the appropriate crewmember ply.

from smoke, carbon dioxide, and other (a) If certification with supplemental harmful gases while on flight deck oxygen equipment is requested, the duty or while combating fires. equipment must meet the requirements (2) The equipment must include— of this section and §§ 25.1443 through 25.1453.

(i) Masks covering the eyes, nose and (b) The oxygen system must be free mouth, or from hazards in itself, in its method of (ii) Masks covering the nose and operation, and in its effect upon other mouth, plus accessory equipment to components.

cover the eyes.

(c) There must be a means to allow (3) Equipment, including portable the crew to readily determine, during equipment, must allow communication flight, the quantity of oxygen available with other crewmembers while in use.

in each source of supply.

Equipment available at flightcrew as- (d) The oxygen flow rate and the oxy- signed duty stations must also enable gen equipment for airplanes for which the flightcrew to use radio equipment.

certification for operation above 40,000 (4) The part of the equipment pro- feet is requested must be approved.

tecting the eyes shall not cause any ap- § 25.1443 Minimum mass flow of sup- preciable adverse effect on vision and plemental oxygen.

must allow corrective glasses to be (a) If continuous flow equipment is worn.

installed for use by flight crew- (5) The equipment must supply pro- members, the minimum mass flow of tective oxygen of 15 minutes duration supplemental oxygen required for each per crewmember at a pressure altitude crewmember may not be less than the of 8,000 feet with a respiratory minute flow required to maintain, during in- volume of 30 liters per minute BTPD.

spiration, a mean tracheal oxygen par- The equipment and system must be de- tial pressure of 149 mm. Hg. when signed to prevent any inward leakage breathing 15 liters per minute, BTPS, to the inside of the device and prevent and with a maximum tidal volume of any outward leakage causing signifi- 700 cc. with a constant time interval cant increase in the oxygen content of between respirations.

the local ambient atmosphere. If a de- (b) If demand equipment is installed mand oxygen system is used, a supply for use by flight crewmembers, the of 300 liters of free oxygen at 70 ° F. and minimum mass flow of supplemental 760 mm. Hg. pressure is considered to oxygen required for each crewmember be of 15-minute duration at the pre- may not be less than the flow required scribed altitude and minute volume. If to maintain, during inspiration, a a continuous flow open circuit protec- mean tracheal oxygen partial pressure tive breathing system is used, a flow of 122 mm. Hg., up to and including a rate of 60 liters per minute at 8,000 feet cabin pressure altitude of 35,000 feet, and 95 percent oxygen between cabin (45 liters per minute at sea level) and a pressure altitudes of 35,000 and 40,000 supply of 600 liters of free oxygen at 70 feet, when breathing 20 liters per ° F. and 760 mm. Hg. pressure is consid- minute BTPS. In addition, there must ered to be of 15-minute duration at the be means to allow the crew to use undi- prescribed altitude and minute volume.

luted oxygen at their discretion.

Continuous flow systems must not in- (c) For passengers and cabin attend- crease the ambient oxygen content of ants, the minimum mass flow of sup- the local atmosphere above that of de- plemental oxygen required for each mand systems. BTPD refers to body person at various cabin pressure alti- temperature conditions (that is, 37 ° C., tudes may not be less than the flow re- at ambient pressure, dry).

quired to maintain, during inspiration (6) The equipment must meet the re- and while using the oxygen equipment quirements of § 25.1441.

(including masks) provided, the fol- lowing mean tracheal oxygen partial [Doc. No. FAA–2002–13859, 69 FR 40528, July 2, 2004] pressures: Federal Aviation Administration, DOT § 25.1447 (1) At cabin pressure altitudes above whom supplemental oxygen is to be 10,000 feet up to and including 18,500 supplied. Units must be designed to feet, a mean tracheal oxygen partial cover the nose and mouth and must be pressure of 100 mm. Hg. when breathing equipped with a suitable means to re- 15 liters per minute, BTPS, and with a tain the unit in position on the face.

tidal volume of 700 cc. with a constant Flight crew masks for supplemental time interval between respirations.

oxygen must have provisions for the (2) At cabin pressure altitudes above use of communication equipment.

18,500 feet up to and including 40,000 (b) If certification for operation up to feet, a mean tracheal oxygen partial and including 25,000 feet is requested, pressure of 83.8 mm. Hg. when breath- an oxygen supply terminal and unit of ing 30 liters per minute, BTPS, and oxygen dispensing equipment for the with a tidal volume of 1,100 cc. with a immediate use of oxygen by each crew- constant time interval between res- member must be within easy reach of pirations.

that crewmember. For any other occu- (d) If first-aid oxygen equipment is pants, the supply terminals and dis- installed, the minimum mass flow of pensing equipment must be located to oxygen to each user may not be less allow the use of oxygen as required by than four liters per minute, STPD.

the operating rules in this chapter.

However, there may be a means to de- (c) If certification for operation crease this flow to not less than two li- above 25,000 feet is requested, there ters per minute, STPD, at any cabin al- must be oxygen dispensing equipment titude. The quantity of oxygen re- meeting the following requirements: quired is based upon an average flow (1) There must be an oxygen dis- rate of three liters per minute per per- son for whom first-aid oxygen is re- pensing unit connected to oxygen sup- quired. ply terminals immediately available to (e) If portable oxygen equipment is each occupant wherever seated, and at installed for use by crewmembers, the least two oxygen dispensing units con- minimum mass flow of supplemental nected to oxygen terminals in each lav- oxygen is the same as specified in para- atory. The total number of dispensing graph (a) or (b) of this section, which- units and outlets in the cabin must ex- ever is applicable.

ceed the number of seats by at least 10 percent. The extra units must be as § 25.1445 Equipment standards for the uniformly distributed throughout the oxygen distributing system.

cabin as practicable. Except as pro- (a) When oxygen is supplied to both vided in paragraph (c)(5) of this sec- crew and passengers, the distribution tion, if certification for operation system must be designed for either— above 30,000 feet is requested, the dis- (1) A source of supply for the flight pensing units providing the required crew on duty and a separate source for oxygen flow must be automatically the passengers and other crewmembers; presented to the occupants before the or cabin pressure altitude exceeds 15,000 (2) A common source of supply with feet. The crewmembers must be pro- means to separately reserve the min- vided with a manual means of making imum supply required by the flight the dispensing units immediately crew on duty.

available in the event of failure of the (b) Portable walk-around oxygen automatic system.

units of the continuous flow, diluter- (2) Each flight crewmember on flight demand, and straight demand kinds deck duty must be provided with a may be used to meet the crew or pas- quick-donning type oxygen dispensing senger breathing requirements.

unit connected to an oxygen supply § 25.1447 Equipment standards for ox- terminal. This dispensing unit must be ygen dispensing units.

immediately available to the flight crewmember when seated at his sta- If oxygen dispensing units are in- tion, and installed so that it: stalled, the following apply: (a) There must be an individual dis- (i) Can be placed on the face from its pensing unit for each occupant for ready position, properly secured, 14 CFR Ch. I (1–1–25 Edition) § 25.1449 sealed, and supplying oxygen upon de- as a device which produces oxygen by mand, with one hand, within five sec- chemical reaction.

onds and without disturbing eyeglasses (b) Each chemical oxygen generator or causing delay in proceeding with must be designed and installed in ac- emergency duties; and cordance with the following require- (ii) Allows, while in place, the per- ments: formance of normal communication (1) Surface temperature developed by functions. the generator during operation may (3) The oxygen dispensing equipment not create a hazard to the airplane or for the flight crewmembers must be: to its occupants.

(i) The diluter demand or pressure de- (2) Means must be provided to relieve mand (pressure demand mask with a any internal pressure that may be haz- diluter demand pressure breathing reg- ardous.

ulator) type, or other approved oxygen (3) Except as provided in SFAR 109, equipment shown to provide the same each chemical oxygen generator instal- degree of protection, for airplanes to be lation must meet the requirements of operated above 25,000 feet.

§ 25.795(d).

(ii) The pressure demand (pressure (c) In addition to meeting the re- demand mask with a diluter demand quirements in paragraph (b) of this sec- pressure breathing regulator) type with tion, each portable chemical oxygen mask-mounted regulator, or other ap- generator that is capable of sustained proved oxygen equipment shown to operation by successive replacement of provide the same degree of protection, a generator element must be placarded for airplanes operated at altitudes to show— where decompressions that are not ex- (1) The rate of oxygen flow, in liters tremely improbable may expose the per minute; flightcrew to cabin pressure altitudes (2) The duration of oxygen flow, in in excess of 34,000 feet.

minutes, for the replaceable generator (4) Portable oxygen equipment must element; and be immediately available for each (3) A warning that the replaceable cabin attendant. The portable oxygen generator element may be hot, unless equipment must have the oxygen dis- the element construction is such that pensing unit connected to the portable the surface temperature cannot exceed oxygen supply.

100 degrees F.

(5) When operating into or out of air- [Amdt. 25–41, 42 FR 36971, July 18, 1977, as ports with elevations above 13,000 feet, amended at 79 FR 13519, Mar. 11, 2014] the dispensing units providing the re- quired oxygen flow must be automati- § 25.1453 Protection of oxygen equip- cally presented to the occupants at ment from rupture.

cabin pressure altitudes no higher than Oxygen pressure tanks, and lines be- 2,000 feet above the airplane’s max- tween tanks and the shutoff means, imum takeoff and landing altitude.

must be— [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) Protected from unsafe tempera- amended by Amdt. 25–41, 42 FR 36971, July 18, tures; and 1977; Amdt. 25–87, 61 FR 28696, June 5, 1996; (b) Located where the probability and Amdt. 25–116, 69 FR 62789, Oct. 27, 2004; Amdt.

hazards of rupture in a crash landing No. 25–151, 88 FR 39161, June 15, 2023; 88 FR are minimized.

44032, July 11, 2023] § 25.1455 Draining of fluids subject to § 25.1449 Means for determining use of freezing.

oxygen.

If fluids subject to freezing may be There must be a means to allow the drained overboard in flight or during crew to determine whether oxygen is ground operation, the drains must be being delivered to the dispensing equip- designed and located to prevent the ment.

formation of hazardous quantities of ice on the airplane as a result of the § 25.1450 Chemical oxygen generators.

drainage.

(a) For the purpose of this section, a chemical oxygen generator is defined [Amdt. 25–23, 35 FR 5680, Apr. 8, 1970] Federal Aviation Administration, DOT § 25.1457 (1) For the first channel, from each § 25.1457 Cockpit voice recorders.

boom, mask, or hand-held microphone, (a) Each cockpit voice recorder re- headset, or speaker used at the first quired by the operating rules of this pilot station.

chapter must be approved and must be (2) For the second channel from each installed so that it will record the fol- boom, mask, or hand-held microphone, lowing: headset, or speaker used at the second (1) Voice communications trans- pilot station.

mitted from or received in the airplane (3) For the third channel—from the by radio.

cockpit-mounted area microphone.

(2) Voice communications of flight (4) For the fourth channel, from— crewmembers on the flight deck.

(i) Each boom, mask, or hand-held (3) Voice communications of flight microphone, headset, or speaker used crewmembers on the flight deck, using at the station for the third and fourth the airplane’s interphone system.

crew members; or (4) Voice or audio signals identifying (ii) If the stations specified in para- navigation or approach aids introduced graph (c)(4)(i) of this section are not re- into a headset or speaker.

quired or if the signal at such a station (5) Voice communications of flight is picked up by another channel, each crewmembers using the passenger loud- microphone on the flight deck that is speaker system, if there is such a sys- used with the passenger loudspeaker tem and if the fourth channel is avail- system, if its signals are not picked up able in accordance with the require- by another channel.

ments of paragraph (c)(4)(ii) of this sec- (5) As far as is practicable all sounds tion.

received by the microphone listed in (6) If datalink communication equip- paragraphs (c)(1), (2), and (4) of this ment is installed, all datalink commu- section must be recorded without nications, using an approved data mes- interruption irrespective of the posi- sage set. Datalink messages must be tion of the interphone-transmitter key recorded as the output signal from the switch. The design shall ensure that communications unit that translates sidetone for the flight crew is produced the signal into usable data.

only when the interphone, public ad- (b) The recording requirements of dress system, or radio transmitters are paragraph (a)(2) of this section must be in use.

met by installing a cockpit-mounted (d) Each cockpit voice recorder must area microphone, located in the best be installed so that— position for recording voice commu- (1)(i) It receives its electrical power nications originating at the first and from the bus that provides the max- second pilot stations and voice commu- imum reliability for operation of the nications of other crewmembers on the cockpit voice recorder without jeopard- flight deck when directed to those sta- izing service to essential or emergency tions. The microphone must be so lo- loads.

cated and, if necessary, the pre- (ii) It remains powered for as long as amplifiers and filters of the recorder possible without jeopardizing emer- must be so adjusted or supplemented, gency operation of the airplane.

that the intelligibility of the recorded (2) There is an automatic means to communications is as high as prac- simultaneously stop the recorder and ticable when recorded under flight prevent each erasure feature from func- cockpit noise conditions and played tioning, within 10 minutes after crash back. Repeated aural or visual play- impact; back of the record may be used in eval- (3) There is an aural or visual means uating intelligibility. for preflight checking of the recorder (c) Each cockpit voice recorder must for proper operation; be installed so that the part of the (4) Any single electrical failure exter- communication or audio signals speci- nal to the recorder does not disable fied in paragraph (a) of this section ob- both the cockpit voice recorder and the tained from each of the following flight data recorder; sources is recorded on a separate chan- (5) It has an independent power nel: source— 14 CFR Ch. I (1–1–25 Edition) § 25.1459 (i) That provides 10 ± 1 minutes of the container which is secured in such electrical power to operate both the manner that they are not likely to be cockpit voice recorder and cockpit- separated during crash impact.

mounted area microphone; [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (ii) That is located as close as prac- amended by Amdt. 25–2, 30 FR 3932, Mar. 26, ticable to the cockpit voice recorder; 1965; Amdt. 25–16, 32 FR 13914, Oct. 6, 1967; and Amdt. 25–41, 42 FR 36971, July 18, 1977; Amdt.

(iii) To which the cockpit voice re- 25–65, 53 FR 26143, July 11, 1988; Amdt. 25–124, 73 FR 12563, Mar. 7, 2008; 74 FR 32800, July 9, corder and cockpit-mounted area 2009] microphone are switched automati- cally in the event that all other power § 25.1459 Flight data recorders.

to the cockpit voice recorder is inter- (a) Each flight recorder required by rupted either by normal shutdown or the operating rules of this chapter by any other loss of power to the elec- must be installed so that— trical power bus; and (1) It is supplied with airspeed, alti- (6) It is in a separate container from tude, and directional data obtained the flight data recorder when both are from sources that meet the accuracy required. If used to comply with only requirements of §§ 25.1323, 25.1325, and the cockpit voice recorder require- 25.1327, as appropriate; ments, a combination unit may be in- stalled. (2) The vertical acceleration sensor is (e) The recorder container must be rigidly attached, and located longitu- located and mounted to minimize the dinally either within the approved cen- probability of rupture of the container ter of gravity limits of the airplane, or as a result of crash impact and con- at a distance forward or aft of these sequent heat damage to the recorder limits that does not exceed 25 percent from fire. of the airplane’s mean aerodynamic (1) Except as provided in paragraph chord; (e)(2) of this section, the recorder con- (3)(i) It receives its electrical power tainer must be located as far aft as from the bus that provides the max- practicable, but need not be outside of imum reliability for operation of the the pressurized compartment, and may flight data recorder without jeopard- not be located where aft-mounted en- izing service to essential or emergency gines may crush the container during loads.

impact. (ii) It remains powered for as long as (2) If two separate combination dig- possible without jeopardizing emer- ital flight data recorder and cockpit gency operation of the airplane.

voice recorder units are installed in- (4) There is an aural or visual means stead of one cockpit voice recorder and for preflight checking of the recorder one digital flight data recorder, the for proper recording of data in the stor- combination unit that is installed to age medium; comply with the cockpit voice recorder (5) Except for recorders powered sole- requirements may be located near the ly by the engine-driven electrical gen- cockpit. erator system, there is an automatic (f) If the cockpit voice recorder has a means to simultaneously stop a re- bulk erasure device, the installation corder that has a data erasure feature must be designed to minimize the prob- and prevent each erasure feature from ability of inadvertent operation and ac- functioning, within 10 minutes after tuation of the device during crash im- crash impact; pact. (6) There is a means to record data (g) Each recorder container must— from which the time of each radio (1) Be either bright orange or bright transmission either to or from ATC can yellow; be determined; (2) Have reflective tape affixed to its (7) Any single electrical failure exter- external surface to facilitate its loca- nal to the recorder does not disable tion under water; and both the cockpit voice recorder and the (3) Have an underwater locating de- flight data recorder; and vice, when required by the operating (8) It is in a separate container from rules of this chapter, on or adjacent to the cockpit voice recorder when both Federal Aviation Administration, DOT § 25.1503 are required. If used to comply with § 25.1461 Equipment containing high only the flight data recorder require- energy rotors.

ments, a combination unit may be in- (a) Equipment containing high en- stalled. If a combination unit is in- ergy rotors must meet paragraph (b), stalled as a cockpit voice recorder to (c), or (d) of this section.

comply with § 25.1457(e)(2), a combina- (b) High energy rotors contained in tion unit must be used to comply with equipment must be able to withstand this flight data recorder requirement.

damage caused by malfunctions, vibra- (b) Each nonejectable record con- tion, abnormal speeds, and abnormal tainer must be located and mounted so temperatures. In addition— as to minimize the probability of con- (1) Auxiliary rotor cases must be able to contain damage caused by the fail- tainer rupture resulting from crash im- ure of high energy rotor blades; and pact and subsequent damage to the (2) Equipment control devices, sys- record from fire. In meeting this re- tems, and instrumentation must rea- quirement the record container must sonably ensure that no operating limi- be located as far aft as practicable, but tations affecting the integrity of high need not be aft of the pressurized com- energy rotors will be exceeded in serv- partment, and may not be where aft- ice.

mounted engines may crush the con- (c) It must be shown by test that tainer upon impact.

equipment containing high energy ro- (c) A correlation must be established tors can contain any failure of a high between the flight recorder readings of energy rotor that occurs at the highest airspeed, altitude, and heading and the speed obtainable with the normal speed corresponding readings (taking into ac- control devices inoperative.

count correction factors) of the first pi- (d) Equipment containing high en- lot’s instruments. The correlation ergy rotors must be located where must cover the airspeed range over rotor failure will neither endanger the which the airplane is to be operated, occupants nor adversely affect contin- the range of altitude to which the air- ued safe flight.

plane is limited, and 360 degrees of [Amdt. 25–41, 42 FR 36971, July 18, 1977] heading. Correlation may be estab- lished on the ground as appropriate.

Subpart G—Operating Limitations (d) Each recorder container must— and Information (1) Be either bright orange or bright yellow; § 25.1501 General.

(2) Have reflective tape affixed to its (a) Each operating limitation speci- external surface to facilitate its loca- fied in §§ 25.1503 through 25.1533 and tion under water; and other limitations and information nec- (3) Have an underwater locating de- essary for safe operation must be es- vice, when required by the operating tablished.

rules of this chapter, on or adjacent to (b) The operating limitations and the container which is secured in such other information necessary for safe a manner that they are not likely to be operation must be made available to separated during crash impact.

the crewmembers as prescribed in (e) Any novel or unique design or §§ 25.1541 through 25.1587.

operational characteristics of the air- [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978] craft shall be evaluated to determine if any dedicated parameters must be re- O PERATING L IMITATIONS corded on flight recorders in addition to or in place of existing requirements.

§ 25.1503 Airspeed limitations: general.

[Amdt. 25–8, 31 FR 127, Jan. 6, 1966, as amend- When airspeed limitations are a func- ed by Amdt. 25–25, 35 FR 13192, Aug. 19, 1970; tion of weight, weight distribution, al- Amdt. 25–37, 40 FR 2577, Jan. 14, 1975; Amdt.

titude, or Mach number, limitations 25–41, 42 FR 36971, July 18, 1977; Amdt. 25–65, corresponding to each critical com- 53 FR 26144, July 11, 1988; Amdt. 25–124, 73 FR bination of these factors must be estab- 12563, Mar. 7, 2008; 74 FR 32800, July 9, 2009] lished.

14 CFR Ch. I (1–1–25 Edition) § 25.1505 tended position, and that determined § 25.1505 Maximum operating limit speed. under § 25.729.

The maximum operating limit speed [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ( V / M airspeed or Mach Number, amended by Amdt. 25–38, 41 FR 55468, Dec. 20, MO MO 1976] whichever is critical at a particular al- titude) is a speed that may not be de- § 25.1516 Other speed limitations.

liberately exceeded in any regime of flight (climb, cruise, or descent), unless Any other limitation associated with a higher speed is authorized for flight speed must be established.

test or pilot training operations. V / MO [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] M must be established so that it is MO not greater than the design cruising § 25.1517 Rough air speed, V RA.

speed V and so that it is sufficiently C (a) A rough air speed, V for use as RA, below V / M or V / M to make it D D DF DF, the recommended turbulence penetra- highly improbable that the latter tion airspeed, and a rough air Mach speeds will be inadvertently exceeded number, M , for use as the rec- RA in operations. The speed margin be- ommended turbulence penetration tween V / M and V / M or V M/ MO MO D D DF DF Mach number, must be established.

may not be less than that determined V /M must be sufficiently less than RA RA under § 25.335(b) or found necessary dur- V /M to ensure that likely speed MO MO ing the flight tests conducted under variation during rough air encounters § 25.253.

will not cause the overspeed warning to [Amdt. 25–23, 35 FR 5680, Apr. 8, 1970] operate too frequently.

(b) At altitudes where V is not lim- MO § 25.1507 Maneuvering speed.

ited by Mach number, in the absence of The maneuvering speed must be es- a rational investigation substantiating tablished so that it does not exceed the the use of other values, V must be RA design maneuvering speed V deter- less than V minus 35 KTAS.

A MO mined under § 25.335(c). (c) At altitudes where V is limited MO by Mach number, M may be chosen RA § 25.1511 Flap extended speed.

to provide an optimum margin between low and high speed buffet boundaries.

The established flap extended speed V must be established so that it does FE [Amdt. 25–141, 79 FR 73469, Dec. 11, 2014, as not exceed the design flap speed V F amended by FAA–2022–1355; Amdt. No. 25–148, chosen under §§ 25.335(e) and 25.345, for 87 FR 75710, Dec. 9, 2022; 88 FR 2813, Jan. 18, the corresponding flap positions and 2023] engine powers.

§ 25.1519 Weight, center of gravity, and weight distribution.

§ 25.1513 Minimum control speed.

The minimum control speed V de- The airplane weight, center of grav- MC termined under § 25.149 must be estab- ity, and weight distribution limita- lished as an operating limitation. tions determined under §§ 25.23 through 25.27 must be established as operating § 25.1515 Landing gear speeds.

limitations.

(a) The established landing gear oper- § 25.1521 Powerplant limitations.

ating speed or speeds, V may not ex- LO, ceed the speed at which it is safe both (a) General. The powerplant limita- to extend and to retract the landing tions prescribed in this section must be gear, as determined under § 25.729 or by established so that they do not exceed flight characteristics. If the extension the corresponding limits for which the speed is not the same as the retraction engines or propellers are type certifi- speed, the two speeds must be des- cated and do not exceed the values on ignated as V and V respec- which compliance with any other re- LO ( EXT ) LO ( RET ) , tively. quirement of this part is based.

(b) The established landing gear ex- (b) Reciprocating engine installations.

tended speed V may not exceed the Operating limitations relating to the LE speed at which it is safe to fly with the following must be established for recip- landing gear secured in the fully ex- rocating engine installations: Federal Aviation Administration, DOT § 25.1529 (1) Horsepower or torque, r.p.m., § 25.1522 Auxiliary power unit limita- manifold pressure, and time at critical tions.

pressure altitude and sea level pressure If an auxiliary power unit is installed altitude for— in the airplane, limitations established (i) Maximum continuous power (re- for the auxiliary power unit, including lating to unsupercharged operation or categories of operation, must be speci- to operation in each supercharger mode fied as operating limitations for the as applicable); and airplane.

(ii) Takeoff power (relating to unsu- percharged operation or to operation in [Amdt. 25–72, 55 FR 29786, July 20, 1990] each supercharger mode as applicable).

§ 25.1523 Minimum flight crew.

(2) Fuel grade or specification.

(3) Cylinder head and oil tempera- The minimum flight crew must be es- tures.

tablished so that it is sufficient for safe (4) Any other parameter for which a operation, considering— limitation has been established as part (a) The workload on individual crew- of the engine type certificate except members; that a limitation need not be estab- (b) The accessibility and ease of oper- lished for a parameter that cannot be ation of necessary controls by the ap- exceeded during normal operation due propriate crewmember; and to the design of the installation or to (c) The kind of operation authorized another established limitation.

under § 25.1525.

(c) Turbine engine installations. Oper- The criteria used in making the deter- ating limitations relating to the fol- minations required by this section are lowing must be established for turbine set forth in appendix D.

engine installations: (1) Horsepower, torque or thrust, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as r.p.m., gas temperature, and time for— amended by Amdt. 25–3, 30 FR 6067, Apr. 29, (i) Maximum continuous power or 1965] thrust (relating to augmented or un- § 25.1525 Kinds of operation.

augmented operation as applicable).

(ii) Takeoff power or thrust (relating The kinds of operation to which the to augmented or unaugmented oper- airplane is limited are established by ation as applicable).

the category in which it is eligible for (2) Fuel designation or specification.

certification and by the installed (3) Maximum time interval between equipment.

engine run-ups from idle, run-up power setting and duration at power for § 25.1527 Ambient air temperature and ground operation in icing conditions, operating altitude.

as defined in § 25.1093(b)(2).

The extremes of the ambient air tem- (4) Any other parameter for which a perature and operating altitude for limitation has been established as part which operation is allowed, as limited of the engine type certificate except by flight, structural, powerplant, func- that a limitation need not be estab- tional, or equipment characteristics, lished for a parameter that cannot be must be established.

exceeded during normal operation due to the design of the installation or to [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] another established limitation.

§ 25.1529 Instructions for Continued (d) Ambient temperature. An ambient Airworthiness.

temperature limitation (including lim- itations for winterization installations, The applicant must prepare Instruc- if applicable) must be established as tions for Continued Airworthiness in the maximum ambient atmospheric accordance with appendix H to this temperature established in accordance part that are acceptable to the Admin- with § 25.1043(b).

istrator. The instructions may be in- complete at type certification if a pro- [Amdt. 25–72, 55 FR 29786, July 20, 1990, as gram exists to ensure their completion amended by Amdt. 25–140, 79 FR 65528, Nov. 4, 2014] prior to delivery of the first airplane or 14 CFR Ch. I (1–1–25 Edition) § 25.1531 issuance of a standard certificate of conditions defined in Appendix O of airworthiness, whichever occurs later. this part for which the airplane has not been certified to safely operate; and [Amdt. 25–54, 45 FR 60173, Sept. 11, 1980] (2) Require exiting all icing condi- § 25.1531 Maneuvering flight load fac- tions if icing conditions defined in Ap- tors.

pendix O of this part are encountered Load factor limitations, not exceed- for which the airplane has not been ing the positive limit load factors de- certified to safely operate.

termined from the maneuvering dia- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as gram in § 25.333(b), must be established.

amended by Amdt. 25–38, 41 FR 55468, Dec. 20, 1976; Amdt. 25–72, 55 FR 29786, July 20, 1990; § 25.1533 Additional operating limita- Amdt. 25–92, 63 FR 8321, Feb. 18, 1998; Amdt.

tions.

25–140, 79 FR 65528, Nov. 4, 2014] (a) Additional operating limitations must be established as follows: § 25.1535 ETOPS approval.

(1) The maximum takeoff weights Except as provided in § 25.3, each ap- must be established as the weights at plicant seeking ETOPS type design ap- which compliance is shown with the proval must comply with the provi- applicable provisions of this part (in- sions of Appendix K of this part.

cluding the takeoff climb provisions of § 25.121(a) through (c), for altitudes and [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, ambient temperatures).

2007] (2) The maximum landing weights must be established as the weights at M ARKINGS AND P LACARDS which compliance is shown with the applicable provisions of this part (in- § 25.1541 General.

cluding the landing and approach climb (a) The airplane must contain— provisions of §§ 25.119 and 25.121(d) for (1) The specified markings and plac- altitudes and ambient temperatures).

ards; and (3) The minimum takeoff distances (2) Any additional information, in- must be established as the distances at strument markings, and placards re- which compliance is shown with the quired for the safe operation if there applicable provisions of this part (in- are unusual design, operating, or han- cluding the provisions of §§ 25.109 and 25.113, for weights, altitudes, tempera- dling characteristics.

tures, wind components, runway sur- (b) Each marking and placard pre- face conditions (dry and wet), and run- scribed in paragraph (a) of this sec- way gradients) for smooth, hard-sur- tion— faced runways. Additionally, at the op- (1) Must be displayed in a con- tion of the applicant, wet runway take- spicuous place; and off distances may be established for (2) May not be easily erased, dis- runway surfaces that have been figured, or obscured.

grooved or treated with a porous fric- tion course, and may be approved for § 25.1543 Instrument markings: gen- use on runways where such surfaces eral.

have been designed constructed, and For each instrument— maintained in a manner acceptable to (a) When markings are on the cover the Administrator.

glass of the instrument, there must be (b) The extremes for variable factors means to maintain the correct align- (such as altitude, temperature, wind, and runway gradients) are those at ment of the glass cover with the face of which compliance with the applicable the dial; and provisions of this part is shown.

