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(iii) The equipment, systems, and installa- 29.81—Landing distance (Ground level sites): tions must be designed so that one display of Category A. the information essential to the safety of 29.85—Balked landing: Category A. flight which is provided by the instruments 29.87(a)—Height-velocity envelope. will remain available to a pilot, without ad- 29.547(a) and (b)—Main and tail rotor struc- ditional crewmember action, after any single ture. failure or combination of failures that is not 29.861(a)—Fire protection of structure, con- shown to be extremely improbable; and trols, and other parts. (iv) For single-pilot configurations, instru- 29.901(c)—Powerplant: Installation. ments which require a static source must be 29.903(b) (c) and (e)—Engines. provided with a means of selecting an alter- 29.908(a)—Cooling fans. nate source and that source must be cali- 29.917(b) and (c)(1)—Rotor drive system: De- brated. sign. IX. Rotorcraft Flight Manual. A Rotorcraft 29.927(c)(1)—Additional tests. Flight Manual or Rotorcraft Flight Manual 29.953(a)—Fuel system independence. IFR Supplement must be provided and must 29.1027(a)—Transmission and gearboxes: Gen- contain— eral. (a) Limitations. The approved IFR flight en- 29.1045(a)(1), (b), (c), (d), and (f)—Climb cool- velope, the IFR flightcrew composition, the ing test procedures. revised kinds of operation, and the steepest 29.1047(a)—Takeoff cooling test procedures. IFR precision approach gradient for which 29.1181(a)—Designated fire zones: Regions in- the helicopter is approved; cluded. (b) Procedures. Required information for 29.1187(e)—Drainage and ventilation of fire proper operation of IFR systems and the rec- zones. ommended procedures in the event of sta- 29.1189(c)—Shutoff means. bility augmentation or electrical system 29.1191(a)(1)—Firewalls. failures; and 29.1193(e)—Cowling and engine compartment (c) Performance. If V differs from V , covering. YI Y 29.1195(a) and (d)—Fire extinguishing sys- climb performance at VYI and with maximum continuous power throughout the ranges of tems (one shot). weight, altitude, and temperature for which 29.1197—Fire extinguishing agents. 29.1199—Extinguishing agent containers. approval is requested. 29.1201—Fire extinguishing system materials. [Amdt. 27–19, 48 FR 4389, Jan. 31, 1983] 29.1305(a) (6) and (b)—Powerplant instru- ments. APPENDIX C TO PART 27—CRITERIA FOR 29.1309(b)(2) (i) and (d)—Equipment, systems, CATEGORY A and installations. 29.1323(c)(1)—Airspeed indicating system. C27.1 General. 29.1331(b)—Instruments using a power supply. A small multiengine rotorcraft may not be 29.1351(d)(2)—Electrical systems and equip- type certificated for Category A operation ment: General (operation without normal unless it meets the design installation and electrical power). performance requirements contained in this 29.1587(a)—Performance information. appendix in addition to the requirements of NOTE: In complying with the paragraphs this part. listed in paragraph C27.2 above, relevant ma- C27.2 Applicable part 29 sections. The fol- terial in the AC ‘‘Certification of Transport lowing sections of part 29 of this chapter Category Rotorcraft’’ should be used. must be met in addition to the requirements [Doc. No. 28008, 61 FR 21907, May 10, 1996] of this part: 29.45(a) and (b)(2)—General. 29.49(a)—Performance at minimum operating PART 29—AIRWORTHINESS STAND- speed. ARDS: TRANSPORT CATEGORY 29.51—Takeoff data: General. ROTORCRAFT 29.53—Takeoff: Category A. 29.55—Takeoff decision point: Category A. Subpart A—General 29.59—Takeoff Path: Category A. 29.60—Elevated heliport takeoff path: Cat- Sec. egory A. 29.1 Applicability. 29.61—Takeoff distance: Category A. 29.2 Special retroactive requirements. 29.62—Rejected takeoff: Category A. 29.64—Climb: General. Subpart B—Flight 29.65(a)—Climb: AEO. 29.67(a)—Climb: OEI. GENERAL 29.75—Landing: General. 29.21 Proof of compliance. 29.77—Landing decision point: Category A. 29.25 Weight limits. 29.79—Landing: Category A. 29.27 Center of gravity limits.

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29.29 Empty weight and corresponding cen- 29.337 Limit maneuvering load factor. ter of gravity. 29.339 Resultant limit maneuvering loads. 29.31 Removable ballast. 29.341 Gust loads. 29.33 Main rotor speed and pitch limits. 29.351 Yawing conditions. 29.361 Engine torque. PERFORMANCE 29.45 General. CONTROL SURFACE AND SYSTEM LOADS 29.49 Performance at minimum operating 29.391 General. speed. 29.395 Control system. 29.51 Takeoff data: general. 29.397 Limit pilot forces and torques. 29.53 Takeoff: Category A. 29.399 Dual control system. 29.55 Takeoff decision point (TDP): Cat- 29.411 Ground clearance: tail rotor guard. egory A. 29.427 Unsymmetrical loads. 29.59 Takeoff path: Category A. 29.60 Elevated heliport takeoff path: Cat- GROUND LOADS egory A. 29.61 Takeoff distance: Category A. 29.471 General. 29.62 Rejected takeoff: Category A. 29.473 Ground loading conditions and as- 29.63 Takeoff: Category B. sumptions. 29.64 Climb: General. 29.475 Tires and shock absorbers. 29.65 Climb: All engines operating. 29.477 Landing gear arrangement. 29.67 Climb: One engine inoperative (OEI). 29.479 Level landing conditions. 29.71 Helicopter angle of glide: Category B. 29.481 Tail-down landing conditions. 29.75 Landing: General. 29.483 One-wheel landing conditions. 29.77 Landing Decision Point (LDP): Cat- 29.485 Lateral drift landing conditions. egory A. 29.493 Braked roll conditions. 29.79 Landing: Category A. 29.497 Ground loading conditions: landing 29.81 Landing distance: Category A. gear with tail wheels. 29.83 Landing: Category B. 29.501 Ground loading conditions: landing 29.85 Balked landing: Category A. gear with skids. 29.87 Height-velocity envelope. 29.505 Ski landing conditions. 29.511 Ground load: unsymmetrical loads on FLIGHT CHARACTERISTICS multiple-wheel units. 29.141 General. 29.143 Controllability and maneuverability. WATER LOADS 29.151 Flight controls. 29.519 Hull type rotorcraft: Water-based and 29.161 Trim control. amphibian. 29.171 Stability: general. 29.521 Float landing conditions. 29.173 Static longitudinal stability. 29.175 Demonstration of static longitudinal MAIN COMPONENT REQUIREMENTS stability. 29.177 Static directional stability. 29.547 Main and tail rotor structure. 29.181 Dynamic stability: Category A rotor- 29.549 Fuselage and rotor pylon structures. craft. 29.551 Auxiliary lifting surfaces.

GROUND AND WATER HANDLING EMERGENCY LANDING CONDITIONS CHARACTERISTICS 29.561 General. 29.231 General. 29.562 Emergency landing dynamic condi- 29.235 Taxiing condition. tions. 29.239 Spray characteristics. 29.563 Structural ditching provisions. 29.241 Ground resonance. FATIGUE EVALUATION MISCELLANEOUS FLIGHT REQUIREMENTS 29.571 Fatigue evaluation of structure. 29.251 Vibration. Subpart D—Design and Construction Subpart C—Strength Requirements GENERAL GENERAL 29.601 Design. 29.301 Loads. 29.602 Critical parts. 29.303 Factor of safety. 29.603 Materials. 29.305 Strength and deformation. 29.605 Fabrication methods. 29.307 Proof of structure. 29.607 Fasteners. 29.309 Design limitations. 29.609 Protection of structure. 29.610 Lightning and static electricity pro- FLIGHT LOADS tection. 29.321 General. 29.611 Inspection provisions.

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29.613 Material strength properties and de- 29.815 Main aisle width. sign values. 29.831 Ventilation. 29.619 Special factors. 29.833 Heaters. 29.621 Casting factors. 29.623 Bearing factors. FIRE PROTECTION 29.625 Fitting factors. 29.851 Fire extinguishers. 29.629 Flutter and divergence. 29.853 Compartment interiors. 29.631 Bird strike. 29.855 Cargo and baggage compartments. 29.859 Combustion heater fire protection. ROTORS 29.861 Fire protection of structure, controls, 29.653 Pressure venting and drainage of and other parts. rotor blades. 29.863 Flammable fluid fire protection. 29.659 Mass balance. 29.661 Rotor blade clearance. EXTERNAL LOADS 29.663 Ground resonance prevention means. 29.865 External loads. CONTROL SYSTEMS MISCELLANEOUS 29.671 General. 29.871 Leveling marks. 29.672 Stability augmentation, automatic, 29.873 Ballast provisions. and power-operated systems. 29.673 Primary flight controls. Subpart E—Powerplant 29.674 Interconnected controls. 29.675 Stops. GENERAL 29.679 Control system locks. 29.681 Limit load static tests. 29.901 Installation. 29.683 Operation tests. 29.903 Engines. 29.685 Control system details. 29.907 Engine vibration. 29.687 Spring devices. 29.908 Cooling fans. 29.691 Autorotation control mechanism. ROTOR DRIVE SYSTEM 29.695 Power boost and power-operated con- trol system. 29.917 Design. 29.921 Rotor brake. LANDING GEAR 29.923 Rotor drive system and control mech- 29.723 Shock absorption tests. anism tests. 29.725 Limit drop test. 29.927 Additional tests. 29.727 Reserve energy absorption drop test. 29.931 Shafting critical speed. 29.729 Retracting mechanism. 29.935 Shafting joints. 29.731 Wheels. 29.939 Turbine engine operating characteris- 29.733 Tires. tics. 29.735 Brakes. FUEL SYSTEM 29.737 Skis. 29.951 General. FLOATS AND HULLS 29.952 Fuel system crash resistance. 29.751 Main float buoyancy. 29.953 Fuel system independence. 29.753 Main float design. 29.954 Fuel system lightning protection. 29.755 Hull buoyancy. 29.955 Fuel flow. 29.757 Hull and auxiliary float strength. 29.957 Flow between interconnected tanks. 29.959 Unusable fuel supply. PERSONNEL AND CARGO ACCOMMODATIONS 29.961 Fuel system hot weather operation. 29.771 Pilot compartment. 29.963 Fuel tanks: general. 29.773 Pilot compartment view. 29.965 Fuel tank tests. 29.775 Windshields and windows. 29.967 Fuel tank installation. 29.777 Cockpit controls. 29.969 Fuel tank expansion space. 29.779 Motion and effect of cockpit controls. 29.971 Fuel tank sump. 29.783 Doors. 29.973 Fuel tank filler connection. 29.785 Seats, berths, litters, safety belts, 29.975 Fuel tank vents and carburetor vapor and harnesses. vents. 29.787 Cargo and baggage compartments. 29.977 Fuel tank outlet. 29.801 Ditching. 29.979 Pressure refueling and fueling provi- 29.803 Emergency evacuation. sions below fuel level. 29.805 Flight crew emergency exits. FUEL SYSTEM COMPONENTS 29.807 Passenger emergency exits. 29.809 Emergency exit arrangement. 29.991 Fuel pumps. 29.811 Emergency exit marking. 29.993 Fuel system lines and fittings. 29.812 Emergency lighting. 29.995 Fuel valves. 29.813 Emergency exit access. 29.997 Fuel strainer or filter.

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29.999 Fuel system drains. 29.1203 Fire detector systems. 29.1001 Fuel jettisoning. Subpart F—Equipment OIL SYSTEM GENERAL 29.1011 Engines: general. 29.1013 Oil tanks. 29.1301 Function and installation. 29.1015 Oil tank tests. 29.1303 Flight and navigation instruments. 29.1017 Oil lines and fittings. 29.1305 Powerplant instruments. 29.1019 Oil strainer or filter. 29.1307 Miscellaneous equipment. 29.1021 Oil system drains. 29.1309 Equipment, systems, and installa- 29.1023 Oil radiators. tions. 29.1025 Oil valves. 29.1027 Transmission and gearboxes: gen- INSTRUMENTS: INSTALLATION eral. 29.1321 Arrangement and visibility. 29.1322 Warning, caution, and advisory COOLING lights. 29.1041 General. 29.1323 Airspeed indicating system. 29.1043 Cooling tests. 29.1325 Static pressure and pressure altim- 29.1045 Climb cooling test procedures. eter systems. 29.1047 Takeoff cooling test procedures. 29.1327 Magnetic direction indicator. 29.1049 Hovering cooling test procedures. 29.1329 Automatic pilot system. 29.1331 Instruments using a power supply. INDUCTION SYSTEM 29.1333 Instrument systems. 29.1091 Air induction. 29.1335 Flight director systems. 29.1093 Induction system icing protection. 29.1337 Powerplant instruments. 29.1101 Carburetor air preheater design. ELECTRICAL SYSTEMS AND EQUIPMENT 29.1103 Induction systems ducts and air duct systems. 29.1351 General. 29.1105 Induction system screens. 29.1353 Electrical equipment and installa- 29.1107 Inter-coolers and after-coolers. tions. 29.1109 Carburetor air cooling. 29.1355 Distribution system. 29.1357 Circuit protective devices. EXHAUST SYSTEM 29.1359 Electrical system fire and smoke 29.1121 General. protection. 29.1123 Exhaust piping. 29.1363 Electrical system tests. 29.1125 Exhaust heat exchangers. LIGHTS POWERPLANT CONTROLS AND ACCESSORIES 29.1381 Instrument lights. 29.1141 Powerplant controls: general. 29.1383 Landing lights. 29.1142 Auxiliary power unit controls. 29.1385 Position light system installation. 29.1143 Engine controls. 29.1387 Position light system dihedral an- 29.1145 Ignition switches. gles. 29.1147 Mixture controls. 29.1389 Position light distribution and in- 29.1151 Rotor brake controls. tensities. 29.1157 Carburetor air temperature controls. 29.1391 Minimum intensities in the hori- 29.1159 Supercharger controls. zontal plane of forward and rear position 29.1163 Powerplant accessories. lights. 29.1165 Engine ignition systems. 29.1393 Minimum intensities in any vertical plane of forward and rear position lights. POWERPLANT FIRE PROTECTION 29.1395 Maximum intensities in overlapping beams of forward and rear position 29.1181 Designated fire zones: regions in- lights. cluded. 29.1397 Color specifications. 29.1183 Lines, fittings, and components. 29.1399 Riding light. 29.1185 Flammable fluids. 29.1401 Anticollision light system. 29.1187 Drainage and ventilation of fire zones. SAFETY EQUIPMENT 29.1189 Shutoff means. 29.1191 Firewalls. 29.1411 General. 29.1193 Cowling and engine compartment 29.1413 Safety belts: passenger warning de- covering. vice. 29.1194 Other surfaces. 29.1415 Ditching equipment. 29.1195 Fire extinguishing systems. 29.1419 Ice protection. 29.1197 Fire extinguishing agents. MISCELLANEOUS EQUIPMENT 29.1199 Extinguishing agent containers. 29.1201 Fire extinguishing system materials. 29.1431 Electronic equipment.

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29.1433 Vacuum systems. Subpart A—General 29.1435 Hydraulic systems. 29.1439 Protective breathing equipment. § 29.1 Applicability. 29.1457 Cockpit voice recorders. (a) This part prescribes airworthiness 29.1459 Flight recorders. standards for the issue of type certifi- 29.1461 Equipment containing high energy rotors. cates, and changes to those certifi- cates, for transport category rotor- Subpart G—Operating Limitations and craft. Information (b) Transport category rotorcraft must be certificated in accordance 29.1501 General. with either the Category A or Category B requirements of this part. A multien- OPERATING LIMITATIONS gine rotorcraft may be type certifi- 29.1503 Airspeed limitations: general. cated as both Category A and Category 29.1505 Never-exceed speed. B with appropriate and different oper- 29.1509 Rotor speed. ating limitations for each category. 29.1517 Limiting height-speed envelope. (c) Rotorcraft with a maximum 29.1519 Weight and center of gravity. weight greater than 20,000 pounds and 29.1521 Powerplant limitations. 10 or more passenger seats must be 29.1522 Auxiliary power unit limitations. type certificated as Category A rotor- 29.1523 Minimum flight crew. craft. 29.1525 Kinds of operations. (d) Rotorcraft with a maximum 29.1527 Maximum operating altitude. weight greater than 20,000 pounds and 29.1529 Instructions for Continued Air- nine or less passenger seats may be worthiness. type certificated as Category B rotor- MARKINGS AND PLACARDS craft provided the Category A require- ments of Subparts C, D, E, and F of 29.1541 General. this part are met. 29.1543 Instrument markings: general. (e) Rotorcraft with a maximum 29.1545 Airspeed indicator. weight of 20,000 pounds or less but with 29.1547 Magnetic direction indicator. 10 or more passenger seats may be type 29.1549 Powerplant instruments. certificated as Category B rotorcraft 29.1551 Oil quantity indicator. provided the Category A requirements 29.1553 Fuel quantity indicator. of §§ 29.67(a)(2), 29.87, 29.1517, and sub- 29.1555 Control markings. parts C, D, E, and F of this part are 29.1557 Miscellaneous markings and plac- ards. met. 29.1559 Limitations placard. (f) Rotorcraft with a maximum 29.1561 Safety equipment. weight of 20,000 pounds or less and nine 29.1565 Tail rotor. or less passenger seats may be type certificated as Category B rotorcraft. ROTORCRAFT FLIGHT MANUAL (g) Each person who applies under 29.1581 General. Part 21 for a certificate or change de- 29.1583 Operating limitations. scribed in paragraphs (a) through (f) of 29.1585 Operating procedures. this section must show compliance 29.1587 Performance information. with the applicable requirements of 29.1589 Loading information. this part. APPENDIX A TO PART 29—INSTRUCTIONS FOR [Amdt. 29–21, 48 FR 4391, Jan. 31, 1983, as CONTINUED AIRWORTHINESS amended by Amdt. 29–39, 61 FR 21898, May 10, APPENDIX B TO PART 29—AIRWORTHINESS CRI- 1996; 61 FR 33963, July 1, 1996] TERIA FOR HELICOPTER INSTRUMENT FLIGHT § 29.2 Special retroactive require- APPENDIX C TO PART 29—ICING CERTIFICATION ments. APPENDIX D TO PART 29—CRITERIA FOR DEM- For each rotorcraft manufactured ONSTRATION OF EMERGENCY EVACUATION after September 16, 1992, each applicant PROCEDURES UNDER §29.803 must show that each occupant’s seat is AUTHORITY: 49 U.S.C. 106(g), 40113, 44701– equipped with a safety belt and shoul- 44702, 44704. der harness that meets the require- SOURCE: Docket No. 5084, 29 FR 16150, Dec. ments of paragraphs (a), (b), and (c) of 3, 1964, unless otherwise noted. this section.

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(a) Each occupant’s seat must have a cannot be reasonably inferred from combined safety belt and shoulder har- combinations investigated. ness with a single-point release. Each [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as pilot’s combined safety belt and shoul- amended by Amdt. 29–24, 49 FR 44435, Nov. 6, der harness must allow each pilot, 1984] when seated with safety belt and shoul- der harness fastened, to perform all § 29.25 Weight limits. functions necessary for flight oper- (a) Maximum weight. The maximum ations. There must be a means to se- weight (the highest weight at which cure belts and harnesses, when not in compliance with each applicable re- use, to prevent interference with the quirement of this part is shown) or, at operation of the rotorcraft and with the option of the applicant, the highest rapid egress in an emergency. weight for each altitude and for each (b) Each occupant must be protected practicably separable operating condi- from serious head injury by a safety tion, such as takeoff, enroute oper- belt plus a shoulder harness that will ation, and landing, must be established prevent the head from contacting any so that it is not more than— (1) The highest weight selected by injurious object. the applicant; (c) The safety belt and shoulder har- (2) The design maximum weight (the ness must meet the static and dynamic highest weight at which compliance strength requirements, if applicable, with each applicable structural loading specified by the rotorcraft type certifi- condition of this part is shown); or cation basis. (3) The highest weight at which com- (d) For purposes of this section, the pliance with each applicable flight re- date of manufacture is either— quirement of this part is shown. (1) The date the inspection accept- (b) Minimum weight. The minimum ance records, or equivalent, reflect weight (the lowest weight at which that the rotorcraft is complete and compliance with each applicable re- meets the FAA-Approved Type Design quirement of this part is shown) must Data; or be established so that it is not less (2) The date that the foreign civil air- than— worthiness authority certifies the (1) The lowest weight selected by the rotorcraft is complete and issues an applicant; original standard airworthiness certifi- (2) The design minimum weight (the cate, or equivalent, in that country. lowest weight at which compliance with each structural loading condition [Doc. No. 26078, 56 FR 41052, Aug. 16, 1991] of this part is shown); or (3) The lowest weight at which com- Subpart B—Flight pliance with each applicable flight re- quirement of this part is shown. GENERAL (c) Total weight with jettisonable exter- nal load. A total weight for the rotor- § 29.21 Proof of compliance. craft with a jettisonable external load Each requirement of this subpart attached that is greater than the max- must be met at each appropriate com- imum weight established under para- graph (a) of this section may be estab- bination of weight and center of grav- lished for any rotorcraft-load combina- ity within the range of loading condi- tion if— tions for which certification is re- (1) The rotorcraft-load combination quested. This must be shown— does not include human external cargo, (a) By tests upon a rotorcraft of the (2) Structural component approval type for which certification is re- for external load operations under ei- quested, or by calculations based on, ther § 29.865 or under equivalent oper- and equal in accuracy to, the results of ational standards is obtained, testing; and (3) The portion of the total weight (b) By systematic investigation of that is greater than the maximum each required combination of weight weight established under paragraph (a) and center of gravity, if compliance of this section is made up only of the

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weight of all or part of the jettisonable with respect to the weights of fuel, oil, external load, coolant, and installed equipment. (4) Structural components of the (Secs. 313(a), 601, 603, 604, and 605 of the Fed- rotorcraft are shown to comply with eral Aviation Act of 1958 (49 U.S.C. 1354(a), the applicable structural requirements 1421, 1423, 1424, and 1425); and sec. 6(c) of the of this part under the increased loads Dept. of Transportation Act (49 U.S.C. and stresses caused by the weight in- 1655(c))) crease over that established under [Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as paragraph (a) of this section, and amended by Amdt. 29–15, 43 FR 2326, Jan. 16, (5) Operation of the rotorcraft at a 1978] total weight greater than the max- § 29.31 Removable ballast. imum certificated weight established under paragraph (a) of this section is Removable ballast may be used in limited by appropriate operating limi- showing compliance with the flight re- tations under § 29.865 (a) and (d) of this quirements of this subpart. part. § 29.33 Main rotor speed and pitch lim- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as its. amended by Amdt. 29–12, 41 FR 55471, Dec. 20, (a) Main rotor speed limits. A range of 1976; Amdt. 29–43, 64 FR 43020, Aug. 6, 1999] main rotor speeds must be established that— § 29.27 Center of gravity limits. (1) With power on, provides adequate The extreme forward and aft centers margin to accommodate the variations of gravity and, where critical, the ex- in rotor speed occurring in any appro- treme lateral centers of gravity must priate maneuver, and is consistent be established for each weight estab- with the kind of governor or synchro- lished under § 29.25. Such an extreme nizer used; and may not lie beyond— (2) With power off, allows each appro- (a) The extremes selected by the ap- priate autorotative maneuver to be plicant; performed throughout the ranges of (b) The extremes within which the airspeed and weight for which certifi- structure is proven; or cation is requested. (b) Normal main rotor high pitch limit (c) The extremes within which com- (power on). For rotorcraft, except heli- pliance with the applicable flight re- copters required to have a main rotor quirements is shown. low speed warning under paragraph (e) [Amdt. 29–3, 33 FR 965, Jan. 26, 1968] of this section, it must be shown, with power on and without exceeding ap- § 29.29 Empty weight and cor- proved engine maximum limitations, responding center of gravity. that main rotor speeds substantially (a) The empty weight and cor- less than the minimum approved main responding center of gravity must be rotor speed will not occur under any determined by weighing the rotorcraft sustained flight condition. This must without the crew and payload, but be met by— with— (1) Appropriate setting of the main (1) Fixed ballast; rotor high pitch stop; (2) Inherent rotorcraft characteris- (2) Unusable fuel; and tics that make unsafe low main rotor (3) Full operating fluids, including— speeds unlikely; or (i) Oil; (3) Adequate means to warn the pilot (ii) Hydraulic fluid; and of unsafe main rotor speeds. (iii) Other fluids required for normal (c) Normal main rotor low pitch limit operation of rotorcraft systems, except (power off). It must be shown, with water intended for injection in the en- power off, that— gines. (1) The normal main rotor low pitch (b) The condition of the rotorcraft at limit provides sufficient rotor speed, in the time of determining empty weight any autorotative condition, under the must be one that is well defined and most critical combinations of weight can be easily repeated, particularly and airspeed; and

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(2) It is possible to prevent over- (2) Without exceptionally favorable speeding of the rotor without excep- conditions. tional piloting skill. (b) Compliance with the performance (d) Emergency high pitch. If the main requirements of this subpart must be rotor high pitch stop is set to meet shown— paragraph (b)(1) of this section, and if (1) For still air at sea level with a that stop cannot be exceeded inadvert- standard atmosphere and; ently, additional pitch may be made (2) For the approved range of atmos- available for emergency use. pheric variables. (e) Main rotor low speed warning for (c) The available power must cor- helicopters. For each single engine respond to engine power, not exceeding helicopter, and each multiengine heli- the approved power, less— copter that does not have an approved (1) Installation losses; and device that automatically increases (2) The power absorbed by the acces- power on the operating engines when sories and services at the values for one engine fails, there must be a main which certification is requested and ap- rotor low speed warning which meets proved. the following requirements: (d) For reciprocating engine-powered (1) The warning must be furnished to rotorcraft, the performance, as affected the pilot in all flight conditions, in- by engine power, must be based on a cluding power-on and power-off flight, relative humidity of 80 percent in a when the speed of a main rotor ap- standard atmosphere. proaches a value that can jeopardize (e) For turbine engine-powered rotor- safe flight. craft, the performance, as affected by (2) The warning may be furnished ei- engine power, must be based on a rel- ther through the inherent aerodynamic ative humidity of— qualities of the helicopter or by a de- (1) 80 percent, at and below standard vice. temperature; and (3) The warning must be clear and (2) 34 percent, at and above standard distinct under all conditions, and must temperature plus 50 °F. be clearly distinguishable from all Between these two temperatures, the other warnings. A visual device that relative humidity must vary linearly. requires the attention of the crew (f) For turbine-engine-power rotor- within the cockpit is not acceptable by craft, a means must be provided to per- itself. mit the pilot to detemine prior to take- (4) If a warning device is used, the de- off that each engine is capable of devel- vice must automatically deactivate oping the power necessary to achieve and reset when the low-speed condition the applicable rotorcraft performance is corrected. If the device has an audi- prescribed in this subpart. ble warning, it must also be equipped with a means for the pilot to manually (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), silence the audible warning before the 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. low-speed condition is corrected. of Transportation Act (49 U.S.C. 1655(c))) (Secs. 313(a), 601, 603, 604, and 605 of the Fed- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as eral Aviation Act of 1958 (49 U.S.C. 1354(a), amended by Amdt. 29–15, 43 FR 2326, Jan. 16, 1421, 1423, 1424, and 1425); and sec. 6(c) of the 1978; Amdt. 29–24, 49 FR 44436, Nov. 6, 1984] Dept. of Transportation Act (49 U.S.C. 1655(c))) § 29.49 Performance at minimum oper- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ating speed. amended by Amdt. 29–3, 33 FR 965, Jan. 26, (a) For each Category A helicopter, 1968; Amdt. 29–15, 43 FR 2326, Jan. 16, 1978] the hovering performance must be de- termined over the ranges of weight, al- PERFORMANCE titude, and temperature for which takeoff data are scheduled— § 29.45 General. (1) With not more than takeoff (a) The performance prescribed in power; this subpart must be determined— (2) With the landing gear extended; (1) With normal piloting skill and; and

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(3) At a height consistent with the (a) Return to, and stop safely on, the procedure used in establishing the takeoff area; or takeoff, climbout, and rejected takeoff (b) Continue the takeoff and paths. climbout, and attain a configuration (b) For each Category B helicopter, and airspeed allowing compliance with the hovering performance must be de- § 29.67(a)(2). termined over the ranges of weight, al- titude, and temperature for which cer- [Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 tification is requested, with— FR 33963, July 1, 1996] (1) Takeoff power; (2) The landing gear extended; and § 29.55 Takeoff decision point (TDP): Category A. (3) The helicopter in ground effect at a height consistent with normal take- (a) The TDP is the first point from off procedures. which a continued takeoff capability is (c) For each helicopter, the out-of- assured under § 29.59 and is the last ground effect hovering performance point in the takeoff path from which a must be determined over the ranges of rejected takeoff is assured within the weight, altitude, and temperature for distance determined under §29.62. which certification is requested with (b) The TDP must be established in takeoff power. relation to the takeoff path using no (d) For rotorcraft other than heli- more than two parameters; e.g., air- copters, the steady rate of climb at the speed and height, to designate the minimum operating speed must be de- TDP. termined over the ranges of weight, al- (c) Determination of the TDP must titude, and temperature for which cer- include the pilot recognition time in- tification is requested with— terval following failure of the critical (1) Takeoff power; and engine. (2) The landing gear extended. [Doc. No. 24802, 61 FR 21899, May 10, 1996] [Doc. No. 24802, 61 FR 21898, May 10, 1996; 61 FR 33963, July 1, 1996] § 29.59 Takeoff path: Category A. § 29.51 Takeoff data: general. (a) The takeoff path extends from the (a) The takeoff data required by point of commencement of the takeoff §§ 29.53, 29.55, 29.59, 29.60, 29.61, 29.62, procedure to a point at which the 29.63, and 29.67 must be determined— rotorcraft is 1,000 feet above the take- (1) At each weight, altitude, and tem- off surface and compliance with perature selected by the applicant; and § 29.67(a)(2) is shown. In addition— (2) With the operating engines within (1) The takeoff path must remain approved operating limitations. clear of the height-velocity envelope (b) Takeoff data must— established in accordance with § 29.87; (1) Be determined on a smooth, dry, (2) The rotorcraft must be flown to hard surface; and the engine failure point; at which (2) Be corrected to assume a level point, the critical engine must be made takeoff surface. inoperative and remain inoperative for (c) No takeoff made to determine the the rest of the takeoff; data required by this section may re- (3) After the critical engine is made quire exceptional piloting skill or inoperative, the rotorcraft must con- alertness, or exceptionally favorable tinue to the takeoff decision point, and conditions. then attain VTOSS; [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (4) Only primary controls may be amended by Amdt. 29–39, 61 FR 21899, May 10, used while attaining VTOSS and while 1996] establishing a positive rate of climb. Secondary controls that are located on § 29.53 Takeoff: Category A. the primary controls may be used after The takeoff performance must be de- a positive rate of climb and VTOSS are termined and scheduled so that, if one established but in no case less than 3 engine fails at any time after the start seconds after the critical engine is of takeoff, the rotorcraft can— made inoperative; and

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(5) After attaining VTOSS and a posi- (b) The scheduled takeoff weight tive rate of a climb, the landing gear must be such that the climb require- may be retracted. ments of § 29.67 (a)(1) and (a)(2) will be (b) During the takeoff path deter- met. mination made in accordance with (c) Takeoff distance will be deter- paragraph (a) of this section and after mined in accordance with § 29.61. attaining V and a positive rate of TOSS [Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 climb, the climb must be continued at FR 33963, July 1, 1996] a speed as close as practicable to, but not less than, VTOSS until the rotorcraft § 29.61 Takeoff distance: Category A. is 200 feet above the takeoff surface. During this interval, the climb per- (a) The normal takeoff distance is formance must meet or exceed that re- the horizontal distance along the take- quired by § 29.67(a)(1). off path from the start of the takeoff to (c) During the continued takeoff, the the point at which the rotorcraft at- rotorcraft shall not descend below 15 tains and remains at least 35 feet above feet above the takeoff surface when the the takeoff surface, attains and main- takeoff decision point is above 15 feet. tains a speed of at least VTOSS, and es- tablishes a positive rate of climb, as- (d) From 200 feet above the takeoff suming the critical engine failure oc- surface, the rotorcraft takeoff path curs at the engine failure point prior to must be level or positive until a height the takeoff decision point. 1,000 feet above the takeoff surface is attained with not less than the rate of (b) For elevated heliports, the take- climb required by § 29.67(a)(2). Any sec- off distance is the horizontal distance ondary or auxiliary control may be along the takeoff path from the start used after attaining 200 feet above the of the takeoff to the point at which the takeoff surface. rotorcraft attains and maintains a (e) Takeoff distance will be deter- speed of at least VTOSS and establishes a mined in accordance with § 29.61. positive rate of climb, assuming the critical engine failure occurs at the en- [Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 gine failure point prior to the takeoff FR 33963, July 1, 1996, as amended by Amdt. decision point. 29–44, 64 FR 45337, Aug. 19, 1999] [Doc. No. 24802, 61 FR 21899, May 10, 1996] § 29.60 Elevated heliport takeoff path: Category A. § 29.62 Rejected takeoff: Category A. (a) The elevated heliport takeoff path The rejected takeoff distance and extends from the point of commence- procedures for each condition where ment of the takeoff procedure to a takeoff is approved will be established point in the takeoff path at which the with— rotorcraft is 1,000 feet above the take- (a) The takeoff path requirements of off surface and compliance with §§ 29.59 and 29.60 being used up to the § 29.67(a)(2) is shown. In addition— TDP where the critical engine failure (1) The requirements of § 29.59(a) is recognized and the rotorcraft is land- must be met; ed and brought to a complete stop on

(2) While attaining VTOSS and a posi- the takeoff surface; tive rate of climb, the rotorcraft may (b) The remaining engines operating descend below the level of the takeoff within approved limits; surface if, in so doing and when clear- (c) The landing gear remaining ex- ing the elevated heliport edge, every tended throughout the entire rejected part of the rotorcraft clears all obsta- takeoff; and cles by at least 15 feet; (d) The use of only the primary con- (3) The vertical magnitude of any de- trols until the rotorcraft is on the scent below the takeoff surface must be ground. Secondary controls located on determined; and the primary control may not be used (4) After attaining VTOSS and a posi- until the rotorcraft is on the ground. tive rate of climb, the landing gear Means other than wheel brakes may be may be retracted. used to stop the rotorcraft if the means

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are safe and reliable and consistent re- gradient of at least 1:6 under standard sults can be expected under normal op- sea level conditions. erating conditions. (Secs. 313(a), 601, 603, 604, and 605 of the Fed- [Doc. No. 24802, 61 FR 21899, May 10, 1996, as eral Aviation Act of 1958 (49 U.S.C. 1354(a), amended by Amdt. 29–44, 64 FR 45337, Aug. 19, 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. 1999] of Transportation Act (49 U.S.C. 1655(c))) [Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as § 29.63 Takeoff: Category B. amended by Amdt. 29–15, 43 FR 2326, Jan. 16, 1978; Amdt. 29–39, 61 FR 21900, May 10, 1996; 61 The horizontal distance required to FR 33963, July 1, 1996] take off and climb over a 50-foot obsta- cle must be established with the most § 29.67 Climb: One engine inoperative unfavorable center of gravity. The (OEI). takeoff may be begun in any manner (a) For Category A rotorcraft, in the if— critical takeoff configuration existing (a) The takeoff surface is defined; along the takeoff path, the following (b) Adequate safeguards are main- apply: tained to ensure proper center of grav- (1) The steady rate of climb without ity and control positions; and ground effect, 200 feet above the take- (c) A landing can be made safely at off surface, must be at least 100 feet per any point along the flight path if an minute for each weight, altitude, and engine fails. temperature for which takeoff data are to be scheduled with— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (i) The critical engine inoperative amended by Amdt. 29–12, 41 FR 55471, Dec. 20, and the remaining engines within ap- 1976] proved operating limitations, except that for rotorcraft for which the use of § 29.64 Climb: General. 30-second/2-minute OEI power is re- Compliance with the requirements of quested, only the 2-minute OEI power §§ 29.65 and 29.67 must be shown at each may be used in showing compliance weight, altitude, and temperature with this paragraph; within the operational limits estab- (ii) The landing gear extended; and lished for the rotorcraft and with the (iii) The takeoff safety speed selected most unfavorable center of gravity for by the applicant. each configuration. Cowl flaps, or other (2) The steady rate of climb without means of controlling the engine-cool- ground effect, 1000 feet above the take- ing air supply, will be in the position off surface, must be at least 150 feet per that provides adequate cooling at the minute, for each weight, altitude, and temperature for which takeoff data are temperatures and altitudes for which to be scheduled with— certification is requested. (i) The critical engine inoperative [Doc. No. 24802, 61 FR 21900, May 10, 1996] and the remaining engines at max- imum continuous power including con- § 29.65 Climb: All engines operating. tinuous OEI power, if approved, or at (a) The steady rate of climb must be 30-minute OEI power for rotorcraft for determined— which certification for use of 30-minute OEI power is requested; (1) With maximum continuous power; (ii) The landing gear retracted; and (2) With the landing gear retracted; (iii) The speed selected by the appli- and cant. (3) At Vy for standard sea level condi- (3) The steady rate of climb (or de- tions and at speeds selected by the ap- scent) in feet per minute, at each alti- plicant for other conditions. tude and temperature at which the (b) For each Category B rotorcraft rotorcraft is expected to operate and at except helicopters, the rate of climb any weight within the range of weights determined under paragraph (a) of this for which certification is requested, section must provide a steady climb must be determined with—

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(i) The critical engine inoperative tendency to bounce, nose over, ground and the remaining engines at max- loop, porpoise, or water loop. imum continuous power including con- (b) The landing data required by tinuous OEI power, if approved, and at §§ 29.77, 29.79, 29.81, 29.83, and 29.85 must 30-minute OEI power for rotorcraft for be determined— which certification for the use of 30- (1) At each weight, altitude, and tem- minute OEI power is requested; perature for which landing data are ap- (ii) The landing gear retracted; and proved; (iii) The speed selected by the appli- (2) With each operating engine within cant. (b) For multiengine Category B approved operating limitations; and rotorcraft meeting the Category A en- (3) With the most unfavorable center gine isolation requirements, the steady of gravity. rate of climb (or descent) must be de- [Doc. No. 24802, 61 FR 21900, May 10, 1996] termined at the speed for best rate of climb (or minimum rate of descent) at § 29.77 Landing Decision Point (LDP): each altitude, temperature, and weight Category A. at which the rotorcraft is expected to operate, with the critical engine inop- (a) The LDP is the last point in the erative and the remaining engines at approach and landing path from which maximum continuous power including a balked landing can be accomplished continuous OEI power, if approved, and in accordance with § 29.85. at 30-minute OEI power for rotorcraft (b) Determination of the LDP must for which certification for the use of 30- include the pilot recognition time in- minute OEI power is requested. terval following failure of the critical engine. [Doc. No. 24802, 61 FR 21900, May 10, 1996; 61 FR 33963, July 1, 1996, as amended by Amdt. [Doc. No. 24802, 64 FR 45338, Aug. 19, 1999] 29–44, 64 FR 45337, Aug. 19, 1999; 64 FR 47563, Aug. 31, 1999] § 29.79 Landing: Category A. § 29.71 Helicopter angle of glide: Cat- (a) For Category A rotorcraft— egory B. (1) The landing performance must be For each category B helicopter, ex- determined and scheduled so that if the cept multiengine helicopters meeting critical engine fails at any point in the the requirements of § 29.67(b) and the approach path, the rotorcraft can ei- powerplant installation requirements ther land and stop safely or climb out of category A, the steady angle of glide and attain a rotorcraft configuration must be determined in autorotation— and speed allowing compliance with (a) At the forward speed for min- the climb requirement of § 29.67(a)(2); imum rate of descent as selected by the (2) The approach and landing paths applicant; must be established with the critical (b) At the forward speed for best glide engine inoperative so that the transi- angle; tion between each stage can be made (c) At maximum weight; and smoothly and safely; (d) At the rotor speed or speeds se- (3) The approach and landing speeds lected by the applicant. must be selected by the applicant and [Amdt. 29–12, 41 FR 55471, Dec. 20, 1976] must be appropriate to the type of rotorcraft; and § 29.75 Landing: General. (4) The approach and landing path (a) For each rotorcraft— must be established to avoid the crit- (1) The corrected landing data must ical areas of the height-velocity enve- be determined for a smooth, dry, hard, lope determined in accordance with and level surface; § 29.87. (2) The approach and landing must (b) It must be possible to make a safe not require exceptional piloting skill landing on a prepared landing surface or exceptionally favorable conditions; after complete power failure occurring and during normal cruise. (3) The landing must be made with- out excessive vertical acceleration or [Doc. No. 24802, 61 FR 21900, May 10, 1996]

