14 CFR Part 27 — Airworthiness Standards: Normal Category Rotorcraft

Federal Aviation Administration, DOT Pt. 27 27.239 Spray characteristics. 27.573 Damage tolerance and fatigue evalua- 27.241 Ground resonance. tion of composite rotorcraft structures.

M ISCELLANEOUS F LIGHT R EQUIREMENTS Subpart D—Design and Construction 27.251 Vibration.

G ENERAL Subpart C—Strength Requirements 27.601 Design.

27.602 Critical parts.

G ENERAL 27.603 Materials.

27.605 Fabrication methods.

27.301 Loads.

27.607 Fasteners.

27.303 Factor of safety.

27.609 Protection of structure.

27.305 Strength and deformation.

27.610 Lightning and static electricity pro- 27.307 Proof of structure.

tection.

27.309 Design limitations.

27.611 Inspection provisions.

FLIGHT L OADS 27.613 Material strength properties and de- sign values.

27.321 General.

27.619 Special factors.

27.337 Limit maneuvering load factor.

27.621 Casting factors.

27.339 Resultant limit maneuvering loads.

27.623 Bearing factors.

27.341 Gust loads.

27.625 Fitting factors.

27.351 Yawing conditions.

27.629 Flutter.

27.361 Engine torque.

R OTORS C ONTROL S URFACE AND S YSTEM L OADS 27.653 Pressure venting and drainage of 27.391 General.

rotor blades.

27.395 Control system.

27.659 Mass balance.

27.397 Limit pilot forces and torques.

27.661 Rotor blade clearance.

27.399 Dual control system.

27.663 Ground resonance prevention means.

27.411 Ground clearance: tail rotor guard.

27.427 Unsymmetrical loads.

C ONTROL S YSTEMS G ROUND L OADS 27.671 General.

27.471 General. 27.672 Stability augmentation, automatic, 27.473 Ground loading conditions and as- and power-operated systems.

sumptions. 27.673 Primary flight control.

27.475 Tires and shock absorbers. 27.674 Interconnected controls.

27.477 Landing gear arrangement. 27.675 Stops.

27.479 Level landing conditions. 27.679 Control system locks.

27.481 Tail-down landing conditions. 27.681 Limit load static tests.

27.483 One-wheel landing conditions. 27.683 Operation tests.

27.485 Lateral drift landing conditions. 27.685 Control system details.

27.493 Braked roll conditions.

27.687 Spring devices.

27.497 Ground loading conditions: landing 27.691 Autorotation control mechanism.

gear with tail wheels.

27.695 Power boost and power-operated con- 27.501 Ground loading conditions: landing trol system.

gear with skids.

L ANDING G EAR 27.505 Ski landing conditions.

27.723 Shock absorption tests.

W ATER L OADS 27.725 Limit drop test.

27.521 Float landing conditions.

27.727 Reserve energy absorption drop test.

27.729 Retracting mechanism.

M AIN C OMPONENT R EQUIREMENTS 27.731 Wheels.

27.733 Tires.

27.547 Main rotor structure.

27.735 Brakes.

27.549 Fuselage, landing gear, and rotor 27.737 Skis.

pylon structures.

MERGENCY L ANDING C ONDITIONS F LOATS AND H ULLS E 27.561 General. 27.751 Main float buoyancy.

27.562 Emergency landing dynamic condi- 27.753 Main float design.

tions. 27.755 Hulls.

27.563 Structural ditching provisions.

P ERSONNEL AND C ARGO A CCOMMODATIONS F ATIGUE E VALUATION 27.771 Pilot compartment.

27.571 Fatigue evaluation of flight struc- 27.773 Pilot compartment view.

ture. 27.775 Windshields and windows.

14 CFR Ch. I (1–1–25 Edition) Pt. 27 27.777 Cockpit controls. 27.995 Fuel valves.

27.779 Motion and effect of cockpit controls. 27.997 Fuel strainer or filter.

27.783 Doors. 27.999 Fuel system drains.

27.785 Seats, berths, litters, safety belts, O IL SYSTEM and harnesses.

27.787 Cargo and baggage compartments.

27.1011 Engines: General.

27.801 Ditching.

27.1013 Oil tanks.

27.805 Flight crew emergency exits.

27.1015 Oil tank tests.

27.807 Emergency exits.

27.1017 Oil lines and fittings.

27.831 Ventilation.

27.1019 Oil strainer or filter.

27.833 Heaters.

27.1021 Oil system drains.

27.1027 Transmissions and gearboxes: Gen- F IRE P ROTECTION eral.

27.853 Compartment interiors.

COOLING 27.855 Cargo and baggage compartments.

27.859 Heating systems.

27.1041 General.

27.861 Fire protection of structure, controls, 27.1043 Cooling tests.

and other parts.

27.1045 Cooling test procedures.

27.863 Flammable fluid fire protection.

I NDUCTION SYSTEM EXTERNAL L OADS 27.1091 Air induction.

27.865 External loads.

27.1093 Induction system icing protection.

M ISCELLANEOUS EXHAUST S YSTEM 27.871 Leveling marks.

27.1121 General.

27.873 Ballast provisions.

27.1123 Exhaust piping.

Subpart E—Powerplant P OWERPLANT CONTROLS AND A CCESSORIES 27.1141 Powerplant controls: general.

G ENERAL 27.1143 Engine controls.

27.901 Installation.

27.1145 Ignition switches.

27.903 Engines.

27.1147 Mixture controls.

27.907 Engine vibration.

27.1151 Rotor brake controls.

27.1163 Powerplant accessories.

ROTOR D RIVE S YSTEM P OWERPLANT F IRE P ROTECTION 27.917 Design.

27.921 Rotor brake.

27.1183 Lines, fittings, and components.

27.923 Rotor drive system and control mech- 27.1185 Flammable fluids.

anism tests.

27.1187 Ventilation and drainage.

27.927 Additional tests.

27.1189 Shutoff means.

27.931 Shafting critical speed.

27.1191 Firewalls.

27.935 Shafting joints.

27.1193 Cowling and engine compartment 27.939 Turbine engine operating characteris- covering.

tics.

27.1194 Other surfaces.

27.1195 Fire detector systems.

F UEL S YSTEM Subpart F—Equipment 27.951 General.

27.952 Fuel system crash resistance.

G ENERAL 27.953 Fuel system independence.

27.954 Fuel system lightning protection.

27.1301 Function and installation.

27.955 Fuel flow.

27.1303 Flight and navigation instruments.

27.959 Unusable fuel supply.

27.1305 Powerplant instruments.

27.961 Fuel system hot weather operation.

27.1307 Miscellaneous equipment.

27.963 Fuel tanks: general.

27.1309 Equipment, systems, and installa- 27.965 Fuel tank tests.

tions.

27.967 Fuel tank installation.

27.1316 Electrical and electronic system 27.969 Fuel tank expansion space.

lightning protection.

27.971 Fuel tank sump. 27.1317 High-intensity Radiated Fields 27.973 Fuel tank filler connection. (HIRF) Protection.

27.975 Fuel tank vents.

INSTRUMENTS : I NSTALLATION 27.977 Fuel tank outlet.

27.1321 Arrangement and visibility.

F UEL S YSTEM C OMPONENTS 27.1322 Warning, caution, and advisory 27.991 Fuel pumps. lights.

27.993 Fuel system lines and fittings. 27.1323 Airspeed indicating system.

Federal Aviation Administration, DOT § 27.2 27.1325 Static pressure systems. 27.1547 Magnetic direction indicator.

27.1327 Magnetic direction indicator.

27.1549 Powerplant instruments.

27.1329 Automatic pilot and flight guidance 27.1551 Oil quantity indicator.

system.

27.1553 Fuel quantity indicator.

27.1337 Powerplant instruments.

27.1555 Control markings.

27.1557 Miscellaneous markings and plac- E LECTRICAL S YSTEMS AND E QUIPMENT ards.

27.1351 General.

27.1559 Limitations placard.

27.1353 Energy storage systems.

27.1561 Safety equipment.

27.1357 Circuit protective devices.

27.1565 Tail rotor.

27.1361 Master switch.

27.1365 Electric cables.

R OTORCRAFT F LIGHT M ANUAL AND A PPROVED 27.1367 Switches.

M ANUAL M ATERIAL L IGHTS 27.1581 General.

27.1583 Operating limitations.

27.1381 Instrument lights.

27.1383 Landing lights. 27.1585 Operating procedures.

27.1385 Position light system installation.

27.1587 Performance information.

27.1387 Position light system dihedral an- 27.1589 Loading information.

gles.

PPENDIX A TO P ART 27—I NSTRUCTIONS FOR A 27.1389 Position light distribution and in- ONTINUED A IRWORTHINESS C tensities.

APPENDIX B TO P ART 27—A IRWORTHINESS C RI - 27.1391 Minimum intensities in the hori- TERIA FOR H ELICOPTER I NSTRUMENT zontal plane of forward and rear position F LIGHT lights.

A PPENDIX C TO P ART 27—C RITERIA FOR C AT - 27.1393 Minimum intensities in any vertical EGORY A plane of forward and rear position lights.

A PPENDIX D TO P ART 27—HIRF E NVIRON - 27.1395 Maximum intensities in overlapping MENTS AND E QUIPMENT HIRF T EST L EV - beams of forward and rear position ELS lights.

27.1397 Color specifications.

A UTHORITY : 49 U.S.C. 106(f), 106(g), 40113, 27.1399 Riding light.

44701–44702, 44704.

27.1401 Anticollision light system.

S OURCE : Docket No. 5074, 29 FR 15695, Nov.

SAFETY EQUIPMENT 24, 1964, unless otherwise noted.

27.1411 General.

27.1413 Safety belts.

Subpart A—General 27.1415 Ditching equipment.

27.1419 Ice protection.

§ 27.1 Applicability.

27.1435 Hydraulic systems.

(a) This part prescribes airworthiness 27.1457 Cockpit voice recorders.

standards for the issue of type certifi- 27.1459 Flight data recorders.

27.1461 Equipment containing high energy cates, and changes to those certifi- rotors.

cates, for normal category rotorcraft with maximum weights of 7,000 pounds Subpart G—Operating Limitations and or less and nine or less passenger seats.

Information (b) Each person who applies under 27.1501 General. Part 21 for such a certificate or change must show compliance with the appli- O PERATING L IMITATIONS cable requirements of this part.

27.1503 Airspeed limitations: general.

(c) Multiengine rotorcraft may be 27.1505 Never-exceed speed.

type certified as Category A provided 27.1509 Rotor speed.

the requirements referenced in appen- 27.1519 Weight and center of gravity.

dix C of this part are met.

27.1521 Powerplant limitations.

27.1523 Minimum flight crew.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as 27.1525 Kinds of operations.

amended by Amdt. 27–33, 61 FR 21906, May 10, 27.1527 Maximum operating altitude.

1996; Amdt. 27–37, 64 FR 45094, Aug. 18, 1999] 27.1529 Instructions for Continued Air- worthiness.

§ 27.2 Special retroactive require- ments.

M ARKINGS AND P LACARDS (a) For each rotorcraft manufactured 27.1541 General.

after September 16, 1992, each applicant 27.1543 Instrument markings: general.

27.1545 Airspeed indicator. must show that each occupant’s seat is 14 CFR Ch. I (1–1–25 Edition) § 27.21 equipped with a safety belt and shoul- ments of this part in effect on October der harness that meets the require- 18, 1999.

ments of paragraphs (a), (b), and (c) of [Doc. No. 26078, 56 FR 41051, Aug. 16, 1991, as this section.

amended by Amdt. 27–37, 64 FR 45094, Aug. 18, (1) Each occupant’s seat must have a 1999] combined safety belt and shoulder har- ness with a single-point release. Each Subpart B—Flight pilot’s combined safety belt and shoul- der harness must allow each pilot, G ENERAL when seated with safety belt and shoul- der harness fastened, to perform all § 27.21 Proof of compliance.

functions necessary for flight oper- Each requirement of this subpart ations. There must be a means to se- must be met at each appropriate com- cure belts and harnesses, when not in bination of weight and center of grav- use, to prevent interference with the ity within the range of loading condi- operation of the rotorcraft and with tions for which certification is re- rapid egress in an emergency.

quested. This must be shown— (2) Each occupant must be protected (a) By tests upon a rotorcraft of the from serious head injury by a safety type for which certification is re- belt plus a shoulder harness that will quested, or by calculations based on, prevent the head from contacting any and equal in accuracy to, the results of injurious object.

testing; and (3) The safety belt and shoulder har- (b) By systematic investigation of ness must meet the static and dynamic each required combination of weight strength requirements, if applicable, and center of gravity if compliance specified by the rotorcraft type certifi- cannot be reasonably inferred from cation basis.

combinations investigated.

(4) For purposes of this section, the [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as date of manufacture is either— amended by Amdt. 27–21, 49 FR 44432, Nov. 6, (i) The date the inspection accept- 1984] ance records, or equivalent, reflect that the rotorcraft is complete and § 27.25 Weight limits.

meets the FAA-Approved Type Design (a) Maximum weight. The maximum Data; or weight (the highest weight at which (ii) The date the foreign civil air- compliance with each applicable re- worthiness authority certifies that the quirement of this part is shown) must rotorcraft is complete and issues an be established so that it is— original standard airworthiness certifi- (1) Not more than— cate, or equivalent, in that country.

(i) The highest weight selected by the (b) For rotorcraft with a certification applicant; basis established prior to October 18, (ii) The design maximum weight (the 1999— highest weight at which compliance (1) The maximum passenger seat ca- with each applicable structural loading pacity may be increased to eight or condition of this part is shown); nine provided the applicant shows com- (iii) The highest weight at which pliance with all the airworthiness re- compliance with each applicable flight quirements of this part in effect on Oc- requirement of this part is shown; or tober 18, 1999.

(iv) The highest weight in which the (2) The maximum weight may be in- provisions of §§ 27.87 or 27.143(c)(1), or creased to greater than 6,000 pounds combinations thereof, are dem- provided— onstrated if the weights and operating (i) The number of passenger seats is conditions (altitude and temperature) not increased above the maximum prescribed by those requirements can- number certificated on October 18, 1999, not be met; and or (2) Not less than the sum of— (ii) The applicant shows compliance (i) The empty weight determined with all of the airworthiness require- under § 27.29; and Federal Aviation Administration, DOT § 27.29 (ii) The weight of usable fuel appro- crease over that established under priate to the intended operation with paragraph (a) of this section, and full payload; (5) Operation of the rotorcraft at a (iii) The weight of full oil capacity; total weight greater than the max- and imum certificated weight established (iv) For each seat, an occupant under paragraph (a) of this section is weight of 170 pounds or any lower limited by appropriate operating limi- weight for which certification is re- tations under § 27.865(a) and (d) of this quested. part.

(b) Minimum weight. The minimum (Secs. 313(a), 601, 603, 604, and 605 of the Fed- weight (the lowest weight at which eral Aviation Act of 1958 (49 U.S.C. 1354(a), compliance with each applicable re- 1421, 1423, 1424, and 1425); and sec. 6(c) of the quirement of this part is shown) must Dept. of Transportation Act (49 U.S.C.

be established so that it is— 1655(c))) (1) Not more than the sum of— [Doc. No. 5074, 29 FR 15695, Nov. 29, 1964, as (i) The empty weight determined amended by Amdt. 27–11, 41 FR 55468, Dec. 20, under § 27.29; and 1976; Amdt. 25–42, 43 FR 2324, Jan. 16, 1978; Amdt. 27–36, 64 FR 43019, Aug. 6, 1999; Amdt.

(ii) The weight of the minimum crew 27–44, 73 FR 10998, Feb. 29, 2008; 73 FR 33876, necessary to operate the rotorcraft, as- June 16, 2008] suming for each crewmember a weight no more than 170 pounds, or any lower § 27.27 Center of gravity limits.

weight selected by the applicant or in- cluded in the loading instructions; and The extreme forward and aft centers (2) Not less than— of gravity and, where critical, the ex- treme lateral centers of gravity must (i) The lowest weight selected by the be established for each weight estab- applicant; lished under § 27.25. Such an extreme (ii) The design minimum weight (the may not lie beyond— lowest weight at which compliance (a) The extremes selected by the ap- with each applicable structural loading plicant; condition of this part is shown); or (b) The extremes within which the (iii) The lowest weight at which com- structure is proven; or pliance with each applicable flight re- quirement of this part is shown. (c) The extremes within which com- pliance with the applicable flight re- (c) Total weight with jettisonable exter- quirements is shown.

nal load. A total weight for the rotor- craft with a jettisonable external load [Amdt. 27–2, 33 FR 962, Jan. 26, 1968] attached that is greater than the max- imum weight established under para- § 27.29 Empty weight and cor- graph (a) of this section may be estab- responding center of gravity.

lished for any rotorcraft-load combina- (a) The empty weight and cor- tion if— responding center of gravity must be (1) The rotorcraft-load combination determined by weighing the rotorcraft does not include human external cargo, without the crew and payload, but (2) Structural component approval with— for external load operations under ei- (1) Fixed ballast; ther § 27.865 or under equivalent oper- (2) Unusable fuel; and ational standards is obtained, (3) Full operating fluids, including— (3) The portion of the total weight (i) Oil; that is greater than the maximum (ii) Hydraulic fluid; and weight established under paragraph (a) of this section is made up only of the (iii) Other fluids required for normal weight of all or part of the jettisonable operation of roto-craft systems, except external load, water intended for injection in the en- gines.

(4) Structural components of the rotorcraft are shown to comply with (b) The condition of the rotorcraft at the applicable structural requirements the time of determining empty weight of this part under the increased loads must be one that is well defined and and stresses caused by the weight in- can be easily repeated, particularly 14 CFR Ch. I (1–1–25 Edition) § 27.31 with respect to the weights of fuel, oil, (2) It is possible to prevent over- coolant, and installed equipment. speeding of the rotor without excep- tional piloting skill.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- (d) Emergency high pitch. If the main eral Aviation Act of 1958 (49 U.S.C. 1354(a), rotor high pitch stop is set to meet 1421, 1423, 1424, and 1425); and sec. 6(c) of the paragraph (b)(1) of this section, and if Dept. of Transportation Act (49 U.S.C.

1655(c))) that stop cannot be exceeded inadvert- ently, additional pitch may be made [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as available for emergency use.

amended by Amdt. 27–14, 43 FR 2324, Jan. 16, (e) Main rotor low speed warning for 1978] helicopters. For each single engine heli- § 27.31 Removable ballast.

copter, and each multiengine heli- copter that does not have an approved Removable ballast may be used in device that automatically increases showing compliance with the flight re- power on the operating engines when quirements of this subpart.

one engine fails, there must be a main § 27.33 Main rotor speed and pitch lim- rotor low speed warning which meets its.

the following requirements: (1) The warning must be furnished to (a) Main rotor speed limits. A range of the pilot in all flight conditions, in- main rotor speeds must be established cluding power-on and power-off flight, that— when the speed of a main rotor ap- (1) With power on, provides adequate proaches a value that can jeopardize margin to accommodate the variations safe flight.

in rotor speed occurring in any appro- (2) The warning may be furnished ei- priate maneuver, and is consistent ther through the inherent aerodynamic with the kind of governor or synchro- qualities of the helicopter or by a de- nizer used; and vice.

(2) With power off, allows each appro- (3) The warning must be clear and priate autorotative maneuver to be distinct under all conditions, and must performed throughout the ranges of be clearly distinguishable from all airspeed and weight for which certifi- other warnings. A visual device that cation is requested.

(b) Normal main rotor high pitch limits requires the attention of the crew (power on). For rotocraft, except heli- within the cockpit is not acceptable by copters required to have a main rotor itself.

low speed warning under paragraph (e) (4) If a warning device is used, the de- of this section. It must be shown, with vice must automatically deactivate power on and without exceeding ap- and reset when the low-speed condition proved engine maximum limitations, is corrected. If the device has an audi- that main rotor speeds substantially ble warning, it must also be equipped less than the minimum approved main with a means for the pilot to manually rotor speed will not occur under any silence the audible warning before the sustained flight condition. This must low-speed condition is corrected.

be met by— (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (1) Appropriate setting of the main eral Aviation Act of 1958 (49 U.S.C. 1354(a), rotor high pitch stop; 1421, 1423, 1424, and 1425); and sec. 6(c) of the (2) Inherent rotorcraft characteris- Dept. of Transportation Act (49 U.S.C.

tics that make unsafe low main rotor 1655(c))) speeds unlikely; or [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (3) Adequate means to warn the pilot amended by Amdt. 27–2, 33 FR 962, Jan. 26, of unsafe main rotor speeds.

1968; Amdt. 27–14, 43 FR 2324, Jan. 16, 1978] (c) Normal main rotor low pitch limits P ERFORMANCE (power off). It must be shown, with power off, that— § 27.45 General.

(1) The normal main rotor low pitch limit provides sufficient rotor speed, in (a) Unless otherwise prescribed, the any autorotative condition, under the performance requirements of this sub- most critical combinations of weight part must be met for still air and a and airspeed; and standard atmosphere.

Federal Aviation Administration, DOT § 27.65 (b) The performance must correspond (2) The hovering ceiling determined to the engine power available under the under paragraph (a)(1) of this section particular ambient atmospheric condi- must be at least— tions, the particular flight condition, (i) For reciprocating engine powered and the relative humidity specified in helicopters, 4,000 feet at maximum paragraphs (d) or (e) of this section, as weight with a standard atmosphere; appropriate. (ii) For turbine engine powered heli- (c) The available power must cor- copters, 2,500 feet pressure altitude at respond to engine power, not exceeding maximum weight at a temperature of the approved power, less— standard plus 22 ° C (standard plus 40 (1) Installation losses; and ° F).

(2) The power absorbed by the acces- (3) The out-of-ground effect hovering sories and services appropriate to the performance must be determined over particular ambient atmospheric condi- the ranges of weight, altitude, and tions and the particular flight condi- temperature for which certification is tion. requested, using takeoff power.

(d) For reciprocating engine-powered (b) For rotorcraft other than heli- rotorcraft, the performance, as affected copters, the steady rate of climb at the by engine power, must be based on a minimum operating speed must be de- relative humidity of 80 percent in a termined over the ranges of weight, al- standard atmosphere. titude, and temperature for which cer- (e) For turbine engine-powered rotor- tification is requested, with— craft, the performance, as affected by (1) Takeoff power; and engine power, must be based on a rel- (2) The landing gear extended.

ative humidity of— [Amdt. 27–44, 73 FR 10998, Feb. 29, 2008] (1) 80 percent, at and below standard temperature; and § 27.51 Takeoff.

(2) 34 percent, at and above standard The takeoff, with takeoff power and temperature plus 50 degrees F. Between r.p.m. at the most critical center of these two temperatures, the relative gravity, and with weight from the max- humidity must vary linearly.

imum weight at sea level to the weight (f) For turbine-engine-powered rotor- for which takeoff certification is re- craft, a means must be provided to per- quested for each altitude covered by mit the pilot to determine prior to this section— takeoff that each engine is capable of (a) May not require exceptional pilot- developing the power necessary to ing skill or exceptionally favorable achieve the applicable rotorcraft per- conditions throughout the ranges of al- formance prescribed in this subpart.

titude from standard sea level condi- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- tions to the maximum altitude for eral Aviation Act of 1958 (49 U.S.C. 1354(a), which takeoff and landing certification 1421, 1423, 1424, and 1425); and sec. 6(c) of the is requested, and Dept. of Transportation Act (49 U.S.C.

(b) Must be made in such a manner 1655(c))) that a landing can be made safely at [Amdt. 27–14, 43 FR 2324, Jan. 16, 1978, as any point along the flight path if an amended by Amdt. 27–21, 49 FR 44432, Nov. 6, engine fails. This must be dem- 1984] onstrated up to the maximum altitude for which takeoff and landing certifi- § 27.49 Performance at minimum oper- cation is requested or 7,000 feet density ating speed.

altitude, whichever is less.

(a) For helicopters— [Amdt. 27–44, 73 FR 10999, Feb. 29, 2008] (1) The hovering ceiling must be de- termined over the ranges of weight, al- § 27.65 Climb: all engines operating.

titude, and temperature for which cer- tification is requested, with— (a) For rotorcraft other than heli- (i) Takeoff power; copters— (ii) The landing gear extended; and (1) The steady rate of climb, at V Y, (iii) The helicopter in-ground effect must be determined— at a height consistent with normal (i) With maximum continuous power takeoff procedures; and on each engine; 14 CFR Ch. I (1–1–25 Edition) § 27.67 (ii) With the landing gear retracted; (2) Continuous OEI power for heli- and copters for which certification for the (iii) For the weights, altitudes, and use of continuous OEI power is re- temperatures for which certification is quested.

requested; and (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (2) The climb gradient, at the rate of eral Aviation Act of 1958 (49 U.S.C. 1354(a), climb determined in accordance with 1421, 1423, 1424, and 1425); and sec. 6(c) of the paragraph (a)(1) of this section, must Dept. of Transportation Act (49 U.S.C.

be either— 1655(c))) (i) At least 1:10 if the horizontal dis- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as tance required to take off and climb amended by Amdt. 27–23, 53 FR 34210, Sept. 2, over a 50-foot obstacle is determined 1988] for each weight, altitude, and tempera- ture within the range for which certifi- § 27.71 Autorotation performance.

cation is requested; or (ii) At least 1:6 under standard sea For single-engine helicopters and level conditions. multiengine helicopters that do not (b) Each helicopter must meet the meet the Category A engine isolation following requirements: requirements of Part 29 of this chapter, (1) V must be determined— Y the minimum rate of descent airspeed (i) For standard sea level conditions; and the best angle-of-glide airspeed (ii) At maximum weight; and must be determined in autorotation (iii) With maximum continuous at— power on each engine.

(a) Maximum weight; and (2) The steady rate of climb must be (b) Rotor speed(s) selected by the ap- determined— plicant.

(i) At the climb speed selected by the applicant at or below V ; NE [Amdt. 27–21, 49 FR 44433, Nov. 6, 1984] (ii) Within the range from sea level up to the maximum altitude for which § 27.75 Landing.

certification is requested; (a) The rotorcraft must be able to be (iii) For the weights and tempera- landed with no excessive vertical accel- tures that correspond to the altitude eration, no tendency to bounce, nose range set forth in paragraph (b)(2)(ii) of over, ground loop, porpoise, or water this section and for which certification loop, and without exceptional piloting is requested; and skill or exceptionally favorable condi- (iv) With maximum continuous power tions, with— on each engine.

(1) Approach or autorotation speeds (Secs. 313(a), 601, 603, 604, and 605 of the Fed- appropriate to the type of rotorcraft eral Aviation Act of 1958 (49 U.S.C. 1354(a), and selected by the applicant; 1421, 1423, 1424, and 1425); and sec. 6(c) of the (2) The approach and landing made Dept. of Transportation Act (49 U.S.C.

1655(c))) with— (i) Power off, for single engine rotor- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–14, 43 FR 2324, Jan. 16, craft and entered from steady state 1978; Amdt. 27–33, 61 FR 21907, May 10, 1996] autorotation; or (ii) One-engine inoperative (OEI) for § 27.67 Climb: one engine inoperative.

multiengine rotorcraft, with each oper- For multiengine helicopters, the ating engine within approved operating steady rate of climb (or descent), at V y limitations, and entered from an estab- (or at the speed for minimum rate of lished OEI approach.

descent), must be determined with— (b) Multiengine rotorcraft must be (a) Maximum weight; able to be landed safely after complete (b) The critical engine inoperative power failure under normal operating and the remaining engines at either— conditions.

(1) Maximum continuous power and, for helicopters for which certification [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as for the use of 30-minute OEI power is amended by Amdt. 27–14, 43 FR 2324, Jan. 16, requested, at 30-minute OEI power; or 1978; Amdt. 27–44, 73 FR 10999, Feb. 29, 2008] Federal Aviation Administration, DOT § 27.143 (1) At the altitudes and temperatures § 27.87 Height-velocity envelope.

expected in operation; (a) If there is any combination of (2) Under any critical loading condi- height and forward velocity (including tion within the range of weights and hover) under which a safe landing can- centers of gravity for which certifi- not be made under the applicable power cation is requested; failure condition in paragraph (b) of (3) For power-on operations, under this section, a limiting height-velocity any condition of speed, power, and envelope must be established (includ- rotor r.p.m. for which certification is ing all pertinent information) for that requested; and condition, throughout the ranges of— (4) For power-off operations, under (1) Altitude, from standard sea level any condition of speed and rotor r.p.m.

conditions to the maximum altitude for which certification is requested capability of the rotorcraft, or 7000 feet that is attainable with the controls density altitude, whichever is less; and rigged in accordance with the approved (2) Weight, from the maximum rigging instructions and tolerances; weight at sea level to the weight se- (b) Be able to maintain any required lected by the applicant for each alti- flight condition and make a smooth tude covered by paragraph (a)(1) of this transition from any flight condition to section. For helicopters, the weight at any other flight condition without ex- altitudes above sea level may not be ceptional piloting skill, alertness, or less than the maximum weight or the strength, and without danger of ex- highest weight allowing hovering out- ceeding the limit load factor under any of-ground effect, whichever is lower.

operating condition probable for the (b) The applicable power failure con- type, including— ditions are— (1) Sudden failure of one engine, for (1) For single-engine helicopters, full multiengine rotorcraft meeting Trans- autorotation; port Category A engine isolation re- (2) For multiengine helicopters, OEI quirements of Part 29 of this chapter; (where engine isolation features ensure (2) Sudden, complete power failure continued operation of the remaining for other rotorcraft; and engines), and the remaining engine(s) (3) Sudden, complete control system within approved limits and at the min- failures specified in § 27.695 of this part; imum installed specification power and available for the most critical com- (c) Have any additional char- bination of approved ambient tempera- acteristic required for night or instru- ture and pressure altitude resulting in ment operation, if certification for 7000 feet density altitude or the max- those kinds of operation is requested.

imum altitude capability of the heli- Requirements for helicopter instru- copter, whichever is less, and ment flight are contained in appendix (3) For other rotorcraft, conditions B of this part.

appropriate to the type.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (Secs. 313(a), 601, 603, 604, Federal Aviation amended by Amdt. 27–2, 33 FR 962, Jan. 26, Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), 1968; Amdt. 27–11, 41 FR 55468, Dec. 20, 1976; sec. 6(c), Dept. of Transportation Act (49 Amdt. 27–19, 48 FR 4389, Jan. 31, 1983; Amdt.

U.S.C. 1655(c))) 27–21, 49 FR 44433, Nov. 6, 1984] [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–14, 43 FR 2324, Jan. 16, § 27.143 Controllability and maneuver- 1978; Amdt. 27–21, 49 FR 44433, Nov. 6, 1984; ability.

Amdt. 27–44, 73 FR 10999, Feb. 29, 2008; Amdt.

(a) The rotorcraft must be safely con- 27–51, 88 FR 8737, Feb. 10, 2023] trollable and maneuverable— F LIGHT C HARACTERISTICS (1) During steady flight; and (2) During any maneuver appropriate § 27.141 General.

to the type, including— The rotorcraft must— (i) Takeoff; (a) Except as specifically required in (ii) Climb; the applicable section, meet the flight (iii) Level flight; characteristics requirements of this (iv) Turning flight; subpart— (v) Autorotation; 14 CFR Ch. I (1–1–25 Edition) § 27.151 (vi) Landing (power on and power time delay for any condition following off); and power failure may be less than— (vii) Recovery to power-on flight (i) For the cruise condition, one sec- from a balked autorotative approach.

ond, or normal pilot reaction time (b) The margin of cyclic control must (whichever is greater); and allow satisfactory roll and pitch con- (ii) For any other condition, normal trol at V with— pilot reaction time.

NE (1) Critical weight; (f) For helicopters for which a V NE (2) Critical center of gravity; (power-off) is established under (3) Critical rotor r.p.m.; and § 27.1505(c), compliance must be dem- onstrated with the following require- (4) Power off (except for helicopters ments with critical weight, critical demonstrating compliance with para- center of gravity, and critical rotor graph (f) of this section) and power on.

r.p.m.: (c) Wind velocities from zero to at least 17 knots, from all azimuths, must (1) The helicopter must be safely be established in which the rotorcraft slowed to V (power-off), without ex- NE can be operated without loss of control ceptional pilot skill, after the last op- on or near the ground in any maneuver erating engine is made inoperative at appropriate to the type (such as cross- power-on V NE.

wind takeoffs, sideward flight, and (2) At a speed of 1.1 V (power-off), NE rearward flight)— the margin of cyclic control must allow satisfactory roll and pitch con- (1) With altitude, from standard sea trol with power off.

level conditions to the maximum take- off and landing altitude capability of (Secs. 313(a), 601, 603, 604, and 605 of the Fed- the rotorcraft or 7000 feet density alti- eral Aviation Act of 1958 (49 U.S.C. 1354(a), tude, whichever is less; with— 1421, 1423, 1424, and 1425); and sec. 6(c) of the (i) Critical Weight; Dept. of Transportation Act (49 U.S.C.

1655(c))) (ii) Critical center of gravity; (iii) Critical rotor r.p.m.; [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (2) For takeoff and landing altitudes amended by Amdt. 27–2, 33 FR 963, Jan. 26, above 7000 feet density altitude with— 1968; Amdt. 27–14, 43 FR 2325, Jan. 16, 1978; Amdt. 27–21, 49 FR 44433, Nov. 6, 1984; Amdt.

(i) Weight selected by the applicant; 27–44, 73 FR 10999, Feb. 29, 2008] (ii) Critical center of gravity; and (iii) Critical rotor r.p.m.

§ 27.151 Flight controls.

(d) Wind velocities from zero to at (a) Longitudinal, lateral, directional, least 17 knots, from all azimuths, must and collective controls may not exhibit be established in which the rotorcraft excessive breakout force, friction, or can be operated without loss of control preload.

out-of-ground-effect, with— (1) Weight selected by the applicant; (b) Control system forces and free play may not inhibit a smooth, direct (2) Critical center of gravity; rotorcraft response to control system (3) Rotor r.p.m. selected by the appli- input.

cant; and (4) Altitude, from standard sea level [Amdt. 27–21, 49 FR 44433, Nov. 6, 1984] conditions to the maximum takeoff and landing altitude capability of the § 27.161 Trim control.

rotorcraft.

The trim control— (e) The rotorcraft, after (1) failure of (a) Must trim any steady longitu- one engine in the case of multiengine dinal, lateral, and collective control rotorcraft that meet Transport Cat- forces to zero in level flight at any ap- egory A engine isolation requirements, propriate speed; and or (2) complete engine failure in the (b) May not introduce any undesir- case of other rotorcraft, must be con- able discontinuities in control force trollable over the range of speeds and gradients.

altitudes for which certification is re- quested when such power failure occurs [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as with maximum continuous power and amended by Amdt. 27–21, 49 FR 44433, Nov. 6, critical weight. No corrective action 1984] Federal Aviation Administration, DOT § 27.177 (4) The landing gear retracted; and § 27.171 Stability: general.

(5) The rotorcraft trimmed at 0.8 V NE The rotorcraft must be able to be or V , whichever is less.

H flown, without undue pilot fatigue or (c) V Static longitudinal stability NE.

strain, in any normal maneuver for a must be shown at speeds from V ¥ 20 NE period of time as long as that expected kt to V with— NE in normal operation. At least three (1) Critical weight; landings and takeoffs must be made (2) Critical center of gravity; during this demonstration.

(3) Power required for level flight at V ¥ 10 kt or maximum continuous NE § 27.173 Static longitudinal stability.

power, whichever is less; (a) The longitudinal control must be (4) The landing gear retracted; and designed so that a rearward movement (5) The rotorcraft trimmed at V ¥ NE of the control is necessary to obtain an 10 kt.

airspeed less than the trim speed, and a (d) Autorotation. Static longitudinal forward movement of the control is stability must be shown in autorota- necessary to obtain an airspeed more tion at— than the trim speed.

(1) Airspeeds from the minimum rate (b) Throughout the full range of alti- of descent airspeed ¥ 10 kt to the min- tude for which certification is re- imum rate of descent airspeed + 10 kt, quested, with the throttle and collec- with— tive pitch held constant during the ma- (i) Critical weight; neuvers specified in § 27.175(a) through (ii) Critical center of gravity; (d), the slope of the control position (iii) The landing gear extended; and versus airspeed curve must be positive.