(b) Each instrument marking must (c) For airplanes certified in accord- be clearly visible to the appropriate ance with § 25.1420(a)(1) or (2), an oper- crewmember.

ating limitation must be established [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as to: amended by Amdt. 25–72, 55 FR 29786, July 20, (1) Prohibit intentional flight, in- 1990] cluding takeoff and landing, into icing Federal Aviation Administration, DOT § 25.1557 § 25.1545 Airspeed limitation informa- § 25.1555 Control markings.

tion.

(a) Each cockpit control, other than The airspeed limitations required by primary flight controls and controls § 25.1583 (a) must be easily read and un- whose function is obvious, must be derstood by the flight crew. plainly marked as to its function and method of operation.

§ 25.1547 Magnetic direction indicator.

(b) Each aerodynamic control must be marked under the requirements of (a) A placard meeting the require- §§ 25.677 and 25.699.

ments of this section must be installed (c) For powerplant fuel controls— on, or near, the magnetic direction in- (1) Each fuel tank selector control dicator.

must be marked to indicate the posi- (b) The placard must show the cali- tion corresponding to each tank and to bration of the instrument in level each existing cross feed position; flight with the engines operating.

(2) If safe operation requires the use (c) The placard must state whether of any tanks in a specific sequence, the calibration was made with radio re- that sequence must be marked on, or ceivers on or off.

adjacent to, the selector for those (d) Each calibration reading must be tanks; and in terms of magnetic heading in not (3) Each valve control for each engine more than 45 degree increments.

must be marked to indicate the posi- tion corresponding to each engine con- § 25.1549 Powerplant and auxiliary trolled.

power unit instruments.

(d) For accessory, auxiliary, and For each required powerplant and emergency controls— auxiliary power unit instrument, as ap- (1) Each emergency control (includ- propriate to the type of instrument— ing each fuel jettisoning and fluid shut- (a) Each maximum and, if applicable, off must be colored red; and minimum safe operating limit must be (2) Each visual indicator required by marked with a red radial or a red line; § 25.729(e) must be marked so that the (b) Each normal operating range pilot can determine at any time when must be marked with a green arc or the wheels are locked in either extreme green line, not extending beyond the position, if retractable landing gear is maximum and minimum safe limits; used.

(c) Each takeoff and precautionary range must be marked with a yellow § 25.1557 Miscellaneous markings and arc or a yellow line; and placards.

(d) Each engine, auxiliary power (a) Baggage and cargo compartments unit, or propeller speed range that is and ballast location. Each baggage and restricted because of excessive vibra- cargo compartment, and each ballast tion stresses must be marked with red location must have a placard stating arcs or red lines.

any limitations on contents, including weight, that are necessary under the [Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] loading requirements. However, § 25.1551 Oil quantity indication.

underseat compartments designed for the storage of carry-on articles weigh- Each oil quantity indicating means ing not more than 20 pounds need not must be marked to indicate the quan- have a loading limitation placard.

tity of oil readily and accurately.

(b) Powerplant fluid filler openings.

[Amdt. 25–72, 55 FR 29786, July 20, 1990] The following apply: (1) Fuel filler openings must be § 25.1553 Fuel quantity indicator.

marked at or near the filler cover If the unusable fuel supply for any with— tank exceeds one gallon, or five per- (i) The word ‘‘fuel’’; cent of the tank capacity, whichever is (ii) For reciprocating engine powered greater, a red arc must be marked on airplanes, the minimum fuel grade; its indicator extending from the cali- (iii) For turbine engine powered air- brated zero reading to the lowest read- planes, the permissible fuel designa- ing obtainable in level flight. tions; and 14 CFR Ch. I (1–1–25 Edition) § 25.1561 (iv) For pressure fueling systems, the A IRPLANE F LIGHT M ANUAL maximum permissible fueling supply § 25.1581 General.

pressure and the maximum permissible defueling pressure. (a) Furnishing information. An Air- plane Flight Manual must be furnished (2) Oil filler openings must be with each airplane, and it must contain marked at or near the filler cover with the following: the word ‘‘oil’’.

(1) Information required by §§ 25.1583 (3) Augmentation fluid filler open- through 25.1587.

ings must be marked at or near the (2) Other information that is nec- filler cover to identify the required essary for safe operation because of de- fluid.

sign, operating, or handling character- (c) Emergency exit placards. Each istics.

emergency exit placard must meet the (3) Any limitation, procedure, or requirements of § 25.811.

other information established as a con- (d) Doors. Each door that must be dition of compliance with the applica- used in order to reach any required ble noise standards of part 36 of this emergency exit must have a suitable chapter.

placard stating that the door is to be (b) Approved information. Each part of latched in the open position during the manual listed in §§ 25.1583 through takeoff and landing. 25.1587, that is appropriate to the air- plane, must be furnished, verified, and [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as approved, and must be segregated, amended by Amdt. 25–32, 37 FR 3972, Feb. 24, identified, and clearly distinguished 1972; Amdt. 25–38, 41 FR 55468, Dec. 20, 1976; from each unapproved part of that Amdt. 25–72, 55 FR 29786, July 20, 1990] manual.

(c) [Reserved] § 25.1561 Safety equipment.

(d) Each Airplane Flight Manual (a) Each safety equipment control to must include a table of contents if the be operated by the crew in emergency, complexity of the manual indicates a such as controls for automatic liferaft need for it.

releases, must be plainly marked as to [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978, as its method of operation.

amended by Amdt. 25–72, 55 FR 29786, July 20, (b) Each location, such as a locker or 1990] compartment, that carries any fire ex- tinguishing, signaling, or other life § 25.1583 Operating limitations.

saving equipment must be marked ac- (a) Airspeed limitations. The following cordingly.

airspeed limitations and any other air- (c) Stowage provisions for required speed limitations necessary for safe op- emergency equipment must be con- eration must be furnished: spicuously marked to identify the con- (1) The maximum operating limit tents and facilitate the easy removal of speed V / M and a statement that MO MO the equipment.

this speed limit may not be delib- (d) Each liferaft must have obviously erately exceeded in any regime of marked operating instructions. flight (climb, cruise, or descent) unless a higher speed is authorized for flight (e) Approved survival equipment test or pilot training.

must be marked for identification and (2) If an airspeed limitation is based method of operation.

upon compressibility effects, a state- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ment to this effect and information as amended by Amdt. 25–46, 43 FR 50598, Oct. 30, to any symptoms, the probable behav- 1978] ior of the airplane, and the rec- ommended recovery procedures.

§ 25.1563 Airspeed placard.

(3) The maneuvering speed estab- A placard showing the maximum air- lished under § 25.1507 and statements, speeds for flap extension for the take- as applicable to the particular design, off, approach, and landing positions explaining that: must be installed in clear view of each (i) Full application of pitch, roll, or pilot. yaw controls should be confined to Federal Aviation Administration, DOT § 25.1585 speeds below the maneuvering speed; termined under § 25.1523 must be fur- and nished.

(e) Kinds of operation. The kinds of (ii) Rapid and large alternating con- operation approved under § 25.1525 must trol inputs, especially in combination be furnished.

with large changes in pitch, roll, or (f) Ambient air temperatures and oper- yaw, and full control inputs in more ating altitudes. The extremes of the am- than one axis at the same time, should bient air temperatures and operating be avoided as they may result in struc- altitudes established under § 25.1527 tural failures at any speed, including must be furnished.

below the maneuvering speed.

(g) [Reserved] (4) The flap extended speed V and FE (h) Additional operating limitations.

the pertinent flap positions and engine The operating limitations established powers.

under § 25.1533 must be furnished.

(5) The landing gear operating speed (i) Maneuvering flight load factors. The or speeds, and a statement explaining positive maneuvering limit load fac- the speeds as defined in § 25.1515(a).

tors for which the structure is proven, (6) The landing gear extended speed described in terms of accelerations, V if greater than V and a state- LE, LO, must be furnished.

ment that this is the maximum speed at which the airplane can be safely [Doc. No. 5066, 29 FR 1891, Dec. 24, 1964, as flown with the landing gear extended.

amended by Amdt. 25–38, 41 FR 55468, Dec. 20, (b) Powerplant limitations. The fol- 1976; Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; Amdt.

lowing information must be furnished: 25–72, 55 FR 29787, July 20, 1990; Amdt. 25–105, (1) Limitations required by § 25.1521 66 FR 34024, June 26, 2001; 75 FR 49818, Aug.

and § 25.1522.

16, 2010] (2) Explanation of the limitations, when appropriate.

§ 25.1585 Operating procedures.

(3) Information necessary for mark- (a) Operating procedures must be fur- ing the instruments required by nished for— §§ 25.1549 through 25.1553.

(1) Normal procedures peculiar to the (c) Weight and loading distribution.

particular type or model encountered The weight and center of gravity limi- in connection with routine operations; tations established under § 25.1519 must (2) Non-normal procedures for mal- be furnished in the Airplane Flight function cases and failure conditions Manual. All of the following informa- involving the use of special systems or tion, including the weight distribution the alternative use of regular systems; limitations established under § 25.1519, and must be presented either in the Air- (3) Emergency procedures for foresee- plane Flight Manual or in a separate able but unusual situations in which weight and balance control and loading immediate and precise action by the document that is incorporated by ref- crew may be expected to substantially erence in the Airplane Flight Manual: reduce the risk of catastrophe.

(1) The condition of the airplane and (b) Information or procedures not di- the items included in the empty weight rectly related to airworthiness or not as defined in accordance with § 25.29.

under the control of the crew, must not (2) Loading instructions necessary to be included, nor must any procedure ensure loading of the airplane within that is accepted as basic airmanship.

the weight and center of gravity limits, (c) Information identifying each op- and to maintain the loading within erating condition in which the fuel sys- these limits in flight.

tem independence prescribed in § 25.953 (3) If certification for more than one is necessary for safety must be fur- center of gravity range is requested, nished, together with instructions for the appropriate limitations, with re- placing the fuel system in a configura- gard to weight and loading procedures, tion used to show compliance with that for each separate center of gravity section.

range.

(d) The buffet onset envelopes, deter- (d) Flight crew. The number and func- mined under § 25.251 must be furnished.

tions of the minimum flight crew de- The buffet onset envelopes presented 14 CFR Ch. I (1–1–25 Edition) § 25.1587 may reflect the center of gravity at (5) An explanation of significant or which the airplane is normally loaded unusual flight or ground handling char- during cruise if corrections for the ef- acteristics of the airplane.

fect of different center of gravity loca- (6) Corrections to indicated values of tions are furnished. airspeed, altitude, and outside air tem- (e) Information must be furnished perature.

that indicates that when the fuel quan- (7) An explanation of operational tity indicator reads ‘‘zero’’ in level landing runway length factors included flight, any fuel remaining in the fuel in the presentation of the landing dis- tank cannot be used safely in flight. tance, if appropriate.

(f) Information on the total quantity [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001, of usable fuel for each fuel tank must as amended by Amdt. 25–108, 67 FR 70828, be furnished.

Nov. 26, 2002] [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] Subpart H—Electrical Wiring § 25.1587 Performance information.

Interconnection Systems (EWIS) (a) Each Airplane Flight Manual must contain information to permit S OURCE : Docket No. FAA–2004–18379, 72 FR conversion of the indicated tempera- 63406, Nov. 8, 2007, unless otherwise noted.

ture to free air temperature if other § 25.1701 Definition.

than a free air temperature indicator is used to comply with the requirements (a) As used in this chapter, electrical of § 25.1303(a)(1).

wiring interconnection system (EWIS) (b) Each Airplane Flight Manual means any wire, wiring device, or com- must contain the performance informa- bination of these, including termi- tion computed under the applicable nation devices, installed in any area of provisions of this part (including the airplane for the purpose of trans- §§ 25.115, 25.123, and 25.125 for the mitting electrical energy, including weights, altitudes, temperatures, wind data and signals, between two or more components, and runway gradients, as intended termination points. This in- applicable) within the operational lim- cludes: its of the airplane, and must contain (1) Wires and cables.

the following: (2) Bus bars.

(1) In each case, the conditions of (3) The termination point on elec- power, configuration, and speeds, and trical devices, including those on re- the procedures for handling the air- lays, interrupters, switches, plane and any system having a signifi- contactors, terminal blocks and circuit cant effect on the performance infor- breakers, and other circuit protection mation.

devices.

determined in accordance (2) V SR (4) Connectors, including feed- with § 25.103.

through connectors.

(3) The following performance infor- (5) Connector accessories.

mation (determined by extrapolation (6) Electrical grounding and bonding and computed for the range of weights devices and their associated connec- between the maximum landing weight tions.

and the maximum takeoff weight): (7) Electrical splices.

(i) Climb in the landing configura- (8) Materials used to provide addi- tion.

tional protection for wires, including (ii) Climb in the approach configura- wire insulation, wire sleeving, and con- tion.

duits that have electrical termination (iii) Landing distance. for the purpose of bonding.

(4) Procedures established under (9) Shields or braids.

§ 25.101(f) and (g) that are related to the (10) Clamps and other devices used to limitations and information required route and support the wire bundle.

by § 25.1533 and by this paragraph (b) in (11) Cable tie devices.

the form of guidance material, includ- (12) Labels or other means of identi- ing any relevant limitations or infor- fication.

mation. (13) Pressure seals.

Federal Aviation Administration, DOT § 25.1707 (14) EWIS components inside shelves, § 25.1705 Systems and functions: EWIS.

panels, racks, junction boxes, distribu- (a) EWIS associated with any system tion panels, and back-planes of equip- required for type certification or by op- ment racks, including, but not limited erating rules must be considered an in- to, circuit board back-planes, wire in- tegral part of that system and must be tegration units, and external wiring of considered in showing compliance with equipment.

the applicable requirements for that (b) Except for the equipment indi- system.

cated in paragraph (a)(14) of this sec- (b) For systems to which the fol- tion, EWIS components inside the fol- lowing rules apply, the components of lowing equipment, and the external EWIS associated with those systems connectors that are part of that equip- must be considered an integral part of ment, are excluded from the definition that system or systems and must be in paragraph (a) of this section: considered in showing compliance with the applicable requirements for that (1) Electrical equipment or avionics system.

that are qualified to environmental (1) § 25.773(b)(2) Pilot compartment conditions and testing procedures when view.

those conditions and procedures are— (2) § 25.981 Fuel tank ignition pre- (i) Appropriate for the intended func- vention.

tion and operating environment, and (3) § 25.1165 Engine ignition systems.

(ii) Acceptable to the FAA.

(4) § 25.1310 Power source capacity (2) Portable electrical devices that and distribution.

are not part of the type design of the (5) § 25.1316 System lightning protec- airplane. This includes personal enter- tion.

tainment devices and laptop com- (6) § 25.1331(a)(2) Instruments using a puters.

power supply.

(3) Fiber optics.

(7) § 25.1351 General.

(8) § 25.1355 Distribution system.

§ 25.1703 Function and installation: (9) § 25.1360 Precautions against in- EWIS.

jury.

(10) § 25.1362 Electrical supplies for (a) Each EWIS component installed emergency conditions.

in any area of the aircraft must: (11) § 25.1365 Electrical appliances, (1) Be of a kind and design appro- motors, and transformers.

priate to its intended function.

(12) § 25.1431(c) and (d) Electronic (2) Be installed according to limita- equipment.

tions specified for the EWIS compo- nents.

§ 25.1707 System separation: EWIS.

(3) Perform the function for which it (a) Each EWIS must be designed and was intended without degrading the installed with adequate physical sepa- airworthiness of the airplane.

ration from other EWIS and airplane (4) Be designed and installed in a way systems so that an EWIS component that will minimize mechanical strain.

failure will not create a hazardous con- (b) Selection of wires must take into dition. Unless otherwise stated, for the account known characteristics of the purposes of this section, adequate wire in relation to each installation physical separation must be achieved and application to minimize the risk of by separation distance or by a barrier wire damage, including any arc track- that provides protection equivalent to ing phenomena.

that separation distance.

(c) The design and installation of the (b) Each EWIS must be designed and main power cables (including generator installed so that any electrical inter- cables) in the fuselage must allow for a ference likely to be present in the air- reasonable degree of deformation and plane will not result in hazardous ef- stretching without failure.

fects upon the airplane or its systems.

(d) EWIS components located in (c) Wires and cables carrying heavy areas of known moisture accumulation current, and their associated EWIS must be protected to minimize any components, must be designed and in- hazardous effects due to moisture. stalled to ensure adequate physical 14 CFR Ch. I (1–1–25 Edition) § 25.1709 separation and electrical isolation so water/waste lines and other water/ that damage to circuits associated waste system components, so that: with essential functions will be mini- (1) An EWIS component failure will mized under fault conditions. not create a hazardous condition.

(d) Each EWIS associated with inde- (2) Any water/waste leakage onto pendent airplane power sources or EWIS components will not create a power sources connected in combina- hazardous condition.

tion must be designed and installed to (i) EWIS must be designed and in- ensure adequate physical separation stalled with adequate physical separa- and electrical isolation so that a fault tion between the EWIS and flight or in any one airplane power source EWIS other mechanical control systems ca- will not adversely affect any other bles and associated system compo- independent power sources. In addition: nents, so that: (1) Airplane independent electrical (1) Chafing, jamming, or other inter- power sources must not share a com- ference are prevented.

mon ground terminating location.

(2) An EWIS component failure will (2) Airplane system static grounds not create a hazardous condition.

must not share a common ground ter- (3) Failure of any flight or other me- minating location with any of the air- chanical control systems cables or sys- plane’s independent electrical power tems components will not damage the sources.

EWIS and create a hazardous condi- (e) Except to the extent necessary to tion.

provide electrical connection to the (j) EWIS must be designed and in- fuel systems components, the EWIS stalled with adequate physical separa- must be designed and installed with tion between the EWIS components adequate physical separation from fuel and heated equipment, hot air ducts, lines and other fuel system compo- and lines, so that: nents, so that: (1) An EWIS component failure will (1) An EWIS component failure will not create a hazardous condition.

not create a hazardous condition.

(2) Any hot air leakage or heat gen- (2) Any fuel leakage onto EWIS com- erated onto EWIS components will not ponents will not create a hazardous create a hazardous condition.

condition.

(k) For systems for which redun- (f) Except to the extent necessary to dancy is required, by certification provide electrical connection to the rules, by operating rules, or as a result hydraulic systems components, EWIS of the assessment required by § 25.1709, must be designed and installed with EWIS components associated with adequate physical separation from hy- those systems must be designed and in- draulic lines and other hydraulic sys- stalled with adequate physical separa- tem components, so that: tion.

(1) An EWIS component failure will (l) Each EWIS must be designed and not create a hazardous condition.

installed so there is adequate physical (2) Any hydraulic fluid leakage onto separation between it and other air- EWIS components will not create a craft components and aircraft struc- hazardous condition.

ture, and so that the EWIS is protected (g) Except to the extent necessary to from sharp edges and corners, to mini- provide electrical connection to the ox- mize potential for abrasion/chafing, vi- ygen systems components, EWIS must bration damage, and other types of me- be designed and installed with ade- chanical damage.

quate physical separation from oxygen lines and other oxygen system compo- § 25.1709 System safety: EWIS.

nents, so that an EWIS component fail- Each EWIS must be designed and in- ure will not create a hazardous condi- stalled so that: tion.

(a) Each catastrophic failure condi- (h) Except to the extent necessary to tion— provide electrical connection to the (1) Is extremely improbable; and water/waste systems components, EWIS must be designed and installed (2) Does not result from a single fail- with adequate physical separation from ure.

Federal Aviation Administration, DOT § 25.1721 (b) Each hazardous failure condition guishing when tested in accordance is extremely remote. with the applicable portions of Appen- dix F, part I, of 14 CFR part 25.

§ 25.1711 Component identification: EWIS.

§ 25.1715 Electrical bonding and pro- tection against static electricity: (a) EWIS components must be labeled EWIS.

or otherwise identified using a con- sistent method that facilitates identi- (a) EWIS components used for elec- fication of the EWIS component, its trical bonding and protection against function, and its design limitations, if static electricity must meet the re- any. quirements of § 25.899.

(b) For systems for which redundancy (b) On airplanes having grounded is required, by certification rules, by electrical systems, electrical bonding operating rules, or as a result of the as- provided by EWIS components must sessment required by § 25.1709, EWIS provide an electrical return path capa- components associated with those sys- ble of carrying both normal and fault tems must be specifically identified currents without creating a shock haz- with component part number, function, ard or damage to the EWIS compo- and separation requirement for bun- nents, other airplane system compo- dles.

nents, or airplane structure.

(1) The identification must be placed along the wire, cable, or wire bundle at § 25.1717 Circuit protective devices: EWIS.

appropriate intervals and in areas of the airplane where it is readily visible Electrical wires and cables must be to maintenance, repair, or alteration designed and installed so they are com- personnel.

patible with the circuit protection de- (2) If an EWIS component cannot be vices required by § 25.1357, so that a fire marked physically, then other means or smoke hazard cannot be created of identification must be provided.

under temporary or continuous fault (c) The identifying markings re- conditions.

quired by paragraphs (a) and (b) of this section must remain legible through- § 25.1719 Accessibility provisions: out the expected service life of the EWIS.

EWIS component.

Access must be provided to allow in- (d) The means used for identifying spection and replacement of any EWIS each EWIS component as required by component as necessary for continued this section must not have an adverse airworthiness.

effect on the performance of that com- ponent throughout its expected service § 25.1721 Protection of EWIS.

life.

(a) No cargo or baggage compartment (e) Identification for EWIS modifica- may contain any EWIS whose damage tions to the type design must be con- or failure may affect safe operation, sistent with the identification scheme unless the EWIS is protected so that: of the original type design.

(1) It cannot be damaged by move- § 25.1713 Fire protection: EWIS.

ment of cargo or baggage in the com- partment.

(a) All EWIS components must meet (2) Its breakage or failure will not the applicable fire and smoke protec- create a fire hazard.

tion requirements of § 25.831(c) of this (b) EWIS must be designed and in- part.

stalled to minimize damage and risk of (b) EWIS components that are lo- damage to EWIS by movement of peo- cated in designated fire zones and are ple in the airplane during all phases of used during emergency procedures flight, maintenance, and servicing.

must be fire resistant.

(c) Insulation on electrical wire and (c) EWIS must be designed and in- electrical cable, and materials used to stalled to minimize damage and risk of provide additional protection for the damage to EWIS by items carried onto wire and cable, installed in any area of the aircraft by passengers or cabin the airplane, must be self-extin- crew.

14 CFR Ch. I (1–1–25 Edition) § 25.1723 § 25.1723 Flammable fluid fire protec- § 25.1731 Powerplant and APU fire de- tion: EWIS. tector system: EWIS.

EWIS components located in each (a) EWIS that are part of each fire or area where flammable fluid or vapors overheat detector system in a fire zone might escape by leakage of a fluid sys- must be fire-resistant.

tem must be considered a potential ig- (b) No EWIS component of any fire or nition source and must meet the re- overheat detector system for any fire quirements of § 25.863.

zone may pass through another fire zone, unless: § 25.1725 Powerplants: EWIS.

(1) It is protected against the possi- (a) EWIS associated with any power- bility of false warnings resulting from plant must be designed and installed so fires in zones through which it passes; that the failure of an EWIS component or will not prevent the continued safe op- (2) Each zone involved is simulta- eration of the remaining powerplants neously protected by the same detector or require immediate action by any crewmember for continued safe oper- and extinguishing system.

ation, in accordance with the require- (c) EWIS that are part of each fire or ments of § 25.903(b).

overheat detector system in a fire zone (b) Design precautions must be taken must meet the requirements of to minimize hazards to the airplane § 25.1203.

due to EWIS damage in the event of a powerplant rotor failure or a fire origi- § 25.1733 Fire detector systems, gen- nating within the powerplant that eral: EWIS.

burns through the powerplant case, in EWIS associated with any installed accordance with the requirements of fire protection system, including those § 25.903(d)(1).

required by §§ 25.854 and 25.858, must be considered an integral part of the sys- § 25.1727 Flammable fluid shutoff tem in showing compliance with the means: EWIS.

applicable requirements for that sys- EWIS associated with each flam- tem.

mable fluid shutoff means and control must be fireproof or must be located and protected so that any fire in a fire Subpart I—Special Federal zone will not affect operation of the Aviation Regulations flammable fluid shutoff means, in ac- cordance with the requirements of S OURCE : Docket No. FAA–2011–0186, Amdt.

§ 25.1189.

25–133, 76 FR 12555, Mar. 8, 2011, unless other- wise noted.

§ 25.1729 Instructions for Continued Airworthiness: EWIS.

§ 25.1801 SFAR No. 111—Lavatory Oxy- The applicant must prepare Instruc- gen Systems.

tions for Continued Airworthiness ap- The requirements of § 121.1500 of this plicable to EWIS in accordance with chapter also apply to this part.

Appendix H sections H25.4 and H25.5 to this part that are approved by the FAA.

Federal Aviation Administration, DOT Pt. 25, App. A A PPENDIX A TO P ART 25 14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. A Federal Aviation Administration, DOT Pt. 25, App. A 14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. A Federal Aviation Administration, DOT Pt. 25, App. A 14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. B A PPENDIX B TO P ART 25 Federal Aviation Administration, DOT Pt. 25, App. B

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. C

by the appropriate factor from figure 3 of A PPENDIX C TO P ART 25 this appendix.

Part I—Atmospheric Icing Conditions (b) Intermittent maximum icing. The inter- mittent maximum intensity of atmospheric (a) Continuous maximum icing. The max- icing conditions (intermittent maximum imum continuous intensity of atmospheric icing) is defined by the variables of the cloud icing conditions (continuous maximum liquid water content, the mean effective di- icing) is defined by the variables of the cloud ameter of the cloud droplets, the ambient air liquid water content, the mean effective di- temperature, and the interrelationship of ameter of the cloud droplets, the ambient air these three variables as shown in figure 4 of temperature, and the interrelationship of this appendix. The limiting icing envelope in these three variables as shown in figure 1 of terms of altitude and temperature is given in this appendix. The limiting icing envelope in figure 5 of this appendix. The inter-relation- terms of altitude and temperature is given in ship of cloud liquid water content with drop figure 2 of this appendix. The inter-relation- diameter and altitude is determined from ship of cloud liquid water content with drop figures 4 and 5. The cloud liquid water con- diameter and altitude is determined from tent for intermittent maximum icing condi- figures 1 and 2. The cloud liquid water con- tions of a horizontal extent, other than 2.6 tent for continuous maximum icing condi- nautical miles, is determined by the value of tions of a horizontal extent, other than 17.4 cloud liquid water content of figure 4 multi- nautical miles, is determined by the value of plied by the appropriate factor in figure 6 of liquid water content of figure 1, multiplied this appendix.

Federal Aviation Administration, DOT Pt. 25, App. C

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. C

Federal Aviation Administration, DOT Pt. 25, App. C

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. C

Federal Aviation Administration, DOT Pt. 25, App. C

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. C

(c) Takeoff maximum icing. The maximum temperature at ground level of minus 9 de- intensity of atmospheric icing conditions for grees Celsius ( ¥ 9 ° C). The takeoff maximum takeoff (takeoff maximum icing) is defined icing conditions extend from ground level to by the cloud liquid water content of 0.35 g/ a height of 1,500 feet above the level of the m3, the mean effective diameter of the cloud takeoff surface.

droplets of 20 microns, and the ambient air

Federal Aviation Administration, DOT Pt. 25, App. D

Part II—Airframe Ice Accretions for Showing (b) In order to reduce the number of ice ac- Compliance With Subpart B. cretions to be considered when dem- onstrating compliance with the require- (a) Ice accretions—General. The most crit- ments of § 25.21(g), any of the ice accretions ical ice accretion in terms of airplane per- defined in paragraph (a) of this section may formance and handling qualities for each be used for any other flight phase if it is flight phase must be used to show compli- shown to be more critical than the specific ance with the applicable airplane perform- ice accretion defined for that flight phase.

ance and handling requirements in icing con- Configuration differences and their effects ditions of subpart B of this part. Applicants on ice accretions must be taken into ac- must demonstrate that the full range of at- count.

mospheric icing conditions specified in part I (c) The ice accretion that has the most ad- of this appendix have been considered, in- verse effect on handling qualities may be cluding the mean effective drop diameter, used for airplane performance tests provided liquid water content, and temperature appro- any difference in performance is conserv- priate to the flight conditions (for example, atively taken into account.

configuration, speed, angle-of-attack, and al- (d) For both unprotected and protected titude). The ice accretions for each flight parts, the ice accretion for the takeoff phase phase are defined as follows: may be determined by calculation, assuming (1) Takeoff ice is the most critical ice accre- the takeoff maximum icing conditions de- tion on unprotected surfaces and any ice ac- fined in appendix C, and assuming that: cretion on the protected surfaces appropriate (1) Airfoils, control surfaces and, if appli- to normal ice protection system operation, cable, propellers are free from frost, snow, or occurring between the end of the takeoff dis- ice at the start of the takeoff; tance and 400 feet above the takeoff surface, (2) The ice accretion starts at the end of assuming accretion starts at the end of the the takeoff distance.

takeoff distance in the takeoff maximum (3) The critical ratio of thrust/power-to- icing conditions defined in part I of this Ap- weight; pendix.