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§ 29.81 Landing distance: Category A. clearance of § 29.60 is maintained and The horizontal distance required to the descent (loss of height) below the land and come to a complete stop (or to landing surface is determined. a speed of approximately 3 knots for [Doc. No. 24802, 64 FR 45338, Aug. 19, 1999] water landings) from a point 50 ft above the landing surface must be de- § 29.87 Height-velocity envelope. termined from the approach and land- (a) If there is any combination of ing paths established in accordance height and forward velocity (including with § 29.79. hover) under which a safe landing can- [Doc. No. 24802, 64 FR 45338, Aug. 19, 1999] not be made after failure of the critical engine and with the remaining engines § 29.83 Landing: Category B. (where applicable) operating within ap- (a) For each Category B rotorcraft, proved limits, a height-velocity enve- the horizontal distance required to lope must be established for— land and come to a complete stop (or to (1) All combinations of pressure alti- a speed of approximately 3 knots for tude and ambient temperature for water landings) from a point 50 feet which takeoff and landing are ap- above the landing surface must be de- proved; and termined with— (2) Weight from the maximum weight (1) Speeds appropriate to the type of (at sea level) to the highest weight ap- rotorcraft and chosen by the applicant proved for takeoff and landing at each to avoid the critical areas of the altitude. For helicopters, this weight height-velocity envelope established need not exceed the highest weight al- under § 29.87; and lowing hovering out-of-ground effect at (2) The approach and landing made each altitude. with power on and within approved (b) For single-engine or multiengine limits. rotorcraft that do not meet the Cat- (b) Each multiengined Category B egory A engine isolation requirements, rotorcraft that meets the powerplant the height-velocity envelope for com- installation requirements for Category plete power failure must be estab- A must meet the requirements of— lished. (1) Sections 29.79 and 29.81; or [Doc. No. 24802, 61 FR 21901, May 10, 1996; 61 (2) Paragraph (a) of this section. FR 33963, July 1, 1996] (c) It must be possible to make a safe landing on a prepared landing surface if FLIGHT CHARACTERISTICS complete power failure occurs during normal cruise. § 29.141 General. [Doc. No. 24802, 61 FR 21900, May 10, 1996; 61 The rotorcraft must— FR 33963, July 1, 1996] (a) Except as specifically required in the applicable section, meet the flight § 29.85 Balked landing: Category A. characteristics requirements of this For Category A rotorcraft, the subpart— balked landing path with the critical (1) At the approved operating alti- engine inoperative must be established tudes and temperatures; so that— (2) Under any critical loading condi- (a) The transition from each stage of tion within the range of weights and the maneuver to the next stage can be centers of gravity for which certifi- made smoothly and safely; cation is requested; and (b) From the LDP on the approach (3) For power-on operations, under path selected by the applicant, a safe any condition of speed, power, and climbout can be made at speeds allow- rotor r.p.m. for which certification is ing compliance with the climb require- requested; and ments of § 29.67(a)(1) and (2); and (4) For power-off operations, under (c) The rotorcraft does not descend any condition of speed, and rotor r.p.m. below 15 feet above the landing surface. for which certification is requested For elevated heliport operations, de- that is attainable with the controls scent may be below the level of the rigged in accordance with the approved landing surface provided the deck edge rigging instructions and tolerances;

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(b) Be able to maintain any required as crosswind takeoffs, sideward flight, flight condition and make a smooth and rearward flight), with— transition from any flight condition to (1) Critical weight; any other flight condition without ex- (2) Critical center of gravity; and ceptional piloting skill, alertness, or (3) Critical rotor r.p.m. strength, and without danger of ex- (d) The rotorcraft, after (1) failure of ceeding the limit load factor under any one engine, in the case of multiengine operating condition probable for the rotorcraft that meet Transport Cat- type, including— egory A engine isolation requirements, (1) Sudden failure of one engine, for or (2) complete power failure in the multiengine rotorcraft meeting Trans- case of other rotorcraft, must be con- port Category A engine isolation re- trollable over the range of speeds and quirements; altitudes for which certification is re- (2) Sudden, complete power failure, quested when such power failure occurs for other rotorcraft; and with maximum continuous power and (3) Sudden, complete control system critical weight. No corrective action failures specified in § 29.695 of this part; time delay for any condition following and power failure may be less than— (c) Have any additional characteris- (i) For the cruise condition, one sec- tics required for night or instrument ond, or normal pilot reaction time operation, if certification for those (whichever is greater); and kinds of operation is requested. Re- (ii) For any other condition, normal quirements for helicopter instrument pilot reaction time. flight are contained in appendix B of (e) For helicopters for which a VNE this part. (power-off) is established under § 29.1505(c), compliance must be dem- [Doc. No. 5084, 29 FR 16150, Dec. 8, 1964, as onstrated with the following require- amended by Amdt. 29–3, 33 FR 905, Jan. 26, ments with critical weight, critical 1968; Amdt. 29–12, 41 FR 55471, Dec. 20, 1976; center of gravity, and critical rotor Amdt. 29–21, 48 FR 4391, Jan. 31, 1983; Amdt. 29–24, 49 FR 44436, Nov. 6, 1984] r.p.m.: (1) The helicopter must be safely § 29.143 Controllability and maneuver- slowed to VNE (power-off), without ex- ability. ceptional pilot skill after the last oper- (a) The rotorcraft must be safely con- ating engine is made inoperative at trollable and maneuverable— power-on VNE. (1) During steady flight; and (2) At a speed of 1.1 VNE (power-off), (2) During any maneuver appropriate the margin of cyclic control must to the type, including— allow satisfactory roll and pitch con- (i) Takeoff; trol with power off. (ii) Climb; (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (iii) Level flight; eral Aviation Act of 1958 (49 U.S.C. 1354(a), (iv) Turning flight; 1421, 1423, 1424, and 1425); and sec. 6(c) of the (v) Glide; and Dept. of Transportation Act (49 U.S.C. (vi) Landing (power on and power 1655(c))) off). [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (b) The margin of cyclic control must amended by Amdt. 29–3, 33 FR 965, Jan. 26, allow satisfactory roll and pitch con- 1968; Amdt. 29–15, 43 FR 2326, Jan. 16, 1978; Amdt. 29–24, 49 FR 44436, Nov. 6, 1984] trol at VNE with— (1) Critical weight; § 29.151 Flight controls. (2) Critical center of gravity; (3) Critical rotor r.p.m.; and (a) Longitudinal, lateral, directional, (4) Power off (except for helicopters and collective controls may not exhibit demonstrating compliance with para- excessive breakout force, friction, or graph (e) of this section) and power on. preload. (c) A wind velocity of not less than 17 (b) Control system forces and free knots must be established in which the play may not inhibit a smooth, direct rotorcraft can be operated without loss rotorcraft response to control system of control on or near the ground in any input. maneuver appropriate to the type (such [Amdt. 29–24, 49 FR 44436, Nov. 6, 1984]

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§ 29.161 Trim control. (5) The rotorcraft trimmed at VY. The trim control— (b) Cruise. Static longitudinal sta- (a) Must trim any steady longitu- bility must be shown in the cruise con- dinal, lateral, and collective control dition at speeds from 0.7 VH or 0.7 VNE, forces to zero in level flight at any ap- whichever is less, to 1.1 VH or 1.1 VNE, propriate speed; and whichever is less, with— (b) May not introduce any undesir- (1) Critical weight; able discontinuities in control force (2) Critical center of gravity; gradients. (3) Power for level flight at 0.9 VH or [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as 0.9 VNE, whichever is less; amended by Amdt. 29–24, 49 FR 44436, Nov. 6, (4) The landing gear retracted, and 1984] (5) The rotorcraft trimmed at 0.9 VH or 0.9 VNE, whichever is less. § 29.171 Stability: general. (c) Autorotation. Static longitudinal The rotorcraft must be able to be stability must be shown in autorota- flown, without undue pilot fatigue or tion at airspeeds from 0.5 times the strain, in any normal maneuver for a speed for minimum rate of descent, or period of time as long as that expected 0.5 times the maximum range glide in normal operation. At least three speed for Category A rotorcraft, to V landings and takeoffs must be made NE or to 1.1 VNE (power-off) if VNE (power- during this demonstration. off) is established under § 29.1505(c), and § 29.173 Static longitudinal stability. with— (1) Critical weight; (a) The longitudinal control must be designed so that a rearward movement (2) Critical center of gravity; of the control is necessary to obtain a (3) Power off; speed less than the trim speed, and a (4) The landing gear—— forward movement of the control is (i) Retracted; and necessary to obtain a speed more than (ii) Extended; and the trim speed. (5) The rotorcraft trimmed at appro- (b) With the throttle and collective priate speeds found necessary by the pitch held constant during the maneu- Administrator to demonstrate stability vers specified in § 29.175 (a) through (c), throughout the prescribed speed range. the slope of the control position versus (d) Hovering. For helicopters, the lon- speed curve must be positive through- gitudinal cyclic control must operate out the full range of altitude for which with the sense, direction of motion, certification is requested. and position as prescribed in § 29.173 be- (c) During the maneuver specified in tween the maximum approved rearward § 29.175(d), the longitudinal control po- speed and a forward speed of 17 knots sition versus speed curve may have a with— negative slope within the specified (1) Critical weight; speed range if the negative motion is not greater than 10 percent of total (2) Critical center of gravity; control travel. (3) Power required to maintain an ap- proximate constant height in ground [Amdt. 29–24, 49 FR 44436, Nov. 6, 1984] effect; § 29.175 Demonstration of static longi- (4) The landing gear extended; and tudinal stability. (5) The helicopter trimmed for hov- (a) Climb. Static longitudinal sta- ering. bility must be shown in the climb con- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- dition at speeds from 0.85 VY, or 15 eral Aviation Act of 1958 (49 U.S.C. 1354(a), knots below VY, whichever is less, to 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. 1.2 VY or 15 knots above VY, whichever of Transportation Act (49 U.S.C. 1655(c))) is greater, with— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (1) Critical weight; amended by Amdt. 29–3, 33 FR 966, Jan. 26, (2) Critical center of gravity; 1968; Amdt. 29–12, 41 FR 55471, Dec. 20, 1976; (3) Maximum continuous power; Amdt. 29–15, 43 FR 2327, Jan. 16, 1978; Amdt. (4) The landing gear retracted; and 29–24, 49 FR 44436, Nov. 6, 1984]

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§ 29.177 Static directional stability. each appropriate speed and power con- dition. Static directional stability must be positive with throttle and collective controls held constant at the trim con- Subpart C—Strength Requirements ditions specified in § 29.175 (a), (b), and GENERAL (c). Sideslip angle must increase stead- ily with directional control deflection § 29.301 Loads. for sideslip angles up to ±10° from trim. (a) Strength requirements are speci- Sufficient cues must accompany side- fied in terms of limit loads (the max- slip to alert the pilot when approach- imum loads to be expected in service) ing sideslip limits. and ultimate loads (limit loads multi- [Amdt. 29–24, 49 FR 44436, Nov. 6, 1984] plied by prescribed factors of safety). Unless otherwise provided, prescribed § 29.181 Dynamic stability: Category A loads are limit loads. rotorcraft. (b) Unless otherwise provided, the Any short-period oscillation occur- specified air, ground, and water loads ring at any speed from VY to VNE must must be placed in equilibrium with in- be positively damped with the primary ertia forces, considering each item of flight controls free and in a fixed posi- mass in the rotorcraft. These loads tion. must be distributed to closely approxi- mate or conservatively represent ac- [Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] tual conditions. (c) If deflections under load would GROUND AND WATER HANDLING significantly change the distribution of CHARACTERISTICS external or internal loads, this redis- § 29.231 General. tribution must be taken into account. The rotorcraft must have satisfac- § 29.303 Factor of safety. tory ground and water handling char- Unless otherwise provided, a factor of acteristics, including freedom from un- safety of 1.5 must be used. This factor controllable tendencies in any condi- applies to external and inertia loads tion expected in operation. unless its application to the resulting § 29.235 Taxiing condition. internal stresses is more conservative. The rotorcraft must be designed to § 29.305 Strength and deformation. withstand the loads that would occur (a) The structure must be able to when the rotorcraft is taxied over the support limit loads without detri- roughest ground that may reasonably mental or permanent deformation. At be expected in normal operation. any load up to limit loads, the defor- mation may not interfere with safe op- § 29.239 Spray characteristics. eration. If certification for water operation is (b) The structure must be able to requested, no spray characteristics support ultimate loads without failure. during taxiing, takeoff, or landing may This must be shown by— obscure the vision of the pilot or dam- (1) Applying ultimate loads to the age the rotors, propellers, or other structure in a static test for at least parts of the rotorcraft. three seconds; or (2) Dynamic tests simulating actual § 29.241 Ground resonance. load application. The rotorcraft may have no dan- gerous tendency to oscillate on the § 29.307 Proof of structure. ground with the rotor turning. (a) Compliance with the strength and deformation requirements of this sub- MISCELLANEOUS FLIGHT REQUIREMENTS part must be shown for each critical loading condition accounting for the § 29.251 Vibration. environment to which the structure Each part of the rotorcraft must be will be exposed in operation. Struc- free from excessive vibration under tural analysis (static or fatigue) may

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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 (1) Dynamic and endurance tests of limitations in the Rotorcraft Flight rotors, rotor drives, and rotor controls; Manual. (2) Limit load tests of the control system, including control surfaces; § 29.337 Limit maneuvering load fac- (3) Operation tests of the control sys- tor. tem; The rotorcraft must be designed for— (4) Flight stress measurement tests; (a) A limit maneuvering load factor (5) Landing gear drop tests; and ranging from a positive limit of 3.5 to (6) Any additional tests required for a negative limit of ¥1.0; or new or unusual design features. (b) Any positive limit maneuvering load factor not less than 2.0 and any (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, 1425) negative limit maneuvering load factor of not less than ¥0.5 for which— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (1) The probability of being exceeded amended by Amdt. 29–4, 33 FR 14106, Sept. 18, is shown by analysis and flight tests to 1968; Amdt. 27–26, 55 FR 8001, Mar. 6, 1990] be extremely remote; and § 29.309 Design limitations. (2) The selected values are appro- priate to each weight condition be- The following values and limitations tween the design maximum and design must be established to show compli- minimum weights. ance with the structural requirements of this subpart: [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (a) The design maximum and design amended by Amdt. 27–26, 55 FR 8002, Mar. 6, minimum weights. 1990] (b) The main rotor r.p.m. ranges, § 29.339 Resultant limit maneuvering power on and power off. loads. (c) The maximum forward speeds for 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 the rotorcraft flight velocity compo- (g) The positive and negative limit nent in the plane of the rotor disc to maneuvering load factors. the rotational tip speed of the rotor FLIGHT LOADS blades, and is expressed as follows:

§ 29.321 General. µ = V cos a (a) The flight load factor must be as- ΩR sumed to act normal to the longitu- where— dinal axis of the rotorcraft, and to be V=The airspeed along the flight path (f.p.s.); equal in magnitude and opposite in di- a=The angle between the projection, in the rection to the rotorcraft inertia load plane of symmetry, of the axis of no feath- factor at the center of gravity. ering and a line perpendicular to the flight

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path (radians, positive when axis is point- § 29.361 Engine torque. ing aft); W=The angular velocity of rotor (radians per The limit engine torque may not be second); and less than the following: R=The rotor radius (ft.). (a) For turbine engines, the highest of— § 29.341 Gust loads. (1) The mean torque for maximum Each rotorcraft must be designed to continuous power multiplied by 1.25; withstand, at each critical airspeed in- (2) The torque required by § 29.923; cluding hovering, the loads resulting (3) The torque required by § 29.927; or from vertical and horizontal gusts of 30 (4) The torque imposed by sudden en- feet per second. gine stoppage due to malfunction or structural failure (such as compressor § 29.351 Yawing conditions. jamming). (a) Each rotorcraft must be designed (b) For reciprocating engines, 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- acting inertia forces; and [Amdt. 29–26, 53 FR 34215, Sept. 2, 1988] (2) Maximum main rotor speed. CONTROL SURFACE AND SYSTEM LOADS (b) To produce the load required in paragraph (a) of this section, in unac- § 29.391 General. celerated flight with zero yaw, at for- Each auxiliary rotor, each fixed or ward speeds from zero up to 0.6 VNE— (1) Displace the cockpit directional movable stabilizing or control surface, control suddenly to the maximum de- and each system operating any flight flection limited by the control stops or control must meet the requirements of by the maximum pilot force specified §§ 29.395 through 29.399, 29.411, and in § 29.397(a); 29.427. (2) Attain a resulting sideslip angle [Amdt. 29–26, 55 FR 8002, Mar. 6, 1990, as or 90°, whichever is less; and amended by Amdt. 29–41, 62 FR 46173, Aug. 29, (3) Return the directional control 1997] suddenly to neutral. (c) To produce the load required in § 29.395 Control system. paragraph (a) of the section, in unac- (a) The reaction to the loads pre- celerated flight with zero yaw, at for- scribed in § 29.397 must be provided by— ward speeds from 0.6 VNE up to VNE or (1) The control stops only; VH, whichever is less— (2) The control locks only; (1) Displace the cockpit directional (3) The irreversible mechanism only control suddenly to the maximum de- (with the mechanism locked and with flection limited by the control stops or the control surface in the critical posi- by the maximum pilot force specified tions for the effective parts of the sys- in § 29.397(a); tem within its limit of motion); (2) Attain a resulting sideslip angle (4) The attachment of the control or 15°, whichever is less, at the lesser system to the rotor blade pitch control speed of VNE or VH; horn only (with the control in the crit- (3) Vary the sideslip angles of para- ical positions for the affected parts of graphs (b)(2) and (c)(2) of this section the system within the limits of its mo- directly with speed; and tion); and (4) Return the directional control (5) The attachment of the control suddenly to neutral. system to the control surface horn [Amdt. 29–26, 55 FR 8002, Mar. 6, 1990, as (with the control in the critical posi- amended by Amdt. 29–41, 62 FR 46173, Aug. 29, tions for the affected parts of the sys- 1997] tem within the limits of its motion).

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(b) Each primary control system, in- (2) Twist controls, 80R inch-pounds. cluding its supporting structure, must [Amdt. 29–12, 41 FR 55471, Dec. 20, 1976, as be designed as follows: amended by Amdt. 29–47, 66 FR 23538, May 9, (1) The system must withstand loads 2001] resulting from the limit pilot forces prescribed in § 29.397; § 29.399 Dual control system. (2) Notwithstanding paragraph (b)(3) Each dual primary flight control sys- of this section, when power-operated actuator controls or power boost con- tem must be able to withstand the trols are used, the system must also loads that result when pilot forces not withstand the loads resulting from the less than 0.75 times those obtained limit pilot forces prescribed in § 29.397 under § 29.395 are applied— in conjunction with the forces output (a) In opposition; and of each normally energized power de- (b) In the same direction. vice, including any single power boost § 29.411 Ground clearance: tail rotor or actuator system failure; guard. (3) If the system design or the normal operating loads are such that a part of (a) It must be impossible for the tail the system cannot react to the limit rotor to contact the landing surface pilot forces prescribed in § 29.397, that during a normal landing. part of the system must be designed to (b) If a tail rotor guard is required to withstand the maximum loads that can show compliance with paragraph (a) of be obtained in normal operation. The this section— minimum design loads must, in any (1) Suitable design loads must be es- case, provide a rugged system for serv- tablished for the guard: and ice use, including consideration of fa- (2) The guard and its supporting tigue, jamming, ground gusts, control structure must be designed to with- inertia, and friction loads. In the ab- stand those loads. sence of a rational analysis, the design loads resulting from 0.60 of the speci- § 29.427 Unsymmetrical loads. fied limit pilot forces are acceptable (a) Horizontal tail surfaces and their minimum design loads; and supporting structure must be designed (4) If operational loads may be ex- for unsymmetrical loads arising from ceeded through jamming, ground gusts, yawing and rotor wake effects in com- control inertia, or friction, the system bination with the prescribed flight con- must withstand the limit pilot forces ditions. specified in § 29.397, without yielding. (b) To meet the design criteria of [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as paragraph (a) of this section, in the ab- amended by Amdt. 29–26, 55 FR 8002, Mar. 6, sence of more rational data, both of the 1990] following must be met: (1) One hundred percent of the max- § 29.397 Limit pilot forces and torques. imum loading from the symmetrical (a) Except as provided in paragraph flight conditions acts on the surface on (b) of this section, the limit pilot one side of the plane of symmetry, and forces are as follows: no loading acts on the other side. (1) For foot controls, 130 pounds. (2) Fifty percent of the maximum (2) For stick controls, 100 pounds fore loading from the symmetrical flight and aft, and 67 pounds laterally. conditions acts on the surface on each (b) For flap, tab, stabilizer, rotor side of the plane of symmetry, in oppo- brake, and landing gear operating con- site directions. trols, the following apply (R=radius in (c) For empennage arrangements inches): where the horizontal tail surfaces are (1) Crank wheel, and lever controls, [1 supported by the vertical tail surfaces, + R]/3 × 50 pounds, but not less than 50 the vertical tail surfaces and sup- pounds nor more than 100 pounds for porting structure must be designed for hand operated controls or 130 pounds the combined vertical and horizontal for foot operated controls, applied at surface loads resulting from each pre- any angle within 20 degrees of the scribed flight condition, considered plane of motion of the control. separately. The flight conditions must

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be selected so that the maximum de- § 29.475 Tires and shock absorbers. sign loads are obtained on each surface. Unless otherwise prescribed, for each In the absence of more rational data, specified landing condition, the tires the unsymmetrical horizontal tail sur- must be assumed to be in their static face loading distributions described in position and the shock absorbers to be this section must be assumed. in their most critical position. [Amdt. 27–26, 55 FR 8002, Mar. 6, 1990, as amended by Amdt. 29–31, 55 FR 38966, Sept. § 29.477 Landing gear arrangement. 21, 1990] Sections 29.235, 29.479 through 29.485, and 29.493 apply to landing gear with GROUND LOADS two wheels aft, and one or more wheels forward, of the center of gravity. § 29.471 General. (a) Loads and equilibrium. For limit § 29.479 Level landing conditions. ground loads— (a) Attitudes. Under each of the load- (1) The limit ground loads obtained ing conditions prescribed in paragraph in the landing conditions in this part (b) of this section, the rotorcraft is as- must be considered to be external loads sumed to be in each of the following that would occur in the rotorcraft level landing attitudes: structure if it were acting as a rigid (1) An attitude in which each wheel body; and contacts the ground simultaneously. (2) In each specified landing condi- (2) An attitude in which the aft tion, the external loads must be placed wheels contact the ground with the for- ward wheels just clear of the ground. in equilibrium with linear and angular (b) Loading conditions. The rotorcraft inertia loads in a rational or conserv- must be designed for the following ative manner. landing loading conditions: (b) Critical centers of gravity. The crit- (1) Vertical loads applied under ical centers of gravity within the range § 29.471. for which certification is requested (2) The loads resulting from a com- must be selected so that the maximum bination of the loads applied under design loads are obtained in each land- paragraph (b)(1) of this section with ing gear element. drag loads at each wheel of not less than 25 percent of the vertical load at § 29.473 Ground loading conditions that wheel. and assumptions. (3) The vertical load at the instant of (a) For specified landing conditions, peak drag load combined with a drag a design maximum weight must be component simulating the forces re- used that is not less than the max- quired to accelerate the wheel rolling imum weight. A rotor lift may be as- assembly up to the specified ground sumed to act through the center of speed, with— gravity throughout the landing impact. (i) The ground speed for determina- This lift may not exceed two-thirds of tion of the spin-up loads being at least the design maximum weight. 75 percent of the optimum forward (b) Unless otherwise prescribed, for flight speed for minimum rate of de- each specified landing condition, the scent in autorotation; and rotorcraft must be designed for a limit (ii) The loading conditions of para- load factor of not less than the limit graph (b) applied to the landing gear inertia load factor substantiated under and its attaching structure only. § 29.725. (4) If there are two wheels forward, a (c) Triggering or actuating devices distribution of the loads applied to for additional or supplementary energy those wheels under paragraphs (b)(1) absorption may not fail under loads es- and (2) of this section in a ratio of tablished in the tests prescribed in 40:60. (c) Pitching mo- §§ 29.725 and 29.727, but the factor of Pitching moments. ments are assumed to be resisted by— safety prescribed in § 29.303 need not be (1) In the case of the attitude in para- used. graph (a)(1) of this section, the forward [Amdt. 29–3, 33 FR 966, Jan. 26, 1968] landing gear; and

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(2) In the case of the attitude in para- § 29.493 Braked roll conditions. graph (a)(2) of this section, the angular Under braked roll conditions with inertia forces. the shock absorbers in their static po- sitions— § 29.481 Tail-down landing conditions. (a) The limit vertical load must be (a) The rotorcraft is assumed to be in based on a load factor of at least— the maximum nose-up attitude allow- (1) 1.33, for the attitude specified in ing ground clearance by each part of § 29.479(a)(1); and the rotorcraft. (2) 1.0, for the attitude specified in (b) In this attitude, ground loads are § 29.479(a)(2); and assumed to act perpendicular to the (b) The structure must be designed to ground. withstand, at the ground contact point of each wheel with brakes, a drag load § 29.483 One-wheel landing conditions. of at least the lesser of— For the one-wheel landing condition, (1) The vertical load multiplied by a the rotorcraft is assumed to be in the coefficient of friction of 0.8; and level attitude and to contact the (2) The maximum value based on lim- ground on one aft wheel. In this atti- iting brake torque. tude— (a) The vertical load must be the § 29.497 Ground loading conditions: same as that obtained on that side landing gear with tail wheels. under § 29.479(b)(1); and (a) General. Rotorcraft with landing (b) The unbalanced external loads gear with two wheels forward and one must be reacted by rotorcraft inertia. wheel aft of the center of gravity must be designed for loading conditions as § 29.485 Lateral drift landing condi- prescribed in this section. tions. (b) Level landing attitude with only the (a) The rotorcraft is assumed to be in forward wheels contacting the ground. In the level landing attitude, with— this attitude— (1) The vertical loads must be applied (1) Side loads combined with one-half under §§ 29.471 through 29.475; of the maximum ground reactions ob- tained in the level landing conditions (2) The vertical load at each axle of § 29.479(b)(1); and must be combined with a drag load at that axle of not less than 25 percent of (2) The loads obtained under para- that vertical load; and graph (a)(1) of this section applied— (3) Unbalanced pitching moments are (i) At the ground contact point; or assumed to be resisted by angular iner- (ii) For full-swiveling gear, at the tia forces. center of the axle. (c) Level landing attitude with all (b) The rotorcraft must be designed wheels contacting the ground simulta- to withstand, at ground contact— neously. In this attitude, the rotorcraft (1) When only the aft wheels contact must be designed for landing loading the ground, side loads of 0.8 times the conditions as prescribed in paragraph vertical reaction acting inward on one (b) of this section. side and 0.6 times the vertical reaction (d) Maximum nose-up attitude with acting outward on the other side, all only the rear wheel contacting the combined with the vertical loads speci- ground. The attitude for this condition fied in paragraph (a) of this section; must be the maximum nose-up attitude and expected in normal operation, includ- (2) When the wheels contact the ing autorotative landings. In this atti- ground simultaneously— tude— (i) For the aft wheels, the side loads (1) The appropriate ground loads specified in paragraph (b)(1) of this sec- specified in paragraph (b)(1) and (2) of tion; and this section must be determined and (ii) For the forward wheels, a side applied, using a rational method to ac- load of 0.8 times the vertical reaction count for the moment arm between the combined with the vertical load speci- rear wheel ground reaction and the fied in paragraph (a) of this section. rotorcraft center of gravity; or

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(2) The probability of landing with (h) Rear wheel turning loads in the initial contact on the rear wheel must static ground attitude. In the static be shown to be extremely remote. ground attitude, and with the shock (e) Level landing attitude with only one absorbers and tires in their static posi- forward wheel contacting the ground. In tions, the rotorcraft must be designed this attitude, the rotorcraft must be for rear wheel turning loads as follows: designed for ground loads as specified (1) A vertical ground reaction equal in paragraph (b)(1) and (3) of this sec- to the static load on the rear wheel tion. must be combined with an equal side (f) Side loads in the level landing atti- load. tude. In the attitudes specified in para- (2) The load specified in paragraph graphs (b) and (c) of this section, the (h)(1) of this section must be applied to following apply: the rear landing gear— (1) The side loads must be combined (i) Through the axle, if there is a at each wheel with one-half of the max- swivel (the rear wheel being assumed imum vertical ground reactions ob- to be swiveled 90 degrees to the longi- tained for that wheel under paragraphs tudinal axis of the rotorcraft); or (b) and (c) of this section. In this condi- tion, the side loads must be— (ii) At the ground contact point if (i) For the forward wheels, 0.8 times there is a lock, steering device or shim- the the vertical reaction (on one side) my damper (the rear wheel being as- acting inward, and 0.6 times the sumed to be in the trailing position). vertical reaction (on the other side) (i) Taxiing condition. The rotorcraft acting outward; and and its landing gear must be designed (ii) For the rear wheel, 0.8 times the for the loads that would occur when vertical reaction. the rotorcraft is taxied over the rough- (2) The loads specified in paragraph est ground that may reasonably be ex- (f)(1) of this section must be applied— pected in normal operation. (i) At the ground contact point with the wheel in the trailing position (for § 29.501 Ground loading conditions: non-full swiveling landing gear or for landing gear with skids. full swiveling landing gear with a lock, (a) General. Rotorcraft with landing steering device, or shimmy damper to gear with skids must be designed for keep the wheel in the trailing posi- the loading conditions specified in this tion); or section. In showing compliance with (ii) At the center of the axle (for full this section, the following apply: swiveling landing gear without a lock, (1) The design maximum weight, cen- steering device, or shimmy damper). ter of gravity, and load factor must be (g) Braked roll conditions in the level determined under §§ 29.471 through landing attitude. In the attitudes speci- 29.475. fied in paragraphs (b) and (c) of this (2) Structural yielding of elastic section, and with the shock absorbers spring members under limit loads is ac- in their static positions, the rotorcraft ceptable. must be designed for braked roll loads (3) Design ultimate loads for elastic as follows: spring members need not exceed those (1) The limit vertical load must be obtained in a drop test of the gear based on a limit vertical load factor of not less than— with— (i) 1.0, for the attitude specified in (i) A drop height of 1.5 times that paragraph (b) of this section; and specified in § 29.725; and (ii) 1.33, for the attitude specified in (ii) An assumed rotor lift of not more paragraph (c) of this section. than 1.5 times that used in the limit (2) For each wheel with brakes, a drop tests prescribed in § 29.725. drag load must be applied, at the (4) Compliance with paragraph (b) ground contact point, of not less than through (e) of this section must be the lesser of— shown with— (i) 0.8 times the vertical load; and (i) The gear in its most critically de- (ii) The maximum based on limiting flected position for the landing condi- brake torque. tion being considered; and

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(ii) The ground reactions rationally (f) Special conditions. In addition to distributed along the bottom of the the conditions specified in paragraphs skid tube. (b) and (c) of this section, the rotor- (b) Vertical reactions in the level land- craft must be designed for the fol- ing attitude. In the level attitude, and lowing ground reactions: with the rotorcraft contacting the (1) A ground reaction load acting up ground along the bottom of both skids, and aft at an angle of 45 degrees to the the vertical reactions must be applied longitudinal axis of the rotorcraft. as prescribed in paragraph (a) of this This load must be— section. (i) Equal to 1.33 times the maximum (c) Drag reactions in the level landing weight; attitude. In the level attitude, and with (ii) Distributed symmetrically among the rotorcraft contacting the ground the skids; along the bottom of both skids, the fol- (iii) Concentrated at the forward end lowing apply: of the straight part of the skid tube; (1) The vertical reactions must be and combined with horizontal drag reac- (iv) Applied only to the forward end tions of 50 percent of the vertical reac- of the skid tube and its attachment to tion applied at the ground. the rotorcraft. (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. [Amdt. 29–3, 33 FR 966, Jan. 26, 1968; as (2) The vertical ground reactions amended by Amdt. 27–26, 55 FR 8002, Mar. 6, must be combined with a horizontal 1990] sideload of 25 percent of their value. (3) The total sideload must be applied § 29.505 Ski landing conditions. equally between skids and along the If certification for ski operation is 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 § 29.473(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

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axis through the centerline of the ped- onstrated that the forward velocity se- estal bearings. lected would not be exceeded in a nor- mal one-engine-out landing. § 29.511 Ground load: unsymmetrical (d) Auxiliary float immersion condition. loads on multiple-wheel units. In addition to the loads from the land- (a) In dual-wheel gear units, 60 per- ing conditions, the auxiliary float, and cent of the total ground reaction for its support and attaching structure in the gear unit must be applied to one the hull, must be designed for the load wheel and 40 percent to the other. developed by a fully immersed float un- (b) To provide for the case of one de- less it can be shown that full immer- flated tire, 60 percent of the specified sion of the float is unlikely, in which load for the gear unit must be applied case the highest likely float buoyancy to either wheel except that the vertical load must be applied that considers ground reaction may not be less than the full static value. loading of the float immersed to create (c) In determining the total load on a restoring moments compensating for gear unit, the transverse shift in the upsetting moments caused by side load centroid, due to unsymmetrical wind, asymmetrical rotorcraft loading, load distribution on the wheels, may be water wave action, and rotorcraft iner- neglected. tia. [Amdt. 29–3, 33 FR 966, Jan. 26, 1968] [Amdt. 29–3, 33 FR 966, Jan. 26, 196; as amend- ed by Amdt. 27–26, 55 FR 8002, Mar. 6, 1990] WATER LOADS § 29.521 Float landing conditions. § 29.519 Hull type rotorcraft: Water- based and amphibian. If certification for float operation (including float amphibian operation) (a) General. For hull type rotorcraft, is requested, the rotorcraft, with the structure must be designed to with- floats, must be designed to withstand stand the water loading set forth in the following loading conditions (where paragraphs (b), (c), and (d) of this sec- the limit load factor is determined tion considering the most severe wave under § 29.473(b) or assumed to be equal heights and profiles for which approval to that determined for wheel landing is desired. The loads for the landing gear): conditions of paragraphs (b) and (c) of this section must be developed and dis- (a) Up-load conditions in which— tributed along and among the hull and (1) A load is applied so that, with the auxiliary floats, if used, in a rational rotorcraft in the static level attitude, and conservative manner, assuming a the resultant water reaction passes rotor lift not exceeding two-thirds of vertically through the center of grav- the rotorcraft weight to act through- ity; and out the landing impact. (2) The vertical load prescribed in (b) Vertical landing conditions. The paragraph (a)(1) of this section is ap- rotorcraft must initially contact the plied simultaneously with an aft com- most critical wave surface at zero for- ponent of 0.25 times the vertical com- ward speed in likely pitch and roll atti- ponent tudes which result in critical design (b) A side load condition in which— loadings. The vertical descent velocity (1) A vertical load of 0.75 times the may not be less than 6.5 feet per second total vertical load specified in para- relative to the mean water surface. graph (a)(1) of this section is divided (c) Forward speed landing conditions. equally among the floats; and The rotorcraft must contact the most (2) For each float, the load share de- critical wave at forward velocities termined under paragraph (b)(1) of this from zero up to 30 knots in likely section, combined with a total side pitch, roll, and yaw attitudes and with load of 0.25 times the total vertical a vertical descent velocity of not less load specified in paragraph (b)(1) of than 6.5 feet per second relative to the this section, is applied to that float mean water surface. A maximum for- only. ward velocity of less than 30 knots may be used in design if it can be dem- [Amdt. 29–3, 33 FR 967, Jan. 26, 1968]

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MAIN COMPONENT REQUIREMENTS § 29.549 Fuselage and rotor pylon structures. § 29.547 Main and tail rotor structure. (a) Each fuselage and rotor pylon (a) A rotor is an assembly of rotating structure must be designed to with- components, which includes the rotor stand— hub, blades, blade dampers, the pitch (1) The critical loads prescribed in control mechanisms, and all other §§ 29.337 through 29.341, and 29.351; parts that rotate with the assembly. (2) The applicable ground loads pre- (b) Each rotor assembly must be de- scribed in §§ 29.235, 29.471 through 29.485, signed as prescribed in this section and 29.493, 29.497, 29.505, and 29.521; and must function safely for the critical (3) The loads prescribed in § 29.547 flight load and operating conditions. A (d)(1) and (e)(1)(i). design assessment must be performed, (b) Auxiliary rotor thrust, the torque including a detailed failure analysis to reaction of each rotor drive system, identify all failures that will prevent and the balancing air and inertia loads continued safe flight or safe landing, occurring under accelerated flight con- and must identify the means to mini- ditions, must be considered. mize the likelihood of their occurrence. (c) Each engine mount and adjacent (c) The rotor structure must be de- fuselage structure must be designed to signed to withstand the following loads withstand the loads occurring under prescribed in §§ 29.337 through 29.341 and accelerated flight and landing condi- 29.351: tions, including engine torque. (d) [Reserved] (1) Critical flight loads. (e) If approval for the use of 21⁄2- (2) Limit loads occurring under nor- minute OEI power is requested, each mal conditions of autorotation. engine mount and adjacent structure (d) The rotor structure must be de- must be designed to withstand the signed to withstand loads simulating— loads resulting from a limit torque (1) For the rotor blades, hubs, and equal to 1.25 times the mean torque for flapping hinges, the impact force of 21⁄2-minute OEI power combined with 1g each blade against its stop during flight loads. ground operation; and (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, (2) Any other critical condition ex- 1425) pected in normal operation. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (e) The rotor structure must be de- amended by Amdt. 29–4, 33 FR 14106, Sept. 18, signed to withstand the limit torque at 1968; Amdt. 29–26, 53 FR 34215, Sept. 2, 1988] any rotational speed, including zero. In addition: § 29.551 Auxiliary lifting surfaces. (1) The limit torque need not be Each auxiliary lifting surface must greater than the torque defined by a be designed to withstand— torque limiting device (where pro- (a) The critical flight loads in §§ 29.337 vided), and may not be less than the through 29.341, and 29.351; greater of— (b) the applicable ground loads in (i) The maximum torque likely to be §§ 29.235, 29.471 through 29.485, 29.493, transmitted to the rotor structure, in 29.505, and 29.521; and either direction, by the rotor drive or (c) Any other critical condition ex- by sudden application of the rotor pected in normal operation. brake; and (ii) For the main rotor, the limit en- EMERGENCY LANDING CONDITIONS gine torque specified in § 29.361. § 29.561 General. (2) The limit torque must be equally and rationally distributed to the rotor (a) The rotorcraft, although it may blades. be damaged in emergency landing con- ditions on land or water, must be de- (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, signed as prescribed in this section to 1425) protect the occupants under those con- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ditions. amended by Amdt. 29–4, 33 FR 14106, Sept. 18, (b) The structure must be designed to 1968; Amdt. 29–40, 61 FR 21907, May 10, 1996] give each occupant every reasonable

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chance of escaping serious injury in a (1) The occupant properly uses the crash landing when— seats, safety belts, and shoulder har- (1) Proper use is made of seats, belts, nesses provided in the design; and and other safety design provisions; (2) The occupant is exposed to loads (2) The wheels are retracted (where equivalent to those resulting from the applicable); and 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 the in- pant simulated by a 170-pound tended 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 factor up to those (1) A change in downward velocity of specified in this paragraph, any item of not less than 30 feet per second when mass above and/or behind the crew and the seat or other seating device is ori- passenger compartment that could in- ented in its nominal position with re- jure 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, transmission, 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. (5) Rearward—1.5g. (2) A change in forward velocity of (d) Any fuselage structure in the area not less than 42 feet per second when of internal fuel tanks below the pas- the seat or other seating device is ori- senger floor level must be designed to ented in its nominal position with re- resist the following ultimate inertial spect to the rotorcraft’s reference sys- factors and loads, and to protect the tem, the rotorcraft’s longitudinal axis ° fuel tanks from rupture, if rupture is is yawed 10 either right or left of the likely when those loads are applied to impact velocity vector (whichever that area: would cause the greatest load on the shoulder harness), the rotorcraft’s lat- (1) Upward—1.5g. eral axis is contained in a horizontal (2) Forward—4.0g. plane containing the impact velocity (3) Sideward—2.0g. vector, and the rotorcraft’s vertical (4) Downward—4.0g. axis is perpendicular to a horizontal [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as plane containing the impact velocity amended by Amdt. 29–29, 54 FR 47319, Nov. 13, vector. Peak floor deceleration must 1989; Amdt. 29–38, 61 FR 10438, Mar. 13, 1996] occur in not more than 0.071 seconds after impact and must reach a min- § 29.562 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 a crash landing, must be tach the seating devices to the air- designed to reasonably protect each oc- frame structure for the conditions of cupant when— this section, the rails or devices must

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be misaligned with respect to each § 29.563 Structural ditching provi- other by at least 10° vertically (i.e., sions. pitch out of parallel) and by at least a If certification with ditching provi- 10° lateral roll, with the directions op- sions is requested, structural strength tional, to account for possible floor for ditching must meet the require- warp. ments of this section and § 29.801(e). (c) Compliance with the following (a) Forward speed landing conditions. must be shown: The rotorcraft must initially contact (1) The seating device system must the most critical wave for reasonably remain intact although it may experi- probable water conditions at forward ence separation intended as part of its velocities from zero up to 30 knots in design. likely pitch, roll, and yaw attitudes. (2) The attachment between the seat- The rotorcraft limit vertical descent ing device and the airframe structure velocity may not be less than 5 feet per second relative to the mean water sur- must remain intact although the struc- face. Rotor lift may be used to act ture may have exceeded its limit load. through the center of gravity through- (3) The ATD’s shoulder harness strap out the landing impact. This lift may or straps must remain on or in the im- not exceed two-thirds of the design mediate vicinity of the ATD’s shoulder maximum weight. A maximum forward during the impact. velocity of less than 30 knots may be (4) The safety belt must remain on used in design if it can be dem- the ATD’s pelvis during the impact. onstrated that the forward velocity se- (5) The ATD’s head either does not lected would not be exceeded in a nor- contact any portion of the crew or pas- mal one-engine-out touchdown. senger compartment or, if contact is (b) Auxiliary or emergency float condi- made, the head impact does not exceed tions—(1) Floats fixed or deployed before a head injury criteria (HIC) of 1,000 as initial water contact. In addition to the determined by this equation. landing loads in paragraph (a) of this section, each auxiliary or emergency  2.5 float, or its support and attaching =−()1 t2 structure in the airframe or fuselage, HIC t21 t  ∫ a(t)dt must be designed for the load devel- ()tt− t1   21  oped by a fully immersed float unless it Where: a(t) is the resultant acceleration at can be shown that full immersion is the center of gravity of the head form ex- unlikely. If full immersion is unlikely, pressed as a multiple of g (the acceleration the highest likely float buoyancy load of gravity) and t2 ¥ t1 is the time duration, must be applied. The highest likely in seconds, of major head impact, not to buoyancy load must include consider- exceed 0.05 seconds. ation of a partially immersed float cre- (6) Loads in individual shoulder har- ating restoring moments to com- ness straps must not exceed 1,750 pensate the upsetting moments caused pounds. If dual straps are used for re- by side wind, unsymmetrical rotorcraft taining the upper torso, the total har- loading, water wave action, rotorcraft ness strap loads must not exceed 2,000 inertia, and probable structural dam- pounds. age and leakage considered under (7) The maximum compressive load § 29.801(d). Maximum roll and pitch an- measured between the pelvis and the gles determined from compliance with § 29.801(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. 29–29, 54 FR 47320, Nov. 13, 1989, as purpose, the design airspeed for limit amended by Amdt. 29–41, 62 FR 46173, Aug. 29, load is the float deployed airspeed op- 1997] erating limit multiplied by 1.11.