(iv) The rotorcraft trimmed at the However, in limited flight conditions minimum rate of descent airspeed.

or modes of operation determined by (2) Airspeeds from best angle-of-glide the Administrator to be acceptable, the airspeed ¥ 10 kt to the best angle-of- slope of the control position versus air- glide airspeed + 10 kt, with— speed curve may be neutral or negative (i) Critical weight; if the rotorcraft possesses flight char- (ii) Critical center of gravity; acteristics that allow the pilot to (iii) The landing gear retracted; and maintain airspeed within ± 5 knots of (iv) The rotorcraft trimmed at the the desired trim airspeed without ex- best angle-of-glide airspeed.

ceptional piloting skill or alertness.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- [Amdt. 27–21, 49 FR 44433, Nov. 6, 1984, as eral Aviation Act of 1958 (49 U.S.C. 1354(a), amended by Amdt. 27–44, 73 FR 10999, Feb. 29, 1421, 1423, 1424, and 1425); and sec. 6(c) of the 2008] Dept. of Transportation Act (49 U.S.C.

1655(c))) § 27.175 Demonstration of static longi- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as tudinal stability.

amended by Amdt. 27–2, 33 FR 963, Jan. 26, (a) Climb. Static longitudinal sta- 1968; Amdt. 27–11, 41 FR 55468, Dec. 20, 1976; bility must be shown in the climb con- Amdt. 27–14, 43 FR 2325, Jan. 16, 1978; Amdt.

dition at speeds from Vy ¥ 10 kt to Vy 27–21, 49 FR 44433, Nov. 6, 1984; Amdt. 27–34, 62 FR 46173, Aug. 29, 1997; Amdt. 27–44, 73 FR + 10 kt with— 10999, Feb. 29, 2008] (1) Critical weight; (2) Critical center of gravity; § 27.177 Static directional stability.

(3) Maximum continuous power; (a) The directional controls must op- (4) The landing gear retracted; and erate in such a manner that the sense (5) The rotorcraft trimmed at V Y.

and direction of motion of the rotor- (b) Cruise. Static longitudinal sta- craft following control displacement bility must be shown in the cruise con- are in the direction of the pedal motion dition at speeds from 0.8 V ¥ 10 kt to NE with the throttle and collective con- 0.8 V + 10 kt or, if V is less than 0.8 NE H trols held constant at the trim condi- V , from V ¥ 10 kt to V + 10 kt, NE H H tions specified in § 27.175(a), (b), and (c).

with— (1) Critical weight; Sideslip angles must increase with (2) Critical center of gravity; steadily increasing directional control (3) Power for level flight at 0.8 V or deflection for sideslip angles up to the NE V , whichever is less; lesser of— H 14 CFR Ch. I (1–1–25 Edition) § 27.231 (1) ± 25 degrees from trim at a speed of M ISCELLANEOUS F LIGHT R EQUIREMENTS 15 knots less than the speed for min- § 27.251 Vibration.

imum rate of descent varying linearly to ± 10 degrees from trim at V ; NE Each part of the rotorcraft must be (2) The steady state sideslip angles free from excessive vibration under established by § 27.351; each appropriate speed and power con- (3) A sideslip angle selected by the dition.

applicant, which corresponds to a sideforce of at least 0.1g; or Subpart C—Strength Requirements (4) The sideslip angle attained by maximum directional control input. G ENERAL (b) Sufficient cues must accompany § 27.301 Loads.

the sideslip to alert the pilot when the aircraft is approaching the sideslip (a) Strength requirements are speci- limits.

fied in terms of limit loads (the max- (c) During the maneuver specified in imum loads to be expected in service) paragraph (a) of this section, the side- and ultimate loads (limit loads multi- slip angle versus directional control plied by prescribed factors of safety).

position curve may have a negative Unless otherwise provided, prescribed slope within a small range of angles loads are limit loads.

around trim, provided the desired head- (b) Unless otherwise provided, the ing can be maintained without excep- specified air, ground, and water loads tional piloting skill or alertness.

must be placed in equilibrium with in- ertia forces, considering each item of [Amdt. 27–44, 73 FR 11000, Feb. 29, 2008] mass in the rotorcraft. These loads must be distributed to closely approxi- G ROUND AND W ATER H ANDLING mate or conservatively represent ac- C HARACTERISTICS tual conditions.

§ 27.231 General.

(c) If deflections under load would significantly change the distribution of The rotorcraft must have satisfac- external or internal loads, this redis- tory ground and water handling char- tribution must be taken into account.

acteristics, including freedom from un- controllable tendencies in any condi- § 27.303 Factor of safety.

tion expected in operation.

Unless otherwise provided, a factor of § 27.235 Taxiing condition.

safety of 1.5 must be used. This factor applies to external and inertia loads The rotorcraft must be designed to unless its application to the resulting withstand the loads that would occur internal stresses is more conservative.

when the rotorcraft is taxied over the roughest ground that may reasonably § 27.305 Strength and deformation.

be expected in normal operation.

(a) The structure must be able to § 27.239 Spray characteristics.

support limit loads without detri- mental or permanent deformation. At If certification for water operation is any load up to limit loads, the defor- requested, no spray characteristics mation may not interfere with safe op- during taxiing, takeoff, or landing may eration.

obscure the vision of the pilot or dam- (b) The structure must be able to age the rotors, propellers, or other support ultimate loads without failure.

parts of the rotorcraft.

This must be shown by— § 27.241 Ground resonance. (1) Applying ultimate loads to the structure in a static test for at least The rotorcraft may have no dan- three seconds; or gerous tendency to oscillate on the (2) Dynamic tests simulating actual ground with the rotor turning.

load application.

Federal Aviation Administration, DOT § 27.339 F LIGHT L OADS § 27.307 Proof of structure.

(a) Compliance with the strength and § 27.321 General.

deformation requirements of this sub- (a) The flight load factor must be as- part must be shown for each critical sumed to act normal to the longitu- loading condition accounting for the dinal axis of the rotorcraft, and to be environment to which the structure equal in magnitude and opposite in di- will be exposed in operation. Struc- rection to the rotorcraft inertia load tural analysis (static or fatigue) may factor at the center of gravity.

be used only if the structure conforms (b) Compliance with the flight load to those structures for which experi- requirements of this subpart must be ence has shown this method to be reli- shown— able. In other cases, substantiating (1) At each weight from the design load tests must be made.

minimum weight to the design max- (b) Proof of compliance with the imum weight; and strength requirements of this subpart (2) With any practical distribution of must include— disposable load within the operating limitations in the Rotorcraft Flight (1) Dynamic and endurance tests of Manual.

rotors, rotor drives, and rotor controls; (2) Limit load tests of the control [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as system, including control surfaces; amended by Amdt. 27–11, 41 FR 55468, Dec. 20, 1976] (3) Operation tests of the control sys- tem; § 27.337 Limit maneuvering load fac- (4) Flight stress measurement tests; tor.

(5) Landing gear drop tests; and The rotorcraft must be designed for— (6) Any additional test required for (a) A limit maneuvering load factor new or unusual design features.

ranging from a positive limit of 3.5 to (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, a negative limit of ¥ 1.0; or 1425) (b) Any positive limit maneuvering load factor not less than 2.0 and any [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as negative limit maneuvering load factor amended by Amdt. 27–3, 33 FR 14105, Sept. 18, of not less than ¥ 0.5 for which— 1968; Amdt. 27–26, 55 FR 7999, Mar. 6, 1990] (1) The probability of being exceeded § 27.309 Design limitations.

is shown by analysis and flight tests to be extremely remote; and The following values and limitations (2) The selected values are appro- must be established to show compli- priate to each weight condition be- ance with the structural requirements tween the design maximum and design of this subpart: minimum weights.

(a) The design maximum weight.

[Amdt. 27–26, 55 FR 7999, Mar. 6, 1990] (b) The main rotor r.p.m. ranges power on and power off.

§ 27.339 Resultant limit maneuvering (c) The maximum forward speeds for loads.

each main rotor r.p.m. within the The loads resulting from the applica- ranges determined under paragraph (b) tion of limit maneuvering load factors of this section.

are assumed to act at the center of (d) The maximum rearward and side- each rotor hub and at each auxiliary ward flight speeds.

lifting surface, and to act in directions, (e) The center of gravity limits cor- and with distributions of load among responding to the limitations deter- the rotors and auxiliary lifting sur- mined under paragraphs (b), (c), and (d) faces, so as to represent each critical of this section.

maneuvering condition, including (f) The rotational speed ratios be- power-on and power-off flight with the tween each powerplant and each con- maximum design rotor tip speed ratio.

nected rotating component.

The rotor tip speed ratio is the ratio of (g) The positive and negative limit the rotorcraft flight velocity compo- maneuvering load factors. nent in the plane of the rotor disc to 14 CFR Ch. I (1–1–25 Edition) § 27.341 the rotational tip speed of the rotor by the maximum pilot force specified blades, and is expressed as follows: in § 27.397(a); (2) Attain a resulting sideslip angle V cos a or 15 ° , whichever is less, at the lesser μ = speed of V or V ; NE H R Ω (3) Vary the sideslip angles of para- where— graphs (b)(2) and (c)(2) of this section V = The airspeed along flight path (f.p.s.); directly with speed; and a = The angle between the projection, in the (4) Return the directional control plane of symmetry, of the axis of no suddenly to neutral.

feathering and a line perpendicular to the flight path (radians, positive when [Amdt. 27–26, 55 FR 7999, Mar. 6, 1990, as axis is pointing aft); amended by Amdt. 27–34, 62 FR 46173, Aug. 29, omega = The angular velocity of rotor (radi- 1997] ans per second); and R = The rotor radius (ft).

§ 27.361 Engine torque.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (a) For turbine engines, the limit amended by Amdt. 27–11, 41 FR 55469, Dec. 20, torque may not be less than the high- 1976] est of— (1) The mean torque for maximum § 27.341 Gust loads.

continuous power multiplied by 1.25; The rotorcraft must be designed to (2) The torque required by § 27.923; withstand, at each critical airspeed in- (3) The torque required by § 27.927; or cluding hovering, the loads resulting (4) The torque imposed by sudden en- from a vertical gust of 30 feet per sec- gine stoppage due to malfunction or ond.

structural failure (such as compressor jamming).

§ 27.351 Yawing conditions.

(b) For reciprocating engines, the (a) Each rotorcraft must be designed limit torque may not be less than the for the loads resulting from the maneu- mean torque for maximum continuous vers specified in paragraphs (b) and (c) power multiplied by— of this section with— (1) 1.33, for engines with five or more (1) Unbalanced aerodynamic mo- cylinders; and ments about the center of gravity (2) Two, three, and four, for engines which the aircraft reacts to in a ration- with four, three, and two cylinders, re- al or conservative manner considering spectively.

the principal masses furnishing the re- [Amdt. 27–23, 53 FR 34210, Sept. 2, 1988] acting inertia forces; and (2) Maximum main rotor speed.

C ONTROL S URFACE AND S YSTEM L OADS (b) To produce the load required in paragraph (a) of this section, in unac- § 27.391 General.

celerated flight with zero yaw, at for- Each auxiliary rotor, each fixed or ward speeds from zero up to 0.6 V — NE movable stabilizing or control surface, (1) Displace the cockpit directional and each system operating any flight control suddenly to the maximum de- control must meet the requirements of flection limited by the control stops or §§ 27.395, 27.397, 27.399, 27.411, and 27.427.

by the maximum pilot force specified in § 27.397(a); [Amdt. 27–26, 55 FR 7999, Mar. 6, 1990, as (2) Attain a resulting sideslip angle amended by Amdt. 27–34, 62 FR 46173, Aug. 29, 1997] or 90 ° , whichever is less; and (3) Return the directional control § 27.395 Control system.

suddenly to neutral.

(c) To produce the load required in (a) The part of each control system paragraph (a) of this section, in unac- from the pilot’s controls to the control celerated flight with zero yaw, at for- stops must be designed to withstand ward speeds from 0.6 V up to V or pilot forces of not less than— NE NE V , whichever is less— (1) The forces specified in § 27.397; or H (1) Displace the cockpit directional (2) If the system prevents the pilot control suddenly to the maximum de- from applying the limit pilot forces to flection limited by the control stops or the system, the maximum forces that Federal Aviation Administration, DOT § 27.427 the system allows the pilot to apply, any angle within 20 degrees of the but not less than 0.60 times the forces plane of motion of the control.

specified in § 27.397. (2) Twist controls, 80R inch-pounds.

(b) Each primary control system, in- [Amdt. 27–11, 41 FR 55469, Dec. 20, 1976, as cluding its supporting structure, must amended by Amdt. 27–40, 66 FR 23538, May 9, be designed as follows: 2001] (1) The system must withstand loads resulting from the limit pilot forces § 27.399 Dual control system.

prescribed in § 27.397.

Each dual primary flight control sys- (2) Notwithstanding paragraph (b)(3) tem must be designed to withstand the of this section, when power-operated loads that result when pilot forces of actuator controls or power boost con- 0.75 times those obtained under § 27.395 trols are used, the system must also are applied— withstand the loads resulting from the (a) In opposition; and force output of each normally ener- (b) In the same direction.

gized power device, including any sin- gle power boost or actuator system § 27.411 Ground clearance: tail rotor guard.

failure.

(3) If the system design or the normal (a) It must be impossible for the tail operating loads are such that a part of rotor to contact the landing surface the system cannot react to the limit during a normal landing.

pilot forces prescribed in § 27.397, that (b) If a tail rotor guard is required to part of the system must be designed to show compliance with paragraph (a) of withstand the maximum loads that can this section— be obtained in normal operation. The (1) Suitable design loads must be es- minimum design loads must, in any tablished for the guard; and case, provide a rugged system for serv- (2) The guard and its supporting ice use, including consideration of fa- structure must be designed to with- tigue, jamming, ground gusts, control stand those loads.

inertia, and friction loads. In the ab- § 27.427 Unsymmetrical loads.

sence of rational analysis, the design loads resulting from 0.60 of the speci- (a) Horizontal tail surfaces and their fied limit pilot forces are acceptable supporting structure must be designed minimum design loads.

for unsymmetrical loads arising from (4) If operational loads may be ex- yawing and rotor wake effects in com- ceeded through jamming, ground gusts, bination with the prescribed flight con- control inertia, or friction, the system ditions.

must withstand the limit pilot forces (b) To meet the design criteria of specified in § 27.397, without yielding.

paragraph (a) of this section, in the ab- sence of more rational data, both of the [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–26, 55 FR 7999, Mar. 6, following must be met: 1990] (1) One hundred percent of the max- imum loading from the symmetrical § 27.397 Limit pilot forces and torques.

flight conditions acts on the surface on one side of the plane of symmetry, and (a) Except as provided in paragraph no loading acts on the other side.

(b) of this section, the limit pilot (2) Fifty percent of the maximum forces are as follows: loading from the symmetrical flight (1) For foot controls, 130 pounds.

(2) For stick controls, 100 pounds fore conditions acts on the surface on each and aft, and 67 pounds laterally. side of the plane of symmetry but in (b) For flap, tab, stabilizer, rotor opposite directions.

brake, and landing gear operating con- (c) For empennage arrangements trols, the follows apply (R = radius in where the horizontal tail surfaces are inches): supported by the vertical tail surfaces, (1) Crank, wheel, and lever controls, the vertical tail surfaces and sup- [1 + R]/3 × 50 pounds, but not less than porting structure must be designed for 50 pounds nor more than 100 pounds for the combined vertical and horizontal hand operated controls or 130 pounds surface loads resulting from each pre- for foot operated controls, applied at scribed flight condition, considered 14 CFR Ch. I (1–1–25 Edition) § 27.471 separately. The flight conditions must § 27.477 Landing gear arrangement.

be selected so the maximum design Sections 27.235, 27.479 through 27.485, loads are obtained on each surface. In and 27.493 apply to landing gear with the absence of more rational data, the two wheels aft, and one or more wheels unsymmetrical horizontal tail surface forward, of the center of gravity.

loading distributions described in this section must be assumed.

§ 27.479 Level landing conditions.

[Amdt. 27–26, 55 FR 7999, Mar. 6, 1990, as (a) Attitudes. Under each of the load- amended by Amdt. 27–27, 55 FR 38966, Sept.

ing conditions prescribed in paragraph 21, 1990] (b) of this section, the rotorcraft is as- sumed to be in each of the following G ROUND LOADS level landing attitudes: (1) An attitude in which all wheels § 27.471 General.

contact the ground simultaneously.

(a) Loads and equilibrium. For limit (2) An attitude in which the aft ground loads— wheels contact the ground with the for- (1) The limit ground loads obtained ward wheels just clear of the ground.

in the landing conditions in this part (b) Loading conditions. The rotorcraft must be considered to be external loads must be designed for the following that would occur in the rotorcraft landing loading conditions: structure if it were acting as a rigid (1) Vertical loads applied under body; and § 27.471.

(2) In each specified landing condi- (2) The loads resulting from a com- tion, the external loads must be placed bination of the loads applied under in equilibrium with linear and angular paragraph (b)(1) of this section with inertia loads in a rational or conserv- drag loads at each wheel of not less ative manner.

than 25 percent of the vertical load at (b) Critical centers of gravity. The crit- that wheel.

ical centers of gravity within the range (3) If there are two wheels forward, a for which certification is requested distribution of the loads applied to must be selected so that the maximum those wheels under paragraphs (b)(1) design loads are obtained in each land- and (2) of this section in a ratio of ing gear element.

40:60.

§ 27.473 Ground loading conditions (c) Pitching moments. Pitching mo- and assumptions.

ments are assumed to be resisted by— (1) In the case of the attitude in para- (a) For specified landing conditions, graph (a)(1) of this section, the forward a design maximum weight must be landing gear; and used that is not less than the max- (2) In the case of the attitude in para- imum weight. A rotor lift may be as- graph (a)(2) of this section, the angular sumed to act through the center of inertia forces.

gravity throughout the landing impact.

This lift may not exceed two-thirds of [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964; 29 the design maximum weight.

FR 17885, Dec. 17, 1964] (b) Unless otherwise prescribed, for each specified landing condition, the § 27.481 Tail-down landing conditions.

rotorcraft must be designed for a limit (a) The rotorcraft is assumed to be in load factor of not less than the limit the maximum nose-up attitude allow- inertia load factor substantiated under ing ground clearance by each part of § 27.725.

the rotorcraft.

(b) In this attitude, ground loads are [Amdt. 27–2, 33 FR 963, Jan. 26, 1968] assumed to act perpendicular to the § 27.475 Tires and shock absorbers.

ground.

Unless otherwise prescribed, for each § 27.483 One-wheel landing conditions.

specified landing condition, the tires must be assumed to be in their static For the one-wheel landing condition, position and the shock absorbers to be the rotorcraft is assumed to be in the in their most critical position. level attitude and to contact the Federal Aviation Administration, DOT § 27.497 ground on one aft wheel. In this atti- § 27.497 Ground loading conditions: tude— landing gear with tail wheels.

(a) The vertical load must be the (a) General. Rotorcraft with landing same as that obtained on that side gear with two wheels forward, and one under § 27.479(b)(1); and wheel aft, of the center of gravity must (b) The unbalanced external loads be designed for loading conditions as must be reacted by rotorcraft inertia.

prescribed in this section.

(b) Level landing attitude with only the § 27.485 Lateral drift landing condi- forward wheels contacting the ground. In tions.

this attitude— (a) The rotorcraft is assumed to be in (1) The vertical loads must be applied the level landing attitude, with— under §§ 27.471 through 27.475; (1) Side loads combined with one-half (2) The vertical load at each axle of the maximum ground reactions ob- must be combined with a drag load at tained in the level landing conditions that axle of not less than 25 percent of of § 27.479 (b)(1); and that vertical load; and (2) The loads obtained under para- (3) Unbalanced pitching moments are graph (a)(1) of this section applied— assumed to be resisted by angular iner- (i) At the ground contact point; or tia forces.

(ii) For full-swiveling gear, at the (c) Level landing attitude with all center of the axle.

wheels contacting the ground simulta- (b) The rotorcraft must be designed neously. In this attitude, the rotorcraft to withstand, at ground contact— must be designed for landing loading (1) When only the aft wheels contact conditions as prescribed in paragraph the ground, side loads of 0.8 times the (b) of this section.

vertical reaction acting inward on one (d) Maximum nose-up attitude with side, and 0.6 times the vertical reaction only the rear wheel contacting the acting outward on the other side, all ground. The attitude for this condition combined with the vertical loads speci- must be the maximum nose-up attitude fied in paragraph (a) of this section; expected in normal operation, includ- and ing autorotative landings. In this atti- (2) When all wheels contact the tude— ground simultaneously— (1) The appropriate ground loads (i) For the aft wheels, the side loads specified in paragraphs (b)(1) and (2) of specified in paragraph (b)(1) of this sec- this section must be determined and tion; and applied, using a rational method to ac- (ii) For the forward wheels, a side count for the moment arm between the load of 0.8 times the vertical reaction rear wheel ground reaction and the combined with the vertical load speci- rotorcraft center of gravity; or fied in paragraph (a) of this section.

(2) The probability of landing with § 27.493 Braked roll conditions.

initial contact on the rear wheel must be shown to be extremely remote.

Under braked roll conditions with (e) Level landing attitude with only one the shock absorbers in their static po- forward wheel contacting the ground. In sitions— this attitude, the rotorcraft must be (a) The limit vertical load must be designed for ground loads as specified based on a load factor of at least— in paragraphs (b)(1) and (3) of this sec- (1) 1.33, for the attitude specified in tion.

§ 27.479(a)(1); and (f) Side loads in the level landing atti- (2) 1.0 for the attitude specified in tude. In the attitudes specified in para- § 27.479(a)(2); and graphs (b) and (c) of this section, the (b) The structure must be designed to following apply: withstand at the ground contact point of each wheel with brakes, a drag load (1) The side loads must be combined at least the lesser of— at each wheel with one-half of the max- (1) The vertical load multiplied by a imum vertical ground reactions ob- coefficient of friction of 0.8; and tained for that wheel under paragraphs (2) The maximum value based on lim- (b) and (c) of this section. In this condi- iting brake torque. tion, the side loads must be— 14 CFR Ch. I (1–1–25 Edition) § 27.501 (i) For the forward wheels, 0.8 times my damper (the rear wheel being as- the vertical reaction (on one side) act- sumed to be in the trailing position).

ing inward, and 0.6 times the vertical (i) Taxiing condition. The rotorcraft reaction (on the other side) acting out- and its landing gear must be designed ward; and for loads that would occur when the (ii) For the rear wheel, 0.8 times the rotorcraft is taxied over the roughest vertical reaction.

ground that may reasonably be ex- (2) The loads specified in paragraph pected in normal operation.

(f)(1) of this section must be applied— (i) At the ground contact point with § 27.501 Ground loading conditions: the wheel in the trailing position (for landing gear with skids.

non-full swiveling landing gear or for (a) General. Rotorcraft with landing full swiveling landing gear with a lock, gear with skids must be designed for steering device, or shimmy damper to the loading conditions specified in this keep the wheel in the trailing posi- section. In showing compliance with tion); or this section, the following apply: (ii) At the center of the axle (for full (1) The design maximum weight, cen- swiveling landing gear without a lock, ter of gravity, and load factor must be steering device, or shimmy damper).

determined under §§ 27.471 through (g) Braked roll conditions in the level 27.475.

landing attitude. In the attitudes speci- (2) Structural yielding of elastic fied in paragraphs (b) and (c) of this spring members under limit loads is ac- section, and with the shock absorbers ceptable.

in their static positions, the rotorcraft (3) Design ultimate loads for elastic must be designed for braked roll loads spring members need not exceed those as follows: obtained in a drop test of the gear (1) The limit vertical load must be based on a limit vertical load factor of with— not less than— (i) A drop height of 1.5 times that (i) 1.0, for the attitude specified in specified in § 27.725; and paragraph (b) of this section; and (ii) An assumed rotor lift of not more (ii) 1.33, for the attitude specified in than 1.5 times that used in the limit paragraph (c) of this section.

drop tests prescribed in § 27.725.

(2) For each wheel with brakes, a (4) Compliance with paragraphs (b) drag load must be applied, at the through (e) of this section must be ground contact point, of not less than shown with— the lesser of— (i) The gear in its most critically de- (i) 0.8 times the vertical load; and flected position for the landing condi- (ii) The maximum based on limiting tion being considered; and brake torque.

(ii) The ground reactions rationally (h) Rear wheel turning loads in the distributed along the bottom of the static ground attitude. In the static skid tube.

ground attitude, and with the shock (b) Vertical reactions in the level land- absorbers and tires in their static posi- ing attitude. In the level attitude, and tions, the rotorcraft must be designed with the rotorcraft contacting the for rear wheel turning loads as follows: ground along the bottom of both skids, (1) A vertical ground reaction equal the vertical reactions must be applied to the static load on the rear wheel as prescribed in paragraph (a) of this must be combined with an equal section.

sideload.

(c) Drag reactions in the level landing (2) The load specified in paragraph attitude. In the level attitude, and with (h)(1) of this section must be applied to the rotorcraft contacting the ground the rear landing gear— along the bottom of both skids, the fol- (i) Through the axle, if there is a lowing apply: swivel (the rear wheel being assumed to be swiveled 90 degrees to the longi- (1) The vertical reactions must be tudinal axis of the rotorcraft); or combined with horizontal drag reac- (ii) At the ground contact point, if tions of 50 percent of the vertical reac- there is a lock, steering device or shim- tion applied at the ground.

Federal Aviation Administration, DOT § 27.521 (2) The resultant ground loads must (2) With the rotorcraft in the level equal the vertical load specified in landing attitude, a vertical ground re- paragraph (b) of this section. action load equal to one-half of the (d) Sideloads in the level landing atti- vertical load determined under para- tude. In the level attitude,and with the graph (b) of this section. This load rotorcraft contacting the ground along must be— the bottom of both skids, the following (i) Applied only to the skid tube and apply: its attachment to the rotorcraft; and (1) The vertical ground reaction must (ii) Distributed equally over 33.3 per- be— cent of the length between the skid (i) Equal to the vertical loads ob- tube attachments and centrally located tained in the condition specified in midway between the skid tube attach- paragraph (b) of this section; and ments.

(ii) Divided equally among the skids.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (2) The vertical ground reactions amended by Amdt. 27–2, 33 FR 963, Jan. 26, must be combined with a horizontal 1968; Amdt. 27–26, 55 FR 8000, Mar. 6, 1990] sideload of 25 percent of their value.

(3) The total sideload must be applied § 27.505 Ski landing conditions.

equally between the skids and along If certification for ski operation is the length of the skids.

requested, the rotorcraft, with skis, (4) The unbalanced moments are as- must be designed to withstand the fol- sumed to be resisted by angular iner- lowing loading conditions (where P is tia.

the maximum static weight on each ski (5) The skid gear must be inves- with the rotorcraft at design maximum tigated for— weight, and n is the limit load factor (i) Inward acting sideloads; and determined under § 27.473(b).

(ii) Outward acting sideloads.

(a) Up-load conditions in which— (e) One-skid landing loads in the level (1) A vertical load of Pn and a hori- attitude. In the level attitude, and with zontal load of Pn/ 4 are simultaneously the rotorcraft contacting the ground applied at the pedestal bearings; and along the bottom of one skid only, the (2) A vertical load of 1.33 P is applied following apply: at the pedestal bearings.

(1) The vertical load on the ground (b) A side-load condition in which a contact side must be the same as that side load of 0.35 Pn is applied at the obtained on that side in the condition pedestal bearings in a horizontal plane specified in paragraph (b) of this sec- perpendicular to the centerline of the tion.

rotorcraft.

(2) The unbalanced moments are as- (c) A torque-load condition in which sumed to be resisted by angular iner- a torque load of 1.33 P (in foot pounds) tia.

is applied to the ski about the vertical (f) Special conditions. In addition to axis through the centerline of the ped- the conditions specified in paragraphs estal bearings.

(b) and (c) of this section, the rotor- craft must be designed for the fol- W ATER L OADS lowing ground reactions: (1) A ground reaction load acting up § 27.521 Float landing conditions.

and aft at an angle of 45 degrees to the If certification for float operation is longitudinal axis of the rotorcraft.

requested, the rotorcraft, with floats, This load must be— must be designed to withstand the fol- (i) Equal to 1.33 times the maximum lowing loading conditions (where the weight; limit load factor is determined under (ii) Distributed symmetrically among § 27.473(b) or assumed to be equal to the skids; that determined for wheel landing (iii) Concentrated at the forward end gear): of the straight part of the skid tube; and (a) Up-load conditions in which— (iv) Applied only to the forward end (1) A load is applied so that, with the of the skid tube and its attachment to rotorcraft in the static level attitude, the rotorcraft. the resultant water reaction passes 14 CFR Ch. I (1–1–25 Edition) § 27.547 vertically through the center of grav- (ii) The limit engine torque specified ity; and in § 27.361.

(2) The limit torque must be distrib- (2) The vertical load prescribed in uted to the rotor blades in a rational paragraph (a)(1) of this section is ap- manner.

plied simultaneously with an aft com- ponent of 0.25 times the vertical com- (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, ponent.

1425) (b) A side-load condition in which— [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (1) A vertical load of 0.75 times the amended by Amdt. 27–3, 33 FR 14105, Sept. 18, total vertical load specified in para- 1968] graph (a)(1) of this section is divided § 27.549 Fuselage, landing gear, and equally among the floats; and rotor pylon structures.

(2) For each float, the load share de- termined under paragraph (b)(1) of this (a) Each fuselage, landing gear, and section, combined with a total side rotor pylon structure must be designed load of 0.25 times the total vertical as prescribed in this section. Resultant load specified in paragraph (b)(1) of rotor forces may be represented as a this section, is applied to that float single force applied at the rotor hub at- only. tachment point.

(b) Each structure must be designed M AIN C OMPONENT R EQUIREMENTS to withstand— (1) The critical loads prescribed in § 27.547 Main rotor structure.

§§ 27.337 through 27.341; (2) The applicable ground loads pre- (a) Each main rotor assembly (in- scribed in §§ 27.235, 27.471 through 27.485, cluding rotor hubs and blades) must be 27.493, 27.497, 27.501, 27.505, and 27.521; designed as prescribed in this section.

and (b) [Reserved] (3) The loads prescribed in § 27.547 (c) The main rotor structure must be (d)(2) and (e).

designed to withstand the following (c) Auxiliary rotor thrust, and the loads prescribed in §§ 27.337 through balancing air and inertia loads occur- 27.341: ring under accelerated flight condi- (1) Critical flight loads.

tions, must be considered.

(2) Limit loads occurring under nor- (d) Each engine mount and adjacent mal conditions of autorotation. For fuselage structure must be designed to this condition, the rotor r.p.m. must be withstand the loads occurring under selected to include the effects of alti- accelerated flight and landing condi- tude.

tions, including engine torque.

(d) The main rotor structure must be (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, designed to withstand loads simu- 1425) lating— [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (1) For the rotor blades, hubs, and amended by Amdt. 27–3, 33 FR 14105, Sept. 18, flapping hinges, the impact force of 1968] each blade against its stop during ground operation; and E MERGENCY L ANDING C ONDITIONS (2) Any other critical condition ex- pected in normal operation. § 27.561 General.

(e) The main rotor structure must be (a) The rotorcraft, although it may designed to withstand the limit torque be damaged in emergency landing con- at any rotational speed, including zero.

ditions on land or water, must be de- In addition: signed as prescribed in this section to (1) The limit torque need not be protect the occupants under those con- greater than the torque defined by a ditions.

torque limiting device (where pro- (b) The structure must be designed to vided), and may not be less than the give each occupant every reasonable greater of— chance of escaping serious injury in a (i) The maximum torque likely to be crash landing when— transmitted to the rotor structure in (1) Proper use is made of seats, belts, either direction; and and other safety design provisions; Federal Aviation Administration, DOT § 27.562 (2) The wheels are retracted (where (2) The occupant is exposed to the applicable); and loads resulting from the conditions prescribed in this section.

(3) Each occupant and each item of (b) Each seat type design or other mass inside the cabin that could injure seating device approved for crew or an occupant is restrained when sub- passenger occupancy during takeoff jected to the following ultimate iner- and landing must successfully com- tial load factors relative to the sur- plete dynamic tests or be demonstrated rounding structure: by rational analysis based on dynamic (i) Upward—4g.

tests of a similar type seat in accord- (ii) Forward—16g.

ance with the following criteria. The (iii) Sideward—8g.

tests must be conducted with an occu- (iv) Downward—20g, after intended pant, simulated by a 170-pound displacement of the seat device.

anthropomorphic test dummy (ATD), (v) Rearward—1.5g.

as defined by 49 CFR 572, subpart B, or (c) The supporting structure must be its equivalent, sitting in the normal designed to restrain, under any ulti- upright position.

mate inertial load up to those specified (1) A change in downward velocity of in this paragraph, any item of mass not less than 30 feet per second when above and/or behind the crew and pas- the seat or other seating device is ori- senger compartment that could injure ented in its nominal position with re- an occupant if it came loose in an spect to the rotorcraft’s reference sys- emergency landing. Items of mass to be tem, the rotorcraft’s longitudinal axis considered include, but are not limited is canted upward 60 ° with respect to to, rotors, transmissions, and engines.

the impact velocity vector, and the The items of mass must be restrained rotorcraft’s lateral axis is perpen- for the following ultimate inertial load dicular to a vertical plane containing factors: the impact velocity vector and the (1) Upward—1.5g.

rotorcraft’s longitudinal axis. Peak (2) Forward—12g.

floor deceleration must occur in not (3) Sideward—6g.

more than 0.031 seconds after impact (4) Downward—12g. and must reach a minimum of 30g’s.

(2) A change in forward velocity of (5) Rearward—1.5g not less than 42 feet per second when (d) Any fuselage structure in the area the seat or other seating device is ori- of internal fuel tanks below the pas- ented in its nominal position with re- senger floor level must be designed to spect to the rotorcraft’s reference sys- resist the following ultimate inertial tem, the rotorcraft’s longitudinal axis factors and loads and to protect the is yawed 10 ° either right or left of the fuel tanks from rupture when those impact velocity vector (whichever loads are applied to that area: would cause the greatest load on the (i) Upward—1.5g.

shoulder harness), the rotorcraft’s lat- (ii) Forward—4.0g.

eral axis is contained in a horizontal (iii) Sideward—2.0g.

plane containing the impact velocity (iv) Downward—4.0g.

vector, and the rotorcraft’s vertical axis is perpendicular to a horizontal [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–25, 54 FR 47318, Nov. 13, plane containing the impact velocity 1989; Amdt. 27–30, 59 FR 50386, Oct. 3, 1994; vector. Peak floor deceleration must Amdt. 27–32, 61 FR 10438, Mar. 13, 1996] occur in not more than 0.071 seconds after impact and must reach a min- § 27.562 Emergency landing dynamic imum of 18.4g’s.

conditions.

(3) Where floor rails or floor or side- (a) The rotorcraft, although it may wall attachment devices are used to at- be damaged in an emergency crash tach the seating devices to the air- landing, must be designed to reason- frame structure for the conditions of ably protect each occupant when— this section, the rails or devices must (1) The occupant properly uses the be misaligned with respect to each seats, safety belts, and shoulder har- other by at least 10 ° vertically (i.e., nesses provided in the design; and pitch out of parallel) and by at least a 14 CFR Ch. I (1–1–25 Edition) § 27.563 10 ° lateral roll, with the directions op- for ditching must meet the require- tional, to account for possible floor ments of this section and § 27.801(e).

warp.

(a) Forward speed landing conditions.

(c) Compliance with the following The rotorcraft must initially contact must be shown: the most critical wave for reasonably (1) The seating device system must probable water conditions at forward remain intact although it may experi- velocities from zero up to 30 knots in ence separation intended as part of its likely pitch, roll, and yaw attitudes.

design.

The rotorcraft limit vertical descent (2) The attachment between the seat- velocity may not be less than 5 feet per ing device and the airframe structure second relative to the mean water sur- must remain intact, although the face. Rotor lift may be used to act structure may have exceeded its limit through the center of gravity through- load.

out the landing impact. This lift may (3) The ATD’s shoulder harness strap not exceed two-thirds of the design or straps must remain on or in the im- maximum weight. A maximum forward mediate vicinity of the ATD’s shoulder velocity of less than 30 knots may be during the impact.

used in design if it can be dem- (4) The safety belt must remain on onstrated that the forward velocity se- the ATD’s pelvis during the impact.

lected would not be exceeded in a nor- (5) The ATD’s head either does not mal one-engine-out touchdown.

contact any portion of the crew or pas- (b) Auxiliary or emergency float condi- senger compartment, or if contact is tions —(1) Floats fixed or deployed before made, the head impact does not exceed initial water contact. In addition to the a head injury criteria (HIC) of 1,000 as landing loads in paragraph (a) of this determined by this equation.

section, each auxiliary or emergency float, of its support and attaching 2.5 structure in the airframe or fuselage, t ⎡ ⎤ must be designed for the load devel- HIC t t a(t)dt = −

( ) ⎢ ⎥

2 1

∫

t

t t − ( ) oped by a fully immersed float unless it

2 1 ⎢ ⎥ ⎣ ⎦ can be shown that full immersion is Where: a(t) is the resultant acceleration at unlikely. If full immersion is unlikely, the center of gravity of the head form ex- the highest likely float buoyancy load pressed as a multiple of g (the accelera- must be applied. The highest likely tion of gravity) and t ¥ t is the time 2 1 buoyancy load must include consider- duration, in seconds, of major head im- pact, not to exceed 0.05 seconds. ation of a partially immersed float cre- ating restoring moments to com- (6) Loads in individual upper torso pensate the upsetting moments caused harness straps must not exceed 1,750 by side wind, unsymmetrical rotorcraft pounds. If dual straps are used for re- loading, water wave action, rotorcraft taining the upper torso, the total har- inertia, and probable structural dam- ness strap loads must not exceed 2,000 age and leakage considered under pounds.