(4) Failure of the critical engine occurs at (2) Final takeoff ice is the most critical ice V EF ; and accretion on unprotected surfaces, and any (5) Crew activation of the ice protection ice accretion on the protected surfaces ap- system is in accordance with a normal oper- propriate to normal ice protection system ating procedure provided in the Airplane operation, between 400 feet and either 1,500 Flight Manual, except that after beginning feet above the takeoff surface, or the height the takeoff roll, it must be assumed that the at which the transition from the takeoff to crew takes no action to activate the ice pro- the en route configuration is completed and tection system until the airplane is at least V is reached, whichever is higher. Ice ac- FTO 400 feet above the takeoff surface.

cretion is assumed to start at the end of the (e) The ice accretion before the ice protec- takeoff distance in the takeoff maximum tion system has been activated and is per- icing conditions of part I, paragraph (c) of forming its intended function is the critical this Appendix.

ice accretion formed on the unprotected and (3) En route ice is the critical ice accretion normally protected surfaces before activa- on the unprotected surfaces, and any ice ac- tion and effective operation of the ice pro- cretion on the protected surfaces appropriate tection system in continuous maximum at- to normal ice protection system operation, mospheric icing conditions. This ice accre- during the en route phase. tion only applies in showing compliance to (4) Holding ice is the critical ice accretion §§ 25.143(j) and 25.207(h), and 25.207(i).

on the unprotected surfaces, and any ice ac- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as cretion on the protected surfaces appropriate amended by Amdt. 25–121, 72 FR 44669, Aug. 8, to normal ice protection system operation, 2007; 72 FR 50467, Aug. 31, 2007; Amdt. 25–129, during the holding flight phase.

74 FR 38340, Aug. 3, 2009; Amdt. 25–140, 79 FR (5) Approach ice is the critical ice accretion 65528, Nov. 4, 2014] on the unprotected surfaces, and any ice ac- cretion on the protected surfaces appropriate A PPENDIX D TO P ART 25 to normal ice protection system operation following exit from the holding flight phase Criteria for determining minimum flight crew.

and transition to the most critical approach The following are considered by the Agency configuration.

in determining the minimum flight crew (6) Landing ice is the critical ice accretion under § 25.1523: on the unprotected surfaces, and any ice ac- (a) Basic workload functions. The following cretion on the protected surfaces appropriate basic workload functions are considered: to normal ice protection system operation (1) Flight path control.

following exit from the approach flight phase (2) Collision avoidance.

and transition to the final landing configura- (3) Navigation.

tion. (4) Communications.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. E

(5) Operation and monitoring of aircraft that each airplane certificated under this engines and systems. Part will operate under IFR conditions.

(6) Command decisions.

[Amdt. 25–3, 30 FR 6067, Apr. 29, 1965] (b) Workload factors. The following work- load factors are considered significant when A PPENDIX E TO P ART 25 analyzing and demonstrating workload for I—Limited Weight Credit For Airplanes minimum flight crew determination: Equipped With Standby Power (1) The accessibility, ease, and simplicity of operation of all necessary flight, power, (a) Each applicant for an increase in the and equipment controls, including emer- maximum certificated takeoff and landing gency fuel shutoff valves, electrical controls, weights of an airplane equipped with a type- electronic controls, pressurization system certificated standby power rocket engine controls, and engine controls.

may obtain an increase as specified in para- (2) The accessibility and conspicuity of all graph (b) if— necessary instruments and failure warning (1) The installation of the rocket engine devices such as fire warning, electrical sys- has been approved and it has been estab- tem malfunction, and other failure or cau- lished by flight test that the rocket engine tion indicators. The extent to which such in- and its controls can be operated safely and struments or devices direct the proper cor- reliably at the increase in maximum weight; rective action is also considered. and (2) The Airplane Flight Manual, or the (3) The number, urgency, and complexity placard, markings or manuals required in of operating procedures with particular con- place thereof, set forth in addition to any sideration given to the specific fuel manage- other operating limitations the Adminis- ment schedule imposed by center of gravity, trator may require, the increased weight ap- structural or other considerations of an air- proved under this regulation and a prohibi- worthiness nature, and to the ability of each tion against the operation of the airplane at engine to operate at all times from a single the approved increased weight when— tank or source which is automatically re- (i) The installed standby power rocket en- plenished if fuel is also stored in other tanks.

gines have been stored or installed in excess (4) The degree and duration of con- of the time limit established by the manu- centrated mental and physical effort in- facturer of the rocket engine (usually sten- volved in normal operation and in diagnosing ciled on the engine casing); or and coping with malfunctions and emer- (ii) The rocket engine fuel has been ex- gencies.

pended or discharged.

(5) The extent of required monitoring of (b) The currently approved maximum take- the fuel, hydraulic, pressurization, elec- off and landing weights at which an airplane trical, electronic, deicing, and other systems is certificated without a standby power rock- while en route.

et engine installation may be increased by (6) The actions requiring a crewmember to an amount that does not exceed any of the be unavailable at his assigned duty station, following: including: observation of systems, emer- (1) An amount equal in pounds to 0.014 IN, gency operation of any control, and emer- where I is the maximum usable impulse in gencies in any compartment.

pounds-seconds available from each standby (7) The degree of automation provided in power rocket engine and N is the number of the aircraft systems to afford (after failures rocket engines installed.

or malfunctions) automatic crossover or iso- (2) An amount equal to 5 percent of the lation of difficulties to minimize the need for maximum certificated weight approved in flight crew action to guard against loss of accordance with the applicable airworthiness hydraulic or electric power to flight controls regulations without standby power rocket or to other essential systems.

engines installed.

(8) The communications and navigation (3) An amount equal to the weight of the workload.

rocket engine installation.

(9) The possibility of increased workload (4) An amount that, together with the cur- associated with any emergency that may rently approved maximum weight, would lead to other emergencies. equal the maximum structural weight estab- lished for the airplane without standby rock- (10) Incapacitation of a flight crewmember et engines installed.

whenever the applicable operating rule re- quires a minimum flight crew of at least two II—Performance Credit for Transport Category pilots.

Airplanes Equipped With Standby Power (c) Kind of operation authorized. The deter- mination of the kind of operation authorized The Administrator may grant performance requires consideration of the operating rules credit for the use of standby power on trans- under which the airplane will be operated. port category airplanes. However, the per- Unless an applicant desires approval for a formance credit applies only to the max- more limited kind of operation. It is assumed imum certificated takeoff and landing

Federal Aviation Administration, DOT Pt. 25, App. E

weights, the takeoff distance, and the take- from the start of the takeoff to the point off paths, and may not exceed that found by where the airplane attains a height of 50 feet the Administrator to result in an overall above the takeoff surface for reciprocating- level of safety in the takeoff, approach, and engine-powered airplanes and a height of 35 landing regimes of flight equivalent to that feet above the takeoff surface for turbine- prescribed in the regulations under which powered airplanes.

the airplane was originally certificated with- (4) Maximum certificated takeoff weights. The out standby power. For the purposes of this maximum certificated takeoff weights must appendix, ‘‘standby power’’ is power or be determined at all altitudes, and at ambi- thrust, or both, obtained from rocket en- ent temperatures, if applicable, at which per- gines for a relatively short period and actu- formance credit is to be applied and may not ated only in cases of emergency. The fol- exceed the weights established in compliance lowing provisions apply: with paragraphs (a) and (b) of this section.

(1) Takeoff; general. The takeoff data pre- (a) The conditions of paragraphs (2)(b) through (d) must be met at the maximum scribed in paragraphs (2) and (3) of this ap- pendix must be determined at all weights certificated takeoff weight.

and altitudes, and at ambient temperatures (b) Without the use of standby power, the if applicable, at which performance credit is airplane must meet all of the en route re- to be applied. quirements of the applicable airworthiness (2) Takeoff path. regulations under which the airplane was (a) The one-engine-inoperative takeoff originally certificated. In addition, turbine- path with standby power in use must be de- powered airplanes without the use of standby termined in accordance with the perform- power must meet the final takeoff climb re- ance requirements of the applicable air- quirements prescribed in the applicable air- worthiness regulations. worthiness regulations.

(b) The one-engine-inoperative takeoff (5) Maximum certificated landing weights.

path (excluding that part where the airplane (a) The maximum certificated landing is on or just above the takeoff surface) deter- weights (one-engine-inoperative approach mined in accordance with paragraph (a) of and all-engine-operating landing climb) must this section must lie above the one-engine- be determined at all altitudes, and at ambi- inoperative takeoff path without standby ent temperatures if applicable, at which per- formance credit is to be applied and must power at the maximum takeoff weight at which all of the applicable air-worthiness re- not exceed that established in compliance quirements are met. For the purpose of this with paragraph (b) of this section.

comparison, the flight path is considered to (b) The flight path, with the engines oper- extend to at least a height of 400 feet above ating at the power or thrust, or both, appro- the takeoff surface. priate to the airplane configuration and with (c) The takeoff path with all engines oper- standby power in use, must lie above the ating, but without the use of standby power, flight path without standby power in use at must reflect a conservatively greater overall the maximum weight at which all of the ap- level of performance than the one-engine-in- plicable airworthiness requirements are met.

operative takeoff path established in accord- In addition, the flight paths must comply ance with paragraph (a) of this section. The with subparagraphs (i) and (ii) of this para- margin must be established by the Adminis- graph.

trator to insure safe day-to-day operations, (i) The flight paths must be established but in no case may it be less than 15 percent.

without changing the appropriate airplane The all-engines-operating takeoff path must configuration.

be determined by a procedure consistent (ii) The flight paths must be carried out for with that established in complying with a minimum height of 400 feet above the point paragraph (a) of this section.

where standby power is actuated.

(d) For reciprocating-engine-powered air- (6) Airplane configuration, speed, and power planes, the takeoff path to be scheduled in and thrust; general. Any change in the air- the Airplane Flight Manual must represent plane’s configuration, speed, and power or the one-engine-operative takeoff path deter- thrust, or both, must be made in accordance mined in accordance with paragraph (a) of with the procedures established by the appli- this section and modified to reflect the pro- cant for the operation of the airplane in cedure (see paragraph (6)) established by the service and must comply with paragraphs (a) applicant for flap retraction and attainment through (c) of this section. In addition, pro- of the en route speed. The scheduled takeoff cedures must be established for the execu- path must have a positive slope at all points tion of balked landings and missed ap- of the airborne portion and at no point must proaches.

it lie above the takeoff path specified in (a) The Administrator must find that the paragraph (a) of this section. procedure can be consistently executed in (3) Takeoff distance. The takeoff distance service by crews of average skill.

must be the horizontal distance along the (b) The procedure may not involve methods one-engine-inoperative take off path deter- or the use of devices which have not been mined in accordance with paragraph (2)(a) proven to be safe and reliable.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(c) Allowances must be made for such time the film travels through ducts, the ducts delays in the execution of the procedures as must meet the requirements of subparagraph may be reasonably expected to occur during (ii) of this paragraph.

service. (iv) Clear plastic windows and signs, parts (7) Installation and operation; standby power. constructed in whole or in part of elas- The standby power unit and its installation tomeric materials, edge lighted instrument must comply with paragraphs (a) and (b) of assemblies consisting of two or more instru- this section. ments in a common housing, seat belts, (a) The standby power unit and its instal- shoulder harnesses, and cargo and baggage lation must not adversely affect the safety of tiedown equipment, including containers, the airplane. bins, pallets, etc., used in passenger or crew (b) The operation of the standby power compartments, may not have an average unit and its control must have proven to be burn rate greater than 2.5 inches per minute safe and reliable.

when tested horizontally in accordance with the applicable portions of this appendix.

[Amdt. 25–6, 30 FR 8468, July 2, 1965] (v) Except for small parts (such as knobs, handles, rollers, fasteners, clips, grommets, A PPENDIX F TO P ART 25 rub strips, pulleys, and small electrical parts) that would not contribute signifi- Part I—Test Criteria and Procedures for cantly to the propagation of a fire and for Showing Compliance With § 25.853 or § 25.855 electrical wire and cable insulation, mate- (a) Material test criteria —(1) Interior com- rials in items not specified in paragraphs partments occupied by crew or passengers. (i) (a)(1)(i), (ii), (iii), or (iv) of part I of this ap- Interior ceiling panels, interior wall panels, pendix may not have a burn rate greater partitions, galley structure, large cabinet than 4.0 inches per minute when tested hori- walls, structural flooring, and materials used zontally in accordance with the applicable in the construction of stowage compart- portions of this appendix.

ments (other than underseat stowage com- (2) Cargo and baggage compartments not oc- partments and compartments for stowing cupied by crew or passengers.

small items such as magazines and maps) (i) [Reserved] must be self-extinguishing when tested (ii) A cargo or baggage compartment de- vertically in accordance with the applicable fined in § 25.857 as Class B or E must have a portions of part I of this appendix. The aver- liner constructed of materials that meet the age burn length may not exceed 6 inches and requirements of paragraph (a)(1)(ii) of part I the average flame time after removal of the of this appendix and separated from the air- flame source may not exceed 15 seconds.

plane structure (except for attachments). In Drippings from the test specimen may not addition, such liners must be subjected to continue to flame for more than an average the 45 degree angle test. The flame may not of 3 seconds after falling.

penetrate (pass through) the material during (ii) Floor covering, textiles (including application of the flame or subsequent to its draperies and upholstery), seat cushions, removal. The average flame time after re- padding, decorative and non-decorative coat- moval of the flame source may not exceed 15 ed fabrics, leather, trays and galley fur- seconds, and the average glow time may not nishings, electrical conduit, air ducting, exceed 10 seconds.

joint and edge covering, liners of Class B and (iii) A cargo or baggage compartment de- E cargo or baggage compartments, floor pan- fined in § 25.857 as Class B, C, E, or F must els of Class B, C, E, or F cargo or baggage have floor panels constructed of materials compartments, cargo covers and trans- which meet the requirements of paragraph parencies, molded and thermoformed parts, (a)(1)(ii) of part I of this appendix and which air ducting joints, and trim strips (decora- are separated from the airplane structure tive and chafing), that are constructed of (except for attachments). Such panels must materials not covered in paragraph (a)(1)(iv) be subjected to the 45 degree angle test. The below, must be self-extinguishing when test- flame may not penetrate (pass through) the ed vertically in accordance with the applica- material during application of the flame or ble portions of part I of this appendix or subsequent to its removal. The average other approved equivalent means. The aver- flame time after removal of the flame source age burn length may not exceed 8 inches, and may not exceed 15 seconds, and the average the average flame time after removal of the glow time may not exceed 10 seconds.

flame source may not exceed 15 seconds. (iv) Insulation blankets and covers used to Drippings from the test specimen may not protect cargo must be constructed of mate- continue to flame for more than an average rials that meet the requirements of para- of 5 seconds after falling. graph (a)(1)(ii) of part I of this appendix. Tie- (iii) Motion picture film must be safety down equipment (including containers, bins, film meeting the Standard Specifications for and pallets) used in each cargo and baggage Safety Photographic Film PHI.25 (available compartment must be constructed of mate- from the American National Standards Insti- rials that meet the requirements of para- tute, 1430 Broadway, New York, NY 10018). If graph (a)(1)(v) of part I of this appendix.

Federal Aviation Administration, DOT Pt. 25, App. F

(3) Electrical system components. Insulation cordance with Federal Test Method Standard on electrical wire or cable installed in any 191 Model 5903 (revised Method 5902) for the area of the fuselage must be self-extin- vertical test, or Method 5906 for horizontal guishing when subjected to the 60 degree test test (available from the General Services Ad- specified in part I of this appendix. The aver- ministration, Business Service Center, Re- age burn length may not exceed 3 inches, and gion 3, Seventh & D Streets SW., Wash- the average flame time after removal of the ington, DC 20407). Specimens which are too flame source may not exceed 30 seconds.

large for the cabinet must be tested in simi- Drippings from the test specimen may not lar draft-free conditions.

continue to flame for more than an average (4) Vertical test. A minimum of three speci- of 3 seconds after falling.

mens must be tested and results averaged.

(b) Test Procedures —(1) Conditioning. Speci- For fabrics, the direction of weave cor- mens must be conditioned to 70 ± 5 F., and at responding to the most critical flammability 50 percent ± 5 percent relative humidity until conditions must be parallel to the longest di- moisture equilibrium is reached or for 24 mension. Each specimen must be supported hours. Each specimen must remain in the vertically. The specimen must be exposed to conditioning environment until it is sub- a Bunsen or Tirrill burner with a nominal ⁄ 8- jected to the flame.

inch I.D. tube adjusted to give a flame of 1 ⁄2 (2) Specimen configuration. Except for small inches in height. The minimum flame tem- parts and electrical wire and cable insula- perature measured by a calibrated thermo- tion, materials must be tested either as sec- couple pyrometer in the center of the flame tion cut from a fabricated part as installed must be 1550 ° F. The lower edge of the speci- in the airplane or as a specimen simulating men must be ⁄ 4 -inch above the top edge of a cut section, such as a specimen cut from a the burner. The flame must be applied to the flat sheet of the material or a model of the center line of the lower edge of the specimen.

fabricated part. The specimen may be cut For materials covered by paragraph (a)(1)(i) from any location in a fabricated part; how- of part I of this appendix, the flame must be ever, fabricated units, such as sandwich pan- applied for 60 seconds and then removed. For els, may not be separated for test. Except as materials covered by paragraph (a)(1)(ii) of noted below, the specimen thickness must be part I of this appendix, the flame must be ap- no thicker than the minimum thickness to plied for 12 seconds and then removed. Flame be qualified for use in the airplane. Test time, burn length, and flaming time of drip- specimens of thick foam parts, such as seat pings, if any, may be recorded. The burn ⁄ 2 -inch in thickness. Test cushions, must be length determined in accordance with sub- specimens of materials that must meet the paragraph (7) of this paragraph must be requirements of paragraph (a)(1)(v) of part I measured to the nearest tenth of an inch.

of this appendix must be no more than ⁄8 - (5) Horizontal test. A minimum of three inch in thickness. Electrical wire and cable specimens must be tested and the results specimens must be the same size as used in averaged. Each specimen must be supported the airplane. In the case of fabrics, both the horizontally. The exposed surface, when in- warp and fill direction of the weave must be stalled in the aircraft, must be face down for tested to determine the most critical flam- the test. The specimen must be exposed to a mability condition. Specimens must be Bunsen or Tirrill burner with a nominal ⁄8- mounted in a metal frame so that the two inch I.D. tube adjusted to give a flame of 1 ⁄2 long edges and the upper edge are held se- inches in height. The minimum flame tem- curely during the vertical test prescribed in perature measured by a calibrated thermo- subparagraph (4) of this paragraph and the couple pyrometer in the center of the flame two long edges and the edge away from the must be 1550 ° F. The specimen must be posi- flame are held securely during the horizontal tioned so that the edge being tested is cen- test prescribed in subparagraph (5) of this tered ⁄4-inch above the top of the burner.

paragraph. The exposed area of the specimen The flame must be applied for 15 seconds and must be at least 2 inches wide and 12 inches then removed. A minimum of 10 inches of long, unless the actual size used in the air- specimen must be used for timing purposes, plane is smaller. The edge to which the burn- approximately 1 ⁄ 2 inches must burn before er flame is applied must not consist of the the burning front reaches the timing zone, finished or protected edge of the specimen and the average burn rate must be recorded.

but must be representative of the actual cross-section of the material or part as in- (6) Forty-five degree test. A minimum of stalled in the airplane. The specimen must three specimens must be tested and the re- be mounted in a metal frame so that all four sults averaged. The specimens must be sup- edges are held securely and the exposed area ported at an angle of 45 ° to a horizontal sur- of the specimen is at least 8 inches by 8 face. The exposed surface when installed in inches during the 45 ° test prescribed in sub- the aircraft must be face down for the test.

paragraph (6) of this paragraph. The specimens must be exposed to a Bunsen (3) Apparatus. Except as provided in sub- or Tirrill burner with a nominal ⁄8-inch I.D.

paragraph (7) of this paragraph, tests must tube adjusted to give a flame of 1 ⁄2 inches in be conducted in a draft-free cabinet in ac- height. The minimum flame temperature

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

measured by a calibrated thermocouple py- Part II—Flammability of Seat Cushions rometer in the center of the flame must be (a) Criteria for Acceptance. Each seat cush- 1550 ° F. Suitable precautions must be taken ion must meet the following criteria: to avoid drafts. The flame must be applied (1) At least three sets of seat bottom and for 30 seconds with one-third contacting the seat back cushion specimens must be tested.

material at the center of the specimen and (2) If the cushion is constructed with a fire then removed. Flame time, glow time, and blocking material, the fire blocking material whether the flame penetrates (passes must completely enclose the cushion foam through) the specimen must be recorded.

core material.

(7) Sixty degree test. A minimum of three (3) Each specimen tested must be fab- specimens of each wire specification (make ricated using the principal components (i.e., and size) must be tested. The specimen of foam core, flotation material, fire blocking wire or cable (including insulation) must be material, if used, and dress covering) and as- placed at an angle of 60 ° with the horizontal sembly processes (representative seams and in the cabinet specified in subparagraph (3) closures) intended for use in the production of this paragraph with the cabinet door open articles. If a different material combination during the test, or must be placed within a is used for the back cushion than for the bot- chamber approximately 2 feet high by 1 foot tom cushion, both material combinations by 1 foot, open at the top and at one vertical must be tested as complete specimen sets, side (front), and which allows sufficient flow each set consisting of a back cushion speci- of air for complete combustion, but which is men and a bottom cushion specimen. If a free from drafts. The specimen must be par- cushion, including outer dress covering, is allel to and approximately 6 inches from the demonstrated to meet the requirements of front of the chamber. The lower end of the this appendix using the oil burner test, the specimen must be held rigidly clamped. The dress covering of that cushion may be re- upper end of the specimen must pass over a placed with a similar dress covering provided pulley or rod and must have an appropriate the burn length of the replacement covering, weight attached to it so that the specimen is as determined by the test specified in held tautly throughout the flammability § 25.853(c), does not exceed the corresponding test. The test specimen span between lower burn length of the dress covering used on the clamp and upper pulley or rod must be 24 cushion subjected to the oil burner test.

inches and must be marked 8 inches from the (4) For at least two-thirds of the total lower end to indicate the central point for number of specimen sets tested, the burn flame application. A flame from a Bunsen or length from the burner must not reach the Tirrill burner must be applied for 30 seconds side of the cushion opposite the burner. The at the test mark. The burner must be mount- burn length must not exceed 17 inches. Burn ed underneath the test mark on the speci- length is the perpendicular distance from the men, perpendicular to the specimen and at inside edge of the seat frame closest to the an angle of 30 ° to the vertical plane of the burner to the farthest evidence of damage to specimen. The burner must have a nominal the test specimen due to flame impingement, bore of ⁄8-inch and be adjusted to provide a including areas of partial or complete con- 3-inch high flame with an inner cone ap- sumption, charring, or embrittlement, but proximately one-third of the flame height.

not including areas sooted, stained, warped, The minimum temperature of the hottest portion of the flame, as measured with a or discolored, or areas where material has calibrated thermocouple pyrometer, may not shrunk or melted away from the heat source.

be less than 1750 ° F. The burner must be posi- (5) The average percentage weight loss tioned so that the hottest portion of the must not exceed 10 percent. Also, at least flame is applied to the test mark on the two-thirds of the total number of specimen wire. Flame time, burn length, and flaming sets tested must not exceed 10 percent time of drippings, if any, must be recorded. weight loss. All droppings falling from the The burn length determined in accordance cushions and mounting stand are to be dis- with paragraph (8) of this paragraph must be carded before the after-test weight is deter- measured to the nearest tenth of an inch. mined. The percentage weight loss for a spec- Breaking of the wire specimens is not consid- imen set is the weight of the specimen set ered a failure. before testing less the weight of the speci- men set after testing expressed as the per- (8) Burn length. Burn length is the distance from the original edge to the farthest evi- centage of the weight before testing.

dence of damage to the test specimen due to (b) Test Conditions. Vertical air velocity flame impingement, including areas of par- should average 25 fpm ± 10 fpm at the top of tial or complete consumption, charring, or the back seat cushion. Horizontal air veloc- embrittlement, but not including areas soot- ity should be below 10 fpm just above the ed, stained, warped, or discolored, nor areas bottom seat cushion. Air velocities should be where material has shrunk or melted away measured with the ventilation hood oper- from the heat source. ating and the burner motor off.

Federal Aviation Administration, DOT Pt. 25, App. F

(c) Test Specimens. (1) For each test, one set (ii) Because crumbling of the insulating of cushion specimens representing a seat bot- board with service can result in misalign- tom and seat back cushion must be used. ment of the calorimeter, the calorimeter (2) The seat bottom cushion specimen must must be monitored and the mounting 1 1 ⁄8 inches (457 ± 3 mm) wide by 20 ± ⁄8 be 18 ± shimmed, as necessary, to ensure that the inches (508 ± 3 mm) deep by 4 ± ⁄ 8 inches (102 calorimeter face is flush with the exposed ± 3 mm) thick, exclusive of fabric closures plane of the insulating board in a plane par- and seam overlap.

allel to the exit of the test burner cone.

(3) The seat back cushion specimen must (4) Thermocouples. The seven thermocouples 1 1 1 1 be 18 ± ⁄8 inches (432 ± 3 mm) wide by 25 ± ⁄8 to be used for testing must be ⁄ 16 - to ⁄8-inch inches (635 ± 3 mm) high by 2 ± ⁄ 8 inches (51 ± 3 metal sheathed, ceramic packed, type K, mm) thick, exclusive of fabric closures and grounded thermocouples with a nominal 22 seam overlap.

to 30 American wire gage (AWG)-size con- (4) The specimens must be conditioned at ductor. The seven thermocouples must be at- 70 ± 5 ° F (21 ± 2 ° C) 55% ± 10% relative humidity tached to a steel angle bracket to form a for at least 24 hours before testing.

thermocouple rake for placement in the test (d) Test Apparatus. The arrangement of the stand during burner calibration, as shown in test apparatus is shown in Figures 1 through Figure 5.

5 and must include the components described (5) Apparatus Arrangement. The test burner in this section. Minor details of the appa- must be mounted on a suitable stand to posi- ratus may vary, depending on the model tion the exit of the burner cone a distance of burner used.

4 ± ⁄8 inches (102 ± 3 mm) from one side of the (1) Specimen Mounting Stand. The mounting specimen mounting stand. The burner stand stand for the test specimens consists of steel should have the capability of allowing the angles, as shown in Figure 1. The length of burner to be swung away from the specimen the mounting stand legs is 12 ± ⁄8 inches (305 mounting stand during warmup periods.

± 3 mm). The mounting stand must be used (6) Data Recording. A recording potentiom- for mounting the test specimen seat bottom eter or other suitable calibrated instrument and seat back, as shown in Figure 2. The with an appropriate range must be used to mounting stand should also include a suit- measure and record the outputs of the calo- able drip pan lined with aluminum foil, dull rimeter and the thermocouples.

side up.

(7) Weight Scale. Weighing Device—A device (2) Test Burner. The burner to be used in must be used that with proper procedures testing must— may determine the before and after test (i) Be a modified gun type; weights of each set of seat cushion specimens (ii) Have an 80-degree spray angle nozzle within 0.02 pound (9 grams). A continuous nominally rated for 2.25 gallons/hour at 100 weighing system is preferred.

psi; (8) Timing Device. A stopwatch or other de- (iii) Have a 12-inch (305 mm) burner cone vice (calibrated to ± 1 second) must be used to installed at the end of the draft tube, with measure the time of application of the burn- an opening 6 inches (152 mm) high and 11 er flame and self-extinguishing time or test inches (280 mm) wide, as shown in Figure 3; duration.

and (e) Preparation of Apparatus. Before calibra- (iv) Have a burner fuel pressure regulator tion, all equipment must be turned on and that is adjusted to deliver a nominal 2.0 gal- the burner fuel must be adjusted as specified lon/hour of # 2 Grade kerosene or equivalent required for the test. in paragraph (d)(2).