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(2) Floats deployed after initial water other procedures must be established contact. Each float must be designed for as necessary to avoid catastrophic fail- full or partial immersion prescribed in ure. These inspections, replacement paragraph (b)(1) of this section. In addi- times, combinations thereof, or other tion, each float must be designed for procedures must be included in the air- combined vertical and drag loads using worthiness limitations section of the a relative limit speed of 20 knots be- Instructions for Continued Airworthi- tween the rotorcraft and the water. ness required by § 29.1529 and section The vertical load may not be less than A29.4 of appendix A of this part. the highest likely buoyancy load deter- (b) Fatigue tolerance evaluation (in- mined under paragraph (b)(1) of this cluding tolerance to flaws). The struc- section. ture must be shown by analysis sup- [Amdt. 27–26, 55 FR 8003, Mar. 6, 1990] ported by test evidence and, if avail- able, service experience to be of fatigue FATIGUE EVALUATION tolerant design. The fatigue tolerance evaluation must include the require- § 29.571 Fatigue evaluation of struc- ments of either paragraph (b)(1), (2), or ture. (3) of this section, or a combination (a) General. An evaluation of the thereof, and also must include a deter- strength of principal elements, detail mination of the probable locations and design points, and fabrication tech- modes of damage caused by fatigue, niques must show that catastrophic considering environmental effects, in- failure due to fatigue, considering the trinsic/discrete flaws, or accidental effects of environment, intrinsic/dis- damage. Compliance with the flaw tol- crete flaws, or accidental damage will erance requirements of paragraph (b)(1) be avoided. Parts to be evaluated in- or (2) of this section is required unless clude, but are not limited to, rotors, the applicant establishes that these fa- rotor drive systems between the en- tigue flaw tolerant methods for a par- gines and rotor hubs, controls, fuse- ticular structure cannot be achieved lage, fixed and movable control sur- within the limitations of geometry, faces, engine and transmission mount- inspectability, or good design practice. ings, landing gear, and their related Under these circumstances, the safe- primary attachments. In addition, the life evaluation of paragraph (b)(3) of following apply: this section is required. (1) Each evaluation required by this (1) Flaw tolerant safe-life evaluation. It section must include— must be shown that the structure, with (i) The identification of principal flaws present, is able to withstand re- structural elements, the failure of peated loads of variable magnitude which could result in catastrophic fail- without detectable flaw growth for the ure of the rotorcraft; following time intervals— (ii) In-flight measurement in deter- mining the loads or stresses for items (i) Life of the rotorcraft; or in paragraph (a)(1)(i) of this section in (ii) Within a replacement time fur- all critical conditions throughout the nished under section A29.4 of appendix range of limitations in § 29.309 (includ- A to this part. ing altitude effects), except that ma- (2) Fail-safe (residual strength after neuvering load factors need not exceed flaw growth) evaluation. It must be the maximum values expected in oper- shown that the structure remaining ations; and after a partial failure is able to with- (iii) Loading spectra as severe as stand design limit loads without fail- those expected in operation based on ure within an inspection period fur- loads or stresses determined under nished under section A29.4 of appendix paragraph (a)(1)(ii) of this section, in- A to this part. Limit loads are defined cluding external load operations, if ap- in § 29.301(a). plicable, and other high frequency (i) The residual strength evaluation power cycle operations. must show that the remaining struc- (2) Based on the evaluations required ture after flaw growth is able to with- by this section, inspections, replace- stand design limit loads without fail- ment times, combinations thereof, or ure within its operational life.

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(ii) Inspection intervals and methods which could adversely affect safety, must be established as necessary to en- must— sure that failures are detected prior to (a) Be established on the basis of ex- residual strength conditions being perience or tests; reached. (b) Meet approved specifications that (iii) If significant changes in struc- ensure their having the strength and tural stiffness or geometry, or both, other properties assumed in the design follow from a structural failure or par- data; and tial failure, the effect on flaw tolerance (c) Take into account the effects of must be further investigated. environmental conditions, such as tem- (3) Safe-life evaluation. It must be perature and humidity, expected in shown that the structure is able to service. withstand repeated loads of variable magnitude without detectable cracks (Secs. 313(a), 601, 603, 604, and 605 of the Fed- for the following time intervals— eral Aviation Act of 1958 (49 U.S.C. 1354(a), (i) Life of the rotorcraft; or 1421, 1423, 1424), and sec. 6(c), Dept. of Trans- (ii) Within a replacement time fur- portation Act (49 U.S.C. 1655(c))) nished under section A29.4 of appendix [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as A to this part. amended by Amdt. 29–12, 41 FR 55471, Dec. 20, [Amdt. 29–28, 54 FR 43930, Oct. 27, 1989] 1976; Amdt. 29–17, 43 FR 50599, Oct. 30, 1978] § 29.605 Fabrication methods. Subpart D—Design and Construction (a) The methods of fabrication used must produce consistently sound struc- GENERAL tures. If a fabrication process (such as gluing, spot welding, or heat-treating) § 29.601 Design. requires close control to reach this ob- (a) The rotorcraft may have no de- jective, the process must be performed sign features or details that experience according to an approved process speci- has shown to be hazardous or unreli- fication. able. (b) Each new aircraft fabrication (b) The suitability of each question- method must be substantiated by a able design detail and part must be es- test program. tablished by tests. (Secs. 313(a), 601, 603, 604, Federal Aviation § 29.602 Critical parts. Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 (a) Critical part. A critical part is a U.S.C. 1655(c))) part, the failure of which could have a catastrophic effect upon the rotocraft, [Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as and for which critical characterists amended by Amdt. 29–17, 43 FR 50599, Oct. 30, 1978] have been identified which must be controlled to ensure the required level § 29.607 Fasteners. of integrity. (b) If the type design includes critical (a) Each removable bolt, screw, nut, parts, a critical parts list shall be es- pin, or other fastener whose loss could tablished. Procedures shall be estab- jeopardize the safe operation of the lished to define the critical design rotorcraft must incorporate two sepa- characteristics, identify processes that rate locking devices. The fastener and affect those characteristics, and iden- its locking devices may not be ad- tify the design change and process versely affected by the environmental change controls necessary for showing conditions associated with the par- compliance with the quality assurance ticular installation. requirements of part 21 of this chapter. (b) No self-locking nut may be used on any bolt subject to rotation in oper- [Doc. No. 29311, 64 FR 46232, Aug. 24, 1999] ation unless a nonfriction locking de- § 29.603 Materials. vice is used in addition to the self-lock- ing device. The suitability and durability of ma- terials used for parts, the failure of [Amdt. 29–5, 33 FR 14533, Sept. 27, 1968]

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§ 29.609 Protection of structure. (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; § 29.613 Material strength properties (2) Corrosion; and and design values. (3) Abrasion; and (a) Material strength properties must (b) Have provisions for ventilation be based on enough tests of material and drainage where necessary to pre- meeting specifications to establish de- vent the accumulation of corrosive, sign values on a statistical basis. flammable, or noxious fluids. (b) Design values must be chosen to minimize the probability of structural § 29.610 Lightning and static elec- tricity protection. failure due to material variability. Ex- cept as provided in paragraphs (d) and (a) The rotorcraft structure must be (e) of this section, compliance with protected against catastrophic effects this paragraph must be shown by se- from lightning. lecting design values that assure mate- (b) For metallic components, compli- ance with paragraph (a) of this section rial strength with the following prob- may be shown by— ability— (1) Electrically bonding the compo- (1) Where applied loads are eventu- nents properly to the airframe; or ally distributed through a single mem- (2) Designing the components so that ber within an assembly, the failure of a strike will not endanger the rotor- which would result in loss of structural craft. integrity of the component, 99 percent (c) For nonmetallic components, probability with 95 percent confidence; compliance with paragraph (a) of this and section may be shown by— (2) For redundant structures, those in (1) Designing the components to which the failure of individual ele- minmize the effect of a strike; or ments would result in applied loads (2) Incorporating acceptable means of being safely distributed to other load- diverting the resulting electrical cur- carrying members, 90 percent prob- rent to not endanger the rotorcraft. ability with 95 percent confidence. (d) The electric bonding and protec- (c) The strength, detail design, and tion against lightning and static elec- fabrication of the structure must mini- tricity must— mize the probability of disastrous fa- (1) Minimize the accumulation of tigue failure, particularly at points of electrostatic charge; stress concentration. (2) Minimize the risk of electric (d) Design values may be those con- shock to crew, passengers, and service tained in the following publications and maintenance personnel using nor- (available from the Naval Publications mal precautions; (3) Provide and electrical return and Forms Center, 5801 Tabor Avenue, path, under both normal and fault con- Philadelphia, PA 19120) or other values ditions, on rotorcraft having grounded approved by the Administrator: electrical systems; and (1) MIL—HDBK–5, ‘‘Metallic Materials and (4) Reduce to an acceptable level the Elements for Flight Vehicle Structure’’. effects of lightning and static elec- (2) MIL—HDBK–17, ‘‘Plastics for Flight Ve- tricity on the functioning of essential hicles’’. electrical and electronic equipment. (3) ANC–18, ‘‘Design of Wood Aircraft Structures’’. [Amdt. 29–24, 49 FR 44437, Nov. 6, 1984; Amdt. (4) MIL—HDBK–23, ‘‘Composite Construc- 29–40, 61 FR 21907, May 10, 1996; 61 FR 33963, tion for Flight Vehicles’’. July 1, 1996] (e) Other design values may be used if § 29.611 Inspection provisions. a selection of the material is made in There must be means to allow close which a specimen of each individual examination of each part that re- item is tested before use and it is de- quires— termined that the actual strength

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properties of that particular item will bearing factor is larger than the appli- equal or exceed those used in design. cable casting factor. (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. 5084, 29 FR 16150, Dec. 3, 1964, as (1) Each critical casting must— amended by Amdt. 29–17, 43 FR 50599, Oct. 30, (i) Have a casting factor of not less 1978; Amdt. 27–26, 55 FR 8003, Mar. 6, 1990] than 1.25; and § 29.619 Special factors. (ii) Receive 100 percent inspection by visual, radiographic, and magnetic par- (a) The special factors prescribed in ticle (for ferromagnetic materials) or §§ 29.621 through 29.625 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) For each critical casting with a (2) Likely to deteriorate in service casting factor less than 1.50, three sam- before normal replacement; or ple castings must be static tested and (3) Subject to appreciable variability shown to meet— due to— (i) The strength requirements of (i) Uncertainties in manufacturing § 29.305 at an ultimate load cor- processes; or responding to a casting factor of 1.25; (ii) Uncertainties in inspection meth- and ods. (ii) The deformation requirements of (b) For each part of the rotorcraft to § 29.305 at a load of 1.15 times the limit which §§ 29.621 through 29.625 apply, the load. factor of safety prescribed in § 29.303 (d) Noncritical castings. For each cast- must be multiplied by a special factor ing other than those specified in para- equal to— graph (c) of this section, the following (1) The applicable special factors pre- apply: scribed in §§ 29.621 through 29.625; or (1) Except as provided in paragraphs (2) Any other factor great enough to (d)(2) and (3) of this section, the casting ensure that the probability of the part factors and corresponding inspections being understrength because of the un- must meet the following table: certainties specified in paragraph (a) of this section is extremely remote. Casting factor Inspection

§ 29.621 Casting factors. 2.0 or greater ...... 100 percent visual. Less than 2.0, greater 100 percent visual, and magnetic (a) General. The factors, tests, and in- than 1.5. particle (ferromagnetic materials), spections specified in paragraphs (b) penetrant (nonferromagnetic ma- terials), or approved equivalent and (c) of this section must be applied inspection methods. in addition to those necessary to estab- 1.25 through 1.50 ...... 100 percent visual, and magnetic lish foundry quality control. The in- particle (ferromagnetic materials), spections must meet approved speci- penetrant (nonferromagnetic ma- terials), and radiographic or ap- fications. Paragraphs (c) and (d) of this proved equivalent inspection section apply to structural castings ex- methods. cept castings that are pressure tested as parts of hydraulic or other fluid sys- (2) The percentage of castings in- tems and do not support structural spected by nonvisual methods may be loads. reduced below that specified in para- (b) Bearing stresses and surfaces. The graph (d)(1) of this section when an ap- casting factors specified in paragraphs proved quality control procedure is es- (c) and (d) of this section— tablished. (1) Need not exceed 1.25 with respect (3) For castings procured to a speci- to bearing stresses regardless of the fication that guarantees the mechan- method of inspection used; and ical properties of the material in the (2) Need not be used with respect to casting and provides for demonstration the bearing surfaces of a part whose of these properties by test of coupons

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cut from the castings on a sampling § 29.561(b)(3) multiplied by a fitting fac- basis— tor of 1.33. (i) A casting factor of 1.0 may be [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as used; and amended by Amdt. 29–42, 63 FR 43285, Aug. 12, (ii) The castings must be inspected as 1998] provided in paragraph (d)(1) of this sec- tion for casting factors of ‘‘1.25 through § 29.629 Flutter and divergence. 1.50’’ and tested under paragraph (c)(2) Each aerodynamic surface of the of this section. rotorcraft must be free from flutter [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as and divergence under each appropriate amended by Amdt. 29–41, 62 FR 46173, Aug. 29, speed and power condition. 1997] [Doc. No. 28008, 61 FR 21907, May 10, 1996] § 29.623 Bearing factors. § 29.631 Bird strike. (a) Except as provided in paragraph (b) of this section, each part that has The rotorcraft must be designed to clearance (free fit), and that is subject ensure capability of continued safe to pounding or vibration, must have a flight and landing (for Category A) or bearing factor large enough to provide safe landing (for Category B) after im- for the effects of normal relative mo- pact with a 2.2-lb (1.0 kg) bird when the tion. velocity of the rotorcraft (relative to (b) No bearing factor need be used on the bird along the flight path of the a part for which any larger special fac- rotorcraft) is equal to VNE or VH tor is prescribed. (whichever is the lesser) at altitudes up to 8,000 feet. Compliance must be § 29.625 Fitting factors. shown by tests or by analysis based on For each fitting (part or terminal tests carried out on sufficiently rep- used to join one structural member to resentative structures of similar de- another) the following apply: sign. (a) For each fitting whose strength is [Doc. No. 28008, 61 FR 21907, May 10, 1996; 61 not proven by limit and ultimate load FR 33963, July 1, 1996] tests in which actual stress conditions are simulated in the fitting and sur- ROTORS rounding structures, a fitting factor of at least 1.15 must be applied to each § 29.653 Pressure venting and drain- part of— age of rotor blades. (1) The fitting; (a) For each rotor blade— (2) The means of attachment; and (1) There must be means for venting (3) The bearing on the joined mem- the internal pressure of the blade; bers. (2) Drainage holes must be provided (b) No fitting factor need be used— for the blade; and (1) For joints made under approved (3) The blade must be designed to pre- practices and based on comprehensive vent water from becoming trapped in test data (such as continuous joints in it. metal plating, welded joints, and scarf (b) Paragraphs (a)(1) and (2) of this joints in wood); and section does not apply to sealed rotor (2) With respect to any bearing sur- blades capable of withstanding the face for which a larger special factor is maximum pressure differentials ex- used. pected in service. (c) For each integral fitting, the part must be treated as a fitting up to the [Amdt. 29–3, 33 FR 967, Jan. 26, 1968] point at which the section properties § 29.659 Mass balance. become typical of the member. (d) Each seat, berth, litter, safety (a) The rotor and blades must be belt, and harness attachment to the mass balanced as necessary to— structure must be shown by analysis, (1) Prevent excessive vibration; and tests, or both, to be able to withstand (2) Prevent flutter at any speed up to the inertia forces prescribed in the maximum forward speed.

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(b) The structural integrity of the power-operated system is necessary to mass balance installation must be sub- show compliance with the flight char- stantiated. acteristics requirements of this part, [Amdt. 29–3, 33 FR 967, Jan. 26, 1968] the system must comply with § 29.671 of this part and the following: § 29.661 Rotor blade clearance. (a) A warning which is clearly distin- There must be enough clearance be- guishable to the pilot under expected tween the rotor blades and other parts flight conditions without requiring the of the structure to prevent the blades pilot’s attention must be provided for from striking any part of the structure any failure in the stability augmenta- during any operating condition. tion system or in any other automatic or power-operated system which could [Amdt. 29–3, 33 FR 967, Jan. 26, 1968] result in an unsafe condition if the § 29.663 Ground resonance prevention pilot is unaware of the failure. Warning means. systems must not activate the control (a) The reliability of the means for systems. preventing ground resonance must be (b) The design of the stability aug- shown either by analysis and tests, or mentation system or of any other auto- reliable service experience, or by show- matic or power-operated system must ing through analysis or tests that mal- allow initial counteraction of failures function or failure of a single means without requiring exceptional pilot will not cause ground resonance. skill or strength, by overriding the (b) The probable range of variations, failure by moving the flight controls in during service, of the damping action the normal sense, and by deactivating of the ground resonance prevention the failed system. means must be established and must be (c) It must be show that after any investigated during the test required single failure of the stability aug- by § 29.241. mentation system or any other auto- [Amdt. 27–26, 55 FR 8003, Mar. 6, 1990] matic or power-operated system— (1) The rotorcraft is safely control- CONTROL SYSTEMS lable when the failure or malfunction occurs at any speed or altitude within § 29.671 General. the approved operating limitations; (a) Each control and control system (2) The controllability and maneuver- must operate with the ease, smooth- ability requirements of this part are ness, and positiveness appropriate to met within a practical operational its function. flight envelope (for example, speed, al- (b) Each element of each flight con- titude, normal acceleration, and rotor- trol system must be designed, or dis- craft configurations) which is described tinctively and permanently marked, to in the Rotorcraft Flight Manual; and minimize the probability of any incor- (3) The trim and stability character- rect assembly that could result in the istics are not impaired below a level malfunction of the system. needed to allow continued safe flight (c) A means must be provided to and landing. allow full control movement of all pri- mary flight controls prior to flight, or [Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] a means must be provided that will allow the pilot to determine that full § 29.673 Primary flight controls. control authority is available prior to Primary flight controls are those flight. used by the pilot for immediate control [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as of pitch, roll, yaw, and vertical motion amended by Amdt. 29–24, 49 FR 44437, Nov. 6, of the rotorcraft. 1984] [Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] § 29.672 Stability augmentation, auto- matic, and power-operated systems. § 29.674 Interconnected controls. If the functioning of stability aug- Each primary flight control system mentation or other automatic or must provide for safe flight and landing

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and operate independently after a mal- (2) Each fitting, pulley, and bracket function, failure, or jam of any auxil- used in attaching the system to the iary interconnected control. main structure is included; (b) Compliance must be shown (by [Amdt. 27–26, 55 FR 8003, Mar. 6, 1990] analyses or individual load tests) with § 29.675 Stops. the special factor requirements for control system joints subject to angu- (a) Each control system must have lar motion. stops that positively limit the range of motionof the pilot’s controls. § 29.683 Operation tests. (b) Each stop must be located in the It must be shown by operation tests system so that the range of travel of that, when the controls are operated its control is not appreciably affected from the pilot compartment with the by— control system loaded to correspond (1) Wear; with loads specified for the system, the (2) Slackness; or system is free from— (3) Takeup adjustments. (a) Jamming; (c) Each stop must be able to with- (b) Excessive friction; and stand the loads corresponding to the (c) Excessive deflection. design conditions for the system. (d) For each main rotor blade— § 29.685 Control system details. (1) Stops that are appropriate to the (a) Each detail of each control sys- blade design must be provided to limit tem must be designed to prevent jam- travel of the blade about its hinge ming, chafing, and interference from points; and cargo, passengers, loose objects, or the (2) There must be means to keep the freezing of moisture. blade from hitting the droop stops dur- (b) There must be means in the cock- ing any operation other than starting pit to prevent the entry of foreign ob- and stopping the rotor. jects into places where they would jam the system. (Secs. 313(a), 601, 603, 604, Federal Aviation (c) There must be means to prevent Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 the slapping of cables or tubes against U.S.C. 1655(c))) other parts. (d) Cable systems must be designed [Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as as follows: amended by Amdt. 29–17, 43 FR 50599, Oct. 30, (1) Cables, cable fittings, turn- 1978] buckles, splices, and pulleys must be of § 29.679 Control system locks. an acceptable kind. (2) The design of cable systems must If there is a device to lock the con- prevent any hazardous change in cable trol system with the rotorcraft on the tension throughout the range of travel ground or water, there must be means under any operating conditions and to— temperature variations. (a) Automatically disengage the lock (3) No cable smaller than 1⁄8 inch di- when the pilot operates the controls in ameter may be used in any primary a normal manner, or limit the oper- control system. ation of the rotorcraft so as to give un- (4) Pulley kinds and sizes must cor- mistakable warning to the pilot before respond to the cables with which they takeoff; and are used. The pulley-cable combina- (b) Prevent the lock from engaging in tions and strength values specified in flight. MIL–HDBK–5 must be used unless they are inapplicable. § 29.681 Limit load static tests. (5) Pulleys must have close fitting (a) Compliance with the limit load guards to prevent the cables from being requirements of this part must be displaced or fouled. shown by tests in which— (6) Pulleys must lie close enough to (1) The direction of the test loads the plane passing through the cable to produces the most severe loading in the prevent the cable from rubbing against control system; and the pulley flange.

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(7) No fairlead may cause a change in source (such as hydrualic pumps), and cable direction of more than three de- such items as valves, lines, and actu- grees. ators. (8) No clevis pin subject to load or (c) The failure of mechanical parts motion and retained only by cotter (such as piston rods and links), and the pins may be used in the control sys- jamming of power cylinders, must be tem. considered unless they are extremely (9) Turnbuckles attached to parts improbable. having angular motion must be in- stalled to prevent binding throughout LANDING GEAR the range of travel. (10) There must be means for visual § 29.723 Shock absorption tests. inspection at each fairlead, pulley, ter- The landing inertia load factor and minal, and turnbuckle. the reserve energy absorption capacity (e) Control system joints subject to of the landing gear must be substan- angular motion must incorporate the tiated by the tests prescribed in following special factors with respect §§ 29.725 and 29.727, respectively. These to the ultimate bearing strength of the tests must be conducted on the com- softest material used as a bearing: plete rotorcraft or on units consisting (1) 3.33 for push-pull systems other of wheel, tire, and shock absorber in than ball and roller bearing systems. their proper relation. (2) 2.0 for cable systems. (f) For control system joints, the § 29.725 Limit drop test. manufacturer’s static, non-Brinell rat- The limit drop test must be con- ing of ball and roller bearings may not ducted as follows: be exceeded. (a) The drop height must be at least 8 inches. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55471, Dec. 20, (b) If considered, the rotor lift speci- 1976] fied in § 29.473(a) must be introduced into the drop test by appropriate en- § 29.687 Spring devices. ergy absorbing devices or by the use of (a) Each control system spring device an effective mass. whose failure could cause flutter or (c) Each landing gear unit must be other unsafe characteristics must be tested in the attitude simulating the reliable. landing condition that is most critical (b) Compliance with paragraph (a) of from the standpoint of the energy to be this section must be shown by tests absorbed by it. simulating service conditions. (d) When an effective mass is used in showing compliance with paragraph (b) § 29.691 Autorotation control mecha- of this section, the following formulae nism. may be used instead of more rational Each main rotor blade pitch control computations. mechanism must allow rapid entry into hd+−()1L autorotation after power failure. WW=× ; and e hd+ § 29.695 Power boost and power-oper- =+We ated control system. nnj L W (a) If a power boost or power-oper- where: ated control system is used, an alter- We=the effective weight to be used in the nate system must be immediately drop test (lbs.). available that allows continued safe W=WM for main gear units (lbs.), equal to the flight and landing in the event of— static reaction on the particular unit with (1) Any single failure in the power the rotorcraft in the most critical atti- portion of the system; or tude. A rational method may be used in (2) The failure of all engines. computing a main gear static reaction, taking into consideration the moment arm (b) Each alternate system may be a between the main wheel reaction and the duplicate power portion or a manually rotorcraft center of gravity. operated mechanical system. The W=WN for nose gear units (lbs.), equal to the power portion includes the power vertical component of the static reaction

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that would exist at the nose wheel, assum- (1) The loads occurring in any ma- ing that the mass of the rotorcraft acts at neuvering condition with the gear re- the center of gravity and exerts a force of tracted; 1.0g downward and 0.25g forward. (2) The combined friction, inertia, W=Wt for tailwheel units (lbs.) equal to whichever of the following is critical— and air loads occurring during retrac- tion and extension at any airspeed up (1) The static weight on the tailwheel with to the design maximum landing gear the rotorcraft resting on all wheels; or operating speed; and (2) The vertical component of the ground reaction that would occur at the tailwheel (3) The flight loads, including those assuming that the mass of the rotorcraft in yawed flight, occurring with the acts at the center of gravity and exerts a gear extended at any airspeed up to the force of 1g downward with the rotorcraft in design maximum landing gear extended the maximum nose-up attitude considered in speed. the nose-up landing conditions. (b) Landing gear lock. A positive h=specified free drop height (inches). means must be provided to keep the L=ratio of assumed rotor lift to the rotor- gear extended. craft weight. (c) Emergency operation. When other d=deflection under impact of the tire (at the than manual power is used to operate proper inflation pressure) plus the vertical the gear, emergency means must be component of the axle travel (inches) rel- provided for extending the gear in the ative to the drop mass. n=limit inertia load factor. event of— nj=the load factor developed, during impact, (1) Any reasonably probable failure in on the mass used in the drop test (i.e., the the normal retraction system; or acceleration dv/dt in g’ s recorded in the (2) The failure of any single source of drop test plus 1.0). hydraulic, electric, or equivalent en- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ergy. amended by Amdt. 29–3, 33 FR 967, Jan. 26. (d) Operation tests. The proper func- 1968] tioning of the retracting mechanism must be shown by operation tests. § 29.727 Reserve energy absorption (e) Position indicator. There must be drop test. means to indicate to the pilot when the The reserve energy absorption drop gear is secured in the extreme posi- test must be conducted as follows: tions. (a) The drop height must be 1.5 times (f) Control. The location and oper- that specified in § 29.725(a). ation of the retraction control must (b) Rotor lift, where considered in a meet the requirements of §§ 29.777 and manner similar to that prescribed in 29.779. § 29.725(b), may not exceed 1.5 times the (g) Landing gear warning. An aural or lift allowed under that paragraph. equally effective landing gear warning (c) The landing gear must withstand device must be provided that functions this test without collapsing. Collapse continuously when the rotorcraft is in of the landing gear occurs when a a normal landing mode and the landing member of the nose, tail, or main gear gear is not fully extended and locked. will not support the rotorcraft in the A manual shutoff capability must be proper attitude or allows the rotorcraft provided for the warning device and the structure, other than landing gear and warning system must automatically external accessories, to impact the reset when the rotorcraft is no longer landing surface. in the landing mode. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 27–26, 55 FR 8003, Mar. 6, amended by Amdt. 29–24, 49 FR 44437, Nov. 6, 1990] 1984]

§ 29.729 Retracting mechanism. § 29.731 Wheels. For rotorcraft with retractable land- (a) Each landing gear wheel must be ing gear, the following apply: approved. (a) Loads. The landing gear, retract- (b) The maximum static load rating ing mechanism, wheel well doors, and of each wheel may not be less than the supporting structure must be designed corresponding static ground reaction for— with—

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(1) Maximum weight; and (2) Hold the rotorcraft parked on a (2) Critical center of gravity. 10-degree slope on a dry, smooth pave- (c) The maximum limit load rating of ment. each wheel must equal or exceed the [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as maximum radial limit load determined amended by Amdt. 29–24, 49 FR 44437, Nov. 6, under the applicable ground load re- 1984] quirements of this part. § 29.737 Skis. § 29.733 Tires. (a) The maximum limit load rating of Each landing gear wheel must have a each ski must equal or exceed the max- tire— imum limit load determined under the (a) That is a proper fit on the rim of applicable ground load requirements of the wheel; and this part. (b) Of a rating that is not exceeded (b) There must be a stabilizing means under— to maintain the ski in an appropriate (1) The design maximum weight; position during flight. This means (2) A load on each main wheel tire must have enough strength to with- equal to the static ground reaction cor- stand the maximum aerodynamic and responding to the critical center of inertia loads on the ski. gravity; and FLOATS AND HULLS (3) A load on nose wheel tires (to be compared with the dynamic rating es- § 29.751 Main float buoyancy. tablished for those tires) equal to the (a) For main floats, the buoyancy reaction obtained at the nose wheel, necessary to support the maximum assuming that the mass of the rotor- weight of the rotorcraft in fresh water craft acts as the most critical center of must be exceeded by— gravity and exerts a force of 1.0 g down- (1) 50 percent, for single floats; and ward and 0.25 g forward, the reactions (2) 60 percent, for multiple floats. being distributed to the nose and main wheels according to the principles of (b) Each main float must have enough water-tight compartments so statics with the drag reaction at the that, with any single main float com- ground applied only at wheels with partment flooded, the mainfloats will brakes. provide a margin of positive stability (c) Each tire installed on a retract- great enough to minimize the prob- able landing gear system must, at the ability of capsizing. maximum size of the tire type expected in service, have a clearance to sur- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as rounding structure and systems that is amended by Amdt. 29–3, 33 FR 967, Jan. 26, adequate to prevent contact between 1968] the tire and any part of the structure § 29.753 Main float design. or systems. (a) Bag floats. Each bag float must be [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as designed to withstand— amended by Amdt. 29–12, 41 FR 55471, Dec. 20, (1) The maximum pressure differen- 1976] tial that might be developed at the § 29.735 Brakes. maximum altitude for which certifi- cation with that float is requested; and For rotorcraft with wheel-type land- (2) The vertical loads prescribed in ing gear, a braking device must be in- § 29.521(a), distributed along the length stalled that is— of the bag over three-quarters of its (a) Controllable by the pilot; projected area. (b) Usable during power-off landings; (b) Rigid floats. Each rigid float must and be able to withstand the vertical, hori- (c) Adequate to— zontal, and side loads prescribed in (1) Counteract any normal unbal- § 29.521. An appropriate load distribu- anced torque when starting or stopping tion under critical conditions must be the rotor; and used.

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§ 29.755 Hull buoyancy. § 29.773 Pilot compartment view. Water-based and amphibian rotorcraft. (a) Nonprecipitation conditions. For The hull and auxiliary floats, if used, nonprecipitation conditions, the fol- must have enough watertight compart- lowing apply: ments so that, with any single com- (1) Each pilot compartment must be partment of the hull or auxiliary floats arranged to give the pilots a suffi- flooded, the buoyancy of the hull and ciently extensive, clear, and undis- auxiliary floats, and wheel tires if torted view for safe operation. used, provides a margin of positive (2) Each pilot compartment must be water stability great enough to mini- free of glare and reflection that could mize the probability of capsizing the interfere with the pilot’s view. If cer- rotorcraft for the worst combination of tification for night operation is re- wave heights and surface winds for quested, this must be shown by night which approval is desired. flight tests. [Amdt. 29–3, 33 FR 967, Jan. 26, 1968; as (b) Precipitation conditions. For pre- amended by Amdt. 27–26, 55 FR 8003, Mar. 6, cipitation conditions, the following 1990] apply: (1) Each pilot must have a suffi- § 29.757 Hull and auxiliary float ciently extensive view for safe oper- strength. ation— The hull, and auxiliary floats if used, (i) In heavy rain at forward speeds up must withstand the water loads pre- to VH; and scribed by § 29.519 with a rational and (ii) In the most severe icing condi- conservative distribution of local and tion for which certification is re- distributed water pressures over the quested. hull and float bottom. (2) The first pilot must have a win- dow that— [Amdt. 29–3, 33 FR 967, Jan. 26, 1968] (i) Is openable under the conditions PERSONNEL AND CARGO prescribed in paragraph (b)(1) of this ACCOMMODATIONS section; and (ii) Provides the view prescribed in § 29.771 Pilot compartment. that paragraph. For each pilot compartment— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (a) The compartment and its equip- amended by Amdt. 29–3, 33 FR 967, Jan. 26, ment must allow each pilot to perform 1968] his duties without unreasonable con- centration or fatigue; § 29.775 Windshields and windows. (b) If there is provision for a second Windshields and windows must be pilot, the rotorcraft must be control- made of material that will not break lable with equal safety from either into dangerous fragments. pilot position. Flight and powerplant [Amdt. 29–31, 55 FR 38966, Sept. 21, 1990] controls must be designed to prevent confusion or inadvertent operation § 29.777 Cockpit controls. when the rotorcraft is piloted from ei- ther position; Cockpit controls must be— (c) The vibration and noise charac- (a) Located to provide convenient op- teristics of cockpit appurtenances may eration and to prevent confusion and not interfere with safe operation; inadvertent operation; and (d) Inflight leakage of rain or snow (b) Located and arranged with re- that could distract the crew or harm spect to the pilot’s seats so that there the structure must be prevented. is full and unrestricted movement of each control without interference from [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as the cockpit structure or the pilot’s amended by Amdt. 29–3, 33 FR 967, Jan. 26, clothing when pilots from 5′2″ to 6′0″ in 1968; Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] height are seated.

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§ 29.779 Motion and effect of cockpit (3) Sideward—2.0g. controls. (4) Downward—4.0g. Cockpit controls must be designed so (e) There must be means for direct that they operate in accordance with visual inspection of the locking mecha- the following movements and actu- nism by crewmembers to determine ation: whether the external doors (including (a) Flight controls, including the col- passenger, crew, service, and cargo lective pitch control, must operate doors) are fully locked. There must be with a sense of motion which cor- visual means to signal to appropriate responds to the effect on the rotor- crewmembers when normally used ex- craft. ternal doors are closed and fully (b) Twist-grip engine power controls locked. must be designed so that, for lefthand (f) For outward opening external operation, the motion of the pilot’s doors usable for entrance or egress, hand is clockwise to increase power there must be an auxiliary safety when the hand is viewed from the edge latching device to prevent the door containing the index finger. Other en- from opening when the primary latch- gine power controls, excluding the col- ing mechanism fails. If the door does lective control, must operate with a not meet the requirements of para- forward motion to increase power. graph (c) of this section with this de- (c) Normal landing gear controls vice in place, suitable operating proce- must operate downward to extend the dures must be established to prevent landing gear. the use of the device during takeoff and landing. [Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] (g) If an integral stair is installed in a passenger entry door that is qualified § 29.783 Doors. as a passenger emergency exit, the (a) Each closed cabin must have at stair must be designed so that under least one adequate and easily acces- the following conditions the effective- sible external door. ness of passenger emergency egress will (b) Each external door must be lo- not be impaired: cated, and appropriate operating proce- (1) The door, integral stair, and oper- dures must be established, to ensure ating mechanism have been subjected that persons using the door will not be to the inertial forces specified in para- endangered by the rotors, propellers, graph (d) of this section, acting sepa- engine intakes, and exhausts when the rately relative to the surrounding operating procedures are used. structure. (c) There must be means for locking (2) The rotorcraft is in the normal crew and external passenger doors and ground attitude and in each of the atti- for preventing their opening in flight tudes corresponding to collapse of one inadvertently or as a result of mechan- or more legs, or primary members, as ical failure. It must be possible to open applicable, of the landing gear. external doors from inside and outside (h) Nonjettisonable doors used as the cabin with the rotorcraft on the ditching emergency exits must have ground even though persons may be means to enable them to be secured in crowded against the door on the inside the open position and remain secure for of the rotorcraft. The means of opening emergency egress in sea state condi- must be simple and obvious and so ar- tions prescribed for ditching. ranged and marked that it can be read- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ily located and operated. amended by Amdt. 29–20, 45 FR 60178, Sept. (d) There must be reasonable provi- 11, 1980; Amdt. 29–29, 54 FR 47320, Nov. 13, sions to prevent the jamming of any 1989; Amdt. 27–26, 55 FR 8003, Mar. 6, 1990; external doors in a minor crash as a re- Amdt. 29–31, 55 FR 38966, Sept. 21, 1990] sult of fuselage deformation under the following ultimate inertial forces ex- § 29.785 Seats, berths, litters, safety cept for cargo or service doors not suit- belts, and harnesses. able for use as an exit in an emergency: (a) Each seat, safety belt, harness, (1) Upward—1.5g. and adjacent part of the rotorcraft at (2) Forward—4.0g. each station designated for occupancy

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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 inertial factors specified in § 29.561(b) strength of the safety belt and shoulder and dynamic conditions specified in harness may not be less than that cor- § 29.562. responding to the inertial forces speci- (b) Each occupant must be protected fied in § 29.561(b), considering the occu- from serious head injury by a safety pant weight of at least 170 pounds, con- belt plus a shoulder harness that will sidering the dimensional characteris- prevent the head from contacting any tics of the restraint system installa- injurious object, except as provided for tion, and using a distribution of at in § 29.562(c)(5). A shoulder harness least a 60-percent load to the safety (upper torso restraint), in combination belt and at least a 40-percent load to with the safety belt, constitutes a the shoulder harness. If the safety belt torso restraint system as described in is capable of being used without the TSO–C114. shoulder harness, the inertial forces (c) Each occupant’s seat must have a specified must be met by the safety combined safety belt and shoulder har- belt alone. ness with a single-point release. Each (h) When a headrest is used, the head- pilot’s combined safety belt and shoul- rest and its supporting structure must der harness must allow each pilot when be designed to resist the inertia forces seated with safety belt and shoulder specified in § 29.561, with a 1.33 fitting harness fastened to perform all func- factor and a head weight of at least 13 pounds. tions necessary for flight operations. (i) Each seating device system in- There must be a means to secure belt cludes the device such as the seat, the and harness when not in use to prevent cushions, the occupant restraint sys- interference with the operation of the tem and attachment devices. rotorcraft and with rapid egress in an (j) Each seating device system may emergency. use design features such as crushing or (d) If seat backs do not have a firm separation of certain parts of the seat handhold, there must be hand grips or in the design to reduce occupant loads rails along each aisle to let the occu- for the emergency landing dynamic pants steady themselves while using conditions of § 29.562; otherwise, the the aisle in moderately rough air. system must remain intact and must (e) Each projecting object that would not interfere with rapid evacuation of injure persons seated or moving about the rotorcraft. in the rotorcraft in normal flight must (k) For purposes of this section, a lit- be padded. ter is defined as a device designed to (f) Each seat and its supporting carry a nonambulatory person, pri- structure must be designed for an occu- marily in a recumbent position, into pant weight of at least 170 pounds, con- and on the rotorcraft. Each berth or sidering the maximum load factors, in- litter must be designed to withstand ertial forces, and reactions between the the load reaction of an occupant occupant, seat, and safety belt or har- weight of at least 170 pounds when the ness corresponding with the applicable occupant is subjected to the forward flight and ground-load conditions, in- inertial factors specified in § 29.561(b). cluding the emergency landing condi- A berth or litter installed within 15° or tions of § 29.561(b). In addition— less of the longitudinal axis of the (1) Each pilot seat must be designed rotorcraft must be provided with a pad- for the reactions resulting from the ap- ded end-board, cloth diaphragm, or plication of the pilot forces prescribed equivalent means that can withstand in § 29.397; and the forward load reaction. A berth or (2) The inertial forces prescribed in litter oriented greater than 15° with § 29.561(b) must be multiplied by a fac- the longitudinal axis of the rotorcraft