§ 27.801(d). Maximum roll and pitch an- (7) The maximum compressive load gles determined from compliance with measured between the pelvis and the § 27.801(d) may be used, if significant, to lumbar column of the ATD must not determine the extent of immersion of exceed 1,500 pounds.

each float. If the floats are deployed in (d) An alternate approach that flight, appropriate air loads derived achieves an equivalent or greater level from the flight limitations with the of occupant protection, as required by floats deployed shall be used in sub- this section, must be substantiated on stantiation of the floats and their at- a rational basis.

tachment to the rotorcraft. For this [Amdt. 27–25, 54 FR 47318, Nov. 13, 1989] purpose, the design airspeed for limit load is the float deployed airspeed op- § 27.563 Structural ditching provi- erating limit multiplied by 1.11.

sions.

(2) Floats deployed after initial water If certification with ditching provi- contact. Each float must be designed for sions is requested, structural strength full or partial immersion perscribed in Federal Aviation Administration, DOT § 27.573 paragraph (b)(1) of this section. In addi- tremely remote within a replacement tion, each float must be designed for time furnished under section A27.4 of combined vertical and drag loads using appendix A.

a relative limit speed of 20 knots be- (d) Fail-safe evaluation. The following tween the rotorcraft and the water.

apply to fail-safe evaluation: The vertical load may not be less than (1) It must be shown that all partial the highest likely buoyancy load deter- failures will become readily detectable mined under paragraph (b)(1) of this under inspection procedures furnished section.

under section A27.4 of appendix A.

(2) The interval between the time [Amdt. 27–26, 55 FR 8000, Mar. 6, 1990] when any partial failure becomes read- F ATIGUE E VALUATION ily detectable under paragraph (d)(1) of this section, and the time when any § 27.571 Fatigue evaluation of flight such failure is expected to reduce the structure.

remaining strength of the structure to (a) General. Each portion of the flight limit or maximum attainable loads structure (the flight structure includes (whichever is less), must be deter- rotors, rotor drive systems between the mined.

engines and the rotor hubs, controls, (3) It must be shown that the interval fuselage, landing gear, and their re- determined under paragraph (d)(2) of lated primary attachments), the failure this section is long enough, in relation of which could be catastrophic, must be to the inspection intervals and related identified and must be evaluated under procedures furnished under section paragraph (b), (c), (d), or (e) of this sec- A27.4 of appendix A, to provide a prob- tion. The following apply to each fa- ability of detection great enough to en- tigue evaluation: sure that the probability of cata- (1) The procedure for the evaluation strophic failure is extremely remote.

must be approved.

(e) Combination of replacement time (2) The locations of probable failure and failsafe evaluations. A component must be determined.

may be evaluated under a combination (3) Inflight measurement must be in- of paragraphs (c) and (d) of this sec- cluded in determining the following: tion. For such component it must be (i) Loads or stresses in all critical shown that the probability of cata- conditions throughout the range of strophic failure is extremely remote limitations in § 27.309, except that ma- with an approved combination of re- neuvering load factors need not exceed placement time, inspection intervals, the maximum values expected in oper- and related procedures furnished under ation.

section A27.4 of appendix A.

(ii) The effect of altitude upon these loads or stresses.

(Secs. 313(a), 601, 603, 604, and 605, 72 Stat. 752, (4) The loading spectra must be as se- 775, and 778, (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425; sec. 6(c), 49 U.S.C. 1655(c))) vere as those expected in operation in- cluding, but not limited to, external [Amdt. 27–3, 33 FR 14106, Sept. 18, 1968, as cargo operations, if applicable, and amended by Amdt. 27–12, 42 FR 15044, Mar. 17, ground-air-ground cycles. The loading 1977; Amdt. 27–18, 45 FR 60177, Sept. 11, 1980; Amdt. 27–26, 55 FR 8000, Mar. 6, 1990] spectra must be based on loads or stresses determined under paragraph § 27.573 Damage Tolerance and Fa- (a)(3) of this section.

tigue Evaluation of Composite (b) Fatigue tolerance evaluation. It Rotorcraft Structures.

must be shown that the fatigue toler- ance of the structure ensures that the (a) Each applicant must evaluate the probability of catastrophic fatigue fail- composite rotorcraft structure under ure is extremely remote without estab- the damage tolerance standards of lishing replacement times, inspection paragraph (d) of this section unless the intervals or other procedures under applicant establishes that a damage section A27.4 of appendix A. tolerance evaluation is impractical (c) Replacement time evaluation. it within the limits of geometry, must be shown that the probability of inspectability, and good design prac- catastrophic fatigue failure is ex- tice. If an applicant establishes that it 14 CFR Ch. I (1–1–25 Edition) § 27.573 is impractical within the limits of ge- conditions throughout the range of ometry, inspectability, and good design limits in § 27.309 (including altitude ef- practice, the applicant must do a fa- fects), except that maneuvering load tigue evaluation in accordance with factors need not exceed the maximum paragraph (e) of this section. values expected in service; (b) The methodology used to estab- (iii) The loading spectra as severe as lish compliance with this section must those expected in service based on be submitted to and approved by the loads or stresses determined under Administrator. paragraph (d)(1)(ii) of this section, in- (c) Definitions: cluding external load operations, if ap- (1) Catastrophic failure is an event plicable, and other operations includ- that could prevent continued safe ing high-torque events; flight and landing. (iv) A threat assessment for all PSEs (2) Principal Structural Elements (PSEs) that specifies the locations, types, and are structural elements that con- sizes of damage, considering fatigue, tribute significantly to the carrying of environmental effects, intrinsic and flight or ground loads, the failure of discrete flaws, and impact or other ac- which could result in catastrophic fail- cidental damage (including the discrete ure of the rotorcraft. source of the accidental damage) that (3) Threat Assessment is an assessment may occur during manufacture or oper- that specifies the locations, types, and ation; and sizes of damage, considering fatigue, (v) An assessment of the residual environmental effects, intrinsic and strength and fatigue characteristics of discrete flaws, and impact or other ac- all PSEs that supports the replacement cidental damage (including the discrete times and inspection intervals estab- source of the accidental damage) that lished under paragraph (d)(2) of this may occur during manufacture or oper- section.

ation. (2) Each applicant must establish re- (d) Damage Tolerance Evaluation: placement times, inspections, or other (1) Each applicant must show that procedures for all PSEs to require the catastrophic failure due to static and repair or replacement of damaged parts fatigue loads, considering the intrinsic before a catastrophic failure. These re- or discrete manufacturing defects or placement times, inspections, or other accidental damage, is avoided through- procedures must be included in the Air- out the operational life or prescribed worthiness Limitations Section of the inspection intervals of the rotorcraft Instructions for Continued Airworthi- by performing damage tolerance eval- ness required by § 27.1529.

uations of the strength of composite (i) Replacement times for PSEs must PSEs and other parts, detail design be determined by tests, or by analysis points, and fabrication techniques. supported by tests, and must show that Each applicant must account for the the structure is able to withstand the effects of material and process varia- repeated loads of variable magnitude bility along with environmental condi- expected in-service. In establishing tions in the strength and fatigue eval- these replacement times, the following uations. Each applicant must evaluate items must be considered: parts that include PSEs of the air- (A) Damage identified in the threat frame, main and tail rotor drive sys- assessment required by paragraph tems, main and tail rotor blades and (d)(1)(iv) of this section; hubs, rotor controls, fixed and movable (B) Maximum acceptable manufac- control surfaces, engine and trans- turing defects and in-service damage mission mountings, landing gear, other ( i.e. , those that do not lower the resid- parts, detail design points, and fabrica- ual strength below ultimate design tion techniques deemed critical by the loads and those that can be repaired to FAA. Each damage tolerance evalua- restore ultimate strength); and tion must include: (C) Ultimate load strength capability (i) The identification of all PSEs; after applying repeated loads.

(ii) In-flight and ground measure- (ii) Inspection intervals for PSEs ments for determining the loads or must be established to reveal any dam- stresses for all PSEs for all critical age identified in the threat assessment

Federal Aviation Administration, DOT § 27.603 required by paragraph (d)(1)(iv) of this Subpart D—Design and section that may occur from fatigue or Construction other in-service causes before such damage has grown to the extent that G ENERAL the component cannot sustain the re- quired residual strength capability. In § 27.601 Design.

establishing these inspection intervals, (a) The rotorcraft may have no de- the following items must be consid- sign features or details that experience ered: has shown to be hazardous or unreli- (A) The growth rate, including no- able.

growth, of the damage under the re- (b) The suitability of each question- peated loads expected in-service deter- able design detail and part must be es- mined by tests or analysis supported tablished by tests.

by tests; (B) The required residual strength for § 27.602 Critical parts.

the assumed damage established after (a) Critical part. A critical part is a considering the damage type, inspec- part, the failure of which could have a tion interval, detectability of damage, catastrophic effect upon the rotocraft, and the techniques adopted for damage detection. The minimum required re- and for which critical characteristics sidual strength is limit load; and have been identified which must be controlled to ensure the required level (C) Whether the inspection will de- of integrity.

tect the damage growth before the minimum residual strength is reached (b) If the type design includes critical and restored to ultimate load capa- parts, a critical parts list shall be es- bility, or whether the component will tablished. Procedures shall be estab- require replacement. lished to define the critical design (3) Each applicant must consider the characteristics, identify processes that effects of damage on stiffness, dynamic affect those characteristics, and iden- behavior, loads, and functional per- tify the design change and process formance on all PSEs when substan- change controls necessary for showing tiating the maximum assumed damage compliance with the quality assurance size and inspection interval.

requirements of part 21 of this chapter.

(e) Fatigue Evaluation: If an appli- [Doc. No. 29311, 64 FR 46232, Aug. 24, 1999] cant establishes that the damage toler- ance evaluation described in paragraph § 27.603 Materials.

(d) of this section is impractical within The suitability and durability of ma- the limits of geometry, inspectability, terials used for parts, the failure of or good design practice, the applicant which could adversely affect safety, must do a fatigue evaluation of the must— particular composite rotorcraft struc- ture and: (a) Be established on the basis of ex- perience or tests; (1) Identify all PSEs considered in (b) Meet approved specifications that the fatigue evaluation; (2) Identify the types of damage for ensure their having the strength and all PSEs considered in the fatigue eval- other properties assumed in the design uation; data; and (3) Establish supplemental proce- (c) Take into account the effects of dures to minimize the risk of cata- environmental conditions, such as tem- strophic failure associated with the perature and humidity, expected in damages identified in paragraph (d) of service.

this section; and (Secs. 313(a), 601, 603, 604, Federal Aviation (4) Include these supplemental proce- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424); dures in the Airworthiness Limitations and sec. 6(c) of the Dept. of Transportation section of the Instructions for Contin- Act (49 U.S.C. 1655(c))) ued Airworthiness required by § 27.1529.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as [Doc. No. FAA–2009–0660, Amdt. 27–47, 76 FR amended by Amdt. 27–11, 41 FR 55469, Dec. 20, 74663, Dec. 1, 2011] 1976; Amdt. 27–16, 43 FR 50599, Oct. 30, 1978] 14 CFR Ch. I (1–1–25 Edition) § 27.605 (1) Electrically bonding the compo- § 27.605 Fabrication methods.

nents properly to the airframe; or (a) The methods of fabrication used (2) Designing the components so that must produce consistently sound struc- a strike will not endanger the rotor- tures. If a fabrication process (such as craft.

gluing, spot welding, or heat-treating) (c) For nonmetallic components, requires close control to reach this ob- compliance with paragraph (a) of this jective, the process must be performed section may be shown by— according to an approved process speci- (1) Designing the components to min- fication.

imize the effect of a strike; or (b) Each new aircraft fabrication (2) Incorporating acceptable means of method must be substantiated by a diverting the resulting electrical cur- test program.

rent so as not to endanger the rotor- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- craft.

eral Aviation Act of 1958 (49 U.S.C. 1354(a), (d) The electrical bonding and protec- 1421, 1423, 1424 and 1425); sec. 6(c) of the Dept.

tion against lightning and static elec- of Transportation Act (49 U.S.C. 1655(c))) tricity must— [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (1) Minimize the accumulation of amended by Amdt. 27–16, 43 FR 50599, Oct. 30, electrostatic charge; 1978] (2) Minimize the risk of electric shock to crew, passengers, and service § 27.607 Fasteners.

and maintenance personnel using nor- (a) Each removable bolt, screw, nut, mal precautions; pin, or other fastener whose loss could (3) Provide an electrical return path, jeopardize the safe operation of the under both normal and fault condi- rotorcraft must incorporate two sepa- tions, on rotorcraft having grounded rate locking devices. The fastener and electrical systems; and its locking devices may not be ad- (4) Reduce to an acceptable level the versely affected by the environmental effects of static electricity on the func- conditions associated with the par- tioning of essential electrical and elec- ticular installation.

tronic equipment.

(b) No self-locking nut may be used [Amdt. 27–21, 49 FR 44433, Nov. 6, 1984, as on any bolt subject to rotation in oper- amended by Amdt. 27–37, 64 FR 45094, Aug. 18, ation unless a nonfriction locking de- 1999; Amdt. 27–46, 76 FR 33135, June 8, 2011] vice is used in addition to the self-lock- ing device.

§ 27.611 Inspection provisions.

[Amdt. 27–4, 33 FR 14533, Sept. 27, 1968] There must be means to allow the close examination of each part that re- § 27.609 Protection of structure.

quires— (a) Recurring inspection; Each part of the structure must— (b) Adjustment for proper alignment (a) Be suitably protected against de- and functioning; or terioration or loss of strength in serv- (c) Lubrication.

ice due to any cause, including— (1) Weathering; § 27.613 Material strength properties (2) Corrosion; and and design values.

(3) Abrasion; and (b) Have provisions for ventilation (a) Material strength properties must and drainage where necessary to pre- be based on enough tests of material vent the accumulation of corrosive, meeting specifications to establish de- flammable, or noxious fluids. sign values on a statistical basis.

(b) Design values must be chosen to § 27.610 Lightning and static elec- minimize the probability of structural tricity protection.

failure due to material variability. Ex- (a) The rotorcraft must be protected cept as provided in paragraphs (d) and against catastrophic effects from light- (e) of this section, compliance with ning. this paragraph must be shown by se- (b) For metallic components, compli- lecting design values that assure mate- ance with paragraph (a) of this section rial strength with the following prob- may be shown by— ability— Federal Aviation Administration, DOT § 27.621 (1) Where applied loads are eventu- (i) Uncertainties in manufacturing ally distributed through a single mem- processes; or ber within an assembly, the failure of (ii) Uncertainties in inspection meth- which would result in loss of structural ods.

integrity of the component, 99 percent (b) For each part to which §§ 27.621 probability with 95 percent confidence; through 27.625 apply, the factor of safe- and ty prescribed in § 27.303 must be multi- (2) For redundant structure, those in plied by a special factor equal to— which the failure of individual ele- (1) The applicable special factors pre- ments would result in applied loads scribed in §§ 27.621 through 27.625; or being safely distributed to other load- (2) Any other factor great enough to carrying members, 90 percent prob- ensure that the probability of the part ability with 95 percent confidence.

being understrength because of the un- (c) The strength, detail design, and certainties specified in paragraph (a) of fabrication of the structure must mini- this section is extremely remote.

mize the probability of disastrous fa- tigue failure, particularly at points of § 27.621 Casting factors.

stress concentration.

(d) Design values may be those con- (a) General. The factors, tests, and in- tained in the following publications spections specified in paragraphs (b) (available from the Naval Publications and (c) of this section must be applied and Forms Center, 5801 Tabor Avenue, in addition to those necessary to estab- Philadelphia, Pennsylvania 19120) or lish foundry quality control. The in- other values approved by the Adminis- spections must meet approved speci- trator: fications. Paragraphs (c) and (d) of this (1) MIL-HDBK-5, ‘‘Metallic Materials section apply to structural castings ex- and Elements for Flight Vehicle Struc- cept castings that are pressure tested ture’’.

as parts of hydraulic or other fluid sys- (2) MIL-HDBK-17, ‘‘Plastics for tems and do not support structural Flight Vehicles’’.

loads.

(3) ANC-18, ‘‘Design of Wood Aircraft (b) Bearing stresses and surfaces. The Structures’’.

casting factors specified in paragraphs (4) MIL-HDBK-23, ‘‘Composite Con- (c) and (d) of this section— struction for Flight Vehicles’’.

(1) Need not exceed 1.25 with respect (e) Other design values may be used if to bearing stresses regardless of the a selection of the material is made in method of inspection used; and which a specimen of each individual (2) Need not be used with respect to item is tested before use and it is de- the bearing surfaces of a part whose termined that the actual strength bearing factor is larger than the appli- properties of that particular item will cable casting factor.

equal or exceed those used in design.

(c) Critical castings. For each casting (Secs. 313(a), 601, 603, 604, Federal Aviation whose failure would preclude continued Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), safe flight and landing of the rotorcraft sec. 6(c), Dept. of Transportation Act (49 or result in serious injury to any occu- U.S.C. 1655(c))) pant, the following apply: [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (1) Each critical casting must— amended by Amdt. 27–16, 43 FR 50599, Oct. 30, (i) Have a casting factor of not less 1978; Amdt. 27–26, 55 FR 8000, Mar. 6, 1990] than 1.25; and § 27.619 Special factors. (ii) Receive 100 percent inspection by visual, radiographic, and magnetic par- (a) The special factors prescribed in ticle (for ferromagnetic materials) or §§ 27.621 through 27.625 apply to each penetrant (for nonferromagnetic mate- part of the structure whose strength rials) inspection methods or approved is— equivalent inspection methods.

(1) Uncertain; (2) Likely to deteriorate in service (2) For each critical casting with a before normal replacement; or casting factor less than 1.50, three sam- (3) Subject to appreciable variability ple castings must be static tested and due to— shown to meet— 14 CFR Ch. I (1–1–25 Edition) § 27.623 (i) The strength requirements of for the effects of normal relative mo- § 27.305 at an ultimate load cor- tion.

responding to a casting factor of 1.25; (b) No bearing factor need be used on and a part for which any larger special fac- tor is prescribed.

(ii) The deformation requirements of § 27.305 at a load of 1.15 times the limit § 27.625 Fitting factors.

load.

For each fitting (part or terminal (d) Noncritical castings. For each cast- used to join one structural member to ing other than those specified in para- another) the following apply: graph (c) of this section, the following (a) For each fitting whose strength is apply: not proven by limit and ultimate load (1) Except as provided in paragraphs tests in which actual stress conditions (d)(2) and (3) of this section, the casting are simulated in the fitting and sur- factors and corresponding inspections rounding structures, a fitting factor of must meet the following table: at least 1.15 must be applied to each Casting factor Inspection part of— (1) The fitting; 2.0 or greater ............... 100 percent visual.

(2) The means of attachment; and Less than 2.0, greater 100 percent visual, and magnetic (3) The bearing on the joined mem- than 1.5. particle (ferromagnetic materials), penetrant (nonferromagnetic ma- bers.

terials), or approved equivalent (b) No fitting factor need be used— inspection methods.

(1) For joints made under approved 1.25 through 1.50 ........ 100 percent visual, and magnetic practices and based on comprehensive particle (ferromagnetic materials).

penetrant (nonferromagnetic ma- test data (such as continuous joints in terials), and radiographic or ap- metal plating, welded joints, and scarf proved equivalent inspection joints in wood); and methods.

(2) With respect to any bearing sur- face for which a larger special factor is (2) The percentage of castings in- used.

spected by nonvisual methods may be (c) For each integral fitting, the part reduced below that specified in para- must be treated as a fitting up to the graph (d)(1) of this section when an ap- point at which the section properties proved quality control procedure is es- become typical of the member.

tablished.

(d) Each seat, berth, litter, safety (3) For castings procured to a speci- belt, and harness attachment to the fication that guarantees the mechan- structure must be shown by analysis, ical properties of the material in the tests, or both, to be able to withstand casting and provides for demonstration the inertia forces prescribed in of these properties by test of coupons § 27.561(b)(3) multiplied by a fitting fac- cut from the castings on a sampling tor of 1.33.

basis— [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (i) A casting factor of 1.0 may be amended by Amdt. 27–35, 63 FR 43285, Aug. 12, used; and 1998] (ii) The castings must be inspected as provided in paragraph (d)(1) of this sec- § 27.629 Flutter.

tion for casting factors of ‘‘1.25 through Each aerodynamic surface of the 1.50’’ and tested under paragraph (c)(2) rotorcraft must be free from flutter of this section.

under each appropriate speed and [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as power condition.

amended by Amdt. 27–34, 62 FR 46173, Aug. 29, [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as 1997] amended by Amdt. 27–26, 55 FR 8000, Mar. 6, 1990] § 27.623 Bearing factors.

(a) Except as provided in paragraph R OTORS (b) of this section, each part that has § 27.653 Pressure venting and drain- clearance (free fit), and that is subject age of rotor blades.

to pounding or vibration, must have a bearing factor large enough to provide (a) For each rotor blade— Federal Aviation Administration, DOT § 27.672 (1) There must be means for venting ness, and positiveness appropriate to the internal pressure of the blade; its function.

(2) Drainage holes must be provided (b) Each element of each flight con- for the blade; and trol system must be designed, or dis- (3) The blade must be designed to pre- tinctively and permanently marked, to minimize the probability of any incor- vent water from becoming trapped in rect assembly that could result in the it.

malfunction of the system.

(b) Paragraphs (a)(1) and (2) of this section does not apply to sealed rotor § 27.672 Stability augmentation, auto- blades capable of withstanding the matic, and power-operated systems.

maximum pressure differentials ex- If the functioning of stability aug- pected in service.

mentation or other automatic or [Amdt. 27–2, 33 FR 963, Jan. 26, 1968] power-operated systems is necessary to show compliance with the flight char- § 27.659 Mass balance.

acteristics requirements of this part, (a) The rotors and blades must be such systems must comply with § 27.671 mass balanced as necessary to— of this part and the following: (1) Prevent excessive vibration; and (a) A warning which is clearly distin- (2) Prevent flutter at any speed up to guishable to the pilot under expected the maximum forward speed. flight conditions without requiring the (b) The structural integrity of the pilot’s attention must be provided for mass balance installation must be sub- any failure in the stability augmenta- stantiated. tion system or in any other automatic or power-operated system which could [Amdt. 27–2, 33 FR 963, Jan. 26, 1968] result in an unsafe condition if the pilot is unaware of the failure. Warning § 27.661 Rotor blade clearance.

systems must not activate the control There must be enough clearance be- systems.

tween the rotor blades and other parts (b) The design of the stability aug- of the structure to prevent the blades mentation system or of any other auto- from striking any part of the structure matic or power-operated system must during any operating condition.

allow initial counteraction of failures [Amdt. 27–2, 33 FR 963, Jan. 26, 1968] without requiring exceptional pilot skill or strength by overriding the fail- § 27.663 Ground resonance prevention ure by movement of the flight controls means.

in the normal sense and deactivating (a) The reliability of the means for the failed system.

preventing ground resonance must be (c) It must be shown that after any shown either by analysis and tests, or single failure of the stability aug- reliable service experience, or by show- mentation system or any other auto- ing through analysis or tests that mal- matic or power-operated system— function or failure of a single means (1) The rotorcraft is safely control- will not cause ground resonance. lable when the failure or malfunction occurs at any speed or altitude within (b) The probable range of variations, the approved operating limitations; during service, of the damping action (2) The controllability and maneuver- of the ground resonance prevention ability requirements of this part are means must be established and must be met within a practical operational investigated during the test required flight envelope (for example, speed, al- by § 27.241.

titude, normal acceleration, and rotor- [Amdt. 27–2, 33 FR 963, Jan. 26, 1968, as craft configurations) which is described amended by Amdt. 27–26, 55 FR 8000, Mar. 6, in the Rotorcraft Flight Manual; and 1990] (3) The trim and stability character- istics are not impaired below a level C ONTROL S YSTEMS needed to permit continued safe flight § 27.671 General. and landing.

(a) Each control and control system [Amdt. 27–21, 49 FR 44433, Nov. 6, 1984; 49 FR must operate with the ease, smooth- 47594, Dec. 6, 1984] 14 CFR Ch. I (1–1–25 Edition) § 27.673 § 27.673 Primary flight control. § 27.681 Limit load static tests.

(a) Compliance with the limit load Primary flight controls are those requirements of this part must be used by the pilot for immediate control shown by tests in which— of pitch, roll, yaw, and vertical motion (1) The direction of the test loads of the rotorcraft.

produces the most severe loading in the [Amdt. 27–21, 49 FR 44434, Nov. 6, 1984] control system; and (2) Each fitting, pulley, and bracket § 27.674 Interconnected controls.

used in attaching the system to the Each primary flight control system main structure is included.

(b) Compliance must be shown (by must provide for safe flight and landing analyses or individual load tests) with and operate independently after a mal- the special factor requirements for function, failure, or jam of any auxil- control system joints subject to angu- iary interconnected control.

lar motion.

[Amdt. 27–26, 55 FR 8001, Mar. 6, 1990] § 27.683 Operation tests.

§ 27.675 Stops.

It must be shown by operation tests (a) Each control system must have that, when the controls are operated stops that positively limit the range of from the pilot compartment with the control system loaded to correspond motion of the pilot’s controls.

with loads specified for the system, the (b) Each stop must be located in the system is free from— system so that the range of travel of (a) Jamming; its control is not appreciably affected (b) Excessive friction; and by— (c) Excessive deflection.

(1) Wear; (2) Slackness; or § 27.685 Control system details.

(3) Takeup adjustments.

(a) Each detail of each control sys- (c) Each stop must be able to with- tem must be designed to prevent jam- stand the loads corresponding to the ming, chafing, and interference from design conditions for the system.

cargo, passengers, loose objects or the (d) For each main rotor blade— freezing of moisture.

(1) Stops that are appropriate to the (b) There must be means in the cock- blade design must be provided to limit pit to prevent the entry of foreign ob- travel of the blade about its hinge jects into places where they would jam points; and the system.

(2) There must be means to keep the (c) There must be means to prevent blade from hitting the droop stops dur- the slapping of cables or tubes against ing any operation other than starting other parts.

and stopping the rotor. (d) Cable systems must be designed as follows: (Secs. 313(a), 601, 603, 604, Federal Aviation (1) Cables, cable fittings, turn- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), buckles, splices, and pulleys must be of sec. 6(c), Dept. of Transportation Act (49 an acceptable kind.

U.S.C. 1655(c))) (2) The design of the cable systems [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as must prevent any hazardous change in amended by Amdt. 27–16, 43 FR 50599, Oct. 30, cable tension throughout the range of 1978] travel under any operating conditions and temperature variations.

§ 27.679 Control system locks.

(3) No cable smaller than three thir- If there is a device to lock the con- ty-seconds of an inch diameter may be trol system with the rotorcraft on the used in any primary control system.

ground or water, there must be means (4) Pulley kinds and sizes must cor- to— respond to the cables with which they (a) Give unmistakable warning to the are used. The pulley cable combina- pilot when the lock is engaged; and tions and strength values which must (b) Prevent the lock from engaging in be used are specified in Military Hand- flight. book MIL-HDBK-5C, Vol. 1 & Vol. 2, Federal Aviation Administration, DOT § 27.725 Metallic Materials and Elements for other unsafe characteristics must be Flight Vehicle Structures, (Sept. 15, reliable.

1976, as amended through December 15, (b) Compliance with paragraph (a) of 1978). This incorporation by reference this section must be shown by tests was approved by the Director of the simulating service conditions.

Federal Register in accordance with 5 § 27.691 Autorotation control mecha- U.S.C. section 552(a) and 1 CFR part 51.

nism.

Copies may be obtained from the Naval Publications and Forms Center, 5801 Each main rotor blade pitch control Tabor Avenue, Philadelphia, Pennsyl- mechanism must allow rapid entry into vania, 19120. Copies may be inspected autorotation after power failure.

at the National Archives and Records Administration (NARA). For informa- § 27.695 Power boost and power-oper- ated control system.

tion on the availability of this mate- rial at NARA, call 202–741–6030, or go (a) If a power boost or power-oper- to: http://www.archives.gov/federal-reg- ated control system is used, an alter- ister/cfr/ibr-locations.html nate system must be immediately (5) Pulleys must have close fitting available that allows continued safe guards to prevent the cables from being flight and landing in the event of— displaced or fouled.

(1) Any single failure in the power (6) Pulleys must lie close enough to portion of the system; or the plane passing through the cable to (2) The failure of all engines.

prevent the cable from rubbing against (b) Each alternate system may be a the pulley flange.

duplicate power portion or a manually (7) No fairlead may cause a change in operated mechanical system. The cable direction of more than 3 ° .

power portion includes the power (8) No clevis pin subject to load or source (such as hydraulic pumps), and motion and retained only by cotter such items as valves, lines, and actu- pins may be used in the control sys- ators.

tem.

(c) The failure of mechanical parts (9) Turnbuckles attached to parts (such as piston rods and links), and the having angular motion must be in- jamming of power cylinders, must be stalled to prevent binding throughout considered unless they are extremely the range of travel.

improbable.

(10) There must be means for visual ANDING G EAR inspection at each fairlead, pulley, ter- L minal, and turnbuckle.

§ 27.723 Shock absorption tests.

(e) Control system joints subject to angular motion must incorporate the The landing inertia load factor and following special factors with respect the reserve energy absorption capacity to the ultimate bearing strength of the of the landing gear must be substan- softest material used as a bearing: tiated by the tests prescribed in (1) 3.33 for push-pull systems other §§ 27.725 and 27.727, respectively. These than ball and roller bearing systems. tests must be conducted on the com- plete rotorcraft or on units consisting (2) 2.0 for cable systems.

of wheel, tire, and shock absorber in (f) For control system joints, the their proper relation.

manufacturer’s static, non-Brinell rat- ing of ball and roller bearings must not § 27.725 Limit drop test.

be exceeded.

The limit drop test must be con- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as ducted as follows: amended by Amdt. 27–11, 41 FR 55469, Dec. 20, (a) The drop height must be— 1976; Amdt. 27–26, 55 FR 8001, Mar. 6, 1990; 69 FR 18803, Apr. 9, 2004; Doc. No. FAA–2018– (1) 13 inches from the lowest point of 0119, Amdt. 27–49, 83 FR 9170, Mar. 5, 2018] the landing gear to the ground; or (2) Any lesser height, not less than § 27.687 Spring devices.

eight inches, resulting in a drop con- (a) Each control system spring device tact velocity equal to the greatest whose failure could cause flutter or probable sinking speed likely to occur 14 CFR Ch. I (1–1–25 Edition) § 27.727 the acceleration dv/dt in g ’s recorded in at ground contact in normal power-off the drop test plus 1.0).

landings.

(b) If considered, the rotor lift speci- § 27.727 Reserve energy absorption fied in § 27.473(a) must be introduced drop test.

into the drop test by appropriate en- The reserve energy absorption drop ergy absorbing devices or by the use of test must be conducted as follows: an effective mass.

(a) The drop height must be 1.5 times (c) Each landing gear unit must be that specified in § 27.725(a).

tested in the attitude simulating the (b) Rotor lift, where considered in a landing condition that is most critical manner similar to that prescribed in from the standpoint of the energy to be § 27.725(b), may not exceed 1.5 times the absorbed by it.

(d) When an effective mass is used in lift allowed under that paragraph.

(c) The landing gear must withstand showing compliance with paragraph (b) this test without collapsing. Collapse of this section, the following formula of the landing gear occurs when a may be used instead of more rational member of the nose, tail, or main gear computations: will not support the rotorcraft in the proper attitude or allows the rotorcraft h d + − 1 L

( )

W W = × and ; structure, other than the landing gear e h d + and external accessories, to impact the landing surface.

W e n n L = + j [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as W amended by Amdt. 27–26, 55 FR 8001, Mar. 6, where: 1990] W e = the effective weight to be used in the drop test (lbs.); § 27.729 Retracting mechanism.

W = W for main gear units (lbs.), equal to M For rotorcraft with retractable land- the static reaction on the particular unit ing gear, the following apply: with the rotorcraft in the most critical attitude. A rational method may be used (a) Loads. The landing gear, retract- in computing a main gear static reac- ing mechansim, wheel-well doors, and tion, taking into consideration the mo- supporting structure must be designed ment arm between the main wheel reac- for— tion and the rotorcraft center of gravity.

(1) The loads occurring in any ma- for nose gear units (lbs.), equal to W = W N neuvering condition with the gear re- the vertical component of the static re- tracted; action that would exist at the nose wheel, assuming that the mass of the (2) The combined friction, inertia, rotorcraft acts at the center of gravity and air loads occurring during retrac- and exerts a force of 1.0 g downward and tion and extension at any airspeed up 0.25 g forward.

to the design maximum landing gear W = W T for tailwheel units (lbs.), equal to operating speed; and whichever of the following is critical: (3) The flight loads, including those (1) The static weight on the tailwheel with in yawed flight, occurring with the the rotorcraft resting on all wheels; or gear extended at any airspeed up to the (2) The vertical component of the ground design maximum landing gear extended reaction that would occur at the tailwheel, speed.

assuming that the mass of the rotorcraft (b) Landing gear lock. A positive acts at the center of gravity and exerts a force of l g downward with the rotorcraft in means must be provided to keep the the maximum nose-up attitude considered in gear extended.

the nose-up landing conditions.

(c) Emergency operation. When other than manual power is used to operate h = specified free drop height (inches).

L = ration of assumed rotor lift to the rotor- the gear, emergency means must be craft weight.

provided for extending the gear in the d = deflection under impact of the tire (at event of— the proper inflation pressure) plus the (1) Any reasonably probable failure in vertical component of the axle travels the normal retraction system; or (inches) relative to the drop mass.

(2) The failure of any single source of n = limit inertia load factor.

hydraulic, electric, or equivalent en- n = the load factor developed, during impact, j on the mass used in the drop test (i.e., ergy.

Federal Aviation Administration, DOT § 27.753 (d) Operation tests. The proper func- the tire and any part of the structure tioning of the retracting mechanism or systems.

must be shown by operation tests.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (e) Position indicator. There must be a amended by Amdt. 27–11, 41 FR 55469, Dec. 20, means to indicate to the pilot when the 1976] gear is secured in the extreme posi- tions. § 27.735 Brakes.

(f) Control. The location and oper- For rotorcraft with wheel-type land- ation of the retraction control must ing gear, a braking device must be in- meet the requirements of §§ 27.777 and stalled that is— 27.779.

(a) Controllable by the pilot; (g) Landing gear warning. An aural or (b) Usable during power-off landings; equally effective landing gear warning and device must be provided that functions (c) Adequate to— continuously when the rotorcraft is in (1) Counteract any normal unbal- a normal landing mode and the landing anced torque when starting or stopping gear is not fully extended and locked.

the rotor; and A manual shutoff capability must be (2) Hold the rotorcraft parked on a provided for the warning device and the 10-degree slope on a dry, smooth pave- warning system must automatically ment.

reset when the rotorcraft is no longer [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as in the landing mode.

amended by Amdt. 27–21, 49 FR 44434, Nov. 6, [Amdt. 27–21, 49 FR 44434, Nov. 6, 1984] 1984] § 27.731 Wheels.

§ 27.737 Skis.

(a) Each landing gear wheel must be The maximum limit load rating of approved.

each ski must equal or exceed the max- (b) The maximum static load rating imum limit load determined under the of each wheel may not be less than the applicable ground load requirements of corresponding static ground reaction this part.

with— F LOATS AND H ULLS (1) Maximum weight; and (2) Critical center of gravity.

§ 27.751 Main float buoyancy.

(c) The maximum limit load rating of (a) For main floats, the buoyancy each wheel must equal or exceed the necessary to support the maximum maximum radial limit load determined weight of the rotorcraft in fresh water under the applicable ground load re- must be exceeded by— quirements of this part.

(1) 50 percent, for single floats; and § 27.733 Tires.

(2) 60 percent, for multiple floats.