(f) Calibration. To ensure the proper ther- Burner models which have been used success- mal output of the burner, the following test fully in testing are the Lennox Model OB–32, must be made: Carlin Model 200 CRD, and Park Model DPL (1) Place the calorimeter on the test stand 3400. FAA published reports pertinent to this as shown in Figure 4 at a distance of 4 ± ⁄8 type of burner are: (1) Powerplant inches (102 ± 3 mm) from the exit of the burn- Enginering Report No. 3A, Standard Fire er cone.

Test Apparatus and Procedure for Flexible (2) Turn on the burner, allow it to run for Hose Assemblies, dated March 1978; and (2) Report No. DOT/FAA/RD/76/213, Reevaluation 2 minutes for warmup, and adjust the burner of Burner Characteristics for Fire Resistance air intake damper to produce a reading of 2 2 Tests, dated January 1977. 10.5 ± 0.5 BTU/ft -sec. (11.9 ± 0.6 w/cm ) on the (3) Calorimeter. calorimeter to ensure steady state condi- (i) The calorimeter to be used in testing tions have been achieved. Turn off the burn- 2 2 must be a (0–15.0 BTU/ft -sec. 0–17.0 W/cm ) er.

calorimeter, accurate ± 3%, mounted in a 6- (3) Replace the calorimeter with the ther- inch by 12-inch (152 by 305 mm) by ⁄4 -inch (19 mocouple rake (Figure 5).

mm) thick calcium silicate insulating board (4) Turn on the burner and ensure that the which is attached to a steel angle bracket for thermocouples are reading 1900 ± 100 ° F (1038 placement in the test stand during burner ± 38 ° C) to ensure steady state conditions calibration, as shown in Figure 4. have been achieved.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(5) If the calorimeter and thermocouples do run for 2 minutes to provide adequate not read within range, repeat steps in para- warmup of the burner cone and flame sta- graphs 1 through 4 and adjust the burner air bilization.

intake damper until the proper readings are (5) To begin the test, swing the burner into obtained. The thermocouple rake and the the test position and simultaneously start calorimeter should be used frequently to the timing device.

maintain and record calibrated test param- (6) Expose the seat bottom cushion speci- eters. Until the specific apparatus has dem- men to the burner flame for 2 minutes and onstrated consistency, each test should be then turn off the burner. Immediately swing calibrated. After consistency has been con- the burner away from the test position. Ter- firmed, several tests may be conducted with minate test 7 minutes after initiating cush- the pre-test calibration before and a calibra- ion exposure to the flame by use of a gaseous tion check after the series. extinguishing agent (i.e., Halon or CO ).

(g) Test Procedure. The flammability of (7) Determine the weight of the remains of each set of specimens must be tested as fol- the seat cushion specimen set left on the lows: mounting stand to the nearest 0.02 pound (9 (1) Record the weight of each set of seat grams) excluding all droppings.

bottom and seat back cushion specimens to (h) Test Report. With respect to all speci- be tested to the nearest 0.02 pound (9 grams). men sets tested for a particular seat cushion (2) Mount the seat bottom and seat back for which testing of compliance is performed, cushion test specimens on the test stand as the following information must be recorded: shown in Figure 2, securing the seat back (1) An identification and description of the cushion specimen to the test stand at the specimens being tested.

top. (2) The number of specimen sets tested.

(3) Swing the burner into position and en- (3) The initial weight and residual weight sure that the distance from the exit of the of each set, the calculated percentage weight burner cone to the side of the seat bottom loss of each set, and the calculated average cushion specimen is 4 ± ⁄8 inches (102 ± 3 mm). percentage weight loss for the total number (4) Swing the burner away from the test of sets tested.

position. Turn on the burner and allow it to (4) The burn length for each set tested.

Federal Aviation Administration, DOT Pt. 25, App. F

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

Part III—Test Method To Determine Flame Pen- such as joints, lamp assemblies, etc., the etration Resistance of Cargo Compartment failure of which would affect the capability Liners. of the liner to safely contain a fire.

(3) There must be no flame penetration of (a) Criteria for Acceptance. (1) At least three any specimen within 5 minutes after applica- specimens of cargo compartment sidewall or tion of the flame source, and the peak tem- ceiling liner panels must be tested.

perature measured at 4 inches above the (2) Each specimen tested must simulate upper surface of the horizontal test sample the cargo compartment sidewall or ceiling must not exceed 400 ° F.

liner panel, including any design features,

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(b) Summary of Method. This method pro- orimeter face is parallel to the exit plane of vides a laboratory test procedure for meas- the test burner cone.

uring the capability of cargo compartment (4) Thermocouples. The seven thermocouples lining materials to resist flame penetration to be used for testing must be ⁄16 inch ce- with a 2 gallon per hour (GPH) #2 Grade ker- ramic sheathed, type K, grounded thermocouples with a nominal 30 American osene or equivalent burner fire source. Ceil- ing and sidewall liner panels may be tested wire gage (AWG) size conductor. The seven individually provided a baffle is used to sim- thermocouples must be attached to a steel ulate the missing panel. Any specimen that angle bracket to form a thermocouple rake passes the test as a ceiling liner panel may for placement in the test stand during burn- be used as a sidewall liner panel. er calibration as shown in Figure 3 of this (c) Test Specimens. (1) The specimen to be part of this appendix.

tested must measure 16 ± ⁄8 inches (406 ± 3 (5) Apparatus Arrangement. The test burner mm) by 24 + ⁄ 8 inches (610 ± 3 mm). must be mounted on a suitable stand to posi- (2) The specimens must be conditioned at tion the exit of the burner cone a distance of 70 ° F. ± 5 ° F. (21 ° C. ± 2 ° C.) and 55% ± 5% humid- 8 inches from the ceiling liner panel and 2 ity for at least 24 hours before testing. inches from the sidewall liner panel. The (d) Test Apparatus. The arrangement of the burner stand should have the capability of test apparatus, which is shown in Figure 3 of allowing the burner to be swung away from Part II and Figures 1 through 3 of this part the test specimen during warm-up periods.

of appendix F, must include the components (6) Instrumentation. A recording potentiom- described in this section. Minor details of the eter or other suitable instrument with an ap- apparatus may vary, depending on the model propriate range must be used to measure and of the burner used. record the outputs of the calorimeter and (1) Specimen Mounting Stand. The mounting the thermocouples.

stand for the test specimens consists of steel (7) Timing Device. A stopwatch or other de- angles as shown in Figure 1. vice must be used to measure the time of (2) Test Burner. The burner to be used in flame application and the time of flame pen- tesing must— etration, if it occurs.

(i) Be a modified gun type. (e) Preparation of Apparatus. Before calibra- (ii) Use a suitable nozzle and maintain fuel tion, all equipment must be turned on and pressure to yield a 2 GPH fuel flow. For ex- allowed to stabilize, and the burner fuel flow ample: an 80 degree nozzle nominally rated must be adjusted as specified in paragraph at 2.25 GPH and operated at 85 pounds per (d)(2).

square inch (PSI) gage to deliver 2.03 GPH. (f) Calibration. To ensure the proper ther- (iii) Have a 12 inch (305 mm) burner exten- mal output of the burner the following test sion installed at the end of the draft tube must be made: with an opening 6 inches (152 mm) high and (1) Remove the burner extension from the 11 inches (280 mm) wide as shown in Figure end of the draft tube. Turn on the blower 3 of Part II of this appendix.

portion of the burner without turning the (iv) Have a burner fuel pressure regulator fuel or igniters on. Measure the air velocity that is adjusted to deliver a nominal 2.0 GPH using a hot wire anemometer in the center of of #2 Grade kerosene or equivalent.

the draft tube across the face of the opening.

Adjust the damper such that the air velocity Burner models which have been used success- is in the range of 1550 to 1800 ft./min. If tabs fully in testing are the Lenox Model OB–32, are being used at the exit of the draft tube, Carlin Model 200 CRD and Park Model DPL.

The basic burner is described in FAA Power- they must be removed prior to this measure- plant Engineering Report No. 3A, Standard ment. Reinstall the draft tube extension Fire Test Apparatus and Procedure for Flexi- cone.

ble Hose Assemblies, dated March 1978; how- (2) Place the calorimeter on the test stand ever, the test settings specified in this ap- as shown in Figure 2 at a distance of 8 inches pendix differ in some instances from those (203 mm) from the exit of the burner cone to specified in the report. simulate the position of the horizontal test (3) Calorimeter. (i) The calorimeter to be specimen.

(3) Turn on the burner, allow it to run for used in testing must be a total heat flux Foil Type Gardon Gage of an appropriate range 2 minutes for warm-up, and adjust the damp- (approximately 0 to 15.0 British thermal unit er to produce a calorimeter reading of 8.0 ± 0.5 2 2 2 2 (BTU) per ft. sec., 0–17.0 watts/cm ). The cal- BTU per ft. sec. (9.1 ± 0.6 Watts/cm ).

orimeter must be mounted in a 6 inch by 12 (4) Replace the calorimeter with the ther- inch (152 by 305 mm) by ⁄ 4 inch (19 mm) thick mocouple rake (see Figure 3).

insulating block which is attached to a steel (5) Turn on the burner and ensure that angle bracket for placement in the test stand each of the seven thermocouples reads 1700 during burner calibration as shown in Figure ° F. ± 100 ° F. (927 ° C. ± 38 ° C.) to ensure steady 2 of this part of this appendix. state conditions have been achieved. If the (ii) The insulating block must be mon- temperature is out of this range, repeat steps itored for deterioration and the mounting 2 through 5 until proper readings are ob- shimmed as necessary to ensure that the cal- tained.

Federal Aviation Administration, DOT Pt. 25, App. F

(6) Turn off the burner and remove the The test may be terminated earlier if flame thermocouple rake. penetration is observed.

(7) Repeat (1) to ensure that the burner is (5) When testing ceiling liner panels, in the correct range.

record the peak temperature measured 4 (g) Test Procedure. (1) Mount a thermo- inches above the sample.

couple of the same type as that used for cali- (6) Record the time at which flame pene- bration at a distance of 4 inches (102 mm) tration occurs if applicable.

above the horizontal (ceiling) test specimen.

(h) Test Report. The test report must in- The thermocouple should be centered over clude the following: the burner cone.

(1) A complete description of the materials (2) Mount the test specimen on the test tested including type, manufacturer, thick- stand shown in Figure 1 in either the hori- ness, and other appropriate data.

zontal or vertical position. Mount the insu- lating material in the other position.

(2) Observations of the behavior of the test (3) Position the burner so that flames will specimens during flame exposure such as not impinge on the specimen, turn the burn- delamination, resin ignition, smoke, ect., in- er on, and allow it to run for 2 minutes. Ro- cluding the time of such occurrence.

tate the burner to apply the flame to the (3) The time at which flame penetration specimen and simultaneously start the tim- occurs, if applicable, for each of the three ing device.

specimens tested.

(4) Expose the test specimen to the flame (4) Panel orientation (ceiling or sidewall).

for 5 minutes and then turn off the burner.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

Part IV—Test Method To Determine the Heat this part IV. The truncated diamond-shaped Release Rate From Cabin Materials Exposed mask of .042 ± .002 inch (1.07 ± .05mm) stainless to Radiant Heat. steel must be added to provide uniform heat flux density over the area occupied by the (a) Summary of Method. Three or more vertical sample.

specimens representing the completed air- (4) Air Distribution System. The air entering craft component are tested. Each test speci- the environmental chamber must be distrib- men is injected into an environmental cham- uted by a .25 inch (6.3 mm) thick aluminum ber through which a constant flow of air plate having eight No. 4 drill-holes, located 2 passes. The specimen’s exposure is deter- inches (51 mm) from sides on 4 inch (102 mm) mined by a radiant heat source adjusted to centers, mounted at the base of the environ- produce, on the specimen, the desired total mental chamber. A second plate of 18 guage heat flux of 3.5 W/cm . The specimen is tested stainless steel having 120, evenly spaced, No.

with the exposed surface vertical. Combus- 28 drill holes must be mounted 6 inches (152 tion is initiated by piloted ignition. The mm) above the aluminum plate. A well-regu- combustion products leaving the chamber lated air supply is required. The air-supply are monitored in order to calculate the re- manifold at the base of the pyramidal sec- lease rate of heat.

tion must have 48, evenly spaced, No. 26 drill (b) Apparatus. The Ohio State University holes located .38 inch (10 mm) from the inner (OSU) rate of heat release apparatus, as de- edge of the manifold, resulting in an airflow scribed below, is used. This is a modified split of approximately three to one within version of the rate of heat release apparatus the apparatus.

standardized by the American Society of (5) Exhaust Stack. An exhaust stack, 5.25 × Testing and Materials (ASTM), ASTM E–906.

2.75 inches (133 × 70 mm) in cross section, and (1) This apparatus is shown in Figures 1A 10 inches (254 mm) long, fabricated from 28 and 1B of this part IV. All exterior surfaces guage stainless steel must be mounted on of the apparatus, except the holding cham- the outlet of the pyramidal section. A. 1.0 × ber, must be insulated with 1 inch (25 mm) 3.0 inch (25 × 76 mm) baffle plate of .018 ± .002 thick, low density, high temperature, fiber- inch (.50 ± .05 mm) stainless steel must be glass board insulation. A gasketed door, centered inside the stack, perpendicular to through which the sample injection rod the air flow, 3 inches (76 mm) above the base slides, must be used to form an airtight clo- of the stack.

sure on the specimen hold chamber.

(6) Specimen Holders. (i) The specimen must (2) Thermopile. The temperature difference be tested in a vertical orientation. The speci- between the air entering the environmental men holder (Figure 3 of this part IV) must chamber and that leaving must be monitored incorporate a frame that touches the speci- by a thermopile having five hot, and five men (which is wrapped with aluminum foil cold, 24-guage Chromel-Alumel junctions.

as required by paragraph (d)(3) of this Part) The hot junctions must be spaced across the along only the .25 inch (6 mm) perimeter. A top of the exhaust stack, .38 inches (10 mm) ‘‘V’’ shaped spring is used to hold the assem- below the top of the chimney. The bly together. A detachable .50 × 50 × 5.91 inch thermocouples must have a .050 ± .010 inch (12 × 12 × 150 mm) drip pan and two .020 inch (1.3 ± .3mm) diameter, ball-type, welded tip.

(.5 mm) stainless steel wires (as shown in One thermocouple must be located in the Figure 3 of this part IV) must be used for geometric center, with the other four located testing materials prone to melting and drip- 1.18 inch (30 mm) from the center along the ping. The positioning of the spring and frame diagonal toward each of the corners (Figure may be changed to accommodate different 5 of this part IV). The cold junctions must be specimen thicknesses by inserting the re- located in the pan below the lower air dis- taining rod in different holes on the speci- tribution plate (see paragraph (b)(4) of this men holder.

part IV). Thermopile hot junctions must be (ii) Since the radiation shield described in cleared of soot deposits as needed to main- ASTM E–906 is not used, a guide pin must be tain the calibrated sensitivity. added to the injection mechanism. This fits (3) Radiation Source. A radiant heat source into a slotted metal plate on the injection incorporating four Type LL silicon carbide mechanism outside of the holding chamber.

elements, 20 inches (508 mm) long by .63 inch It can be used to provide accurate posi- (16 mm) O.D., must be used, as shown in Fig- tioning of the specimen face after injection.

ures 2A and 2B of this part IV. The heat The front surface of the specimen must be 3.9 source must have a nominal resistance of 1.4 inches (100 mm) from the closed radiation ohms and be capable of generating a flux up doors after injection.

to 100 kW/m . The silicone carbide elements (iii) The specimen holder clips onto the must be mounted in the stainless steel panel mounted bracket (Figure 3 of this part IV).

box by inserting them through .63 inch (16 The mounting bracket must be attached to mm) holes in .03 inch (1 mm) thick ceramic the injection rod by three screws that pass fiber or calcium-silicate millboard. Loca- through a wide-area washer welded onto a ⁄ 2 - tions of the holes in the pads and stainless inch (13 mm) nut. The end of the injection steel cover plates are shown in Figure 2B of rod must be threaded to screw into the nut,

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

and a .020 inch (5.1 mm) thick wide area the tubing, .50 inch (13 mm) apart, for gas washer must be held between two ⁄2-inch (13 ports, all radiating in the same direction.

mm) nuts that are adjusted to tightly cover The first hole must be .50 inch (13 mm) from the hole in the radiation doors through the closed end of the tubing. The tube must which the injection rod or calibration calo- be positioned above the specimen holder so rimeter pass. that the holes are placed above the specimen (7) Calorimeter. A total-flux type calo- as shown in Figure 1B of this part IV. The rimeter must be mounted in the center of a fuel supplied to the burner must be methane ⁄2 -inch Kaowool ‘‘M’’ board inserted in the mixed with air in a ratio of approximately sample holder to measure the total heat flux. 50/50 by volume. The total gas flow must be The calorimeter must have a view angle of adjusted to produce flame lengths of 1 inch 180 degrees and be calibrated for incident (25 mm). When the gas/air ratio and the flow flux. The calorimeter calibration must be ac- rate are properly adjusted, approximately .25 ceptable to the Administrator. inch (6 mm) of the flame length appears yel- (8) Pilot-Flame Positions. Pilot ignition of low in color.

the specimen must be accomplished by si- (c) Calibration of Equipment —(1) Heat Re- multaneously exposing the specimen to a lease Rate. A calibration burner, as shown in lower pilot burner and an upper pilot burner, Figure 4, must be placed over the end of the as described in paragraph (b)(8)(i) and lower pilot flame tubing using a gas tight (b)(8)(ii) or (b)(8)(iii) of this part IV, respec- connection. The flow of gas to the pilot tively. Since intermittent pilot flame extin- flame must be at least 99 percent methane guishment for more than 3 seconds would in- and must be accurately metered. Prior to validate the test results, a spark ignitor may usage, the wet test meter must be properly be installed to ensure that the lower pilot leveled and filled with distilled water to the burner remains lighted.

tip of the internal pointer while no gas is (i) Lower Pilot Burner. The pilot-flame tub- flowing. Ambient temperature and pressure ing must be .25 inch (6.3 mm) O.D., .03 inch of the water are based on the internal wet (0.8mm) wall, stainless steel tubing. A mix- test meter temperature. A baseline flow rate 3 3 ture of 120 cm /min. of methane and 850 cm / of approximately 1 liter/min. must be set and min. of air must be fed to the lower pilot increased to higher preset flows of 4, 6, 8, 6 flame burner. The normal position of the end and 4 liters/min. Immediately prior to re- of the pilot burner tubing is .40 inch (10 mm) cording methane flow rates, a flow rate of 8 from and perpendicular to the exposed liters/min. must be used for 2 minutes to pre- vertical surface of the specimen. The center- condition the chamber. This is not recorded line at the outlet of the burner tubing must as part of calibration. The rate must be de- intersect the vertical centerline of the sam- termined by using a stopwatch to time a ple at a point .20 inch (5 mm) above the lower complete revolution of the wet test meter for exposed edge of the specimen. both the baseline and higher flow, with the (ii) Standard Three-Hole Upper Pilot Burner. flow returned to baseline before changing to The pilot burner must be a straight length of the next higher flow. The thermopile base- .25 inch (6.3 mm) O.D., .03 inch (0.8 mm) wall, line voltage must be measured. The gas flow stainless steel tubing that is 14 inches (360 to the burner must be increased to the high- mm) long. One end of the tubing must be er preset flow and allowed to burn for 2.0 closed, and three No. 40 drill holes must be minutes, and the thermopile voltage must be drilled into the tubing, 2.38 inch (60 mm) measured. The sequence must be repeated apart, for gas ports, all radiating in the same until all five values have been determined.

direction. The first hole must be .19 inch (5 The average of the five values must be used mm) from the closed end of the tubing. The as the calibration factor. The procedure tube must be positioned .75 inch (19 mm) must be repeated if the percent relative above and .75 inch (19 mm) behind the ex- standard deviation is greater than 5 percent.

posed upper edge of the specimen. The mid- Calculations are shown in paragraph (f) of dle hole must be in the vertical plane perpen- this part IV.

dicular to the exposed surface of the speci- (2) Flux Uniformity. Uniformity of flux over men which passes through its vertical cen- the specimen must be checked periodically terline and must be pointed toward the radi- and after each heating element change to de- ation source. The gas supplied to the burner termine if it is within acceptable limits of must be methane and must be adjusted to plus or minus 5 percent.

produce flame lengths of 1 inch (25 mm). (3) As noted in paragraph (b)(2) of this part (iii) Optional Fourteen-Hole Upper Pilot IV, thermopile hot junctions must be cleared Burner. This burner may be used in lieu of of soot deposits as needed to maintain the the standard three-hole burner described in calibrated sensitivity.

paragraph (b)(8)(ii) of this part IV. The pilot (d) Preparation of Test Specimens. (1) The burner must be a straight length of .25 inch test specimens must be representative of the (6.3 mm) O.D., .03 inch (0.8 mm) wall, stain- aircraft component in regard to materials less steel tubing that is 15.75 inches (400 mm) and construction methods. The standard size long. One end of the tubing must be closed, for the test specimens is 5.91 ± .03 × 5.91 ± .03 and 14 No. 59 drill holes must be drilled into inches (149 ± 1 × 149 ± 1 mm). The thickness of

Federal Aviation Administration, DOT Pt. 25, App. F

the specimen must be the same as that of the (4) The specimen must be placed in the aircraft component it represents up to a hold chamber with the radiation doors maximum thickness of 1.75 inches (45 mm). closed. The airtight outer door must be se- Test specimens representing thicker compo- cured, and the recording devices must be nents must be 1.75 inches (45 mm). started. The specimen must be retained in (2) Conditioning. Specimens must be condi- the hold chamber for 60 seconds, plus or tioned as described in Part 1 of this appen- minus 10 seconds, before injection. The ther- dix.

mopile ‘‘zero’’ value must be determined dur- (3) Mounting. Each test specimen must be ing the last 20 seconds of the hold period.

wrapped tightly on all sides of the specimen, The sample must not be injected before com- except for the one surface that is exposed pletion of the ‘‘zero’’ value determination.

with a single layer of .001 inch (.025 mm) alu- (5) When the specimen is to be injected, the minum foil.

radiation doors must be opened. After the (e) Procedure. (1) The power supply to the specimen is injected into the environmental radiant panel must be set to produce a radi- chamber, the radiation doors must be closed ant flux of 3.5 ± .05 W/cm , as measured at the behind the specimen.

point the center of the specimen surface will (6) [Reserved] occupy when positioned for the test. The ra- (7) Injection of the specimen and closure of diant flux must be measured after the air the inner door marks time zero. A record of flow through the equipment is adjusted to the thermopile output with at least one data the desired rate.

point per second must be made during the (2) After the pilot flames are lighted, their time the specimen is in the environmental position must be checked as described in chamber.

paragraph (b)(8) of this part IV.

(8) The test duration is five minutes. The (3) Air flow through the apparatus must be lower pilot burner and the upper pilot burner controlled by a circular plate orifice located must remain lighted for the entire duration in a 1.5 inch (38.1 mm) I.D. pipe with two of the test, except that there may be inter- pressure measuring points, located 1.5 inches mittent flame extinguishment for periods (38 mm) upstream and .75 inches (19 mm) that do not exceed 3 seconds. Furthermore, if downstream of the orifice plate. The pipe the optional three-hole upper burner is used, must be connected to a manometer set at a at least two flamelets must remain lighted pressure differential of 7.87 inches (200 mm) for the entire duration of the test, except of Hg. (See Figure 1B of this part IV.) The that there may be intermittent flame extin- total air flow to the equipment is approxi- guishment of all three flamelets for periods mately .04 m /seconds. The stop on the that do not exceed 3 seconds.

vertical specimen holder rod must be ad- (9) A minimum of three specimens must be justed so that the exposed surface of the tested.

specimen is positioned 3.9 inches (100 mm) from the entrance when injected into the en- (f) Calculations. (1) The calibration factor is vironmental chamber. calculated as follows:

F F −

k P P mole CH STP WATT kw ( ) − ( ) − 210 8 22 273 4 . min

O 1 cal v

K = × × × × × ×

h V V mole T kcal w − 760 22 01433 1000 .41 .

( )

O a 1 F = flow of methane at baseline (1pm) K = calibration factor (kw/mv) 0 h F = higher preset flow of methane (1pm) (3) The integral of the heat release rate is V = thermopile voltage at baseline (mv) the total heat release as a function of time V = thermopile voltage at higher flow (mv) and is calculated by multiplying the rate by T a = Ambient temperature (K) the data sampling frequency in minutes and P = Ambient pressure (mm Hg) summing the time from zero to two minutes.

P = Water vapor pressure (mm Hg) v (g) Criteria. The total positive heat release (2) Heat release rates may be calculated over the first two minutes of exposure for from the reading of the thermopile output each of the three or more samples tested voltage at any instant of time as: must be averaged, and the peak heat release rate for each of the samples must be aver-

V V K −

( )

m b n aged. The average total heat release must

HRR =

not exceed 65 kilowatt-minutes per square

m .02323

meter, and the average peak heat release HRR = heat release rate (kw/m ) rate must not exceed 65 kilowatts per square V = baseline voltage (mv) b meter.

V = measured thermopile voltage (mv) m

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(h) Report. The test report must include (4) If melting, sagging, delaminating, or the following for each specimen tested: other behavior that affects the exposed sur- (1) Description of the specimen.

face area or the mode of burning occurs, (2) Radiant heat flux to the specimen, ex- these behaviors must be reported, together pressed in W/cm .

with the time at which such behaviors were (3) Data giving release rates of heat (in kW/ observed.

) as a function of time, either graphically m (5) The peak heat release and the 2-minute or tabulated at intervals no greater than 10 integrated heat release rate must be re- ) must be seconds. The calibration factor (k n ported.

recorded.

Federal Aviation Administration, DOT Pt. 25, App. F

F IGURES TO P ART IV OF A PPENDIX F

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

Part V. Test Method To Determine the Smoke Part VI—Test Method To Determine the Flam- Emission Characteristics of Cabin Materials mability and Flame Propagation Characteris- tics of Thermal/Acoustic Insulation Mate- (a) Summary of Method. The specimens rials must be constructed, conditioned, and tested Use this test method to evaluate the flam- in the flaming mode in accordance with mability and flame propagation characteris- American Society of Testing and Materials tics of thermal/acoustic insulation when ex- (ASTM) Standard Test Method ASTM F814– posed to both a radiant heat source and a 83.

flame.

(b) Acceptance Criteria. The specific optical (a) Definitions.

smoke density (D ), which is obtained by s ‘‘Flame propagation’’ means the furthest averaging the reading obtained after 4 min- distance of the propagation of visible flame utes with each of the three specimens, shall towards the far end of the test specimen, not exceed 200. measured from the midpoint of the ignition source flame. Measure this distance after initially applying the ignition source and be- fore all flame on the test specimen is extin- guished. The measurement is not a deter- mination of burn length made after the test.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

‘‘Radiant heat source’’ means an electric material encapsulated by a film covering and or air propane panel. foams.

‘‘Thermal/acoustic insulation’’ means a ‘‘Zero point’’ means the point of applica- material or system of materials used to pro- tion of the pilot burner to the test specimen.

vide thermal and/or acoustic protection. Ex- (b) Test apparatus.

amples include fiberglass or other batting (1) Radiant panel test chamber. Conduct window to provide access to the movable tests in a radiant panel test chamber (see specimen platform holder. The bottom of the figure 1 above). Place the test chamber under test chamber must be a sliding steel plat- an exhaust hood to facilitate clearing the form that has provision for securing the test chamber of smoke after each test. The radi- specimen holder in a fixed and level position.

ant panel test chamber must be an enclosure The chamber must have an internal chimney 55 inches (1397 mm) long by 19.5 (495 mm) with exterior dimensions of 5.1 inches (129 deep by 28 (710 mm) to 30 inches (maximum) mm) wide, by 16.2 inches (411 mm) deep by 13 (762 mm) above the test specimen. Insulate inches (330 mm) high at the opposite end of the sides, ends, and top with a fibrous ce- the chamber from the radiant energy source.

TM ramic insulation, such as Kaowool M The interior dimensions must be 4.5 inches board. On the front side, provide a 52 by 12- (114 mm) wide by 15.6 inches (395 mm) deep.

inch (1321 by 305 mm) draft-free, high-tem- The chimney must extend to the top of the perature, glass window for viewing the sam- chamber (see figure 2).

ple during testing. Place a door below the

Federal Aviation Administration, DOT Pt. 25, App. F

(2) Radiant heat source. Mount the radiant temperatures up to 1300 ° F (704 ° C). An air heat energy source in a cast iron frame or propane panel must be made of a porous re- equivalent. An electric panel must have six, fractory material and have a radiation sur- 3-inch wide emitter strips. The emitter strips face of 12 by 18 inches (305 by 457 mm). The must be perpendicular to the length of the panel must be capable of operating at tem- panel. The panel must have a radiation sur- peratures up to 1,500 ° F (816 ° C). See figures 3a 7 1 ⁄8 by 18 ⁄ 2 inches (327 by 470 mm).

face of 12 and 3b.