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must be equipped with appropriate re- so as to prevent contact between lamp straints, such as straps or safety belts, bulb and cargo. to withstand the forward reaction. In addition— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55472, Dec. 20, (1) The berth or litter must have a re- 1976; Amdt. 29–31, 55 FR 38966, Sept. 21, 1990] straint system and must not have cor- ners or other protuberances likely to § 29.801 Ditching. cause serious injury to a person occu- pying it during emergency landing con- (a) If certification with ditching pro- ditions; and visions is requested, the rotorcraft (2) The berth or litter attachment must meet the requirements of this and the occupant restraint system at- section and §§ 29.807(d), 29.1411 and tachments to the structure must be de- 29.1415. signed to withstand the critical loads (b) Each practicable design measure, resulting from flight and ground load compatible with the general character- conditions and from the conditions pre- istics of the rotorcraft, must be taken scribed in § 29.561(b). The fitting factor to minimize the probability that in an required by § 29.625(d) shall be applied. emergency landing on water, the be- havior of the rotorcraft would cause [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as immediate injury to the occupants or amended by Amdt. 29–24, 49 FR 44437, Nov. 6, 1984; Amdt. 29–29, 54 FR 47320, Nov. 13, 1989; would make it impossible for them to Amdt. 29–42, 63 FR 43285, Aug. 12, 1998] escape. (c) The probable behavior of the § 29.787 Cargo and baggage compart- rotorcraft in a water landing must be ments. investigated by model tests or by com- (a) Each cargo and baggage compart- parison with rotorcraft of similar con- ment must be designed for its plac- figuration for which the ditching char- arded maximum weight of contents and acteristics are known. Scoops, flaps, for the critical load distributions at projections, and any other factors like- the appropriate maximum load factors ly to affect the hydrodynamic charac- corresponding to the specified flight teristics of the rotorcraft must be con- and ground load conditions, except the sidered. emergency landing conditions of (d) It must be shown that, under rea- § 29.561. sonably probable water conditions, the (b) There must be means to prevent flotation time and trim of the rotor- the contents of any compartment from craft will allow the occupants to leave becoming a hazard by shifting under the rotorcraft and enter the liferafts the loads specified in paragraph (a) of required by § 29.1415. If compliance with this section. this provision is shown by bouyancy (c) Under the emergency landing con- and trim computations, appropriate al- ditions of § 29.561, cargo and baggage lowances must be made for probable compartments must— structural damage and leakage. If the (1) Be positioned so that if the con- rotorcraft has fuel tanks (with fuel jet- tents break loose they are unlikely to tisoning provisions) that can reason- cause injury to the occupants or re- ably be expected to withstand a ditch- strict any of the escape facilities pro- ing without leakage, the jettisonable vided for use after an emergency land- volume of fuel may be considered as ing; or bouyancy volume. (2) Have sufficient strength to with- (e) Unless the effects of the collapse stand the conditions specified in of external doors and windows are ac- § 29.561, including the means of re- counted for in the investigation of the straint and their attachments required probable behavior of the rotorcraft in a by paragraph (b) of this section. Suffi- water landing (as prescribed in para- cient strength must be provided for the graphs (c) and (d) of this section), the maximum authorized weight of cargo external doors and windows must be and baggage at the critical loading dis- designed to withstand the probable tribution. maximum local pressures. (d) If cargo compartment lamps are installed, each lamp must be installed [Amdt. 29–12, 41 FR 55472, Dec. 20, 1976]

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§ 29.803 Emergency evacuation. located so as to allow rapid evacuation (a) Each crew and passenger area of the flight crew. This must be shown must have means for rapid evacuation by test. in a crash landing, with the landing (c) Each exit must not be obstructed gear (1) extended and (2) retracted, con- by water or flotation devices after a sidering the possibility of fire. ditching. This must be shown by test, (b) Passenger entrance, crew, and demonstration, or analysis. service doors may be considered as [Amdt. 29–3, 33 FR 968, Jan. 26, 1968; as emergency exits if they meet the re- amended by Amdt. 27–26, 55 FR 8004, Mar. 6, quirements of this section and of 1990] §§ 29.805 through 29.815. (c) [Reserved] § 29.807 Passenger emergency exits. (d) Except as provided in paragraph (a) Type. For the purpose of this part, (e) of this section, the following cat- the types of passenger emergency exit egories of rotorcraft must be tested in are as follows: accordance with the requirements of (1) Type I. This type must have a rec- appendix D of this part to demonstrate tangular opening of not less than 24 that the maximum seating capacity, inches wide by 48 inches high, with cor- including the crewmembers required by ner radii not greater than one-third the the operating rules, can be evacuated width of the exit, in the passenger area from the rotorcraft to the ground with- in the side of the fuselage at floor level in 90 seconds: and as far away as practicable from (1) Rotorcraft with a seating capacity areas that might become potential fire of more than 44 passengers. hazards in a crash. (2) Rotorcraft with all of the fol- (2) Type II. This type is the same as lowing: Type I, except that the opening must (i) Ten or more passengers per pas- be at least 20 inches wide by 44 inches senger exit as determined under high. § 29.807(b). (3) Type III. This type is the same as (ii) No main aisle, as described in Type I, except that— § 29.815, for each row of passenger seats. (i) The opening must be at least 20 (iii) Access to each passenger exit for inches wide by 36 inches high; and each passenger by virtue of design fea- (ii) The exits need not be at floor tures of seats, such as folding or break- level. over seat backs or folding seats. (4) Type IV. This type must have a (e) A combination of analysis and rectangular opening of not less than 19 tests may be used to show that the inches wide by 26 inches high, with cor- rotorcraft is capable of being evacu- ner radii not greater than one-third the ated within 90 seconds under the condi- width of the exit, in the side of the fu- tions specified in § 29.803(d) if the Ad- selage with a step-up inside the rotor- ministrator finds that the combination craft of not more than 29 inches. of analysis and tests will provide data, with respect to the emergency evacu- Openings with dimensions larger than ation capability of the rotorcraft, those specified in this section may be equivalent to that which would be ob- used, regardless of shape, if the base of tained by actual demonstration. the opening has a flat surface of not less than the specified width. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (b) Passenger emergency exits; side-of- amended by Amdt. 29–3, 33 FR 967, Jan. 26, fuselage. Emergency exits must be ac- 1968; Amdt. 27–26, 55 FR 8004, Mar. 6, 1990] cessible to the passengers and, except § 29.805 Flight crew emergency exits. as provided in paragraph (d) of this sec- tion, must be provided in accordance (a) For rotorcraft with passenger with the following table: emergency exits that are not conven- ient to the flight crew, there must be Emergency exits for each side of flight crew emergency exits, on both Passenger seating ca- the fuselage sides of the rotorcraft or as a top pacity Type I Type II Type III Type hatch, in the flight crew area. IV (b) Each flight crew emergency exit 1 through 10 ...... 1 must be of sufficient size and must be 11 through 19 ...... 1 or 2

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Emergency exits for each side of installed instead in the ramp of floor Passenger seating ca- the fuselage ramp rotorcraft if— pacity (1) Its installation in the side of the Type I Type II Type III Type IV fuselage is impractical; and 20 through 39 ...... 1 ...... 1 (2) Its installation in the ramp meets 40 through 59 ...... 1 ...... 1 § 29.813. 60 through 79 ...... 1 ...... 1 or 2 (f) Tests. The proper functioning of each emergency exit must be shown by (c) Passenger emergency exits; other test. than side-of-fuselage. In addition to the requirements of paragraph (b) of this [Amdt. 29–3, 33 FR 968, Jan. 26, 1968, as section— amended by Amdt. 29–12, 41 FR 55472, Dec. 20, (1) There must be enough openings in 1976; Amdt. 27–26, 55 FR 8004, Mar. 6, 1990] the top, bottom, or ends of the fuselage § 29.809 Emergency exit arrangement. to allow evacuation with the rotorcraft on its side; or (a) Each emergency exit must consist (2) The probability of the rotorcraft of a movable door or hatch in the ex- coming to rest on its side in a crash ternal walls of the fuselage and must landing must be extremely remote. provide an unobstructed opening to the (d) Ditching emergency exits for pas- outside. sengers. If certification with ditching (b) Each emergency exit must be provisions is requested, ditching emer- openable from the inside and from the gency exits must be provided in accord- outside. ance with the following requirements (c) The means of opening each emer- and must be proven by test, demonstra- gency exit must be simple and obvious tion, or analysis unless the emergency and may not require exceptional effort. exits required by paragraph (b) of this (d) There must be means for locking section already meet these require- each emergency exit and for preventing ments. opening in flight inadvertently or as a (1) For rotorcraft that have a pas- result of mechanical failure. senger seating configuration, excluding (e) There must be means to minimize pilots seats, of nine seats or less, one the probability of the jamming of any exit above the waterline in each side of emergency exit in a minor crash land- the rotorcraft, meeting at least the di- ing as a result of fuselage deformation mensions of a Type IV exit. under the ultimate inertial forces in (2) For rotorcraft that have a pas- § 29.783(d). senger seating configuration, excluding (f) Except as provided in paragraph pilots seats, of 10 seats or more, one (h) of this section, each land-based exit above the waterline in a side of the rotorcraft emergency exit must have rotorcraft meeting at least the dimen- an approved slide as stated in para- sions of a Type III exit, for each unit graph (g) of this section, or its equiva- (or part of a unit) of 35 passenger seats, lent, to assist occupants in descending but no less than two such exits in the to the ground from each floor level exit passenger cabin, with one on each side and an approved rope, or its equivalent, of the rotorcraft. However, where it for all other exits, if the exit threshold has been shown through analysis, is more that 6 feet above the ground— ditching demonstrations, or any other (1) With the rotorcraft on the ground tests found necessary by the Adminis- and with the landing gear extended; trator, that the evacuation capability (2) With one or more legs or part of of the rotorcraft during ditching is im- the landing gear collapsed, broken, or proved by the use of larger exits, or by not extended; and other means, the passenger seat to exit (3) With the rotorcraft resting on its ratio may be increased. side, if required by § 29.803(d). (3) Flotation devices, whether stowed (g) The slide for each passenger emer- or deployed, may not interfere with or gency exit must be a self-supporting obstruct the exits. slide or equivalent, and must be de- (e) Ramp exits. One Type I exit only, signed to meet the following require- or one Type II exit only, that is re- ments: quired in the side of the fuselage under (1) It must be automatically de- paragraph (b) of this section, may be ployed, and deployment must begin

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during the interval between the time (1) Withstand a 400-pound static load; the exit opening means is actuated and from inside the rotorcraft and the time (2) Attach to the fuselage structure the exit is fully opened. However, each at or above the top of the emergency passenger emergency exit which is also exit opening, or at another approved a passenger entrance door or a service location if the stowed rope would re- door must be provided with means to duce the pilot’s view in flight. prevent deployment of the slide when [Amdt. 29–3, 33 FR 968, Jan. 26, 1968, as the exit is opened from either the in- amended by Amdt. 29–29, 54 FR 47321, Nov. 13, side or the outside under non- 1989; Amdt. 27–26, 55 FR 8004, Mar. 6, 1990] emergency conditions for normal use. (2) It must be automatically erected § 29.811 Emergency exit marking. within 10 seconds after deployment is (a) Each passenger emergency exit, begun. its means of access, and its means of (3) It must be of such length after full opening must be conspicuously marked deployment that the lower end is self- for the guidance of occupants using the supporting on the ground and provides exits in daylight or in the dark. Such safe evacuation of occupants to the markings must be designed to remain ground after collapse of one or more visible for rotorcraft equipped for legs or part of the landing gear. overwater flights if the rotorcraft is (4) It must have the capability, in 25- capsized and the cabin is submerged. knot winds directed from the most (b) The identity and location of each critical angle, to deploy and, with the passenger emergency exit must be rec- assistance of only one person, to re- ognizable from a distance equal to the main usable after full deployment to width of the cabin. evacuate occupants safely to the (c) The location of each passenger ground. emergency exit must be indicated by a (5) Each slide installation must be sign visible to occupants approaching qualified by five consecutive deploy- along the main passenger aisle. There ment and inflation tests conducted (per must be a locating sign— exit) without failure, and at least three (1) Next to or above the aisle near tests of each such five-test series must each floor emergency exit, except that be conducted using a single representa- one sign may serve two exits if both ex- tive sample of the device. The sample ists can be seen readily from that sign; devices must be deployed and inflated and by the system’s primary means after (2) On each bulkhead or divider that being subjected to the inertia forces prevents fore and aft vision along the specified in § 29.561(b). If any part of the passenger cabin, to indicate emergency system fails or does not function prop- exits beyond and obscured by it, except erly during the required tests, the that if this is not possible the sign may cause of the failure or malfunction be placed at another appropriate loca- must be corrected by positive means tion. and after that, the full series of five (d) Each passenger emergency exit consecutive deployment and inflation marking and each locating sign must tests must be conducted without fail- have white letters 1 inch high on a red ure. background 2 inches high, be self or (h) For rotorcraft having 30 or fewer electrically illuminated, and have a passenger seats and having an exit minimum luminescence (brightness) of threshold more than 6 feet above the at least 160 microlamberts. The colors ground, a rope or other assist means may be reversed if this will increase may be used in place of the slide speci- the emergency illumination of the pas- fied in paragraph (f) of this section, senger compartment. provided an evacuation demonstration (e) The location of each passenger is accomplished as prescribed in emergency exit operating handle and § 29.803(d) or (e). instructions for opening must be (i) If a rope, with its attachment, is shown— used for compliance with paragraph (f), (1) For each emergency exit, by a (g), or (h) of this section, it must— marking on or near the exit that is

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readable from a distance of 30 inches; (a) A source of light with its power and supply independent of the main light- (2) For each Type I or Type II emer- ing system must be installed to— gency exit with a locking mechanism (1) Illuminate each passenger emer- released by rotary motion of the han- gency exit marking and locating sign; dle, by— and (i) A red arrow, with a shaft at least (2) Provide enough general lighting three-fourths inch wide and a head in the passenger cabin so that the aver- twice the width of the shaft, extending age illumination, when measured at 40- along at least 70 degrees of arc at a ra- inch intervals at seat armrest height dius approximately equal to three- on the center line of the main pas- fourths of the handle length; and senger aisle, is at least 0.05 foot-candle. (b) Exterior emergency lighting must (ii) The word ‘‘open’’ in red letters 1 be provided at each emergency exit. inch high, placed horizontally near the The illumination may not be less than head of the arrow. 0.05 foot-candle (measured normal to (f) Each emergency exit, and its the direction of incident light) for min- means of opening, must be marked on imum width on the ground surface, the outside of the rotorcraft. In addi- with landing gear extended, equal to tion, the following apply: the width of the emergency exit where (1) There must be a 2-inch colored an evacuee is likely to make first con- band outlining each passenger emer- tact with the ground outside the cabin. gency exit, except small rotorcraft The exterior emergency lighting may with a maximum weight of 12,500 be provided by either interior or exte- pounds or less may have a 2-inch col- rior sources with light intensity meas- ored band outlining each exit release urements made with the emergency lever or device of passenger emergency exits open. exits which are normally used doors. (c) Each light required by paragraph (2) Each outside marking, including (a) or (b) of this section must be oper- the band, must have color contrast to able manually from the cockpit station be readily distinguishable from the sur- and from a point in the passenger com- rounding fuselage surface. The contrast partment that is readily accessible. must be such that, if the reflectance of The cockpit control device must have the darker color is 15 percent or less, an ‘‘on,’’ ‘‘off,’’ and ‘‘armed’’ position the reflectance of the lighter color so that when turned on at the cockpit must be at least 45 percent. ‘‘Reflec- or passenger compartment station or tance’’ is the ratio of the luminous flux when armed at the cockpit station, the reflected by a body to the luminous emergency lights will either illuminate flux it receives. When the reflectance or remain illuminated upon interrup- of the darker color is greater than 15 tion of the rotorcraft’s normal electric percent, at least a 30 percent difference power. between its reflectance and the reflec- (d) Any means required to assist the tance of the lighter color must be pro- occupants in descending to the ground vided. must be illuminated so that the erect- (g) Exits marked as such, though in ed assist means is visible from the excess of the required number of exits, rotorcraft. must meet the requirements for emer- (1) The assist means must be pro- vided with an illumination of not less gency exits of the particular type. than 0.03 foot-candle (measured normal Emergency exits need only be marked to the direction of the incident light) with the word ‘‘Exit.’’ at the ground end of the erected assist [Amdt. 29–3, 33 FR 968, Jan. 26, 1968, as means where an evacuee using the es- amended by Amdt. 29–24, 49 FR 44438, Nov. 6, tablished escape route would normally 1984; Amdt. 27–26, 55 FR 8004, Mar. 6, 1990; make first contact with the ground, Amdt. 29–31, 55 FR 38967, Sept. 21, 1990] with the rotorcraft in each of the atti- tudes corresponding to the collapse of § 29.812 Emergency lighting. one or more legs of the landing gear. For transport Category A rotorcraft, (2) If the emergency lighting sub- the following apply: system illuminating the assist means

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is independent of the rotorcraft’s main to maintain the effectiveness of the emergency lighting system, it— exit. (i) Must automatically be activated [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as when the assist means is erected; amended by Amdt. 29–12, 41 FR 55472, Dec. 20, (ii) Must provide the illumination re- 1976] quired by paragraph (d)(1); and (iii) May not be adversely affected by § 29.815 Main aisle width. stowage. The main passenger aisle width be- (e) The energy supply to each emer- tween seats must equal or exceed the gency lighting unit must provide the values in the following table:

required level of illumination for at Minimum main passenger least 10 minutes at the critical ambient aisle width conditions after an emergency landing. Passenger seating capacity Less than 25 Inches (f) If storage batteries are used as the 25 inches and more from floor from floor energy supply for the emergency light- (inches) (inches) ing system, they may be recharged 10 or less ...... 12 15 from the rotorcraft’s main electrical 11 through 19 ...... 12 20 power system provided the charging 20 or more ...... 15 20 circuit is designed to preclude inad- 1 A narrower width not less than 9 inches may be approved vertent battery discharge into charg- when substantiated by tests found necessary by the ing circuit faults. Administrator. [Amdt. 29–24, 49 FR 44438, Nov. 6, 1984] [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55472, Dec. 20, § 29.813 Emergency exit access. 1976] (a) Each passageway between pas- § 29.831 Ventilation. senger compartments, and each pas- (a) Each passenger and crew compart- sageway leading to Type I and Type II ment must be ventilated, and each emergency exits, must be— crew compartment must have enough (1) Unobstructed; and fresh air (but not less than 10 cu. ft. per (2) At least 20 inches wide. minute per crewmember) to let crew- (b) For each emergency exit covered members perform their duties without by § 29.809(f), there must be enough undue discomfort or fatigue. space adjacent to that exit to allow a (b) Crew and passenger compartment crewmember to assist in the evacu- air must be free from harmful or haz- ation of passengers without reducing ardous concentrations of gases or va- the unobstructed width of the passage- pors. way below that required for that exit. (c) The concentration of carbon mon- (c) There must be access from each oxide may not exceed one part in 20,000 aisle to each Type III and Type IV exit, parts of air during forward flight. If the concentration exceeds this value under and other conditions, there must be suit- (1) For rotorcraft that have a pas- able operating restrictions. senger seating configuration, excluding (d) There must be means to ensure pilot seats, of 20 or more, the projected compliance with paragraphs (b) and (c) opening of the exit provided must not of this section under any reasonably be obstructed by seats, berths, or other probable failure of any ventilating, protrusions (including seatbacks in any heating, or other system or equipment. position) for a distance from that exit of not less than the width of the nar- § 29.833 Heaters. rowest passenger seat installed on the Each combustion heater must be ap- rotorcraft; proved. (2) For rotorcraft that have a pas- senger seating configuration, excluding FIRE PROTECTION pilot seats, of 19 or less, there may be minor obstructions in the region de- § 29.851 Fire extinguishers. scribed in paragraph (c)(1) of this sec- (a) Hand fire extinguishers. For hand tion, if there are compensating factors fire extinguishers the following apply:

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(1) Each hand fire extinguisher must conduit, thermal and acoustical insula- be approved. tion and insulation covering, air duct- (2) The kinds and quantities of each ing, joint and edge covering, cargo extinguishing agent used must be ap- compartment liners, insulation blan- propriate to the kinds of fires likely to kets, cargo covers, and transparencies, occur where that agent is used. molded and thermoformed parts, air (3) Each extinguisher for use in a per- ducting joints, and trim strips (decora- sonnel compartment must be designed tive and chafing) that are constructed to minimize the hazard of toxic gas of materials not covered in paragraph concentrations. (a)(3) of this section, must be self ex- (b) Built-in fire extinguishers. If a tinguishing when tested vertically in built-in fire extinguishing system is re- accordance with the applicable portion quired— of appendix F of Part 25 of this chapter, (1) The capacity of each system, in or other approved equivalent methods. relation to the volume of the compart- The average burn length may not ex- ment where used and the ventilation ceed 8 inches and the average flame rate, must be adequate for any fire time after removal of the flame source likely to occur in that compartment. may not exceed 15 seconds. Drippings (2) Each system must be installed so from the test specimen may not con- that— tinue to flame for more than an aver- (i) No extinguishing agent likely to age of 5 seconds after falling. enter personnel compartments will be (3) Acrylic windows and signs, parts present in a quantity that is hazardous constructed in whole or in part of to the occupants; and elastometric materials, edge lighted (ii) No discharge of the extinguisher can cause structural damage. instrument assemblies consisting of two or more instruments in a common § 29.853 Compartment interiors. housing, seat belts, shoulder harnesses, and cargo and baggage tiedown equip- For each compartment to be used by ment, including containers, bins, pal- the crew or passengers— lets, etc., used in passenger or crew (a) The materials (including finishes compartments, may not have an aver- or decorative surfaces applied to the materials) must meet the following age burn rate greater than 2.5 inches test criteria as applicable: per minute when tested horizontally in (1) Interior ceiling panels, interior accordance with the applicable por- wall panels, partitions, galley struc- tions of appendix F of Part 25 of this ture, large cabinet walls, structural chapter, or other approved equivalent flooring, and materials used in the con- methods. struction of stowage compartments (4) Except for electrical wire and (other than underseat stowage com- cable insulation, and for small parts partments and compartments for stow- (such as knobs, handles, rollers, fas- ing small items such as magazines and teners, clips, grommets, rub strips, pul- maps) must be self-extinguishing when leys, and small electrical parts) that tested vertically in accordance with the Administrator finds would not con- the applicable portions of appendix F tribute significantly to the propaga- of Part 25 of this chapter, or other ap- tion of a fire, materials in items not proved equivalent methods. The aver- specified in paragraphs (a)(1), (a)(2), or age burn length may not exceed 6 (a)(3) of this section may not have a inches and the average flame time burn rate greater than 4 inches per after removal of the flame source may minute when tested horizontally in ac- not exceed 15 seconds. Drippings from cordance with the applicable portions the test specimen may not continue to of appendix F of Part 25 of this chapter, flame for more than an average of 3 or other approved equivalent methods. seconds after falling. (b) In addition to meeting the re- (2) Floor covering, textiles (including quirements of paragraph (a)(2), seat draperies and upholstery), seat cush- cushions, except those on flight crew- ions, padding, decorative and non- member seats, must meet the test re- decorative coated fabrics, leather, quirements of Part II of appendix F of trays and galley furnishings, electrical Part 25 of this chapter, or equivalent.

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(c) If smoking is to be prohibited, for cargo or baggage compartments in there must be a placard so stating, and which— if smoking is to be allowed— (i) The presence of a compartment (1) There must be an adequate num- fire would be easily discovered by a ber of self-contained, removable ash- crewmember while at the crew- trays; and member’s station; (2) Where the crew compartment is (ii) Each part of the compartment is separated from the passenger compart- easily accessible in flight; ment, there must be at least one illu- (iii) The compartment has a volume minated sign (using either letters or of 200 cubic feet or less; and symbols) notifying all passengers when (iv) Notwithstanding § 29.1439(a), pro- smoking is prohibited. Signs which no- tective breathing equipment is not re- tify when smoking is prohibited must— quired. (i) When illuminated, be legible to each passenger seated in the passenger (b) No compartment may contain any cabin under all probable lighting condi- controls, wiring, lines, equipment, or tions; and accessories whose damage or failure (ii) Be so constructed that the crew would affect safe operation, unless can turn the illumination on and off. those items are protected so that— (d) Each receptacle for towels, paper, (1) They cannot be damaged by the or waste must be at least fire-resistant movement of cargo in the compart- and must have means for containing ment; and possible fires; (2) Their breakage or failure will not (e) There must be a hand fire extin- create a fire hazard. guisher for the flight crewmembers; (c) The design and sealing of inacces- and sible compartments must be adequate (f) At least the following number of to contain compartment fires until a hand fire extinguishers must be con- landing and safe evacuation can be veniently located in passenger com- made. partments: (d) Each cargo and baggage compart-

Fire extin- ment that is not sealed so as to contain Passenger capacity guishers cargo compartment fires completely without endangering the safety of a 7 through 30 ...... 1 31 through 60 ...... 2 rotorcraft or its occupants must be de- 61 or more ...... 3 signed, or must have a device, to en- sure detection of fires or smoke by a (Secs. 313(a), 601, 603, 604, Federal Aviation crewmember while at his station and Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), to prevent the accumulation of harm- sec. 6(c), Dept. of Transportation Act (49 ful quantities of smoke, flame, extin- U.S.C. 1655(c))) guishing agents, and other noxious [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as gases in any crew or passenger com- amended by Amdt. 29–3, 33 FR 969, Jan. 26, partment. This must be shown in 1968; Amdt. 29–17, 43 FR 50600, Oct. 30, 1978; flight. Amdt 29–18, 45 FR 7756, Feb. 4, 1980; Amdt. 29– (e) For rotorcraft used for the car- 23, 49 FR 43200, Oct. 26, 1984] riage of cargo only, the cabin area may § 29.855 Cargo and baggage compart- be considered a cargo compartment ments. and, in addition to paragraphs (a) (a) Each cargo and baggage compart- through (d) of this section, the fol- ment must be construced of or lined lowing apply: with materials in accordance with the (1) There must be means to shut off following: the ventilating airflow to or within the (1) For accessible and inaccessible compartment. Controls for this purpose compartments not occupied by pas- must be accessible to the flight crew in sengers or crew, the material must be the crew compartment. at least fire resistant. (2) Required crew emergency exits (2) Materials must meet the require- must be accessible under all cargo ments in § 29.853(a)(1), (a)(2), and (a)(3) loading conditions.

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(3) Sources of heat within each com- stream unless flames from backfires or partment must be shielded and insu- reverse burning cannot enter the ven- lated to prevent igniting the cargo. tilating airstream under any operating condition, including reverse flow or [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–3, 33 FR 969, Jan 26, malfunction of the heater or its associ- 1968; Amdt. 29–24, 49 FR 44438, Nov. 6, 1984; ated components; and Amdt. 27–26, 55 FR 8004, Mar. 6, 1990] (2) No combustion air duct may re- strict the prompt relief of any backfire § 29.859 Combustion heater fire pro- that, if so restricted, could cause heat- tection. er failure. (a) Combustion heater fire zones. The (d) Heater controls; general. There following combustion heater fire zones must be means to prevent the haz- must be protected against fire under ardous accumulation of water or ice on the applicable provisions of §§ 29.1181 or in any heater control component, through 29.1191, and 29.1195 through control system tubing, or safety con- 29.1203: trol. (1) The region surrounding any heat- (e) Heater safety controls. For each er, if that region contains any flam- combustion heater, safety control mable fluid system components (in- means must be provided as follows: cluding the heater fuel system), that (1) Means independent of the compo- could— nents provided for the normal contin- (i) Be damaged by heater malfunc- uous control of air temperature, air- tioning; or flow, and fuel flow must be provided, (ii) Allow flammable fluids or vapors for each heater, to automatically shut to reach the heater in case of leakage. off the ignition and fuel supply of that (2) Each part of any ventilating air heater at a point remote from that passage that— heater when any of the following oc- (i) Surrounds the combustion cham- curs: ber; and (i) The heat exchanger temperature (ii) Would not contain (without dam- exceeds safe limits. age to other rotorcraft components) (ii) The ventilating air temperature any fire that may occur within the pas- exceeds safe limits. sage. (iii) The combustion airflow becomes (b) Ventilating air ducts. Each ven- inadequate for safe operation. tilating air duct passing through any (iv) The ventilating airflow becomes fire zone must be fireproof. In addi- inadequate for safe operation. tion— (2) The means of complying with (1) Unless isolation is provided by paragraph (e)(1) of this section for any fireproof valves or by equally effective individual heater must— means, the ventilating air duct down- (i) Be independent of components stream of each heater must be fireproof serving any other heater whose heat for a distance great enough to ensure output is essential for safe operation; that any fire originating in the heater and can be contained in the duct; and (ii) Keep the heater off until re- (2) Each part of any ventilating duct started by the crew. passing through any region having a (3) There must be means to warn the flammable fluid system must be so crew when any heater whose heat out- constructed or isolated from that sys- put is essential for safe operation has tem that the malfunctioning of any been shut off by the automatic means component of that system cannot in- prescribed in paragraph (e)(1) of this troduce flammable fluids or vapors section. into the ventilating airstream. (f) Air intakes. Each combustion and (c) Combustion air ducts. Each com- ventilating air intake must be where bustion air duct must be fireproof for a no flammable fluids or vapors can distance great enough to prevent dam- enter the heater system under any op- age from backfiring or reverse flame erating condition— propagation. In addition— (1) During normal operation; or (1) No combustion air duct may com- (2) As a result of the malfunction of municate with the ventilating air- any other component.

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(g) Heater exhaust. Each heater ex- § 29.863 Flammable fluid fire protec- haust system must meet the require- tion. ments of §§ 29.1121 and 29.1123. In addi- (a) In each area where flammable tion— fluids or vapors might escape by leak- (1) Each exhaust shroud must be age of a fluid system, there must be sealed so that no flammable fluids or means to minimize the probability of hazardous quantities of vapors can ignition of the fluids and vapors, and reach the exhaust systems through the resultant hazards if ignition does joints; and occur. (2) No exhaust system may restrict (b) Compliance with paragraph (a) of the prompt relief of any backfire that, this section must be shown by analysis if so restricted, could cause heater fail- or tests, and the following factors must ure. be considered: (1) Possible sources and paths of fluid (h) Heater fuel systems. Each heater leakage, and means of detecting leak- fuel system must meet the powerplant age. fuel system requirements affecting safe (2) Flammability characteristics of heater operation. Each heater fuel sys- fluids, including effects of any combus- tem component in the ventilating air- tible or absorbing materials. stream must be protected by shrouds (3) Possible ignition sources, includ- so that no leakage from those compo- ing electrical faults, overheating of nents can enter the ventilating air- equipment, and malfunctioning of pro- stream. tective devices. (i) Drains. There must be means for (4) Means available for controlling or safe drainage of any fuel that might ac- extinguishing a fire, such as stopping cumulate in the combustion chamber flow of fluids, shutting down equip- or the heat exchanger. In addition— ment, fireproof containment, or use of (1) Each part of any drain that oper- extinguishing agents. (5) Ability of rotorcraft components ates at high temperatures must be pro- that are critical to safety of flight to tected in the same manner as heater withstand fire and heat. exhausts; and (c) If action by the flight crew is re- (2) Each drain must be protected quired to prevent or counteract a fluid against hazardous ice accumulation fire (e.g. equipment shutdown or actu- under any operating condition. ation of a fire extinguisher), quick act- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ing means must be provided to alert amended by Amdt. 29–2, 32 FR 6914, May 5, the crew. 1967] (d) Each area where flammable fluids or vapors might escape by leakage of a § 29.861 Fire protection of structure, fluid system must be identified and de- controls, and other parts. fined. Each part of the structure, controls, (Secs. 313(a), 601, 603, 604, Federal Aviation and the rotor mechanism, and other Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), parts essential to controlled landing sec. 6(c), Dept. of Transportation Act (49 U.S.C. 1655(c))) and (for category A) flight that would be affected by powerplant fires must be [Amdt. 29–17, 43 FR 50600, Oct. 30, 1978] isolated under § 29.1191, or must be— EXTERNAL LOADS (a) For category A rotorcraft, fire- proof; and § 29.865 External loads. (b) For Category B rotorcraft, fire- (a) It must be shown by analysis, proof or protected so that they can per- test, or both, that the rotorcraft exter- form their essential functions for at nal load attaching means for rotor- least 5 minutes under any foreseeable craft-load combinations to be used for powerplant fire conditions. nonhuman external cargo applications can withstand a limit static load equal [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 27–26, 55 FR 8005, Mar. 6, to 2.5, or some lower load factor ap- 1990] proved under §§ 29.337 through 29.341, multiplied by the maximum external

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load for which authorization is re- (i) Be reliable, durable, and function quested. It must be shown by analysis, properly with all external loads up to test, or both that the rotorcraft exter- and including the maximum external nal load attaching means and cor- limit load for which authorization is responding personnel carrying device requested. system for rotorcraft-load combina- (ii) Be protected against electro- tions to be used for human external magnetic interference (EMI) from ex- cargo applications can withstand a ternal and internal sources and against limit static load equal to 3.5 or some lightning to prevent inadvertent load lower load factor, not less than 2.5, ap- release. proved under §§ 29.337 through 29.341, (A) The minimum level of protection multiplied by the maximum external required for jettisonable rotorcraft- load for which authorization is re- load combinations used for nonhuman quested. The load for any rotorcraft- external cargo is a radio frequency load combination class, for any exter- field strength of 20 volts per meter. nal cargo type, must be applied in the (B) The minimum level of protection vertical direction. For jettisonable ex- required for jettisonable rotorcraft- ternal loads of any applicable external load combinations used for human ex- cargo type, the load must also be ap- ternal cargo is a radio frequency field plied in any direction making the max- strength of 200 volts per meter. imum angle with the vertical that can (iii) Be protected against any failure be achieved in service but not less than that could be induced by a failure mode 30°. However, the 30° angle may be re- of any other electrical or mechanical duced to a lesser angle if— rotorcraft system. (1) An operating limitation is estab- (c) For rotorcraft-load combinations lished limiting external load oper- to be used for human external cargo ations to such angles for which compli- applications, the rotorcraft must— ance with this paragraph has been (1) For jettisonable external loads, shown; or have a quick-release system that meets (2) It is shown that the lesser angle the requirements of paragraph (b) of can not be exceeded in service. this section and that— (b) The external load attaching (i) Provides a dual actuation device means, for jettisonable rotorcraft-load for the primary quick release sub- combinations, must include a quick-re- system, and lease system to enable the pilot to re- (ii) Provides a separate dual actu- lease the external load quickly during ation device for the backup quick re- flight. The quick-release system must lease subsystem; consist of a primary quick release sub- (2) Have a reliable, approved per- system and a backup quick release sub- sonnel carrying device system that has system that are isolated from one an- the structural capability and personnel other. The quick release system, and safety features essential for external the means by which it is controlled, occupant safety; must comply with the following: (3) Have placards and markings at all (1) A control for the primary quick appropriate locations that clearly state release subsystem must be installed ei- the essential system operating instruc- ther on one of the pilot’s primary con- tions and, for the personnel carrying trols or in an equivalently accessible device system, ingress and egress in- location and must be designed and lo- structions; cated so that it may be operated by ei- (4) Have equipment to allow direct ther the pilot or a crewmember with- intercommunication among required out hazardously limiting the ability to crewmembers and external occupants; control the rotorcraft during an emer- (5) Have the appropriate limitations gency situation. and procedures incorporated in the (2) A control for the backup quick re- flight manual for conducting human lease subsystem, readily accessible to external cargo operations; and either the pilot or another crew- (6) For human external cargo applica- member, must be provided. tions requiring use of Category A (3) Both the primary and backup rotorcraft, have one-engine-inoperative quick release subsystems must— hover performance data and procedures

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in the flight manual for the weights, Subpart E—Powerplant altitudes, and temperatures for which external load approval is requested. GENERAL (d) The critically configured jettison- able external loads must be shown by a § 29.901 Installation. combination of analysis, ground tests, (a) For the purpose of this part, the and flight tests to be both transport- powerplant installation includes each able and releasable throughout the ap- part of the rotorcraft (other than the proved operational envelope without main and auxiliary rotor structures) hazard to the rotorcraft during normal that— flight conditions. In addition, these ex- (1) Is necessary for propulsion; ternal loads—must be shown to be re- (2) Affects the control of the major leasable without hazard to the rotor- propulsive units; or craft during emergency flight condi- (3) Affects the safety of the major tions. propulsive units between normal in- (e) A placard or marking must be in- spections or overhauls. (b) For each powerplant installa- stalled next to the external-load at- tion— taching means clearly stating any (1) The installation must comply operational limitations and the max- with— imum authorized external load as dem- (i) The installation instructions pro- onstrated under § 29.25 and this section. vided under § 33.5 of this chapter; and (f) The fatigue evaluation of § 29.571 (ii) The applicable provisions of this of this part does not apply to rotor- subpart. craft-load combinations to be used for (2) Each component of the installa- nonhuman external cargo except for tion must be constructed, arranged, the failure of critical structural ele- and installed to ensure its continued ments that would result in a hazard to safe operation between normal inspec- the rotorcraft. For rotorcraft-load tions or overhauls for the range of tem- combinations to be used for human ex- perature and altitude for which ap- ternal cargo, the fatigue evaluation of proval is requested. § 29.571 of this part applies to the entire (3) Accessibility must be provided to quick release and personnel carrying allow any inspection and maintenance device structural systems and their at- necessary for continued airworthiness; tachments. and (4) Electrical interconnections must [Amdt. 29–12, 41 FR 55472, Dec. 20, 1976, as be provided to prevent differences of amended by Amdt. 27–26, 55 FR 8005, Mar. 6, potential between major components of 1990; Amdt. 29–43, 64 FR 43020, Aug. 6, 1999] the installation and the rest of the MISCELLANEOUS rotorcraft. (5) Axial and radial expansion of tur- § 29.871 Leveling marks. bine engines may not affect the safety of the installation. There must be reference marks for (6) Design precautions must be taken leveling the rotorcraft on the ground. to minimize the possibility of incorrect assembly of components and equipment § 29.873 Ballast provisions. essential to safe operation of the rotor- Ballast provisions must be designed craft, except where operation with the and constructed to prevent inadvertent incorrect assembly can be shown to be shifting of ballast in flight. extremely improbable.