(b) Each main float must have (a) Each landing gear wheel must enough water-tight compartments so have a tire— that, with any single main float com- (1) That is a proper fit on the rim of partment flooded, the main floats will the wheel; and provide a margin of positive stability (2) Of the proper rating.

great enough to minimize the prob- (b) The maximum static load rating ability of capsizing.

of each tire must equal or exceed the static ground reaction obtained at its [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as wheel, assuming— amended by Amdt. 27–2, 33 FR 963, Jan. 26, 1968] (1) The design maximum weight; and (2) The most unfavorable center of § 27.753 Main float design.

gravity.

(c) Each tire installed on a retract- (a) Bag floats. Each bag float must be able landing gear system must, at the designed to withstand— maximum size of the tire type expected (1) The maximum pressure differen- in service, have a clearance to sur- tial that might be developed at the rounding structure and systems that is maximum altitude for which certifi- adequate to prevent contact between cation with that float is requested; and 14 CFR Ch. I (1–1–25 Edition) § 27.755 (2) The vertical loads prescribed in pilot’s outside field of view, such as a § 27.521(a), distributed along the length head up-display, head mounted display, of the bag over three-quarters of its or other equivalent display, must meet projected area. the following requirements: (1) While the vision system display is (b) Rigid floats. Each rigid float must in operation, it must compensate for be able to withstand the vertical, hori- interference with the pilot’s outside zontal, and side loads prescribed in field of view such that the combination § 27.521. These loads may be distributed of what is visible in the display and along the length of the float.

what remains visible through and § 27.755 Hulls.

around it, allows the pilot compart- ment to satisfy the requirements of For each rotorcraft, with a hull and paragraphs (a)(1) and (b) of this sec- auxiliary floats, that is to be approved tion.

for both taking off from and landing on (2) The pilot’s view of the external water, the hull and auxiliary floats scene may not be distorted by the must have enough watertight compart- transparent display surface or by the ments so that, with any single com- vision system imagery. When the vi- partment flooded, the buoyancy of the sion system displays imagery or any hull and auxiliary floats (and wheel symbology that is referenced to the im- tires if used) provides a margin of posi- agery and outside scene topography, tive stability great enough to minimize including attitude symbology, flight the probability of capsizing.

path vector, and flight path angle ref- erence cue, that imagery and sym- P ERSONNEL AND C ARGO bology must be aligned with, and A CCOMMODATIONS scaled to, the external scene.

§ 27.771 Pilot compartment.

(3) The vision system must provide a means to allow the pilot using the dis- For each pilot compartment— play to immediately deactivate and re- (a) The compartment and its equip- activate the vision system imagery, on ment must allow each pilot to perform demand, without removing the pilot’s his duties without unreasonable con- hands from the primary flight and centration or fatigue; power controls, or their equivalent.

(b) If there is provision for a second (4) When the vision system is not in pilot, the rotorcraft must be control- operation it must permit the pilot lable with equal safety from either compartment to satisfy the require- pilot seat; and ments of paragraphs (a)(1) and (b) of (c) The vibration and noise charac- this section.

teristics of cockpit appurtenances may not interfere with safe operation.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Docket FAA–2013–0485, Amdt. 27– § 27.773 Pilot compartment view.

48, 81 FR 90170, Dec. 13, 2016; Docket FAA– 2016–9275, Amdt. 27–50, 83 FR 9423, Mar. 6, (a) Each pilot compartment must be 2018] free from glare and reflections that could interfere with the pilot’s view, § 27.775 Windshields and windows.

and designed so that— Windshields and windows must be (1) Each pilot’s view is sufficiently made of material that will not break extensive, clear, and undistorted for into dangerous fragments.

safe operation; and (2) Each pilot is protected from the [Amdt. 27–27, 55 FR 38966, Sept. 21, 1990] elements so that moderate rain condi- § 27.777 Cockpit controls.

tions do not unduly impair his view of the flight path in normal flight and Cockpit controls must be— while landing. (a) Located to provide convenient op- (b) If certification for night oper- eration and to prevent confusion and ation is requested, compliance with inadvertent operation; and paragraph (a) of this section must be (b) Located and arranged with re- shown by ground or night flight tests. spect to the pilots’ seats so that there (c) A vision system with a trans- is full and unrestricted movement of parent display surface located in the each control without interference from Federal Aviation Administration, DOT § 27.785 the cockpit structure or the pilot’s § 27.561(b) and dynamic conditions spec- clothing when pilots from 5 ′ 2 ″ to 6 ′ 0 ″ in ified in § 27.562.

height are seated. (b) Each occupant must be protected from serious head injury by a safety § 27.779 Motion and effect of cockpit belt plus a shoulder harness that will controls.

prevent the head from contacting any Cockpit controls must be designed so injurious object except as provided for that they operate in accordance with in § 27.562(c)(5). A shoulder harness the following movements and actu- (upper torso restraint), in combination ation: with the safety belt, constitutes a torso restraint system as described in (a) Flight controls, including the col- TSO-C114.

lective pitch control, must operate (c) Each occupant’s seat must have a with a sense of motion which cor- combined safety belt and shoulder har- responds to the effect on the rotor- ness with a single-point release. Each craft.

pilot’s combined safety belt and shoul- (b) Twist-grip engine power controls der harness must allow each pilot when must be designed so that, for lefthand seated with safety belt and shoulder operation, the motion of the pilot’s harness fastened to perform all func- hand is clockwise to increase power tions necessary for flight operations.

when the hand is viewed from the edge There must be a means to secure belts containing the index finger. Other en- and harnesses, when not in use, to pre- gine power controls, excluding the col- vent interference with the operation of lective control, must operate with a the rotorcraft and with rapid egress in forward motion to increase power.

an emergency.

(c) Normal landing gear controls (d) If seat backs do not have a firm must operate downward to extend the handhold, there must be hand grips or landing gear.

rails along each aisle to enable the oc- [Amdt. 27–21, 49 FR 44434, Nov. 6, 1984] cupants to steady themselves while using the aisle in moderately rough § 27.783 Doors.

air.

(a) Each closed cabin must have at (e) Each projecting object that could least one adequate and easily acces- injure persons seated or moving about sible external door.

in the rotorcraft in normal flight must (b) Each external door must be lo- be padded.

cated where persons using it will not be (f) Each seat and its supporting endangered by the rotors, propellers, structure must be designed for an occu- engine intakes, and exhausts when ap- pant weight of at least 170 pounds con- propriate operating procedures are sidering the maximum load factors, in- used. If opening procedures are re- ertial forces, and reactions between oc- quired, they must be marked inside, on cupant, seat, and safety belt or harness or adjacent to the door opening device.

corresponding with the applicable flight and ground load conditions, in- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as cluding the emergency landing condi- amended by Amdt. 27–26, 55 FR 8001, Mar. 6, 1990] tions of § 27.561(b). In addition— (1) Each pilot seat must be designed § 27.785 Seats, berths, litters, safety for the reactions resulting from the ap- belts, and harnesses.

plication of the pilot forces prescribed (a) Each seat, safety belt, harness, in § 27.397; and and adjacent part of the rotorcraft at (2) The inertial forces prescribed in each station designated for occupancy § 27.561(b) must be multiplied by a fac- during takeoff and landing must be free tor of 1.33 in determining the strength of potentially injurious objects, sharp of the attachment of— edges, protuberances, and hard surfaces (i) Each seat to the structure; and and must be designed so that a person (ii) Each safety belt or harness to the making proper use of these facilities seat or structure.

will not suffer serious injury in an (g) When the safety belt and shoulder emergency landing as a result of the harness are combined, the rated static inertial load factors specified in strength of the safety belt and shoulder 14 CFR Ch. I (1–1–25 Edition) § 27.787 harness may not be less than that cor- pying it during emergency landing con- responding to the inertial forces speci- ditions; and fied in § 27.561(b), considering the occu- (2) The berth or litter attachment pant weight of at least 170 pounds, con- and the occupant restraint system at- sidering the dimensional characteris- tachments to the structure must be de- signed to withstand the critical loads tics of the restraint system installa- resulting from flight and ground load tion, and using a distribution of at conditions and from the conditions pre- least a 60-percent load to the safety scribed in § 27.561(b). The fitting factor belt and at least a 40-percent load to required by § 27.625(d) shall be applied.

the shoulder harness. If the safety belt is capable of being used without the [Amdt. 27–21, 49 FR 44434, Nov. 6, 1984, as shoulder harness, the inertial forces amended by Amdt. 27–25, 54 FR 47319, Nov. 13, specified must be met by the safety 1989; Amdt. 27–35, 63 FR 43285, Aug. 12, 1998] belt alone.

§ 27.787 Cargo and baggage compart- (h) When a headrest is used, the head- ments.

rest and its supporting structure must (a) Each cargo and baggage compart- be designed to resist the inertia forces ment must be designed for its plac- specified in § 27.561, with a 1.33 fitting arded maximum weight of contents and factor and a head weight of at least 13 for the critical load distributions at pounds.

the appropriate maximum load factors (i) Each seating device system in- corresponding to the specified flight cludes the device such as the seat, the and ground load conditions, except the cushions, the occupant restraint sys- emergency landing conditions of tem, and attachment devices.

§ 27.561.

(j) Each seating device system may (b) There must be means to prevent use design features such as crushing or the contents of any compartment from separation of certain parts of the seats becoming a hazard by shifting under to reduce occupant loads for the emer- the loads specified in paragraph (a) of gency landing dynamic conditions of this section.

§ 27.562; otherwise, the system must re- (c) Under the emergency landing con- main intact and must not interfere ditions of § 27.561, cargo and baggage with rapid evacuation of the rotorcraft.

compartments must— (k) For the purposes of this section, a (1) Be positioned so that if the con- litter is defined as a device designed to tents break loose they are unlikely to carry a nonambulatory person, pri- cause injury to the occupants or re- marily in a recumbent position, into strict any of the escape facilities pro- and on the rotorcraft. Each berth or vided for use after an emergency land- litter must be designed to withstand ing; or the load reaction of an occupant (2) Have sufficient strength to with- weight of at least 170 pounds when the stand the conditions specified in § 27.561 occupant is subjected to the forward including the means of restraint, and inertial factors specified in § 27.561(b). their attachments, required by para- A berth or litter installed within 15 ° or graph (b) of this section. Sufficient less of the longitudinal axis of the strength must be provided for the max- rotorcraft must be provided with a pad- imum authorized weight of cargo and ded end-board, cloth diaphram, or baggage at the critical loading dis- equivalent means that can withstand tribution.

(d) If cargo compartment lamps are the forward load reaction. A berth or installed, each lamp must be installed litter oriented greater than 15 ° with so as to prevent contact between lamp the longitudinal axis of the rotorcraft bulb and cargo.

must be equipped with appropriate re- straints, such as straps or safety belts, [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as to withstand the forward load reaction.

amended by Amdt. 27–11, 41 FR 55469, Dec. 20, In addition— 1976; Amdt. 27–27, 55 FR 38966, Sept. 21, 1990] (1) The berth or litter must have a re- § 27.801 Ditching.

straint system and must not have cor- ners or other protuberances likely to (a) If certification with ditching pro- cause serious injury to a person occu- visions is requested, the rotorcraft

Federal Aviation Administration, DOT § 27.807 must meet the requirements of this of the flight crew. This must be shown section and §§ 27.807(d), 27.1411 and by test.

27.1415.

(c) Each flight crew emergency exit (b) Each practicable design measure, must not be obstructed by water or flo- compatible with the general character- tation devices after an emergency istics of the rotorcraft, must be taken landing on water. This must be shown to minimize the probability that in an by test, demonstration, or analysis.

emergency landing on water, the be- [Doc. No. 29247, 64 FR 45094, Aug. 18, 1999] havior of the rotorcraft would cause immediate injury to the occupants or § 27.807 Emergency exits.

would make it impossible for them to escape.

(a) Number and location. (1) There (c) The probable behavior of the must be at least one emergency exit on rotorcraft in a water landing must be each side of the cabin readily acces- investigated by model tests or by com- sible to each passenger. One of these parison with rotorcraft of similar con- exits must be usable in any probable figuration for which the ditching char- attitude that may result from a crash; acteristics are known. Scoops, flaps, (2) Doors intended for normal use projections, and any other factor likely may also serve as emergency exits, pro- to affect the hydrodynamic character- vided that they meet the requirements istics of the rotorcraft must be consid- of this section; and ered.

(3) If emergency flotation devices are (d) It must be shown that, under rea- installed, there must be an emergency sonably probable water conditions, the exit accessible to each passenger on flotation time and trim of the rotor- each side of the cabin that is shown by craft will allow the occupants to leave test, demonstration, or analysis to; the rotorcraft and enter the life rafts (i) Be above the waterline; and required by § 27.1415. If compliance with (ii) Open without interference from this provision is shown by buoyancy and trim computations, appropriate al- flotation devices, whether stowed or lowances must be made for probable deployed.

structural damage and leakage. If the (b) Type and operation. Each emer- rotorcraft has fuel tanks (with fuel jet- gency exit prescribed by paragraph (a) tisoning provisions) that can reason- of this section must— ably be expected to withstand a ditch- (1) Consist of a movable window or ing without leakage, the jettisonable panel, or additional external door, pro- volume of fuel may be considered as viding an unobstructed opening that buoyancy volume.

will admit a 19-by 26-inch ellipse; (e) Unless the effects of the collapse (2) Have simple and obvious methods of external doors and windows are ac- of opening, from the inside and from counted for in the investigation of the the outside, which do not require ex- probable behavior of the rotorcraft in a ceptional effort; water landing (as prescribed in para- (3) Be arranged and marked so as to graphs (c) and (d) of this section), the be readily located and opened even in external doors and windows must be darkness; and designed to withstand the probable (4) Be reasonably protected from maximum local pressures.

jamming by fuselage deformation.

[Amdt. 27–11, 41 FR 55469, Dec. 20, 1976] (c) Tests. The proper functioning of each emergency exit must be shown by § 27.805 Flight crew emergency exits.

test.

(a) For rotorcraft with passenger (d) Ditching emergency exits for pas- emergency exits that are not conven- sengers. If certification with ditching ient to the flight crew, there must be provisions is requested, the markings flight crew emergency exits, on both required by paragraph (b)(3) of this sec- sides of the rotorcraft or as a top hatch tion must be designed to remain visible in the flight crew area.

if the rotorcraft is capsized and the (b) Each flight crew emergency exit cabin is submerged.

must be of sufficient size and must be located so as to allow rapid evacuation [Doc. No. 29247, 64 FR 45094, Aug. 18, 1999] 14 CFR Ch. I (1–1–25 Edition) § 27.831 (2) Fire resistant, in the case of other § 27.831 Ventilation.

compartments.

(a) The ventilating system for the (b) No compartment may contain any pilot and passenger compartments controls, wiring, lines, equipment, or must be designed to prevent the pres- accessories whose damage or failure ence of excessive quantities of fuel would affect safe operation, unless fumes and carbon monoxide.

those items are protected so that— (b) The concentration of carbon mon- (1) They cannot be damaged by the oxide may not exceed one part in 20,000 movement of cargo in the compart- parts of air during forward flight or ment; and hovering in still air. If the concentra- (2) Their breakage or failure will not tion exceeds this value under other create a fire hazard.

conditions, there must be suitable op- erating restrictions.

§ 27.859 Heating systems.

§ 27.833 Heaters.

(a) General. For each heating system that involves the passage of cabin air Each combustion heater must be ap- over, or close to, the exhaust manifold, proved.

there must be means to prevent carbon [Amdt. 27–23, 53 FR 34210, Sept. 2, 1988] monoxide from entering any cabin or pilot compartment.

F IRE P ROTECTION (b) Heat exchangers. Each heat ex- § 27.853 Compartment interiors. changer must be— (1) Of suitable materials; For each compartment to be used by (2) Adequately cooled under all con- the crew or passengers— ditions; and (a) The materials must be at least (3) Easily disassembled for inspec- flame-resistant; tion.

(b) [Reserved] (c) Combustion heater fire protection.

(c) If smoking is to be prohibited, Except for heaters which incorporate there must be a placard so stating, and designs to prevent hazards in the event if smoking is to be allowed— of fuel leakage in the heater fuel sys- (1) There must be an adequate num- tem, fire within the ventilating air pas- ber of self-contained, removable ash- sage, or any other heater malfunction, trays; and each heater zone must incorporate the (2) Where the crew compartment is fire protection features of the applica- separated from the passenger compart- ble requirements of §§ 27.1183, 27.1185, ment, there must be at least one illu- 27.1189, 27.1191, and be provided with— minated sign (using either letters or (1) Approved, quick-acting fire detec- symbols) notifying all passengers when tors in numbers and locations ensuring smoking is prohibited. Signs which no- prompt detection of fire in the heater tify when smoking is prohibited must— region.

(i) When illuminated, be legible to each passenger seated in the passenger (2) Fire extinguisher systems that cabin under all probable lighting condi- provide at least one adequate discharge tions; and to all areas of the heater region.

(ii) Be so constructed that the crew (3) Complete drainage of each part of can turn the illumination on and off. each zone to minimize the hazards re- sulting from failure or malfunction of [Amdt. 27–17, 45 FR 7755, Feb. 4, 1980, as any component containing flammable amended by Amdt. 27–37, 64 FR 45095, Aug. 18, fluids. The drainage means must be— 1999] (i) Effective under conditions ex- § 27.855 Cargo and baggage compart- pected to prevail when drainage is ments.

needed; and (ii) Arranged so that no discharged (a) Each cargo and baggage compart- fluid will cause an additional fire haz- ment must be constructed of, or lined ard.

with, materials that are at least— (1) Flame resistant, in the case of (4) Ventilation, arranged so that no compartments that are readily acces- discharged vapors will cause an addi- sible to a crewmember in flight; and tional fire hazard.

Federal Aviation Administration, DOT § 27.859 (d) Ventilating air ducts. Each ven- (iv) The ventilating airflow becomes tilating air duct passing through any inadequate for safe operation.

heater region must be fireproof. (2) The means of complying with paragraph (g)(1) of this section for any (1) Unless isolation is provided by individual heater must— fireproof valves or by equally effective (i) Be independent of components means, the ventilating air duct down- serving any other heater, the heat out- stream of each heater must be fireproof put of which is essential for safe oper- for a distance great enough to ensure ation; and that any fire originating in the heater (ii) Keep the heater off until re- can be contained in the duct.

started by the crew.

(2) Each part of any ventilating duct (3) There must be means to warn the passing through any region having a crew when any heater, the heat output flammable fluid system must be so of which is essential for safe operation, constructed or isolated from that sys- has been shut off by the automatic tem that the malfunctioning of any means prescribed in paragraph (g)(1) of component of that system cannot in- this section.

troduce flammable fluids or vapors (h) Air intakes. Each combustion and into the ventilating airstream.

ventilating air intake must be located (e) Combustion air ducts. Each com- so that no flammable fluids or vapors bustion air duct must be fireproof for a can enter the heater system— distance great enough to prevent dam- (1) During normal operation; or age from backfiring or reverse flame (2) As a result of the malfunction of propagation.

any other component.

(1) No combustion air duct may con- (i) Heater exhaust. Each heater ex- nect with the ventilating airstream un- haust system must meet the require- less flames from backfires or reverse ments of §§ 27.1121 and 27.1123.

burning cannot enter the ventilating (1) Each exhaust shroud must be airstream under any operating condi- sealed so that no flammable fluids or tion, including reverse flow or mal- hazardous quantities of vapors can function of the heater or its associated reach the exhaust system through components.

joints.

(2) No combustion air duct may re- (2) No exhaust system may restrict strict the prompt relief of any backfire the prompt relief of any backfire that, that, if so restricted, could cause heat- if so restricted, could cause heater fail- er failure.

ure.

(f) Heater control: General. There must (j) Heater fuel systems. Each heater be means to prevent the hazardous ac- fuel system must meet the powerplant cumulation of water or ice on or in any fuel system requirements affecting safe heater control component, control sys- heater operation. Each heater fuel sys- tem tubing, or safety control.

tem component in the ventilating air- (g) Heater safety controls. For each stream must be protected by shrouds combustion heater, safety control so that no leakage from those compo- means must be provided as follows: nents can enter the ventilating air- (1) Means independent of the compo- stream.

nents provided for the normal contin- (k) Drains. There must be means for uous control of air temperature, air- safe drainage of any fuel that might ac- flow, and fuel flow must be provided for cumulate in the combustion chamber each heater to automatically shut off or the heat exchanger.

the ignition and fuel supply of that (1) Each part of any drain that oper- heater at a point remote from that ates at high temperatures must be pro- heater when any of the following oc- tected in the same manner as heater curs: exhausts.

(i) The heat exchanger temperature (2) Each drain must be protected exceeds safe limits. against hazardous ice accumulation under any operating condition.

(ii) The ventilating air temperature exceeds safe limits.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (iii) The combustion airflow becomes amended by Amdt. 27–23, 53 FR 34211, Sept. 2, inadequate for safe operation. 1988] 14 CFR Ch. I (1–1–25 Edition) § 27.861 E XTERNAL L OADS § 27.861 Fire protection of structure, controls, and other parts.

§ 27.865 External loads.

Each part of the structure, controls, (a) It must be shown by analysis, rotor mechanism, and other parts es- test, or both, that the rotorcraft exter- sential to a controlled landing that nal load attaching means for rotor- would be affected by powerplant fires craft-load combinations to be used for must be fireproof or protected so they nonhuman external cargo applications can perform their essential functions can withstand a limit static load equal for at least 5 minutes under any fore- to 2.5, or some lower load factor ap- seeable powerplant fire conditions.

proved under §§ 27.337 through 27.341, [Amdt. 27–26, 55 FR 8001, Mar. 6, 1990] multiplied by the maximum external load for which authorization is re- § 27.863 Flammable fluid fire protec- quested. It must be shown by analysis, tion.

test, or both that the rotorcraft exter- (a) In each area where flammable nal load attaching means and cor- fluids or vapors might escape by leak- responding personnel carrying device age of a fluid system, there must be system for rotorcraft-load combina- means to minimize the probability of tions to be used for human external ignition of the fluids and vapors, and cargo applications can withstand a the resultant hazards if ignition does limit static load equal to 3.5 or some occur.

lower load factor, not less than 2.5, ap- (b) Compliance with paragraph (a) of proved under §§ 27.337 through 27.341, this section must be shown by analysis multiplied by the maximum external or tests, and the following factors must load for which authorization is re- be considered: quested. The load for any rotorcraft- (1) Possible sources and paths of fluid load combination class, for any exter- leakage, and means of detecting leak- nal cargo type, must be applied in the age.

vertical direction. For jettisonable ex- (2) Flammability characteristics of ternal loads of any applicable external fluids, including effects of any combus- cargo type, the load must also be ap- tible or absorbing materials.

plied in any direction making the max- (3) Possible ignition sources, includ- imum angle with the vertical that can ing electrical faults, overheating of be achieved in service but not less than equipment, and malfunctioning of pro- 30 ° . However, the 30 ° angle may be re- tective devices.

duced to a lesser angle if— (4) Means available for controlling or (1) An operating limitation is estab- extinguishing a fire, such as stopping lished limiting external load oper- flow of fluids, shutting down equip- ations to such angles for which compli- ment, fireproof containment, or use of ance with this paragraph has been extinguishing agents.

shown; or (5) Ability of rotorcraft components (2) It is shown that the lesser angle that are critical to safety of flight to can not be exceeded in service.

withstand fire and heat.

(b) The external load attaching (c) If action by the flight crew is re- means, for jettisonable rotorcraft-load quired to prevent or counteract a fluid combinations, must include a quick-re- fire (e.g. equipment shutdown or actu- lease system to enable the pilot to re- ation of a fire extinguisher) quick act- lease the external load quickly during ing means must be provided to alert flight. The quick-release system must the crew.

consist of a primary quick release sub- (d) Each area where flammable fluids system and a backup quick release sub- or vapors might escape by leakage of a system that are isolated from one an- fluid system must be identified and de- other. The quick-release system, and fined.

the means by which it is controlled, (Secs. 313(a), 601, 603, 604, Federal Aviation must comply with the following: Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), (1) A control for the primary quick sec. 6(c), Dept. of Transportation Act (49 release subsystem must be installed ei- U.S.C. 1655(c))) ther on one of the pilot’s primary con- [Amdt. 27–16, 43 FR 50599, Oct. 30, 1978] trols or in an equivalently accessible Federal Aviation Administration, DOT § 27.873 location and must be designed and lo- device system, the ingress and egress cated so that it may be operated by ei- instructions; ther the pilot or a crewmember with- (4) Have equipment to allow direct out hazardously limiting the ability to intercommunication among required control the rotorcraft during an emer- crewmembers and external occupants; gency situation.

and (2) A control for the backup quick re- (5) Have the appropriate limitations lease subsystem, readily accessible to and procedures incorporated in the either the pilot or another crew- flight manual for conducting human member, must be provided.

external cargo operations.

(3) Both the primary and backup (d) The critically configured jettison- quick release subsystems must— able external loads must be shown by a (i) Be reliable, durable, and function combination of analysis, ground tests, properly with all external loads up to and flight tests to be both transport- and including the maximum external able and releasable throughout the ap- limit load for which authorization is proved operational envelope without requested.

hazard to the rotorcraft during normal (ii) Be protected against electro- flight conditions. In addition, these ex- magnetic interference (EMI) from ex- ternal loads must be shown to be re- ternal and internal sources and against leasable without hazard to the rotor- lightning to prevent inadvertent load craft during emergency flight condi- release.

tions.

(A) The minimum level of protection (e) A placard or marking must be in- required for jettisonable rotorcraft- stalled next to the external-load at- load combinations used for nonhuman taching means clearly stating any external cargo is a radio frequency operational limitations and the max- field strength of 20 volts per meter.

imum authorized external load as dem- (B) The minimum level of protection onstrated under § 27.25 and this section.

required for jettisonable rotorcraft- load combinations used for human ex- (f) The fatigue evaluation of § 27.571 ternal cargo is a radio frequency field of this part does not apply to rotor- strength of 200 volts per meter.

craft-load combinations to be used for (iii) Be protected against any failure nonhuman external cargo except for that could be induced by a failure mode the failure of critical structural ele- of any other electrical or mechanical ments that would result in a hazard to rotorcraft system.

the rotorcraft. For rotorcraft-load (c) For rotorcraft-load combinations combinations to be used for human ex- to be used for human external cargo ternal cargo, the fatigue evaluation of applications, the rotorcraft must— § 27.571 of this part applies to the entire (1) For jettisonable external loads, quick release and personnel carrying have a quick-release system that meets device structural systems and their at- the requirements of paragraph (b) of tachments.

this section and that— [Amdt. 27–11, 41 FR 55469, Dec. 20, 1976, as (i) Provides a dual actuation device amended by Amdt. 27–26, 55 FR 8001, Mar. 6, for the primary quick release sub- 1990; Amdt. 27–36, 64 FR 43019, Aug. 6, 1999] system, and (ii) Provides a separate dual actu- M ISCELLANEOUS ation device for the backup quick re- lease subsystem; § 27.871 Leveling marks.

(2) Have a reliable, approved per- There must be reference marks for sonnel carrying device system that has leveling the rotorcraft on the ground.

the structural capability and personnel safety features essential for external § 27.873 Ballast provisions.

occupant safety; Ballast provisions must be designed (3) Have placards and markings at all and constructed to prevent inadvertent appropriate locations that clearly state the essential system operating instruc- shifting of ballast in flight.

tions and, for the personnel carrying

14 CFR Ch. I (1–1–25 Edition) § 27.901 § 27.903 Engines.

Subpart E—Powerplant (a) Engine type certification. Each en- G ENERAL gine must have an approved type cer- tificate. Reciprocating engines for use § 27.901 Installation.

in helicopters must be qualified in ac- (a) For the purpose of this part, the cordance with § 33.49(d) of this chapter powerplant installation includes each or be otherwise approved for the in- part of the rotorcraft (other than the tended usage.

main and auxiliary rotor structures) (b) Engine or drive system cooling fan that— blade protection. (1) If an engine or rotor (1) Is necessary for propulsion; drive system cooling fan is installed, (2) Affects the control of the major there must be a means to protect the propulsive units; or rotorcraft and allow a safe landing if a (3) Affects the safety of the major fan blade fails. This must be shown by propulsive units between normal in- showing that— spections or overhauls.

(i) The fan blades are contained in (b) For each powerplant installa- case of failure; tion— (ii) Each fan is located so that a fail- (1) Each component of the installa- ure will not jeopardize safety; or tion must be constructed, arranged, (iii) Each fan blade can withstand an and installed to ensure its continued ultimate load of 1.5 times the cen- safe operation between normal inspec- trifugal force resulting from operation tions or overhauls for the range of tem- limited by the following: perature and altitude for which ap- (A) For fans driven directly by the proval is requested; engine— (2) Accessibility must be provided to ( 1 ) The terminal engine r.p.m. under allow any inspection and maintenance uncontrolled conditions; or necessary for continued airworthiness; ( 2 ) An overspeed limiting device.

(3) Electrical interconnections must (B) For fans driven by the rotor drive be provided to prevent differences of system, the maximum rotor drive sys- potential between major components of tem rotational speed to be expected in the installation and the rest of the service, including transients.

rotorcraft; (2) Unless a fatigue evaluation under (4) Axial and radial expansion of tur- § 27.571 is conducted, it must be shown bine engines may not affect the safety that cooling fan blades are not oper- of the installation; and ating at resonant conditions within the (5) Design precautions must be taken operating limits of the rotorcraft.

to minimize the possibility of incorrect (c) Turbine engine installation. For assembly of components and equipment turbine engine installations, the pow- essential to safe operation of the rotor- erplant systems associated with engine craft, except where operation with the control devices, systems, and instru- incorrect assembly can be shown to be mentation must be designed to give extremely improbable.

reasonable assurance that those engine (c) The installation must comply operating limitations that adversely with— affect turbine rotor structural integ- (1) The installation instructions pro- rity will not be exceeded in service.

vided under § 33.5 of this chapter; and (d) Restart capability. (1) A means to (2) The applicable provisions of this restart any engine in flight must be subpart.

provided.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 (2) Except for the in-flight shutdown U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 of all engines, engine restart capability U.S.C. 1655(c)) must be demonstrated throughout a [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as flight envelope for the rotorcraft.

amended by Amdt. 27–2, 33 FR 963, Jan. 26, 1968; Amdt. 27–12, 42 FR 15044, Mar. 17, 1977; Amdt. 27–23, 53 FR 34211, Sept. 2, 1988] Federal Aviation Administration, DOT § 27.923 (3) Following the in-flight shutdown pendently of the engine, any limita- of all engines, in-flight engine restart tions on the use of that means must be capability must be provided. specified, and the control for that means must be guarded to prevent in- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as advertent operation.

amended by Amdt. 27–11, 41 FR 55469, Dec. 20, 1976; Amdt. 27–23, 53 FR 34211, Sept. 2, 1988; § 27.923 Rotor drive system and con- Amdt. 27–44, 73 FR 11000, Feb. 29, 2008; Amdt.

trol mechanism tests.

27–51, 88 FR 8737, Feb. 10, 2023] (a) Each part tested as prescribed in § 27.907 Engine vibration.

this section must be in a serviceable (a) Each engine must be installed to condition at the end of the tests. No in- prevent the harmful vibration of any tervening disassembly which might af- part of the engine or rotorcraft. fect test results may be conducted.

(b) The addition of the rotor and the (b) Each rotor drive system and con- rotor drive system to the engine may trol mechanism must be tested for not not subject the principal rotating parts less than 100 hours. The test must be of the engine to excessive vibration conducted on the rotorcraft, and the stresses. This must be shown by a vi- torque must be absorbed by the rotors bration investigation. to be installed, except that other (c) No part of the rotor drive system ground or flight test facilities with may be subjected to excessive vibra- other appropriate methods of torque tion stresses. absorption may be used if the condi- tions of support and vibration closely R OTOR D RIVE S YSTEM simulate the conditions that would exist during a test on the rotorcraft.

§ 27.917 Design.

(c) A 60-hour part of the test pre- (a) Each rotor drive system must in- scribed in paragraph (b) of this section corporate a unit for each engine to must be run at not less than maximum automatically disengage that engine continuous torque and the maximum from the main and auxiliary rotors if speed for use with maximum contin- that engine fails.

uous torque. In this test, the main (b) Each rotor drive system must be rotor controls must be set in the posi- arranged so that each rotor necessary tion that will give maximum longitu- for control in autorotation will con- dinal cyclic pitch change to simulate tinue to be driven by the main rotors forward flight. The auxiliary rotor con- after disengagement of the engine from trols must be in the position for nor- the main and auxiliary rotors.

mal operation under the conditions of (c) If a torque limiting device is used the test.

in the rotor drive system, it must be (d) A 30-hour or, for rotorcraft for located so as to allow continued con- which the use of either 30-minute OEI trol of the rotorcraft when the device power or continuous OEI power is re- is operating.

quested, a 25-hour part of the test pre- (d) The rotor drive system includes scribed in paragraph (b) of this section any part necessary to transmit power must be run at not less than 75 percent from the engines to the rotor hubs.

of maximum continuous torque and the This includes gear boxes, shafting, uni- minimum speed for use with 75 percent versal joints, couplings, rotor brake as- of maximum continuous torque. The semblies, clutches, supporting bearings main and auxiliary rotor controls must for shafting, any attendant accessory be in the position for normal operation pads or drives, and any cooling fans under the conditions of the test.

that are a part of, attached to, or (e) A 10-hour part of the test pre- mounted on the rotor drive system.

scribed in paragraph (b) of this section must be run at not less than takeoff [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–11, 41 FR 55469, Dec. 20, torque and the maximum speed for use 1976] with takeoff torque. The main and aux- iliary rotor controls must be in the § 27.921 Rotor brake.

normal position for vertical ascent.

If there is a means to control the ro- (1) For multiengine rotorcraft for tation of the rotor drive system inde- which the use of 2 ⁄2 minute OEI power 14 CFR Ch. I (1–1–25 Edition) § 27.923 is requested, 12 runs during the 10-hour flight. The part of the test prescribed test must be conducted as follows: in paragraph (e) of this section must be (i) Each run must consist of at least conducted in intervals of not less than one period of 2 ⁄ 2 minutes with takeoff five minutes.

torque and the maximum speed for use (g) At intervals of not more than five with takeoff torque on all engines.

hours during the tests prescribed in (ii) Each run must consist of at least paragraphs (c), (d), and (e) of this sec- one period for each engine in sequence, tion, the engine must be stopped rap- during which that engine simulates a idly enough to allow the engine and power failure and the remaining en- rotor drive to be automatically dis- gines are run at 2 ⁄2 minute OEI torque engaged from the rotors.

and the maximum speed for use with (h) Under the operating conditions 1 1 2 ⁄2 minute OEI torque for 2 ⁄2 minutes.

specified in paragraph (c) of this sec- (2) For multiengine turbine-powered tion, 500 complete cycles of lateral con- rotorcraft for which the use of 30-sec- trol, 500 complete cycles of longitu- ond and 2-minute OEI power is re- dinal control of the main rotors, and quested, 10 runs must be conducted as 500 complete cycles of control of each follows: auxiliary rotor must be accomplished.

(i) Immediately following a takeoff A ‘‘complete cycle’’ involves movement run of at least 5 minutes, each power of the controls from the neutral posi- source must simulate a failure, in turn, and apply the maximum torque and the tion, through both extreme positions, maximum speed for use with 30-second and back to the neutral position, ex- OEI power to the remaining affected cept that control movements need not drive system power inputs for not less produce loads or flapping motions ex- than 30 seconds, followed by applica- ceeding the maximum loads or motions tion of the maximum torque and the encountered in flight. The cycling may maximum speed for use with 2-minute be accomplished during the testing pre- OEI power for not less than 2 minutes.

scribed in paragraph (c) of this section.

At least one run sequence must be con- (i) At least 200 start-up clutch en- ducted from a simulated ‘‘flight idle’’ gagements must be accomplished— condition. When conducted on a bench (1) So that the shaft on the driven test, the test sequence must be con- side of the clutch is accelerated; and ducted following stabilization at take- (2) Using a speed and method selected off power.

by the applicant.

(ii) For the purpose of this para- (j) For multiengine rotorcraft for graph, an affected power input includes which the use of 30-minute OEI power all parts of the rotor drive system is requested, five runs must be made at which can be adversely affected by the 30-minute OEI torque and the max- application of higher or asymmetric torque and speed prescribed by the imum speed for use with 30-minute OEI test. torque, in which each engine, in se- (iii) This test may be conducted on a quence, is made inoperative and the re- representative bench test facility when maining engine(s) is run for a 30- engine limitations either preclude re- minute period.

peated use of this power or would re- (k) For multiengine rotorcraft for sult in premature engine removal dur- which the use of continuous OEI power ing the test. The loads, the vibration is requested, five runs must be made at frequency, and the methods of applica- continuous OEI torque and the max- tion to the affected rotor drive system imum speed for use with continuous components must be representative of OEI torque, in which each engine, in rotorcraft conditions. Test components sequence, is made inoperative and the must be those used to show compliance with the remainder of this section.