The panel must be capable of operating at

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

(i) Electric radiant panel. The radiant panel (3) Specimen holding system. (i) The sliding must be 3-phase and operate at 208 volts. A platform serves as the housing for test speci- men placement. Brackets may be attached single-phase, 240 volt panel is also accept- (via wing nuts) to the top lip of the platform able. Use a solid-state power controller and in order to accommodate various thicknesses microprocessor-based controller to set the of test specimens. Place the test specimens electric panel operating parameters.

TM on a sheet of Kaowool M board or 1260 (ii) Gas radiant panel. Use propane (liquid Standard Board (manufactured by Thermal petroleum gas—2.1 UN 1075) for the radiant Ceramics and available in Europe), or equiv- panel fuel. The panel fuel system must con- alent, either resting on the bottom lip of the sist of a venturi-type aspirator for mixing sliding platform or on the base of the brack- gas and air at approximately atmospheric ets. It may be necessary to use multiple pressure. Provide suitable instrumentation sheets of material based on the thickness of for monitoring and controlling the flow of the test specimen (to meet the sample height fuel and air to the panel. Include an air flow requirement). Typically, these non-combus- gauge, an air flow regulator, and a gas pres- 1 tible sheets of material are available in ⁄ 4 sure gauge.

inch (6 mm) thicknesses. See figure 4. A slid- (iii) Radiant panel placement. Mount the ing platform that is deeper than the 2-inch panel in the chamber at 30 ° to the horizontal (50.8mm) platform shown in figure 4 is also specimen plane, and 7 ⁄2 inches above the zero acceptable as long as the sample height re- point of the specimen. quirement is met.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(ii) Attach a ⁄ 2 inch (13 mm) piece of (iii) Place the test specimen horizontally TM Kaowool M board or other high tempera- on the non-combustible board(s). Place a 1 1 ture material measuring 41 ⁄ 2 by 8 ⁄ 4 inches steel retaining/securing frame fabricated of (1054 by 210 mm) to the back of the platform. mild steel, having a thickness of ⁄ 8 inch (3.2 This board serves as a heat retainer and pro- mm) and overall dimensions of 23 by 13 ⁄8 tects the test specimen from excessive inches (584 by 333 mm) with a specimen open- preheating. The height of this board must ing of 19 by 10 ⁄4 inches (483 by 273 mm) over not impede the sliding platform movement the test specimen. The front, back, and right (in and out of the test chamber). If the plat- portions of the top flange of the frame must form has been fabricated such that the back rest on the top of the sliding platform, and side of the platform is high enough to pre- the bottom flanges must pinch all 4 sides of vent excess preheating of the specimen when the test specimen. The right bottom flange the sliding platform is out, a retainer board must be flush with the sliding platform. See is not necessary. figure 5.

Federal Aviation Administration, DOT Pt. 25, App. F

(4) Pilot Burner. The pilot burner used to (19 mm). A ⁄4 inch (19 mm) guide (such as a ignite the specimen must be a thin strip of metal) may be soldered to the TM Bernzomatic commercial propane venturi top of the burner to aid in setting the flame torch with an axially symmetric burner tip height. The overall flame length must be ap- and a propane supply tube with an orifice di- proximately 5 inches long (127 mm). Provide ameter of 0.006 inches (0.15 mm). The length a way to move the burner out of the ignition of the burner tube must be 2 ⁄8 inches (71 position so that the flame is horizontal and mm). The propane flow must be adjusted via at least 2 inches (50 mm) above the specimen gas pressure through an in-line regulator to plane. See figure 6.

produce a blue inner cone length of ⁄4 inch

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(5) Thermocouples. Install a 24 American (E) The graphite plate must be electrically Wire Gauge (AWG) Type K (Chromel-Alumel) heated, have a clear surface area on each thermocouple in the test chamber for tem- side of the plate of at least 2 by 2 inches (51 1 1 perature monitoring. Insert it into the by 51 mm), and be ⁄8 inch ± ⁄16 inch thick (3.2 chamber through a small hole drilled ± 1.6 mm).

through the back of the chamber. Place the (F) Center the 2 transducers on opposite thermocouple so that it extends 11 inches sides of the plates at equal distances from (279 mm) out from the back of the chamber the plate.

wall, 11 ⁄2 inches (292 mm) from the right side (G) The distance of the calorimeter to the of the chamber wall, and is 2 inches (51 mm) plate must be no less than 0.0625 inches (1.6 below the radiant panel. The use of other mm), nor greater than 0.375 inches (9.5 mm).

(H) The range used in calibration must be thermocouples is optional.

at least 0–3.5 BTUs/ft second (0–3.9 Watts/ (6) Calorimeter. The calorimeter must be a 2 2 cm ) and no greater than 0–5.7 BTUs/ft sec- one-inch cylindrical water-cooled, total heat ond (0–6.4 Watts/cm ).

flux density, foil type Gardon Gage that has 2 (I) The recording device used must record a range of 0 to 5 BTU/ft -second (0 to 5.7 2 the 2 transducers simultaneously or at least Watts/cm ).

within ⁄ 10 of each other.

(7) Calorimeter calibration specification and (8) Calorimeter fixture. With the sliding plat- procedure.

form pulled out of the chamber, install the (i) Calorimeter specification.

calorimeter holding frame and place a sheet (A) Foil diameter must be 0.25 ± 0.005 inches of non-combustible material in the bottom (6.35 ± 0.13 mm).

of the sliding platform adjacent to the hold- (B) Foil thickness must be 0.0005 ± 0.0001 ing frame. This will prevent heat losses dur- inches (0.013 ± 0.0025 mm).

ing calibration. The frame must be 13 ⁄8 (C) Foil material must be thermocouple inches (333 mm) deep (front to back) by 8 grade Constantan.

inches (203 mm) wide and must rest on the (D) Temperature measurement must be a top of the sliding platform. It must be fab- Copper Constantan thermocouple.

ricated of ⁄8 inch (3.2 mm) flat stock steel (E) The copper center wire diameter must 1 and have an opening that accommodates a ⁄2 be 0.0005 inches (0.013 mm).

inch (12.7 mm) thick piece of refractory (F) The entire face of the calorimeter must board, which is level with the top of the slid- be lightly coated with ‘‘Black Velvet’’ paint ing platform. The board must have three 1- having an emissivity of 96 or greater.

inch (25.4 mm) diameter holes drilled (ii) Calorimeter calibration. (A) The calibra- through the board for calorimeter insertion.

tion method must be by comparison to a like The distance to the radiant panel surface standardized transducer.

from the centerline of the first hole (‘‘zero’’ 1 1 (B) The standardized transducer must meet position) must be 7 ⁄2 ± ⁄8 inches (191 ± 3 mm).

the specifications given in paragraph VI(b)(6) The distance between the centerline of the of this appendix.

first hole to the centerline of the second hole (C) Calibrate the standard transducer must be 2 inches (51 mm). It must also be the against a primary standard traceable to the same distance from the centerline of the sec- National Institute of Standards and Tech- ond hole to the centerline of the third hole.

nology (NIST). See figure 7. A calorimeter holding frame (D) The method of transfer must be a heat- that differs in construction is acceptable as ed graphite plate. long as the height from the centerline of the

Federal Aviation Administration, DOT Pt. 25, App. F

first hole to the radiant panel and the dis- tance between holes is the same as described in this paragraph.

(9) Instrumentation. Provide a calibrated re- cover material is over-cut enough to be cording device with an appropriate range or drawn down the sides without compressing a computerized data acquisition system to the core material. The fastening means measure and record the outputs of the calo- should be as continuous as possible along the rimeter and the thermocouple. The data ac- length of the seams. The specimen thickness quisition system must be capable of record- must be of the same thickness as installed in ing the calorimeter output every second dur- the airplane.

ing calibration.

(3) Specimen Dimensions. To facilitate prop- (10) Timing device. Provide a stopwatch or er placement of specimens in the sliding other device, accurate to ± 1 second/hour, to platform housing, cut non-rigid core mate- measure the time of application of the pilot rials, such as fiberglass, 12 ⁄2 inches (318mm) burner flame. wide by 23 inches (584mm) long. Cut rigid 1 1 (c) Test specimens. (1) Specimen preparation. materials, such as foam, 11 ⁄2 ± ⁄ 4 inches (292 Prepare and test a minimum of three test mm ± 6mm) wide by 23 inches (584mm) long in specimens. If an oriented film cover material order to fit properly in the sliding platform is used, prepare and test both the warp and housing and provide a flat, exposed surface fill directions. equal to the opening in the housing.

(d) Specimen conditioning. Condition the (2) Construction. Test specimens must in- test specimens at 70 ± 5 ° F (21 ± 2 ° C) and 55% clude all materials used in construction of ± 10% relative humidity, for a minimum of 24 the insulation (including batting, film, hours prior to testing.

scrim, tape etc.). Cut a piece of core material such as foam or fiberglass, and cut a piece of (e) Apparatus Calibration. (1) With the slid- film cover material (if used) large enough to ing platform out of the chamber, install the cover the core material. Heat sealing is the calorimeter holding frame. Push the plat- preferred method of preparing fiberglass form back into the chamber and insert the samples, since they can be made without calorimeter into the first hole (‘‘zero’’ posi- compressing the fiberglass (‘‘box sample’’). tion). See figure 7. Close the bottom door lo- Cover materials that are not heat sealable cated below the sliding platform. The dis- may be stapled, sewn, or taped as long as the tance from the centerline of the calorimeter

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

to the radiant panel surface at this point up to 1 hour). The pilot burner must be off 1 1 must be 7. ⁄ 2 inches ± ⁄8 (191 mm ± 3). Prior to and in the down position during this time.

igniting the radiant panel, ensure that the (3) After steady-state conditions have been calorimeter face is clean and that there is reached, move the calorimeter 2 inches (51 water running through the calorimeter. mm) from the ‘‘zero’’ position (first hole) to (2) Ignite the panel. Adjust the fuel/air position 1 and record the heat flux. Move the mixture to achieve 1.5 BTUs/ft -second ± 5% calorimeter to position 2 and record the heat (1.7 Watts/cm ± 5%) at the ‘‘zero’’ position. If flux. Allow enough time at each position for using an electric panel, set the power con- the calorimeter to stabilize. Table 1 depicts troller to achieve the proper heat flux. Allow typical calibration values at the three posi- the unit to reach steady state (this may take tions.

T ABLE 1—C ALIBRATION T ABLE 2 2 Position BTU’s/ft sec Watts/cm ‘‘Zero’’ Position ...................................................................................................... 1.5 1.7 Position 1 ............................................................................................................... 1.51–1.50–1.49 1.71–1.70–1.69 Position 2 ............................................................................................................... 1.43–1.44 1.62–1.63 (4) Open the bottom door, remove the calo- make a slit in the film cover to purge any air rimeter and holder fixture. Use caution as inside. This allows the operator to maintain the fixture is very hot. the proper test specimen position (level with (f) Test Procedure. (1) Ignite the pilot burn- the top of the platform) and to allow ventila- er. Ensure that it is at least 2 inches (51 mm) tion of gases during testing. A longitudinal above the top of the platform. The burner slit, approximately 2 inches (51mm) in must not contact the specimen until the test length, must be centered 3 inches ± ⁄2 inch begins. (76mm ± 13mm) from the left flange of the se- (2) Place the test specimen in the sliding curing frame. A utility knife is acceptable platform holder. Ensure that the test sample for slitting the film cover.

surface is level with the top of the platform. (4) Immediately push the sliding platform At ‘‘zero’’ point, the specimen surface must into the chamber and close the bottom door.

1 1 be 7 ⁄ 2 inches ± ⁄8 inch (191 mm ± 3) below the (5) Bring the pilot burner flame into con- radiant panel. tact with the center of the specimen at the (3) Place the retaining/securing frame over ‘‘zero’’ point and simultaneously start the the test specimen. It may be necessary (due timer. The pilot burner must be at a 27 ° to compression) to adjust the sample (up or angle with the sample and be approximately down) in order to maintain the distance from ⁄2 inch (12 mm) above the sample. See figure the sample to the radiant panel (7 ⁄2 inches 7. A stop, as shown in figure 8, allows the op- ± ⁄8 inch (191 mm ± 3) at ‘‘zero’’ position). With erator to position the burner correctly each film/fiberglass assemblies, it is critical to time.

Federal Aviation Administration, DOT Pt. 25, App. F

(6) Leave the burner in position for 15 sec- to the left of the centerline of the pilot onds and then remove to a position at least flame application.

(2) The flame time after removal of the 2 inches (51 mm) above the specimen.

pilot burner may not exceed 3 seconds on any (g) Report. (1) Identify and describe the specimen.

test specimen.

(2) Report any shrinkage or melting of the Part VII—Test Method To Determine the test specimen.

Burnthrough Resistance of Thermal/Acoustic (3) Report the flame propagation distance.

Insulation Materials If this distance is less than 2 inches, report Use the following test method to evaluate this as a pass (no measurement required).

the burnthrough resistance characteristics (4) Report the after-flame time.

of aircraft thermal/acoustic insulation mate- (h) Requirements. (1) There must be no rials when exposed to a high intensity open flame propagation beyond 2 inches (51 mm) flame.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(a) Definitions. the test rig, at an angle of 30 ° with respect to Burnthrough time means the time, in sec- vertical.

onds, for the burner flame to penetrate the Specimen set means two insulation blanket test specimen, and/or the time required for specimens. Both specimens must represent the heat flux to reach 2.0 Btu/ft sec (2.27 W/ the same production insulation blanket con- cm ) on the inboard side, at a distance of 12 struction and materials, proportioned to cor- inches (30.5 cm) from the front surface of the respond to the specimen size.

insulation blanket test frame, whichever is (b) Apparatus. (1) The arrangement of the sooner. The burnthrough time is measured at test apparatus is shown in figures 1 and 2 and the inboard side of each of the insulation must include the capability of swinging the blanket specimens.

burner away from the test specimen during Insulation blanket specimen means one of warm-up.

two specimens positioned in either side of

Federal Aviation Administration, DOT Pt. 25, App. F

(2) Test burner. The test burner must be a eters such as fuel pressure, nozzle depth, sta- modified gun-type such as the Park Model tor position, and intake airflow must be DPL 3400. Flame characteristics are highly properly adjusted to achieve the correct dependent on actual burner setup. Param- flame output.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(i) Nozzle. A nozzle must maintain the fuel (ii) Fuel Rail. The fuel rail must be ad- pressure to yield a nominal 6.0 gal/hr (0.378 L/ justed to position the fuel nozzle at a depth min) fuel flow. A Monarch-manufactured 80 ° of 0.3125 inch (8 mm) from the end plane of PL (hollow cone) nozzle nominally rated at the exit stator, which must be mounted in 6.0 gal/hr at 100 lb/in (0.71 MPa) delivers a the end of the draft tube.

proper spray pattern. (iii) Internal Stator. The internal stator, lo- cated in the middle of the draft tube, must

Federal Aviation Administration, DOT Pt. 25, App. F

be positioned at a depth of 3.75 inches (95 (vi) Fuel. Use JP–8, Jet A, or their inter- mm) from the tip of the fuel nozzle. The sta- national equivalent, at a flow rate of 6.0 ± 0.2 tor must also be positioned such that the in- gal/hr (0.378 ± 0.0126 L/min). If this fuel is un- tegral igniters are located at an angle mid- available, ASTM K2 fuel (Number 2 grade way between the 10 and 11 o’clock position, kerosene) or ASTM D2 fuel (Number 2 grade when viewed looking into the draft tube.

fuel oil or Number 2 diesel fuel) are accept- Minor deviations to the igniter angle are ac- able if the nominal fuel flow rate, tempera- ceptable if the temperature and heat flux re- ture, and heat flux measurements conform to quirements conform to the requirements of the requirements of paragraph VII(e) of this paragraph VII(e) of this appendix.

appendix.

(iv) Blower Fan. The cylindrical blower fan (vii) Fuel pressure regulator. Provide a fuel used to pump air through the burner must pressure regulator, adjusted to deliver a measure 5.25 inches (133 mm) in diameter by nominal 6.0 gal/hr (0.378 L/min) flow rate. An 3.5 inches (89 mm) in width.

operating fuel pressure of 100 lb/in (0.71 (v) Burner cone. Install a 12 + 0.125-inch (305 MPa) for a nominally rated 6.0 gal/hr 80 ° ± 3 mm) burner extension cone at the end of spray angle nozzle (such as a PL type) deliv- the draft tube. The cone must have an open- ing 6 ± 0.125-inch (152 ± 3 mm) high and 11 ers 6.0 ± 0.2 gal/hr (0.378 ± 0.0126 L/min).

± 0.125-inch (280 ± 3 mm) wide (see figure 3).

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(3) Calibration rig and equipment. (i) Con- perature. Position the calibration rigs to struct individual calibration rigs to incor- allow movement of the burner from the test porate a calorimeter and thermocouple rake rig position to either the heat flux or tem- for the measurement of heat flux and tem- perature position with minimal difficulty.

Federal Aviation Administration, DOT Pt. 25, App. F

(ii) Calorimeter. The calorimeter must be a ± 3 mm) by 0.75 ± 0.125 inch (19 mm ± 3 mm) total heat flux, foil type Gardon Gage of an thick insulating block which is attached to appropriate range such as 0–20 Btu/ft -sec (0– the heat flux calibration rig during calibra- 22.7 W/cm ), accurate to ± 3% of the indicated tion (figure 4). Monitor the insulating block reading. The heat flux calibration method for deterioration and replace it when nec- must be in accordance with paragraph essary. Adjust the mounting as necessary to VI(b)(7) of this appendix.

ensure that the calorimeter face is parallel (iii) Calorimeter mounting. Mount the calo- to the exit plane of the test burner cone.

rimeter in a 6- by 12- ± 0.125 inch (152- by 305-

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

Federal Aviation Administration, DOT Pt. 25, App. F

(iv) Thermocouples. Provide seven ⁄8-inch Wire Gauge (AWG) size conductor for cali- (3.2 mm) ceramic packed, metal sheathed, bration. Attach the thermocouples to a steel type K (Chromel-alumel), grounded junction angle bracket to form a thermocouple rake thermocouples with a nominal 24 American

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

for placement in the calibration rig during inch (3.2 mm) thick steel as shown in figure burner calibration (figure 5). 1, except for the center vertical former, (v) Air velocity meter. Use a vane-type air which should be ⁄ 4-inch (6.4 mm) thick to velocity meter to calibrate the velocity of minimize warpage. The specimen mounting air entering the burner. An Omega Engineer- frame stringers (horizontal) should be bolted ing Model HH30A is satisfactory. Use a suit- to the test frame formers (vertical) such that able adapter to attach the measuring device the expansion of the stringers will not cause to the inlet side of the burner to prevent air the entire structure to warp. Use the mount- from entering the burner other than through ing frame for mounting the two insulation the measuring device, which would produce blanket test specimens as shown in figure 2.

erroneously low readings. Use a flexible duct, (5) Backface calorimeters. Mount two total measuring 4 inches wide (102 mm) by 20 feet heat flux Gardon type calorimeters behind long (6.1 meters), to supply fresh air to the the insulation test specimens on the back burner intake to prevent damage to the air side (cold) area of the test specimen mount- velocity meter from ingested soot. An op- ing frame as shown in figure 6. Position the tional airbox permanently mounted to the calorimeters along the same plane as the burner intake area can effectively house the air velocity meter and provide a mounting burner cone centerline, at a distance of 4 port for the flexible intake duct. inches (102 mm) from the vertical centerline (4) Test specimen mounting frame. Make the of the test frame.

⁄ 8 - mounting frame for the test specimens of

Federal Aviation Administration, DOT Pt. 25, App. F

(i) The calorimeters must be a total heat The heat flux calibration method must com- flux, foil type Gardon Gage of an appropriate ply with paragraph VI(b)(7) of this appendix.

2 2 range such as 0–5 Btu/ft -sec (0–5.7 W/cm ), accurate to ± 3% of the indicated reading.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

(6) Instrumentation. Provide a recording po- (i) Fire barrier material. If the insulation tentiometer or other suitable calibrated in- blanket is constructed with a fire barrier strument with an appropriate range to meas- material, place the fire barrier material in a ure and record the outputs of the calo- manner reflective of the installed arrange- rimeter and the thermocouples. ment For example, if the material will be (7) Timing device. Provide a stopwatch or placed on the outboard side of the insulation other device, accurate to ± 1%, to measure material, inside the moisture film, place it the time of application of the burner flame the same way in the test specimen.

and burnthrough time. (ii) Insulation material. Blankets that uti- (8) Test chamber. Perform tests in a suitable lize more than one variety of insulation chamber to reduce or eliminate the possi- (composition, density, etc.) must have speci- bility of test fluctuation due to air move- men sets constructed that reflect the insula- ment. The chamber must have a minimum tion combination used. If, however, several floor area of 10 by 10 feet (305 by 305 cm). blanket types use similar insulation com- (i) Ventilation hood. Provide the test cham- binations, it is not necessary to test each ber with an exhaust system capable of re- combination if it is possible to bracket the moving the products of combustion expelled various combinations.

during tests. (iii) Moisture barrier film. If a production (c) Test Specimens. (1) Specimen preparation. blanket construction utilizes more than one Prepare a minimum of three specimen sets of type of moisture barrier film, perform sepa- the same construction and configuration for rate tests on each combination. For example, testing. if a polyimide film is used in conjunction (2) Insulation blanket test specimen. with an insulation in order to enhance the (i) For batt-type materials such as fiber- burnthrough capabilities, also test the same glass, the constructed, finished blanket spec- insulation when used with a polyvinyl fluo- imen assemblies must be 32 inches wide by 36 ride film.

inches long (81.3 by 91.4 cm), exclusive of (iv) Installation on test frame. Attach the heat sealed film edges. blanket test specimens to the test frame (ii) For rigid and other non-conforming using 12 steel spring type clamps as shown in types of insulation materials, the finished figure 7. Use the clamps to hold the blankets test specimens must fit into the test rig in in place in both of the outer vertical such a manner as to replicate the actual in- formers, as well as the center vertical former service installation. (4 clamps per former). The clamp surfaces (3) Construction. Make each of the speci- should measure 1 inch by 2 inches (25 by 51 mens tested using the principal components mm). Place the top and bottom clamps 6 ( i.e., insulation, fire barrier material if used, inches (15.2 cm) from the top and bottom of and moisture barrier film) and assembly the test frame, respectively. Place the mid- processes (representative seams and clo- dle clamps 8 inches (20.3 cm) from the top sures). and bottom clamps.

Federal Aviation Administration, DOT Pt. 25, App. F

(Note: For blanket materials that cannot blower. Measure the airflow of the test be installed in accordance with figure 7 chamber using a vane anemometer or equiv- above, the blankets must be installed in a alent measuring device. The vertical air ve- manner approved by the FAA.) locity just behind the top of the upper insu- (v) Conditioning. Condition the specimens lation blanket test specimen must be 100 ± 50 at 70 ° ± 5 ° F (21 ° ± 2 ° C) and 55% ± 10% relative ft/min (0.51 ± 0.25 m/s). The horizontal air ve- humidity for a minimum of 24 hours prior to locity at this point must be less than 50 ft/ testing. min (0.25 m/s).

(d) Preparation of apparatus. (1) Level and (3) If a calibrated flow meter is not avail- center the frame assembly to ensure align- able, measure the fuel flow rate using a grad- ment of the calorimeter and/or thermocouple uated cylinder of appropriate size. Turn on rake with the burner cone. the burner motor/fuel pump, after insuring (2) Turn on the ventilation hood for the that the igniter system is turned off. Collect test chamber. Do not turn on the burner the fuel via a plastic or rubber tube into the

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. F

graduated cylinder for a 2-minute period. De- rimeter face. Ensure that the horizontal cen- termine the flow rate in gallons per hour. terline of the burner cone is offset 1 inch The fuel flow rate must be 6.0 ± 0.2 gallons per below the horizontal centerline of the calo- hour (0.378 ± 0.0126 L/min). rimeter (figure 8). Without disturbing the (e) Calibration. (1) Position the burner in calorimeter position, rotate the burner in front of the calorimeter so that it is centered front of the thermocouple rake, such that and the vertical plane of the burner cone exit the middle thermocouple (number 4 of 7) is is 4 ± 0.125 inches (102 ± 3 mm) from the calo- centered on the burner cone.

Federal Aviation Administration, DOT Pt. 25, App. F

Ensure that the horizontal centerline of the average temperature of each thermo- the burner cone is also offset 1 inch below couple over this 30-second period and record.

the horizontal centerline of the thermo- The average temperature of each of the 7 couple tips. Re-check measurements by ro- thermocouples should be 1900 ° F ± 100 ° F (1038 tating the burner to each position to ensure ± 56 ° C).

proper alignment between the cone and the (6) If either the heat flux or the tempera- calorimeter and thermocouple rake. (Note: tures are not within the specified range, ad- The test burner mounting system must in- just the burner intake air velocity and re- corporate ‘‘detents’’ that ensure proper cen- peat the procedures of paragraphs (4) and (5) tering of the burner cone with respect to above to obtain the proper values. Ensure both the calorimeter and the thermocouple that the inlet air velocity is within the rakes, so that rapid positioning of the burner range of 2150 ft/min ± 50 ft/min (10.92 ± 0.25 m/ can be achieved during the calibration proce- s).

dure.)

(7) Calibrate prior to each test until con- (2) Position the air velocity meter in the adapter or airbox, making certain that no sistency has been demonstrated. After con- gaps exist where air could leak around the sistency has been confirmed, several tests air velocity measuring device. Turn on the may be conducted with calibration con- blower/motor while ensuring that the fuel so- ducted before and after a series of tests.

lenoid and igniters are off. Adjust the air in- (f) Test procedure. (1) Secure the two insula- take velocity to a level of 2150 ft/min, (10.92 tion blanket test specimens to the test m/s) then turn off the blower/motor. (Note: frame. The insulation blankets should be at- The Omega HH30 air velocity meter meas- tached to the test rig center vertical former ures 2.625 inches in diameter. To calculate using four spring clamps positioned as shown the intake airflow, multiply the cross-sec- in figure 7 (according to the criteria of para- tional area (0.03758 ft ) by the air velocity 3 graph paragraph (c)(3)(iv) of this part of this (2150 ft/min) to obtain 80.80 ft /min. An air appendix).

velocity meter other than the HH30 unit can (2) Ensure that the vertical plane of the be used, provided the calculated airflow of 3 3 burner cone is at a distance of 4 ± 0.125 inch 80.80 ft /min (2.29 m /min) is equivalent.)

(102 ± 3 mm) from the outer surface of the (3) Rotate the burner from the test posi- tion to the warm-up position. Prior to light- horizontal stringers of the test specimen ing the burner, ensure that the calorimeter frame, and that the burner and test frame face is clean of soot deposits, and there is are both situated at a 30 ° angle with respect water running through the calorimeter. Ex- to vertical.

amine and clean the burner cone of any evi- (3) When ready to begin the test, direct the dence of buildup of products of combustion, burner away from the test position to the soot, etc. Soot buildup inside the burner warm-up position so that the flame will not cone may affect the flame characteristics impinge on the specimens prematurely. Turn and cause calibration difficulties. Since the on and light the burner and allow it to sta- burner cone may distort with time, dimen- bilize for 2 minutes.

sions should be checked periodically.

(4) To begin the test, rotate the burner into (4) While the burner is still rotated to the the test position and simultaneously start warm-up position, turn on the blower/motor, the timing device.

igniters and fuel flow, and light the burner.

(5) Expose the test specimens to the burner Allow it to warm up for a period of 2 min- flame for 4 minutes and then turn off the utes. Move the burner into the calibration position and allow 1 minute for calorimeter burner. Immediately rotate the burner out of stabilization, then record the heat flux once the test position.

every second for a period of 30 seconds. Turn (6) Determine (where applicable) the off burner, rotate out of position, and allow burnthrough time, or the point at which the 2 2 to cool. Calculate the average heat flux over heat flux exceeds 2.0 Btu/ft -sec (2.27 W/cm ).

this 30-second duration. The average heat (g) Report. (1) Identify and describe the flux should be 16.0 ± 0.8 Btu/ft sec (18.2 ± 0.9 W/ specimen being tested.

cm ).

(2) Report the number of insulation blan- (5) Position the burner in front of the ther- ket specimens tested.

mocouple rake. After checking for proper (3) Report the burnthrough time (if any), alignment, rotate the burner to the warm-up and the maximum heat flux on the back face position, turn on the blower/motor, igniters of the insulation blanket test specimen, and and fuel flow, and light the burner. Allow it the time at which the maximum occurred.

to warm up for a period of 2 minutes. Move (h) Requirements. (1) Each of the two insula- the burner into the calibration position and tion blanket test specimens must not allow allow 1 minute for thermocouple stabiliza- fire or flame penetration in less than 4 min- tion, then record the temperature of each of utes.

the 7 thermocouples once every second for a period of 30 seconds. Turn off burner, rotate (2) Each of the two insulation blanket test out of position, and allow to cool. Calculate specimens must not allow more than 2.0 Btu/

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. H

2 2 ft -sec (2.27 W/cm ) on the cold side of the in- (3) Basic control and operation information sulation specimens at a point 12 inches (30.5 describing how the airplane components and cm) from the face of the test rig. systems are controlled and how they oper- ate, including any special procedures and [Amdt. 25–32, 37 FR 3972, Feb. 24, 1972] limitations that apply.