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(c) For each powerplant and auxiliary (2) Duplicate means must be avail- power unit installation, it must be es- able for stopping the engine and the tablished that no single failure or mal- controls must be where all are not like- function or probable combination of ly to be damaged at the same time in failures will jeopardize the safe oper- case of fire. ation of the rotorcraft except that the (d) Turbine engine installation. For failure of structural elements need not turbine engine installations— be considered if the probability of any (1) Design precautions must be taken such failure is extremely remote. to minimize the hazards to the rotor- (d) Each auxiliary power unit instal- craft in the event of an engine rotor lation must meet the applicable provi- failure; and sions of this subpart. (2) The powerplant systems associ- ated with engine control devices, sys- (Secs. 313(a), 601, 603, 604, Federal Aviation tems, and instrumentation must be de- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 signed to give reasonable assurance U.S.C. 1655(c))) that those engine operating limitations that adversely affect engine rotor [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as structural integrity will not be exceed- amended by Amdt. 29–3, 33 FR 969, Jan. 26, ed in service. 1968; Amdt, 29–13, 42 FR 15046, Mar. 17, 1977; Amdt. 29–17, 43 FR 50600, Oct. 30, 1978; Amdt. (e) Restart capability. (1) A means to 29–26, 53 FR 34215, Sept. 2, 1988; Amdt. 29–36, restart any engine in flight must be 60 FR 55776, Nov. 2, 1995] provided. (2) Except for the in-flight shutdown § 29.903 Engines. of all engines, engine restart capability (a) Engine type certification. Each en- must be demonstrated throughout a gine must have an approved type cer- flight envelope for the rotorcraft. tificate. Reciprocating engines for use (3) Following the in-flight shutdown in helicopters must be qualified in ac- of all engines, in-flight engine restart cordance with § 33.49(d) of this chapter capability must be provided. or be otherwise approved for the in- (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 tended usage. U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 (b) Category A; engine isolation. For U.S.C. 1655(c)) each category A rotorcraft, the power- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as plants must be arranged and isolated amended by Amdt. 29–12, 41 FR 55472, Dec. 20, from each other to allow operation, in 1976; Amdt. 29–26, 53 FR 34215, Sept. 2, 1988; at least one configuration, so that the Amdt. 29–31, 55 FR 38967, Sept. 21, 1990; 55 FR failure or malfunction of any engine, or 41309, Oct. 10, 1990; Amdt. 29–36, 60 FR 55776, the failure of any system that can af- Nov. 2, 1995] fect any engine, will not— (1) Prevent the continued safe oper- § 29.907 Engine vibration. ation of the remaining engines; or (a) Each engine must be installed to (2) Require immediate action, other prevent the harmful vibration of any than normal pilot action with primary part of the engine or rotorcraft. flight controls, by any crewmember to (b) The addition of the rotor and the maintain safe operation. rotor drive system to the engine may (c) Category A; control of engine rota- not subject the principal rotating parts tion. For each Category A rotorcraft, of the engine to excessive vibration there must be a means for stopping the stresses. This must be shown by a vi- rotation of any engine individually in bration investigation. flight, except that, for turbine engine installations, the means for stopping § 29.908 Cooling fans. the engine need be provided only where For cooling fans that are a part of a necessary for safety. In addition— powerplant installation the following (1) Each component of the engine apply: stopping system that is located on the (a) Category A. For cooling fans in- engine side of the firewall, and that stalled in Category A rotorcraft, it might be exposed to fire, must be at must be shown that a fan blade failure least fire resistant; or will not prevent continued safe flight

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either because of damage caused by the (c) Arrangement. Rotor drive systems failed blade or loss of cooling air. must be arranged as follows: (b) Category B. For cooling fans in- (1) Each rotor drive system of multi- stalled in category B rotorcraft, there engine rotorcraft must be arranged so must be means to protect the rotor- that each rotor necessary for operation craft and allow a safe landing if a fan and control will continue to be driven blade fails. It must be shown that— by the remaining engines if any engine (1) The fan blade would be contained fails. in the case of a failure; (2) For single-engine rotorcraft, each (2) Each fan is located so that a fan rotor drive system must be so arranged blade failure will not jeopardize safety; that each rotor necessary for control in or autorotation will continue to be driven (3) Each fan blade can withstand an by the main rotors after disengage- ultimate load of 1.5 times the cen- ment of the engine from the main and trifugal force expected in service, lim- auxiliary rotors. ited by either— (3) Each rotor drive system must in- (i) The highest rotational speeds corporate a unit for each engine to achievable under uncontrolled condi- automatically disengage that engine tions; or from the main and auxiliary rotors if (ii) An overspeed limiting device. that engine fails. (c) Fatigue evaluation. Unless a fa- (4) If a torque limiting device is used tigue evaluation under § 29.571 is con- in the rotor drive system, it must be ducted, it must be shown that cooling located so as to allow continued con- fan blades are not operating at reso- trol of the rotorcraft when the device nant conditions within the operating is operating. limits of the rotorcraft. (5) If the rotors must be phased for intermeshing, each system must pro- (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 vide constant and positive phase rela- U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 U.S.C. 1655 (c)) tionship under any operating condi- tion. [Amdt. 29–13, 42 FR 15046, Mar. 17, 1977, as (6) If a rotor dephasing device is in- amended by Amdt. 29–26, 53 FR 34215, Sept. 2, corporated, there must be means to 1988] keep the rotors locked in proper phase ROTOR DRIVE SYSTEM before operation. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as § 29.917 Design. amended by Amdt. 29–12, 41 FR 55472, Dec. 20, (a) General. The rotor drive system 1976; Amdt. 29–40, 61 FR 21908, May 10, 1996] includes any part necessary to trans- mit power from the engines to the § 29.921 Rotor brake. rotor hubs. This includes gear boxes, If there is a means to control the ro- shafting, universal joints, couplings, tation of the rotor drive system inde- rotor brake assemblies, clutches, sup- pendently of the engine, any limita- porting bearings for shafting, any at- tions on the use of that means must be tendant accessory pads or drives, and specified, and the control for that any cooling fans that are a part of, at- means must be guarded to prevent in- tached to, or mounted on the rotor advertent operation. drive system. (b) Design assessment. A design assess- § 29.923 Rotor drive system and con- ment must be performed to ensure that trol mechanism tests. the rotor drive system functions safely (a) Endurance tests, general. Each over the full range of conditions for rotor drive system and rotor control which certification is sought. The de- mechanism must be tested, as pre- sign assessment must include a de- scribed in paragraphs (b) through (n) tailed failure analysis to identify all and (p) of this section, for at least 200 failures that will prevent continued hours plus the time required to meet safe flight or safe landing and must the requirements of paragraphs (b)(2), identify the means to minimize the (b)(3), and (k) of this section. These likelihood of their occurrence. tests must be conducted as follows:

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(1) Ten-hour test cycles must be used, (3) For multiengine, turbine-powered except that the test cycle must be ex- rotorcraft for which the use of 30-sec- tended to include the OEI test of para- ond/2-minute OEI power is requested, graphs (b)(2) and (k), of this section if the takeoff run must be conducted as OEI ratings are requested. prescribed in paragraph (b)(1) of this (2) The tests must be conducted on section except for the following: the rotorcraft. (i) Immediately following any one 5- (3) The test torque and rotational minute power-on run required by para- speed must be— graph (b)(1) of this section, simulate a (i) Determined by the powerplant failure for each power source in turn, limitations; and and apply the maximum torque and the (ii) Absorbed by the rotors to be ap- maximum speed for use with 30-second proved for the rotorcraft. OEI power to the remaining affected (b) Endurance tests; takeoff run. The drive system power inputs for not less takeoff run must be conducted as fol- than 30 seconds. Each application of 30- lows: second OEI power must be followed by (1) Except as prescribed in para- two applications of the maximum graphs (b)(2) and (b)(3) of this section, torque and the maximum speed for use the takeoff torque run must consist of with the 2 minute OEI power for not 1 hour of alternate runs of 5 minutes at less than 2 minutes each; the second takeoff torque and the maximum speed application must follow a period at sta- for use with takeoff torque, and 5 min- bilized continuous or 30 minute OEI utes at as low an engine idle speed as power (whichever is requested by the practicable. The engine must be de- applicant). At least one run sequence clutched from the rotor drive system, must be conducted from a simulated and the rotor brake, if furnished and so ‘‘flight idle’’ condition. When con- intended, must be applied during the ducted on a bench test, the test se- first minute of the idle run. During the quence must be conducted following remaining 4 minutes of the idle run, stabilization at take-off power. the clutch must be engaged so that the (ii) For the purpose of this para- engine drives the rotors at the min- graph, an affected power input includes imum practical r.p.m. The engine and all parts of the rotor drive system the rotor drive system must be acceler- which can be adversely affected by the ated at the maximum rate. When de- application of higher or asymmetric clutching the engine, it must be decel- torque and speed prescribed by the erated rapidly enough to allow the op- test. eration of the overrunning clutch. (iii) This test may be conducted on a (2) For helicopters for which the use representative bench test facility when of a 21⁄2-minute OEI rating is requested, engine limitations either preclude re- the takeoff run must be conducted as peated use of this power or would re- prescribed in paragraph (b)(1) of this sult in premature engine removals dur- section, except for the third and sixth ing the test. The loads, the vibration runs for which the takeoff torque and frequency, and the methods of applica- the maximum speed for use with take- tion to the affected rotor drive system off torque are prescribed in that para- components must be representative of graph. For these runs, the following rotorcraft conditions. Test components apply: must be those used to show compliance (i) Each run must consist of at least with the remainder of this section. one period of 21⁄2 minutes with takeoff (c) Endurance tests; maximum contin- torque and the maximum speed for use uous run. Three hours of continuous op- with takeoff torque on all engines. eration at maximum continuous torque (ii) Each run must consist of at least and the maximum speed for use with one period, for each engine in sequence, maximum continuous torque must be during which that engine simulates a conducted as follows: power failure and the remaining en- (1) The main rotor controls must be gines are run at the 21⁄2-minute OEI operated at a minimum of 15 times torque and the maximum speed for use each hour through the main rotor pitch with 21⁄2-minute OEI torque for 21⁄2 min- positions of maximum vertical thrust, utes. maximum forward thrust component,

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maximum aft thrust component, max- must be conducted in accordance with imum left thrust component, and max- the run schedule of paragraph (b)(1) of imum right thrust component, except this section without consideration of that the control movements need not paragraph (b)(2) of this section. produce loads or blade flapping motion (h) Endurance tests; overspeed run. One exceeding the maximum loads of mo- hour of continuous operation must be tions encountered in flight. conducted at maximum continuous (2) The directional controls must be torque and the maximum power-on operated at a minimum of 15 times overspeed expected in service, assum- each hour through the control ex- ing that speed and torque limiting de- tremes of maximum right turning vices, if any, function properly. torque, neutral torque as required by (i) Endurance tests; rotor control posi- the power applied to the main rotor, tions. When the rotor controls are not and maximum left turning torque. being cycled during the tie-down tests, (3) Each maximum control position the rotor must be operated, using the must be held for at least 10 seconds, procedures prescribed in paragraph (c) and the rate of change of control posi- of this section, to produce each of the tion must be at least as rapid as that maximum thrust positions for the fol- for normal operation. lowing percentages of test time (except (d) Endurance tests; 90 percent of max- that the control positions need not imum continuous run. One hour of con- produce loads or blade flapping motion tinuous operation at 90 percent of max- exceeding the maximum loads or mo- imum continuous torque and the max- tions encountered in flight): imum speed for use with 90 percent of (1) For full vertical thrust, 20 per- maximum continuous torque must be cent. conducted. (2) For the forward thrust compo- (e) Endurance tests; 80 percent of max- nent, 50 percent. imum continuous run. One hour of con- (3) For the right thrust component, tinuous operation at 80 percent of max- 10 percent. imum continuous torque and the min- (4) For the left thrust component, 10 imum speed for use with 80 percent of percent. maximum continuous torque must be (5) For the aft thrust component, 10 conducted. percent. (f) Endurance tests; 60 percent of max- (j) Endurance tests, clutch and brake imum continuous run. Two hours or, for engagements. A total of at least 400 helicopters for which the use of either clutch and brake engagements, includ- 30-minute OEI power or continuous OEI ing the engagements of paragraph (b) power is requested, 1 hour of contin- of this section, must be made during uous operation at 60 percent of max- the takeoff torque runs and, if nec- imum continuous torque and the min- essary, at each change of torque and imum speed for use with 60 percent of speed throughout the test. In each maximum continuous torque must be clutch engagement, the shaft on the conducted. driven side of the clutch must be accel- (g) Endurance tests; engine malfunc- erated from rest. The clutch engage- tioning run. It must be determined ments must be accomplished at the whether malfunctioning of compo- speed and by the method prescribed by nents, such as the engine fuel or igni- the applicant. During deceleration tion systems, or whether unequal en- after each clutch engagement, the en- gine power can cause dynamic condi- gines must be stopped rapidly enough tions detrimental to the drive system. to allow the engines to be automati- If so, a suitable number of hours of op- cally disengaged from the rotors and eration must be accomplished under rotor drives. If a rotor brake is in- those conditions, 1 hour of which must stalled for stopping the rotor, the be included in each cycle, and the re- clutch, during brake engagements, maining hours of which must be ac- must be disengaged above 40 percent of complished at the end of the 20 cycles. maximum continuous rotor speed and If no detrimental condition results, an the rotors allowed to decelerate to 40 additional hour of operation in compli- percent of maximum continuous rotor ance with paragraph (b) of this section speed, at which time the rotor brake

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must be applied. If the clutch design (o) Each part tested as prescribed in does not allow stopping the rotors with this section must be in a serviceable the engine running, or if no clutch is condition at the end of the tests. No in- provided, the engine must be stopped tervening disassembly which might af- before each application of the rotor fect test results may be conducted. brake, and then immediately be started (p) Endurance tests; operating lubri- after the rotors stop. cants. To be approved for use in rotor (k) Endurance tests; OEI power run. (1) drive and control systems, lubricants 30-minute OEI power run. For rotorcraft must meet the specifications of lubri- for which the use of 30-minute OEI cants used during the tests prescribed power is requested, a run at 30-minute by this section. Additional or alternate OEI torque and the maximum speed for lubricants may be qualified by equiva- use with 30-minute OEI torque must be lent testing or by comparative analysis conducted as follows: For each engine, of lubricant specifications and rotor in sequence, that engine must be inop- drive and control system characteris- erative and the remaining engines tics. In addition— must be run for a 30-minute period. (1) At least three 10-hour cycles re- (2) Continuous OEI power run. For quired by this section must be con- rotorcraft for which the use of contin- ducted with transmission and gearbox uous OEI power is requested, a run at lubricant temperatures, at the location continuous OEI torque and the max- prescribed for measurement, not lower imum speed for use with continuous than the maximum operating tempera- OEI torque must be conducted as fol- ture for which approval is requested; lows: For each engine, in sequence, (2) For pressure lubricated systems, that engine must be inoperative and at least three 10-hour cycles required the remaining engines must be run for by this section must be conducted with 1 hour. the lubricant pressure, at the location (3) The number of periods prescribed prescribed for measurement, not higher in paragraph (k)(1) or (k)(2) of this sec- than the minimum operating pressure tion may not be less than the number for which approval is requested; and of engines, nor may it be less than two. (3) The test conditions of paragraphs (l) [Reserved] (p)(1) and (p)(2) of this section must be applied simultaneously and must be ex- (m) Any components that are af- tended to include operation at any one- fected by maneuvering and gust loads engine-inoperative rating for which ap- must be investigated for the same proval is requested. flight conditions as are the main ro- tors, and their service lives must be de- (Secs. 313(a), 601, 603, 604, Federal Aviation termined by fatigue tests or by other Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), acceptable methods. In addition, a sec. 6(c), Dept. of Transportation Act (49 level of safety equal to that of the U.S.C. 1655(c))) main rotors must be provided for— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (1) Each component in the rotor drive amended by Amdt. 29–1, 30 FR 8778, July 13, system whose failure would cause an 1965; Amdt. 29–17, 43 FR 50600, Oct. 30, 1978; uncontrolled landing; Amdt. 29–26, 53 FR 34215, Sept. 2, 1988; Amdt. 29–31, 55 FR 38967, Sept. 21, 1990; Amdt. 29–34, (2) Each component essential to the 59 FR 47768, Sept. 16, 1994; Amdt. 29–40, 61 FR phasing of rotors on multirotor rotor- 21908, May 10, 1996; Amdt. 29–42, 63 FR 43285, craft, or that furnishes a driving link Aug. 12, 1998] for the essential control of rotors in autorotation; and § 29.927 Additional tests. (3) Each component common to two (a) Any additional dynamic, endur- or more engines on multiengine rotor- ance, and operational tests, and vibra- craft. tory investigations necessary to deter- (n) Special tests. Each rotor drive sys- mine that the rotor drive mechanism is tem designed to operate at two or more safe, must be performed. gear ratios must be subjected to special (b) If turbine engine torque output to testing for durations necessary to sub- the transmission can exceed the high- stantiate the safety of the rotor drive est engine or transmission torque system. limit, and that output is not directly

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controlled by the pilot under normal shown to be reliable, their rotational operating conditions (such as where speed limits need not be exceeded. the primary engine power control is ac- These runs must be conducted as fol- complished through the flight control), lows: the following test must be made: (1) Overspeed runs must be alternated (1) Under conditions associated with with stabilizing runs of from 1 to 5 all engines operating, make 200 appli- minutes duration each at 60 to 80 per- cations, for 10 seconds each, of torque cent of maximum continuous speed. that is at least equal to the lesser of— (2) Acceleration and deceleration (i) The maximum torque used in must be accomplished in a period not meeting § 29.923 plus 10 percent; or longer than 10 seconds (except where (ii) The maximum torque attainable maximum engine acceleration rate will under probable operating conditions, require more than 10 seconds), and the assuming that torque limiting devices, time for changing speeds may not be if any, function properly. deducted from the specified time for (2) For multiengine rotorcraft under the overspeed runs. conditions associated with each engine, (3) Overspeed runs must be made with in turn, becoming inoperative, apply to the rotors in the flattest pitch for the remaining transmission torque in- smooth operation. puts the maximum torque attainable (e) The tests prescribed in paragraphs under probable operating conditions, (b) and (d) of this section must be con- assuming that torque limiting devices, ducted on the rotorcraft and the torque if any, function properly. Each trans- must be absorbed by the rotors to be mission input must be tested at this installed, except that other ground or maximum torque for at least fifteen flight test facilities with other appro- minutes. priate methods of torque absorption (c) Lubrication system failure. For lu- may be used if the conditions of sup- brication systems required for proper port and vibration closely simulate the operation of rotor drive systems, the conditions that would exist during a following apply: test on the rotorcraft. (1) Category A. Unless such failures (f) Each test prescribed by this sec- are extremely remote, it must be tion must be conducted without inter- shown by test that any failure which vening disassembly and, except for the results in loss of lubricant in any nor- lubrication system failure test re- mal use lubrication system will not quired by paragraph (c) of this section, prevent continued safe operation, al- each part tested must be in a service- though not necessarily without dam- able condition at the conclusion of the age, at a torque and rotational speed test. prescribed by the applicant for contin- ued flight, for at least 30 minutes after (Secs. 313(a), 601, 603, 604, Federal Aviation perception by the flightcrew of the lu- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423 1424), brication system failure or loss of lu- sec. 6(c), Dept. of Transportation Act (49 U.S.C. 1655(c))) bricant. (2) Category B. The requirements of [Amdt. 29–3, 33 FR 969, Jan. 26, 1968, as Category A apply except that the rotor amended by Amdt. 29–17, 43 FR 50601, Oct. 30, drive system need only be capable of 1978; Amdt. 29–26, 53 FR 34216, Sept. 2, 1988] operating under autorotative condi- tions for at least 15 minutes. § 29.931 Shafting critical speed. (d) Overspeed test. The rotor drive sys- (a) The critical speeds of any shafting tem must be subjected to 50 overspeed must be determined by demonstration runs, each 30±3 seconds in duration, at except that analytical methods may be not less than either the higher of the used if reliable methods of analysis are rotational speed to be expected from an available for the particular design. engine control device failure or 105 per- (b) If any critical speed lies within, cent of the maximum rotational speed, or close to, the operating ranges for including transients, to be expected in idling, power-on, and autorotative con- service. If speed and torque limiting ditions, the stresses occurring at that devices are installed, are independent speed must be within safe limits. This of the normal engine control, and are must be shown by tests.

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(c) If analytical methods are used and (2) There are means to prevent intro- show that no critical speed lies within ducing air into the system. the permissible operating ranges, the (c) Each fuel system for a turbine en- margins between the calculated crit- gine must be capable of sustained oper- ical speeds and the limits of the allow- ation throughout its flow and pressure able operating ranges must be adequate range with fuel initially saturated with to allow for possible variations be- water at 80 degrees F. and having 0.75cc tween the computed and actual values. of free water per gallon added and cooled to the most critical condition [Amdt. 29–12, 41 FR 55472, Dec. 20, 1976] for icing likely to be encountered in § 29.935 Shafting joints. operation. Each universal joint, slip joint, and [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as other shafting joints whose lubrication amended by Amdt. 29–10, 39 FR 35462, Oct. 1, is necessary for operation must have 1974; Amdt. 29–12, 41 FR 55473, Dec. 20, 1976] provision for lubrication. § 29.952 Fuel system crash resistance. § 29.939 Turbine engine operating Unless other means acceptable to the characteristics. Administrator are employed to mini- (a) Turbine engine operating charac- mize the hazard of fuel fires to occu- teristics must be investigated in flight pants following an otherwise surviv- to determine that no adverse charac- able impact (crash landing), the fuel teristics (such as stall, surge, of flame- systems must incorporate the design out) are present, to a hazardous degree, features of this section. These systems during normal and emergency oper- must be shown to be capable of sus- ation within the range of operating taining the static and dynamic decel- limitations of the rotorcraft and of the eration loads of this section, consid- engine. ered as ultimate loads acting alone, (b) The turbine engine air inlet sys- measured at the system component’s tem may not, as a result of airflow dis- center of gravity without structural tortion during normal operation, cause damage to the system components, fuel vibration harmful to the engine. tanks, or their attachments that would (c) For governor-controlled engines, leak fuel to an ignition source. it must be shown that there exists no (a) Drop test requirements. Each tank, hazardous torsional instability of the or the most critical tank, must be drive system associated with critical drop-tested as follows: combinations of power, rotational (1) The drop height must be at least speed, and control displacement. 50 feet. (2) The drop impact surface must be [Amdt. 29–2, 32 FR 6914, May 5, 1967, as amended by Amdt. 29–12, 41 FR 55473, Dec. 20, nondeforming. 1976] (3) The tanks must be filled with water to 80 percent of the normal, full FUEL SYSTEM capacity. (4) The tank must be enclosed in a § 29.951 General. surrounding structure representative (a) Each fuel system must be con- of the installation unless it can be es- structed and arranged to ensure a flow tablished that the surrounding struc- of fuel at a rate and pressure estab- ture is free of projections or other de- lished for proper engine and auxiliary sign features likely to contribute to power unit functioning under any like- upture of the tank. ly operating conditions, including the (5) The tank must drop freely and im- maneuvers for which certification is pact in a horizontal position ±10°. requested and during which the engine (6) After the drop test, there must be or auxiliary power unit is permitted to no leakage. be in operation. (b) Fuel tank load factors. Except for (b) Each fuel system must be ar- fuel tanks located so that tank rupture ranged so that— with fuel release to either significant (1) No engine or fuel pump can draw ignition sources, such as engines, heat- fuel from more than one tank at a ers, and auxiliary power units, or occu- time; or pants is extremely remote, each fuel

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tank must be designed and installed to (iv) All breakaway couplings must in- retain its contents under the following corporate design provisions to prevent ultimate inertial load factors, acting uncoupling or unintended closing due alone. to operational shocks, vibrations, or (1) For fuel tanks in the cabin: accelerations. (i) Upward—4g. (v) No breakaway coupling design (ii) Forward—16g. may allow the release of fuel once the (iii) Sideward—8g. coupling has performed its intended (iv) Downward—20g. function. (2) For fuel tanks located above or (2) All individual breakaway cou- behind the crew or passenger compart- plings, coupling fuel feed systems, or ment that, if loosened, could injure an equivalent means must be designed, occupant in an emergency landing: tested, installed, and maintained so in- (i) Upward—1.5g. advertent fuel shutoff in flight is im- (ii) Forward—8g. probable in accordance with § 29.955(a) (iii) Sideward—2g. and must comply with the fatigue eval- (iv) Downward—4g. uation requirements of § 29.571 without (3) For fuel tanks in other areas: leaking. (i) Upward—1.5g. (3) Alternate, equivalent means to (ii) Forward—4g. the use of breakaway couplings must (iii) Sideward—2g. not create a survivable impact-induced (iv) Downward—4g. load on the fuel line to which it is in- (c) Fuel line self-sealing breakaway stalled greater than 25 to 50 percent of couplings. Self-sealing breakaway cou- the ultimate load (strength) of the plings must be installed unless haz- weakest component in the line and ardous relative motion of fuel system must comply with the fatigue require- components to each other or to local ments of § 29.571 without leaking. rotorcraft structure is demonstrated to (d) Frangible or deformable structural be extremely improbable or unless attachments. Unless hazardous relative other means are provided. The cou- motion of fuel tanks and fuel system plings or equivalent devices must be components to local rotorcraft struc- installed at all fuel tank-to-fuel line ture is demonstrated to be extremely connections, tank-to-tank intercon- improbable in an otherwise survivable nects, and at other points in the fuel impact, frangible or locally deformable system where local structural deforma- attachments of fuel tanks and fuel sys- tion could lead to the release of fuel. tem components to local rotorcraft (1) The design and construction of structure must be used. The attach- self-sealing breakaway couplings must ment of fuel tanks and fuel system incorporate the following design fea- components to local rotorcraft struc- tures: ture, whether frangible or locally de- (i) The load necessary to separate a formable, must be designed such that breakaway coupling must be between its separation or relative local defor- 25 to 50 percent of the minimum ulti- mation will occur without rupture or mate failure load (ultimate strength) local tear-out of the fuel tank or fuel of the weakest component in the fluid- system component that will cause fuel carrying line. The separation load leakage. The ultimate strength of fran- must in no case be less than 300 pounds, gible or deformable attachments must regardless of the size of the fluid line. be as follows: (ii) A breakaway coupling must sepa- (1) The load required to separate a rate whenever its ultimate load (as de- frangible attachment from its support fined in paragraph (c)(1)(i) of this sec- structure, or deform a locally deform- tion) is applied in the failure modes able attachment relative to its support most likely to occur. structure, must be between 25 and 50 (iii) All breakaway couplings must percent of the minimum ultimate load incorporate design provisions to vis- (ultimate strength) of the weakest ually ascertain that the coupling is component in the attached system. In locked together (leak-free) and is open no case may the load be less than 300 during normal installation and service. pounds.

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(2) A frangible or locally deformable § 29.955 Fuel flow. attachment must separate or locally (a) General. The fuel system for each deform as intended whenever its ulti- engine must provide the engine with at mate load (as defined in paragraph least 100 percent of the fuel required (d)(1) of this section) is applied in the under all operating and maneuvering modes most likely to occur. conditions to be approved for the rotor- (3) All frangible or locally deformable craft, including, as applicable, the fuel attachments must comply with the fa- required to operate the engines under tigue requirements of § 29.571. the test conditions required by § 29.927. (e) Separation of fuel and ignition Unless equivalent methods are used, sources. To provide maximum crash re- compliance must be shown by test dur- sistance, fuel must be located as far as ing which the following provisions are practicable from all occupiable areas met, except that combinations of con- and from all potential ignition sources. ditions which are shown to be improb- (f) Other basic mechanical design cri- able need not be considered. teria. Fuel tanks, fuel lines, electrical (1) The fuel pressure, corrected for wires, and electrical devices must be accelerations (load factors), must be designed, constructed, and installed, as within the limits specified by the en- far as practicable, to be crash resist- gine type certificate data sheet. ant. (2) The fuel level in the tank may not exceed that established as the unusable (g) Rigid or semirigid fuel tanks. Rigid fuel supply for that tank under § 29.959, or semirigid fuel tank or bladder walls plus that necessary to conduct the must be impact and tear resistant. test. [Doc. No. 26352, 59 FR 50387, Oct. 3, 1994] (3) The fuel head between the tank and the engine must be critical with § 29.953 Fuel system independence. respect to rotorcraft flight attitudes. (a) For category A rotorcraft— (4) The fuel flow transmitter, if in- stalled, and the critical fuel pump (for (1) The fuel system must meet the re- pump-fed systems) must be installed to quirements of § 29.903(b); and produce (by actual or simulated fail- (2) Unless other provisions are made ure) the critical restriction to fuel flow to meet paragraph (a)(1) of this section, to be expected from component failure. the fuel system must allow fuel to be (5) Critical values of engine rota- supplied to each engine through a sys- tional speed, electrical power, or other tem independent of those parts of each sources of fuel pump motive power system supplying fuel to other engines. must be applied. (b) Each fuel system for a multien- (6) Critical values of fuel properties gine category B rotorcraft must meet which adversely affect fuel flow are ap- the requirements of paragraph (a)(2) of plied during demonstrations of fuel this section. However, separate fuel flow capability. tanks need not be provided for each en- (7) The fuel filter required by § 29.997 gine. is blocked to the degree necessary to simulate the accumulation of fuel con- § 29.954 Fuel system lightning protec- tamination required to activate the in- tion. dicator required by § 29.1305(a)(17). The fuel system must be designed (b) Fuel transfer system. If normal op- and arranged to prevent the ignition of eration of the fuel system requires fuel fuel vapor within the system by— to be transferred to another tank, the (a) Direct lightning strikes to areas transfer must occur automatically via having a high probability of stroke at- a system which has been shown to tachment; maintain the fuel level in the receiving (b) Swept lightning strokes to areas tank within acceptable limits during where swept strokes are highly prob- flight or surface operation of the rotor- able; and craft. (c) Multiple fuel tanks. If an engine (c) Corona and streamering at fuel can be supplied with fuel from more vent outlets. than one tank, the fuel system, in addi- [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] tion to having appropriate manual

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switching capability, must be designed tion, inertia, fluid, and structural loads to prevent interruption of fuel flow to to which it may be subjected in oper- that engine, without attention by the ation. flightcrew, when any tank supplying (b) Each flexible fuel tank bladder or fuel to that engine is depleted of usable liner must be approved or shown to be fuel during normal operation and any suitable for the particular application other tank that normally supplies fuel and must be puncture resistant. Punc- to that engine alone contains usable ture resistance must be shown by fuel. meeting the TSO–C80, paragraph 16.0, requirements using a minimum punc- [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] ture force of 370 pounds. § 29.957 Flow between interconnected (c) Each integral fuel tank must have tanks. facilities for inspection and repair of its interior. (a) Where tank outlets are inter- (d) The maximum exposed surface connected and allow fuel to flow be- temperature of all components in the tween them due to gravity or flight ac- fuel tank must be less by a safe margin celerations, it must be impossible for than the lowest expected autoignition fuel to flow between tanks in quan- temperature of the fuel or fuel vapor in tities great enough to cause overflow the tank. Compliance with this re- from the tank vent in any sustained quirement must be shown under all op- flight condition. erating conditions and under all nor- (b) If fuel can be pumped from one mal or malfunction conditions of all tank to another in flight— components inside the tank. (1) The design of the vents and the (e) Each fuel tank installed in per- fuel transfer system must prevent sonnel compartments must be isolated structural damage to tanks from over- by fume-proof and fuel-proof enclosures filling; and that are drained and vented to the ex- (2) There must be means to warn the terior of the rotorcraft. The design and crew before overflow through the vents construction of the enclosures must occurs. provide necessary protection for the § 29.959 Unusable fuel supply. tank, must be crash resistant during a survivable impact in accordance with The unusable fuel supply for each § 29.952, and must be adequate to with- tank must be established as not less stand loads and abrasions to be ex- than the quantity at which the first pected in personnel compartments. evidence of malfunction occurs under the most adverse fuel feed condition [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as occurring under any intended oper- amended by Amdt. 29–26, 53 FR 34217, Sept. 2, 1988; Amdt. 29–35, 59 FR 50388, Oct. 3, 1994] ations and flight maneuvers involving that tank. § 29.965 Fuel tank tests. § 29.961 Fuel system hot weather oper- (a) Each fuel tank must be able to ation. withstand the applicable pressure tests Each suction lift fuel system and in this section without failure or leak- other fuel systems conducive to vapor age. If practicable, test pressures may formation must be shown to operate be applied in a manner simulating the satisfactorily (within certification lim- pressure distribution in service. (b) Each conventional metal tank, its) when using fuel at the most crit- each nonmetallic tank with walls that ical temperature for vapor formation are not supported by the rotorcraft under critical operating conditions in- structure, and each integral tank must cluding, if applicable, the engine oper- be subjected to a pressure of 3.5 p.s.i. ating conditions defined by § 29.927(b)(1) unless the pressure developed during and (b)(2). maximum limit acceleration or emer- [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] gency deceleration with a full tank ex- ceeds this value, in which case a hydro- § 29.963 Fuel tanks: general. static head, or equivalent test, must be (a) Each fuel tank must be able to applied to duplicate the acceleration withstand, without failure, the vibra- loads as far as possible. However, the

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pressure need not exceed 3.5 p.s.i. on in the normal operating range of en- surfaces not exposed to the accelera- gine or rotor system speeds is critical, tion loading. the most critical of these frequencies (c) Each nonmetallic tank with walls must be the test frequency. supported by the rotorcraft structure (4) Under paragraph (d)(3)(ii) and (iii), must be subjected to the following the time of test must be adjusted to ac- tests: complish the same number of vibration (1) A pressure test of at least 2.0 p.s.i. cycles as would be accomplished in 25 This test may be conducted on the hours at the frequency specified in tank alone in conjunction with the test paragraph (d)(3)(i) of this section. specified in paragraph (c)(2) of this sec- (5) During the test, the tank assem- tion. bly must be rocked at the rate of 16 to (2) A pressure test, with the tank 20 complete cycles per minute through mounted in the rotorcraft structure, an angle of 15 degrees on both sides of equal to the load developed by the re- the horizontal (30 degrees total), about action of the contents, with the tank the most critical axis, for 25 hours. If full, during maximum limit accelera- motion about more than one axis is tion or emergency deceleration. How- likely to be critical, the tank must be ever, the pressure need not exceed 2.0 rocked about each critical axis for 121⁄2 p.s.i. on surfaces faces not exposed to hours. the acceleration loading. (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 (d) Each tank with large unsupported U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 or unstiffened flat areas, or with other U.S.C. 1655 (c)) features whose failure or deformation [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as could cause leakage, must be subjected amended by Amdt. 29–13, 42 FR 15046, Mar. 17, to the following test or its equivalent: 1977] (1) Each complete tank assembly and its supports must be vibration tested § 29.967 Fuel tank installation. while mounted to simulate the actual (a) Each fuel tank must be supported installation. so that tank loads are not con- (2) The tank assembly must be vi- centrated on unsupported tank sur- brated for 25 hours while two-thirds faces. In addition— full of any suitable fluid. The ampli- (1) There must be pads, if necessary, tude of vibration may not be less than to prevent chafing between each tank one thirty-second of an inch, unless and its supports; otherwise substantiated. (2) The padding must be non- (3) The test frequency of vibration absorbent or treated to prevent the ab- must be as follows: sorption of fuel; (i) If no frequency of vibration result- (3) If flexible tank liners are used, ing from any r.p.m. within the normal they must be supported so that they operating range of engine or rotor sys- are not required to withstand fluid tem speeds is critical, the test fre- loads; and quency of vibration, in number of cy- (4) Each interior surface of tank com- cles per minute, must, unless a fre- partments must be smooth and free of quency based on a more rational analy- projections that could cause wear of sis is used, be the number obtained by the liner, unless— averaging the maximum and minimum (i) There are means for protection of power-on engine speeds (r.p.m.) for re- the liner at those points; or ciprocating engine powered rotorcraft (ii) The construction of the liner or 2,000 c.p.m. for turbine engine pow- itself provides such protection. ered rotorcraft. (b) Any spaces adjacent to tank sur- (ii) If only one frequency of vibration faces must be adequately ventilated to resulting from any r.p.m. within the avoid accumulation of fuel or fumes in normal operating range of engine or those spaces due to minor leakage. If rotor system speeds is critical, that the tank is in a sealed compartment, frequency of vibration must be the test ventilation may be limited to drain frequency. holes that prevent clogging and that (iii) If more than one frequency of vi- prevent excessive pressure resulting bration resulting from any r.p.m. with- from altitude changes. If flexible tank

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liners are installed, the venting ar- ations and must be crash resistant dur- rangement for the spaces between the ing a survivable impact in accordance liner and its container must maintain with § 29.952(c). In addition— the proper relationship to tank vent (1) Each filler must be marked as pre- pressures for any expected flight condi- scribed in § 29.1557(c)(1); tion. (2) Each recessed filler connection (c) The location of each tank must that can retain any appreciable quan- meet the requirements of § 29.1185(b) tity of fuel must have a drain that dis- and (c). charges clear of the entire rotorcraft; (d) No rotorcraft skin immediately and adjacent to a major air outlet from the (3) Each filler cap must provide a engine compartment may act as the fuel-tight seal under the fluid pressure wall of an integral tank. expected in normal operation and in a [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as survivable impact. amended by Amdt. 29–26, 53 FR 34217, Sept. 2, (b) Each filler cap or filler cap cover 1988; Amdt. 29–35, 59 FR 50388, Oct. 3, 1994] must warn when the cap is not fully locked or seated on the filler connec- § 29.969 Fuel tank expansion space. tion. Each fuel tank or each group of fuel [Doc. No. 26352, 59 FR 50388, Oct. 3, 1994] tanks with interconnected vent sys- tems must have an expansion space of § 29.975 Fuel tank vents and carbu- not less than 2 percent of the combined retor vapor vents. tank capacity. It must be impossible to (a) Fuel tank vents. Each fuel tank fill the fuel tank expansion space inad- must be vented from the top part of the vertently with the rotorcraft in the expansion space so that venting is ef- normal ground attitude. fective under normal flight conditions. [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] In addition— (1) The vents must be arranged to § 29.971 Fuel tank sump. avoid stoppage by dirt or ice forma- (a) Each fuel tank must have a sump tion; with a capacity of not less than the (2) The vent arrangement must pre- greater of— vent siphoning of fuel during normal (1) 0.10 per cent of the tank capacity; operation; or (3) The venting capacity and vent (2) 1⁄16 gallon. pressure levels must maintain accept- (b) The capacity prescribed in para- able differences of pressure between graph (a) of this section must be effec- the interior and exterior of the tank, tive with the rotorcraft in any normal during— attitude, and must be located so that (i) Normal flight operation; the sump contents cannot escape (ii) Maximum rate of ascent and de- through the tank outlet opening. scent; and (c) Each fuel tank must allow drain- (iii) Refueling and defueling (where age of hazardous quantities of water applicable); from each part of the tank to the sump (4) Airspaces of tanks with inter- with the rotorcraft in any ground atti- connected outlets must be inter- tude to be expected in service. connected; (d) Each fuel tank sump must have a (5) There may be no point in any vent drain that allows complete drainage of line where moisture can accumulate the sump on the ground. with the rotorcraft in the ground atti- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as tude or the level flight attitude, unless amended by Amdt. 29–12, 41 FR 55473, Dec. 20, drainage is provided; 1976; Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] (6) No vent or drainage provision may end at any point— § 29.973 Fuel tank filler connection. (i) Where the discharge of fuel from (a) Each fuel tank filler connection the vent outlet would constitute a fire must prevent the entrance of fuel into hazard; or any part of the rotorcraft other than (ii) From which fumes could enter the tank itself during normal oper- personnel compartments; and

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(7) The venting system must be de- (c) The rotorcraft pressure fueling signed to minimize spillage of fuel system (not fuel tanks and fuel tank through the vents to an ignition source vents) must withstand an ultimate in the event of a rollover during land- load that is 2.0 times the load arising ing, ground operations, or a survivable from the maximum pressure, including impact. surge, that is likely to occur during (b) Carburetor vapor vents. Each car- fueling. The maximum surge pressure buretor with vapor elimination connec- must be established with any combina- tions must have a vent line to lead va- tion of tank valves being either inten- pors back to one of the fuel tanks. In tionally or inadvertently closed. addition— (d) The rotorcraft defueling system (1) Each vent system must have (not including fuel tanks and fuel tank means to avoid stoppage by ice; and vents) must withstand an ultimate (2) If there is more than one fuel load that is 2.0 times the load arising tank, and it is necessary to use the from the maximum permissible tanks in a definite sequence, each defueling pressure (positive or nega- vapor vent return line must lead back tive) at the rotorcraft fueling connec- to the fuel tank used for takeoff and tion. landing. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55473, Dec. 20, amended by Amdt. 29–26, 53 FR 34217, Sept. 2, 1976] 1988; Amdt. 29–35, 59 FR 50388, Oct. 3, 1994; Amdt. 29–42, 63 FR 43285, Aug. 12, 1998] FUEL SYSTEM COMPONENTS

§ 29.977 Fuel tank outlet. § 29.991 Fuel pumps. (a) There must be a fuel strainer for (a) Compliance with § 29.955 must not the fuel tank outlet or for the booster be jeopardized by failure of— pump. This strainer must— (1) Any one pump except pumps that (1) For reciprocating engine powered are approved and installed as parts of a airplanes, have 8 to 16 meshes per inch; type certificated engine; or and (2) Any component required for pump (2) For turbine engine powered air- operation except the engine served by planes, prevent the passage of any ob- that pump. ject that could restrict fuel flow or (b) The following fuel pump installa- damage any fuel system component. tion requirements apply: (b) The clear area of each fuel tank (1) When necessary to maintain the outlet strainer must be at least five proper fuel pressure— times the area of the outlet line. (i) A connection must be provided to (c) The diameter of each strainer transmit the carburetor air intake must be at least that of the fuel tank static pressure to the proper fuel pump outlet. relief valve connection; and (d) Each finger strainer must be ac- (ii) The gauge balance lines must be cessible for inspection and cleaning. independently connected to the carbu- retor inlet pressure to avoid incorrect [Amdt. 29–12, 41 FR 55473, Dec. 20, 1976] fuel pressure readings. § 29.979 Pressure refueling and fueling (2) The installation of fuel pumps provisions below fuel level. having seals or diaphragms that may leak must have means for draining (a) Each fueling connection below the leaking fuel. fuel level in each tank must have (3) Each drain line must discharge means to prevent the escape of haz- where it will not create a fire hazard. ardous quantities of fuel from that tank in case of malfunction of the fuel [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] entry valve. (b) For systems intended for pressure § 29.993 Fuel system lines and fittings. refueling, a means in addition to the (a) Each fuel line must be installed normal means for limiting the tank and supported to prevent excessive vi- content must be installed to prevent bration and to withstand loads due to damage to the tank in case of failure of fuel pressure, valve actuation, and ac- the normal means. celerated flight conditions.

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(b) Each fuel line connected to com- rotorcraft or engine fuel system com- ponents of the rotorcraft between ponents required for proper rotorcraft which relative motion could exist must or engine fuel system operation. have provisions for flexibility. [Amdt. No. 29–10, 39 FR 35462, Oct. 1, 1974, as (c) Each flexible connection in fuel amended by Amdt. 29–22, 49 FR 6850, Feb. 23, lines that may be under pressure or 1984; Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] subjected to axial loading must use flexible hose assemblies. § 29.999 Fuel system drains. (d) Flexible hose must be approved. (a) There must be at least one acces- (e) No flexible hose that might be ad- sible drain at the lowest point in each versely affected by high temperatures fuel system to completely drain the may be used where excessive tempera- system with the rotorcraft in any tures will exist during operation or ground attitude to be expected in serv- after engine shutdown. ice. (b) Each drain required by paragraph § 29.995 Fuel valves. (a) of this section including the drains In addition to meeting the require- prescribed in § 29.971 must— ments of § 29.1189, each fuel valve (1) Discharge clear of all parts of the must— rotorcraft; (a) [Reserved] (2) Have manual or automatic means (b) Be supported so that no loads re- to ensure positive closure in the off po- sulting from their operation or from sition; and accelerated flight conditions are trans- (3) Have a drain valve— mitted to the lines attached to the (i) That is readily accessible and valve. which can be easily opened and closed; (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 and U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 (ii) That is either located or pro- U.S.C. 1655 (c)) tected to prevent fuel spillage in the [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as event of a landing with landing gear re- amended by Amdt. 29–13, 42 FR 15046, Mar. 17, tracted. 1977] [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55473, Dec. 20, § 29.997 Fuel strainer or filter. 1976; Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] There must be a fuel strainer or filter between the fuel tank outlet and the § 29.1001 Fuel jettisoning. inlet of the first fuel system compo- If a fuel jettisoning system is in- nent which is susceptible to fuel con- stalled, the following apply: tamination, including but not limited (a) Fuel jettisoning must be safe dur- to the fuel metering device or an en- ing all flight regimes for which jetti- gine positive displacement pump, soning is to be authorized. whichever is nearer the fuel tank out- (b) In showing compliance with para- let. This fuel strainer or filter must— graph (a) of this section, it must be (a) Be accessible for draining and shown that— cleaning and must incorporate a screen (1) The fuel jettisoning system and or element which is easily removable; its operation are free from fire hazard; (b) Have a sediment trap and drain, (2) No hazard results from fuel or fuel except that it need not have a drain if vapors which impinge on any part of the strainer or filter is easily remov- the rotorcraft during fuel jettisoning; able for drain purposes; and (c) Be mounted so that its weight is (3) Controllability of the rotorcraft not supported by the connecting lines remains satisfactory throughout the or by the inlet or outlet connections of fuel jettisoning operation. the strainer or filter inself, unless ade- (c) Means must be provided to auto- quate strengh margins under all load- matically prevent jettisoning fuel ing conditions are provided in the lines below the level required for an all-en- and connections; and gine climb at maximum continuous (d) Provide a means to remove from power from sea level to 5,000 feet alti- the fuel any contaminant which would tude and cruise thereafter for 30 min- jeopardize the flow of fuel through utes at maximum range engine power.