(f) The parts of the test prescribed in paragraphs (c) and (d) of this section must be conducted in intervals of not less than 30 minutes and may be ac- complished either on the ground or in Federal Aviation Administration, DOT § 27.939 remaining engine(s) is run for a 1-hour would exist during a test on the rotor- period. craft.

(c) It must be shown by tests that the (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 rotor drive system is capable of oper- U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 ating under autorotative conditions for U.S.C. 1655(c)) 15 minutes after the loss of pressure in [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as the rotor drive primary oil system.

amended by Amdt. 27–2, 33 FR 963, Jan. 26, 1968; Amdt. 27–12, 42 FR 15044, Mar. 17, 1977; (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 Amdt. 27–23, 53 FR 34212, Sept. 2, 1988; Amdt. U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 27–29, 59 FR 47767, Sept. 16, 1994] U.S.C. 1655(c)) [Amdt. 27–2, 33 FR 963, Jan. 26, 1968, as § 27.927 Additional tests.

amended by Amdt. 27–12, 42 FR 15045, Mar. 17, (a) Any additional dynamic, endur- 1977; Amdt. 27–23, 53 FR 34212, Sept. 2, 1988] ance, and operational tests, and vibra- § 27.931 Shafting critical speed.

tory investigations necessary to deter- mine that the rotor drive mechanism is (a) The critical speeds of any shafting safe, must be performed.

must be determined by demonstration (b) If turbine engine torque output to except that analytical methods may be the transmission can exceed the high- used if reliable methods of analysis are est engine or transmission torque rat- available for the particular design.

ing limit, and that output is not di- (b) If any critical speed lies within, rectly controlled by the pilot under or close to, the operating ranges for normal operating conditions (such as idling, power on, and autorotative con- where the primary engine power con- ditions, the stresses occurring at that trol is accomplished through the flight speed must be within safe limits. This control), the following test must be must be shown by tests.

made: (c) If analytical methods are used and (1) Under conditions associated with show that no critical speed lies within all engines operating, make 200 appli- the permissible operating ranges, the margins between the calculated crit- cations, for 10 seconds each, or torque ical speeds and the limits of the allow- that is at least equal to the lesser of— able operating ranges must be adequate (i) The maximum torque used in to allow for possible variations be- meeting § 27.923 plus 10 percent; or (ii) The maximum attainable torque tween the computed and actual values.

output of the engines, assuming that § 27.935 Shafting joints.

torque limiting devices, if any, func- tion properly.

Each universal joint, slip joint, and (2) For multiengine rotorcraft under other shafting joints whose lubrication conditions associated with each engine, is necessary for operation must have in turn, becoming inoperative, apply to provision for lubrication.

the remaining transmission torque in- § 27.939 Turbine engine operating puts the maximum torque attainable characteristics.

under probable operating conditions, assuming that torque limiting devices, (a) Turbine engine operating charac- if any, function properly. Each trans- teristics must be investigated in flight mission input must be tested at this to determine that no adverse charac- maximum torque for at least 15 min- teristics (such as stall, surge, or flame- utes. out) are present, to a hazardous degree, (3) The tests prescribed in this para- during normal and emergency oper- graph must be conducted on the rotor- ation within the range of operating craft at the maximum rotational speed limitations of the rotorcraft and of the intended for the power condition of the engine.

test and the torque must be absorbed (b) The turbine engine air inlet sys- by the rotors to be installed, except tem may not, as a result of airflow dis- that other ground or flight test facili- tortion during normal operation, cause ties with other appropriate methods of vibration harmful to the engine.

torque absorption may be used if the (c) For governor-controlled engines, conditions of support and vibration it must be shown that there exists no closely simulate the conditions that hazardous torsional instability of the 14 CFR Ch. I (1–1–25 Edition) § 27.951 drive system associated with critical (3) The tank must be filled with combinations of power, rotational water to 80 percent of the normal, full speed, and control displacement. capacity.

(4) The tank must be enclosed in a [Amdt. 27–1, 32 FR 6914, May 5, 1967, as surrounding structure representative amended by Amdt. 27–11, 41 FR 55469, Dec. 20, of the installation unless it can be es- 1976] tablished that the surrounding struc- F UEL S YSTEM ture is free of projections or other de- sign features likely to contribute to § 27.951 General.

rupture of the tank.

(a) Each fuel system must be con- (5) The tank must drop freely and im- structed and arranged to ensure a flow pact in a horizontal position ± 10 ° .

of fuel at a rate and pressure estab- (6) After the drop test, there must be lished for proper engine functioning no leakage.

under any likely operating condition, (b) Fuel tank load factors. Except for including the maneuvers for which cer- fuel tanks located so that tank rupture tification is requested.

with fuel release to either significant (b) Each fuel system must be ar- ignition sources, such as engines, heat- ranged so that— ers, and auxiliary power units, or occu- (1) No fuel pump can draw fuel from pants is extremely remote, each fuel more than one tank at a time; or tank must be designed and installed to (2) There are means to prevent intro- retain its contents under the following ducing air into the system.

ultimate inertial load factors, acting (c) Each fuel system for a turbine en- alone.

gine must be capable of sustained oper- (1) For fuel tanks in the cabin: ation throughout its flow and pressure (i) Upward—4g.

range with fuel initially saturated with (ii) Forward—16g.

water at 80 ° F. and having 0.75cc of free (iii) Sideward—8g.

water per gallon added and cooled to (iv) Downward—20g.

the most critical condition for icing (2) For fuel tanks located above or likely to be encountered in operation.

behind the crew or passenger compart- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as ment that, if loosened, could injure an amended by Amdt. 27–9, 39 FR 35461, Oct. 1, occupant in an emergency landing: 1974] (i) Upward—1.5g.

(ii) Forward—8g.

§ 27.952 Fuel system crash resistance.

(iii) Sideward—2g.

Unless other means acceptable to the (iv) Downward—4g.

Administrator are employed to mini- (3) For fuel tanks in other areas: mize the hazard of fuel fires to occu- (i) Upward—1.5g.

pants following an otherwise surviv- (ii) Forward—4g.

able impact (crash landing), the fuel (iii) Sideward—2g.

systems must incorporate the design (iv) Downward—4g.

features of this section. These systems (c) Fuel line self-sealing breakaway must be shown to be capable of sus- couplings. Self-sealing breakaway cou- taining the static and dynamic decel- plings must be installed unless haz- eration loads of this section, consid- ardous relative motion of fuel system ered as ultimate loads acting alone, components to each other or to local measured at the system component’s rotorcraft structure is demonstrated to center of gravity, without structural be extremely improbable or unless damage to system components, fuel other means are provided. The cou- tanks, or their attachments that would plings or equivalent devices must be leak fuel to an ignition source.

(a) Drop test requirements. Each tank, installed at all fuel tank-to-fuel line or the most critical tank, must be connections, tank-to-tank intercon- drop-tested as follows: nects, and at other points in the fuel (1) The drop height must be at least system where local structural deforma- 50 feet. tion could lead to the release of fuel.

(2) The drop impact surface must be (1) The design and construction of nondeforming. self-sealing breakaway couplings must Federal Aviation Administration, DOT § 27.953 incorporate the following design fea- ture, whether frangible or locally de- tures: formable, must be designed such that (i) The load necessary to separate a its separation or relative local defor- breakaway coupling must be between mation will occur without rupture or 25 to 50 percent of the minimum ulti- local tear-out of the fuel tank or fuel system components that will cause fuel mate failure load (ultimate strength) leakage. The ultimate strength of fran- of the weakest component in the fluid- gible or deformable attachments must carrying line. The separation load be as follows: must in no case be less than 300 pounds, (1) The load required to separate a regardless of the size of the fluid line.

frangible attachment from its support (ii) A breakaway coupling must sepa- structure, or deform a locally deform- rate whenever its ultimate load (as de- able attachment relative to its support fined in paragraph (c)(1)(i) of this sec- structure, must be between 25 and 50 tion) is applied in the failure modes percent of the minimum ultimate load most likely to occur.

(ultimate strength) of the weakest (iii) All breakaway couplings must component in the attached system. In incorporate design provisions to vis- no case may the load be less than 300 ually ascertain that the coupling is pounds.

locked together (leak-free) and is open (2) A frangible or locally deformable during normal installation and service.

attachment must separate or locally (iv) All breakaway couplings must in- deform as intended whenever its ulti- corporate design provisions to prevent mate load (as defined in paragraph uncoupling or unintended closing due (d)(1) of this section) is applied in the to operational shocks, vibrations, or modes most likely to occur.

accelerations.

(3) All frangible or locally deformable (v) No breakaway coupling design attachments must comply with the fa- may allow the release of fuel once the tigue requirements of § 27.571.

coupling has performed its intended (e) Separation of fuel and ignition function.

sources. To provide maximum crash re- (2) All individual breakaway cou- sistance, fuel must be located as far as plings, coupling fuel feed systems, or practicable from all occupiable areas equivalent means must be designed, and from all potential ignition sources.

tested, installed, and maintained so (f) Other basic mechanical design cri- that inadvertent fuel shutoff in flight teria. Fuel tanks, fuel lines, electrical is improbable in accordance with wires, and electrical devices must be § 27.955(a) and must comply with the fa- designed, constructed, and installed, as tigue evaluation requirements of far as practicable, to be crash resist- § 27.571 without leaking.

ant.

(3) Alternate, equivalent means to (g) Rigid or semirigid fuel tanks. Rigid the use of breakaway couplings must or semirigid fuel tank or bladder walls not create a survivable impact-induced must be impact and tear resistant.

load on the fuel line to which it is in- stalled greater than 25 to 50 percent of [Doc. No. 26352, 59 FR 50386, Oct. 3, 1994] the ultimate load (strength) of the § 27.953 Fuel system independence.

weakest component in the line and must comply with the fatigue require- (a) Each fuel system for multiengine ments of § 27.571 without leaking. rotorcraft must allow fuel to be sup- (d) Frangible or deformable structural plied to each engine through a system attachments. Unless hazardous relative independent of those parts of each sys- motion of fuel tanks and fuel system tem supplying fuel to other engines.

components to local rotorcraft struc- However, separate fuel tanks need not ture is demonstrated to be extremely be provided for each engine.

improbable in an otherwise survivable (b) If a single fuel tank is used on a impact, frangible or locally deformable multiengine rotorcraft, the following attachments of fuel tanks and fuel sys- must be provided: tem components to local rotorcraft (1) Independent tank outlets for each structure must be used. The attach- engine, each incorporating a shutoff ment of fuel tanks and fuel system valve at the tank. This shutoff valve components to local rotorcraft struc- may also serve as the firewall shutoff 14 CFR Ch. I (1–1–25 Edition) § 27.954 valve required by § 27.995 if the line be- critical with respect to rotorcraft tween the valve and the engine com- flight attitudes.

partment does not contain a hazardous (4) The critical fuel pump (for pump- amount of fuel that can drain into the fed systems) is installed to produce (by actual or simulated failure) the critical engine compartment.

(2) At least two vents arranged to restriction to fuel flow to be expected minimize the probability of both vents from pump failure.

becoming obstructed simultaneously. (5) Critical values of engine rotation (3) Filler caps designed to minimize speed, electrical power, or other the probability of incorrect installa- sources of fuel pump motive power tion or inflight loss. must be applied.

(4) A fuel system in which those parts (6) Critical values of fuel properties of the system from each tank outlet to which adversely affect fuel flow must any engine are independent of each be applied.

part of each system supplying fuel to (7) The fuel filter required by § 27.997 other engines. must be blocked to the degree nec- essary to simulate the accumulation of § 27.954 Fuel system lightning protec- fuel contamination required to acti- tion.

vate the indicator required by The fuel system must be designed § 27.1305(q).

(b) Fuel transfer systems. If normal op- and arranged to prevent the ignition of fuel vapor within the system by— eration of the fuel system requires fuel (a) Direct lightning strikes to areas to be transferred to an engine feed having a high probability of stroke at- tank, the transfer must occur auto- tachment; matically via a system which has been (b) Swept lightning strokes to areas shown to maintain the fuel level in the where swept strokes are highly prob- engine feed tank within acceptable able; or limits during flight or surface oper- (c) Corona and streamering at fuel ation of the rotorcraft.

vent outlets. (c) Multiple fuel tanks. If an engine can be supplied with fuel from more [Amdt. 27–23, 53 FR 34212, Sept. 2, 1988] than one tank, the fuel systems must, in addition to having appropriate man- § 27.955 Fuel flow.

ual switching capability, be designed to (a) General. The fuel system for each prevent interruption of fuel flow to engine must be shown to provide the that engine, without attention by the engine with at least 100 percent of the flightcrew, when any tank supplying fuel required under each operating and fuel to that engine is depleted of usable maneuvering condition to be approved fuel during normal operation, and any for the rotorcraft including, as applica- other tank that normally supplies fuel ble, the fuel required to operate the en- to the engine alone contains usable gine(s) under the test conditions re- fuel.

quired by § 27.927. Unless equivalent [Amdt. 27–23, 53 FR 34212, Sept. 2, 1988] methods are used, compliance must be shown by test during which the fol- § 27.959 Unusable fuel supply.

lowing provisions are met except that The unusable fuel supply for each combinations of conditions which are tank must be established as not less shown to be improbable need not be than the quantity at which the first considered.

(1) The fuel pressure, corrected for evidence of malfunction occurs under critical accelerations, must be within the most adverse fuel feed condition the limits specified by the engine type occurring under any intended oper- certificate data sheet. ations and flight maneuvers involving (2) The fuel level in the tank may not that tank.

exceed that established as the unusable § 27.961 Fuel system hot weather oper- fuel supply for that tank under § 27.959, ation.

plus the minimum additional fuel nec- essary to conduct the test. Each suction lift fuel system and (3) The fuel head between the tank other fuel systems with features condu- outlet and the engine inlet must be cive to vapor formation must be shown Federal Aviation Administration, DOT § 27.965 by test to operate satisfactorily (with- (g) Each flexible fuel tank bladder or in certification limits) when using fuel liner must be approved or shown to be at a temperature of 110 ° F under crit- suitable for the particular application ical operating conditions including, if and must be puncture resistant. Punc- applicable, the engine operating condi- ture resistance must be shown by tions defined by § 27.927 (b)(1) and (b)(2). meeting the TSO-C80, paragraph 16.0, requirements using a minimum punc- [Amdt. 27–23, 53 FR 34212, Sept. 2, 1988] ture force of 370 pounds.

(h) Each integral fuel tank must have § 27.963 Fuel tanks: general.

provisions for inspection and repair of (a) Each fuel tank must be able to its interior.

withstand, without failure, the vibra- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as tion, inertia, fluid, and structural loads amended by Amdt. 27–23, 53 FR 34213, Sept. 2, to which it may be subjected in oper- 1988; Amdt. 27–30, 59 FR 50387, Oct. 3, 1994] ation.

(b) Each fuel tank of 10 gallons or § 27.965 Fuel tank tests.

greater capacity must have internal (a) Each fuel tank must be able to baffles, or must have external support withstand the applicable pressure tests to resist surging.

in this section without failure or leak- (c) Each fuel tank must be separated age. If practicable, test pressures may from the engine compartment by a be applied in a manner simulating the firewall. At least one-half inch of clear pressure distribution in service.

airspace must be provided between the (b) Each conventional metal tank, tank and the firewall.

nonmetallic tank with walls that are (d) Spaces adjacent to the surfaces of not supported by the rotorcraft struc- fuel tanks must be ventilated so that ture, and integral tank must be sub- fumes cannot accumulate in the tank jected to a pressure of 3.5 p.s.i. unless compartment in case of leakage. If two the pressure developed during max- or more tanks have interconnected imum limit acceleration or emergency outlets, they must be considered as one deceleration with a full tank exceeds tank, and the airspaces in those tanks this value, in which case a hydrostatic must be interconnected to prevent the head, or equivalent test, must be ap- flow of fuel from one tank to another plied to duplicate the acceleration as a result of a difference in pressure loads as far as possible. However, the between those airspaces.

pressure need not exceed 3.5 p.s.i. on (e) The maximum exposed surface surfaces not exposed to the accelera- temperature of any component in the tion loading.

fuel tank must be less, by a safe mar- (c) Each nonmetallic tank with walls gin as determined by the Adminis- supported by the rotorcraft structure trator, than the lowest expected must be subjected to the following autoignition temperature of the fuel or tests: fuel vapor in the tank. Compliance with this requirement must be shown (1) A pressure test of at least 2.0 p.s.i.

under all operating conditions and This test may be conducted on the under all failure or malfunction condi- tank alone in conjunction with the test tions of all components inside the specified in paragraph (c)(2) of this sec- tank. tion.

(2) A pressure test, with the tank (f) Each fuel tank installed in per- mounted in the rotorcraft structure, sonnel compartments must be isolated equal to the load developed by the re- by fume-proof and fuel-proof enclosures action of the contents, with the tank that are drained and vented to the ex- full, during maximum limit accelera- terior of the rotorcraft. The design and tion or emergency deceleration. How- construction of the enclosures must ever, the pressure need not exceed 2.0 provide necessary protection for the p.s.i. on surfaces not exposed to the ac- tank, must be crash resistant during a celeration loading.

survivable impact in accordance with § 27.952, and must be adequate to with- (d) Each tank with large unsupported stand loads and abrasions to be ex- or unstiffened flat areas, or with other pected in personnel compartments. features whose failure or deformation 14 CFR Ch. I (1–1–25 Edition) § 27.967 could cause leakage, must be subjected rocked about each critical axis for 12 ⁄2 to the following test or its equivalent: hours.

(1) Each complete tank assembly and (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 its support must be vibration tested U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 while mounted to simulate the actual U.S.C. 1655(c)) installation.

[Amdt. 27–12, 42 FR 15045, Mar. 17, 1977] (2) The tank assembly must be vi- brated for 25 hours while two-thirds § 27.967 Fuel tank installation.

full of any suitable fluid. The ampli- (a) Each fuel tank must be supported tude of vibration may not be less than so that tank loads are not con- one thirty-second of an inch, unless centrated on unsupported tank sur- otherwise substantiated.

faces. In addition— (3) The test frequency of vibration (1) There must be pads, if necessary, must be as follows: to prevent chafing between each tank (i) If no frequency of vibration result- and its supports; ing from any r.p.m. within the normal (2) The padding must be non- operating range of engine or rotor sys- absorbent or treated to prevent the ab- tem speeds is critical, the test fre- sorption of fuel; quency of vibration, in number of cy- (3) If flexible tank liners are used, cles per minute must, unless a fre- they must be supported so that it is not necessary for them to withstand quency based on a more rational cal- fluid loads; and culation is used, be the number ob- tained by averaging the maximum and (4) Each interior surface of tank com- partments must be smooth and free of minimum power-on engine speeds projections that could cause wear of (r.p.m.) for reciprocating engine pow- the liner unless— ered rotorcraft or 2,000 c.p.m. for tur- (i) There are means for protection of bine engine powered rotorcraft.

the liner at those points; or (ii) If only one frequency of vibration (ii) The construction of the liner resulting from any r.p.m. within the itself provides such protection.

normal operating range of engine or (b) Any spaces adjacent to tank sur- rotor system speeds is critical, that faces must be adequately ventilated to frequency of vibration must be the test avoid accumulation of fuel or fumes in frequency.

those spaces due to minor leakage. If (iii) If more than one frequency of vi- the tank is in a sealed compartment, bration resulting from any r.p.m. with- ventilation may be limited to drain in the normal operating range of en- holes that prevent clogging and exces- gine or rotor system speeds is critical, sive pressure resulting from altitude the most critical of these frequencies changes. If flexible tank liners are in- must be the test frequency.

stalled, the venting arrangement for (4) Under paragraphs (d)(3)(ii) and the spaces between the liner and its (iii) of this section, the time of test container must maintain the proper re- must be adjusted to accomplish the lationship to tank vent pressures for same number of vibration cycles as any expected flight condition.

would be accomplished in 25 hours at (c) The location of each tank must the frequency specified in paragraph meet the requirements of § 27.1185 (a) and (c).

(d)(3)(i) of this section.

(d) No rotorcraft skin immediately (5) During the test, the tank assem- adjacent to a major air outlet from the bly must be rocked at the rate of 16 to engine compartment may act as the 20 complete cycles per minute through wall of the integral tank.

an angle of 15 degrees on both sides of the horizontal (30 degrees total), about [Doc. No. 26352, 59 FR 50387, Oct. 3, 1994] the most critical axis, for 25 hours. If motion about more than one axis is § 27.969 Fuel tank expansion space.

likely to be critical, the tank must be Each fuel tank or each group of fuel tanks with interconnected vent sys- tems must have an expansion space of Federal Aviation Administration, DOT § 27.993 not less than 2 percent of the tank ca- space so that venting is effective under pacity. It must be impossible to fill the all normal flight conditions. Each vent fuel tank expansion space inadvert- must minimize the probability of stop- ently with the rotorcraft in the normal page by dirt or ice.

ground attitude. (b) The venting system must be de- signed to minimize spillage of fuel [Amdt. 27–23, 53 FR 34213, Sept. 2, 1988] through the vents to an ignition source in the event of a rollover during land- § 27.971 Fuel tank sump.

ing, ground operation, or a survivable (a) Each fuel tank must have a drain- impact.

able sump with an effective capacity in any ground attitude to be expected in [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–23, 53 FR 34213, Sept. 2, service of 0.25 percent of the tank ca- 1988; Amdt. 27–30, 59 FR 50387, Oct. 3, 1994; pacity or ⁄ 16 gallon, whichever is great- Amdt. 27–35, 63 FR 43285, Aug. 12, 1998] er, unless— (1) The fuel system has a sediment § 27.977 Fuel tank outlet.

bowl or chamber that is accessible for (a) There must be a fuel stainer for preflight drainage and has a minimum the fuel tank outlet or for the booster capacity of 1 ounce for every 20 gallons pump. This strainer must— of fuel tank capacity; and (1) For reciprocating engine powered (2) Each fuel tank drain is located so rotorcraft, have 8 to 16 meshes per that in any ground attitude to be ex- inch; and pected in service, water will drain from (2) For turbine engine powered rotor- all parts of the tank to the sediment craft, prevent the passage of any object bowl or chamber.

that could restrict fuel flow or damage (b) Each sump, sediment bowl, and any fuel system component.

sediment chamber drain required by (b) The clear area of each fuel tank this section must comply with the outlet strainer must be at least five drain provisions of § 27.999(b).

times the area of the outlet line.

[Amdt. 27–23, 53 FR 34213, Sept. 2, 1988] (c) The diameter of each strainer must be at least that of the fuel tank § 27.973 Fuel tank filler connection.

outlet.

(a) Each fuel tank filler connection (d) Each finger strainer must be ac- must prevent the entrance of fuel into cessible for inspection and cleaning.

any part of the rotorcraft other than [Amdt. 27–11, 41 FR 55470, Dec. 20, 1976] the tank itself during normal oper- ations and must be crash resistant dur- F UEL S YSTEM C OMPONENTS ing a survivable impact in accordance with § 27.952(c). In addition— § 27.991 Fuel pumps.

(1) Each filler must be marked as pre- Compliance with § 27.955 may not be scribed in § 27.1557(c)(1); jeopardized by failure of— (2) Each recessed filler connection (a) Any one pump except pumps that that can retain any appreciable quan- are approved and installed as parts of a tity of fuel must have a drain that dis- type certificated engine; or charges clear of the entire rotorcraft; (b) Any component required for pump and operation except, for engine driven (3) Each filler cap must provide a pumps, the engine served by that fuel-tight seal under the fluid pressure pump.

expected in normal operation and in a survivable impact.

[Amdt. 27–23, 53 FR 34213, Sept. 2, 1988] (b) Each filler cap or filler cap cover must warn when the cap is not fully § 27.993 Fuel system lines and fittings.

locked or seated on the filler connec- (a) Each fuel line must be installed tion.

and supported to prevent excessive vi- [Doc. No. 26352, 59 FR 50387, Oct. 3, 1994] bration and to withstand loads due to fuel pressure and accelerated flight § 27.975 Fuel tank vents.

conditions.

(a) Each fuel tank must be vented (b) Each fuel line connected to com- from the top part of the expansion ponents of the rotorcraft between 14 CFR Ch. I (1–1–25 Edition) § 27.995 which relative motion could exist must fuel system or engine fuel system oper- have provisions for flexibility. ation.

(c) Flexible hose must be approved.

[Amdt. 27–9, 39 FR 35461, Oct. 1, 1974, as (d) Each flexible connection in fuel amended by Amdt. 27–20, 49 FR 6849, Feb. 23, lines that may be under pressure or 1984; Amdt. 27–23, 53 FR 34213, Sept. 2, 1988] subjected to axial loading must use flexible hose assemblies.

§ 27.999 Fuel system drains.

(e) No flexible hose that might be ad- (a) There must be at least one acces- versely affected by high temperatures sible drain at the lowest point in each may be used where excessive tempera- fuel system to completely drain the tures will exist during operation or after engine shutdown. system with the rotorcraft in any ground attitude to be expected in serv- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as ice.

amended by Amdt. 27–2, 33 FR 964, Jan. 26, 1968] (b) Each drain required by paragraph (a) of this section must— § 27.995 Fuel valves.

(1) Discharge clear of all parts of the (a) There must be a positive, quick- rotorcraft; acting valve to shut off fuel to each en- (2) Have manual or automatic means gine individually.

to assure positive closure in the off po- (b) The control for this valve must be sition; and within easy reach of appropriate crew- (3) Have a drain valve— members.

(i) That is readily accessible and (c) Where there is more than one which can be easily opened and closed; source of fuel supply there must be and means for independent feeding from (ii) That is either located or pro- each source.

tected to prevent fuel spillage in the (d) No shutoff valve may be on the event of a landing with landing gear re- engine side of any firewall.

tracted.

§ 27.997 Fuel strainer or filter.

[Doc. No. 574, 29 FR 15695, Nov. 24, 1964, as There must be a fuel strainer or filter amended by Amdt. 27–11, 41 FR 55470, Dec. 20, between the fuel tank outlet and the 1976; Amdt. 27–23, 53 FR 34213, Sept. 2, 1988] inlet of the first fuel system compo- nent which is susceptible to fuel con- O IL S YSTEM tamination, including but not limited § 27.1011 Engines: General.

to the fuel metering device or an en- gine positive displacement pump, (a) Each engine must have an inde- whichever is nearer the fuel tank out- pendent oil system that can supply it let. This fuel strainer or filter must— with an appropriate quantity of oil at a (a) Be accessible for draining and temperature not above that safe for cleaning and must incorporate a screen continuous operation.

or element which is easily removable; (b) The usable oil capacity of each (b) Have a sediment trap and drain system may not be less than the prod- except that it need not have a drain if uct of the endurance of the rotorcraft the strainer or filter is easily remov- under critical operating conditions and able for drain purposes; the maximum oil consumption of the (c) Be mounted so that its weight is engine under the same conditions, plus not supported by the connecting lines or by the inlet or outlet connections of a suitable margin to ensure adequate the strainer or filter itself, unless ade- circulation and cooling. Instead of a ra- quate strength margins under all load- tional analysis of endurance and con- ing conditions are provided in the lines sumption, a usable oil capacity of one and connections; and gallon for each 40 gallons of usable fuel (d) Provide a means to remove from may be used.

the fuel any contaminant which would (c) The oil cooling provisions for each jeopardize the flow of fuel through engine must be able to maintain the oil rotorcraft or engine fuel system com- inlet temperature to that engine at or ponents required for proper rotorcraft Federal Aviation Administration, DOT § 27.1021 below the maximum established value.

§ 27.1019 Oil strainer or filter.

This must be shown by flight tests.

(a) Each turbine engine installation [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as must incorporate an oil strainer or fil- amended by Amdt. 27–23, 53 FR 34213, Sept. 2, ter through which all of the engine oil 1988] flows and which meets the following re- quirements: § 27.1013 Oil tanks.

(1) Each oil strainer or filter that has Each oil tank must be designed and a bypass must be constructed and in- installed so that— stalled so that oil will flow at the nor- (a) It can withstand, without failure, mal rate through the rest of the sys- each vibration, inertia, fluid, and tem with the strainer or filter com- structural load expected in operation; pletely blocked.

(b) [Reserved] (2) The oil strainer or filter must (c) Where used with a reciprocating have the capacity (with respect to op- engine, it has an expansion space of not erating limitations established for the less than the greater of 10 percent of engine) to ensure that engine oil sys- the tank capacity or 0.5 gallon, and tem functioning is not impaired when where used with a turbine engine, it has an expansion space of not less than the oil is contaminated to a degree 10 percent of the tank capacity. (with respect to particle size and den- (d) It is impossible to fill the tank sity) that is greater than that estab- expansion space inadvertently with the lished for the engine under Part 33 of rotorcraft in the normal ground atti- this chapter.

tude; (3) The oil strainer or filter, unless it (e) Adequate venting is provided; and is installed at an oil tank outlet, must (f) There are means in the filler open- incorporate a means to indicate con- ing to prevent oil overflow from enter- tamination before it reaches the capac- ing the oil tank compartment.

ity established in accordance with paragraph (a)(2) of this section.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–9, 39 FR 35461, Oct. 1, (4) The bypass of a strainer or filter 1974] must be constructed and installed so that the release of collected contami- § 27.1015 Oil tank tests.

nants is minimized by appropriate lo- Each oil tank must be designed and cation of the bypass to ensure that col- installed so that it can withstand, lected contaminants are not in the by- without leakage, an internal pressure pass flow path.

of 5 p.s.i., except that each pressurized (5) An oil strainer or filter that has oil tank used with a turbine engine no bypass, except one that is installed must be designed and installed so that at an oil tank outlet, must have a it can withstand, without leakage, an means to connect it to the warning internal pressure of 5 p.s.i., plus the system required in § 27.1305(r).

maximum operating pressure of the (b) Each oil strainer or filter in a tank.

powerplant installation using recipro- [Amdt. 27–9, 39 FR 35462, Oct. 1, 1974] cating engines must be constructed and installed so that oil will flow at the § 27.1017 Oil lines and fittings.

normal rate through the rest of the (a) Each oil line must be supported to system with the strainer or filter ele- prevent excessive vibration.

ment completely blocked.

(b) Each oil line connected to compo- [Amdt. 27–9, 39 FR 35462, Oct. 1, 1974, as nents of the rotorcraft between which amended by Amdt. 27–20, 49 FR 6849, Feb. 23, relative motion could exist must have 1984; Amdt. 27–23, 53 FR 34213, Sept. 2, 1988] provisions for flexibility.

(c) Flexible hose must be approved.

§ 27.1021 Oil system drains.

(d) Each oil line must have an inside A drain (or drains) must be provided diameter of not less than the inside di- to allow safe drainage of the oil sys- ameter of the engine inlet or outlet. No tem. Each drain must— line may have splices between connec- tions. (a) Be accessible; and 14 CFR Ch. I (1–1–25 Edition) § 27.1027 (b) Have manual or automatic means (e) Splash-type lubrication systems for positive locking in the closed posi- for rotor drive system gearboxes must comply with §§ 27.1021 and 27.1337(d).

tion.

[Amdt. 27–23, 53 FR 34213, Sept. 2, 1988, as [Amdt. 27–20, 49 FR 6849, Feb. 23, 1984] amended by Amdt. 27–37, 64 FR 45095, Aug. 18, 1999] § 27.1027 Transmissions and gear- boxes: General.

C OOLING (a) The lubrication system for com- § 27.1041 General.

ponents of the rotor drive system that require continuous lubrication must be (a) Each powerplant cooling system sufficiently independent of the lubrica- must be able to maintain the tempera- tion systems of the engine(s) to ensure tures of powerplant components within the limits established for these compo- lubrication during autorotation.

nents under critical surface (ground or (b) Pressure lubrication systems for water) and flight operating conditions transmissions and gearboxes must for which certification is required and comply with the engine oil system re- after normal shutdown. Powerplant quirements of §§ 27.1013 (except para- components to be considered include graph (c)), 27.1015, 27.1017, 27.1021, and but may not be limited to engines, 27.1337(d).

rotor drive system components, auxil- (c) Each pressure lubrication system iary power units, and the cooling or lu- must have an oil strainer or filter bricating fluids used with these compo- through which all of the lubricant nents.

flows and must— (b) Compliance with paragraph (a) of (1) Be designed to remove from the this section must be shown in tests lubricant any contaminant which may conducted under the conditions pre- damage transmission and drive system scribed in that paragraph.

components or impede the flow of lu- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as bricant to a hazardous degree; amended by Amdt. 27–23, 53 FR 34213, Sept. 2, (2) Be equipped with a means to indi- 1988] cate collection of contaminants on the § 27.1043 Cooling tests.

filter or strainer at or before opening of the bypass required by paragraph (a) General. For the tests prescribed (c)(3) of this section; and in § 27.1041(b), the following apply: (3) Be equipped with a bypass con- (1) If the tests are conducted under structed and installed so that— conditions deviating from the max- imum ambient atmospheric tempera- (i) The lubricant will flow at the nor- ture specified in paragraph (b) of this mal rate through the rest of the sys- section, the recorded powerplant tem- tem with the strainer or filter com- peratures must be corrected under pletely blocked; and paragraphs (c) and (d) of this section (ii) The release of collected contami- unless a more rational correction nants is minimized by appropriate lo- method is applicable.

cation of the bypass to ensure that col- (2) No corrected temperature deter- lected contaminants are not in the by- mined under paragraph (a)(1) of this pass flowpath.

section may exceed established limits.

(d) For each lubricant tank or sump (3) For reciprocating engines, the fuel outlet supplying lubrication to rotor used during the cooling tests must be drive systems and rotor drive system of the minimum grade approved for the components, a screen must be provided engines, and the mixture settings must to prevent entrance into the lubrica- be those normally used in the flight tion system of any object that might stages for which the cooling tests are obstruct the flow of lubricant from the conducted.

outlet to the filter required by para- (4) The test procedures must be as graph (c) of this section. The require- prescribed in § 27.1045.

ments of paragraph (c) do not apply to (b) Maximum ambient atmospheric tem- screens installed at lubricant tank or perature. A maximum ambient atmos- sump outlets. pheric temperature corresponding to Federal Aviation Administration, DOT § 27.1091 sea level conditions of at least 100 de- engine fluid temperature stabilization grees F. must be established. The as- rules apply: sumed temperature lapse rate is 3.6 de- (1) For each rotorcraft, and for each grees F. per thousand feet of altitude stage of flight— above sea level until a temperature of (i) The temperatures must be sta- ¥ 69.7 degrees F. is reached, above bilized under the conditions from which altitude the temperature is con- which entry is made into the stage of sidered constant at ¥ 69.7 degrees F.

flight being investigated; or However, for winterization installa- (ii) If the entry condition normally tions, the applicant may select a max- does not allow temperatures to sta- imum ambient atmospheric tempera- bilize, operation through the full entry ture corresponding to sea level condi- condition must be conducted before tions of less than 100 degrees F.

entry into the stage of flight being in- (c) Correction factor (except cylinder vestigated in order to allow the tem- barrels). Unless a more rational correc- peratures to attain their natural levels tion applies, temperatures of engine at the time of entry.

fluids and power-plant components (ex- (2) For each helicopter during the cept cylinder barrels) for which tem- takeoff stage of flight, the climb at perature limits are established, must takeoff power must be preceded by a be corrected by adding to them the dif- period of hover during which the tem- ference between the maximum ambient peratures are stabilized.

atmospheric temperature and the tem- (c) Duration of test. For each stage of perature of the ambient air at the time flight the tests must be continued of the first occurrence of the maximum until— component or fluid temperature re- (1) The temperatures stabilize or 5 corded during the cooling test.

minutes after the occurrence of the (d) Correction factor for cylinder barrel highest temperature recorded, as ap- temperatures. Cylinder barrel tempera- propriate to the test condition; tures must be corrected by adding to (2) That stage of flight is completed; them 0.7 times the difference between or the maximum ambient atmospheric (3) An operating limitation is temperature and the temperature of reached.

the ambient air at the time of the first [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as occurrence of the maximum cylinder amended by Amdt. 27–23, 53 FR 34214, Sept. 2, barrel temperature recorded during the 1988] cooling test.

I NDUCTION S YSTEM (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), § 27.1091 Air induction.

1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C.