(4) Servicing information that covers de- E DITORIAL N OTE : For F EDERAL R EGISTER ci- tails regarding servicing points, capacities of tations affecting appendix F to Part 25, see tanks, reservoirs, types of fluids to be used, the List of CFR Sections Affected, which ap- pressures applicable to the various systems, pears in the Finding Aids section of the location of access panels for inspection and printed volume and at www.govinfo.gov.

servicing, locations of lubrication points, lu- bricants to be used, equipment required for servicing, tow instructions and limitations, mooring, jacking, and leveling information.

A PPENDIX H TO P ART 25—I NSTRUCTIONS (b) Maintenance instructions. (1) Scheduling information for each part of the airplane and FOR C ONTINUED A IRWORTHINESS its engines, auxiliary power units, propellers, H25.1 General. accessories, instruments, and equipment that provides the recommended periods at (a) This appendix specifies requirements which they should be cleaned, inspected, ad- for preparation of Instructions for Continued justed, tested, and lubricated, and the degree Airworthiness as required by §§ 25.1529, of inspection, the applicable wear tolerances, 25.1729, and applicable provisions of parts 21 and work recommended at these periods.

and 26 of this chapter.

However, the applicant may refer to an ac- (b) The Instructions for Continued Air- cessory, instrument, or equipment manufac- worthiness for each airplane must include turer as the source of this information if the the Instructions for Continued Airworthiness applicant shows that the item has an excep- for each engine and propeller (hereinafter tionally high degree of complexity requiring designated ‘‘products’’), for each appliance specialized maintenance techniques, test required by this chapter, and any required equipment, or expertise. The recommended information relating to the interface of overhaul periods and necessary cross ref- those appliances and products with the air- erences to the Airworthiness Limitations plane. If Instructions for Continued Air- section of the manual must also be included.

worthiness are not supplied by the manufac- In addition, the applicant must include an turer of an appliance or product installed in inspection program that includes the fre- the airplane, the Instructions for Continued quency and extent of the inspections nec- Airworthiness for the airplane must include essary to provide for the continued air- the information essential to the continued worthiness of the airplane.

airworthiness of the airplane.

(2) Troubleshooting information describing (c) The applicant must submit to the FAA probable malfunctions, how to recognize a program to show how changes to the In- those malfunctions, and the remedial action structions for Continued Airworthiness made for those malfunctions.

by the applicant or by the manufacturers or (3) Information describing the order and products and appliances installed in the air- method of removing and replacing products plane will be distributed.

and parts with any necessary precautions to be taken.

H25.2 Format.

(4) Other general procedural instructions (a) The Instructions for Continued Air- including procedures for system testing dur- worthiness must be in the form of a manual ing ground running, symmetry checks, or manuals as appropriate for the quantity weighing and determining the center of grav- of data to be provided.

ity, lifting and shoring, and storage limita- (b) The format of the manual or manuals tions.

must provide for a practical arrangement.

(c) Diagrams of structural access plates H25.3 Content. and information needed to gain access for in- The contents of the manual or manuals spections when access plates are not pro- must be prepared in the English language. vided.

The Instructions for Continued Airworthi- (d) Details for the application of special in- ness must contain the following manuals or spection techniques including radiographic sections, as appropriate, and information: and ultrasonic testing where such processes are specified.

(a) Airplane maintenance manual or section.

(e) Information needed to apply protective (1) Introduction information that includes an treatments to the structure after inspection.

explanation of the airplane’s features and (f) All data relative to structural fasteners data to the extent necessary for mainte- such as identification, discard recommenda- nance or preventive maintenance.

tions, and torque values.

(2) A description of the airplane and its (g) A list of special tools needed.

systems and installations including its en- gines, propellers, and appliances. H25.4 Airworthiness Limitations section.

Federal Aviation Administration, DOT Pt. 25, App. I

(a) The Instructions for Continued Air- (iv) Identification of each zone in which worthiness must contain a section titled Air- EWIS is in close proximity to both primary worthiness Limitations that is segregated and back-up hydraulic, mechanical, or elec- and clearly distinguishable from the rest of trical flight controls and lines.

the document. This section must set forth— (v) Identification of— (1) Each mandatory modification time, re- (A) Tasks, and the intervals for performing placement time, structural inspection inter- those tasks, that will reduce the likelihood val, and related structural inspection proce- of ignition sources and accumulation of com- dure approved under § 25.571.

bustible material, and (2) Each mandatory replacement time, in- (B) Procedures, and the intervals for per- spection interval, related inspection proce- forming those procedures, that will effec- dure, and all critical design configuration tively clean the EWIS components of com- control limitations approved under § 25.981 bustible material if there is not an effective for the fuel tank system.

task to reduce the likelihood of combustible (3) Any mandatory replacement time of material accumulation.

EWIS components as defined in section (vi) Instructions for protections and cau- 25.1701.

tion information that will minimize con- (4) A limit of validity of the engineering tamination and accidental damage to EWIS, data that supports the structural mainte- as applicable, during performance of mainte- nance program (LOV), stated as a total num- nance, alteration, or repairs.

ber of accumulated flight cycles or flight (2) Acceptable EWIS maintenance prac- hours or both, approved under § 25.571. Until tices in a standard format.

the full-scale fatigue testing is completed (3) Wire separation requirements as deter- and the FAA has approved the LOV, the mined under § 25.1707.

number of cycles accumulated by the air- (4) Information explaining the EWIS iden- plane cannot be greater than ⁄ 2 the number tification method and requirements for iden- of cycles accumulated on the fatigue test ar- tifying any changes to EWIS under § 25.1711.

ticle.

(5) Electrical load data and instructions for (5) Each mandatory replacement time, in- updating that data.

spection interval, and related inspection and (b) The EWIS ICA developed in accordance test procedure, and each critical design con- with the requirements of H25.5(a)(1) must be figuration control limitation for each light- in the form of a document appropriate for ning protection feature approved under the information to be provided, and they § 25.954.

must be easily recognizable as EWIS ICA.

(6) Each certification maintenance require- This document must either contain the re- ment established to comply with any of the quired EWIS ICA or specifically reference applicable provisions of part 25.

other portions of the ICA that contain this (b) If the Instructions for Continued Air- information.

worthiness consist of multiple documents, the section required by this paragraph must [Amdt. 25–54, 45 FR 60177, Sept. 11, 1980, as be included in the principal manual. This amended by Amdt. 25–68, 54 FR 34329, Aug. 18, section must contain a legible statement in 1989; Amdt. 25–102, 66 FR 23130, May 7, 2001; a prominent location that reads: ‘‘The Air- Amdt. 25–123, 72 FR 63408, Nov. 8, 2007; Amdt.

worthiness Limitations section is FAA-ap- 25–132, 75 FR 69782, Nov. 15, 2010; Doc. No.

proved and specifies maintenance required FAA–2014–1027, Amdt. No. 25–146, 83 FR 47557, under §§ 43.16 and 91.403 of the Federal Avia- Sept. 20, 2018; Doc. No. FAA–2022–1544, 89 FR tion Regulations, unless an alternative pro- 68735, Aug. 27, 2024] gram has been FAA approved.’’ H25.5 Electrical Wiring Interconnection Sys- A PPENDIX I TO P ART 25—I NSTALLATION tem (EWIS) Instructions for Continued Air- OF AN A UTOMATIC T AKEOFF T HRUST worthiness.

C ONTROL S YSTEM (ATTCS) (a) The applicant must prepare Instruc- tions for Continued Airworthiness (ICA) ap- I25.1 General.

plicable to EWIS as defined by § 25.1701 that (a) This appendix specifies additional re- are approved by the FAA and include the fol- quirements for installation of an engine lowing: power control system that automatically (1) Maintenance and inspection require- resets thrust or power on operating engine(s) ments for the EWIS developed with the use in the event of any one engine failure during of an enhanced zonal analysis procedure that takeoff.

includes: (b) With the ATTCS and associated sys- (i) Identification of each zone of the air- tems functioning normally as designed, all plane.

applicable requirements of Part 25, except as (ii) Identification of each zone that con- provided in this appendix, must be met with- tains EWIS.

out requiring any action by the crew to in- (iii) Identification of each zone containing crease thrust or power.

EWIS that also contains combustible mate- rials. I25.2 Definitions.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. I

(a) Automatic Takeoff Thrust Control System (b) Critical Time Interval. When conducting (ATTCS). An ATTCS is defined as the entire an ATTCS takeoff, the critical time interval is between V minus 1 second and a point on automatic system used on takeoff, including 1 the minimum performance, all-engine flight all devices, both mechanical and electrical, path where, assuming a simultaneous occur- that sense engine failure, transmit signals, rence of an engine and ATTCS failure, the actuate fuel controls or power levers or in- resulting minimum flight path thereafter crease engine power by other means on oper- intersects the Part 25 required actual flight ating engines to achieve scheduled thrust or path at no less than 400 feet above the take- power increases, and furnish cockpit infor- off surface. This time interval is shown in mation on system operation.

the following illustration: I25.3 Performance and System Reliability Re- (1) Shall not prevent the insertion of the quirements. maximum approved takeoff thrust or power, or must be shown to be an improbable event.

The applicant must comply with the per- (2) Shall not result in a significant loss or formance and ATTCS reliability require- reduction in thrust or power, or must be ments as follows: shown to be an extremely improbable event.

(a) An ATTCS failure or a combination of failures in the ATTCS during the critical time interval:

Federal Aviation Administration, DOT Pt. 25, App. J

(b) The concurrent existence of an ATTCS In addition to the requirements of § 25.1305: failure and an engine failure during the crit- (a) A means must be provided to indicate ical time interval must be shown to be ex- when the ATTCS is in the armed or ready tremely improbable. condition; and (c) All applicable performance require- (b) If the inherent flight characteristics of ments of Part 25 must be met with an engine the airplane do not provide adequate warn- failure occurring at the most critical point ing that an engine has failed, a warning sys- during takeoff with the ATTCS system func- tem that is independent of the ATTCS must tioning. be provided to give the pilot a clear warning of any engine failure during takeoff.

I25.4 Thrust Setting.

The initial takeoff thrust or power setting [Amdt. 25–62, 52 FR 43156, Nov. 9, 1987] on each engine at the beginning of the take- off roll may not be less than any of the fol- A PPENDIX J TO P ART 25—E MERGENCY lowing: E VACUATION (a) Ninety (90) percent of the thrust or The following test criteria and procedures power set by the ATTCS (the maximum must be used for showing compliance with takeoff thrust or power approved for the air- § 25.803: plane under existing ambient conditions); (a) The emergency evacuation must be con- (b) That required to permit normal oper- ducted with exterior ambient light levels of ation of all safety-related systems and equip- no greater than 0.3 foot-candles prior to the ment dependent upon engine thrust or power activation of the airplane emergency light- lever position; or ing system. The source(s) of the initial exte- (c) That shown to be free of hazardous en- rior ambient light level may remain active gine response characteristics when thrust or or illuminated during the actual demonstra- power is advanced from the initial takeoff tion. There must, however, be no increase in thrust or power to the maximum approved the exterior ambient light level except for takeoff thrust or power.

that due to activation of the airplane emer- I25.5 Powerplant Controls.

gency lighting system.

(a) In addition to the requirements of (b) The airplane must be in a normal atti- § 25.1141, no single failure or malfunction, or tude with landing gear extended.

probable combination thereof, of the ATTCS, (c) Unless the airplane is equipped with an including associated systems, may cause the off-wing descent means, stands or ramps may failure of any powerplant function necessary be used for descent from the wing to the for safety.

ground. Safety equipment such as mats or (b) The ATTCS must be designed to: inverted life rafts may be placed on the floor (1) Apply thrust or power on the operating or ground to protect participants. No other engine(s), following any one engine failure equipment that is not part of the emergency during takeoff, to achieve the maximum ap- evacuation equipment of the airplane may be proved takeoff thrust or power without ex- used to aid the participants in reaching the ceeding engine operating limits; ground.

(2) Permit manual decrease or increase in (d) Except as provided in paragraph (a) of thrust or power up to the maximum takeoff this appendix, only the airplane’s emergency thrust or power approved for the airplane lighting system may provide illumination.

under existing conditions through the use of (e) All emergency equipment required for the power lever. For airplanes equipped with the planned operation of the airplane must limiters that automatically prevent engine be installed.

operating limits from being exceeded under (f) Each internal door or curtain must be existing ambient conditions, other means in the takeoff configuration.

may be used to increase the thrust or power (g) Each crewmember must be seated in in the event of an ATTCS failure provided the normally assigned seat for takeoff and the means is located on or forward of the must remain in the seat until receiving the power levers; is easily identified and oper- signal for commencement of the demonstra- ated under all operating conditions by a sin- tion. Each crewmember must be a person gle action of either pilot with the hand that having knowledge of the operation of exits is normally used to actuate the power levers; and emergency equipment and, if compliance and meets the requirements of § 25.777 (a), with § 121.291 is also being demonstrated, (b), and (c); each flight attendant must be a member of a (3) Provide a means to verify to the regularly scheduled line crew.

flightcrew before takeoff that the ATTCS is (h) A representative passenger load of per- in a condition to operate; and sons in normal health must be used as fol- (4) Provide a means for the flightcrew to lows: deactivate the automatic function. This (1) At least 40 percent of the passenger load means must be designed to prevent inad- must be female.

vertent deactivation.

(2) At least 35 percent of the passenger load I25.6 Powerplant Instruments. must be over 50 years of age.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. K

(3) At least 15 percent of the passenger load Administrator. At least one floor level exit must be female and over 50 years of age. must be used.

(4) Three life-size dolls, not included as (q) Except as provided in paragraph (c) of part of the total passenger load, must be car- this section, all evacuees must leave the air- ried by passengers to simulate live infants 2 plane by a means provided as part of the air- years old or younger. plane’s equipment.

(5) Crewmembers, mechanics, and training (r) The applicant’s approved procedures personnel, who maintain or operate the air- must be fully utilized, except the flightcrew plane in the normal course of their duties, must take no active role in assisting others may not be used as passengers. inside the cabin during the demonstration.

(i) No passenger may be assigned a specific (s) The evacuation time period is com- seat except as the Administrator may re- pleted when the last occupant has evacuated quire. Except as required by subparagraph the airplane and is on the ground. Provided (g) of this paragraph, no employee of the ap- that the acceptance rate of the stand or plicant may be seated next to an emergency ramp is no greater than the acceptance rate exit. of the means available on the airplane for de- (j) Seat belts and shoulder harnesses (as re- scent from the wing during an actual crash quired) must be fastened. situation, evacuees using stands or ramps al- (k) Before the start of the demonstration, lowed by paragraph (c) of this appendix are approximately one-half of the total average considered to be on the ground when they are amount of carry-on baggage, blankets, pil- on the stand or ramp.

lows, and other similar articles must be dis- [Amdt. 25–72, 55 FR 29788, July 20, 1990, as tributed at several locations in aisles and amended by Amdt. 25–79, Aug. 26, 1993; Amdt.

emergency exit access ways to create minor 25–117, 69 FR 67499, Nov. 17, 2004] obstructions.

(l) No prior indication may be given to any A PPENDIX K TO P ART 25—E XTENDED crewmember or passenger of the particular O PERATIONS (ETOPS) exits to be used in the demonstration.

(m) The applicant may not practice, re- This appendix specifies airworthiness re- hearse, or describe the demonstration for the quirements for the approval of an airplane- participants nor may any participant have engine combination for extended operations taken part in this type of demonstration (ETOPS). For two-engine airplanes, the ap- within the preceding 6 months.

plicant must comply with sections K25.1 and (n) Prior to entering the demonstration K25.2 of this appendix. For airplanes with aircraft, the passengers may also be advised more than two engines, the applicant must to follow directions of crewmembers but may comply with sections K25.1 and K25.3 of this not be instructed on the procedures to be fol- appendix.

lowed in the demonstration, except with re- K25.1 Design requirements.

spect to safety procedures in place for the K25.1.1 Part 25 compliance.

demonstration or which have to do with the The airplane-engine combination must demonstration site. Prior to the start of the comply with the requirements of part 25 con- demonstration, the pre-takeoff passenger sidering the maximum flight time and the briefing required by § 121.571 may be given.

longest diversion time for which the appli- Flight attendants may assign demonstration cant seeks approval.

subjects to assist persons from the bottom of K25.1.2 Human factors.

a slide, consistent with their approved train- An applicant must consider crew workload, ing program.

operational implications, and the crew’s and (o) The airplane must be configured to pre- passengers’ physiological needs during con- vent disclosure of the active emergency exits tinued operation with failure effects for the to demonstration participants in the air- longest diversion time for which it seeks ap- plane until the start of the demonstration.

proval.

(p) Exits used in the demonstration must K25.1.3 Airplane systems.

consist of one exit from each exit pair. The (a) Operation in icing conditions.

demonstration may be conducted with the (1) The airplane must be certificated for escape slides, if provided, inflated and the operation in icing conditions in accordance exits open at the beginning of the dem- with § 25.1419.

onstration. In this case, all exits must be (2) The airplane must be able to safely con- configured such that the active exits are not duct an ETOPS diversion with the most crit- disclosed to the occupants. If this method is ical ice accretion resulting from: used, the exit preparation time for each exit (i) Icing conditions encountered at an alti- utilized must be accounted for, and exits tude that the airplane would have to fly fol- that are not to be used in the demonstration lowing an engine failure or cabin decompres- must not be indicated before the demonstra- sion.

tion has started. The exits to be used must (ii) A 15-minute hold in the continuous be representative of all of the emergency maximum icing conditions specified in Ap- exits on the airplane and must be designated pendix C of this part with a liquid water con- by the applicant, subject to approval by the tent factor of 1.0.

Federal Aviation Administration, DOT Pt. 25, App. K

(iii) Ice accumulated during approach and (2) If it is necessary that the APU be able landing in the icing conditions specified in to start in flight, it is able to start at any al- Appendix C of this part. titude up to the maximum operating altitude (b) Electrical power supply. The airplane of the airplane, or 45,000 feet, whichever is must be equipped with at least three inde- lower, and run for the remainder of any pendent sources of electrical power. flight .

(c) Time limited systems. The applicant must (c) Engine oil tank design. The engine oil define the system time capability of each tank filler cap must comply with § 33.71(c)(4) ETOPS significant system that is time-lim- of this chapter.

ited. K25.1.5 Engine-condition monitoring.

K25.1.4 Propulsion systems. Procedures for engine-condition moni- (a) Fuel system design. Fuel necessary to toring must be specified and validated in ac- complete an ETOPS flight (including a diver- cordance with Part 33, Appendix A, para- sion for the longest time for which the appli- graph A33.3(c) of this chapter.

cant seeks approval) must be available to the K25.1.6 Configuration, maintenance, and operating engines at the pressure and fuel- procedures.

flow required by § 25.955 under any airplane The applicant must list any configuration, failure condition not shown to be extremely operating and maintenance requirements, improbable. Types of failures that must be hardware life limits, MMEL constraints, and considered include, but are not limited to: ETOPS approval in a CMP document.

crossfeed valve failures, automatic fuel man- K25.1.7 Airplane flight manual.

agement system failures, and normal elec- The airplane flight manual must contain trical power generation failures. the following information applicable to the (1) If the engine has been certified for lim- ETOPS type design approval: ited operation with negative engine-fuel- (a) Special limitations, including any limi- pump-inlet pressures, the following require- tation associated with operation of the air- ments apply: plane up to the maximum diversion time (i) Airplane demonstration-testing must being approved.

cover worst case cruise and diversion condi- (b) Required markings or placards.

tions involving: (c) The airborne equipment required for ex- (A) Fuel grade and temperature.

tended operations and flightcrew operating (B) Thrust or power variations. procedures for this equipment.

(C) Turbulence and negative G. (d) The system time capability for the fol- (D) Fuel system components degraded lowing: within their approved maintenance limits. (1) The most limiting fire suppression sys- (ii) Unusable-fuel quantity in the suction tem for Class C cargo or baggage compart- feed configuration must be determined in ac- ments.

cordance with § 25.959.

(2) The most limiting ETOPS significant (2) For two-engine airplanes to be certifi- system other than fire suppression systems cated for ETOPS beyond 180 minutes, one for Class C cargo or baggage compartments.

fuel boost pump in each main tank and at (e) This statement: ‘‘The type-design reli- least one crossfeed valve, or other means for ability and performance of this airplane-en- transferring fuel, must be powered by an gine combination has been evaluated under independent electrical power source other 14 CFR 25.1535 and found suitable for (iden- than the three power sources required to tify maximum approved diversion time) ex- comply with section K25.1.3(b) of this appen- tended operations (ETOPS) when the con- dix. This requirement does not apply if the figuration, maintenance, and procedures normal fuel boost pressure, crossfeed valve standard contained in (identify the CMP doc- actuation, or fuel transfer capability is not ument) are met. The actual maximum ap- provided by electrical power.

proved diversion time for this airplane may (3) An alert must be displayed to the be less based on its most limiting system flightcrew when the quantity of fuel avail- time capability. This finding does not con- able to the engines falls below the level re- stitute operational approval to conduct quired to fly to the destination. The alert ETOPS.’’ must be given when there is enough fuel re- K25.2. Two-engine airplanes.

maining to safely complete a diversion. This An applicant for ETOPS type design ap- alert must account for abnormal fuel man- proval of a two-engine airplane must use one agement or transfer between tanks, and pos- of the methods described in section K25.2.1, sible loss of fuel. This paragraph does not K25.2.2, or K25.2.3 of this appendix.

apply to airplanes with a required flight en- K25.2.1 Service experience method.

gineer. An applicant for ETOPS type design ap- (b) APU design. If an APU is needed to com- proval using the service experience method ply with this appendix, the applicant must must comply with sections K25.2.1(a) and demonstrate that: K25.2.1(b) of this appendix before conducting (1) The reliability of the APU is adequate the assessments specified in sections to meet those requirements; and K25.2.1(c) and K25.2.1(d) of this appendix, and

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. K

the flight test specified in section K25.2.1(e) (B) Engine configuration, and major alter- of this appendix. ation history; (a) Service experience. The world fleet for (C) Engine position; the airplane-engine combination must accu- (D) Circumstances leading up to the engine mulate a minimum of 250,000 engine-hours.

shutdown or occurrence; The FAA may reduce this number of hours if (E) Phase of flight or ground operation; the applicant identifies compensating fac- (F) Weather and other environmental con- tors that are acceptable to the FAA. The ditions; and compensating factors may include experi- (G) Cause of engine shutdown or occur- ence on another airplane, but experience on rence.

the candidate airplane must make up a sig- (ii) A history of unscheduled engine re- nificant portion of the total service experi- moval rates since introduction into service ence.

(using 6- and 12-month rolling averages), (b) In-flight shutdown (IFSD) rates. The with a summary of the major causes for the demonstrated 12-month rolling average IFSD removals.

rate for the world fleet of the airplane-en- (iii) A list of all propulsion system events gine combination must be commensurate (whether or not caused by maintenance or with the level of ETOPS approval being flightcrew error), including dispatch delays, sought.

cancellations, aborted takeoffs, turnbacks, (1) For type design approval up to and in- diversions, and flights that continue to des- cluding 120 minutes: An IFSD rate of 0.05 or tination after the event.

less per 1,000 world-fleet engine-hours, unless otherwise approved by the FAA. Unless the (iv) The total number of engine hours and IFSD rate is 0.02 or less per 1,000 world-fleet cycles, the number of hours for the engine engine-hours, the applicant must provide a with the highest number of hours, the num- list of corrective actions in the CMP docu- ber of cycles for the engine with the highest ment specified in section K25.1.6 of this ap- number of cycles, and the distribution of pendix, that, when taken, would result in an hours and cycles.

IFSD rate of 0.02 or less per 1,000 fleet en- (v) The mean time between failures gine-hours.

(MTBF) of propulsion system components (2) For type design approval up to and in- that affect reliability.

cluding 180 minutes: An IFSD rate of 0.02 or (vi) A history of the IFSD rates since in- less per 1,000 world-fleet engine-hours, unless troduction into service using a 12-month otherwise approved by the FAA. If the air- rolling average.

plane-engine combination does not meet this (2) The cause or potential cause of each rate by compliance with an existing 120- item listed in K25.2.1(c)(1)(i) must have a cor- minute CMP document, then new or addi- rective action or actions that are shown to tional CMP requirements that the applicant be effective in preventing future occur- has demonstrated would achieve this IFSD rences. Each corrective action must be iden- rate must be added to the CMP document.

tified in the CMP document specified in sec- (3) For type design approval beyond 180 tion K25.1.6. A corrective action is not re- minutes: An IFSD rate of 0.01 or less per 1,000 quired: fleet engine-hours unless otherwise approved (i) For an item where the manufacturer is by the FAA. If the airplane-engine combina- unable to determine a cause or potential tion does not meet this rate by compliance cause.

with an existing 120-minute or 180-minute (ii) For an event where it is technically CMP document, then new or additional CMP unfeasible to develop a corrective action.

requirements that the applicant has dem- (iii) If the world-fleet IFSD rate— onstrated would achieve this IFSD rate must (A) Is at or below 0.02 per 1,000 world-fleet be added to the CMP document.

engine-hours for approval up to and includ- (c) Propulsion system assessment. (1) The ap- ing 180-minute ETOPS; or plicant must conduct a propulsion system (B) Is at or below 0.01 per 1,000 world-fleet assessment based on the following data col- engine-hours for approval greater than 180- lected from the world-fleet of the airplane- minute ETOPS.

engine combination: (d) Airplane systems assessment. The appli- (i) A list of all IFSD’s, unplanned ground cant must conduct an airplane systems as- engine shutdowns, and occurrences (both sessment. The applicant must show that the ground and in-flight) when an engine was not airplane systems comply with § 25.1309(b) shut down, but engine control or the desired using available in-service reliability data for thrust or power level was not achieved, in- cluding engine flameouts. Planned IFSD’s ETOPS significant systems on the candidate performed during flight training need not be airplane-engine combination. Each cause or included. For each item, the applicant must potential cause of a relevant design, manu- provide— facturing, operational, and maintenance (A) Each airplane and engine make, model, problem occurring in service must have a and serial number; corrective action or actions that are shown

Federal Aviation Administration, DOT Pt. 25, App. K

to be effective in preventing future occur- (i) An IFSD rate of 0.02 or less per 1,000 rences. Each corrective action must be iden- world-fleet engine-hours for type design ap- tified in the CMP document specified in sec- proval up to and including 180 minutes.

(ii) An IFSD rate of 0.01 or less per 1,000 tion K25.1.6 of this appendix. A corrective ac- world-fleet engine-hours for type design ap- tion is not required if the problem would not proval beyond 180 minutes.

significantly impact the safety or reliability (c) Maintenance and operational procedures.

of the airplane system involved. A relevant The applicant must validate all maintenance problem is a problem with an ETOPS group and operational procedures for ETOPS sig- 1 significant system that has or could result nificant systems. The applicant must iden- in, an IFSD or diversion. The applicant must tify, track, and resolve any problems found include in this assessment relevant problems during the validation in accordance with the with similar or identical equipment installed problem tracking and resolution system on other types of airplanes to the extent specified in section K25.2.2(h) of this appen- such information is reasonably available.

dix.