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(d) The controls for any fuel jetti- (b) Expansion space. Oil tank expan- soning system must be designed to sion space must be provided so that— allow flight personnel (minimum crew) (1) Each oil tank used with a recipro- to safely interrupt fuel jettisoning dur- cating engine has an expansion space of ing any part of the jettisoning oper- not less than the greater of 10 percent ation. of the tank capacity or 0.5 gallon, and (e) The fuel jettisoning system must each oil tank used with a turbine en- be designed to comply with the power- gine has an expansion space of not less plant installation requirements of than 10 percent of the tank capacity; § 29.901(c). (2) Each reserve oil tank not directly (f) An auxiliary fuel jettisoning sys- connected to any engine has an expan- tem which meets the requirements of sion space of not less than two percent paragraphs (a), (b), (d), and (e) of this of the tank capacity; and section may be installed to jettison ad- (3) It is impossible to fill the expan- ditional fuel provided it has separate sion space inadvertently with the and independent controls. rotorcraft in the normal ground atti- [Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] tude. (c) Filler connections. Each recessed OIL SYSTEM oil tank filler connection that can re- tain any appreciable quantity of oil § 29.1011 Engines: general. must have a drain that discharges clear (a) Each engine must have an inde- of the entire rotorcraft. In addition— pendent oil system that can supply it (1) Each oil tank filler cap must pro- with an appropriate quantity of oil at a vide an oil-tight seal under the pres- temperature not above that safe for sure expected in operation; continuous operation. (2) For category A rotorcraft, each (b) The usable oil capacity of each oil tank filler cap or filler cap cover system may not be less than the prod- must incorporate features that provide uct of the endurance of the rotorcraft a warning when caps are not fully under critical operating conditions and locked or seated on the filler connec- the maximum allowable oil consump- tion; and tion of the engine under the same con- (3) Each oil filler must be marked ditions, plus a suitable margin to en- under § 29.1557(c)(2). sure adequate circulation and cooling. (d) Vent. Oil tanks must be vented as Instead of a rational analysis of endur- follows: ance and consumption, a usable oil ca- (1) Each oil tank must be vented pacity of one gallon for each 40 gallons from the top part of the expansion of usable fuel may be used for recipro- space to that venting is effective under cating engine installations. all normal flight conditions. (c) Oil-fuel ratios lower than those (2) Oil tank vents must be arranged prescribed in paragraph (c) of this sec- so that condensed water vapor that tion may be used if they are substan- might freeze and obstruct the line can- tiated by data on the oil consumption not accumulate at any point; of the engine. (e) Outlet. There must be means to (d) The ability of the engine and oil prevent entrance into the tank itself, cooling provisions to maintain the oil or into the tank outlet, of any object temperature at or below the maximum that might obstruct the flow of oil established value must be shown under through the system. No oil tank outlet the applicable requirements of §§ 29.1041 may be enclosed by a screen or guard through 29.1049. that would reduce the flow of oil below [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as a safe value at any operating tempera- amended by Amdt. 29–26, 53 FR 34218, Sept. 2, ture. There must be a shutoff valve at 1988] the outlet of each oil tank used with a turbine engine unless the external por- § 29.1013 Oil tanks. tion of the oil system (including oil (a) Installation. Each oil tank instal- tank supports) is fireproof. lation must meet the requirements of (f) Flexible liners. Each flexible oil § 29.967. tank liner must be approved or shown

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to be suitable for the particular instal- the oil is contaminated to a degree lation. (with respect to particle size and den- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as sity) that is greater than that estab- amended by Amdt. 29–10, 39 FR 35462, Oct. 1, lished for the engine under Part 33 of 1974] this chapter. (3) The oil strainer or filter, unless it § 29.1015 Oil tank tests. is installed at an oil tank outlet, must Each oil tank must be designed and incorporate a means to indicate con- installed so that— tamination before it reaches the capac- (a) It can withstand, without failure, ity established in accordance with any vibration, inertia, and fluid loads paragraph (a)(2) of this section. to which it may be subjected in oper- (4) The bypass of a strainer or filter ation; and must be constructed and installed so (b) It meets the requirements of that the release of collected contami- § 29.965, except that instead of the pres- sure specified in § 29.965(b)— nants is minimized by appropriate lo- (1) For pressurized tanks used with a cation of the bypass to ensure that col- turbine engine, the test pressure may lected contaminants are not in the by- not be less than 5 p.s.i. plus the max- pass flow path. imum operating pressure of the tank; (5) An oil strainer or filter that has and no bypass, except one that is installed (2) For all other tanks, the test pres- at an oil tank outlet, must have a sure may not be less than 5 p.s.i. means to connect it to the warning [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as system required in § 29.1305(a)(18). amended by Amdt. 29–10, 39 FR 35462, Oct. 1, (b) Each oil strainer or filter in a 1974] powerplant installation using recipro- cating engines must be constructed and § 29.1017 Oil lines and fittings. installed so that oil will flow at the (a) Each oil line must meet the re- normal rate through the rest of the quirements of § 29.993. system with the strainer or filter ele- (b) Breather lines must be arranged ment completely blocked. so that— (1) Condensed water vapor that might [Amdt. 29–10, 39 FR 35463, Oct. 1, 1974, as freeze and obstruct the line cannot ac- amended by Amdt. 29–22, 49 FR 6850, Feb. 23, cumulate at any point; 1984; Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] (2) The breather discharge will not constitute a fire hazard if foaming oc- § 29.1021 Oil system drains. curs, or cause emitted oil to strike the A drain (or drains) must be provided pilot’s windshield; and to allow safe drainage of the oil sys- (3) The breather does not discharge tem. Each drain must— into the engine air induction system. (a) Be accessible; and (b) Have manual or automatic means § 29.1019 Oil strainer or filter. for positive locking in the closed posi- (a) Each turbine engine installation tion. must incorporate an oil strainer or fil- ter through which all of the engine oil [Amdt. 29–22, 49 FR 6850, Feb. 23, 1984] flows and which meets the following re- quirements: § 29.1023 Oil radiators. (1) Each oil strainer or filter that has (a) Each oil radiator must be able to a bypass must be constructed and in- withstand any vibration, inertia, and stalled so that oil will flow at the nor- oil pressure loads to which it would be mal rate through the rest of the sys- subjected in operation. tem with the strainer or filter com- (b) Each oil radiator air duct must be pletely blocked. (2) The oil strainer or filter must located, or equipped, so that, in case of have the capacity (with respect to op- fire, and with the airflow as it would be erating limitations established for the with and without the engine operating, engine) to ensure that engine oil sys- flames cannot directly strike the radi- tem functioning is not impaired when ator.

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§ 29.1025 Oil valves. flow of lubricant from the outlet to the (a) Each oil shutoff must meet the re- filter required by paragraph (b)(1) of quirements of § 29.1189. this section. The requirements of para- (b) The closing of oil shutoffs may graph (b)(1) of this section do not apply not prevent autorotation. to screens installed at lubricant tank (c) Each oil valve must have positive or sump outlets. stops or suitable index provisions in (c) Splash type lubrication systems the ‘‘on’’ and ‘‘off’’ positions and must for rotor drive system gearboxes must be supported so that no loads resulting comply with §§ 29.1021 and 29.1337(d). from its operation or from accelerated [Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] flight conditions are transmitted to the lines attached to the valve. COOLING

§ 29.1027 Transmission and gearboxes: § 29.1041 General. general. (a) The powerplant and auxiliary (a) The oil system for components of power unit cooling provisions must be the rotor drive system that require able to maintain the temperatures of continuous lubrication must be suffi- powerplant components, engine fluids, ciently independent of the lubrication and auxiliary power unit components systems of the engine(s) to ensure— and fluids within the temperature lim- (1) Operation with any engine inoper- its established for these components ative; and and fluids, under ground, water, and (2) Safe autorotation. flight operating conditions for which (b) Pressure lubrication systems for certification is requested, and after transmissions and gearboxes must normal engine or auxiliary power unit comply with the requirements of shutdown, or both. §§ 29.1013, paragraphs (c), (d), and (f) (b) There must be cooling provisions only, 29.1015, 29.1017, 29.1021, 29.1023, and to maintain the fluid temperatures in 29.1337(d). In addition, the system must any power transmission within safe have— values under any critical surface (1) An oil strainer or filter through (ground or water) and flight operating which all the lubricant flows, and conditions. must— (c) Except for ground-use-only auxil- (i) Be designed to remove from the iary power units, compliance with lubricant any contaminant which may paragraphs (a) and (b) of this section damage transmission and drive system must be shown by flight tests in which components or impede the flow of lu- the temperatures of selected power- bricant to a hazardous degree; and plant component and auxiliary power (ii) Be equipped with a bypass con- unit component, engine, and trans- structed and installed so that— mission fluids are obtained under the (A) The lubricant will flow at the conditions prescribed in those para- normal rate through the rest of the graphs. system with the strainer or filter com- pletely blocked; and [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (B) The release of collected contami- amended by Amdt. 29–26, 53 FR 34218, Sept. 2, nants is minimized by appropriate lo- 1988] cation of the bypass to ensure that col- lected contaminants are not in the by- § 29.1043 Cooling tests. pass flowpath; (a) General. For the tests prescribed (iii) Be equipped with a means to in- in § 29.1041(c), the following apply: dicate collection of contaminants on (1) If the tests are conducted under the filter or strainer at or before open- conditions deviating from the max- ing of the bypass; imum ambient atmospheric tempera- (2) For each lubricant tank or sump ture specified in paragraph (b) of this outlet supplying lubrication to rotor section, the recorded powerplant tem- drive systems and rotor drive system peratures must be corrected under components, a screen to prevent en- paragraphs (c) and (d) of this section, trance into the lubrication system of unless a more rational correction any object that might obstruct the method is applicable.

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(2) No corrected temperature deter- barrel temperature recorded during the mined under paragraph (a)(1) of this cooling test. section may exceed established limits. (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (3) The fuel used during the cooling eral Aviation Act of 1958 (49 U.S.C. 1354(a), tests must be of the minimum grade 1421, 1423, 1424, and 1425); and sec. 6(c) of the approved for the engines, and the mix- Dept. of Transportation Act (49 U.S.C. ture settings must be those used in 1655(c))) normal operation. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (4) The test procedures must be as amended by Amdt. 29–12, 41 FR 55473, Dec. 20, prescribed in §§ 29.1045 through 29.1049. 1976; Amdt. 29–15, 43 FR 2327, Jan. 16, 1978; (5) For the purposes of the cooling Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] tests, a temperature is ‘‘stabilized’’ ° § 29.1045 Climb cooling test proce- when its rate of change is less than 2 F dures. per minute. (b) Maximum ambient atmospheric tem- (a) Climb cooling tests must be con- perature. A maximum ambient atmos- ducted under this section for— pheric temperature corresponding to (1) Category A rotorcraft; and sea level conditions of at least 100 de- (2) Multiengine category B rotorcraft grees F. must be established. The as- for which certification is requested sumed temperature lapse rate is 3.6 de- under the category A powerplant in- grees F. per thousand feet of altitude stallation requirements, and under the above sea level until a temperature of requirements of § 29.861(a) at the steady rate of climb or descent established ¥69.7 degrees F. is reached, above under § 29.67(b). which altitude the temperature is con- sidered constant at ¥69.7 degrees F. (b) The climb or descent cooling tests must be conducted with the engine in- However, for winterization installa- operative that produces the most ad- tions, the applicant may select a max- verse cooling conditions for the re- imum ambient atmospheric tempera- maining engines and powerplant com- ture corresponding to sea level condi- ponents. tions of less than 100 degrees F. (c) Each operating engine must— (c) Correction factor (except cylinder (1) For helicopters for which the use Unless a more rational correc- barrels). of 30-minute OEI power is requested, be tion applies, temperatures of engine at 30-minute OEI power for 30 minutes, fluids and powerplant components (ex- and then at maximum continuous cept cylinder barrels) for which tem- power (or at full throttle when above perature limits are established, must the critical altitude); be corrected by adding to them the dif- (2) For helicopters for which the use ference between the maximum ambient of continuous OEI power is requested, atmospheric temperature and the tem- be at continuous OEI power (or at full perature of the ambient air at the time throttle when above the critical alti- of the first occurrence of the maximum tude); and component or fluid temperature re- (3) For other rotorcraft, be at max- corded during the cooling test. imum continuous power (or at full (d) Correction factor for cylinder barrel throttle when above the critical alti- temperatures. Cylinder barrel tempera- tude). tures must be corrected by adding to (d) After temperatures have sta- them 0.7 times the difference between bilized in flight, the climb must be— the maximum ambient atmospheric (1) Begun from an altitude not great- temperature and the temperature of er than the lower of— the ambient air at the time of the first (i) 1,000 feet below the engine critcal occurrence of the maximum cylinder altitude; and

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(ii) 1,000 feet below the maximum al- section, the power must be changed to titude at which the rate of climb is 150 that used in meeting § 29.67(a)(2) and f.p.m; and the climb must be continued for— (2) Continued for at least five min- (i) Thirty minutes, if 30-minute OEI utes after the occurrence of the highest power is used; or temperature recorded, or until the (ii) At least 5 minutes after the oc- rotorcraft reaches the maximum alti- currence of the highest temperature re- tude for which certification is re- corded, if continuous OEI power or quested. maximum continuous power is used. (e) For category B rotorcraft without (5) The speeds must be those used in a positive rate of climb, the descent determining the takeoff flight path must begin at the all-engine-critical under § 29.59. altitude and end at the higher of— (b) Category B. For each category B (1) The maximum altitude at which rotorcraft, cooling must be shown dur- level flight can be maintained with one ing takeoff and subsequent climb as engine operative; and follows: (2) Sea level. (1) Each temperature must be sta- (f) The climb or descent must be con- bilized while hovering in ground effect ducted at an airspeed representing a with— normal operational practice for the (i) The power necessary for hovering; configuration being tested. However, if (ii) The appropriate cowl flap and the cooling provisions are sensitive to shutter settings; and rotorcraft speed, the most critical air- (iii) The maximum weight. speed must be used, but need not ex- (2) After the temperatures have sta- ceed the speeds established under bilized, a climb must be started at the § 29.67(a)(2) or § 29.67(b). The climb cool- lowest practicable altitude with ing test may be conducted in conjunc- takeoff power. tion with the takeoff cooling test of (3) Takeoff power must be used for § 29.1047. the same time interval as takeoff power is used in determining the [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–26, 53 FR 34218, Sept. 2, takeoff flight path under § 29.63. 1988] (4) At the end of the time interval prescribed in paragraph (a)(3) of this § 29.1047 Takeoff cooling test proce- section, the power must be reduced to dures. maximum continuous power and the (a) Category A. For each category A climb must be continued for at least rotorcraft, cooling must be shown dur- five minutes after the occurence of the ing takeoff and subsequent climb as highest temperature recorded. follows: (5) The cooling test must be con- (1) Each temperature must be sta- ducted at an airspeed corresponding to bilized while hovering in ground effect normal operating practice for the con- with— figuration being tested. However, if the (i) The power necessary for hovering; cooling provisions are sensitive to (ii) The appropriate cowl flap and rotorcraft speed, the most critical air- shutter settings; and speed must be used, but need not ex- (iii) The maximum weight. ceed the speed for best rate of climb (2) After the temperatures have sta- with maximum continuous power. bilized, a climb must be started at the [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as lowest practicable altitude and must be amended by Amdt. 29–1, 30 FR 8778, July 13, conducted with one engine inoperative. 1965; Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] (3) The operating engines must be at the greatest power for which approval § 29.1049 Hovering cooling test proce- is sought (or at full throttle when dures. above the critical altitude) for the The hovering cooling provisions must same period as this power is used in de- be shown— termining the takeoff climbout path (a) At maximum weight or at the under § 29.59. greatest weight at which the rotorcraft (4) At the end of the time interval can hover (if less), at sea level, with prescribed in paragraph (b)(3) of this the power required to hover but not

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more than maximum continuous § 29.1093 Induction system icing pro- power, in the ground effect in still air, tection. until at least five minutes after the oc- (a) Reciprocating engines. Each recip- currence of the highest temperature re- rocating engine air induction system corded; and must have means to prevent and elimi- (b) With maximum continuous power, nate icing. Unless this is done by other maximum weight, and at the altitude means, it must be shown that, in air resulting in zero rate of climb for this free of visible moisture at a tempera- configuration, until at least five min- ture of 30 °F., and with the engines at utes after the occurrence of the highest 60 percent of maximum continuous temperature recorded. power— (1) Each rotorcraft with sea level en- INDUCTION SYSTEM gines using conventional venturi car- buretors has a preheater that can pro- § 29.1091 Air induction. vide a heat rise of 90 °F.; (a) The air induction system for each (2) Each rotorcraft with sea level en- gines using carburetors tending to pre- engine and auxiliary power unit must vent icing has a preheater that can supply the air required by that engine provide a heat rise of 70 °F.; and auxiliary power unit under the op- (3) Each rotorcraft with altitude en- erating conditions for which certifi- gines using conventional venturi car- cation is requested. buretors has a preheater that can pro- (b) Each engine and auxiliary power vide a heat rise of 120 °F.; and unit air induction system must provide (4) Each rotorcraft with altitude en- air for proper fuel metering and mix- gines using carburetors tending to pre- ture distribution with the induction vent icing has a preheater that can system valves in any position. provide a heat rise of 100 °F. (c) No air intake may open within (b) Turbine engines. (1) It must be the engine accessory section or within shown that each turbine engine and its other areas of any powerplant compart- air inlet system can operate through- ment where emergence of backfire out the flight power range of the en- flame would constitute a fire hazard. gine (including idling)— (d) Each reciprocating engine must (i) Without accumulating ice on en- have an alternate air source. gine or inlet system components that (e) Each alternate air intake must be would adversely affect engine oper- located to prevent the entrance of rain, ation or cause a serious loss of power ice, or other foreign matter. under the icing conditions specified in appendix C of this Part; and (f) For turbine engine powered rotor- (ii) In snow, both falling and blowing, craft and rotorcraft incorporating aux- without adverse effect on engine oper- iliary power units— ation, within the limitations estab- (1) There must be means to prevent lished for the rotorcraft. hazardous quantities of fuel leakage or (2) Each turbine engine must idle for overflow from drains, vents, or other 30 minutes on the ground, with the air components of flammable fluid systems bleed available for engine icing protec- from entering the engine or auxiliary tion at its critical condition, without power unit intake system; and adverse effect, in an atmosphere that is (2) The air inlet ducts must be lo- at a temperature between 15° and 30 °F cated or protected so as to minimize (between ¥9° and ¥1 °C) and has a liq- the ingestion of foreign matter during uid water content not less than 0.3 takeoff, landing, and taxiing. grams per cubic meter in the form of drops having a mean effective diameter (Secs. 313(a), 601, 603, 604, Federal Aviation not less than 20 microns, followed by Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), momentary operation at takeoff power sec. 6(c), Dept. of Transportation Act (49 or thrust. During the 30 minutes of idle U.S.C. 1655(c))) operation, the engine may be run up [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as periodically to a moderate power or amended by Amdt. 29–3, 33 FR 969, Jan. 26, thrust setting in a manner acceptable 1968; Amdt. 29–17, 43 FR 50601, Oct. 30, 1978] to the Administrator.

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(c) Supercharged reciprocating engines. units must be fireproof within the aux- For each engine having a supercharger iliary power unit fire zone. to pressurize the air before it enters (e) Each auxiliary power unit induc- the carburetor, the heat rise in the air tion system duct must be fireproof for caused by that supercharging at any a sufficient distance upstream of the altitude may be utilized in determining auxiliary power unit compartment to compliance with paragraph (a) of this prevent hot gas reverse flow from burn- section if the heat rise utilized is that ing through auxiliary power unit ducts which will be available, automatically, and entering any other compartment for the applicable altitude and oper- or area of the rotorcraft in which a ation condition because of super- hazard would be created resulting from charging. the entry of hot gases. The materials used to form the remainder of the in- (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 duction system duct and plenum cham- U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 ber of the auxiliary power unit must be U.S.C. 1655 (c)) capable of resisting the maximum heat [Amdt. No. 29–3, 33 FR 969, Jan. 26, 1968, as conditions likely to occur. amended by Amdt. 29–12, 41 FR 55473, Dec. 20, (f) Each auxiliary power unit induc- 1976; Amdt. 29–13, 42 FR 15046, Mar. 17, 1977; tion system duct must be constructed Amdt. 29–22, 49 FR 6850, Feb. 23, 1984; Amdt. of materials that will not absorb or 29–26, 53 FR 34219, Sept. 2, 1988] trap hazardous quantities of flammable § 29.1101 Carburetor air preheater de- fluids that could be ignited in the sign. event of a surge or reverse flow condi- tion. Each carburetor air preheater must be designed and constructed to— (Secs. 313(a), 601, 603, 604, Federal Aviation (a) Ensure ventilation of the pre- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), heater when the engine is operated in sec. 6(c), Dept. of Transportation Act (49 U.S.C. 1655(c))) cold air; (b) Allow inspection of the exhaust [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as manifold parts that it surrounds; and amended by Amdt. 29–17, 43 FR 50602, Oct. 30, 1978] (c) Allow inspection of critical parts of the preheater itself. § 29.1105 Induction system screens. § 29.1103 Induction systems ducts and If induction system screens are air duct systems. used— (a) Each screen must be upstream of (a) Each induction system duct up- the carburetor; stream of the first stage of the engine (b) No screen may be in any part of supercharger and of the auxiliary the induction system that is the only power unit compressor must have a passage through which air can reach drain to prevent the hazardous accu- the engine, unless it can be deiced by mulation of fuel and moisture in the heated air; ground attitude. No drain may dis- (c) No screen may be deiced by alco- charge where it might cause a fire haz- hol alone; and ard. (d) It must be impossible for fuel to (b) Each duct must be strong enough strike any screen. to prevent induction system failure from normal backfire conditions. § 29.1107 Inter-coolers and after-cool- (c) Each duct connected to compo- ers. nents between which relative motion Each inter-cooler and after-cooler could exist must have means for flexi- must be able to withstand the vibra- bility. tion, inertia, and air pressure loads to (d) Each duct within any fire zone for which it would be subjected in oper- which a fire-extinguishing system is re- ation. quired must be at least— (1) Fireproof, if it passes through any § 29.1109 Carburetor air cooling. firewall; or It must be shown under § 29.1043 that (2) Fire resistant, for other ducts, ex- each installation using two-stage su- cept that ducts for auxiliary power perchargers has means to maintain the

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air temperature, at the carburetor mulation after the failure of an at- inlet, at or below the maximum estab- tempted engine start. lished value. (Secs. 313(a), 601, and 603, 72 Stat. 752, 755, 49 U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 EXHAUST SYSTEM U.S.C. 1655 (c)) § 29.1121 General. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–3, 33 FR 970, Jan. 26, For powerplant and auxiliary power 1968; Amdt. 29–13, 42 FR 15046, Mar. 17, 1977] unit installations the following apply: (a) Each exhaust system must ensure § 29.1123 Exhaust piping. safe disposal of exhaust gases without (a) Exhaust piping must be heat and fire hazard or carbon monoxide con- corrosion resistant, and must have pro- tamination in any personnel compart- visions to prevent failure due to expan- ment. sion by operating temperatures. (b) Each exhaust system part with a (b) Exhaust piping must be supported surface hot enough to ignite flammable to withstand any vibration and inertia fluids or vapors must be located or loads to which it would be subjected in shielded so that leakage from any sys- operation. tem carrying flammable fluids or va- (c) Exhaust piping connected to com- pors will not result in a fire caused by ponents between which relative motion impingement of the fluids or vapors on could exist must have provisions for flexibility. any part of the exhaust system includ- ing shields for the exhaust system. § 29.1125 Exhaust heat exchangers. (c) Each component upon which hot For reciprocating engine powered exhaust gases could impinge, or that rotorcraft the following apply: could be subjected to high tempera- (a) Each exhaust heat exchanger tures from exhaust system parts, must must be constructed and installed to be fireproof. Each exhaust system com- withstand the vibration, inertia, and ponent must be separated by a fire- other loads to which it would be sub- proof shield from adjacent parts of the jected in operation. In addition— rotorcraft that are outside the engine (1) Each exchanger must be suitable and auxiliary power unit compart- for continued operation at high tem- ments. peratures and resistant to corrosion (d) No exhaust gases may discharge from exhaust gases; so as to cause a fire hazard with re- (2) There must be means for inspect- spect to any flammable fluid vent or ing the critical parts of each ex- drain. changer; (e) No exhaust gases may discharge (3) Each exchanger must have cooling where they will cause a glare seriously provisions wherever it is subject to affecting pilot vision at night. contact with exhaust gases; and (f) Each exhaust system component (4) No exhaust heat exchanger or must be ventilated to prevent points of muff may have stagnant areas or liquid traps that would increase the prob- excessively high temperature. ability of ignition of flammable fluids (g) Each exhaust shroud must be ven- or vapors that might be present in case tilated or insulated to avoid, during of the failure or malfunction of compo- normal operation, a temperature high nents carrying flammable fluids. enough to ignite any flammable fluids (b) If an exhaust heat exchanger is or vapors outside the shroud. used for heating ventilating air used by (h) If significant traps exist, each personnel— turbine engine exhaust system must (1) There must be a secondary heat have drains discharging clear of the exchanger between the primary ex- rotorcraft, in any normal ground and haust gas heat exchanger and the ven- flight attitudes, to prevent fuel accu- tilating air system; or

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(2) Other means must be used to pre- gency shutdown of each installed auxil- vent harmful contamination of the iary power unit. ventilating air. (Secs. 313(a), 601, 603, 604, Federal Aviation [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), amended by Amdt. 29–12, 41 FR 55473, Dec. 20, sec. 6(c), Dept. of Transportation Act (49 1976; Amdt. 29–41, 62 FR 46173, Aug. 29, 1997] U.S.C. 1655(c))) [Amdt. 29–17, 43 FR 50602, Oct. 30, 1978] POWERPLANT CONTROLS AND ACCESSORIES § 29.1143 Engine controls. § 29.1141 Powerplant controls: general. (a) There must be a separate power control for each engine. (a) Powerplant controls must be lo- (b) Power controls must be arranged cated and arranged under § 29.777 and to allow ready synchronization of all marked under § 29.1555. engines by— (b) Each control must be located so (1) Separate control of each engine; that it cannot be inadvertently oper- and ated by persons entering, leaving, or (2) Simultaneous control of all en- moving normally in the cockpit. gines. (c) Each flexible powerplant control (c) Each power control must provide must be approved. a positive and immediately responsive (d) Each control must be able to means of controlling its engine. maintain any set position without— (d) Each fluid injection control other (1) Constant attention; or than fuel system control must be in (2) Tendency to creep due to control the corresponding power control. How- loads or vibration. ever, the injection system pump may (e) Each control must be able to have a separate control. withstand operating loads without ex- (e) If a power control incorporates a cessive deflection. fuel shutoff feature, the control must (f) Controls of powerplant valves re- have a means to prevent the inad- quired for safety must have— vertent movement of the control into (1) For manual valves, positive stops the shutoff position. The means must— or in the case of fuel valves suitable (1) Have a positive lock or stop at the index provisions, in the open and closed idle position; and position; and (2) Require a separate and distinct (2) For power-assisted valves, a operation to place the control in the means to indicate to the flight crew shutoff position. when the valve— (f) For rotorcraft to be certificated (i) Is in the fully open or fully closed for a 30-second OEI power rating, a position; or means must be provided to automati- (ii) Is moving between the fully open cally activate and control the 30-sec- and fully closed position. ond OEI power and prevent any engine from exceeding the installed engine (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 limits associated with the 30-second U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 OEI power rating approved for the U.S.C. 1655(c)) rotorcraft. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as [Amdt. 29–26, 53 FR 34219, Sept. 2, 1988, as amended by Amdt. 29–13, 42 FR 15046, Mar. 17, amended by Amdt. 29–34, 59 FR 47768, Sept. 1977; Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] 16, 1994]

§ 29.1142 Auxiliary power unit con- § 29.1145 Ignition switches. trols. (a) Ignition switches must control Means must be provided on the flight each ignition circuit on each engine. deck for starting, stopping, and emer- (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition con- trol.

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(c) Each group of ignition switches, (b) Electrical equipment subject to except ignition switches for turbine en- arcing or sparking must be installed, gines for which continuous ignition is to minimize the probability of igniting not required, and each master ignition flammable fluids or vapors. control must have a means to prevent (c) If continued rotation of an engine- its inadvertent operation. driven cabin supercharger or any re- (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 mote accessory driven by the engine U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 will be a hazard if they malfunction, U.S.C. 1655 (c)) there must be means to prevent their hazardous rotation without interfering [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as with the continued operation of the en- amended by Amdt. 29–13, 42 FR 15046, Mar. 17, 1977] gine. (d) Unless other means are provided, § 29.1147 Mixture controls. torque limiting means must be pro- (a) If there are mixture controls, vided for accessory drives located on each engine must have a separate con- any component of the transmission and trol, and the controls must be arranged rotor drive system to prevent damage to allow— to these components from excessive ac- (1) Separate control of each engine; cessory load. and [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (2) Simultaneous control of all en- amended by Amdt. 29–22, 49 FR 6850, Feb. 23, gines. 1984; Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] (b) Each intermediate position of the mixture controls that corresponds to a § 29.1165 Engine ignition systems. normal operating setting must be iden- (a) Each battery ignition system tifiable by feel and sight. must be supplemented with a generator that is automatically available as an § 29.1151 Rotor brake controls. alternate source of electrical energy to (a) It must be impossible to apply the allow continued engine operation if rotor brake inadvertently in flight. any battery becomes depleted. (b) There must be means to warn the (b) The capacity of batteries and gen- crew if the rotor brake has not been erators must be large enough to meet completely released before takeoff. the simultaneous demands of the en- gine ignition system and the greatest § 29.1157 Carburetor air temperature demands of any electrical system com- controls. ponents that draw from the same There must be a separate carburetor source. air temperature control for each en- (c) The design of the engine ignition gine. system must account for— (1) The condition of an inoperative § 29.1159 Supercharger controls. generator; Each supercharger control must be (2) The condition of a completely de- accessible to— pleted battery with the generator run- (a) The pilots; or ning at its normal operating speed; and (b) (If there is a separate flight engi- (3) The condition of a completely de- neer station with a control panel) the pleted battery with the generator oper- flight engineer. ating at idling speed, if there is only one battery. § 29.1163 Powerplant accessories. (d) Magneto ground wiring (for sepa- (a) Each engine mounted accessory rate ignition circuits) that lies on the must— engine side of any firewall must be in- (1) Be approved for mounting on the stalled, located, or protected, to mini- engine involved; mize the probability of the simulta- (2) Use the provisions on the engine neous failure of two or more wires as a for mounting; and result of mechanical damage, electrical (3) Be sealed in such a way as to pre- fault, or other cause. vent contamination of the engine oil (e) No ground wire for any engine system and the accessory system. may be routed through a fire zone of

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another engine unless each part of that gine fire conditions and each compo- wire within that zone is fireproof. nent which conveys or contains flam- (f) Each ignition system must be mable fluid in a designated fire zone independent of any electrical circuit must be fire resistant, except that that is not used for assisting, control- flammable fluid tanks and supports in ling, or analyzing the operation of that a designated fire zone must be fireproof system. or be enclosed by a fireproof shield un- (g) There must be means to warn ap- less damage by fire to any non-fire- propriate crewmembers if the malfunc- proof part will not cause leakage or tioning of any part of the electrical spillage of flammable fluid. Compo- system is causing the continuous dis- nents must be shielded or located so as charge of any battery necessary for en- to safeguard against the ignition of gine ignition. leaking flammable fluid. An integral [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as oil sump of less than 25-quart capacity amended by Amdt. 29–12, 41 FR 55473, Dec. 20, on a reciprocating engine need not be 1976] fireproof nor be enclosed by a fireproof shield. POWERPLANT FIRE PROTECTION (b) Paragraph (a) of this section does § 29.1181 Designated fire zones: re- not apply to— gions included. (1) Lines, fittings, and components (a) Designated fire zones are— which are already approved as part of a (1) The engine power section of recip- type certificated engine; and rocating engines; (2) Vent and drain lines, and their fit- (2) The engine accessory section of tings, whose failure will not result in reciprocating engines; or add to, a fire hazard. (3) Any complete powerplant com- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as partment in which there is no isolation amended by Amdt. 29–2, 32 FR 6914, May 5, between the engine power section and 1967; Amdt. 29–10, 39 FR 35463, Oct. 1, 1974; the engine accessory section, for recip- Amdt. 29–22, 49 FR 6850, Feb. 23, 1984] rocating engines; (4) Any auxiliary power unit com- § 29.1185 Flammable fluids. partment; (a) No tank or reservoir that is part (5) Any fuel-burning heater and other of a system containing flammable combustion equipment installation de- fluids or gases may be in a designated scribed in § 29.859; fire zone unless the fluid contained, the (6) The compressor and accessory sec- design of the system, the materials tions of turbine engines; and (7) The combustor, turbine, and tail- used in the tank and its supports, the pipe sections of turbine engine instal- shutoff means, and the connections, lations except sections that do not con- lines, and controls provide a degree of tain lines and components carrying safety equal to that which would exist flammable fluids or gases and are iso- if the tank or reservoir were outside lated from the designated fire zone pre- such a zone. scribed in paragraph (a)(6) of this sec- (b) Each fuel tank must be isolated tion by a firewall that meets § 29.1191. from the engines by a firewall or (b) Each designated fire zone must shroud. meet the requirements of §§ 29.1183 (c) There must be at least one-half through 29.1203. inch of clear airspace between each tank or reservoir and each firewall or [Amdt. 29–3, 33 FR 970, Jan. 26, 1968, as amended by Amdt. 29–26, 53 FR 34219, Sept. 2, shroud isolating a designated fire zone, 1988] unless equivalent means are used to prevent heat transfer from the fire § 29.1183 Lines, fittings, and compo- zone to the flammable fluid. nents. (d) Absorbent material close to flam- (a) Except as provided in paragraph mable fluid system components that (b) of this section, each line, fitting, might leak must be covered or treated and other component carrying flam- to prevent the absorption of hazardous mable fluid in any area subject to en- quantities of fluids.

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§ 29.1187 Drainage and ventilation of (c) For category A rotorcraft, no haz- fire zones. ardous quantity of flammable fluid (a) There must be complete drainage may drain into any designated fire of each part of each designated fire zone after shutoff has been accom- zone to minimize the hazards resulting plished, nor may the closing of any fuel from failure or malfunction of any shutoff valve for an engine make fuel component containing flammable unavailable to the remaining engines. fluids. The drainage means must be— (d) The operation of any shutoff may not interfere with the later emergency (1) Effective under conditions ex- operation of any other equipment, such pected to prevail when drainage is as the means for declutching the en- needed; and gine from the rotor drive. (2) Arranged so that no discharged (e) Each shutoff valve and its control fluid will cause an additional fire haz- must be designed, located, and pro- ard. tected to function properly under any (b) Each designated fire zone must be condition likely to result from fire in a ventilated to prevent the accumulation designated fire zone. of flammable vapors. (f) Except for ground-use-only auxil- (c) No ventilation opening may be iary power unit installations, there where it would allow the entry of flam- must be means to prevent inadvertent mable fluids, vapors, or flame from operation of each shutoff and to make other zones. it possible to reopen it in flight after it (d) Ventilation means must be ar- has been closed. ranged so that no discharged vapors will cause an additional fire hazard. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (e) For category A rotorcraft, there amended by Amdt. 29–12, 41 FR 55473, Dec. 20, 1976; Amdt. 29–22, 49 FR 6850, Feb. 23, 1984; must be means to allow the crew to Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] shut off the sources of forced ventila- tion in any fire zone (other than the § 29.1191 Firewalls. engine power section of the powerplant (a) Each engine, including the com- compartment) unless the amount of ex- bustor, turbine, and tailpipe sections of tinguishing agent and the rate of dis- turbine engine installations, must be charge are based on the maximum air- isolated by a firewall, shroud, or equiv- flow through that zone. alent means, from personnel compart- § 29.1189 Shutoff means. ments, structures, controls, rotor mechanisms, and other parts that are— (a) There must be means to shut off (1) Essential to controlled flight and or otherwise prevent hazardous quan- landing; and tities of fuel, oil, de-icing fluid, and (2) Not protected under § 29.861. other flammable fluids from flowing (b) Each auxiliary power unit, com- into, within, or through any designated bustion heater, and other combustion fire zone, except that this means need equipment to be used in flight, must be not be provided— isolated from the rest of the rotorcraft (1) For lines, fittings, and compo- by firewalls, shrouds, or equivalent nents forming an integral part of an means. engine; (c) Each firewall or shroud must be (2) For oil systems for turbine engine constructed so that no hazardous quan- installations in which all components tity of air, fluid, or flame can pass of the system, including oil tanks, are from any engine compartment to other fireproof or located in areas not subject parts of the rotorcraft. to engine fire conditions; or (d) Each opening in the firewall or (3) For engine oil systems in category shroud must be sealed with close-fit- B rotorcraft using reciprocating en- ting fireproof grommets, bushings, or gines of less than 500 cubic inches dis- firewall fittings. placement. (e) Each firewall and shroud must be (b) The closing of any fuel shutoff fireproof and protected against corro- valve for any engine may not make sion. fuel unavailable to the remaining en- (f) In meeting this section, account gines. must be taken of the probable path of

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a fire as affected by the airflow in nor- (2) Fire in a fire zone, if such fire mal flight and in autorotation. could adversely affect the normal means of retention. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–3, 33 FR 970, Jan. 26, (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 1968] U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 U.S.C. 1655(c)) § 29.1193 Cowling and engine compart- ment covering. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–3, 33 FR 970, Jan. 26, (a) Each cowling and engine compart- 1968; Amdt. 29–13, 42 FR 15046, Mar. 17, 1977; ment covering must be constructed and Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] supported so that it can resist the vi- bration, inertia, and air loads to which § 29.1194 Other surfaces. it may be subjected in operation. All surfaces aft of, and near, engine (b) Cowling must meet the drainage compartments and designated fire and ventilation requirements of zones, other than tail surfaces not sub- § 29.1187. ject to heat, flames, or sparks ema- (c) On rotorcraft with a diaphragm nating from a designated fire zone or isolating the engine power section from engine compartment, must be at least the engine accessory section, each part fire resistant. of the accessory section cowling sub- [Amdt. 29–3, 33 FR 970, Jan. 26, 1968] ject to flame in case of fire in the en- gine power section of the powerplant § 29.1195 Fire extinguishing systems. must— (1) Be fireproof; and (a) Each turbine engine powered (2) Meet the requirements of § 29.1191. rotorcraft and Category A recipro- cating engine powered rotorcraft, and (d) Each part of the cowling or engine each Category B reciprocating engine compartment covering subject to high powered rotorcraft with engines of temperatures due to its nearness to ex- more than 1,500 cubic inches must have haust system parts or exhaust gas im- a fire extinguishing system for the des- pingement must be fireproof. ignated fire zones. The fire extin- (e) Each rotorcraft must— guishing system for a powerplant must (1) Be designated and constructed so be able to simultaneously protect all that no fire originating in any fire zone zones of the powerplant compartment can enter, either through openings or for which protection is provided. by burning through external skin, any (b) For multiengine powered rotor- other zone or region where it would craft, the fire extinguishing system, create additional hazards; the quantity of extinguishing agent, (2) Meet the requirements of para- and the rate of discharge must— graph (e)(1) of this section with the (1) For each auxiliary power unit and landing gear retracted (if applicable); combustion equipment, provide at least and one adequate discharge; and (3) Have fireproof skin in areas sub- (2) For each other designated fire ject to flame if a fire starts in or burns zone, provide two adequate discharges. out of any designated fire zone. (c) For single engine rotorcraft, the (f) A means of retention for each quantity of extinguishing agent and openable or readily removable panel, the rate of discharge must provide at cowling, or engine or rotor drive sys- least one adequate discharge for the tem covering must be provided to pre- engine compartment. clude hazardous damage to rotors or (d) It must be shown by either actual critical control components in the or simulated flight tests that under event of— critical airflow conditions in flight the (1) Structural or mechanical failure discharge of the extinguishing agent in of the normal retention means, unless each designated fire zone will provide such failure is extremely improbable; an agent concentration capable of ex- or tinguishing fires in that zone and of