(a) The air induction system for each 1655(c))) engine must supply the air required by [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as that engine under the operating condi- amended by Amdt. 27–11, 41 FR 55470, Dec. 20, tions and maneuvers for which certifi- 1976; Amdt. 27–14, 43 FR 2325, Jan. 16, 1978] cation is requested.

(b) Each cold air induction system § 27.1045 Cooling test procedures.

opening must be outside the cowling if (a) General. For each stage of flight, backfire flames can emerge.

the cooling tests must be conducted (c) If fuel can accumulate in any air with the rotorcraft— induction system, that system must have drains that discharge fuel— (1) In the configuration most critical for cooling; and (1) Clear of the rotorcraft; and (2) Under the conditions most critical (2) Out of the path of exhaust flames.

for cooling.

(d) For turbine engine powered rotor- (b) Temperature stabilization. For the craft— purpose of the cooling tests, a tempera- (1) There must be means to prevent ture is ‘‘stabilized’’ when its rate of hazardous quantities of fuel leakage or change is less than two degrees F. per overflow from drains, vents, or other minute. The following component and components of flammable fluid systems 14 CFR Ch. I (1–1–25 Edition) § 27.1093 from entering the engine intake sys- ation, within the limitations estab- tem; and lished for the rotorcraft.

(2) The air inlet ducts must be lo- (2) Each turbine engine must idle for cated or protected so as to minimize 30 minutes on the ground, with the air the ingestion of foreign matter during bleed available for engine icing protec- takeoff, landing, and taxiing.

tion at its critical condition, without adverse effect, in an atmosphere that is [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as at a temperature between 15 ° and 30 ° F amended by Amdt. 27–2, 33 FR 964, Jan. 26, (between ¥ 9 ° and ¥ 1 ° C) and has a liq- 1968; Amdt. 27–23, 53 FR 34214, Sept. 2, 1988] uid water content not less than 0.3 § 27.1093 Induction system icing pro- gram per cubic meter in the form of tection.

drops having a mean effective diameter (a) Reciprocating engines. Each recip- not less than 20 microns, followed by rocating engine air induction system momentary operation at takeoff power must have means to prevent and elimi- or thrust. During the 30 minutes of idle nate icing. Unless this is done by other operation, the engine may be run up means, it must be shown that, in air periodically to a moderate power or free of visible moisture at a tempera- thrust setting in a manner acceptable ture of 30 degrees F., and with the en- to the Administrator.

gines at 75 percent of maximum contin- (c) Supercharged reciprocating engines.

uous power— For each engine having superchargers (1) Each rotorcraft with sea level en- to pressurize the air before it enters gines using conventional venturi car- the carburetor, the heat rise in the air buretors has a preheater that can pro- caused by that supercharging at any vide a heat rise of 90 degrees F.; altitude may be utilized in determining (2) Each rotorcraft with sea level en- compliance with paragraph (a) of this gines using carburetors tending to pre- section if the heat rise utilized is that vent icing has a sheltered alternate which will be available, automatically, source of air, and that the preheat sup- for the applicable altitude and oper- plied to the alternate air intake is not ating condition because of super- less than that provided by the engine charging.

cooling air downstream of the cyl- (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 inders; U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 (3) Each rotorcraft with altitude en- U.S.C. 1655(c)) gines using conventional venturi car- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as buretors has a preheater capable of amended by Amdt. 27–11, 41 FR 55470, Dec. 20, providing a heat rise of 120 degrees F.; 1976; Amdt. 27–12, 42 FR 15045, Mar. 17, 1977; and Amdt. 27–20, 49 FR 6849, Feb. 23, 1984; Amdt.

(4) Each rotorcraft with altitude en- 27–23, 53 FR 34214, Sept. 2, 1988] gines using carburetors tending to pre- vent icing has a preheater that can E XHAUST S YSTEM provide a heat rise of— (i) 100 degrees F.; or § 27.1121 General.

(ii) If a fluid deicing system is used, For each exhaust system— at least 40 degrees F.

(a) There must be means for thermal (b) Turbine engine. (1) It must be expansion of manifolds and pipes; shown that each turbine engine and its (b) There must be means to prevent air inlet system can operate through- local hot spots; out the flight power range of the en- (c) Exhaust gases must discharge gine (including idling)— clear of the engine air intake, fuel sys- (i) Without accumulating ice on en- tem components, and drains; gine or inlet system components that would adversely affect engine oper- (d) Each exhaust system part with a ation or cause a serious loss of power surface hot enough to ignite flammable under the icing conditions specified in fluids or vapors must be located or appendix C of Part 29 of this chapter; shielded so that leakage from any sys- and tem carrying flammable fluids or va- (ii) In snow, both falling and blowing, pors will not result in a fire caused by without adverse effect on engine oper- impingement of the fluids or vapors on Federal Aviation Administration, DOT § 27.1143 any part of the exhaust system includ- (2) For power-assisted valves, a ing shields for the exhaust system; means to indicate to the flight crew (e) Exhaust gases may not impair when the valve— pilot vision at night due to glare; (i) Is in the fully open or fully closed (f) If significant traps exist, each tur- position; or bine engine exhaust system must have (ii) Is moving between the fully open drains discharging clear of the rotor- and fully closed position.

craft, in any normal ground and flight (e) For turbine engine powered rotor- attitudes, to prevent fuel accumulation craft, no single failure or malfunction, after the failure of an attempted en- or probable combination thereof, in gine start; any powerplant control system may (g) Each exhaust heat exchanger cause the failure of any powerplant must incorporate means to prevent function necessary for safety.

blockage of the exhaust port after any (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 internal heat exchanger failure.

U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1655(c)) U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as U.S.C. 1655(c)) amended by Amdt. 27–12, 42 FR 15045, Mar. 17, [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as 1977; Amdt. 27–23, 53 FR 34214, Sept. 2, 1988; amended by Amdt. 27–12, 42 FR 15045, Mar. 17, Amdt. 27–33, 61 FR 21907, May 10, 1996] 1977] § 27.1143 Engine controls.

§ 27.1123 Exhaust piping.

(a) There must be a separate power (a) Exhaust piping must be heat and control for each engine.

corrosion resistant, and must have pro- (b) Power controls must be grouped visions to prevent failure due to expan- and arranged to allow— sion by operating temperatures.

(1) Separate control of each engine; (b) Exhaust piping must be supported and to withstand any vibration and inertia (2) Simultaneous control of all en- loads to which it would be subjected in gines.

operations.

(c) Each power control must provide (c) Exhaust piping connected to com- a positive and immediately responsive ponents between which relative motion could exist must have provisions for means of controlling its engine.

flexibility.

(d) If a power control incorporates a fuel shutoff feature, the control must [Amdt. 27–11, 41 FR 55470, Dec. 20, 1976] have a means to prevent the inad- vertent movement of the control into P OWERPLANT CONTROLS AND the shutoff position. The means must— A CCESSORIES (1) Have a positive lock or stop at the § 27.1141 Powerplant controls: general.

idle position; and (2) Require a separate and distinct (a) Powerplant controls must be lo- operation to place the control in the cated and arranged under § 27.777 and shutoff position.

marked under § 27.1555.

(e) For rotorcraft to be certificated (b) Each flexible powerplant control for a 30-second OEI power rating, a must be approved.

means must be provided to automati- (c) Each control must be able to cally activate and control the 30-sec- maintain any set position without— ond OEI power and prevent any engine (1) Constant attention; or from exceeding the installed engine (2) Tendency to creep due to control limits associated with the 30-second loads or vibration.

OEI power rating approved for the (d) Controls of powerplant valves re- rotorcraft.

quired for safety must have— (1) For manual valves, positive stops [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as or in the case of fuel valves suitable amended by Amdt. 27–11, 41 FR 55470, Dec. 20, index provisions, in the open and closed 1976; Amdt. 27–23, 53 FR 34214, Sept. 2, 1988; position; and Amdt. 27–29, 59 FR 47767, Sept. 16, 1994] 14 CFR Ch. I (1–1–25 Edition) § 27.1145 P OWERPLANT F IRE P ROTECTION § 27.1145 Ignition switches.

(a) There must be means to quickly § 27.1183 Lines, fittings, and compo- shut off all ignition by the grouping of nents.

switches or by a master ignition con- (a) Except as provided in paragraph trol.

(b) of this section, each line, fitting, (b) Each group of ignition switches, and other component carrying flam- except ignition switches for turbine en- mable fluid in any area subject to en- gines for which continuous ignition is gine fire conditions must be fire resist- not required, and each master ignition ant, except that flammable fluid tanks control must have a means to prevent and supports which are part of and at- its inadvertent operation.

tached to the engine must be fireproof (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 or be enclosed by a fireproof shield un- U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 less damage by fire to any non-fire- U.S.C. 1655(c)) proof part will not cause leakage or spillage of flammable fluid. Compo- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–12, 42 FR 15045, Mar. 17, nents must be shielded or located so as 1977] to safeguard against the ignition of leaking flammable fluid. An integral § 27.1147 Mixture controls.

oil sump of less than 25-quart capacity on a reciprocating engine need not be If there are mixture controls, each fireproof nor be enclosed by a fireproof engine must have a separate control shield.

and the controls must be arranged to (b) Paragraph (a) does not apply to— allow— (a) Separate control of each engine; (1) Lines, fittings, and components and which are already approved as part of a type certificated engine; and (b) Simultaneous control of all en- (2) Vent and drain lines, and their fit- gines.

tings, whose failure will not result in, § 27.1151 Rotor brake controls.

or add to, a fire hazard.

(c) Each flammable fluid drain and (a) It must be impossible to apply the vent must discharge clear of the induc- rotor brake inadvertently in flight.

tion system air inlet.

(b) There must be means to warn the crew if the rotor brake has not been [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as completely released before takeoff.

amended by Amdt. 27–1, 32 FR 6914, May 5, 1967; Amdt. 27–9, 39 FR 35462, Oct. 1, 1974; [Doc. No. 28008, 61 FR 21907, May 10, 1996] Amdt. 27–20, 49 FR 6849, Feb. 23, 1984] § 27.1163 Powerplant accessories.

§ 27.1185 Flammable fluids.

(a) Each engine-mounted accessory (a) Each fuel tank must be isolated must— from the engines by a firewall or (1) Be approved for mounting on the shroud.

engine involved; (b) Each tank or reservoir, other (2) Use the provisions on the engine than a fuel tank, that is part of a sys- for mounting; and tem containing flammable fluids or (3) Be sealed in such a way as to pre- gases must be isolated from the engine vent contamination of the engine oil by a firewall or shroud, unless the de- system and the accessory system.

sign of the system, the materials used (b) Unless other means are provided, in the tank and its supports, the shut- torque limiting means must be pro- off means, and the connections, lines vided for accessory drives located on and controls provide a degree of safety any component of the transmission and equal to that which would exist if the rotor drive system to prevent damage tank or reservoir were isolated from to these components from excessive ac- the engines.

cessory load.

(c) There must be at least one-half inch of clear airspace between each [Amdt. 27–2, 33 FR 964, Jan. 26, 1968, as tank and each firewall or shroud iso- amended by Amdt. 27–20, 49 FR 6849, Feb. 23, 1984; Amdt. 27–23, 53 FR 34214, Sept. 2, 1988] lating that tank, unless equivalent Federal Aviation Administration, DOT § 27.1193 means are used to prevent heat trans- § 27.1191 Firewalls.

fer from each engine compartment to (a) Each engine, including the com- the flammable fluid.

bustor, turbine, and tailpipe sections of (d) Absorbent materials close to turbine engines must be isolated by a flammable fluid system components firewall, shroud, or equivalent means, that might leak must be covered or from personnel compartments, struc- treated to prevent the absorption of tures, controls, rotor mechanisms, and hazardous quantities of fluids.

other parts that are— (1) Essential to a controlled landing: [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as and amended by Amdt. 27–2, 33 FR 964, Jan. 26, (2) Not protected under § 27.861.

1968; Amdt. 27–11, 41 FR 55470, Dec. 20, 1976; (b) Each auxiliary power unit and Amdt. 27–37, 64 FR 45095, Aug. 18, 1999] combustion heater, and any other com- bustion equipment to be used in flight, § 27.1187 Ventilation and drainage.

must be isolated from the rest of the Each compartment containing any rotorcraft by firewalls, shrouds, or part of the powerplant installation equivalent means.

must have provision for ventilation (c) In meeting paragraphs (a) and (b) and drainage of flammable fluids. The of this section, account must be taken drainage means must be— of the probable path of a fire as af- (a) Effective under conditions ex- fected by the airflow in normal flight pected to prevail when drainage is and in autorotation.

needed, and (d) Each firewall and shroud must be (b) Arranged so that no discharged constructed so that no hazardous quan- fluid will cause an additional fire haz- tity of air, fluids, or flame can pass ard. from any engine compartment to other parts of the rotorcraft.

[Doc. No. 29247, 64 FR 45095, Aug. 18, 1999] (e) Each opening in the firewall or shroud must be sealed with close-fit- § 27.1189 Shutoff means.

ting, fireproof grommets, bushings, or (a) There must be means to shut off firewall fittings.

each line carrying flammable fluids (f) Each firewall and shroud must be into the engine compartment, except— fireproof and protected against corro- sion.

(1) Lines, fittings, and components forming an intergral part of an engine; [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (2) For oil systems for which all com- amended by Amdt. 27–2, 22 FR 964, Jan. 26, ponents of the system, including oil 1968] tanks, are fireproof or located in areas § 27.1193 Cowling and engine compart- not subject to engine fire conditions; ment covering.

and (a) Each cowling and engine compart- (3) For reciprocating engine installa- ment covering must be constructed and tions only, engine oil system lines in supported so that it can resist the vi- installation using engines of less than bration, inertia, and air loads to which 500 cu. in. displacement.

it may be subjected in operation.

(b) There must be means to guard (b) There must be means for rapid against inadvertent operation of each and complete drainage of each part of shutoff, and to make it possible for the the cowling or engine compartment in crew to reopen it in flight after it has the normal ground and flight attitudes.

been closed.

(c) No drain may discharge where it (c) Each shutoff valve and its control might cause a fire hazard.

must be designed, located, and pro- (d) Each cowling and engine compart- tected to function properly under any ment covering must be at least fire re- condition likely to result from an en- sistant.

gine fire.

(e) Each part of the cowling or engine compartment covering subject to high [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as temperatures due to its nearness to ex- amended by Amdt. 27–2, 33 FR 964, Jan. 26, haust system parts or exhaust gas im- 1968; Amdt. 27–20, 49 FR 6850, Feb. 23, 1984; Amdt. 27–23, 53 FR 34214, Sept. 2, 1988] pingement must be fireproof.

14 CFR Ch. I (1–1–25 Edition) § 27.1194 (f) A means of retaining each open- § 27.1305 Powerplant instruments.

able or readily removable panel, cowl- The following are the required power- ing, or engine or rotor drive system plant instruments: covering must be provided to preclude (a) A carburetor air temperature in- hazardous damage to rotors or critical dicator, for each engine having a pre- control components in the event of heater that can provide a heat rise in structural or mechanical failure of the excess of 60 ° F.

normal retention means, unless such (b) A cylinder head temperature indi- failure is extremely improbable.

cator, for each— [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (1) Air cooled engine; amended by Amdt. 27–23, 53 FR 34214, Sept. 2, (2) Rotorcraft with cooling shutters; 1988] and § 27.1194 Other surfaces.

(3) Rotorcraft for which compliance with § 27.1043 is shown in any condition All surfaces aft of, and near, power- other than the most critical flight con- plant compartments, other than tail dition with respect to cooling.

surfaces not subject to heat, flames, or (c) A fuel pressure indicator, for each sparks emanating from a powerplant pump-fed engine.

compartment, must be at least fire re- (d) A fuel quantity indicator, for each sistant.

fuel tank.

[Amdt. 27–2, 33 FR 964, Jan. 26, 1968] (e) A means to indicate manifold pressure for each altitude engine.

§ 27.1195 Fire detector systems.

(f) An oil temperature warning device Each turbine engine powered rotor- to indicate when the temperature ex- craft must have approved quick-acting ceeds a safe value in each main rotor fire detectors in numbers and locations drive gearbox (including any gearboxes insuring prompt detection of fire in the essential to rotor phasing) having an engine compartment which cannot be oil system independent of the engine readily observed in flight by the pilot oil system.

in the cockpit.

(g) An oil pressure warning device to [Amdt. 27–5, 36 FR 5493, Mar. 24, 1971] indicate when the pressure falls below a safe value in each pressure-lubricated main rotor drive gearbox (including Subpart F—Equipment any gearboxes essential to rotor phas- G ENERAL ing) having an oil system independent of the engine oil system.

§ 27.1301 Function and installation.

(h) An oil pressure indicator for each engine.

Each item of installed equipment (i) An oil quantity indicator for each must— oil tank.

(a) Be of a kind and design appro- priate to its intended function; (j) An oil temperature indicator for each engine.

(b) Be labeled as to its identification, function, or operating limitations, or (k) A means to indicate the r.p.m. of any applicable combination of these each engine and at least one tachom- factors; eter, as applicable, for: (c) Be installed according to limita- (1) The r.p.m. of the single main tions specified for that equipment; and rotor; (d) Function properly when installed.

(2) The common r.p.m. of any main rotors whose speeds cannot vary appre- § 27.1303 Flight and navigation instru- ciably with respect to each other; or ments.

(3) The r.p.m. of each main rotor The following are the required flight whose speed can vary appreciably with and navigation instruments: respect to that of another main rotor.

(a) An airspeed indicator.

(l) A low fuel warning device for each (b) An altimeter.

fuel tank which feeds an engine. This (c) A magnetic direction indicator. device must— Federal Aviation Administration, DOT § 27.1309 (1) Provide a warning to the netic particles are detected by the chip flightcrew when approximately 10 min- detector required by § 27.1337(e).

utes of usable fuel remains in the tank; [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as and amended by Amdt. 27–9, 39 FR 35462, Oct. 1, (2) Be independent of the normal fuel 1974; Amdt. 27–23, 53 FR 34214, Sept. 2, 1988; Amdt. 27–29, 59 FR 47767, Sept. 16, 1994; Amdt.

quantity indicating system.

27–37, 64 FR 45095, Aug. 18, 1999; 64 FR 47563, (m) Means to indicate to the Aug. 31, 1999; Amdt. 27–51, 88 FR 8737, Feb. 10, flightcrew the failure of any fuel pump 2023] installed to show compliance with § 27.955.

§ 27.1307 Miscellaneous equipment.

(n) A means to indicate the gas tem- The following is the required mis- perature for each turbine engine.

cellaneous equipment: (o) A means to enable the pilot to de- (a) An approved seat for each occu- termine the torque of each turbine en- pant.

gine, if a torque limitation is estab- (b) An approved safety belt for each lished for that engine under § 27.1521(e).

occupant.

(p) For each turbine engine, an indi- (c) A master switch arrangement.

cator to indicate the functioning of the (d) An adequate source of electrical powerplant ice protection system.

energy, where electrical energy is nec- (q) An indicator for the fuel filter re- essary for operation of the rotorcraft.

quired by § 27.997 to indicate the occur- (e) Electrical protective devices.

rence of contamination of the filter at § 27.1309 Equipment, systems, and in- the degree established by the applicant stallations.

in compliance with § 27.955.

(r) For each turbine engine, a warn- The equipment, systems, and instal- ing means for the oil strainer or filter lations whose functioning is required required by § 27.1019, if it has no bypass, by this subchapter must be designed and installed to ensure that they per- to warn the pilot of the occurrence of form their intended functions under contamination of the strainer or filter any foreseeable operating condition.

before it reaches the capacity estab- For any item of equipment or system lished in accordance with § 27.1019(a)(2).

whose failure has not been specifically (s) An indicator to indicate the func- addressed by another requirement in tioning of any selectable or control- this chapter, the following require- lable heater used to prevent ice clog- ments also apply: ging of fuel system components.

(a) The design of each item of equip- (t) For rotorcraft for which a 30-sec- ment, system, and installation must be ond/2-minute OEI power rating is re- analyzed separately and in relation to quested, a means must be provided to other rotorcraft systems and installa- alert the pilot when the engine is at tions to determine and identify any the 30-second and the 2-minute OEI failure that would affect the capability power levels, when the event begins, of the rotorcraft or the ability of the and when the time interval expires.

crew to perform their duties in all op- (u) For each turbine engine utilizing erating conditions.

30-second/2-minute OEI power, a device (b) Each item of equipment, system, or system must be provided for use by and installation must be designed and ground personnel which— installed so that: (1) Automatically records each usage (1) The occurrence of any cata- and duration of power at the 30-second strophic failure condition is extremely and 2-minute OEI levels; improbable; (2) Permits retrieval of the recorded (2) The occurrence of any major fail- data; ure condition is no more than improb- (3) Can be reset only by ground main- able; and tenance personnel; and (3) For the occurrence of any other (4) Has a means to verify proper oper- failure condition between major and ation of the system or device.

catastrophic, the probability of the (v) Warning or caution devices to sig- failure condition must be inversely nal to the flight crew when ferromag- proportional to its consequences.

14 CFR Ch. I (1–1–25 Edition) § 27.1316 (c) A means to alert the crew in the § 27.1317 High-intensity Radiated event of a failure must be provided Fields (HIRF) Protection.

when an unsafe system operating con- (a) Except as provided in paragraph dition exists and to enable them to (d) of this section, each electrical and take corrective action. Systems, con- electronic system that performs a func- trols, and associated monitoring and tion whose failure would prevent the crew alerting means must be designed continued safe flight and landing of the to minimize crew errors that could cre- rotorcraft must be designed and in- ate additional hazards. stalled so that— (1) The function is not adversely af- (d) Compliance with the require- fected during and after the time the ments of this section must be shown by rotorcraft is exposed to HIRF environ- analysis and, where necessary, by ment I, as described in appendix D to ground, flight, or simulator tests. The this part; analysis must account for: (2) The system automatically recov- (1) Possible modes of failure, includ- ers normal operation of that function, ing malfunctions and misleading data in a timely manner, after the rotor- and input from external sources; craft is exposed to HIRF environment (2) The effect of multiple failures and I, as described in appendix D to this latent failures; part, unless this conflicts with other (3) The resulting effects on the rotor- operational or functional requirements craft and occupants, considering the of that system; stage of flight and operating condi- (3) The system is not adversely af- tions; and fected during and after the time the (4) The crew alerting cues and the rotorcraft is exposed to HIRF environ- corrective action required. ment II, as described in appendix D to this part; and [Amdt. 27–51, 88 FR 8737, Feb. 10, 2023] (4) Each function required during op- eration under visual flight rules is not § 27.1316 Electrical and electronic sys- adversely affected during and after the tem lightning protection.

time the rotorcraft is exposed to HIRF (a) Each electrical and electronic environment III, as described in appen- system that performs a function, for dix D to this part.

which failure would prevent the contin- (b) Each electrical and electronic ued safe flight and landing of the rotor- system that performs a function whose craft, must be designed and installed so failure would significantly reduce the that— capability of the rotorcraft or the abil- (1) The function is not adversely af- ity of the flightcrew to respond to an fected during and after the time the adverse operating condition must be designed and installed so the system is rotorcraft is exposed to lightning; and not adversely affected when the equip- (2) The system automatically recov- ment providing these functions is ex- ers normal operation of that function posed to equipment HIRF test level 1 in a timely manner after the rotorcraft or 2, as described in appendix D to this is exposed to lightning.

part.

(b) For rotorcraft approved for in- (c) Each electrical and electronic sys- strument flight rules operation, each tem that performs a function whose electrical and electronic system that failure would reduce the capability of performs a function, for which failure the rotorcraft or the ability of the would reduce the capability of the flightcrew to respond to an adverse op- rotorcraft or the ability of the erating condition, must be designed flightcrew to respond to an adverse op- and installed so the system is not ad- erating condition, must be designed versely affected when the equipment and installed so that the function re- providing these functions is exposed to covers normal operation in a timely equipment HIRF test level 3, as de- manner after the rotorcraft is exposed scribed in appendix D to this part.

to lightning.

(d) Before December 1, 2012, an elec- trical or electronic system that per- [Doc. No. FAA–2010–0224, Amdt. 27–46, 76 FR 33135, June 8, 2011] forms a function whose failure would Federal Aviation Administration, DOT § 27.1325 prevent the continued safe flight and (d) Any other color, including white, landing of a rotorcraft may be designed for lights not described in paragraphs and installed without meeting the pro- (a) through (c) of this section, provided visions of paragraph (a) provided— the color differs sufficiently from the (1) The system has previously been colors prescribed in paragraphs (a) shown to comply with special condi- through (c) of this section to avoid pos- tions for HIRF, prescribed under § 21.16, sible confusion.

issued before December 1, 2007; [Amdt. 27–11, 41 FR 55470, Dec. 20, 1976] (2) The HIRF immunity characteris- tics of the system have not changed § 27.1323 Airspeed indicating system.

since compliance with the special con- (a) Each airspeed indicating instru- ditions was demonstrated; and ment must be calibrated to indicate (3) The data used to demonstrate true airspeed (at sea level with a stand- compliance with the special conditions ard atmosphere) with a minimum prac- is provided.

ticable instrument calibration error [Doc. No. FAA–2006–23657, 72 FR 44026, Aug. 6, when the corresponding pitot and stat- 2007] ic pressures are applied.

(b) The airspeed indicating system I NSTRUMENTS : I NSTALLATION must be calibrated in flight at forward speeds of 20 knots and over.

§ 27.1321 Arrangement and visibility.

(c) At each forward speed above 80 (a) Each flight, navigation, and pow- percent of the climbout speed, the air- erplant instrument for use by any pilot speed indicator must indicate true air- must be easily visible to him.

speed, at sea level with a standard at- (b) For each multiengine rotorcraft, mosphere, to within an allowable in- identical powerplant instruments must stallation error of not more than the be located so as to prevent confusion as greater of— to which engine each instrument re- (1) ± 3 percent of the calibrated air- lates.

speed; or (c) Instrument panel vibration may (2) Five knots.

not damage, or impair the readability or accuracy of, any instrument.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- (d) If a visual indicator is provided to eral Aviation Act of 1958 (49 U.S.C. 1354(a), indicate malfunction of an instrument, 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C.

it must be effective under all probable 1655(c))) cockpit lighting conditions.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (Secs. 313(a), 601, 603, 604, and 605 of the Fed- amended by Amdt. 27–13, 42 FR 36972, July 18, eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1977] 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C.

§ 27.1325 Static pressure systems.

1655(c))) (a) Each instrument with static air [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964; 29 case connections must be vented so FR 17885, Dec. 17, 1964, as amended by Amdt.

that the influence of rotorcraft speed, 27–13, 42 FR 36971, July 18, 1977] the opening and closing of windows, § 27.1322 Warning, caution, and advi- airflow variation, and moisture or sory lights.

other foreign matter does not seriously If warning, caution or advisory lights affect its accuracy.

are installed in the cockpit, they must, (b) Each static pressure port must be unless otherwise approved by the Ad- designed and located in such manner ministrator, be— that the correlation between air pres- (a) Red, for warning lights (lights in- sure in the static pressure system and dicating a hazard which may require true ambient atmospheric static pres- immediate corrective action): sure is not altered when the rotorcraft (b) Amber, for caution lights (lights encounters icing conditions. An anti- indicating the possible need for future icing means or an alternate source of corrective action); static pressure may be used in showing (c) Green, for safe operation lights; compliance with this requirement. If and the reading of the altimeter, when on 14 CFR Ch. I (1–1–25 Edition) § 27.1327 the alternate static pressure system, be placarded in accordance with differs from the reading of the altim- § 27.1547(e).

eter when on the primary static system (Secs. 313(a), 601, 603, 604, and 605 of the Fed- by more than 50 feet, a correction card eral Aviation Act of 1958 (49 U.S.C. 1354(a), must be provided for the alternate 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C.

static system.

1655(c))) (c) Except as provided in paragraph (d) of this section, if the static pressure [Amdt. 27–13, 42 FR 36972, July 18, 1977] system incorporates both a primary § 27.1329 Automatic pilot and flight and an alternate static pressure source, guidance system.

the means for selecting one or the For the purpose of this subpart, an other source must be designed so automatic pilot and flight guidance that— system may consist of an autopilot, (1) When either source is selected, the flight director, or a component that other is blocked off; and interacts with stability augmentation (2) Both sources cannot be blocked or trim.

off simultaneously.

(a) Each automatic pilot and flight (d) For unpressurized rotorcraft, guidance system must be designed so paragraph (c)(1) of this section does not that it: apply if it can be demonstrated that (1) Can be overpowered by one pilot the static pressure system calibration, to allow control of the rotorcraft; when either static pressure source is (2) Provides a means to disengage the selected is not changed by the other system, or any malfunctioning compo- static pressure source being open or nent of the system, by each pilot to blocked.

prevent it from interfering with the control of the rotorcraft; and (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (3) Provides a means to indicate to eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the the flight crew its current mode of op- Dept. of Transportation Act (49 U.S.C. eration. Selector switch position is not 1655(c))) acceptable as a means of indication.

(b) Unless there is automatic syn- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as chronization, each system must have a amended by Amdt. 27–13, 42 FR 36972, July 18, means to readily indicate to the pilot 1977] the alignment of the actuating device § 27.1327 Magnetic direction indicator. in relation to the control system it op- erates.

(a) Except as provided in paragraph (c) Each manually operated control (b) of this section— for the system’s operation must be (1) Each magnetic direction indicator readily accessible to the pilots.

must be installed so that its accuracy (d) The system must be designed so is not excessively affected by the that, within the range of adjustment rotorcraft’s vibration or magnetic available to the pilot, it cannot fields; and produce hazardous loads on the rotor- (2) The compensated installation may craft, or create hazardous deviations in not have a deviation, in level flight, the flight path, under any flight condi- greater than 10 degrees on any heading.

tion appropriate to its use or in the (b) A magnetic nonstabilized direc- event of a malfunction, assuming that tion indicator may deviate more than corrective action begins within a rea- 10 degrees due to the operation of elec- sonable period of time.

trically powered systems such as elec- (e) If the automatic pilot and flight trically heated windshields if either a guidance system integrates signals magnetic stabilized direction indi- from auxiliary controls or furnishes cator, which does not have a deviation signals for operation of other equip- in level flight greater than 10 degrees ment, there must be a means to pre- on any heading, or a gyroscopic direc- vent improper operation.

tion indicator, is installed. Deviations (f) If the automatic pilot system can of a magnetic nonstabilized direction be coupled to airborne navigation indicator of more than 10 degrees must equipment, means must be provided to Federal Aviation Administration, DOT § 27.1351 indicate to the pilots the current mode (e) Rotor drive system transmissions of operation. Selector switch position and gearboxes utilizing ferromagnetic is not acceptable as a means of indica- materials must be equipped with chip tion. detectors designed to indicate the pres- ence of ferromagnetic particles result- [Amdt. 27–21, 49 FR 44435, Nov. 6, 1984, as ing from damage or excessive wear.

amended by Amdt. 27–35, 63 FR 43285, Aug. 12, Chip detectors must— 1998; Amdt. 27–51, 88 FR 8738, Feb. 10, 2023] (1) Be designed to provide a signal to § 27.1337 Powerplant instruments. the device required by § 27.1305(v) and be provided with a means to allow (a) Instruments and instrument lines.

crewmembers to check, in flight, the (1) Each powerplant instrument line function of each detector electrical cir- must meet the requirements of §§ 27.- cuit and signal.

961 and 27.993.

(2) [Reserved] (2) Each line carrying flammable fluids under pressure must— (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 (i) Have restricting orifices or other U.S.C. 1354(a), 1421, and 1423; sec. 6(c) 49 safety devices at the source of pressure U.S.C. 1655(c)) to prevent the escape of excessive fluid [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as if the line fails; and amended by Amdt. 27–12, 42 FR 15046, Mar. 17, (ii) Be installed and located so that 1977; Amdt. 27–23, 53 FR 34214, Sept. 2, 1988; the escape of fluids would not create a Amdt. 27–37, 64 FR 45095, Aug. 18, 1999] hazard.

E LECTRICAL S YSTEMS AND E QUIPMENT (3) Each powerplant instrument that utilizes flammable fluids must be in- § 27.1351 General.

stalled and located so that the escape (a) Electrical system capacity. Elec- of fluid would not create a hazard.

trical equipment must be adequate for (b) Fuel quantity indicator. Each fuel its intended use. In addition— quantity indicator must be installed to (1) Electric power sources, their clearly indicate to the flight crew the transmission cables, and their associ- quantity of fuel in each tank in flight.

ated control and protective devices In addition— must be able to furnish the required (1) Each fuel quantity indicator must power at the proper voltage to each be calibrated to read ‘‘zero’’ during load circuit essential for safe oper- level flight when the quantity of fuel ation; and remaining in the tank is equal to the (2) Compliance with paragraph (a)(1) unusable fuel supply determined under of this section must be shown by an § 27.959; electrical load analysis, or by elec- (2) When two or more tanks are close- trical measurements that take into ac- ly interconnected by a gravity feed sys- count the electrical loads applied to tem and vented, and when it is impos- the electrical system, in probable com- sible to feed from each tank sepa- binations and for probable durations.

rately, at least one fuel quantity indi- (b) Function. For each electrical sys- cator must be installed; and tem, the following apply: (3) Each exposed sight gauge used as (1) Each system, when installed, a fuel quantity indicator must be pro- must be— tected against damage.

(i) Free from hazards in itself, in its (c) Fuel flowmeter system. If a fuel method of operation, and in its effects flowmeter system is installed, each on other parts of the rotorcraft; and metering component must have a (ii) Protected from fuel, oil, water, means for bypassing the fuel supply if other detrimental substances, and me- malfunction of that component se- chanical damage.

verely restricts fuel flow.

(d) Oil quantity indicator. There must (2) Electric power sources must func- be means to indicate the quantity of tion properly when connected in com- oil in each tank— bination or independently.

(1) On the ground (including during (3) No failure or malfunction of any the filling of each tank); and source may impair the ability of any (2) In flight, if there is an oil transfer remaining source to supply load cir- system or reserve oil supply system. cuits essential for safe operation.

14 CFR Ch. I (1–1–25 Edition) § 27.1353 (4) Each electric power source control (a) Energy storage systems must pro- must allow the independent operation vide automatic protective features for of each source. any conditions that could prevent con- tinued safe flight and landing.

(c) Generating system. There must be (b) Energy storage systems must not at least one generator if the system emit any flammable, explosive, or supplies power to load circuits essen- toxic gases, smoke, or fluids that could tial for safe operation. In addition— accumulate in hazardous quantities (1) Each generator must be able to within the rotorcraft.

deliver its continuous rated power; (c) Corrosive fluids or gases that es- (2) Generator voltage control equip- cape from the system must not damage ment must be able to dependably regu- surrounding structures, adjacent equip- late each generator output within ment, or systems necessary for contin- rated limits; ued safe flight and landing.

(3) Each generator must have a re- (d) The maximum amount of heat verse current cutout designed to dis- and pressure that can be generated dur- connect the generator from the battery ing any operation or under any failure and from the other generators when condition of the energy storage system enough reverse current exists to dam- or its individual components must not age that generator; and result in any hazardous effect on rotor- (4) Each generator must have an craft structure, equipment, or systems overvoltage control designed and in- necessary for continued safe flight and stalled to prevent damage to the elec- landing.

trical system, or to equipment supplied (e) Energy storage system installa- by the electrical system, that could re- tions required for continued safe flight sult if that generator were to develop and landing of the rotorcraft must an overvoltage condition.

have monitoring features and a means (d) Instruments. There must be means to indicate to the pilot the status of all to indicate to appropriate crew- critical system parameters.

members the electric power system [Amdt. 27–51, 88 FR 8738, Feb. 10, 2023] quantities essential for safe operation of the system. In addition— § 27.1357 Circuit protective devices.

(1) For direct current systems, an (a) Protective devices, such as fuses ammeter that can be switched into or circuit breakers, must be installed each generator feeder may be used; and in each electrical circuit other than— (2) If there is only one generator, the (1) The main circuits of starter mo- ammeter may be in the battery feeder.

tors; and (e) External power. If provisions are (2) Circuits in which no hazard is pre- made for connecting external power to sented by their omission.

the rotorcraft, and that external power (b) A protective device for a circuit can be electrically connected to equip- essential to flight safety may not be ment other than that used for engine used to protect any other circuit.

starting, means must be provided to (c) Each resettable circuit protective ensure that no external power supply device (‘‘trip free’’ device in which the having a reverse polarity, or a reverse tripping mechanism cannot be over- phase sequence, can supply power to ridden by the operating control) must the rotorcraft’s electrical system.

be designed so that— (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (1) A manual operation is required to eral Aviation Act of 1958 (49 U.S.C. 1354(a), restore service after trippling; and 1421, 1423, 1424, and 1425); and sec. 6(c) of the (2) If an overload or circuit fault ex- Dept. of Transportation Act (49 U.S.C.

ists, the device will open the circuit re- 1655(c))) gardless of the position of the oper- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as ating control.

amended by Amdt. 27–11, 41 FR 55470, Dec. 20, (d) If the ability to reset a circuit 1976; Amdt. 27–13, 42 FR 36972, July 18, 1977] breaker or replace a fuse is essential to safety in flight, that circuit breaker or § 27.1353 Energy storage systems.

fuse must be located and identified so Energy storage systems must be de- that it can be readily reset or replaced signed and installed as follows: in flight.