(e) Airplane flight test. The applicant must (d) Propulsion system validation test. (1) The conduct a flight test to validate the installed engine configuration for which ap- flightcrew’s ability to safely conduct an proval is being sought must comply with ETOPS diversion with an inoperative engine § 33.201(c) of this chapter. The test engine and worst-case ETOPS Significant System must be configured with a complete airplane failures and malfunctions that could occur in nacelle package, including engine-mounted service. The flight test must validate the air- equipment, except for any configuration dif- plane’s flying qualities and performance ferences necessary to accommodate test with the demonstrated failures and malfunc- stand interfaces with the engine nacelle tions.

package. At the conclusion of the test, the K25.2.2 Early ETOPS method.

propulsion system must be— An applicant for ETOPS type design ap- (i) Visually inspected according to the ap- proval using the Early ETOPS method must plicant’s on-wing inspection recommenda- comply with the following requirements: tions and limits; and (ii) Completely disassembled and the pro- (a) Assessment of relevant experience with air- pulsion system hardware inspected to deter- planes previously certificated under part 25.

mine whether it meets the service limits The applicant must identify specific correc- specified in the Instructions for Continued tive actions taken on the candidate airplane Airworthiness submitted in compliance with to prevent relevant design, manufacturing, § 25.1529.

operational, and maintenance problems ex- (2) The applicant must identify, track, and perienced on airplanes previously certifi- resolve each cause or potential cause of cated under part 25 manufactured by the ap- IFSD, loss of thrust control, or other power plicant. Specific corrective actions are not loss encountered during this inspection in required if the nature of a problem is such accordance with the problem tracking and that the problem would not significantly im- resolution system specified in section K25.2.2 pact the safety or reliability of the airplane (h) of this appendix.

system involved. A relevant problem is a (e) New technology testing. Technology new problem with an ETOPS group 1 significant to the applicant, including substantially new system that has or could result in an IFSD manufacturing techniques, must be tested to or diversion. The applicant must include in substantiate its suitability for the airplane this assessment relevant problems of sup- design.

plier-provided ETOPS group 1 significant (f) APU validation test. If an APU is needed systems and similar or identical equipment to comply with this appendix, one APU of used on airplanes built by other manufactur- the type to be certified with the airplane ers to the extent such information is reason- must be tested for 3,000 equivalent airplane ably available.

operational cycles. Following completion of (b) Propulsion system design. (1) The engine the test, the APU must be disassembled and used in the applicant’s airplane design must inspected. The applicant must identify, be approved as eligible for Early ETOPS in track, and resolve each cause or potential accordance with § 33.201 of this chapter.

cause of an inability to start or operate the (2) The applicant must design the propul- APU in flight as intended in accordance with sion system to preclude failures or malfunc- the problem tracking and resolution system tions that could result in an IFSD. The ap- specified in section K25.2.2(h) of this appen- plicant must show compliance with this re- dix.

quirement by analysis, test, in-service expe- (g) Airplane demonstration. For each air- rience on other airplanes, or other means ac- plane-engine combination to be approved for ceptable to the FAA. If analysis is used, the ETOPS, the applicant must flight test at applicant must show that the propulsion sys- least one airplane to demonstrate that the tem design will minimize failures and mal- airplane, and its components and equipment functions with the objective of achieving the are capable of functioning properly during following IFSD rates: ETOPS flights and diversions of the longest

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. K

duration for which the applicant seeks ap- (4) At the completion of the airplane dem- proval. This flight testing may be performed onstration flight test program, each ETOPS in conjunction with, but may not substitute significant system must undergo an on-wing for the flight testing required by § 21.35(b)(2) inspection or test in accordance with the of this chapter. tasks defined in the proposed Instructions (1) The airplane demonstration flight test for Continued Airworthiness to establish its program must include: condition for continued safe operation. Each (i) Flights simulating actual ETOPS, in- engine must also undergo a gas path inspec- cluding flight at normal cruise altitude, step tion. These inspections must be conducted in climbs, and, if applicable, APU operation.

a manner to identify abnormal conditions (ii) Maximum duration flights with max- that could result in an IFSD or diversion.

imum duration diversions.

The applicant must identify, track and re- (iii) Maximum duration engine-inoperative solve any abnormal conditions in accordance diversions distributed among the engines in- with the problem tracking and resolution stalled on the airplanes used for the airplane system specified in section K25.2.2(h) of this demonstration flight test program. At least appendix.

two one-engine-inoperative diversions must (h) Problem tracking and resolution system.

be conducted at maximum continuous thrust (1) The applicant must establish and main- or power using the same engine.

tain a problem tracking and resolution sys- (iv) Flights under non-normal conditions tem. The system must: to demonstrate the flightcrew’s ability to (i) Contain a process for prompt reporting safely conduct an ETOPS diversion with to the FAA office responsible for the design worst-case ETOPS significant system fail- approval of each occurrence reportable under ures or malfunctions that could occur in § 21.4(a)(6) encountered during the phases of service.

airplane and engine development used to as- (v) Diversions to airports that represent sess Early ETOPS eligibility.

airports of the types used for ETOPS diver- (ii) Contain a process for notifying the sions.

FAA office responsible for the design ap- (vi) Repeated exposure to humid and in- proval of each proposed corrective action clement weather on the ground followed by a that the applicant determines necessary for long-duration flight at normal cruise alti- each problem identified from the occurrences tude.

reported under section K25.2.2. (h)(1)(i) of (2) The airplane demonstration flight test this appendix. The timing of the notification program must validate the adequacy of the must permit appropriate FAA review before airplane’s flying qualities and performance, taking the proposed corrective action.

and the flightcrew’s ability to safely conduct an ETOPS diversion under the conditions (2) If the applicant is seeking ETOPS type specified in section K25.2.2(g)(1) of this ap- design approval of a change to an airplane- pendix. engine combination previously approved for (3) During the airplane demonstration ETOPS, the problem tracking and resolution flight test program, each test airplane must system need only address those problems be operated and maintained using the appli- specified in the following table, provided the cant’s recommended operating and mainte- applicant obtains prior authorization from nance procedures. the FAA: If the change does not require a new airplane type certificiate Then the Problem Tracking and Resolution System must ad- and . . . dress . . .

(i) Requires a new engine type certificate .................................. All problems applicable to the new engine installation, and for the remainder of the airplane, problems in changed systems only.

(ii) Does not require a new engine type certificate ..................... Problems in changed systems only.

(i) Acceptance criteria. The type and fre- An applicant for ETOPS type design ap- quency of failures and malfunctions on proval using the combined service experience ETOPS significant systems that occur dur- and Early ETOPS method must comply with ing the airplane flight test program and the the following requirements.

airplane demonstration flight test program (a) A service experience requirement of not specified in section K25.2.2(g) of this appen- less than 15,000 engine-hours for the world dix must be consistent with the type and fre- fleet of the candidate airplane-engine com- quency of failures and malfunctions that bination.

would be expected to occur on currently cer- (b) The Early ETOPS requirements of tificated airplanes approved for ETOPS. K25.2.2, except for the airplane demonstra- K25.2.3. Combined service experience and tion specified in section K25.2.2(g) of this ap- Early ETOPS method. pendix; and

Federal Aviation Administration, DOT Pt. 25, App. K

(c) The flight test requirement of section (a) Maintenance and operational procedures.

K25.2.1(e) of this appendix. The applicant must validate all maintenance K25.3. Airplanes with more than two engines. and operational procedures for ETOPS sig- nificant systems. The applicant must iden- An applicant for ETOPS type design ap- tify, track and resolve any problems found proval of an airplane with more than two en- during the validation in accordance with the gines must use one of the methods described problem tracking and resolution system in section K25.3.1, K25.3.2, or K25.3.3 of this specified in section K25.3.2(e) of this appen- appendix.

dix.

K25.3.1 Service experience method.

(b) New technology testing. Technology new An applicant for ETOPS type design ap- to the applicant, including substantially new proval using the service experience method manufacturing techniques, must be tested to must comply with section K25.3.1(a) of this substantiate its suitability for the airplane appendix before conducting the airplane sys- design.

tems assessment specified in K25.3.1(b), and (c) APU validation test. If an APU is needed the flight test specified in section K25.3.1(c) to comply with this appendix, one APU of of this appendix.

the type to be certified with the airplane (a) Service experience. The world fleet for must be tested for 3,000 equivalent airplane the airplane-engine combination must accu- operational cycles. Following completion of mulate a minimum of 250,000 engine-hours.

the test, the APU must be disassembled and The FAA may reduce this number of hours if inspected. The applicant must identify, the applicant identifies compensating fac- track, and resolve each cause or potential tors that are acceptable to the FAA. The cause of an inability to start or operate the compensating factors may include experi- APU in flight as intended in accordance with ence on another airplane, but experience on the problem tracking and resolution system the candidate airplane must make up a sig- specified in section K25.3.2(e) of this appen- nificant portion of the total required service dix.

experience.

(d) Airplane demonstration. For each air- (b) Airplane systems assessment. The appli- plane-engine combination to be approved for cant must conduct an airplane systems as- ETOPS, the applicant must flight test at sessment. The applicant must show that the least one airplane to demonstrate that the airplane systems comply with the § 25.1309(b) airplane, and its components and equipment using available in-service reliability data for are capable of functioning properly during ETOPS significant systems on the candidate ETOPS flights and diversions of the longest airplane-engine combination. Each cause or duration for which the applicant seeks ap- potential cause of a relevant design, manu- proval. This flight testing may be performed facturing, operational or maintenance prob- in conjunction with, but may not substitute lem occurring in service must have a correc- for the flight testing required by § 21.35(b)(2).

tive action or actions that are shown to be (1) The airplane demonstration flight test effective in preventing future occurrences.

program must include: Each corrective action must be identified in (i) Flights simulating actual ETOPS in- the CMP document specified in section cluding flight at normal cruise altitude, step K25.1.6 of this appendix. A corrective action climbs, and, if applicable, APU operation.

is not required if the problem would not sig- (ii) Maximum duration flights with max- nificantly impact the safety or reliability of imum duration diversions.

the airplane system involved. A relevant (iii) Maximum duration engine-inoperative problem is a problem with an ETOPS group diversions distributed among the engines in- 1 significant system that has or could result stalled on the airplanes used for the airplane in an IFSD or diversion. The applicant must demonstration flight test program. At least include in this assessment relevant problems two one engine-inoperative diversions must with similar or identical equipment installed be conducted at maximum continuous thrust on other types of airplanes to the extent or power using the same engine.

such information is reasonably available.

(iv) Flights under non-normal conditions (c) Airplane flight test. The applicant must to validate the flightcrew’s ability to safely conduct a flight test to validate the conduct an ETOPS diversion with worst-case flightcrew’s ability to safely conduct an ETOPS significant system failures or mal- ETOPS diversion with an inoperative engine functions that could occur in service.

and worst-case ETOPS significant system (v) Diversions to airports that represent failures and malfunctions that could occur in airports of the types used for ETOPS diver- service. The flight test must validate the air- sions.

plane’s flying qualities and performance (vi) Repeated exposure to humid and in- with the demonstrated failures and malfunc- clement weather on the ground followed by a tions.

long duration flight at normal cruise alti- K25.3.2 Early ETOPS method. tude.

An applicant for ETOPS type design ap- (2) The airplane demonstration flight test proval using the Early ETOPS method must program must validate the adequacy of the comply with the following requirements: airplane’s flying qualities and performance,

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. L

and the flightcrew’s ability to safely conduct tain a problem tracking and resolution sys- an ETOPS diversion under the conditions tem. The system must: specified in section K25.3.2(d)(1) of this ap- (i) Contain a process for prompt reporting pendix.

to the FAA office responsible for the design (3) During the airplane demonstration approval of each occurrence reportable under flight test program, each test airplane must § 21.4(a)(6) encountered during the phases of be operated and maintained using the appli- airplane and engine development used to as- cant’s recommended operating and mainte- sess Early ETOPS eligibility.

nance procedures.

(ii) Contain a process for notifying the (4) At the completion of the airplane dem- FAA office responsible for the design ap- onstration, each ETOPS significant system proval of each proposed corrective action must undergo an on-wing inspection or test that the applicant determines necessary for in accordance with the tasks defined in the each problem identified from the occurrences proposed Instructions for Continued Air- reported under section K25.3.2(h)(1)(i) of this worthiness to establish its condition for con- appendix. The timing of the notification tinued safe operation. Each engine must also must permit appropriate FAA review before undergo a gas path inspection. These inspec- taking the proposed corrective action.

tions must be conducted in a manner to iden- (2) If the applicant is seeking ETOPS type tify abnormal conditions that could result in design approval of a change to an airplane- an IFSD or diversion. The applicant must engine combination previously approved for identify, track and resolve any abnormal ETOPS, the problem tracking and resolution conditions in accordance with the problem system need only address those problems tracking and resolution system specified in specified in the following table, provided the section K25.3.2(e) of this appendix.

applicant obtains prior authorization from (e) Problem tracking and resolution system.

the FAA: (1) The applicant must establish and main- If the change does not require a new airplane type certificate Then the Problem Tracking and Resolution System must ad- and . . . dress . . .

(i) Requires a new engine type certificate .................................. All problems applicable to the new engine installation, and for the remainder of the airplane, problems in changed systems only.

(ii) Does not require a new engine type certificate ..................... Problems in changed systems only.

(f) Acceptance criteria. The type and fre- A PPENDIX L TO P ART 25—HIRF E NVI - quency of failures and malfunctions on RONMENTS AND E QUIPMENT HIRF ETOPS significant systems that occur dur- T EST L EVELS ing the airplane flight test program and the airplane demonstration flight test program This appendix specifies the HIRF environ- specified in section K25.3.2(d) of this appen- ments and equipment HIRF test levels for electrical and electronic systems under dix must be consistent with the type and fre- § 25.1317. The field strength values for the quency of failures and malfunctions that HIRF environments and equipment HIRF would be expected to occur on currently cer- test levels are expressed in root-mean-square tificated airplanes approved for ETOPS.

units measured during the peak of the modu- K25.3.3 Combined service experience and lation cycle.

Early ETOPS method.

(a) HIRF environment I is specified in the An applicant for ETOPS type design ap- following table: proval using the Early ETOPS method must comply with the following requirements: T ABLE I.—HIRF E NVIRONMENT I (a) A service experience requirement of less than 15,000 engine-hours for the world Field strength (volts/meter) fleet of the candidate airplane-engine com- Frequency bination; Peak Average (b) The Early ETOPS requirements of sec- 10 kHz–2 MHz ................................... 50 50 tion K25.3.2 of this appendix, except for the 2 MHz–30 MHz ................................. 100 100 airplane demonstration specified in section 30 MHz–100 MHz ............................. 50 50 K25.3.2(d) of this appendix; and 100 MHz–400 MHz ........................... 100 100 (c) The flight test requirement of section 400 MHz–700 MHz ........................... 700 50 K25.3.1(c) of this appendix.

700 MHz–1 GHz ................................ 700 100 1 GHz–2 GHz .................................... 2,000 200 [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, 2 GHz–6 GHz .................................... 3,000 200 2007, as amended by Doc. No. FAA–2018–0119, 6 GHz–8 GHz .................................... 1,000 200 Amdt. 25–145, 83 FR 9169, Mar. 5, 2018] 8 GHz–12 GHz .................................. 3,000 300 12 GHz–18 GHz ................................ 2,000 200

Federal Aviation Administration, DOT Pt. 25, App. M

ation curves. Testing must cover the fre- T ABLE I.—HIRF E NVIRONMENT I—Continued quency band of 10 kHz to 8 GHz.

Field strength (e) Equipment HIRF Test Level 3. (1) From 10 (volts/meter) Frequency kHz to 400 MHz, use conducted susceptibility Peak Average tests, starting at a minimum of 0.15 mA at 10 kHz, increasing 20 dB per frequency decade 18 GHz–40 GHz ................................ 600 200 to a minimum of 7.5 mA at 500 kHz.

In this table, the higher field strength applies at the fre- (2) From 500 kHz to 40 MHz, use conducted quency band edges.

susceptibility tests at a minimum of 7.5 mA.

(b) HIRF environment II is specified in the (3) From 40 MHz to 400 MHz, use conducted following table: susceptibility tests, starting at a minimum of 7.5 mA at 40 MHz, decreasing 20 dB per fre- T ABLE II.–HIRF E NVIRONMENT II quency decade to a minimum of 0.75 mA at 400 MHz.

Field strength (4) From 100 MHz to 8 GHz, use radiated (volts/meter) Frequency susceptibility tests at a minimum of 5 V/m.

Peak Average [Doc. No. FAA–2006–23657, 72 FR 44026, Aug. 6, 10 kHz–500 kHz ................................ 20 20 2007] 500 kHz–2 MHz ................................. 30 30 2 MHz–30 MHz ................................. 100 100 A PPENDIX M TO P ART 25—F UEL T ANK 30 MHz–100 MHz ............................. 10 10 100 MHz–200 MHz ........................... 30 10 S YSTEM F LAMMABILITY R EDUCTION 200 MHz–400 MHz ........................... 10 10 M EANS 400 MHz–1 GHz ................................ 700 40 1 GHz–2 GHz .................................... 1,300 160 M25.1 Fuel tank flammability exposure re- 2 GHz–4 GHz .................................... 3,000 120 quirements.

4 GHz–6 GHz .................................... 3,000 160 6 GHz–8 GHz .................................... 400 170 (a) The Fleet Average Flammability Expo- 8 GHz–12 GHz .................................. 1,230 230 sure of each fuel tank, as determined in ac- 12 GHz–18 GHz ................................ 730 190 cordance with Appendix N of this part, may 18 GHz–40 GHz ................................ 600 150 not exceed 3 percent of the Flammability Ex- In this table, the higher field strength applies at the fre- posure Evaluation Time (FEET), as defined quency band edges.

in Appendix N of this part. As a portion of this 3 percent, if flammability reduction (c) Equipment HIRF Test Level 1. (1) From 10 kilohertz (kHz) to 400 megahertz (MHz), use means (FRM) are used, each of the following conducted susceptibility tests with contin- time periods may not exceed 1.8 percent of uous wave (CW) and 1 kHz square wave mod- the FEET: ulation with 90 percent depth or greater. The (1) When any FRM is operational but the conducted susceptibility current must start fuel tank is not inert and the tank is flam- at a minimum of 0.6 milliamperes (mA) at 10 mable; and kHz, increasing 20 decibels (dB) per fre- (2) When any FRM is inoperative and the quency decade to a minimum of 30 mA at 500 tank is flammable.

kHz.

(b) The Fleet Average Flammability Expo- (2) From 500 kHz to 40 MHz, the conducted sure, as defined in Appendix N of this part, of susceptibility current must be at least 30 each fuel tank may not exceed 3 percent of mA.

the portion of the FEET occurring during ei- (3) From 40 MHz to 400 MHz, use conducted ther ground or takeoff/climb phases of flight susceptibility tests, starting at a minimum during warm days. The analysis must con- of 30 mA at 40 MHz, decreasing 20 dB per fre- sider the following conditions.

quency decade to a minimum of 3 mA at 400 (1) The analysis must use the subset of MHz.

those flights that begin with a sea level (4) From 100 MHz to 400 MHz, use radiated ground ambient temperature of 80 ° F (stand- susceptibility tests at a minimum of 20 volts ard day plus 21 ° F atmosphere) or above, per meter (V/m) peak with CW and 1 kHz from the flammability exposure analysis square wave modulation with 90 percent done for overall performance.

depth or greater.

(2) For the ground and takeoff/climb phases (5) From 400 MHz to 8 gigahertz (GHz), use of flight, the average flammability exposure radiated susceptibility tests at a minimum must be calculated by dividing the time dur- of 150 V/m peak with pulse modulation of 4 ing the specific flight phase the fuel tank is percent duty cycle with a 1 kHz pulse repeti- flammable by the total time of the specific tion frequency. This signal must be switched flight phase.

on and off at a rate of 1 Hz with a duty cycle (3) Compliance with this paragraph may be of 50 percent.

(d) Equipment HIRF Test Level 2. Equipment shown using only those flights for which the HIRF test level 2 is HIRF environment II in airplane is dispatched with the flammability table II of this appendix reduced by accept- reduction means operational.

able aircraft transfer function and attenu- M25.2 Showing compliance.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. N

(a) The applicant must provide data from (b) Unless alternative reporting procedures analysis, ground testing, and flight testing, are approved by the responsible Aircraft Cer- or any combination of these, that: tification Service office, as defined in part 26 (1) Validate the parameters used in the of this subchapter, provide a report to the analysis required by paragraph M25.1 of this FAA every six months for the first five years appendix; after service introduction. After that period, continued reporting every six months may (2) Substantiate that the FRM is effective be replaced with other reliability tracking at limiting flammability exposure in all methods found acceptable to the FAA or compartments of each tank for which the eliminated if it is established that the reli- FRM is used to show compliance with para- ability of the FRM meets, and will continue graph M25.1 of this appendix; and to meet, the exposure requirements of para- (3) Describe the circumstances under which graph M25.1 of this appendix.

the FRM would not be operated during each phase of flight. (c) Develop service instructions or revise (b) The applicant must validate that the the applicable airplane manual, according to FRM meets the requirements of paragraph a schedule approved by the responsible Air- M25.1 of this appendix with any airplane or craft Certification Service office, as defined engine configuration affecting the perform- in part 26 of this subchapter, to correct any ance of the FRM for which approval is failures of the FRM that occur in service sought. that could increase any fuel tank’s Fleet Av- M25.3 Reliability indications and mainte- erage Flammability Exposure to more than nance access. that required by paragraph M25.1 of this ap- (a) Reliability indications must be pro- pendix.

vided to identify failures of the FRM that [Doc. No. FAA–2005–22997, 73 FR 42494, July would otherwise be latent and whose identi- 21, 2008, as amended by Doc. No. FAA–2018– fication is necessary to ensure the fuel tank 0119, Amdt. 25–145, 83 FR 9169, Mar. 5, 2018] with an FRM meets the fleet average flam- mability exposure requirements listed in A PPENDIX N TO P ART 25—F UEL T ANK paragraph M25.1 of this appendix, including FLAMMABILITY E XPOSURE AND R ELI- when the FRM is inoperative.

ABILITY A NALYSIS (b) Sufficient accessibility to FRM reli- ability indications must be provided for N25.1 General.

maintenance personnel or the flightcrew.

(a) This appendix specifies the require- (c) The access doors and panels to the fuel ments for conducting fuel tank fleet average tanks with FRMs (including any tanks that flammability exposure analyses required to communicate with a tank via a vent sys- meet § 25.981(b) and Appendix M of this part.

tem), and to any other confined spaces or en- For fuel tanks installed in aluminum wings, closed areas that could contain hazardous at- a qualitative assessment is sufficient if it mosphere under normal conditions or failure substantiates that the tank is a conven- conditions, must be permanently stenciled, tional unheated wing tank.

marked, or placarded to warn maintenance (b) This appendix defines parameters af- personnel of the possible presence of a poten- fecting fuel tank flammability that must be tially hazardous atmosphere.

used in performing the analysis. These in- M25.4 Airworthiness limitations and proce- clude parameters that affect all airplanes dures.

within the fleet, such as a statistical dis- (a) If FRM is used to comply with para- tribution of ambient temperature, fuel flash graph M25.1 of this appendix, Airworthiness point, flight lengths, and airplane descent Limitations must be identified for all main- rate. Demonstration of compliance also re- tenance or inspection tasks required to iden- quires application of factors specific to the tify failures of components within the FRM airplane model being evaluated. Factors that that are needed to meet paragraph M25.1 of need to be included are maximum range, this appendix.

cruise mach number, typical altitude where (b) Maintenance procedures must be devel- the airplane begins initial cruise phase of oped to identify any hazards to be considered flight, fuel temperature during both ground during maintenance of the FRM. These pro- and flight times, and the performance of a cedures must be included in the instructions flammability reduction means (FRM) if in- for continued airworthiness (ICA).

stalled.

M25.5 Reliability reporting.

(c) The following definitions, input vari- The effects of airplane component failures ables, and data tables must be used in the on FRM reliability must be assessed on an program to determine fleet average flamma- on-going basis. The applicant/holder must do bility exposure for a specific airplane model.

the following: N25.2 Definitions.

(a) Demonstrate effective means to ensure collection of FRM reliability data. The (a) Bulk Average Fuel Temperature means means must provide data affecting FRM reli- the average fuel temperature within the fuel ability, such as component failures. tank or different sections of the tank if the

Federal Aviation Administration, DOT Pt. 25, App. N

tank is subdivided by baffles or compart- feet to 14.5 percent at 40,000 feet altitude, ments. and extrapolated linearly above that alti- (b) Flammability Exposure Evaluation Time tude.

(FEET). The time from the start of preparing (i) Inerting. A process where a noncombus- the airplane for flight, through the flight tible gas is introduced into the ullage of a and landing, until all payload is unloaded, fuel tank so that the ullage becomes non- and all passengers and crew have dis- flammable.

embarked. In the Monte Carlo program, the (j) Monte Carlo Analysis. The analytical flight time is randomly selected from the method that is specified in this appendix as Flight Length Distribution (Table 2), the the compliance means for assessing the fleet pre-flight times are provided as a function of average flammability exposure time for a the flight time, and the post-flight time is a fuel tank.

constant 30 minutes.

(k) Oxygen evolution occurs when oxygen (c) Flammable. With respect to a fluid or dissolved in the fuel is released into the gas, flammable means susceptible to igniting ullage as the pressure and temperature in readily or to exploding (14 CFR Part 1, Defi- the fuel tank are reduced.

nitions). A non-flammable ullage is one (l) Standard deviation is a statistical meas- where the fuel-air vapor is too lean or too ure of the dispersion or variation in a dis- rich to burn or is inert as defined below. For tribution, equal to the square root of the the purposes of this appendix, a fuel tank arithmetic mean of the squares of the devi- that is not inert is considered flammable ations from the arithmetic means.

when the bulk average fuel temperature (m) Transport Effects. For purposes of this within the tank is within the flammable appendix, transport effects are the change in range for the fuel type being used. For any fuel vapor concentration in a fuel tank fuel tank that is subdivided into sections by caused by low fuel conditions and fuel con- baffles or compartments, the tank is consid- densation and vaporization.

ered flammable when the bulk average fuel (n) Ullage. The volume within the fuel tank temperature within any section of the tank, not occupied by liquid fuel.

that is not inert, is within the flammable N25.3 Fuel tank flammability exposure anal- range for the fuel type being used.

ysis.

(d) Flash Point. The flash point of a flam- (a) A flammability exposure analysis must mable fluid means the lowest temperature at be conducted for the fuel tank under evalua- which the application of a flame to a heated tion to determine fleet average flammability sample causes the vapor to ignite momen- exposure for the airplane and fuel types tarily, or ‘‘flash.’’ Table 1 of this appendix under evaluation. For fuel tanks that are provides the flash point for the standard fuel subdivided by baffles or compartments, an to be used in the analysis.

analysis must be performed either for each (e) Fleet average flammability exposure is the section of the tank, or for the section of the percentage of the flammability exposure tank having the highest flammability expo- evaluation time (FEET) each fuel tank sure. Consideration of transport effects is ullage is flammable for a fleet of an airplane not allowed in the analysis. The analysis type operating over the range of flight must be done in accordance with the meth- lengths in a world-wide range of environ- ods and procedures set forth in the Fuel mental conditions and fuel properties as de- Tank Flammability Assessment Method fined in this appendix.

User’s Manual, dated May 2008, document (f) Gaussian Distribution is another name number DOT/FAA/AR–05/8 (incorporated by for the normal distribution, a symmetrical reference, see § 25.5). The parameters speci- frequency distribution having a precise fied in sections N25.3(b) and (c) of this appen- mathematical formula relating the mean dix must be used in the fuel tank flamma- and standard deviation of the samples.

bility exposure ‘‘Monte Carlo’’ analysis.

Gaussian distributions yield bell-shaped fre- (b) The following parameters are defined in quency curves having a preponderance of val- the Monte Carlo analysis and provided in ues around the mean with progressively paragraph N25.4 of this appendix: fewer observations as the curve extends out- (1) Cruise Ambient Temperature, as de- ward.

fined in this appendix.

(g) Hazardous atmosphere. An atmosphere (2) Ground Ambient Temperature, as de- that may expose maintenance personnel, fined in this appendix.

passengers or flight crew to the risk of death, incapacitation, impairment of ability (3) Fuel Flash Point, as defined in this ap- to self-rescue (that is, escape unaided from a pendix.

confined space), injury, or acute illness. (4) Flight Length Distribution, as defined (h) Inert. For the purpose of this appendix, in Table 2 of this appendix.

the tank is considered inert when the bulk (5) Airplane Climb and Descent Profiles, as average oxygen concentration within each defined in the Fuel Tank Flammability As- compartment of the tank is 12 percent or less sessment Method User’s Manual, dated May from sea level up to 10,000 feet altitude, then 2008, document number DOT/FAA/AR–05/8 linearly increasing from 12 percent at 10,000 (incorporated by reference in § 25.5).

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. N

(c) Parameters that are specific to the par- patch limit unless an alternative period has ticular airplane model under evaluation that been approved by the Administrator), must be provided as inputs to the Monte (3) Frequency and duration of time periods Carlo analysis are: of FRM inoperability, substantiated by test (1) Airplane cruise altitude. or analysis acceptable to the FAA, caused by (2) Fuel tank quantities. If fuel quantity latent or known failures, including airplane affects fuel tank flammability, inputs to the system shut-downs and failures that could Monte Carlo analysis must be provided that cause the FRM to shut down or become inop- represent the actual fuel quantity within the erative.

fuel tank or compartment of the fuel tank (4) Effects of failures of the FRM that throughout each of the flights being evalu- could increase the flammability exposure of ated. Input values for this data must be ob- the fuel tank.

tained from ground and flight test data or (5) If an FRM is used that is affected by ox- ygen concentrations in the fuel tank, the the approved FAA fuel management proce- dures. time periods when oxygen evolution from the (3) Airplane cruise mach number. fuel results in the fuel tank or compartment (4) Airplane maximum range. exceeding the inert level. The applicant (5) Fuel tank thermal characteristics. If must include any times when oxygen evo- fuel temperature affects fuel tank flamma- lution from the fuel in the tank or compart- bility, inputs to the Monte Carlo analysis ment under evaluation would result in a must be provided that represent the actual flammable fuel tank. The oxygen evolution bulk average fuel temperature within the rate that must be used is defined in the Fuel fuel tank at each point in time throughout Tank Flammability Assessment Method each of the flights being evaluated. For fuel User’s Manual, dated May 2008, document tanks that are subdivided by baffles or com- number DOT/FAA/AR–05/8 (incorporated by partments, bulk average fuel temperature reference in § 25.5).

inputs must be provided for each section of (6) If an inerting system FRM is used, the the tank. Input values for these data must be effects of any air that may enter the fuel obtained from ground and flight test data or tank following the last flight of the day due a thermal model of the tank that has been to changes in ambient temperature, as de- validated by ground and flight test data. fined in Table 4, during a 12-hour overnight (6) Maximum airplane operating tempera- period.

ture limit, as defined by any limitations in (e) The applicant must submit to the re- the airplane flight manual. sponsible Aircraft Certification Service (7) Airplane Utilization. The applicant officefor approval the fuel tank flammability must provide data supporting the number of analysis, including the airplane-specific pa- flights per day and the number of hours per rameters identified under paragraph N25.3(c) flight for the specific airplane model under of this appendix and any deviations from the evaluation. If there is no existing airplane parameters identified in paragraph N25.3(b) fleet data to support the airplane being eval- of this appendix that affect flammability ex- uated, the applicant must provide substan- posure, substantiating data, and any air- tiation that the number of flights per day worthiness limitations and other conditions and the number of hours per flight for that assumed in the analysis.

airplane model is consistent with the exist- N25.4 Variables and data tables.