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minimizing the probability of reigni- (d) The temperature of each con- tion. tainer must be maintained, under in- tended operating conditions, to prevent (Secs. 313(a), 601, 603, 604, Federal Aviation the pressure in the container from— Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 (1) Falling below that necessary to U.S.C. 1655(c))) provide an adequate rate of discharge; or [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (2) Rising high enough to cause pre- amended by Amdt. 29–3, 33 FR 970, Jan. 26, mature discharge. 1968; Amdt. 29–13, 42 FR 15047, Mar. 17, 1977; Amdt. 29–17, 43 FR 50602, Oct. 30, 1978] (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 § 29.1197 Fire extinguishing agents. U.S.C. 1655 (c)) (a) Fire extinguishing agents must— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (1) Be capable of extinguishing amended by Amdt. 29–13, 42 FR 15047, Mar. 17, flames emanating from any burning of 1977] fluids or other combustible materials § 29.1201 Fire extinguishing system in the area protected by the fire extin- materials. guishing system; and (2) Have thermal stability over the (a) No materials in any fire extin- temperature range likely to be experi- guishing system may react chemically enced in the compartment in which with any extinguishing agent so as to they are stored. create a hazard. (b) If any toxic extinguishing agent is (b) Each system component in an en- used, it must be shown by test that gine compartment must be fireproof. entry of harmful concentrations of § 29.1203 Fire detector systems. fluid or fluid vapors into any personnel compartment (due to leakage during (a) For each turbine engine powered normal operation of the rotorcraft, or rotorcraft and Category A recipro- discharge on the ground or in flight) is cating engine powered rotorcraft, and prevented, even though a defect may for each Category B reciprocating en- exist in the extinguishing system. gine powered rotorcraft with engines of more than 900 cubic inches displace- (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 ment, there must be approved, quick- U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 acting fire detectors in designated fire U.S.C. 1655(c)) zones and in the combustor, turbine, [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as and tailpipe sections of turbine instal- amended by Amdt. 29–12, 41 FR 55473, Dec. 20, lations (whether or not such sections 1976; Amdt. 29–13, 42 FR 15047, Mar. 17, 1977] are designated fire zones) in numbers and locations ensuring prompt detec- § 29.1199 Extinguishing agent con- tion of fire in those zones. tainers. (b) Each fire detector must be con- (a) Each extinguishing agent con- structed and installed to withstand any tainer must have a pressure relief to vibration, inertia, and other loads to prevent bursting of the container by which it would be subjected in oper- excessive internal pressures. ation. (b) The discharge end of each dis- (c) No fire detector may be affected charge line from a pressure relief con- by any oil, water, other fluids, or nection must be located so that dis- fumes that might be present. charge of the fire extinguishing agent (d) There must be means to allow would not damage the rotorcraft. The crewmembers to check, in flight, the line must also be located or protected functioning of each fire detector sys- to prevent clogging caused by ice or tem electrical circuit. other foreign matter. (e) The writing and other components (c) There must be a means for each of each fire detector system in an en- fire extinguishing agent container to gine compartment must be at least fire indicate that the container has dis- resistant. charged or that the charging pressure (f) No fire detector system compo- is below the established minimum nec- nent for any fire zone may pass essary for proper functioning. through another fire zone, unless—

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(1) It is protected against the possi- (2) Is powered from a source inde- bility of false warnings resulting from pendent of the electrical generating fires in zones through which it passes; system; or (3) Continues reliable operation for a (2) The zones involved are simulta- minimum of 30 minutes after total fail- neously protected by the same detector ure of the electrical generating system; and extinguishing systems. (4) Operates independently of any other altitude indicating system; [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–3, 33 FR 970, Jan. 26, (5) Is operative without selection 1968] after total failure of the electrical gen- erating system; Subpart F—Equipment (6) Is located on the instrument panel in a position acceptable to the Admin- GENERAL istrator that will make it plainly visi- ble to and useable by any pilot at his § 29.1301 Function and installation. station; and Each item of installed equipment (7) Is appropriately lighted during all must— phases of operation. (a) Be of a kind and design appro- (h) A gyroscopic direction indicator. priate to its intended function; (i) A rate-of-climb (vertical speed) in- (b) Be labeled as to its identification, dicator. function, or operating limitations, or (j) For Category A rotorcraft, a speed any applicable combination of these warning device when VNE is less than factors; the speed at which unmistakable over- (c) Be installed according to limita- speed warning is provided by other tions specified for that equipment; and pilot cues. The speed warning device (d) Function properly when installed. must give effective aural warning (dif- fering distinctively from aural warn- § 29.1303 Flight and navigation instru- ings used for other purposes) to the pi- ments. lots whenever the indicated speed ex- The following are required flight and ceeds VNE plus 3 knots and must oper- navigational instruments: ate satisfactorily throughout the ap- (a) An airspeed indicator. For Cat- proved range of altitudes and tempera- tures. egory A rotorcraft with VNE less than a speed at which unmistakable pilot cues (Secs. 313(a), 601, 603, 604, and 605 of the Fed- provide overspeed warning, a maximum eral Aviation Act of 1958 (49 U.S.C. 1354(a), allowable airspeed indicator must be 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. provided. If maximum allowable air- of Transportation Act (49 U.S.C. 1655(c))) speed varies with weight, altitude, [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as temperature, or r.p.m., the indicator amended by Amdt. 29–12, 41 FR 55474, Dec. 20, must show that variation. 1976; Amdt. 29–14, 42 FR 36972, July 18, 1977; (b) A sensitive altimeter. Amdt. 29–24, 49 FR 44438, Nov. 6, 1984] (c) A magnetic direction indicator. (d) A clock displaying hours, min- § 29.1305 Powerplant instruments. utes, and seconds with a sweep-second The following are required power- pointer or digital presentation. plant instruments: (e) A free-air temperature indicator. (a) For each rotorcraft— (f) A non-tumbling gyroscopic bank (1) A carburetor air temperature indi- and pitch indicator. cator for each reciprocating engine; (g) A gyroscopic rate-of-turn indi- (2) A cylinder head temperature indi- cator combined with an integral slip- cator for each air-cooled reciprocating skid indicator (turn-and-bank indi- engine, and a coolant temperature indi- cator) except that only a slip-skid indi- cator for each liquid-cooled recipro- cator is required on rotorcraft with a cating engine; third altitude instrument system (3) A fuel quantity indicator for each that— fuel tank; (1) Is useable through flight altitudes (4) A low fuel warning device for each of ± 80 degrees of pitch and ± 120 de- fuel tank which feeds an engine. This grees of roll; device must—

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(i) Provide a warning to the crew contamination of the strainer or filter when approximately 10 minutes of usa- before it reaches the capacity estab- ble fuel remains in the tank; and lished in accordance with § 29.1019(a)(2); (ii) Be independent of the normal fuel (20) An indicator to indicate the func- quantity indicating system. tioning of any selectable or control- (5) A manifold pressure indicator, for lable heater used to prevent ice clog- each reciprocating engine of the alti- ging of fuel system components; tude type; (21) An individual fuel pressure indi- (6) An oil pressure indicator for each cator for each engine, unless the fuel pressure-lubricated gearbox. system which supplies that engine does (7) An oil pressure warning device for not employ any pumps, filters, or other each pressure-lubricated gearbox to in- components subject to degradation or dicate when the oil pressure falls below failure which may adversely affect fuel a safe value; pressure at the engine; (8) An oil quantity indicator for each (22) A means to indicate to the oil tank and each rotor drive gearbox, flightcrew the failure of any fuel pump if lubricant is self-contained; installed to show compliance with (9) An oil temperature indicator for § 29.955; each engine; (23) Warning or caution devices to (10) An oil temperature warning de- signal to the flightcrew when ferro- vice to indicate unsafe oil tempera- magnetic particles are detected by the tures in each main rotor drive gearbox, chip detector required by § 29.1337(e); including gearboxes necessary for rotor and phasing; (24) For auxiliary power units, an in- (11) A gas temperature indicator for dividual indicator, warning or caution each turbine engine; device, or other means to advise the (12) A gas producer rotor tachometer flightcrew that limits are being exceed- for each turbine engine; ed, if exceeding these limits can be haz- (13) A tachometer for each engine ardous, for— that, if combined with the applicable (i) Gas temperature; instrument required by paragraph (ii) Oil pressure; and (a)(14) of this section, indicates rotor (iii) Rotor speed. r.p.m. during autorotation. (14) At least one tachometer to indi- (25) For rotorcraft for which a 30-sec- cate, as applicable— ond/2-minute OEI power rating is re- (i) The r.p.m. of the single main quested, a means must be provided to rotor; alert the pilot when the engine is at (ii) The common r.p.m. of any main the 30-second and 2-minute OEI power rotors whose speeds cannot vary appre- levels, when the event begins, and ciably with respect to each other; and when the time interval expires. (iii) The r.p.m. of each main rotor (26) For each turbine engine utilizing whose speed can vary appreciably with 30-second/2-minute OEI power, a device respect to that of another main rotor; or system must be provided for use by (15) A free power turbine tachometer ground personnel which— for each turbine engine; (i) Automatically records each usage (16) A means, for each turbine engine, and duration of power at the 30-second to indicate power for that engine; and 2-minute OEI levels; (17) For each turbine engine, an indi- (ii) Permits retrieval of the recorded cator to indicate the functioning of the data; powerplant ice protection system; (iii) Can be reset only by ground (18) An indicator for the filter re- maintenance personnel; and quired by § 29.997 to indicate the occur- (iv) Has a means to verify proper op- rence of contamination of the filter to eration of the system or device. the degree established in compliance (b) For category A rotorcraft— with § 29.955; (1) An individual oil pressure indi- (19) For each turbine engine, a warn- cator for each engine, and either an ing means for the oil strainer or filter independent warning device for each required by § 29.1019, if it has no bypass, engine or a master warning device for to warn the pilot of the occurrence of the engines with means for isolating

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the individual warning circuit from the (ii) The occurrence of any other fail- master warning device; ure conditions which would reduce the (2) An independent fuel pressure capability of the rotorcraft or the abil- warning device for each engine or a ity of the crew to cope with adverse op- master warning device for all engines erating conditions is improbable. with provision for isolating the indi- (c) Warning information must be pro- vidual warning device from the master vided to alert the crew to unsafe sys- warning device; and tem operating conditions and to enable (3) Fire warning indicators. them to take appropriate corrective (c) For category B rotorcraft— action. Systems, controls, and associ- (1) An individual oil pressure indi- ated monitoring and warning means cator for each engine; and must be designed to minimize crew er- (2) Fire warning indicators, when fire rors which could create additional haz- detection is required. ards. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (d) Compliance with the require- amended by Amdt. 29–3, 33 FR 970, Jan. 26, ments of paragraph (b)(2) of this sec- 1968; Amdt. 29–10, 39 FR 35463, Oct. 1, 1974; tion must be shown by analysis and, Amdt. 29–26, 53 FR 34219, Sept. 2, 1988; Amdt. where necessary, by appropriate 29–34, 59 FR 47768, Sept. 16, 1994; Amdt. 29–40, ground, flight, or simulator tests. The 61 FR 21908, May 10, 1996; 61 FR 43952, Aug. 27, analysis must consider— 1996] (1) Possible modes of failure, includ- § 29.1307 Miscellaneous equipment. ing malfunctions and damage from ex- ternal sources; The following is required miscella- neous equipment: (2) The probability of multiple fail- (a) An approved seat for each occu- ures and undetected failures; pant. (3) The resulting effects on the rotor- (b) A master switch arrangement for craft and occupants, considering the electrical circuits other than ignition. stage of flight and operating condi- (c) Hand fire extinguishers. tions; and (d) A windshield wiper or equivalent (4) The crew warning cues, corrective device for each pilot station. action required, and the capability of (e) A two-way radio communication detecting faults. system. (e) For Category A rotorcraft, each installation whose functioning is re- [Amdt. 29–12, 41 FR 55473, Dec. 20, 1976] quired by this subchapter and which re- § 29.1309 Equipment, systems, and in- quires a power supply is an ‘‘essential stallations. load’’ on the power supply. The power sources and the system must be able to (a) The equipment, systems, and in- supply the following power loads in stallations whose functioning is re- probable operating combinations and quired by this subchapter must be de- for probable durations: signed and installed to ensure that they perform their intended functions (1) Loads connected to the system under any foreseeable operating condi- with the system functioning normally. tion. (2) Essential loads, after failure of (b) The rotorcraft systems and asso- any one prime mover, power converter, ciated components, considered sepa- or energy storage device. rately and in relation to other systems, (3) Essential loads, after failure of— must be designed so that— (i) Any one engine, on rotorcraft with (1) For Category B rotorcraft, the two engines; and equipment, systems, and installations (ii) Any two engines, on rotorcraft must be designed to prevent hazards to with three or more engines. the rotorcraft if they malfunction or (f) In determining compliance with fail; or paragraphs (e)(2) and (3) of this section, (2) For Category A rotorcraft— the power loads may be assumed to be (i) The occurrence of any failure con- reduced under a monitoring procedure dition which would prevent the contin- consistent with safety in the kinds of ued safe flight and landing of the rotor- operations authorized. Loads not re- craft is extremely improbable; and quired for controlled flight need not be

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considered for the two-engine-inoper- (2) The instrument that most effec- ative condition on rotorcraft with tively indicates direction of flight three or more engines. must be adjacent to and directly below (g) In showing compliance with para- the attitude instrument; graphs (a) and (b) of this section with (3) The instrument that most effec- regard to the electrical system and to tively indicates airspeed must be adja- equipment design and installation, cent to and to the left of the attitude critical environmental conditions must instrument; and be considered. For electrical genera- (4) The instrument that most effec- tion, distribution, and utilization tively indicates altitude or is most fre- equipment required by or used in com- quently utilized in control of altitude plying with this subchapter, except must be adjacent to and to the right of equipment covered by Technical Stand- the attitude instrument. ard Orders containing environmental (c) Other required powerplant instru- test procedures, the ability to provide ments must be closely grouped on the continuous, safe service under foresee- instrument panel. able environmental conditions may be (d) Identical powerplant instruments shown by environmental tests, design for the engines must be located so as to analysis, or reference to previous com- prevent any confusion as to which en- parable service experience on other air- gine each instrument relates. craft. (h) In showing compliance with para- (e) Each powerplant instrument vital graphs (a) and (b) of this section, the to safe operation must be plainly visi- effects of lightning strikes on the ble to appropriate crewmembers. rotorcraft must be considered. (f) Instrument panel vibration may not damage, or impair the readability (Secs. 313(a), 601, 603, 604, and 605 of the Fed- or accuracy of, any instrument. eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. (g) If a visual indicator is provided to of Transportation Act (49 U.S.C. 1655(c))) indicate malfunction of an instrument, it must be effective under all probable [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as cockpit lighting conditions. amended by Amdt. 29–14, 42 FR 36972, July 18, 1977; Amdt. 29–24, 49 FR 44438, Nov. 6, 1984; (Secs. 313(a), 601, 603, 604, and 605 of the Fed- Amdt. 29–40, 61 FR 21908, May 10, 1996] eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. INSTRUMENTS: INSTALLATION of Transportation Act (49 U.S.C. 1655(c))) § 29.1321 Arrangement and visibility. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–14, 42 FR 36972, July 18, (a) Each flight, navigation, and pow- 1977; Amdt. 29–21, 48 FR 4391, Jan. 31, 1983] erplant instrument for use by any pilot must be easily visible to him from his § 29.1322 Warning, caution, and advi- station with the minimum practicable sory lights. deviation from his normal position and If warning, caution or advisory lights line of vision when he is looking for- are installed in the cockpit they must, ward along the flight path. unless otherwise approved by the Ad- (b) Each instrument necessary for ministrator, be— safe operation, including the airspeed indicator, gyroscopic direction indi- (a) Red, for warning lights (lights in- cator, gyroscopic bank-and-pitch indi- dicating a hazard which may require cator, slip-skid indicator, altimeter, immediate corrective action); rate-of-climb indicator, rotor tachom- (b) Amber, for caution lights (lights eters, and the indicator most rep- indicating the possible need for future resentative of engine power, must be corrective action); grouped and centered as nearly as prac- (c) Green, for safe operation lights; ticable about the vertical plane of the and pilot’s forward vision. In addition, for (d) Any other color, including white, rotorcraft approved for IFR flight— for lights not described in paragraphs (1) The instrument that most effec- (a) through (c) of this section, provided tively indicates attitude must be on the color differs sufficiently from the the panel in the top center position; colors prescribed in paragraphs (a)

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through (c) of this section to avoid pos- (f) Each system must have a heated sible confusion. pitot tube or an equivalent means of preventing malfunction due to icing. [Amdt. 29–12, 41 FR 55474, Dec. 20, 1976] [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964 as § 29.1323 Airspeed indicating system. amended by Amdt. 29–3, 33 FR 970, Jan. 26, 1968; Amdt. 29–24, 49 FR 44439, Nov. 6, 1984; For each airspeed indicating system, Amdt. 29–39, 61 FR 21901, May 10, 1996; Amdt. the following apply: 29–44, 64 FR 45338, Aug. 19, 1999] (a) Each airspeed indicating instru- ment must be calibrated to indicate § 29.1325 Static pressure and pressure true airspeed (at sea level with a stand- altimeter systems. ard atmosphere) with a minimum prac- (a) Each instrument with static air ticable instrument calibration error case connections must be vented to the when the corresponding pitot and stat- outside atmosphere through an appro- ic pressures are applied. priate piping system. (b) Each system must be calibrated (b) Each vent must be located where to determine system error excluding its orifices are least affected by airflow airspeed instrument error. This cali- variation, moisture, or foreign matter. bration must be determined— (c) Each static pressure port must be designed and located in such manner (1) In level flight at speeds of 20 that the correlation between air pres- knots and greater, and over an appro- sure in the static pressure system and priate range of speeds for flight condi- true ambient atmospheric static pres- tions of climb and autorotation; and sure is not altered when the rotorcraft (2) During takeoff, with repeatable encounters icing conditions. An anti- and readable indications that ensure— icing means or an alternate source of (i) Consistent realization of the field static pressure may be used in showing lengths specified in the Rotorcraft compliance with this requirement. If Flight Manual; and the reading of the altimeter, when on (ii) Avoidance of the critical areas of the alternate static pressure system, the height-velocity envelope as estab- differs from the reading of altimeter lished under § 29.87. when on the primary static system by (c) For Category A rotorcraft— more than 50 feet, a correction card (1) The indication must allow con- must be provided for the alternate sistent definition of the takeoff deci- static system. sion point; and (d) Except for the vent into the at- (2) The system error, excluding the mosphere, each system must be air- airspeed instrument calibration error, tight. may not exceed— (e) Each pressure altimeter must be (i) Three percent or 5 knots, which- approved and calibrated to indicate ever is greater, in level flight at speeds pressure altitude in a standard atmos- phere with a minimum practicable above 80 percent of takeoff safety calibration error when the cor- speed; and responding static pressures are applied. (ii) Ten knots in climb at speeds from (f) Each system must be designed and 10 knots below takeoff safety speed to installed so that an error in indicated 10 knots above VY. pressure altitude, at sea level, with a (d) For Category B rotorcraft, the standard atmosphere, excluding instru- system error, excluding the airspeed ment calibration error, does not result instrument calibration error, may not in an error of more than ±30 feet per 100 exceed 3 percent or 5 knots, whichever knots speed. However, the error need is greater, in level flight at speeds not be less than ±30 feet. above 80 percent of the climbout speed (g) Except as provided in paragraph attained at 50 feet when complying (h) of this section, if the static pressure with § 29.63. system incorporates both a primary (e) Each system must be arranged, so and an alternate static pressure source, far as practicable, to prevent malfunc- the means for selecting one or the tion or serious error due to the entry of other source must be designed so moisture, dirt, or other substances. that—

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(1) When either source is selected, the the event of a malfunction, assuming other is blocked off; and that corrective action begins within a (2) Both sources cannot be blocked reasonable period of time. off simultaneously. (e) If the automatic pilot integrates (h) For unpressurized rotorcraft, signals from auxiliary controls or fur- paragraph (g)(1) of this section does not nishes signals for operation of other apply if it can be demonstrated that equipment, there must be positive the static pressure system calibration, interlocks and sequencing of engage- when either static pressure source is ment to prevent improper operation. selected, is not changed by the other (f) If the automatic pilot system can static pressure source being open or be coupled to airborne navigation blocked. equipment, means must be provided to indicate to the pilots the current mode (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), of operation. Selector switch position 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. is not acceptable as a means of indica- of Transportation Act (49 U.S.C. 1655(c))) tion. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–14, 42 FR 36972, July 18, amended by Amdt. 29–24, 49 FR 44439, Nov. 6, 1977; Amdt. 29–24, 49 FR 44439, Nov. 6, 1984] 1984; Amdt. 29–24, 49 FR 47594, Dec. 6, 1984; Amdt. 29–42, 63 FR 43285, Aug. 12, 1998] § 29.1327 Magnetic direction indicator. (a) Each magnetic direction indicator § 29.1331 Instruments using a power supply. must be installed so that its accuracy is not excessively affected by the For category A rotorcraft— rotorcraft’s vibration or magnetic (a) Each required flight instrument fields. using a power supply must have— (b) The compensated installation (1) Two independent sources of power; may not have a deviation, in level (2) A means of selecting either power flight, greater than 10 degrees on any source; and heading. (3) A visual means integral with each instrument to indicate when the power § 29.1329 Automatic pilot system. adequate to sustain proper instrument (a) Each automatic pilot system performance is not being supplied. The must be designed so that the automatic power must be measured at or near the pilot can— point where it enters the instrument. (1) Be sufficiently overpowered by For electrical instruments, the power one pilot to allow control of the rotor- is considered to be adequate when the craft; and voltage is within the approved limits; (2) Be readily and positively dis- and engaged by each pilot to prevent it (b) The installation and power supply from interfering with the control of the system must be such that failure of rotorcraft. any flight instrument connected to one (b) Unless there is automatic syn- source, or of the energy supply from chronization, each system must have a one source, or a fault in any part of the means to readily indicate to the pilot power distribution system does not the alignment of the actuating device interfere with the proper supply of en- in relation to the control system it op- ergy from any other source. erates. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (c) Each manually operated control amended by Amdt. 29–24, 49 FR 44439, Nov. 6, for the system’s operation must be 1984] readily accessible to the pilots. (d) The system must be designed and § 29.1333 Instrument systems. adjusted so that, within the range of For systems that operate the re- adjustment available to the pilot, it quired flight instruments which are lo- cannot produce hazardous loads on the cated at each pilot’s station, the fol- rotorcraft, or create hazardous devi- lowing apply: ations in the flight path, under any (a) Only the required flight instru- flight condition appropriate to its use, ments for the first pilot may be con- either during normal operation or in nected to that operating system.

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(b) The equipment, systems, and in- or equivalent units, of usable fuel in stallations must be designed so that each tank during flight. In addition— one display of the information essen- (1) Each fuel quantity indicator must tial to the safety of flight which is pro- be calibrated to read ‘‘zero’’ during vided by the flight instruments re- level flight when the quantity of fuel mains available to a pilot, without ad- remaining in the tank is equal to the ditional crewmember action, after any unusable fuel supply determined under single failure or combination of fail- § 29.959; ures that are not shown to be ex- (2) When two or more tanks are close- tremely improbable. ly interconnected by a gravity feed sys- (c) Additional instruments, systems, tem and vented, and when it is impos- or equipment may not be connected to sible to feed from each tank sepa- the operating system for a second pilot rately, at least one fuel quantity indi- unless provisions are made to ensure cator must be installed; the continued normal functioning of (3) Tanks with interconnected outlets the required flight instruments in the and airspaces may be treated as one event of any malfunction of the addi- tank and need not have separate indi- tional instruments, systems, or equip- cators; and ment which is not shown to be ex- (4) Each exposed sight gauge used as tremely improbable. a fuel quantity indicator must be pro- tected against damage. [Amdt. 29–24, 49 FR 44439, Nov. 6, 1984] (c) Fuel flowmeter system. If a fuel flowmeter system is installed, each § 29.1335 Flight director systems. metering component must have a If a flight director system is in- means for bypassing the fuel supply if stalled, means must be provided to in- malfunction of that component se- dicate to the flight crew its current verely restricts fuel flow. mode of operation. Selector switch po- (d) Oil quantity indicator. There must sition is not acceptable as a means of be a stick gauge or equivalent means indication. to indicate the quantity of oil— (1) In each tank; and (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), (2) In each transmission gearbox. 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. (e) Rotor drive system transmissions of Transportation Act (49 U.S.C. 1655(c))) and gearboxes utilizing ferromagnetic materials must be equipped with chip [Amdt. 29–14, 42 FR 36973, July 18, 1977] detectors designed to indicate the pres- ence of ferromagnetic particles result- § 29.1337 Powerplant instruments. ing from damage or excessive wear (a) Instruments and instrument lines. within the transmission or gearbox. (1) Each powerplant and auxiliary Each chip detector must— power unit instrument line must meet (1) Be designed to provide a signal to the requirements of §§ 29.993 and 29.1183. the indicator required by (2) Each line carrying flammable § 29.1305(a)(22); and fluids under pressure must— (2) Be provided with a means to allow (i) Have restricting orifices or other crewmembers to check, in flight, the safety devices at the source of pressure function of each detector electrical cir- to prevent the escape of excessive fluid cuit and signal. if the line fails; and (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 (ii) Be installed and located so that U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 the escape of fluids would not create a U.S.C. 1655(c)) hazard. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (3) Each powerplant and auxiliary amended by Amdt. 29–13, 42 FR 15047, Mar. 17, power unit instrument that utilizes 1977; Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] flammable fluids must be installed and located so that the escape of fluid ELECTRICAL SYSTEMS AND EQUIPMENT would not create a hazard. (b) Fuel quantity indicator. There § 29.1351 General. must be means to indicate to the flight (a) Electrical system capacity. The re- crew members the quantity, in gallons quired generating capacity and the

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number and kind of power sources sources excluding the battery) inoper- must— ative, with critical type fuel (from the (1) Be determined by an electrical standpoint of flameout and restart ca- load analysis; and pability), and with the rotorcraft ini- (2) Meet the requirements of § 29.1309. tially at the maximum certificated al- (b) Generating system. The generating titude. Parts of the electrical system system includes electrical power may remain on if— sources, main power busses, trans- (i) A single malfunction, including a mission cables, and associated control, wire bundle or junction box fire, can- regulation, and protective devices. It not result in loss of the part turned off must be designed so that— and the part turned on; (1) Power sources function properly (ii) The parts turned on are elec- when independent and when connected trically and mechanically isolated in combination; from the parts turned off; and (2) No failure or malfunction of any (2) Additional requirements for Cat- power source can create a hazard or egory A Rotorcraft. impair the ability of remaining sources (i) Unless it can be shown that the to supply essential loads; loss of the normal electrical power gen- (3) The system voltage and frequency erating system is extremely improb- (as applicable) at the terminals of es- able, an emergency electrical power sential load equipment can be main- system, independent of the normal tained within the limits for which the electrical power generating system, equipment is designed, during any must be provided, with sufficient ca- probable operating condition; pacity to power all systems necessary (4) System transients due to switch- for continued safe flight and landing. ing, fault clearing, or other causes do (ii) Failures, including junction box, not make essential loads inoperative, control panel, or wire bundle fires, and do not cause a smoke or fire haz- which would result in the loss of the ard; normal and emergency systems, must (5) There are means accessible in be shown to be extremely improbable. flight to appropriate crewmembers for (iii) Systems necessary for imme- the individual and collective dis- diate safety must continue to operate connection of the electrical power following the loss of the normal elec- sources from the main bus; and trical power generating system, with- (6) There are means to indicate to ap- out the need for flight crew action. propriate crewmembers the generating system quantities essential for the safe (Secs. 313(a), 601, 603, 604, and 605 of the Fed- operation of the system, such as the eral Aviation Act of 1958 (49 U.S.C. 1354(a), voltage and current supplied by each 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. generator. of Transportation Act (49 U.S.C. 1655(c))) (c) External power. If provisions are [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as made for connecting external power to amended by Amdt. 29–14, 42 FR 36973, July 18, the rotorcraft, and that external power 1977; Amdt. 29–40, 61 FR 21908, May 10, 1996; can be electrically connected to equip- Amdt. 29–42, 63 FR 43285, Aug. 12, 1998] ment other than that used for engine starting, means must be provided to § 29.1353 Electrical equipment and in- ensure that no external power supply stallations. having a reverse polarity, or a reverse (a) Electrical equipment, controls, phase sequence, can supply power to and wiring must be installed so that the rotorcraft’s electrical system. operation of any one unit or system of (d) Operation with the normal elec- units will not adversely affect the si- trical power generating system inoper- multaneous operation of any other ative. electrical unit or system essential to (1) It must be shown by analysis, safe operation. tests, or both, that the rotorcraft can (b) Cables must be grouped, routed, be operated safely in VFR conditions and spaced so that damage to essential for a period of not less than 5 minutes, circuits will be minimized if there are with the normal electrical power gen- faults in heavy current-carrying ca- erating system (electrical power bles.

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(c) Storage batteries must be de- ing source in the event of battery fail- signed and installed as follows: ure. (1) Safe cell temperatures and pres- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- sures must be maintained during any eral Aviation Act of 1958 (49 U.S.C. 1354(a), probable charging and discharging con- 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. dition. No uncontrolled increase in cell of Transportation Act (49 U.S.C. 1655(c))) temperature may result when the bat- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as tery is recharged (after previous com- amended by Amdt. 29–14, 42 FR 36973, July 18, plete discharge)— 1977; Amdt. 29–15, 43 FR 2327, Jan. 16, 1978] (i) At maximum regulated voltage or power; § 29.1355 Distribution system. (ii) During a flight of maximum dura- (a) The distribution system includes tion; and the distribution busses, their associ- (iii) Under the most adverse cooling ated feeders, and each control and pro- condition likely in service. tective device. (2) Compliance with paragraph (a)(1) (b) If two independent sources of of this section must be shown by test electrical power for particular equip- unless experience with similar bat- ment or systems are required by this teries and installations has shown that chapter, in the event of the failure of maintaining safe cell temperatures and one power source for such equipment or pressures presents no problem. system, another power source (includ- (3) No explosive or toxic gases emit- ing its separate feeder) must be pro- ted by any battery in normal oper- vided automatically or be manually se- ation, or as the result of any probable lectable to maintain equipment or sys- malfunction in the charging system or tem operation. battery installation, may accumulate (Secs. 313(a), 601, 603, 604, and 605 of the Fed- in hazardous quantities within the eral Aviation Act of 1958 (49 U.S.C. 1354(a), rotorcraft. 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. (4) No corrosive fluids or gases that of Transportation Act (49 U.S.C. 1655(c))) may escape from the battery may dam- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as age surrounding structures or adjacent amended by Amdt. 29–14, 42 FR 36973, July 18, essential equipment. 1977; Amdt. 29–24, 49 FR 44439, Nov. 6, 1984] (5) Each nickel cadmium battery in- § 29.1357 Circuit protective devices. stallation capable of being used to start an engine or auxiliary power unit (a) Automatic protective devices must have provisions to prevent any must be used to minimize distress to hazardous effect on structure or essen- the electrical system and hazard to the tial systems that may be caused by the rotorcraft system and hazard to the maximum amount of heat the battery rotorcraft in the event of wiring faults or serious malfunction of the system or can generate during a short circuit of connected equipment. the battery or of its individual cells. (b) The protective and control de- (6) Nickel cadmium battery installa- vices in the generating system must be tions capable of being used to start an designed to de-energize and disconnect engine or auxiliary power unit must faulty power sources and power trans- have— mission equipment from their associ- (i) A system to control the charging ated buses with sufficient rapidity to rate of the battery automatically so as provide protection from hazardous to prevent battery overheating; overvoltage and other malfunctioning. (ii) A battery temperature sensing (c) Each resettable circuit protective and over-temperature warning system device must be designed so that, when with a means for disconnecting the an overload or circuit fault exists, it battery from its charging source in the will open the circuit regardless of the event of an over-temperature condi- position of the operating control. tion; or (d) If the ability to reset a circuit (iii) A battery failure sensing and breaker or replace a fuse is essential to warning system with a means for dis- safety in flight, that circuit breaker or connecting the battery from its charg- fuse must be located and identified so

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that it can be readily reset or replaced laboratory or by ground tests on the in flight. rotorcraft, flight tests must be made. (e) Each essential load must have in- dividual circuit protection. However, LIGHTS individual protection for each circuit in an essential load system (such as § 29.1381 Instrument lights. each position light circuit in a system) The instrument lights must— is not required. (a) Make each instrument, switch, (f) If fuses are used, there must be and other device for which they are spare fuses for use in flight equal to at provided easily readable; and least 50 percent of the number of fuses (b) Be installed so that— of each rating required for complete (1) Their direct rays are shielded circuit protection. from the pilot’s eyes; and (g) Automatic reset circuit breakers (2) No objectionable reflections are may be used as integral protectors for visible to the pilot. electrical equipment provided there is circuit protection for the cable sup- § 29.1383 Landing lights. plying power to the equipment. (a) Each required landing or hovering [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as light must be approved. amended by Amdt. 29–24, 49 FR 44440, Nov. 6, (b) Each landing light must be in- 1984] stalled so that— (1) No objectionable glare is visible § 29.1359 Electrical system fire and to the pilot; smoke protection. (2) The pilot is not adversely affected (a) Components of the electrical sys- by halation; and tem must meet the applicable fire and (3) It provides enough light for night smoke protection provisions of §§ 29.831 operation, including hovering and land- and 29.863. ing. (b) Electrical cables, terminals, and (c) At least one separate switch must equipment, in designated fire zones, be provided, as applicable— and that are used in emergency proce- (1) For each separately installed dures, must be at least fire resistant. landing light; and (c) Insulation on electrical wire and (2) For each group of landing lights cable installed in the rotorcraft must installed at a common location. be self-extinguishing when tested in ac- cordance with Appendix F, Part I(a)(3), § 29.1385 Position light system installa- of part 25 of this chapter. tion. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (a) General. Each part of each posi- amended by Amdt. 29–42, 63 FR 43285, Aug. 12, tion light system must meet the appli- 1998] cable requirements of this section and each system as a whole must meet the § 29.1363 Electrical system tests. requirements of §§ 29.1387 through (a) When laboratory tests of the elec- 29.1397. trical system are conducted— (b) Forward position lights. Forward (1) The tests must be performed on a position lights must consist of a red mock-up using the same generating and a green light spaced laterally as equipment used in the rotorcraft; far apart as practicable and installed (2) The equipment must simulate the forward on the rotorcraft so that, with electrical characteristics of the dis- the rotorcraft in the normal flying po- tribution wiring and connected loads to sition, the red light is on the left side, the extent necessary for valid test re- and the green light is on the right side. sults; and Each light must be approved. (3) Laboratory generator drives must (c) Rear position light. The rear posi- simulate the prime movers on the tion light must be a white light mount- rotorcraft with respect to their reac- ed as far aft as practicable, and must tion to generator loading, including be approved. loading due to faults. (d) Circuit. The two forward position (b) For each flight condition that lights and the rear position light must cannot be simulated adequately in the make a single circuit.

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(e) Light covers and color filters. Each filters in place. Intensities must be de- light cover or color filter must be at termined with the light source oper- least flame resistant and may not ating at a steady value equal to the av- change color or shape or lose any ap- erage luminous output of the source at preciable light transmission during the normal operating voltage of the normal use. rotorcraft. The light distribution and intensity of each position light must § 29.1387 Position light system dihe- meet the requirements of paragraph (b) dral angles. 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- must be expressed in terms of min- dral angles described in this section. imum intensities in the horizontal (b) Dihedral angle L (left) is formed plane, minimum intensities in any by two intersecting vertical planes, the vertical plane, and maximum inten- first parallel to the longitudinal axis of sities in overlapping beams, within di- the rotorcraft, and the other at 110 de- hedral angles, L, R, and A, and must grees to the left of the first, as viewed meet the following requirements: when looking forward along the longi- (1) Intensities in the horizontal plane. tudinal axis. Each intensity in the horizontal plane (c) Dihedral angle R (right) is formed (the plane containing the longitudinal by two intersecting vertical planes, the axis of the rotorcraft and perpendicular first parallel to the longitudinal axis of to the plane of symmetry of the rotor- the rotorcraft, and the other at 110 de- craft), must equal or exceed the values grees to the right of the first, as viewed in § 29.1391. when looking forward along the longi- (2) Intensities in any vertical plane. tudinal axis. Each intensity in any vertical plane (d) Dihedral angle A (aft) is formed 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 § 29.1393 where I is vertical plane passing through the lon- the minimum intensity prescribed in gitudinal axis, as viewed when looking § 29.1391 for the corresponding angles in aft along the longitudinal axis. the horizontal plane. (e) If the rear position light, when (3) Intensities in overlaps between adja- mounted as far aft as practicable in ac- cent signals. No intensity in any over- cordance with § 29.1385(c), cannot show lap between adjacent signals may ex- unbroken light within dihedral angle A ceed the values in § 29.1395, except that (as defined in paragraph (d) of this sec- higher intensities in overlaps may be tion), a solid angle or angles of ob- used with the use of main beam inten- structed visibility totaling not more sities substantially greater than the than 0.04 steradians is allowable within minima specified in §§ 29.1391 and that dihedral angle, if such solid angle 29.1393 if the overlap intensities in rela- is within a cone whose apex is at the tion to the main beam intensities do rear position light and whose elements not adversely affect signal clarity. make an angle of 30° with a vertical line passing through the rear position § 29.1391 Minimum intensities in the horizontal plane of forward and light. rear position lights. (49 U.S.C. 1655(c)) Each position light intensity must [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as equal or exceed the applicable values in amended by Amdt. 29–9, 36 FR 21279, Nov. 5, the following table: 1971] Angle from right or left § 29.1389 Position light distribution Dihedral angle (light in- of longitudinal axis, Intensity cluded) measured from dead (candles) and intensities. ahead (a) General. The intensities prescribed L and R (forward red 0° to 10° ...... 40 in this section must be provided by new and green). 10° to 20° ...... 30 equipment with light covers and color 20° to 110° ...... 5

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Angle from right or left z is not greater than 0.002. Dihedral angle (light in- of longitudinal axis, Intensity cluded) measured from dead (candles) (b) Aviation green— ahead x is not greater than 0.440¥0.320y; A (rear white) ...... 110° to 180° ...... 20 x is not greater than y¥0.170; and y is not less than 0.390¥0.170x. § 29.1393 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 Each position light intensity must 0.540; y is not less than x¥0.040 or yc¥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¥0.400x; Angle above or below the horizontal plane Intensity, I Where Ye is the y coordinate of the Planck- 0° ...... 1.00 ian radiator for the value of x considered. ° ° 0 to 5 ...... 90 [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as 5° to 10° ...... 80 10° to 15° ...... 70 amended by Amdt. 29–7, 36 FR 12972, July 10, 15° to 20° ...... 50 1971] 20° to 30° ...... 30 30° to 40° ...... 10 § 29.1399 Riding light. 40° to 90° ...... 05 (a) Each riding light required for water operation must be installed so § 29.1395 Maximum intensities in over- lapping beams of forward and rear that it can— position lights. (1) Show a white light for at least two miles at night under clear atmos- No position light intensity may ex- pheric 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 § 29.1389(b)(3). water. Maximum intensity (b) Externally hung lights may be used. Overlaps Area A Area B (candles) (candles) § 29.1401 Anticollision light system. Green in dihedral angle L ...... 10 1 Red in dihedral angle R ...... 10 1 (a) General. If certification for night Green in dihedral angle A ...... 5 1 operation is requested, the rotorcraft Red in dihedral angle A ...... 5 1 must have an anticollision light sys- Rear white in dihedral angle L .. 5 1 Rear white in dihedral angle R 51tem that— (1) Consists of one or more approved Where— anticollision lights located so that (a) Area A includes all directions in their emitted light will not impair the the adjacent dihedral angle that pass crew’s vision or detract from the con- through the light source and intersect spicuity of the position lights; and the common boundary plane at more (2) Meets the requirements of para- than 10 degrees but less than 20 de- graphs (b) through (f) of this section. grees; and (b) Field of coverage. The system must (b) Area B includes all directions in consist of enough lights to illuminate the adjacent dihedral angle that pass the vital areas around the rotorcraft, through the light source and intersect considering the physical configuration the common boundary plane at more and flight characteristics of the rotor- than 20 degrees. craft. The field of coverage must ex- tend in each direction within at least § 29.1397 Color specifications. 30 degrees above and 30 degrees below the horizontal plane of the rotorcraft, Each position light color must have except that there may be solid angles the applicable International Commis- of obstructed visibility totaling not sion on Illumination chromaticity co- more than 0.5 steradians. ordinates as follows: (c) Flashing characteristics. The ar- (a) Aviation red— rangement of the system, that is, the y is not greater than 0.335; and number of light sources, beam width,

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speed of rotation, and other character- emergency, such as automatic liferaft istics, must give an effective flash fre- releases, must be readily accessible. quency of not less than 40, nor more (b) Stowage provisions. Stowage provi- than 100, cycles per minute. The effec- sions for required emergency equip- tive flash frequency is the frequency at ment must be furnished 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 inadvertent damage. than one light source. In overlaps, (c) Emergency exit descent device. The flash frequencies may exceed 100, but stowage provisions for the emergency not 180, cycles per minute. exit descent device required by (d) Color. Each anticollision light § 29.809(f) must be at the exits for which must be aviation red and must meet they are intended. the applicable requirements of § 29.1397. (d) Liferafts. Liferafts must be stowed (e) Light intensity. The minimum near exits through which the rafts can light intensities in any vertical plane, be launched during an unplanned ditch- measured with the red filter (if used) ing. Rafts automatically or remotely and expressed in terms of ‘‘effective’’ released outside the rotorcraft must be intensities must meet the require- attached to the rotorcraft by the static ments of paragraph (f) of this section. line prescribed in § 29.1415. The following relation must be as- (e) Long-range signaling device. The sumed: stowage provisions for the long-range signaling device required by § 29.1415 t2 must be near an exit available during ∫ Itdt() t an unplanned ditching. I = 1 (f) Life preservers. Each life preserver e +− 02.(tt21 ) must be within easy reach of each oc- where: cupant while seated.