Federal Aviation Administration, DOT § 27.1385 (e) If fuses are used, there must be (a) Make each instrument, switch, one spare of each rating, or 50 percent and other devices for which they are spare fuses of each rating, whichever is provided easily readable; and greater. (b) Be installed so that— (1) Their direct rays are shielded (Secs. 313(a), 601, 603, 604, and 605 of the Fed- from the pilot’s eyes; and eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the (2) No objectionable reflections are Dept. of Transportation Act (49 U.S.C.

visible to the pilot.

1655(c))) § 27.1383 Landing lights.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964; 29 FR 17885, Dec. 17, 1964, as amended by Amdt.

(a) Each required landing or hovering 27–13, 42 FR 36972, July 18, 1977] light must be approved.

(b) Each landing light must be in- § 27.1361 Master switch.

stalled so that— (a) There must be a master switch ar- (1) No objectionable glare is visible rangement to allow ready disconnec- to the pilot; tion of each electric power source from (2) The pilot is not adversely affected the main bus. The point of disconnec- by halation; and tion must be adjacent to the sources (3) It provides enough light for night controlled by the switch.

operation, including hovering and land- (b) Load circuits may be connected so ing.

that they remain energized after the (c) At least one separate switch must switch is opened, if they are protected be provided, as applicable— by circuit protective devices, rated at (1) For each separately installed five amperes or less, adjacent to the landing light; and electric power source.

(2) For each group of landing lights (c) The master switch or its controls installed at a common location.

must be installed so that the switch is easily discernible and accessible to a § 27.1385 Position light system installa- crewmember in flight.

tion.

§ 27.1365 Electric cables.

(a) General. Each part of each posi- tion light system must meet the appli- (a) Each electric connecting cable cable requirements of this section, and must be of adequate capacity.

each system as a whole must meet the (b) Each cable that would overheat in requirements of §§ 27.1387 through the event of circuit overload or fault 27.1397.

must be at least flame resistant and (b) Forward position lights. Forward may not emit dangerous quantities of position lights must consist of a red toxic fumes.

and a green light spaced laterally as (c) Insulation on electrical wire and far apart as practicable and installed cable installed in the rotorcraft must forward on the rotorcraft so that, with be self-extinguishing when tested in ac- the rotorcraft in the normal flying po- cordance with appendix F, part I(a)(3), sition, the red light is on the left side of part 25 of this chapter.

and the green light is on the right side.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as Each light must be approved.

amended by Amdt. 27–35, 63 FR 43285, Aug. 12, (c) Rear position light. The rear posi- 1998] tion light must be a white light mount- § 27.1367 Switches. ed as far aft as practicable, and must be approved.

Each switch must be— (d) Circuit. The two forward position (a) Able to carry its rated current; lights and the rear position light must (b) Accessible to the crew; and make a single circuit.

(c) Labeled as to operation and the (e) Light covers and color filters. Each circuit controlled.

light cover or color filter must be at L IGHTS least flame resistant and may not change color or shape or lose any ap- § 27.1381 Instrument lights.

preciable light transmission during The instrument lights must— normal use.

14 CFR Ch. I (1–1–25 Edition) § 27.1387 intensity of each position light must § 27.1387 Position light system dihe- dral angles. meet the requirements of paragraph (b) of this section.

(a) Except as provided in paragraph (b) Forward and rear position lights.

(e) of this section, each forward and The light distribution and intensities rear position light must, as installed, of forward and rear position lights show unbroken light within the dihe- dral angles described in this section. must be expressed in terms of min- (b) Dihedral angle L (left) is formed imum intensities in the horizontal by two intersecting vertical planes, the plane, minimum intensities in any first parallel to the longitudinal axis of vertical plane, and maximum inten- the rotorcraft, and the other at 110 de- sities in overlapping beams, within di- grees to the left of the first, as viewed hedral angles L, R, and A, and must when looking forward along the longi- meet the following requirements: tudinal axis.

(1) Intensities in the horizontal plane.

(c) Dihedral angle R (right) is formed Each intensity in the horizontal plane by two intersecting vertical planes, the (the plane containing the longitudinal first parallel to the longitudinal axis of axis of the rotorcraft and perpendicular the rotorcraft, and the other at 110 de- to the plane of symmetry of the rotor- grees to the right of the first, as viewed craft) must equal or exceed the values when looking forward along the longi- in § 27.1391.

tudinal axis.

(2) Intensities in any vertical plane.

(d) Dihedral angle A (aft) is formed Each intensity in any vertical plane by two intersecting vertical planes (the plane perpendicular to the hori- making angles of 70 degrees to the zontal plane) must equal or exceed the right and to the left, respectively, to a appropriate value in § 27.1393, where I is vertical plane passing through the lon- the minimum intensity prescribed in gitudinal axis, as viewed when looking § 27.1391 for the corresponding angles in aft along the longitudinal axis.

the horizontal plane.

(e) If the rear position light, when mounted as far aft as practicable in ac- (3) Intensities in overlaps between adja- cordance with § 25.1385(c), cannot show cent signals. No intensity in any over- unbroken light within dihedral angle A lap between adjacent signals may ex- (as defined in paragraph (d) of this sec- ceed the values in § 27.1395, except that tion), a solid angle or angles of ob- higher intensities in overlaps may be structed visibility totaling not more used with main beam intensities sub- than 0.04 steradians is allowable within stantially greater than the minima that dihedral angle, if such solid angle specified in §§ 27.1391 and 27.1393, if the is within a cone whose apex is at the overlap intensities in relation to the rear position light and whose elements main beam intensities do not adversely make an angle of 30 ° with a vertical affect signal clarity. When the peak in- line passing through the rear position tensity of the forward position lights is light.

greater than 100 candles, the maximum (49 U.S.C. 1655(c)) overlap intensities between them may exceed the values in § 27.1395 if the [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–7, 36 FR 21278, Nov. 5, overlap intensity in Area A is not more 1971] than 10 percent of peak position light intensity and the overlap intensity in § 27.1389 Position light distribution Area B is not more than 2.5 percent of and intensities.

peak position light intensity.

(a) General. the intensities prescribed in this section must be provided by new § 27.1391 Minimum intensities in the equipment with light covers and color horizontal plane of forward and filters in place. Intensities must be de- rear position lights.

termined with the light source oper- Each position light intensity must ating at a steady value equal to the av- equal or exceed the applicable values in erage luminous output of the source at the following table: the normal operating voltage of the rotorcraft. The light distribution and Federal Aviation Administration, DOT § 27.1401 Angle from right or left (a) Aviation red — Dihedral angle (light in- of longitudinal axis, Intensity cluded) measured from dead (candles) y is not greater than 0.335; and ahead z is not greater than 0.002.

(b) Aviation green — x is not greater than 0.440 ¥ 0.320 y ; x is not greater than y ¥ 0.170; and y is not less than 0.390 ¥ 0.170 x.

§ 27.1393 Minimum intensities in any (c) Aviation white — vertical plane of forward and rear position lights.

x is not less than 0.300 and not greater than 0.540; Each position light intensity must y is not less than x ¥ 0.040’’ or y c ¥ 0.010, equal or exceed the applicable values in whichever is the smaller; and the following table: y is not greater than x + 0.020 nor 0.636 ¥ 0.400 x ; Angle above or below the horizontal plane Intensity, l Where y is the y coordinate of the Planck- c 0 ° ......................................................................... 1.00 ian radiator for the value of x considered.

0 ° to 5 ° ................................................................ 0.90 5 ° to 10 ° .............................................................. 0.80 [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as 10 ° to 15 ° ............................................................ 0.70 amended by Amdt. 27–6, 36 FR 12972, July 10, 15 ° to 20 ° ............................................................ 0.50 1971] 20 ° to 30 ° ............................................................ 0.30 30 ° to 40 ° ............................................................ 0.10 § 27.1399 Riding light.

40 ° to 90 ° ............................................................ 0.05 (a) Each riding light required for water operation must be installed so § 27.1395 Maximum intensities in over- that it can— lapping beams of forward and rear position lights. (1) Show a white light for at least two nautical miles at night under clear No position light intensity may ex- atmospheric conditions; and ceed the applicable values in the fol- (2) Show a maximum practicable un- lowing table, except as provided in broken light with the rotorcraft on the § 27.1389(b)(3).

water.

Maximum Intensity (b) Externally hung lights may be Overlaps used.

Area A Area B (candles) (candles) [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as Green in dihedral angle L ............. 10 1 amended by Amdt. 27–2, 33 FR 964, Jan. 26, Red in dihedral angle R ................ 10 1 1968] Green in dihedral angle A ............. 5 1 Red in dihedral angle A ................ 5 1 § 27.1401 Anticollision light system.

Rear white in dihedral angle L ...... 5 1 Rear white in dihedral angle R ..... 5 1 (a) General. If certification for night operation is requested, the rotorcraft Where— must have an anticollision light sys- (a) Area A includes all directions in tem that— the adjacent dihedral angle that pass (1) Consists of one or more approved through the light source and intersect anticollision lights located so that the common boundary plane at more their emitted light will not impair the than 10 degrees but less than 20 de- crew’s vision or detract from the con- grees, and spicuity of the position lights; and (b) Area B includes all directions in (2) Meets the requirements of para- the adjacent dihedral angle that pass graphs (b) through (f) of this section.

through the light source and intersect (b) Field of coverage. The system must the common boundary plane at more consist of enough lights to illuminate than 20 degrees.

the vital areas around the rotorcraft, considering the physical configuration § 27.1397 Color specifications.

and flight characteristics of the rotor- Each position light color must have craft. The field of coverage must ex- the applicable International Commis- tend in each direction within at least sion on Illumination chromaticity co- 30 degrees below the horizontal plane of ordinates as follows: the rotorcraft, except that there may 14 CFR Ch. I (1–1–25 Edition) § 27.1411 be solid angles of obstructed visibility S AFETY E QUIPMENT totaling not more than 0.5 steradians.

§ 27.1411 General.

(c) Flashing characteristics. The ar- rangement of the system, that is, the (a) Required safety equipment to be number of light sources, beam width, used by the crew in an emergency, such speed of rotation, and other character- as flares and automatic liferaft re- istics, must give an effective flash fre- leases, must be readily accessible.

quency of not less than 40, nor more (b) Stowage provisions for required than 100, cycles per minute. The effec- safety equipment must be furnished tive flash frequency is the frequency at and must— which the rotorcraft’s complete anti- (1) Be arranged so that the equip- collision light system is observed from ment is directly accessible and its loca- a distance, and applies to each sector tion is obvious; and of light including any overlaps that (2) Protect the safety equipment exist when the system consists of more from damage caused by being subjected than one light source. In overlaps, to the inertia loads specified in § 27.561.

flash frequencies may exceed 100, but not 180, cycles per minute. [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–11, 41 FR 55470, Dec. 20, (d) Color. Each anticollision light 1976] must be aviation red and must meet the applicable requirements of § 27.1397.

§ 27.1413 Safety belts.

(e) Light intensity. The minimum Each safety belt must be equipped light intensities in any vertical plane, with a metal to metal latching device.

measured with the red filter (if used) and expressed in terms of ‘‘effective’’ (Secs. 313, 314, and 601 through 610 of the Fed- intensities, must meet the require- eral Aviation Act of 1958 (49 U.S.C. 1354, 1355, ments of paragraph (f) of this section.

and 1421 through 1430) and sec. 6(c), Dept. of The following relation must be as- Transportation Act (49 U.S.C. 1655(c))) sumed: [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–15, 43 FR 46233, Oct. 5, t 1978; Amdt. 27–21, 49 FR 44435, Nov. 6, 1984] I t dt ( )

∫

t I = e § 27.1415 Ditching equipment.

t t + − . 0 2

( )

2 1 (a) Emergency flotation and sig- where: naling equipment required by any oper- I = effective intensity (candles).

e ating rule in this chapter must meet I(t) = instantaneous intensity as a function the requirements of this section.

of time.

(b) Each raft and each life preserver t ¥ t = flash time interval (seconds).

2 1 must be approved and must be installed Normally, the maximum value of effective so that it is readily available to the intensity is obtained when t 2 and t 1 are cho- crew and passengers. The storage pro- sen so that the effective intensity is equal to visions for life preservers must accom- the instantaneous intensity at t 2 and t 1 .

modate one life preserver for each oc- (f) Minimum effective intensities for cupant for which certification for anticollision light. Each anticollision ditching is requested.

light effective intensity must equal or (c) Each raft released automatically exceed the applicable values in the fol- or by the pilot must be attached to the lowing table: rotorcraft by a line to keep it alongside Effective the rotorcraft. This line must be weak Angle above or below the horizontal plane intensity enough to break before submerging the (candles) empty raft to which it is attached.

(d) Each signaling device must be free from hazard in its operation and must be installed in an accessible loca- tion.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–6, 36 FR 12972, July 10, amended by Amdt. 27–11, 41 FR 55470, Dec. 20, 1971; Amdt. 27–10, 41 FR 5290, Feb. 5, 1976] 1976] Federal Aviation Administration, DOT § 27.1457 out yielding, any structural loads ex- § 27.1419 Ice protection.

pected in addition to hydraulic loads.

(a) To obtain certification for flight (b) Tests. Each system must be sub- into icing conditions, compliance with stantiated by proof pressure tests.

this section must be shown.

When proof tested, no part of any sys- (b) It must be demonstrated that the tem may fail, malfunction, or experi- rotorcraft can be safely operated in the ence a permanent set. The proof load of continuous maximum and intermittent each system must be at least 1.5 times maximum icing conditions determined the maximum operating pressure of under appendix C of Part 29 of this that system.

chapter within the rotorcraft altitude (c) Accumulators. No hydraulic accu- envelope. An analysis must be per- mulator or pressurized reservoir may formed to establish, on the basis of the be installed on the engine side of any rotorcraft’s operational needs, the ade- firewall unless it is an integral part of quacy of the ice protection system for an engine.

the various components of the rotor- craft.

§ 27.1457 Cockpit voice recorders.

(c) In addition to the analysis and physical evaluation prescribed in para- (a) Each cockpit voice recorder re- graph (b) of this section, the effective- quired by the operating rules of this ness of the ice protection system and chapter must be approved, and must be its components must be shown by installed so that it will record the fol- flight tests of the rotorcraft or its com- lowing: ponents in measured natural atmos- (1) Voice communications trans- pheric icing conditions and by one or mitted from or received in the rotor- more of the following tests as found craft by radio.

necessary to determine the adequacy of (2) Voice communications of flight the ice protection system: crewmembers on the flight deck.

(1) Laboratory dry air or simulated (3) Voice communications of flight icing tests, or a combination of both, of crewmembers on the flight deck, using the components or models of the com- the rotorcraft’s interphone system.

ponents.

(4) Voice or audio signals identifying (2) Flight dry air tests of the ice pro- navigation or approach aids introduced tection system as a whole, or its indi- into a headset or speaker.

vidual components.

(5) Voice communications of flight (3) Flight tests of the rotorcraft or crewmembers using the passenger loud- its components in measured simulated speaker system, if there is such a sys- icing conditions.

tem, and if the fourth channel is avail- (d) The ice protection provisions of able in accordance with the require- this section are considered to be appli- ments of paragraph (c)(4)(ii) of this sec- cable primarily to the airframe. Power- tion.

plant installation requirements are (6) If datalink communication equip- contained in Subpart E of this part.

ment is installed, all datalink commu- (e) A means must be indentified or nications, using an approved data mes- provided for determining the formation sage set. Datalink messages must be of ice on critical parts of the rotor- recorded as the output signal from the craft. Unless otherwise restricted, the communications unit that translates means must be available for nighttime the signal into usable data.

as well as daytime operation. The (b) The recording requirements of rotorcraft flight manual must describe paragraph (a)(2) of this section may be the means of determining ice forma- met: tion and must contain information nec- (1) By installing a cockpit-mounted essary for safe operation of the rotor- area microphone located in the best po- craft in icing conditions.

sition for recording voice communica- [Amdt. 27–19, 48 FR 4389, Jan. 31, 1983] tions originating at the first and sec- ond pilot stations and voice commu- § 27.1435 Hydraulic systems.

nications of other crewmembers on the (a) Design. Each hydraulic system flight deck when directed to those sta- and its elements must withstand, with- tions; or 14 CFR Ch. I (1–1–25 Edition) § 27.1457 (2) By installing a continually ener- (ii) It remains powered for as long as gized or voice-actuated lip microphone possible without jeopardizing emer- at the first and second pilot stations. gency operation of the rotorcraft.

The microphone specified in this (2) There is an automatic means to paragraph must be so located and, if simultaneously stop the recorder and necessary, the preamplifiers and filters prevent each erasure feature from func- of the recorder must be adjusted or tioning, within 10 minutes after crash supplemented so that the recorded impact; communications are intelligible when (3) There is an aural or visual means recorded under flight cockpit noise for preflight checking of the recorder conditions and played back. The level for proper operation; of intelligibility must be approved by (4) Whether the cockpit voice re- the Administrator. Repeated aural or corder and digital flight data recorder visual playback of the record may be are installed in separate boxes or in a used in evaluating intelligibility.

combination unit, no single electrical (c) Each cockpit voice recorder must failure external to the recorder may be installed so that the part of the disable both the cockpit voice recorder communication or audio signals speci- and the digital flight data recorder; fied in paragraph (a) of this section ob- and tained from each of the following (5) It has an independent power sources is recorded on a separate chan- source— nel: (i) That provides 10 ± 1 minutes of (1) For the first channel, from each electrical power to operate both the microphone, headset, or speaker used cockpit voice recorder and cockpit- at the first pilot station.

mounted area microphone; (2) For the second channel, from each (ii) That is located as close as prac- microphone, headset, or speaker used ticable to the cockpit voice recorder; at the second pilot station.

and (3) For the third channel, from the (iii) To which the cockpit voice re- cockpit-mounted area microphone, or corder and cockpit-mounted area the continually energized or voice-ac- microphone are switched automati- tuated lip microphone at the first and cally in the event that all other power second pilot stations.

to the cockpit voice recorder is inter- (4) For the fourth channel, from: rupted either by normal shutdown or (i) Each microphone, headset, or by any other loss of power to the elec- speaker used at the stations for the trical power bus.

third and fourth crewmembers; or (e) The record container must be lo- (ii) If the stations specified in para- cated and mounted to minimize the graph (c)(4)(i) of this section are not re- probability of rupture of the container quired or if the signal at such a station as a result of crash impact and con- is picked up by another channel, each sequent heat damage to the record microphone on the flight deck that is from fire.

used with the passenger loudspeaker (f) If the cockpit voice recorder has a system if its signals are not picked up bulk erasure device, the installation by another channel.

must be designed to minimize the prob- (iii) Each microphone on the flight ability of inadvertent operation and ac- deck that is used with the rotorcraft’s tuation of the device during crash im- loudspeaker system if its signals are pact.

not picked up by another channel.

(g) Each recorder container must be (d) Each cockpit voice recorder must either bright orange or bright yellow.

be installed so that: (h) When both a cockpit voice re- (1)(i) It receives its electrical power corder and a flight data recorder are from the bus that provides the max- required by the operating rules, one imum reliability for operation of the combination unit may be installed, cockpit voice recorder without jeopard- provided that all other requirements of izing service to essential or emergency this section and the requirements for loads.

Federal Aviation Administration, DOT § 27.1461 flight data recorders under this part lot’s instruments. This correlation are met. must cover the airspeed range over which the aircraft is to be operated, [Amdt. 27–22, 53 FR 26144, July 11, 1988, as the range of altitude to which the air- amended by Amdt. 27–43, 73 FR 12563, Mar. 7, craft is limited, and 360 degrees of 2008; 74 FR 32800, July 9, 2009; Amdt. 27–45, 75 heading. Correlation may be estab- FR 17045, Apr. 5, 2010] lished on the ground as appropriate.

§ 27.1459 Flight data recorders.

(d) Each recorder container must: (1) Be either bright orange or bright (a) Each flight recorder required by yellow; the operating rules of Subchapter G of (2) Have a reflective tape affixed to this chapter must be installed so that: its external surface to facilitate its lo- (1) It is supplied with airspeed, alti- cation under water; and tude, and directional data obtained (3) Have an underwater locating de- from sources that meet the accuracy vice, when required by the operating requirements of §§ 27.1323, 27.1325, and rules of this chapter, on or adjacent to 27.1327 of this part, as applicable; the container which is secured in such (2) The vertical acceleration sensor is a manner that they are not likely to be rigidly attached, and located longitu- separated during crash impact.

dinally within the approved center of (e) When both a cockpit voice re- gravity limits of the rotorcraft; corder and a flight data recorder are (3)(i) It receives its electrical power required by the operating rules, one from the bus that provides the max- combination unit may be installed, imum reliability for operation of the provided that all other requirements of flight data recorder without jeopard- this section and the requirements for izing service to essential or emergency cockpit voice recorders under this part loads.

are met.

(ii) It remains powered for as long as possible without jeopardizing emer- [Amdt. 27–22, 53 FR 26144, July 11, 1988, as gency operation of the rotorcraft.

amended by Amdt. 27–43, 73 FR 12564, Mar. 7, (4) There is an aural or visual means 2008; 74 FR 32800, July 9, 2009; Amdt. 27–45, 75 for preflight checking of the recorder FR 17045, Apr. 5, 2010] for proper recording of data in the stor- § 27.1461 Equipment containing high age medium; energy rotors.

(5) Except for recorders powered sole- ly by the engine-driven electrical gen- (a) Equipment containing high en- erator system, there is an automatic ergy rotors must meet paragraph (b), means to simultaneously stop a re- (c), or (d) of this section.

corder that has a data erasure feature (b) High energy rotors contained in and prevent each erasure feature from equipment must be able to withstand functioning, within 10 minutes after damage caused by malfunctions, vibra- any crash impact; and tion, abnormal speeds, and abnormal (6) Whether the cockpit voice re- temperatures. In addition— corder and digital flight data recorder (1) Auxiliary rotor cases must be able are installed in separate boxes or in a to contain damage caused by the fail- combination unit, no single electrical ure of high energy rotor blades; and failure external to the recorder may (2) Equipment control devices, sys- disable both the cockpit voice recorder tems, and instrumentation must rea- and the digital flight data recorder. sonably ensure that no operating limi- (b) Each nonejectable recorder con- tations affecting the integrity of high tainer must be located and mounted so energy rotors will be exceeded in serv- as to minimize the probability of con- ice.

tainer rupture resulting from crash im- (c) It must be shown by test that pact and subsequent damage to the equipment containing high energy ro- record from fire. tors can contain any failure of a high (c) A correlation must be established energy rotor that occurs at the highest between the flight recorder readings of speed obtainable with the normal speed airspeed, altitude, and heading and the control devices inoperative.

corresponding readings (taking into ac- (d) Equipment containing high en- count correction factors) of the first pi- ergy rotors must be located where 14 CFR Ch. I (1–1–25 Edition) § 27.1501 rotor failure will neither endanger the (2) The ranges of these variables (or occupants nor adversely affect contin- of the indications on instruments inte- ued safe flight. grating more than one of these vari- ables) are large enough to allow an [Amdt. 27–2, 33 FR 964, Jan. 26, 1968] operationally practical and safe vari- ation of V .

NE Subpart G—Operating Limitations (c) For helicopters, a stabilized and Information power-off V denoted as V (power- NE NE off) may be established at a speed less § 27.1501 General.

than V established pursuant to para- NE (a) Each operating limitation speci- graph (a) of this section, if the fol- fied in §§ 27.1503 through 27.1525 and lowing conditions are met: other limitations and information nec- (1) V (power-off) is not less than a NE essary for safe operation must be es- speed midway between the power-on tablished.

V and the speed used in meeting the NE (b) The operating limitations and requirements of— other information necessary for safe (i) § 27.65(b) for single engine heli- operation must be made available to copters; and the crewmembers as prescribed in (ii) § 27.67 for multiengine heli- §§ 27.1541 through 27.1589.

copters.

(2) V (power-off) is— NE (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (i) A constant airspeed; eral Aviation Act of 1958 (49 U.S.C. 1354(a), (ii) A constant amount less than 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. power-on V ; or NE 1655(c))) (iii) A constant airspeed for a portion of the altitude range for which certifi- [Amdt. 27–14, 43 FR 2325, Jan. 16, 1978] cation is requested, and a constant O PERATING L IMITATIONS amount less than power-on V for the NE remainder of the altitude range.

§ 27.1503 Airspeed limitations: general.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- (a) An operating speed range must be eral Aviation Act of 1958 (49 U.S.C. 1354(a), established.

1421, 1423, 1424, and 1425); and sec. 6(c) of the (b) When airspeed limitations are a Dept. of Transportation Act (49 U.S.C.

1655(c))) function of weight, weight distribution, altitude, rotor speed, power, or other [Amdt. 27–2, 33 FR 964, Jan. 26, 1968, and factors, airspeed limitations cor- Amdt. 27–14, 43 FR 2325, Jan. 16, 1978; Amdt.

responding with the critical combina- 27–21, 49 FR 44435, Nov. 6, 1984] tions of these factors must be estab- § 27.1509 Rotor speed.

lished.

(a) Maximum power-off (autorotation).

§ 27.1505 Never-exceed speed.

The maximum power-off rotor speed (a) The never-exceed speed, V must must be established so that it does not NE, be established so that it is— exceed 95 percent of the lesser of— (1) Not less than 40 knots (CAS); and (1) The maximum design r.p.m. deter- (2) Not more than the lesser of— mined under § 27.309(b); and (i) 0.9 times the maximum forward (2) The maximum r.p.m. shown dur- speeds established under § 27.309; ing the type tests.

(ii) 0.9 times the maximum speed (b) Minimum power off. The minimum shown under §§ 27.251 and 27.629; or power-off rotor speed must be estab- (iii) 0.9 times the maximum speed lished so that it is not less than 105 substantiated for advancing blade tip percent of the greater of— mach number effects. (1) The minimum shown during the (b) V may vary with altitude, type tests; and NE r.p.m., temperature, and weight, if— (2) The minimum determined by de- (1) No more than two of these vari- sign substantiation.

ables (or no more than two instru- (c) Minimum power on. The minimum ments integrating more than one of power-on rotor speed must be estab- these variables) are used at one time; lished so that it is— and (1) Not less than the greater of— Federal Aviation Administration, DOT § 27.1521 (i) The minimum shown during the ating and atmospheric conditions for type tests; and which certification is requested.

(ii) The minimum determined by de- (d) Fuel grade or designation. The min- sign substantiation; and imum fuel grade (for reciprocating en- gines), or fuel designation (for turbine (2) Not more than a value determined engines), must be established so that it under § 27.33(a)(1) and (b)(1).

is not less than that required for the § 27.1519 Weight and center of gravity.

operation of the engines within the limitations in paragraphs (b) and (c) of The weight and center of gravity lim- this section.

itations determined under §§ 27.25 and (e) Turboshaft engine torque. For 27.27, respectively, must be established rotorcraft with main rotors driven by as operating limitations.

turboshaft engines, and that do not [Amdt. 27–2, 33 FR 965, Jan. 26, 1968, as have a torque limiting device in the amended by Amdt. 27–21, 49 FR 44435, Nov. 6, transmission system, the following 1984] apply: (1) A limit engine torque must be es- § 27.1521 Powerplant limitations.

tablished if the maximum torque that (a) General. The powerplant limita- the engine can exert is greater than— tions prescribed in this section must be (i) The torque that the rotor drive established so that they do not exceed system is designed to transmit; or the corresponding limits for which the (ii) The torque that the main rotor engines are type certificated.

assembly is designed to withstand in (b) Takeoff operation. The powerplant showing compliance with § 27.547(e).

takeoff operation must be limited by— (2) The limit engine torque estab- (1) The maximum rotational speed, lished under paragraph (e)(1) of this which may not be greater than— section may not exceed either torque (i) The maximum value determined specified in paragraph (e)(1)(i) or (ii) of by the rotor design; or this section.

(ii) The maximum value shown dur- (f) Ambient temperature. For turbine ing the type tests; engines, ambient temperature limita- (2) The maximum allowable manifold tions (including limitations for winter- pressure (for reciprocating engines); ization installations, if applicable) (3) The time limit for the use of the must be established as the maximum power corresponding to the limitations ambient atmospheric temperature at established in paragraphs (b)(1) and (2) which compliance with the cooling pro- of this section; visions of §§ 27.1041 through 27.1045 is (4) If the time limit in paragraph shown.

(b)(3) of this section exceeds two min- (g) Two and one-half-minute OEI power utes, the maximum allowable cylinder operation. Unless otherwise authorized, head, coolant outlet, or oil tempera- the use of 2 ⁄2 -minute OEI power must tures; be limited to engine failure operation (5) The gas temperature limits for of multiengine, turbine-powered rotor- turbine engines over the range of oper- craft for not longer than 2 ⁄2 minutes ating and atmospheric conditions for after failure of an engine. The use of which certification is requested.

2 ⁄ 2 -minute OEI power must also be lim- (c) Continuous operation. The contin- ited by— uous operation must be limited by— (1) The maximum rotational speed, (1) The maximum rotational speed which may not be greater than— which may not be greater than— (i) The maximum value determined (i) The maximum value determined by the rotor design; or by the rotor design; or (ii) The maximum demonstrated dur- (ii) The maximum value shown dur- ing the type tests; ing the type tests; (2) The maximum allowable gas tem- (2) The minimum rotational speed perature; and shown under the rotor speed require- (3) The maximum allowable torque.

ments in § 27.1509(c); and (h) Thirty-minute OEI power operation.

(3) The gas temperature limits for Unless otherwise authorized, the use of turbine engines over the range of oper- 30-minute OEI power must be limited 14 CFR Ch. I (1–1–25 Edition) § 27.1523 to multiengine, turbine-powered rotor- (k) Rated 2-minute OEI power oper- craft for not longer than 30 minutes ation. Rated 2-minute OEI power is per- after failure of an engine. The use of 30- mitted only on multiengine, turbine- minute OEI power must also be limited powered rotorcraft, also certificated by— for the use of rated 30-second OEI (1) The maximum rotational speed, power, and can only be used for contin- which may not be greater than— ued operation of the remaining en- (i) The maximum value determined gine(s) after a failure or precautionary by the rotor design; or shutdown of an engine. It must be (ii) The maximum value dem- shown that following application of 2- onstrated during the type tests; minute OEI power, any damage will be (2) The maximum allowable gas tem- readily detectable by the applicable in- perature; and spections and other related procedures (3) The maximum allowable torque.

furnished in accordance with Section (i) Continuous OEI power operation.

A27.4 of appendix A of this part and Unless otherwise authorized, the use of Section A33.4 of appendix A of part 33.

continuous OEI power must be limited The use of 2-minute OEI power must be to multiengine, turbine-powered rotor- limited to not more than 2 minutes for craft for continued flight after failure any period in which that power is used, of an engine. The use of continuous and by— OEI power must also be limited by— (1) The maximum rotational speed, (1) The maximum rotational speed, which may not be greater than— which may not be greater than— (i) The maximum value determined (i) The maximum value determined by the rotor design; or by the rotor design; or (ii) The maximum value dem- (ii) The maximum value dem- onstrated during the type tests; onstrated during the type tests; (2) The maximum allowable gas tem- (2) The maximum allowable gas tem- perature; and perature; and (3) The maximum allowable torque.

(3) The maximum allowable torque.

(j) Rated 30-second OEI power oper- ation. Rated 30-second OEI power is (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), permitted only on multiengine, tur- 1421, 1423, 1424, and 1425); and sec. 6(c) of the bine-powered rotorcraft, also certifi- Dept. of Transportation Act (49 U.S.C.

cated for the use of rated 2-minute OEI 1655(c))) power, and can only be used for contin- ued operation of the remaining en- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as gine(s) after a failure or precautionary amended by Amdt. 27–14, 43 FR 2325, Jan. 16, 1978; Amdt. 27–23, 53 FR 34214, Sept. 2, 1988; shutdown of an engine. It must be Amdt. 27–29, 59 FR 47767, Sept. 16, 1994] shown that following application of 30- second OEI power, any damage will be § 27.1523 Minimum flight crew.

readily detectable by the applicable in- spections and other related procedures The minimum flight crew must be es- furnished in accordance with Section tablished so that it is sufficient for safe A27.4 of appendix A of this part and operation, considering— Section A33.4 of appendix A of part 33.

(a) The workload on individual crew- The use of 30-second OEI power must be members; limited to not more than 30 seconds for (b) The accessibility and ease of oper- any period in which that power is used, ation of necessary controls by the ap- and by— propriate crewmember; and (1) The maximum rotational speed, (c) The kinds of operation authorized which may not be greater than— under § 27.1525.

(i) The maximum value determined by the rotor design; or § 27.1525 Kinds of operations.

(ii) The maximum value dem- The kinds of operations (such as onstrated during the type tests; VFR, IFR, day, night, or icing) for (2) The maximum allowable gas tem- perature; and which the rotorcraft is approved are es- (3) The maximum allowable torque. tablished by demonstrated compliance Federal Aviation Administration, DOT § 27.1547 with the applicable certification re- ment of the glass cover with the face of quirements and by the installed equip- the dial; and ment. (b) Each arc and line must be wide enough, and located, to be clearly visi- [Amdt. 27–21, 49 FR 44435, Nov. 6, 1984] ble to the pilot.

§ 27.1527 Maximum operating altitude.

§ 27.1545 Airspeed indicator.

The maximum altitude up to which (a) Each airspeed indicator must be operation is allowed, as limited by marked as specified in paragraph (b) of flight, structural, powerplant, func- this section, with the marks located at tional, or equipment characteristics, the corresponding indicated airspeeds.

must be established.

(b) The following markings must be made: (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (1) A red line— eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the (i) For rotorcraft other than heli- Dept. of Transportation Act (49 U.S.C.

copters, at V .

NE 1655(c))) (ii) For helicopters, at V (power- NE on).

[Amdt. 27–14, 43 FR 2325, Jan. 16, 1978] (iii) For helicopters, at V (power- NE § 27.1529 Instructions for Continued off). If V (power-off) is less than V NE NE Airworthiness.

(power-on) and both are simulta- neously displayed, the red line at V NE The applicant must prepare Instruc- (power-off) must be clearly distinguish- tions for Continued Airworthiness in able from the red line at V (power- NE accordance with appendix A to this on).

part that are acceptable to the Admin- (2) [Reserved] istrator. The instructions may be in- (3) For the caution range, a yellow complete at type certification if a pro- range.

gram exists to ensure their completion (4) For the normal operating range, a prior to delivery of the first rotorcraft green or unmarked range.

or issuance of a standard certificate of airworthiness, whichever occurs later.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), [Amdt. 27–18, 45 FR 60177, Sept. 11, 1980] 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C.

M ARKINGS AND P LACARDS 1655(c))) § 27.1541 General. [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–14, 43 FR 2325, Jan. 16, (a) The rotorcraft must contain— 1978; 43 FR 3900, Jan. 30, 1978; Amdt. 27–16, 43 (1) The markings and placards speci- FR 50599, Oct. 30, 1978; Amdt. 27–51, 88 FR fied in §§ 27.1545 through 27.1565, and 8738, Feb. 10, 2023] (2) Any additional information, in- § 27.1547 Magnetic direction indicator.

strument markings, and placards re- quired for the safe operation of rotor- (a) A placard meeting the require- craft with unusual design, operating or ments of this section must be installed handling characteristics.

on or near the magnetic direction indi- (b) Each marking and placard pre- cator.

scribed in paragraph (a) of this sec- (b) The placard must show the cali- tion— bration of the instrument in level (1) Must be displayed in a con- flight with the engines operating.

spicuous place; and (c) The placard must state whether (2) May not be easily erased, dis- the calibration was made with radio re- figured, or obscured.

ceivers on or off.

(d) Each calibration reading must be § 27.1543 Instrument markings: gen- in terms of magnetic heading in not eral.

more than 45 degree increments.