The following data must be used when con- ing fleet data they propose to use.

(d) Fuel Tank FRM Model. If FRM is used, ducting a flammability exposure analysis to an FAA approved Monte Carlo program must determine the fleet average flammability ex- be used to show compliance with the flam- posure. Variables used to calculate fleet mability requirements of § 25.981 and Appen- flammability exposure must include atmos- dix M of this part. The program must deter- pheric ambient temperatures, flight length, mine the time periods during each flight flammability exposure evaluation time, fuel phase when the fuel tank or compartment flash point, thermal characteristics of the with the FRM would be flammable. The fol- fuel tank, overnight temperature drop, and lowing factors must be considered in estab- oxygen evolution from the fuel into the lishing these time periods: ullage.

(1) Any time periods throughout the flam- (a) Atmospheric Ambient Temperatures mability exposure evaluation time and under and Fuel Properties.

the full range of expected operating condi- (1) In order to predict flammability expo- tions, when the FRM is operating properly sure during a given flight, the variation of but fails to maintain a non-flammable fuel ground ambient temperatures, cruise ambi- tank because of the effects of the fuel tank ent temperatures, and a method to compute vent system or other causes, the transition from ground to cruise and (2) If dispatch with the system inoperative back again must be used. The variation of under the Master Minimum Equipment List the ground and cruise ambient temperatures (MMEL) is requested, the time period as- and the flash point of the fuel is defined by sumed in the reliability analysis (60 flight a Gaussian curve, given by the 50 percent hours must be used for a 10-day MMEL dis- value and a ± 1-standard deviation value.

Federal Aviation Administration, DOT Pt. 25, App. N

(2) Ambient Temperature: Under the pro- sure analysis is a function of the flash point gram, the ground and cruise ambient tem- of the fuel selected by the Monte Carlo for a peratures are linked by a set of assumptions given flight. The flammability envelope for on the atmosphere. The temperature varies the fuel is defined by the upper flammability with altitude following the International limit (UFL) and lower flammability limit Standard Atmosphere (ISA) rate of change (LFL) as follows: from the ground ambient temperature until (A) LFL at sea level = flash point tempera- the cruise temperature for the flight is ture of the fuel at sea level minus 10 ° F. LFL reached. Above this altitude, the ambient decreases from sea level value with increas- temperature is fixed at the cruise ambient ing altitude at a rate of 1 ° F per 808 feet.

temperature. This results in a variation in (B) UFL at sea level = flash point tempera- the upper atmospheric temperature. For cold ture of the fuel at sea level plus 63.5 ° F. UFL days, an inversion is applied up to 10,000 feet, decreases from the sea level value with in- and then the ISA rate of change is used.

creasing altitude at a rate of 1 ° F per 512 (3) Fuel properties: feet.

(i) For Jet A fuel, the variation of flash (4) For each flight analyzed, a separate point of the fuel is defined by a Gaussian random number must be generated for each curve, given by the 50 percent value and a ± 1- of the three parameters (ground ambient standard deviation, as shown in Table 1 of temperature, cruise ambient temperature, this appendix.

and fuel flash point) using the Gaussian dis- (ii) The flammability envelope of the fuel that must be used for the flammability expo- tribution defined in Table 1 of this appendix.

T ABLE 1.—G AUSSIAN D ISTRIBUTION FOR G ROUND AMBIENT T EMPERATURE , C RUISE A MBIENT T EMPERATURE , AND F UEL F LASH P OINT Temperature in deg F Parameter Ground ambient Cruise ambient Fuel flash point temperature temperature (FP) Mean Temp ........................................................................................... 59.95 ¥ 70 120 Neg 1 std dev ........................................................................................ 20.14 8 8 Pos 1 std dev ........................................................................................ 17.28 8 8 (b) The Flight Length Distribution defined in Table 2 must be used in the Monte Carlo analysis.

T ABLE 2.—F LIGHT L ENGTH D ISTRIBUTION Flight length (NM) Airplane maximum range—nautical miles (NM) From To 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000 Distribution of flight lengths (percentage of total) 0 200 11.7 7.5 6.2 5.5 4.7 4.0 3.4 3.0 2.6 2.3 200 400 27.3 19.9 17.0 15.2 13.2 11.4 9.7 8.5 7.5 6.7 400 600 46.3 40.0 35.7 32.6 28.5 24.9 21.2 18.7 16.4 14.8 600 800 10.3 11.6 11.0 10.2 9.1 8.0 6.9 6.1 5.4 4.8 800 1000 4.4 8.5 8.6 8.2 7.4 6.6 5.7 5.0 4.5 4.0 1000 1200 0.0 4.8 5.3 5.3 4.8 4.3 3.8 3.3 3.0 2.7 1200 1400 0.0 3.6 4.4 4.5 4.2 3.8 3.3 3.0 2.7 2.4 1400 1600 0.0 2.2 3.3 3.5 3.3 3.1 2.7 2.4 2.2 2.0 1600 1800 0.0 1.2 2.3 2.6 2.5 2.4 2.1 1.9 1.7 1.6 1800 2000 0.0 0.7 2.2 2.6 2.6 2.5 2.2 2.0 1.8 1.7 2000 2200 0.0 0.0 1.6 2.1 2.2 2.1 1.9 1.7 1.6 1.4 2200 2400 0.0 0.0 1.1 1.6 1.7 1.7 1.6 1.4 1.3 1.2 2400 2600 0.0 0.0 0.7 1.2 1.4 1.4 1.3 1.2 1.1 1.0 2600 2800 0.0 0.0 0.4 0.9 1.0 1.1 1.0 0.9 0.9 0.8 2800 3000 0.0 0.0 0.2 0.6 0.7 0.8 0.7 0.7 0.6 0.6 3000 3200 0.0 0.0 0.0 0.6 0.8 0.8 0.8 0.8 0.7 0.7 3200 3400 0.0 0.0 0.0 0.7 1.1 1.2 1.2 1.1 1.1 1.0 3400 3600 0.0 0.0 0.0 0.7 1.3 1.6 1.6 1.5 1.5 1.4 3600 3800 0.0 0.0 0.0 0.9 2.2 2.7 2.8 2.7 2.6 2.5 3800 4000 0.0 0.0 0.0 0.5 2.0 2.6 2.8 2.8 2.7 2.6 4000 4200 0.0 0.0 0.0 0.0 2.1 3.0 3.2 3.3 3.2 3.1 4200 4400 0.0 0.0 0.0 0.0 1.4 2.2 2.5 2.6 2.6 2.5 4400 4600 0.0 0.0 0.0 0.0 1.0 2.0 2.3 2.5 2.5 2.4 4600 4800 0.0 0.0 0.0 0.0 0.6 1.5 1.8 2.0 2.0 2.0

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. N

T ABLE 2.—F LIGHT L ENGTH D ISTRIBUTION —Continued Flight length (NM) Airplane maximum range—nautical miles (NM) From To 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000 4800 5000 0.0 0.0 0.0 0.0 0.2 1.0 1.4 1.5 1.6 1.5 5000 5200 0.0 0.0 0.0 0.0 0.0 0.8 1.1 1.3 1.3 1.3 5200 5400 0.0 0.0 0.0 0.0 0.0 0.8 1.2 1.5 1.6 1.6 5400 5600 0.0 0.0 0.0 0.0 0.0 0.9 1.7 2.1 2.2 2.3 5600 5800 0.0 0.0 0.0 0.0 0.0 0.6 1.6 2.2 2.4 2.5 5800 6000 0.0 0.0 0.0 0.0 0.0 0.2 1.8 2.4 2.8 2.9 6000 6200 0.0 0.0 0.0 0.0 0.0 0.0 1.7 2.6 3.1 3.3 6200 6400 0.0 0.0 0.0 0.0 0.0 0.0 1.4 2.4 2.9 3.1 6400 6600 0.0 0.0 0.0 0.0 0.0 0.0 0.9 1.8 2.2 2.5 6600 6800 0.0 0.0 0.0 0.0 0.0 0.0 0.5 1.2 1.6 1.9 6800 7000 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.8 1.1 1.3 7000 7200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.4 0.7 0.8 7200 7400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.3 0.5 0.7 7400 7600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.5 0.6 7600 7800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 0.5 0.7 7800 8000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 0.6 0.8 8000 8200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 0.8 8200 8400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 1.0 8400 8600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.6 1.3 8600 8800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.4 1.1 8800 9000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.8 9000 9200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 9200 9400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 9400 9600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 9600 9800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 9800 10000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 (c) Overnight Temperature Drop. For air- T ABLE 4.—O UTSIDE A IR T EMPERATURE (OAT) planes on which FRM is installed, the over- D ROP night temperature drop for this appendix is defined using: OAT drop Parameter temperature ° F (1) A temperature at the beginning of the overnight period that equals the landing Mean Temp ................................................... 12.0 temperature of the previous flight that is a 1 std dev ........................................................ 6.0 random value based on a Gaussian distribu- tion; and (d) Number of Simulated Flights Required (2) An overnight temperature drop that is in Analysis. In order for the Monte Carlo a random value based on a Gaussian distribu- analysis to be valid for showing compliance tion.

with the fleet average and warm day flam- (3) For any flight that will end with an mability exposure requirements, the appli- overnight ground period (one flight per day cant must run the analysis for a minimum out of an average number of flights per day, number of flights to ensure that the fleet av- depending on utilization of the particular erage and warm day flammability exposure airplane model being evaluated), the landing for the fuel tank under evaluation meets the outside air temperature (OAT) is to be cho- applicable flammability limits defined in sen as a random value from the following Table 5 of this appendix.

Gaussian curve: T ABLE 5.—F LAMMABILITY E XPOSURE LIMIT T ABLE 3.—L ANDING O UTSIDE A IR T EMPERATURE Maximum Maximum acceptable Monte acceptable Monte Landing outside Parameter Carlo average fuel Carlo average fuel air temperature ° F Minimum number of tank flammability tank flammability flights in Monte exposure exposure Mean Temperature .................................. 58.68 Carlo analysis (percent) to meet (percent) to meet negative 1 std dev ................................... 20.55 3 percent 7 percent part 26 positive 1 std dev .................................... 13.21 requirements requirements 10,000 .................... 2.91 6.79 (4) The outside ambient air temperature 100,000 .................. 2.98 6.96 (OAT) overnight temperature drop is to be 1,000,000 ............... 3.00 7.00 chosen as a random value from the following Gaussian curve:

Federal Aviation Administration, DOT Pt. 25, App. O

[Doc. No. FAA–2005–22997, 73 FR 42495, July N OTE : Liquid water content (LWC) in 21, 2008, as amended by Doc. No. FAA–2018– grams per cubic meter (g/m ) based on hori- 0119, Amdt. 25–145, 83 FR 9169, Mar. 5, 2018] zontal extent standard distance of 17.4 nau- tical miles.

A PPENDIX O TO P ART 25—S UPERCOOLED (5) Drop diameter distribution: Figure 2.

L ARGE D ROP ICING C ONDITIONS (6) Altitude and temperature envelope: Figure 3.

This Appendix consists of two parts. Part I (b) Freezing Rain (Conditions with spectra defines this Appendix as a description of maximum drop diameters greater than 500 supercooled large drop icing conditions in μ m): which the drop median volume diameter (1) Pressure altitude range: 0 to 12,000 ft (MVD) is less than or greater than 40 μ m, the MSL.

maximum mean effective drop diameter (2) Maximum vertical extent: 7,000 ft.

(MED) of Appendix C of this part continuous (3) Horizontal extent: Standard distance of maximum (stratiform clouds) icing condi- 17.4 nautical miles.

tions. For this Appendix, supercooled large drop icing conditions consist of freezing driz- (4) Total liquid water content.

zle and freezing rain occurring in and/or N OTE : LWC in grams per cubic meter (g/m ) below stratiform clouds. Part II defines ice based on horizontal extent standard distance accretions used to show compliance with the of 17.4 nautical miles.

airplane performance and handling qualities (5) Drop Diameter Distribution: Figure 5.

requirements of subpart B of this part.

(6) Altitude and temperature envelope: Figure 6.

PART I—METEOROLOGY (c) Horizontal extent.

In this Appendix icing conditions are de- The liquid water content for freezing driz- fined by the parameters of altitude, vertical zle and freezing rain conditions for hori- and horizontal extent, temperature, liquid zontal extents other than the standard 17.4 water content, and water mass distribution nautical miles can be determined by the as a function of drop diameter distribution.

value of the liquid water content determined (a) Freezing Drizzle (Conditions with spec- from Figure 1 or Figure 4, multiplied by the tra maximum drop diameters from 100 μ m to factor provided in Figure 7, which is defined 500 μ m): by the following equation: (1) Pressure altitude range: 0 to 22,000 feet S = 1.266 ¥ 0.213 log10(H) MSL.

Where: (2) Maximum vertical extent: 12,000 feet.

S = Liquid Water Content Scale Factor (3) Horizontal extent: Standard distance of (dimensionless) and 17.4 nautical miles.

(4) Total liquid water content. H = horizontal extent in nautical miles

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. O

Federal Aviation Administration, DOT Pt. 25, App. O

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. O

Federal Aviation Administration, DOT Pt. 25, App. O

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. O

Federal Aviation Administration, DOT Pt. 25, App. O

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. O

PART II—AIRFRAME ICE ACCRETIONS in part II, paragraph (c) of this Appendix, FOR SHOWING COMPLIANCE WITH SUB- only the portion of part I of this Appendix in PART B OF THIS PART which the airplane is capable of operating safely must be considered.

(a) General. The most critical ice accretion (3) For an airplane certified in accordance in terms of airplane performance and han- with § 25.1420(a)(3), the ice accretions for each dling qualities for each flight phase must be used to show compliance with the applicable flight phase are defined in part II, paragraph airplane performance and handling qualities (c) of this Appendix.

requirements for icing conditions contained (b) Ice accretions for airplanes certified in in subpart B of this part. Applicants must accordance with § 25.1420(a)(1) or (2).

demonstrate that the full range of atmos- (1) En route ice is the en route ice as de- pheric icing conditions specified in part I of fined by part II, paragraph (c)(3), of this Ap- this Appendix have been considered, includ- pendix, for an airplane certified in accord- ing drop diameter distributions, liquid water ance with § 25.1420(a)(2), or defined by part II, content, and temperature appropriate to the paragraph (a)(3), of Appendix C of this part, flight conditions (for example, configuration, for an airplane certified in accordance with speed, angle of attack, and altitude).

§ 25.1420(a)(1), plus: (1) For an airplane certified in accordance (i) Pre-detection ice as defined by part II, with § 25.1420(a)(1), the ice accretions for each paragraph (b)(5), of this Appendix; and flight phase are defined in part II, paragraph (ii) The ice accumulated during the transit (b) of this Appendix.

of one cloud with a horizontal extent of 17.4 (2) For an airplane certified in accordance nautical miles in the most critical of the with § 25.1420(a)(2), the most critical ice ac- icing conditions defined in part I of this Ap- cretion for each flight phase defined in part II, paragraphs (b) and (c) of this Appendix, pendix and one cloud with a horizontal ex- must be used. For the ice accretions defined tent of 17.4 nautical miles in the continuous

Federal Aviation Administration, DOT Pt. 25, App. O

maximum icing conditions defined in Appen- (b)(2), of this Appendix, or the ice calculated dix C of this part. in the applicable paragraphs (b)(4)(i) or (ii) of (2) Holding ice is the holding ice defined by part II of this Appendix: part II, paragraph (c)(4), of this Appendix, for (i) For an airplane certified in accordance an airplane certified in accordance with with § 25.1420(a)(2), the ice accretion defined § 25.1420(a)(2), or defined by part II, paragraph by part II, paragraph (c)(5)(i), of this Appen- (a)(4), of Appendix C of this part, for an air- dix, plus a descent from 2,000 feet above the plane certified in accordance with landing surface to a height of 200 feet above § 25.1420(a)(1), plus: the landing surface with a transition to the (i) Pre-detection ice as defined by part II, landing configuration in the icing conditions paragraph (b)(5), of this Appendix; and defined in part I of this Appendix, plus: (ii) The ice accumulated during the transit (A) Pre-detection ice, as defined in part II, of one cloud with a 17.4 nautical miles hori- paragraph (b)(5), of this Appendix; and zontal extent in the most critical of the (B) The ice accumulated during an exit icing conditions defined in part I of this Ap- maneuver, beginning with the minimum pendix and one cloud with a horizontal ex- climb gradient required by § 25.119, from a tent of 17.4 nautical miles in the continuous height of 200 feet above the landing surface maximum icing conditions defined in Appen- through one cloud with a horizontal extent dix C of this part.

of 17.4 nautical miles in the most critical of (iii) Except the total exposure to holding the icing conditions defined in part I of this ice conditions does not need to exceed 45 Appendix and one cloud with a horizontal ex- minutes.

tent of 17.4 nautical miles in the continuous (3) Approach ice is the more critical of the maximum icing conditions defined in Appen- holding ice defined by part II, paragraph dix C of this part.

(b)(2), of this Appendix, or the ice calculated in the applicable paragraphs (b)(3)(i) or (ii) of (ii) For an airplane certified in accordance part II, of this Appendix: with § 25.1420(a)(1), the ice accumulated in (i) For an airplane certified in accordance the maximum continuous icing conditions with § 25.1420(a)(2), the ice accumulated dur- defined in Appendix C of this part, during a ing descent from the maximum vertical ex- descent from the maximum vertical extent tent of the icing conditions defined in part I of the icing conditions defined in Appendix C of this Appendix to 2,000 feet above the land- of this part, to 2,000 feet above the landing ing surface in the cruise configuration, plus surface in the cruise configuration, plus transition to the approach configuration, transition to the approach configuration and plus: flying for 15 minutes at 2,000 feet above the (A) Pre-detection ice, as defined by part II, landing surface, plus a descent from 2,000 paragraph (b)(5), of this Appendix; and feet above the landing surface to a height of (B) The ice accumulated during the transit 200 feet above the landing surface with a at 2,000 feet above the landing surface of one transition to the landing configuration, plus: cloud with a horizontal extent of 17.4 nau- (A) Pre-detection ice, as described by part tical miles in the most critical of the icing II, paragraph (b)(5), of this Appendix; and conditions defined in part I of this Appendix (B) The ice accumulated during an exit and one cloud with a horizontal extent of 17.4 maneuver, beginning with the minimum nautical miles in the continuous maximum climb gradient required by § 25.119, from a icing conditions defined in Appendix C of height of 200 feet above the landing surface this part.

through one cloud with a horizontal extent (ii) For an airplane certified in accordance of 17.4 nautical miles in the most critical of with § 25.1420(a)(1), the ice accumulated dur- the icing conditions defined in part I of this ing descent from the maximum vertical ex- Appendix and one cloud with a horizontal ex- tent of the maximum continuous icing condi- tent of 17.4 nautical miles in the continuous tions defined in part I of Appendix C to 2,000 maximum icing conditions defined in Appen- feet above the landing surface in the cruise dix C of this part.

configuration, plus transition to the ap- (5) Pre-detection ice is the ice accretion be- proach configuration, plus: fore detection of flight conditions in this Ap- (A) Pre-detection ice, as defined by part II, pendix that require exiting per § 25.1420(a)(1) paragraph (b)(5), of this Appendix; and and (2). It is the pre-existing ice accretion (B) The ice accumulated during the transit that may exist from operating in icing condi- at 2,000 feet above the landing surface of one cloud with a horizontal extent of 17.4 nau- tions in which the airplane is approved to op- tical miles in the most critical of the icing erate prior to encountering the icing condi- conditions defined in part I of this Appendix tions requiring an exit, plus the ice accumu- and one cloud with a horizontal extent of 17.4 lated during the time needed to detect the nautical miles in the continuous maximum icing conditions, followed by two minutes of icing conditions defined in Appendix C of further ice accumulation to take into ac- this part. count the time for the flightcrew to take ac- (4) Landing ice is the more critical of the tion to exit the icing conditions, including holding ice as defined by part II, paragraph coordination with air traffic control.

14 CFR Ch. I (1–1–25 Edition) Pt. 25, App. O

(i) For an airplane certified in accordance flying for 15 minutes at 2,000 feet above the with § 25.1420(a)(1), the pre-existing ice accre- landing surface; or tion must be based on the icing conditions (ii) Holding ice as defined by part II, para- defined in Appendix C of this part.

graph (c)(4), of this Appendix.

(ii) For an airplane certified in accordance (6) Landing ice is the ice accretion on the with § 25.1420(a)(2), the pre-existing ice accre- unprotected surfaces, and any ice accretion tion must be based on the more critical of on the protected surfaces appropriate to nor- the icing conditions defined in Appendix C of mal ice protection system operation, result- this part, or the icing conditions defined in ing from the more critical of the: part I of this Appendix in which the airplane (i) Ice accretion defined by part II, para- is capable of safely operating.

graph (c)(5)(i), of this Appendix, plus ice ac- (c) Ice accretions for airplanes certified in ac- cumulated in the icing conditions defined in cordance with §§ 25.1420(a)(2) or (3). For an air- part I of this Appendix during a descent from plane certified in accordance with 2,000 feet above the landing surface to a § 25.1420(a)(2), only the portion of the icing height of 200 feet above the landing surface conditions of part I of this Appendix in with a transition to the landing configura- which the airplane is capable of operating tion, followed by a go-around at the min- safely must be considered.

imum climb gradient required by § 25.119, (1) Takeoff ice is the most critical ice accre- from a height of 200 feet above the landing tion on unprotected surfaces, and any ice ac- surface to 2,000 feet above the landing sur- cretion on the protected surfaces, occurring face, flying for 15 minutes at 2,000 feet above between the end of the takeoff distance and the landing surface in the approach configu- 400 feet above the takeoff surface, assuming ration, and a descent to the landing surface accretion starts at the end of the takeoff dis- (touchdown) in the landing configuration; or tance in the icing conditions defined in part (ii) Holding ice as defined by part II, para- I of this Appendix.

graph (c)(4), of this Appendix.

(2) Final takeoff ice is the most critical ice (7) For both unprotected and protected accretion on unprotected surfaces, and any parts, the ice accretion for the takeoff phase ice accretion on the protected surfaces ap- must be determined for the icing conditions propriate to normal ice protection system defined in part I of this Appendix, using the operation, between 400 feet and either 1,500 following assumptions: feet above the takeoff surface, or the height at which the transition from the takeoff to (i) The airfoils, control surfaces, and, if ap- the en route configuration is completed and plicable, propellers are free from frost, snow, V FTO is reached, whichever is higher. Ice ac- or ice at the start of takeoff; cretion is assumed to start at the end of the (ii) The ice accretion starts at the end of takeoff distance in the icing conditions de- the takeoff distance; fined in part I of this Appendix.

(iii) The critical ratio of thrust/power-to- (3) En route ice is the most critical ice ac- weight; cretion on the unprotected surfaces, and any (iv) Failure of the critical engine occurs at ice accretion on the protected surfaces ap- V ; and EF propriate to normal ice protection system (v) Crew activation of the ice protection operation, during the en route flight phase in system is in accordance with a normal oper- the icing conditions defined in part I of this ating procedure provided in the airplane Appendix.

flight manual, except that after beginning (4) Holding ice is the most critical ice ac- the takeoff roll, it must be assumed that the cretion on the unprotected surfaces, and any crew takes no action to activate the ice pro- ice accretion on the protected surfaces ap- tection system until the airplane is at least propriate to normal ice protection system 400 feet above the takeoff surface.

operation, resulting from 45 minutes of flight (d) The ice accretion before the ice protec- within a cloud with a 17.4 nautical miles hor- tion system has been activated and is per- izontal extent in the icing conditions defined forming its intended function is the critical in part I of this Appendix, during the holding ice accretion formed on the unprotected and phase of flight.

normally protected surfaces before activa- (5) Approach ice is the ice accretion on the tion and effective operation of the ice pro- unprotected surfaces, and any ice accretion tection system in the icing conditions de- on the protected surfaces appropriate to nor- fined in part I of this Appendix. This ice ac- mal ice protection system operation, result- cretion only applies in showing compliance ing from the more critical of the: to §§ 25.143(j) and 25.207(h).

(i) Ice accumulated in the icing conditions (e) In order to reduce the number of ice ac- defined in part I of this Appendix during a descent from the maximum vertical extent cretions to be considered when dem- of the icing conditions defined in part I of onstrating compliance with the require- this Appendix, to 2,000 feet above the landing ments of § 25.21(g), any of the ice accretions surface in the cruise configuration, plus defined in this Appendix may be used for any transition to the approach configuration and other flight phase if it is shown to be at least

Federal Aviation Administration, DOT § 26.5 as critical as the specific ice accretion de- Subpart A—General fined for that flight phase. Configuration dif- ferences and their effects on ice accretions § 26.1 Purpose and scope.

must be taken into account.

(f) The ice accretion that has the most ad- (a) This part establishes require- verse effect on handling qualities may be ments for support of the continued air- used for airplane performance tests provided worthiness of and safety improvements any difference in performance is conserv- for transport category airplanes. These atively taken into account.

requirements may include performing [Amdt. 25–140, 79 FR 65528, Nov. 4, 2014] assessments, developing design changes, developing revisions to In- PART 26—CONTINUED AIRWORTHI- structions for Continued Airworthiness (ICA), and making necessary docu- NESS AND SAFETY IMPROVE- mentation available to affected per- MENTS FOR TRANSPORT CAT- sons. Requirements of this part that EGORY AIRPLANES establish standards for design changes and revisions to the ICA are considered Subpart A—General airworthiness requirements.

Sec.

(b) Except as provided in paragraph 26.1 Purpose and scope.

(c) of this section, this part applies to 26.3 [Reserved] the following persons, as specified in 26.5 Applicability table.

each subpart of this part: (1) Holders of type certificates and Subpart B—Enhanced Airworthiness supplemental type certificates.

Program for Airplane Systems (2) Applicants for type certificates 26.11 Electrical wiring interconnection sys- and supplemental type certificates and tems (EWIS) maintenance program.

changes to those certificates (including service bulletins describing design Subpart C—Aging Airplane Safety— changes).

Widespread Fatigue Damage (3) Persons seeking design approval 26.21 Limit of validity.

for airplane repairs, alterations, or 26.23 Extended limit of validity.

modifications that may affect air- worthiness.

Subpart D—Fuel Tank Flammability (4) Holders of type certificates and their licensees producing new air- 26.31 Definitions.

26.33 Holders of type certificates: Fuel tank planes.

flammability.

(c) An applicant for approval of a de- 26.35 Changes to type certificates affecting sign change is not required to comply fuel tank flammability.

with any applicable airworthiness re- 26.37 Pending type certification projects: quirement of this part if the applicant Fuel tank flammability.

elects or is required to comply with a 26.39 Newly produced airplanes: Fuel tank corresponding amendment to part 25 of flammability.

this chapter that is adopted concur- Subpart E—Aging Airplane Safety—Dam- rently or after that airworthiness re- age Tolerance Data for Repairs and quirement.

Alterations (d) For the purposes of this part, the word ‘‘type certificate’’ does not in- 26.41 Definitions.

clude supplemental type certificates.

26.43 Holders of and applicants for type cer- tificates—Repairs.

§ 26.3 [Reserved] 26.45 Holders of type certificates—Alter- ations and repairs to alterations.

§ 26.5 Applicability table.

26.47 Holders of and applicants for a supple- mental type certificate—Alterations and Table 1 of this section provides an repairs to alterations.

overview of the applicability of this 26.49 Compliance plan.

part. It provides guidance in identi- A UTHORITY : 49 U.S.C. 106(g), 40113, 44701, fying what sections apply to various 44702 and 44704.

types of entities. The specific applica- bility of each subpart and section is S OURCE : Docket No. FAA–2004–18379, 72 FR 63409, Nov. 8, 2007, unless otherwise noted. specified in the regulatory text.