For each instrument— (e) If a magnetic nonstabilized direc- (a) When markings are on the cover tion indicator can have a deviation of glass of the instrument, there must be more than 10 degrees caused by the op- means to maintain the correct align- eration of electrical equipment, the 14 CFR Ch. I (1–1–25 Edition) § 27.1549 placard must state which electrical § 27.1555 Control markings.

loads, or combination of loads, would (a) Each cockpit control, other than cause a deviation of more than 10 de- primary flight controls or control grees when turned on.

whose function is obvious, must be (Secs. 313(a), 601, 603, 604, and 605 of the Fed- plainly marked as to its function and eral Aviation Act of 1958 (49 U.S.C. 1354(a), method of operation.

1421, 1423, 1424, and 1425); and sec. 6(c) of the (b) For powerplant fuel controls— Dept. of Transportation Act (49 U.S.C.

(1) Each fuel tank selector control 1655(c))) must be marked to indicate the posi- [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as tion corresponding to each tank and to amended by Amdt. 27–13, 42 FR 36972, July 18, each existing cross feed position; 1977] (2) If safe operation requires the use of any tanks in a specific sequence, § 27.1549 Powerplant instruments.

that sequence must be marked on, or For each required powerplant instru- adjacent to, the selector for those ment, as appropriate to the type of in- tanks; and strument— (3) Each valve control for any engine (a) Each maximum and, if applicable, of a multiengine rotorcraft must be minimum safe operating limit must be marked to indicate the position cor- marked with a red line; responding to each engine controlled.

(b) Each normal operating range (c) Usable fuel capacity must be must be marked as a green or un- marked as follows: marked range; (1) For fuel systems having no selec- (c) Each takeoff and precautionary tor controls, the usable fuel capacity of range must be marked with a yellow the system must be indicated at the range or yellow line; fuel quantity indicator unless it is: (d) Each engine or rotor range that is (i) Provided by another system or restricted because of excessive vibra- equipment readily accessible to the tion stresses must be marked with red pilot; and ranges or red lines; and (ii) Contained in the limitations sec- (e) Each OEI limit or approved oper- tion of the rotorcraft flight manual.

ating range must be marked to be (2) For fuel systems having selector clearly differentiated from the mark- controls, the usable fuel capacity ings of paragraphs (a) through (d) of available at each selector control posi- this section except that no marking is tion must be indicated near the selec- normally required for the 30-second tor control.

OEI limit.

(d) For accessory, auxiliary, and [Amdt. 27–11, 41 FR 55470, Dec. 20, 1976, as emergency controls— amended by Amdt. 27–23, 53 FR 34215, Sept. 2, (1) Each essential visual position in- 1988; Amdt. 27–29, 59 FR 47768, Sept. 16, 1994; dicator, such as those showing rotor Amdt. 27–51, 88 FR 8738, Feb. 10, 2023] pitch or landing gear position, must be marked so that each crewmember can § 27.1551 Oil quantity indicator.

determine at any time the position of Each oil quantity indicator must be the unit to which it relates; and marked with enough increments to in- (2) Each emergency control must be dicate readily and accurately the quan- red and must be marked as to method tity of oil.

of operation.

(e) For rotorcraft incorporating re- § 27.1553 Fuel quantity indicator.

tractable landing gear, the maximum If the unusable fuel supply for any landing gear operating speed must be tank exceeds one gallon, or five per- displayed in clear view of the pilot.

cent of the tank capacity, whichever is [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as greater, a red arc must be marked on amended by Amdt. 27–11, 41 FR 55470, Dec. 20, its indicator extending from the cali- 1976; Amdt. 27–21, 49 FR 44435, Nov. 6, 1984; brated zero reading to the lowest read- Amdt. 27–51, 88 FR 8738, Feb. 10, 2023] ing obtainable in level flight.

Federal Aviation Administration, DOT § 27.1583 (b) Each location, such as a locker or § 27.1557 Miscellaneous markings and placards. compartment, that carries any fire ex- tinguishing, signaling, or other life (a) Baggage and cargo compartments, saving equipment, must be so marked.

and ballast location. Each baggage and cargo compartment, and each ballast § 27.1565 Tail rotor.

location must have a placard stating Each tail rotor must be marked so any limitations on contents, including that its disc is conspicuous under nor- weight, that are necessary under the mal daylight ground conditions.

loading requirements.

(b) Seats. If the maximum allowable [Amdt. 27–2, 33 FR 965, Jan. 26, 1968] weight to be carried in a seat is less R OTORCRAFT F LIGHT M ANUAL AND than 170 pounds, a placard stating the A PPROVED M ANUAL M ATERIAL lesser weight must be permanently at- tached to the seat structure.

§ 27.1581 General.

(c) Fuel and oil filler openings. The fol- lowing apply: (a) Furnishing information. A Rotor- (1) Fuel filler openings must be craft Flight Manual must be furnished marked at or near the filler cover with each rotorcraft, and it must con- with— tain the following: (i) The word ‘‘fuel’’; (1) Information required by §§ 27.1583 through 27.1589.

(ii) For reciprocating engine powered (2) Other information that is nec- rotorcraft, the minimum fuel grade; essary for safe operation because of de- (iii) For turbine engine powered sign, operating, or handling character- rotorcraft, the permissible fuel des- istics.

ignations; and (b) Approved information. Each part of (iv) For pressure fueling systems, the the manual listed in §§ 27.1583 through maximum permissible fueling supply 27.1589, that is appropriate to the rotor- pressure and the maximum permissible craft, must be furnished, verified, and defueling pressure.

approved, and must be segregated, (2) Oil filler openings must be identified, and clearly distinguished marked at or near the filler cover with from each unapproved part of that the word ‘‘oil’’.

manual.

(d) Emergency exit placards. Each (c) [Reserved] placard and operating control for each (d) Table of contents. Each Rotorcraft emergency exit must be red. A placard Flight Manual must include a table of must be near each emergency exit con- contents if the complexity of the man- trol and must clearly indicate the loca- ual indicates a need for it.

tion of that exit and its method of op- eration.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as 1421, 1423, 1424, and 1425); and sec. 6(c) of the amended by Amdt. 27–11, 41 FR 55471, Dec. 20, Dept. of Transportation Act (49 U.S.C.

1976] 1655(c))) § 27.1559 Limitations placard. [Amdt. 27–14, 43 FR 2325, Jan. 16, 1978] There must be a placard in clear view § 27.1583 Operating limitations.

of the pilot that specifies the kinds of (a) Airspeed and rotor limitations. In- operations (such as VFR, IFR, day, formation necessary for the marking of night, or icing) for which the rotorcraft airspeed and rotor limitations on, or is approved.

near, their respective indicators must [Amdt. 27–21, 49 FR 44435, Nov. 6, 1984] be furnished. The significance of each limitation and of the color coding must § 27.1561 Safety equipment.

be explained.

(a) Each safety equipment control to (b) Powerplant limitations. The fol- be operated by the crew in emergency, lowing information must be furnished: such as controls for automatic liferaft (1) Limitations required by § 27.1521.

releases, must be plainly marked as to (2) Explanation of the limitations, its method of operation. when appropriate.

14 CFR Ch. I (1–1–25 Edition) § 27.1585 (3) Information necessary for mark- (c) For helicopters for which a V NE ing the instruments required by (power-off) is established under §§ 27.1549 through 27.1553. § 27.1505(c), information must be fur- nished to explain the V (power-off) (c) Weight and loading distribution. NE and the procedures for reducing air- The weight and center of gravity limits speed to not more than the V (power- required by §§ 27.25 and 27.27, respec- NE off) following failure of all engines.

tively, must be furnished. If the vari- (d) For each rotorcraft showing com- ety of possible loading conditions war- pliance with § 27.1353 (g)(2) or (g)(3), the rants, instructions must be included to operating procedures for disconnecting allow ready observance of the limita- the battery from its charging source tions.

must be furnished.

(d) Flight crew. When a flight crew of (e) If the unusable fuel supply in any more than one is required, the number tank exceeds five percent of the tank and functions of the minimum flight capacity, or one gallon, whichever is crew determined under § 27.1523 must be greater, information must be furnished furnished.

which indicates that when the fuel (e) Kinds of operation. Each kind of quantity indicator reads ‘‘zero’’ in operation for which the rotorcraft and level flight, any fuel remaining in the its equipment installations are ap- fuel tank cannot be used safely in proved must be listed.

flight.

(f) [Reserved] (f) Information on the total quantity (g) Altitude. The altitude established of usable fuel for each fuel tank must under § 27.1527 and an explanation of be furnished.

the limiting factors must be furnished.

(g) The airspeeds and rotor speeds for (Secs. 313(a), 601, 603, 604, and 605 of the Fed- minimum rate of descent and best glide eral Aviation Act of 1958 (49 U.S.C. 1354(a), angle as prescribed in § 27.71 must be 1421, 1423, 1424, and 1425); and sec. 6(c) of the provided.

Dept. of Transportation Act (49 U.S.C.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- 1655(c))) eral Aviation Act of 1958 (49 U.S.C. 1354(a), [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as 1421, 1423, 1424, and 1425); and sec. 6(c) of the amended by Amdt. 27–2, 33 FR 965, Jan. 26, Dept. of Transportation Act (49 U.S.C.

1968; Amdt. 27–14, 43 FR 2325, Jan. 16, 1978; 1655(c))) Amdt. 27–16, 43 FR 50599, Oct. 30, 1978] [Amdt. 27–1, 32 FR 6914, May 5, 1967, as amended by Amdt. 27–14, 43 FR 2326, Jan. 16, § 27.1585 Operating procedures.

1978; Amdt. 27–16, 43 FR 50599, Oct. 30, 1978; (a) Parts of the manual containing Amdt. 27–21, 49 FR 44435, Nov. 6, 1984] operating procedures must have infor- § 27.1587 Performance information.

mation concerning any normal and emergency procedures and other infor- (a) The Rotorcraft Flight Manual mation necessary for safe operation, must contain the following informa- including takeoff and landing proce- tion, determined in accordance with dures and associated airspeeds. The §§ 27.49 through 27.87 and 27.143(c) and manual must contain any pertinent in- (d): formation including— (1) Enough information to determine (1) The kind of takeoff surface used the limiting height-velocity envelope.

in the tests and each appropriate (2) Information relative to— climbout speed; and (i) The steady rates of climb and de- (2) The kind of landing surface used scent, in-ground effect and out-of- in the tests and appropriate approach ground effect hovering ceilings, to- and glide airspeeds.

gether with the corresponding air- (b) For multiengine rotorcraft, infor- speeds and other pertinent information mation identifying each operating con- including the calculated effects of alti- dition in which the fuel system inde- tude and temperatures; pendence prescribed in § 27.953 is nec- (ii) The maximum weight for each al- essary for safety must be furnished, to- titude and temperature condition at gether with instructions for placing which the rotorcraft can safely hover the fuel system in a configuration used in-ground effect and out-of-ground ef- to show compliance with that section. fect in winds of not less than 17 knots Federal Aviation Administration, DOT Pt. 27, App. A (a) This appendix specifies requirements from all azimuths. These data must be for the preparation of Instructions for Con- clearly referenced to the appropriate tinued Airworthiness as required by § 27.1529.

hover charts. In addition, if there are (b) The Instructions for Continued Air- other combinations of weight, altitude worthiness for each rotorcraft must include and temperature for which perform- the Instructions for Continued Airworthiness ance information is provided and at for each engine and rotor (hereinafter des- which the rotorcraft cannot land and ignated ‘products’), for each appliance re- quired by this chapter, and any required in- take off safely with the maximum wind formation relating to the interface of those value, those portions of the operating appliances and products with the rotorcraft.

envelope and the appropriate safe wind If Instructions for Continued Airworthiness conditions must be stated in the Rotor- are not supplied by the manufacturer of an craft Flight Manual; appliance or product installed in the rotor- (iii) For reciprocating engine-pow- craft, the Instructions for Continued Air- ered rotorcraft, the maximum atmos- worthiness for the rotorcraft must include the information essential to the continued pheric temperature at which compli- airworthiness of the rotorcraft.

ance with the cooling provisions of (c) The applicant must submit to the FAA §§ 27.1041 through 27.1045 is shown; and a program to show how changes to the In- (iv) Glide distance as a function of al- structions for Continued Airworthiness made titude when autorotating at the speeds by the applicant or by the manufacturers of and conditions for minimum rate of de- products and appliances installed in the scent and best glide as determined in rotorcraft will be distributed.

§ 27.71. A27.2 Format.

(b) The Rotorcraft Flight Manual (a) The Instructions for Continued Air- must contain— worthiness must be in the form of a manual or manuals as appropriate for the quantity (1) In its performance information of data to be provided.

section any pertinent information con- (b) The format of the manual or manuals cerning the takeoff weights and alti- must provide for a practical arrangement.

tudes used in compliance with § 27.51; A27.3 Content.

and The contents of the manual or manuals (2) The horizontal takeoff distance must be prepared in the English language.

determined in accordance with The Instructions for Continued Airworthi- § 27.65(a)(2)(i).

ness must contain the following manuals or sections, as appropriate, and information: (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (a) Rotorcraft maintenance manual or section.

eral Aviation Act of 1958 (49 U.S.C. 1354(a), (1) Introduction information that includes an 1421, 1423, 1424, and 1425); and sec. 6(c) of the explanation of the rotorcraft’s features and Dept. of Transportation Act (49 U.S.C.

data to the extent necessary for mainte- 1655(c))) nance or preventive maintenance.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as (2) A description of the rotorcraft and its amended by Amdt. 27–14, 43 FR 2326, Jan. 16, systems and installations including its en- 1978; Amdt. 27–21, 49 FR 44435, Nov. 6, 1984; gines, rotors, and appliances.

(3) Basic control and operation information Amdt. 27–44, 73 FR 11000, Feb. 29, 2008; 73 FR describing how the rotorcraft components 33876, June 16, 2008; Amdt. 27–51, 88 FR 8739, and systems are controlled and how they op- Feb. 10, 2023] erate, including any special procedures and limitations that apply.

§ 27.1589 Loading information.

(4) Servicing information that covers de- There must be loading instructions tails regarding servicing points, capacities of for each possible loading condition be- tanks, reservoirs, types of fluids to be used, tween the maximum and minimum pressures applicable to the various systems, location of access panels for inspection and weights determined under § 27.25 that servicing, locations of lubrication points, the can result in a center of gravity beyond lubricants to be used, equipment required for any extreme prescribed in § 27.27, as- servicing, tow instructions and limitations, suming any probable occupant weights.

mooring, jacking, and leveling information.

(b) Maintenance instructions. (1) Scheduling information for each part of the rotorcraft and its engines, auxiliary power units, ro- tors, accessories, instruments and equipment A PPENDIX A TO P ART 27—I NSTRUCTIONS that provides the recommended periods at FOR C ONTINUED A IRWORTHINESS which they should be cleaned, inspected, ad- A27.1 General. justed, tested, and lubricated, and the degree 14 CFR Ch. I (1–1–25 Edition) Pt. 27, App. B of inspection, the applicable wear tolerances, Regulations unless an alternative program and work recommended at these periods. has been FAA approved.’’ However, the applicant may refer to an ac- [Amdt. 27–18, 45 FR 60177, Sept. 11, 1980, as cessory, instrument, or equipment manufac- amended by Amdt. 27–24, 54 FR 34329, Aug. 18, turer as the source of this information if the 1989; Amdt. 27–47, 76 FR 74663, Dec. 1, 2011] applicant shows the item has an exception- ally high degree of complexity requiring spe- A PPENDIX B TO P ART 27—A IRWORTHI - cialized maintenance techniques, test equip- NESS C RITERIA FOR H ELICOPTER I N- ment, or expertise. The recommended over- STRUMENT F LIGHT haul periods and necessary cross references to the Airworthiness Limitations section of I. General. A normal category helicopter the manual must also be included. In addi- may not be type certificated for operation tion, the applicant must include an inspec- under the instrument flight rules (IFR) of tion program that includes the frequency this chapter unless it meets the design and and extent of the inspections necessary to installation requirements contained in this provide for the continued airworthiness of appendix.

the rotorcraft.

II. Definitions. (a) V means instrument YI (2) Troubleshooting information describing climb speed, utilized instead of V for com- Y problem malfunctions, how to recognize pliance with the climb requirements for in- those malfunctions, and the remedial action strument flight.

for those malfunctions.

means instrument flight never ex- (b) V NEI (3) Information describing the order and for com- ceed speed, utilized instead of V NE method of removing and replacing products pliance with maximum limit speed require- and parts with any necessary precautions to ments for instrument flight.

be taken.

(c) V means instrument flight min- MINI (4) Other general procedural instructions imum speed, utilized in complying with min- including procedures for system testing dur- imum limit speed requirements for instru- ing ground running, symmetry checks, ment flight.

weighing and determining the center of grav- III. Trim. It must be possible to trim the ity, lifting and shoring, and storage limita- cyclic, collective, and directional control tions. forces to zero at all approved IFR airspeeds, power settings, and configurations appro- (c) Diagrams of structural access plates priate to the type.

and information needed to gain access for in- spections when access plates are not pro- IV. Static longitudinal stability. (a) General.

vided. The helicopter must possess positive static longitudinal control force stability at crit- (d) Details for the application of special in- ical combinations of weight and center of spection techniques including radiographic gravity at the conditions specified in para- and ultrasonic testing where such processes graph IV (b) or (c) of this appendix, as appro- are specified.

priate. The stick force must vary with speed (e) Information needed to apply protective so that any substantial speed change results treatments to the structure after inspection.

in a stick force clearly perceptible to the (f) All data relative to structural fasteners pilot. For single-pilot approval, the airspeed such as identification, discarded rec- must return to within 10 percent of the trim ommendations, and torque values.

speed when the control force is slowly re- (g) A list of special tools needed.

leased for each trim condition specified in A27.4 Airworthiness Limitations section.

paragraph IV(b) of the this appendix.

(b) For single-pilot approval: The Instructions for Continued Airworthi- ness must contain a section, titled Air- (1) Climb. Stability must be shown in climb worthiness Limitations that is segregated throughout the speed range 20 knots either side of trim with— and clearly distinguishable from the rest of (i) The helicopter trimmed at V ; the document. This section must set forth YI (ii) Landing gear retracted (if retractable); each mandatory replacement time, struc- and tural inspection interval, and related struc- tural inspection procedure required for type (iii) Power required for limit climb rate (at certification. If the Instructions for Contin- least 1,000 fpm) at V or maximum contin- YI ued Airworthiness consist of multiple docu- uous power, whichever is less.

ments, the section required by this para- (2) Cruise. Stability must be shown graph must be included in the principal man- throughout the speed range from 0.7 to 1.1 V H ual. This section must contain a legible or V , whichever is lower, not to exceed ± 20 NEI knots from trim with— statement in a prominent location that reads: ‘‘The Airworthiness Limitations sec- (i) The helicopter trimmed and power ad- tion is FAA approved and specifies inspec- justed for level flight at 0.9 V H or 0.9 V NEI , tions and other maintenance required under whichever is lower; and §§ 43.16 and 91.403 of the Federal Aviation (ii) Landing gear retracted (if retractable).

Federal Aviation Administration, DOT Pt. 27, App. B (3) Slow cruise. Stability must be shown (3) Any oscillation having a period of 10 throughout the speed range from 0.9 V to seconds or more but less than 20 seconds MINI 1.3 V or 20 knots above trim speed, which- must be damped.

MINI ever is greater, with— (4) Any oscillation having a period of 20 seconds or more may not achieve double am- (i) the helicopter trimmed and power ad- plitude in less than 20 seconds.

justed for level flight at 1.1 V ; and MINI (5) Any aperiodic response may not achieve (ii) Landing gear retracted (if retractable).

double amplitude in less than 6 seconds.

(4) Descent. Stability must be shown (b) For helicopters approved with a min- throughout the speed range 20 knots either imum crew of two pilots— side of trim with— (1) Any oscillation having a period of less (i) The helicopter trimmed at 0.8 V or 0.8 H ⁄2 amplitude in than 5 seconds must damp to V (or 0.8 V for the landing gear extended NEI LE not more than two cycles.

case), whichever is lower; (2) Any oscillation having a period of 5 sec- (ii) Power required for 1,000 fpm descent at onds or more but less than 10 seconds must trim speed; and be damped.

(iii) Landing gear extended and retracted, (3) Any oscillation having a period of 10 if applicable.

seconds or more may not achieve double am- (5) Approach. Stability must be shown plitude in less than 10 seconds.

throughout the speed range from 0.7 times VII. Stability Augmentation System (SAS).

the minimum recommended approach speed (a) If a SAS is used, the reliability of the to 20 knots above the maximum rec- SAS must be related to the effects of its fail- ommended approach speed with— ure. Any SAS failure condition that would prevent continued safe flight and landing (i) The helicopter trimmed at the rec- must be extremely improbable. It must be ommended approach speed or speeds; shown that, for any failure condition of the (ii) Landing gear extended and retracted, if SAS that is not shown to be extremely im- applicable; and probable— (iii) Power required to maintain a 3 ° glide (1) The helicopter is safely controllable path and power required to maintain the when the failure or malfunction occurs at steepest approach gradient for which ap- any speed or altitude within the approved proval is requested.

IFR operating limitations; and (c) Helicopters approved for a minimum (2) The overall flight characteristics of the crew of two pilots must comply with the pro- helicopter allow for prolonged instrument visions of paragraphs IV(b)(2) and IV(b)(5) of flight without undue pilot effort. Additional this appendix.

unrelated probable failures affecting the con- V. Static Lateral Directional Stability. (a) trol system must be considered. In addi- Static directional stability must be positive tion— throughout the approved ranges of airspeed, (i) The controllability and maneuver- power, and vertical speed. In straight and ability requirements in Subpart B of this steady sideslips up to ± 10 ° from trim, direc- part must be met throughout a practical tional control position must increase with- flight envelope; out discontinuity with the angle of sideslip, (ii) The flight control, trim, and dynamic except for a small range of sideslip angles stability characteristics must not be im- around trim. At greater angles up to the paired below a level needed to allow contin- maximum sideslip angle appropriate to the ued safe flight and landing; and type, increased directional control position (iii) The static longitudinal and static di- must produce an increased angle of sideslip.

rectional stability requirements of Subpart It must be possible to maintain balanced B must be met throughout a practical flight flight without exceptional pilot skill or envelope.

alertness.

(b) The SAS must be designed so that it (b) During sideslips up to ± 10 ° from trim cannot create a hazardous deviation in flight throughout the approved ranges of airspeed, path or produce hazardous loads on the heli- power, and vertical speed, there must be no copter during normal operation or in the negative dihedral stability perceptible to the event of malfunction or failure, assuming pilot through lateral control motion or corrective action begins within an appro- force. Longitudinal cyclic movement with priate period of time. Where multiple sys- sideslip must not be excessive.

tems are installed, subsequent malfunction VI. Dynamic stability. (a) For single-pilot conditions must be considered in sequence approval— unless their occurrence is shown to be im- (1) Any oscillation having a period of less probable.

than 5 seconds must damp to ⁄ 2 amplitude in VIII. Equipment, systems, and installation.

not more than one cycle.

The basic equipment and installation must (2) Any oscillation having a period of 5 sec- comply with §§ 29.1303, 29.1431, and 29.1433, onds or more but less than 10 seconds must with the following exceptions and additions: damp to ⁄2 amplitude in not more than two (a) Flight and Navigation Instruments. (1) A cycles. magnetic gyro-stablized direction indicator 14 CFR Ch. I (1–1–25 Edition) Pt. 27, App. C instead of a gyroscopic direction indicator (a) Limitations. The approved IFR flight en- required by § 29.1303(h); and velope, the IFR flightcrew composition, the (2) A standby attitude indicator which revised kinds of operation, and the steepest meets the requirements of §§ 29.1303(g)(1) IFR precision approach gradient for which through (7) instead of a rate-of-turn indi- the helicopter is approved; cator required by § 29.1303(g). For two-pilot (b) Procedures. Required information for configurations, one pilot’s primary indicator proper operation of IFR systems and the rec- may be designated for this purpose. If stand- ommended procedures in the event of sta- by batteries are provided, they may be bility augmentation or electrical system charged from the aircraft electrical system failures; and if adequate isolation is incorporated. (c) Performance. If V YI differs from V Y , climb performance at V YI and with maximum (b) Miscellaneous requirements. (1) Instru- continuous power throughout the ranges of ment systems and other systems essential weight, altitude, and temperature for which for IFR flight that could be adversely af- approval is requested.

fected by icing must be adequately protected X. Electrical and electronic system light- when exposed to the continuous and inter- ning protection. For regulations concerning mittent maximum icing conditions defined lightning protection for electrical and elec- in appendix C of Part 29 of this chapter, tronic systems, see § 27.1316.

whether or not the rotorcraft is certificated for operation in icing conditions.

[Amdt. 27–19, 48 FR 4389, Jan. 31, 1983, as (2) There must be means in the generating amended by Amdt. 27–44, 73 FR 11000, Feb. 29, system to automatically de-energize and dis- 2008; Amdt. 27–46, 76 FR 33135, June 8, 2011; connect from the main bus any power source Amdt. 27–51, 88 FR 8739, Feb. 10, 2023] developing hazardous overvoltage.

(3) Each required flight instrument using a A PPENDIX C TO P ART 27—C RITERIA FOR power supply (electric, vacuum, etc.) must C ATEGORY A have a visual means integral with the instru- ment to indicate the adequacy of the power C27.1 General.

being supplied.

A small multiengine rotorcraft may not be (4) When multiple systems performing like type certificated for Category A operation functions are required, each system must be unless it meets the design installation and grouped, routed, and spaced so that physical performance requirements contained in this separation between systems is provided to appendix in addition to the requirements of ensure that a single malfunction will not ad- this part.

versely affect more than one system.

C27.2 Applicable part 29 sections. The fol- (5) For systems that operate the required lowing sections of part 29 of this chapter flight instruments at each pilot’s station— must be met in addition to the requirements (i) For pneumatic systems, only the re- of this part: quired flight instruments for the first pilot 29.45(a) and (b)(2)—General.

may be connected to that operating system; 29.49(a)—Performance at minimum operating (ii) Additional instruments, systems, or speed.

equipment may not be connected to an oper- 29.51—Takeoff data: General.

ating system for a second pilot unless provi- 29.53—Takeoff: Category A.

sions are made to ensure the continued nor- 29.55—Takeoff decision point: Category A.

mal functioning of the required instruments 29.59—Takeoff Path: Category A.

in the event of any malfunction of the addi- 29.60—Elevated heliport takeoff path: Cat- tional instruments, systems, or equipment egory A.

which is not shown to be extremely improb- 29.61—Takeoff distance: Category A.

able; 29.62—Rejected takeoff: Category A.

(iii) The equipment, systems, and installa- 29.64—Climb: General.

tions must be designed so that one display of 29.65(a)—Climb: AEO.

the information essential to the safety of 29.67(a)—Climb: OEI.

flight which is provided by the instruments 29.75—Landing: General.

will remain available to a pilot, without ad- 29.77—Landing decision point: Category A.

ditional crewmember action, after any single 29.79—Landing: Category A.

failure or combination of failures that is not 29.81—Landing distance (Ground level sites): shown to be extremely improbable; and Category A.

(iv) For single-pilot configurations, instru- 29.85—Balked landing: Category A.

ments which require a static source must be 29.87(a)—Height-velocity envelope.

provided with a means of selecting an alter- 29.547(a) and (b)—Main and tail rotor struc- nate source and that source must be cali- ture.

brated. 29.861(a)—Fire protection of structure, con- IX. Rotorcraft Flight Manual. A Rotorcraft trols, and other parts.

Flight Manual or Rotorcraft Flight Manual 29.901(c)—Powerplant: Installation.

IFR Supplement must be provided and must 29.903 (b) and (c)—Engines.

contain— 29.908(a)—Cooling fans.

Federal Aviation Administration, DOT Pt. 27, App. D 29.917(b) and (c)(1)—Rotor drive system: De- T ABLE I.—HIRF E NVIRONMENT I—Continued sign.

Field strength 29.927(c)(1)—Additional tests.

(volts/meter) 29.953(a)—Fuel system independence. Frequency 29.1027(a)—Transmission and gearboxes: Gen- Peak Average eral.

29.1045(a)(1), (b), (c), (d), and (f)—Climb cool- 1 GHz–2 GHz .................................... 2,000 200 2 GHz–6 GHz .................................... 3,000 200 ing test procedures.

6 GHz–8 GHz .................................... 1,000 200 29.1047(a)—Takeoff cooling test procedures.

8 GHz–12 GHz .................................. 3,000 300 29.1181(a)—Designated fire zones: Regions in- 12 GHz–18 GHz ................................ 2,000 200 cluded.

18 GHz–40 GHz ................................ 600 200 29.1187(e)—Drainage and ventilation of fire In this table, the higher field strength applies at the fre- zones.

quency band edges.

29.1189(c)—Shutoff means.

29.1191(a)(1)—Firewalls.

(b) HIRF environment II is specified in the 29.1193(e)—Cowling and engine compartment following table: covering.

29.1195(a) and (d)—Fire extinguishing sys- T ABLE II.—HIRF E NVIRONMENT II tems (one shot).

29.1197—Fire extinguishing agents.

Field strength (volts/meter) 29.1199—Extinguishing agent containers.

Frequency 29.1201—Fire extinguishing system materials.

Peak Average 29.1305(a) (6) and (b)—Powerplant instru- ments. 10 kHz–500 kHz ................................ 20 20 500 kHz–2 MHz ................................. 30 30 29.1309(b)(2) (i) and (d)—Equipment, systems, 2 MHz–30 MHz ................................. 100 100 and installations.

30 MHz–100 MHz ............................. 10 10 29.1323(c)(1)—Airspeed indicating system.

100 MHz–200 MHz ........................... 30 10 29.1331(b)—Instruments using a power supply.

200 MHz–400 MHz ........................... 10 10 29.1351(d)(2)—Electrical systems and equip- 400 MHz–1 GHz ................................ 700 40 ment: General (operation without normal 1 GHz–2 GHz .................................... 1,300 160 electrical power).

2 GHz–4 GHz .................................... 3,000 120 29.1587(a)—Performance information. 4 GHz–6 GHz .................................... 3,000 160 6 GHz–8 GHz .................................... 400 170 OTE : In complying with the paragraphs N 8 GHz–12 GHz .................................. 1,230 230 listed in paragraph C27.2 above, relevant ma- 12 GHz–18 GHz ................................ 730 190 terial in the AC ‘‘Certification of Transport 18 GHz–40 GHz ................................ 600 150 Category Rotorcraft’’ should be used.

In this table, the higher field strength applies at the fre- [Doc. No. 28008, 61 FR 21907, May 10, 1996, as quency band edges.

amended by Amdt. 27–51, 88 FR 8739, Feb. 10, (c) HIRF environment III is specified in the 2023] following table: A PPENDIX D TO P ART 27—HIRF E NVI - T ABLE III.—HIRF E NVIRONMENT III RONMENTS AND E QUIPMENT HIRF T EST L EVELS Field strength (volts/meter) Frequency This appendix specifies the HIRF environ- Peak Average ments and equipment HIRF test levels for electrical and electronic systems under 10 kHz–100 kHz ................................ 150 150 § 27.1317. The field strength values for the 100 kHz–400 MHz ............................. 200 200 HIRF environments and laboratory equip- 400 MHz–700 MHz ........................... 730 200 ment HIRF test levels are expressed in root- 700 MHz–1 GHz ................................ 1,400 240 1 GHz–2 GHz .................................... 5,000 250 mean-square units measured during the peak 2 GHz–4 GHz .................................... 6,000 490 of the modulation cycle.

4 GHz–6 GHz .................................... 7,200 400 (a) HIRF environment I is specified in the 6 GHz–8 GHz .................................... 1,100 170 following table: 8 GHz–12 GHz .................................. 5,000 330 12 GHz–18 GHz ................................ 2,000 330 T ABLE I.—HIRF E NVIRONMENT I 18 GHz–40 GHz ................................ 1,000 420 In this table, the higher field strength applies at the fre- Field strength quency band edges.

(volts/meter) Frequency (d) Equipment HIRF Test Level 1. (1) From 10 Peak Average kilohertz (kHz) to 400 megahertz (MHz), use 10 kHz–2 MHz ................................... 50 50 conducted susceptibility tests with contin- 2 MHz–30 MHz ................................. 100 100 uous wave (CW) and 1 kHz square wave mod- 30 MHz–100 MHz ............................. 50 50 ulation with 90 percent depth or greater. The 100 MHz–400 MHz ........................... 100 100 conducted susceptibility current must start 400 MHz–700 MHz ........................... 700 50 at a minimum of 0.6 milliamperes (mA) at 10 700 MHz–1 GHz ................................ 700 100 14 CFR Ch. I (1–1–25 Edition) Pt. 29 kHz, increasing 20 decibels (dB) per fre- 29.33 Main rotor speed and pitch limits.

quency decade to a minimum of 30 mA at 500 P ERFORMANCE kHz.

(2) From 500 kHz to 40 MHz, the conducted 29.45 General.

susceptibility current must be at least 30 29.49 Performance at minimum operating mA.

speed.

(3) From 40 MHz to 400 MHz, use conducted 29.51 Takeoff data: general.

susceptibility tests, starting at a minimum 29.53 Takeoff: Category A.

of 30 mA at 40 MHz, decreasing 20 dB per fre- 29.55 Takeoff decision point (TDP): Cat- quency decade to a minimum of 3 mA at 400 egory A.

MHz.

29.59 Takeoff path: Category A.

(4) From 100 MHz to 400 MHz, use radiated 29.60 Elevated heliport takeoff path: Cat- susceptibility tests at a minimum of 20 volts egory A.

per meter (V/m) peak with CW and 1 kHz 29.61 Takeoff distance: Category A.

square wave modulation with 90 percent 29.62 Rejected takeoff: Category A.

depth or greater.

29.63 Takeoff: Category B.

(5) From 400 MHz to 8 gigahertz (GHz), use 29.64 Climb: General.

radiated susceptibility tests at a minimum 29.65 Climb: All engines operating.

of 150 V/m peak with pulse modulation of 4 29.67 Climb: One engine inoperative (OEI).

percent duty cycle with a 1 kHz pulse repeti- 29.71 Helicopter angle of glide: Category B.

tion frequency. This signal must be switched 29.75 Landing: General.

on and off at a rate of 1 Hz with a duty cycle 29.77 Landing Decision Point (LDP): Cat- of 50 percent.

egory A.

(e) Equipment HIRF Test Level 2. Equipment 29.79 Landing: Category A.

HIRF test level 2 is HIRF environment II in 29.81 Landing distance: Category A.

table II of this appendix reduced by accept- 29.83 Landing: Category B.

able aircraft transfer function and attenu- 29.85 Balked landing: Category A.

ation curves. Testing must cover the fre- 29.87 Height-velocity envelope.

quency band of 10 kHz to 8 GHz.

(f) Equipment HIRF Test Level 3. (1) From 10 FLIGHT C HARACTERISTICS kHz to 400 MHz, use conducted susceptibility 29.141 General.

tests, starting at a minimum of 0.15 mA at 10 29.143 Controllability and maneuverability.

kHz, increasing 20 dB per frequency decade 29.151 Flight controls.

to a minimum of 7.5 mA at 500 kHz.

29.161 Trim control.

(2) From 500 kHz to 40 MHz, use conducted 29.171 Stability: general.

susceptibility tests at a minimum of 7.5 mA.

29.173 Static longitudinal stability.

(3) From 40 MHz to 400 MHz, use conducted 29.175 Demonstration of static longitudinal susceptibility tests, starting at a minimum stability.

of 7.5 mA at 40 MHz, decreasing 20 dB per fre- 29.177 Static directional stability.

quency decade to a minimum of 0.75 mA at 29.181 Dynamic stability: Category A rotor- 400 MHz.

craft.

(4) From 100 MHz to 8 GHz, use radiated susceptibility tests at a minimum of 5 V/m.

G ROUND AND W ATER H ANDLING C HARACTERISTICS [Doc. No. FAA–2006–23657, 72 FR 44027, Aug. 6, 2007] 29.231 General.

29.235 Taxiing condition.

29.239 Spray characteristics.

PART 29—AIRWORTHINESS STAND- 29.241 Ground resonance.

ARDS: TRANSPORT CATEGORY ROTORCRAFT M ISCELLANEOUS F LIGHT R EQUIREMENTS 29.251 Vibration.

Subpart A—General Subpart C—Strength Requirements Sec.

29.1 Applicability.

G ENERAL 29.2 Special retroactive requirements.

29.301 Loads.

29.303 Factor of safety.

Subpart B—Flight 29.305 Strength and deformation.

G ENERAL 29.307 Proof of structure.

29.309 Design limitations.

29.21 Proof of compliance.

29.25 Weight limits.

FLIGHT L OADS 29.27 Center of gravity limits.

29.29 Empty weight and corresponding cen- 29.321 General.

ter of gravity. 29.337 Limit maneuvering load factor.

29.31 Removable ballast. 29.339 Resultant limit maneuvering loads.