Ie=effective intensity (candles). I(t)=instantaneous intensity as a function of § 29.1413 Safety belts: passenger warn- time. ing device. t2¥tl=flash time interval (seconds). (a) If there are means to indicate to Normally, the maximum value of effective the passengers when safety belts intensity is obtained when t2 and t1 are cho- should be fastened, they must be in- sen so that the effective intensity is equal to stalled to be operated from either pilot the instantaneous intensity at t2 and t1. seat. (f) Minimum effective intensities for (b) Each safety belt must be equipped anticollision light. Each anticollision with a metal to metal latching device. light effective intensity must equal or exceed the applicable values in the fol- (Secs. 313, 314, and 601 through 610 of the Fed- lowing table: eral Aviation Act of 1958 (49 U.S.C. 1354, 1355, and 1421 through 1430) and sec. 6(c), Dept. of Effective Transportation Act (49 U.S.C. 1655(c))) Angle above or below the horizontal plane intensity (candles) [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–16 43 FR 46233, Oct. 5, 0° to 5° ...... 150 1978] 5° to 10° ...... 90 10° to 20° ...... 30 20° to 30° ...... 15 § 29.1415 Ditching equipment. (a) Emergency flotation and sig- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as naling equipment required by any oper- amended by Amdt. 29–7, 36 FR 12972, July 10, ating rule of this chapter must meet 1971; Amdt. 29–11, 41 FR 5290, Feb. 5, 1976] the requirements of this section. (b) Each liferaft and each life pre- SAFETY EQUIPMENT server must be approved. In addition— (1) Provide not less than two rafts, of § 29.1411 General. an approximately equal rated capacity (a) Accessibility. Required safety and buoyancy to accommodate the oc- equipment to be used by the crew in an cupants of the rotorcraft; and

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(2) Each raft must have a trailing of ice on critical parts of the rotor- line, and must have a static line de- craft. Unless otherwise restricted, the signed to hold the raft near the rotor- means must be available for nighttime craft but to release it if the rotorcraft as well as daytime operation. The becomes totally submerged. rotorcraft flight manual must describe (c) Approved survival equipment the means of determining ice forma- must be attached to each liferaft. tion and must contain information nec- (d) There must be an approved sur- essary for safe operation of the rotor- vival type emergency locator trans- craft in icing conditions. mitter for use in one life raft. [Amdt. 29–21, 48 FR 4391, Jan. 31, 1983] [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–8, 36 FR 18722, Sept. 21, MISCELLANEOUS EQUIPMENT 1971; Amdt. 29–19, 45 FR 38348, June 9, 1980; Amdt. 27–26, 55 FR 8005, Mar. 6, 1990; Amdt. § 29.1431 Electronic equipment. 29–33, 59 FR 32057, June 21, 1994] (a) Radio communication and naviga- tion equipment installations must be § 29.1419 Ice protection. free from hazards in themselves, in (a) To obtain certification for flight their method of operation, and in their into icing conditions, compliance with effects on other components, under any this section must be shown. critical environmental conditions. (b) It must be demonstrated that the (b) Radio communication and naviga- rotorcraft can be safely operated in the tion equipment, controls, and wiring continuous maximum and intermittent must be installed so that operation of maximum icing conditions determined any one unit or system of units will under appendix C of this part within not adversely affect the simultaneous the rotorcraft altitude envelope. An operation of any other radio or elec- analysis must be performed to estab- tronic unit, or system of units, re- lish, on the basis of the rotorcraft’s quired by this chapter. operational needs, the adequacy of the ice protection system for the various § 29.1433 Vacuum systems. components of the rotorcraft. (a) There must be means, in addition (c) In addition to the analysis and to the normal pressure relief, to auto- physical evaluation prescribed in para- matically relieve the pressure in the graph (b) of this section, the effective- discharge lines from the vacuum air ness of the ice protection system and pump when the delivery temperature of its components must be shown by the air becomes unsafe. flight tests of the rotorcraft or its com- (b) Each vacuum air system line and ponents in measured natural atmos- fitting on the discharge side of the pheric icing conditions and by one or pump that might contain flammable more of the following tests as found vapors or fluids must meet the require- necessary to determine the adequacy of ments of § 29.1183 if they are in a des- the ice protection system: ignated fire zone. (1) Laboratory dry air or simulated (c) Other vacuum air system compo- icing tests, or a combination of both, of nents in designated fire zones must be the components or models of the com- at least fire resistant. ponents. (2) Flight dry air tests of the ice pro- § 29.1435 Hydraulic systems. tection system as a whole, or its indi- (a) Design. Each hydraulic system vidual components. must be designed as follows: (3) Flight tests of the rotorcraft or (1) Each element of the hydraulic its components in measured simulated system must be designed to withstand, icing conditions. without detrimental, permanent defor- (d) The ice protection provisions of mation, any structural loads that may this section are considered to be appli- be imposed simultaneously with the cable primarily to the airframe. Power- maximum operating hydraulic loads. plant installation requirements are (2) Each element of the hydraulic contained in Subpart E of this part. system must be designed to withstand (e) A means must be identified or pressures sufficiently greater than provided for determining the formation those prescribed in paragraph (b) of

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this section to show that the system (i) Masks covering the eyes, nose, and will not rupture under service condi- mouth; or tions. (ii) Masks covering the nose and (3) There must be means to indicate mouth, plus accessory equipment to the pressure in each main hydraulic protect the eyes; and power system. (3) That equipment must supply pro- (4) There must be means to ensure tective oxygen of 10 minutes duration that no pressure in any part of the sys- per crewmember at a pressure altitude tem will exceed a safe limit above the of 8,000 feet with a respiratory minute maximum operating pressure of the volume of 30 liters per minute BTPD. system, and to prevent excessive pres- sures resulting from any fluid volu- § 29.1457 Cockpit voice recorders. metric change in lines likely to remain (a) Each cockpit voice recorder re- closed long enough for such a change to quired by the operating rules of this take place. The possibility of detri- chapter must be approved, and must be mental transient (surge) pressures dur- installed so that it will record the fol- ing operation must be considered. lowing: (5) Each hydraulic line, fitting, and (1) Voice communications trans- component must be installed and sup- mitted from or received in the rotor- ported to prevent excessive vibration craft by radio. and to withstand inertia loads. Each (2) Voice communications of flight element of the installation must be crewmembers on the flight deck. protected from abrasion, corrosion, and (3) Voice communications of flight mechanical damage. crewmembers on the flight deck, using (6) Means for providing flexibility the rotorcraft’s interphone system. must be used to connect points, in a (4) Voice or audio signals identifying hydraulic fluid line, between which rel- navigation or approach aids introduced ative motion or differential vibration into a headset or speaker. exists. (5) Voice communications of flight (b) Tests. Each element of the system crewmembers using the passenger loud- must be tested to a proof pressure of 1.5 speaker system, if there is such a sys- times the maximum pressure to which tem, and if the fourth channel is avail- that element will be subjected in nor- able in accordance with the require- mal operation, without failure, mal- ments of paragraph (c)(4)(ii) of this sec- function, or detrimental deformation tion. of any part of the system. (b) The recording requirements of (c) Fire protection. Each hydraulic paragraph (a)(2) of this section may be system using flammable hydraulic met— fluid must meet the applicable require- (1) By installing a cockpit-mounted ments of §§ 29.861, 29.1183, 29.1185, and area microphone, located in the best 29.1189. position for recording voice commu- nications originating at the first and § 29.1439 Protective breathing equip- second pilot stations and voice commu- ment. nications of other crewmembers on the (a) If one or more cargo or baggage flight deck when directed to those sta- compartments are to be accessible in tions; or flight, protective breathing equipment (2) By installing a continually ener- must be available for an appropriate gized or voice-actuated lip microphone crewmember. at the first and second pilot stations. (b) For protective breathing equip- The microphone specified in this para- ment required by paragraph (a) of this graph must be so located and, if nec- section or by any operating rule of this essary, the preamplifiers and filters of chapter— the recorder must be so adjusted or (1) That equipment must be designed supplemented, that the recorded com- to protect the crew from smoke, carbon munications are intelligible when re- dioxide, and other harmful gases while corded under flight cockpit noise con- on flight deck duty; ditions and played back. The level of (2) That equipment must include— intelligibility must be approved by the

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Administrator. Repeated aural or vis- (f) If the cockpit voice recorder has a ual playback of the record may be used bulk erasure device, the installation in evaluating intelligibility. must be designed to minimize the prob- (c) Each cockpit voice recorder must ability of inadvertent operation and ac- be installed so that the part of the tuation of the device during crash im- communication or audio signals speci- pact. fied in paragraph (a) of this section ob- (g) Each recorder container must be tained from each of the following either bright orange or bright yellow. sources is recorded on a separate chan- [Amdt. 29–6, 35 FR 7293, May 9, 1970] nel: (1) For the first channel, from each § 29.1459 Flight recorders. microphone, headset, or speaker used (a) Each flight recorder required by at the first pilot station. the operating rules of Subchapter G of (2) For the second channel, from each this chapter must be installed so that: microphone, headset, or speaker used (1) It is supplied with airspeed, alti- at the second pilot station. tude, and directional data obtained (3) For the third channel, from the from sources that meet the accuracy cockpit-mounted area microphone, or requirements of §§ 29.1323, 29.1325, and the continually energized or voice-ac- 29.1327 of this part, as applicable; tuated lip microphones at the first and (2) The vertical acceleration sensor is second pilot stations. rigidly attached, and located longitu- (4) For the fourth channel, from— dinally within the approved center of (i) Each microphone, headset, or gravity limits of the rotorcraft; speaker used at the stations for the (3) It receives its electrical power third and fourth crewmembers; or from the bus that provides the max- (ii) If the stations specified in para- imum reliability for operation of the graph (c)(4)(i) of this section are not re- flight recorder without jeopardizing quired or if the signal at such a station service to essential or emergency is picked up by another channel, each loads; microphone on the flight deck that is (4) There is an aural or visual means used with the passenger loudspeaker for perflight checking of the recorder system if its signals are not picked up for proper recording of data in the stor- by another channel. age medium; and (iii) Each microphone on the flight (5) Except for recorders powered sole- deck that is used with the rotorcraft’s ly by the engine-drive electrical gener- loudspeaker system if its signals are ator system, there is an automatic not picked up by another channel. means to simultaneously stop a re- (d) Each cockpit voice recorder must corder that has a data erasure feature be installed so that— and prevent each erasure feature from (1) It receives its electric power from functioning, within 10 minutes after the bus that provides the maximum re- any crash impact. liability for operation of the cockpit (b) Each nonejectable recorder con- voice recorder without jeopardizing tainer must be located and mounted so service to essential or emergency as to minimize the probability of con- loads; tainer rupture resulting from crash im- (2) There is an automatic means to pact and subsequent damage to the simultaneously stop the recorder and record from fire. prevent each erasure feature from func- (c) A correlation must be established tioning, within 10 minutes after crash between the flight recorder readings of impact; and airspeed, altitude, and heading and the (3) There is an aural or visual means corresponding readings (taking into ac- for preflight checking of the recorder count correction factors) of the first pi- for proper operation. lot’s instruments. This correlation (e) The record container must be lo- must cover the airspeed range over cated and mounted to minimize the which the aircraft is to be operated, probability of rupture of the container the range of altitude to which the air- as a result of crash impact and con- craft is limited, and 360 degrees of sequent heat damage to the record heading. Correlation may be estab- from fire. lished on the ground as appropriate.

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(d) Each recorder container must: operation must be made available to (1) Be either bright orange or bright the crewmembers as prescribed in yellow; §§ 29.1541 through 29.1589. (2) Have a reflective tape affixed to its external surface to facilitate its lo- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), cation under water; and 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. (3) Have an underwater locating de- of Transportation Act (49 U.S.C. 1655(c))) vice, when required by the operating rules of this chapter, on or adjacent to [Amdt. 29–15, 43 FR 2327, Jan. 16, 1978] the container which is secured in such OPERATING LIMITATIONS a manner that it is not likely to be sep- arated during crash impact. § 29.1503 Airspeed limitations: general. [Amdt. 29–25, 53 FR 26145, July 11, 1988; 53 FR (a) An operating speed range must be 26144, July 11, 1988] established. (b) When airspeed limitations are a § 29.1461 Equipment containing high energy rotors. function of weight, weight distribution, altitude, rotor speed, power, or other (a) Equipment containing high en- factors, airspeed limitations cor- ergy rotors must meet paragraph (b), responding with the critical combina- (c), or (d) of this section. tions of these factors must be estab- (b) High energy rotors contained in lished. equipment must be able to withstand damage caused by malfunctions, vibra- § 29.1505 Never-exceed speed. tion, abnormal speeds, and abnormal (a) The never-exceed speed, V must temperatures. In addition— NE, be established so that it is— (1) Auxiliary rotor cases must be able (1) Not less than 40 knots (CAS); and to contain damage caused by the fail- ure of high energy rotor blades; and (2) Not more than the lesser of— (2) Equipment control devices, sys- (i) 0.9 times the maximum forward tems, and instrumentation must rea- speeds established under § 29.309; sonably ensure that no operating limi- (ii) 0.9 times the maximum speed tations affecting the integrity of high shown under §§ 29.251 and 29.629; or energy rotors will be exceeded in serv- (iii) 0.9 times the maximum speed ice. substantiated for advancing blade tip (c) It must be shown by test that mach number effects under critical al- equipment containing high energy ro- titude conditions. tors can contain any failure of a high (b) VNE may vary with altitude, energy rotor that occurs at the highest r.p.m., temperature, and weight, if— speed obtainable with the normal speed (1) No more than two of these vari- control devices inoperative. ables (or no more than two instru- (d) Equipment containing high en- ments integrating more than one of ergy rotors must be located where these variables) are used at one time; rotor failure will neither endanger the and occupants nor adversely affect contin- (2) The ranges of these variables (or ued safe flight. of the indications on instruments inte- grating more than one of these vari- [Amdt. 29–3, 33 FR 971, Jan. 26, 1968] ables) are large enough to allow an operationally practical and safe vari- Subpart G—Operating Limitations ation of VNE. and Information (c) For helicopters, a stabilized power-off VNE denoted as VNE (power- § 29.1501 General. off) may be established at a speed less (a) Each operating limitation speci- than VNE established pursuant to para- fied in §§ 29.1503 through 29.1525 and graph (a) of this section, if the fol- other limitations and information nec- lowing conditions are met: essary for safe operation must be es- (1) VNE (power-off) is not less than a tablished. speed midway between the power-on (b) The operating limitations and VNE and the speed used in meeting the other information necessary for safe requirements of—

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(i) § 29.67(a)(3) for Category A heli- to make a safe landing following power copters; failure, the range of heights and its (ii) § 29.65(a) for Category B heli- variation with forward speed must be copters, except multi-engine heli- established, together with any other copters meeting the requirements of pertinent information, such as the kind § 29.67(b); and of landing surface. (iii) § 29.67(b) for multi-engine Cat- [Amdt. 29–21, 48 FR 4391, Jan. 31, 1983] egory B helicopters meeting the re- quirements of § 29.67(b). § 29.1519 Weight and center of gravity. (2) V (power-off) is— NE The weight and center of gravity lim- (i) A constant airspeed; itations determined under §§ 29.25 and (ii) A constant amount less than 29.27, respectively, must be established power-on V or NE; as operating limitations. (iii) A constant airspeed for a portion of the altitude range for which certifi- § 29.1521 Powerplant limitations. cation is requested, and a constant (a) General. The powerplant limita- amount less than power-on VNE for the remainder of the altitude range. tions prescribed in this section must be established so that they do not exceed (Secs. 313(a), 601, 603, 604, and 605 of the Fed- the corresponding limits for which the eral Aviation Act of 1958 (49 U.S.C. 1354(a), engines are type certificated. 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. (b) Takeoff operation. The powerplant of Transportation Act (49 U.S.C. 1655(c))) takeoff operation must be limited by— [Amdt. 29–3, 33 FR 971, Jan. 26, 1968, as (1) The maximum rotational speed, amended by Amdt. 29–15, 43 FR 2327, Jan. 16, which may not be greater than— 1978; Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] (i) The maximum value determined by the rotor design; or § 29.1509 Rotor speed. (ii) The maximum value shown dur- (a) Maximum power-off (autorotation). ing the type tests; The maximum power-off rotor speed (2) The maximum allowable manifold must be established so that it does not pressure (for reciprocating engines); exceed 95 percent of the lesser of— (3) The maximum allowable turbine (1) The maximum design r.p.m. deter- inlet or turbine outlet gas temperature mined under § 29.309(b); and (for turbine engines); (2) The maximum r.p.m. shown dur- (4) The maximum allowable power or ing the type tests. torque for each engine, considering the (b) Minimum power-off. The minimum power input limitations of the trans- power-off rotor speed must be estab- mission with all engines operating; lished so that it is not less than 105 (5) The maximum allowable power or percent of the greater of— torque for each engine considering the (1) The minimum shown during the power input limitations of the trans- type tests; and mission with one engine inoperative; (2) The minimum determined by de- (6) The time limit for the use of the sign substantiation. power corresponding to the limitations (c) Minimum power-on. The minimum established in paragraphs (b)(1) power-on rotor speed must be estab- through (5) of this section; and lished so that it is— (7) If the time limit established in (1) Not less than the greater of— paragraph (b)(6) of this section exceeds (i) The minimum shown during the 2 minutes— type tests; and (i) The maximum allowable cylinder (ii) The minimum determined by de- head or coolant outlet temperature (for sign substantiation; and reciprocating engines); and (2) Not more than a value determined (ii) The maximum allowable engine under § 29.33 (a)(1) and (c)(1). and transmission oil temperatures. (c) Continuous operation. The contin- § 29.1517 Limiting height-speed enve- uous operation must be limited by— lope. (1) The maximum rotational speed, For Category A rotorcraft, if a range which may not be greater than— of heights exists at any speed, includ- (i) The maximum value determined ing zero, within which it is not possible by the rotor design; or

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(ii) The maximum value shown dur- (3) The maximum allowable torque; ing the type tests; and (2) The minimum rotational speed (4) The maximum allowable oil tem- shown under the rotor speed require- perature. ments in § 29.1509(c). (g) Thirty-minute OEI power operation. (3) The maximum allowable manifold Unless otherwise authorized, the use of pressure (for reciprocating engines); 30-minute OEI power must be limited (4) The maximum allowable turbine to multiengine, turbine-powered rotor- inlet or turbine outlet gas temperature craft for not longer than 30 minutes (for turbine engines); after failure of an engine. The use of 30- (5) The maximum allowable power or minute OEI power must also be limited torque for each engine, considering the by— power input limitations of the trans- (1) The maximum rotational speed, mission with all engines operating; which may not be greater than— (6) The maximum allowable power or (i) The maximum value determined torque for each engine, considering the by the rotor design; or power input limitations of the trans- (ii) The maximum value shown dur- mission with one engine inoperative; ing the type tests; and (2) The maximum allowable gas tem- (7) The maximum allowable tempera- perature; tures for— (3) The maximum allowable torque; and (i) The cylinder head or coolant out- (4) The maximum allowable oil tem- let (for reciprocating engines); perature. (ii) The engine oil; and (h) Continuous OEI power operation. (iii) The transmission oil. Unless otherwise authorized, the use of (d) Fuel grade or designation. The min- continuous OEI power must be limited imum fuel grade (for reciprocating en- to multiengine, turbine-powered rotor- gines) or fuel designation (for turbine craft for continued flight after failure engines) must be established so that it of an engine. The use of continuous is not less than that required for the OEI power must also be limited by— operation of the engines within the (1) The maximum rotational speed, limitations in paragraphs (b) and (c) of which may not be greater than— this section. (i) The maximum value determined (e) Ambient temperature. Ambient by the rotor design; or temperature limitations (including (ii) The maximum value shown dur- limitations for winterization installa- ing the type tests. tions if applicable) must be established (2) The maximum allowable gas tem- as the maximum ambient atmospheric perature; temperature at which compliance with (3) The maximum allowable torque; the cooling provisions of §§ 29.1041 and through 29.1049 is shown. (4) The maximum allowable oil tem- (f) Two and one-half minute OEI power perature. operation. Unless otherwise authorized, (i) Rated 30-second OEI power oper- the use of 21⁄2-minute OEI power must ation. Rated 30-second OEI power is be limited to engine failure operation permitted only on multiengine, tur- of multiengine, turbine-powered rotor- bine-powered rotorcraft, also certifi- craft for not longer than 21⁄2 minutes cated for the use of rated 2-minute OEI for any period in which that power is power, and can only be used for contin- used. The use of 21⁄2-minute OEI power ued operation of the remaining en- must also be limited by— gine(s) after a failure or precautionary (1) The maximum rotational speed, shutdown of an engine. It must be which may not be greater than— shown that following application of 30- (i) The maximum value determined second OEI power, any damage will be by the rotor design; or readily detectable by the applicable in- (ii) The maximum value shown dur- spections and other related procedures ing the type tests; furnished in accordance with Section (2) The maximum allowable gas tem- A29.4 of appendix A of this part and perature; Section A33.4 of appendix A of part 33.

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The use of 30-second OEI power must be power unit under the TSO including limited to not more than 30 seconds for the categories of operation must be any period in which that power is used, specified as operating limitations for and by— the rotorcraft. (1) The maximum rotational speed (Secs. 313(a), 601, 603, 604, Federal Aviation which may not be greater than— Act of 1958 (49 U.S.C. 1354(a), 1421, 1423), sec. (i) The maximum value determined 6(c), Dept. of Transportation Act (49 U.S.C. by the rotor design; or 1655(c))) (ii) The maximum value dem- onstrated during the type tests; [Amdt. 29–17, 43 FR 50602, Oct. 30, 1978] (2) The maximum allowable gas tem- § 29.1523 Minimum flight crew. perature; and (3) The maximum allowable torque. The minimum flight crew must be es- (j) Rated 2-minute OEI power oper- tablished so that it is sufficient for safe ation. Rated 2-minute OEI power is per- operation, considering— mitted only on multiengine, turbine- (a) The workload on individual crew- powered rotorcraft, also certificated members; for the use of rated 30-second OEI (b) The accessibility and ease of oper- power, and can only be used for contin- ation of necessary controls by the ap- ued operation of the remaining en- propriate crewmember; and gine(s) after a failure or precautionary (c) The kinds of operation authorized shutdown of an engine. It must be under § 29.1525. shown that following application of 2- minute OEI power, any damage will be § 29.1525 Kinds of operations. readily detectable by the applicable in- The kinds of operations (such as spections and other related procedures VFR, IFR, day, night, or icing) for furnished in accordance with Section which the rotorcraft is approved are es- A29.4 of appendix a of this part and tablished by demonstrated compliance Section A33.4 of appendix A of part 33. with the applicable certification re- The use of 2-minute OEI power must be quirements and by the installed equip- limited to not more than 2 minutes for ment. any period in which that power is used, [Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] and by— (1) The maximum rotational speed, § 29.1527 Maximum operating altitude. which may not be greater than— (i) The maximum value determined The maximum altitude up to which by the rotor design; or operation is allowed, as limited by (ii) The maximum value dem- flight, structural, powerplant, func- onstrated during the type tests; tional, or equipment characteristics, (2) The maximum allowable gas tem- must be established. perature; and (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (3) The maximum allowable torque. eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. (Secs. 313(a), 601, 603, 604, and 605 of the Fed- of Transportation Act (49 U.S.C. 1655(c))) eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. [Amdt. 29–15, 43 FR 2327, Jan. 16, 1978] of Transportation Act (49 U.S.C. 1655(c))) § 29.1529 Instructions for Continued [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–1, 30 FR 8778, July 13, Airworthiness. 1965; Amdt. 29–3, 33 FR 971, Jan. 26, 1968; The applicant must prepare Instruc- Amdt. 29–15, 43 FR 2327, Jan. 16, 1978; Amdt. tions for Continued Airworthiness in 29–26, 53 FR 34220, Sept. 2, 1988; Amdt. 29–34, accordance with appendix A to this 59 FR 47768, Sept. 16, 1994; Amdt. 29–41, 62 FR part that are acceptable to the Admin- 46173, Aug. 29, 1997] istrator. The instructions may be in- § 29.1522 Auxiliary power unit limita- complete at type certification if a pro- tions. gram exists to ensure their completion If an auxiliary power unit that meets prior to delivery of the first rotorcraft the requirements of TSO–C77 is in- or issuance of a standard certificate of stalled in the rotorcraft, the limita- airworthiness, whichever occurs later. tions established for that auxiliary [Amdt. 29–20, 45 FR 60178, Sept. 11, 1980]

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MARKINGS AND PLACARDS § 29.1547 Magnetic direction indicator. § 29.1541 General. (a) A placard meeting the require- ments of this section must be installed (a) The rotorcraft must contain— on or near the magnetic direction indi- (1) The markings and placards speci- cator. fied in §§ 29.1545 through 29.1565; and (b) The placard must show the cali- (2) Any additional information, in- bration of the instrument in level strument markings, and placards re- flight with the engines operating. quired for the safe operation of the (c) The placard must state whether rotorcraft if it has unusual design, op- the calibration was made with radio re- erating or handling characteristics. ceivers on or off. (b) Each marking and placard pre- (d) Each calibration reading must be scribed in paragraph (a) of this sec- in terms of magnetic heading in not tion— more than 45 degree increments. (1) Must be displayed in a con- spicuous place; and § 29.1549 Powerplant instruments. (2) May not be easily erased, disfig- For each required powerplant instru- ured, or obscured. ment, as appropriate to the type of in- struments— § 29.1543 Instrument markings: gen- (a) Each maximum and, if applicable, eral. minimum safe operating limit must be For each instrument— marked with a red radial or a red line; (a) When markings are on the cover (b) Each normal operating range glass of the instrument there must be must be marked with a green arc or means to maintain the correct align- green line, not extending beyond the ment of the glass cover with the face of maximum and minimum safe limits; the dial; and (c) Each takeoff and precautionary (b) Each arc and line must be wide range must be marked with a yellow enough, and located to be clearly visi- arc or yellow line; ble to the pilot. (d) Each engine or propeller range that is restricted because of excessive § 29.1545 Airspeed indicator. vibration stresses must be marked with (a) Each airspeed indicator must be red arcs or red lines; and marked as specified in paragraph (b) of (e) Each OEI limit or approved oper- this section, with the marks located at ating range must be marked to be the corresponding indicated airspeeds. clearly differentiated from the mark- (b) The following markings must be ings of paragraphs (a) through (d) of made: this section except that no marking is (1) A red radial line— normally required for the 30-second (i) For rotorcraft other than heli- OEI limit. copters, at VNE; and [Amdt. 29–12, 41 FR 55474, Dec. 20, 1976, as (ii) For helicopters, at a VNE (power- amended by Amdt. 29–26, 53 FR 34220, Sept. 2, on). 1988; Amdt. 29–34, 59 FR 47769, Sept. 16, 1994] (2) A red, cross-hatched radial line at VNE (power-off) for helicopters, if VNE § 29.1551 Oil quantity indicator. (power-off) is less than VNE (power-on). Each oil quantity indicator must be (3) For the caution range, a yellow marked with enough increments to in- arc. dicate readily and accurately the quan- (4) For the safe operating range, a tity of oil. green arc. § 29.1553 Fuel quantity indicator. (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), If the unusable fuel supply for any 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. tank exceeds one gallon, or five per- of Transportation Act (49 U.S.C. 1655(c))) cent of the tank capacity, whichever is [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as greater, a red arc must be marked on amended by Amdt. 29–15, 43 FR 2327, Jan. 16, its indicator extending from the cali- 1978; 43 FR 3900, Jan. 30, 1978; Amdt. 29–17, 43 brated zero reading to the lowest read- FR 50602, Oct. 30, 1978] ing obtainable in level flight.

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§ 29.1555 Control markings. (b) Seats. If the maximum allowable (a) Each cockpit control, other than weight to be carried in a seat is less primary flight controls or control than 170 pounds, a placard stating the whose function is obvious, must be lesser weight must be permanently at- plainly marked as to its function and tached to the seat structure. method of operation. (c) Fuel and oil filler openings. The fol- (b) For powerplant fuel controls— lowing apply: (1) Each fuel tank selector valve con- (1) Fuel filler openings must be trol must be marked to indicate the po- marked at or near the filler cover sition corresponding to each tank and with— to each existing cross feed position; (i) The word ‘‘fuel’’; (2) If safe operation requires the use (ii) For reciprocating engine powered of any tanks in a specific sequence, rotorcraft, the minimum fuel grade; that sequence must be marked on, or (iii) For turbine-engine-powered adjacent to, the selector for those rotorcraft, the permissible fuel des- tanks; and (3) Each valve control for any engine ignations, except that if impractical, of a multiengine rotorcraft must be this information may be included in marked to indicate the position cor- the rotorcraft flight manual, and the responding to each engine controlled. fuel filler may be marked with an ap- (c) Usable fuel capacity must be propriate reference to the flight man- marked as follows: ual; and (1) For fuel systems having no selec- (iv) For pressure fueling systems, the tor controls, the usable fuel capacity of maximum permissible fueling supply the system must be indicated at the pressure and the maximum permissible fuel quantity indicator. defueling pressure. (2) For fuel systems having selector (2) Oil filler openings must be controls, the usable fuel capacity marked at or near the filler cover with available at each selector control posi- the word ‘‘oil’’. tion must be indicated near the selec- (d) Emergency exit placards. Each tor control. placard and operating control for each (d) For accessory, auxiliary, and emergency exit must differ in color emergency controls— from the surrounding fuselage surface (1) Each essential visual position in- as prescribed in § 29.811(h)(2). A placard dicator, such as those showing rotor must be near each emergency exit con- pitch or landing gear position, must be trol and must clearly indicate the loca- marked so that each crewmember can tion of that exit and its method of op- determine at any time the position of eration. the unit to which it relates; and (2) Each emergency control must be [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as red and must be marked as to method amended by Amdt. 29–3, 33 FR 971, Jan. 26, of operation. 1968; Amdt. 29–12, 41 FR 55474, Dec. 20, 1976; (e) For rotorcraft incorporating re- Amdt. 29–26, 53 FR 34220, Sept. 2, 1988] tractable landing gear, the maximum landing gear operating speed must be § 29.1559 Limitations placard. displayed in clear view of the pilot. There must be a placard in clear view [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as of the pilot that specifies the kinds of amended by Amdt. 29–12, 41 FR 55474, Dec. 20, operations (VFR, IFR, day, night, or 1976; Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] icing) for which the rotorcraft is ap- proved. § 29.1557 Miscellaneous markings and placards. [Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] (a) Baggage and cargo compartments, § 29.1561 Safety equipment. and ballast location. Each baggage and cargo compartment, and each ballast (a) Each safety equipment control to location must have a placard stating be operated by the crew in emergency, any limitations on contents, including such as controls for automatic liferaft weight, that are necessary under the releases, must be plainly marked as to loading requirements. its method of operation.

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(b) Each location, such as a locker or be furnished. The significance of each compartment, that carries any fire ex- limitation and of the color coding must tinguishing, signaling, or other life be explained. saving equipment, must be so marked. (b) Powerplant limitations. The fol- (c) Stowage provisions for required lowing information must be furnished: emergency equipment must be con- (1) Limitations required by § 29.1521. spicuously marked to identify the con- (2) Explanation of the limitations, tents and facilitate removal of the when appropriate. equipment. (3) Information necessary for mark- (d) Each liferaft must have obviously ing the instruments required by marked operating instructions. §§ 29.1549 through 29.1553. (e) Approved survival equipment must be marked for identification and (c) Weight and loading distribution. method of operation. The weight and center of gravity limits required by §§ 29.25 and 29.27, respec- § 29.1565 Tail rotor. tively, must be furnished. If the vari- ety of possible loading conditions war- Each tail rotor must be marked so that its disc is conspicuous under nor- rants, instructions must be included to mal daylight ground conditions. allow ready observance of the limita- tions. [Amdt. 29–3, 33 FR 971, Jan. 26, 1968] (d) Flight crew. When a flight crew of more than one is required, the number ROTORCRAFT FLIGHT MANUAL and functions of the minimum flight § 29.1581 General. crew determined under § 29.1523 must be furnished. (a) Furnishing information. A Rotor- (e) Kinds of operation. Each kind of craft Flight Manual must be furnished operation for which the rotorcraft and with each rotorcraft, and it must con- its equipment installations are ap- tain the following: proved must be listed. (1) Information required by §§ 29.1583 through 29.1589. (f) Limiting heights. Enough informa- (2) Other information that is nec- tion must be furnished to allow compli- essary for safe operation because of de- ance with § 29.1517. sign, operating, or handling character- (g) Maximum allowable wind. For Cat- istics. egory A rotorcraft, the maximum al- (b) Approved information. Each part of lowable wind for safe operation near the manual listed in §§ 29.1583 through the ground must be furnished. 29.1589 that is appropriate to the rotor- (h) Altitude. The altitude established craft, must be furnished, verified, and under § 29.1527 and an explanation of approved, and must be segregated, the limiting factors must be furnished. indentified, and clearly distinguished (i) Ambient temperature. Maximum from each unapproved part of that and minimum ambient temperature manual. limitations must be furnished. (c) [Reserved] (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (d) Table of contents. Each Rotorcraft eral Aviation Act of 1958 (49 U.S.C. 1354(a), Flight Manual must include a table of 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. contents if the complexity of the man- of Transportation Act (49 U.S.C. 1655(c))) ual indicates a need for it. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (Secs. 313(a), 601, 603, 604, and 605 of the Fed- amended by Amdt. 29–3, 33 FR 971, Jan. 26, eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1968; Amdt. 29–15, 43 FR 2327, Jan. 16, 1978; 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. Amdt. 29–17, 43 FR 50602, Oct. 30, 1978; Amdt. of Transportation Act (49 U.S.C. 1655(c))) 29–24, 49 FR 44440, Nov. 6, 1984] [Amdt. 29–15, 43 FR 2327, Jan. 16, 1978] § 29.1585 Operating procedures. § 29.1583 Operating limitations. (a) The parts of the manual con- (a) Airspeed and rotor limitations. In- taining operating procedures must formation necessary for the marking of have information concerning any nor- airspeed and rotor limitations on or mal and emergency procedures, and near their respective indicators must other information necessary for safe

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operation, including the applicable pro- shown, the limits must be clearly indi- cedures, such as those involving min- cated. The following must be provided: imum speeds, to be followed if an en- (a) Category A. For each category A gine fails. rotorcraft, the Rotorcraft Flight Man- (b) For multiengine rotorcraft, infor- ual must contain a summary of the mation identifying each operating con- performance data, including data nec- dition in which the fuel system inde- essary for the application of any oper- pendence prescribed in § 29.953 is nec- ating rule of this chapter, together essary for safety must be furnished, to- with descriptions of the conditions, gether with instructions for placing such as airspeeds, under which this the fuel system in a configuration used data was determined, and must con- to show compliance with that section. tain— (c) For helicopters for which a VNE (1) The indicated airspeeds cor- (power-off) is established under responding with those determined for § 29.1505(c), information must be fur- takeoff, and the procedures to be fol- nished to explain the VNE (power-off) lowed if the critical engine fails during and the procedures for reducing air- takeoff; speed to not more than the VNE (power- (2) The airspeed calibrations; off) following failure of all engines. (3) The techniques, associated air- (d) For each rotorcraft showing com- speeds, and rates of descent for autoro- pliance with § 29.1353 (c)(6)(ii) or tative landings; (c)(6)(iii), the operating procedures for disconnecting the battery from its (4) The rejected takeoff distance de- charging source must be furnished. termined under § 29.62 and the takeoff distance determined under § 29.61; (e) If the unusable fuel supply in any tank exceeds 5 percent of the tank ca- (5) The landing data determined pacity, or 1 gallon, whichever is great- under § 29.81 and § 29.85; er, information must be furnished (6) The steady gradient of climb for which indicates that when the fuel each weight, altitude, and temperature quantity indicator reads ‘‘zero’’ in for which takeoff data are to be sched- level flight, any fuel remaining in the uled, along the takeoff path deter- fuel tank cannot be used safely in mined in the flight conditions required flight. in § 29.67(a)(1) and (a)(2): (f) Information on the total quantity (i) In the flight conditions required in of usable fuel for each fuel tank must § 29.67(a)(1) between the end of the be furnished. takeoff distance and the point at which (g) For Category B rotorcraft, the the rotorcraft is 200 feet above the airspeeds and corresponding rotor takeoff surface (or 200 feet above the speeds for minimum rate of descent lowest point of the takeoff profile for and best glide angle as prescribed in elevated heliports); § 29.71 must be provided. (ii) In the flight conditions required in § 29.67(a)(2) between the points at (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), which the rotorcraft is 200 and 1000 feet 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. above the takeoff surface (or 200 and of Transportation Act (49 U.S.C. 1655(c))) 1000 feet above the lowest point of the takeoff profile for elevated heliports); [Amdt. 29–2, 32 FR 6914, May 5, 1967, as amended by Amdt. 29–15, 43 FR 2328, Jan. 16, and 1978; Amdt. 29–17, 43 FR 50602, Oct. 30, 1978; (7) Out-of-ground effect hover per- Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] formance determined under § 29.49 and the maximum safe wind demonstrated § 29.1587 Performance information. under the ambient conditions for data Flight manual performance informa- presented. tion which exceeds any operating limi- (b) Category B. For each category B tation may be shown only to the extent rotorcraft, the Rotorcraft Flight Man- necessary for presentation clarity or to ual must contain— determine the effects of approved op- (1) The takeoff distance and the tional equipment or procedures. When climbout speed together with the perti- data beyond operating limits are nent information defining the flight

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path with respect to autorotative land- the Instructions for Continued Airworthiness ing if an engine fails, including the cal- for each engine and rotor (hereinafter des- culated effects of altitude and tempera- ignated ‘‘products’’), for each applicance re- quired by this chapter, and any required in- ture; formation relating to the interface of those (2) The steady rates of climb and hov- appliances and products with the rotorcraft. ering ceiling, together with the cor- If Instructions for Continued Airworthiness responding airspeeds and other perti- are not supplied by the manufacturer of an nent information, including the cal- appliance or product installed in the rotor- culated effects of altitude and tempera- craft, the Instructions for Continued Air- ture; worthiness for the rotorcraft must include (3) The landing distance, appropriate the information essential to the continued airworthiness of the rotorcraft. airspeed, and type of landing surface, (c) The applicant must submit to the FAA together with all pertinent information a program to show how changes to the In- that might affect this distance, includ- structions for Continued Airworthiness made ing the effects of weight, altitude, and by the applicant or by the manufacturers of temperature; products and appliances installed in the (4) The maximum safe wind for oper- rotorcraft will be distributed. ation near the ground; A29.2 FORMAT (5) The airspeed calibrations; (6) The height-speed envelope except (a) The Instructions for Continued Air- for rotorcraft incorporating this as an worthiness must be in the form of a manual operating limitation; or manuals as appropriate for the quantity (7) Glide distance as a function of al- of data to be provided. titude when autorotating at the speeds (b) The format of the manual or manuals must provide for a practical arrangement. and conditions for minimum rate of de- scent and best glide angle, as deter- A29.3 CONTENT mined in § 29.71; The contents of the manual or manuals (8) Out-of-ground effect hover per- must be prepared in the English language. formance determined under § 29.49 and The Instructions for Continued Airworthi- the maximum safe wind demonstrated ness must contain the following manuals or under the ambient conditions for data sections, as appropriate, and information: presented; and (a) Rotorcraft maintenance manual or section. (9) Any additional performance data (1) Introduction information that includes an explanation of the rotorcraft’s features and necessary for the application of any op- data to the extent necessary for mainte- erating rule in this chapter. nance or preventive maintenance. [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (2) A description of the rotorcraft and its amended by Amdt. 29–21, 48 FR 4392, Jan. 31, systems and installations including its en- 1983; Amdt. 29–24, 49 FR 44440, Nov. 6, 1984; gines, rotors, and appliances. Amdt. 29–39, 61 FR 21901, May 10, 1996; Amdt. (3) Basic control and operation information 29–40, 61 FR 21908, May 10, 1996; Amdt. 29–44, describing how the rotorcraft components 64 FR 45338, Aug. 19, 1999] and systems are controlled and how they op- erate, including any special procedures and § 29.1589 Loading information. limitations that apply. (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, weights determined under § 29.25 that location of access panels for inspection and can result in a center of gravity beyond servicing, locations of lubrication points, the any extreme prescribed in § 29.27, as- lubricants to be used, equipment required for servicing, tow instructions and limitations, suming any probable occupant weights. mooring, jacking, and leveling information. (b) Maintenance Instructions. (1) Scheduling APPENDIX A TO PART 29—INSTRUCTIONS information for each part of the rotorcraft FOR CONTINUED AIRWORTHINESS and its engines, auxiliary power units, ro- tors, accessories, instruments, and equip- A29.1 GENERAL ment that provides the recommended periods (a) This appendix specifies requirements at which they should be cleaned, inspected, for the preparation of Instructions for Con- adjusted, tested, and lubricated, and the de- tinued Airworthiness as required by § 29.1529. gree of inspection, the applicable wear toler- (b) The Instructions for Continued Air- ances, and work recommended at these peri- worthiness for each rotorcraft must include ods. However, the applicant may refer to an

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accessory, instrument, or equipment manu- APPENDIX B TO PART 29—AIRWORTHI- facturer as the source of this information if NESS CRITERIA FOR HELICOPTER IN- the applicant shows that the item has an ex- STRUMENT FLIGHT ceptionally high degree of complexity requir- ing specialized maintenance techniques, test I. General. A transport category helicopter equipment, or expertise. The recommended may not be type certificated for operation overhaul periods and necessary cross ref- under the instrument flight rules (IFR) of erences to the Airworthiness Limitations this chapter unless it meets the design and section of the manual must also be included. installation requirements contained in this In addition, the applicant must include an appendix. inspection program that includes the fre- II. Definitions. (a) VYI means instrument quency and extent of the inspections nec- climb speed, utilized instead of V for com- essary to provide for the continued air- Y pliance with the climb requirements for in- worthiness of the rotorcraft. strument flight. (2) Troubleshooting information describing probable malfunctions, how to recognize (b) VNEI means instrument flight never ex- those malfunctions, and the remedial action ceed speed, utilized instead of VNE for com- for those malfunctions. pliance with maximum limit speed require- (3) Information describing the order and ments for instrument flight. method of removing and replacing products (c) VMINI means instrument flight min- and parts with any necessary precautions to imum speed, utilized in complying with min- be taken. imum limit speed requirements for instru- (4) Other general procedural instructions ment flight. including procedures for system testing dur- III. Trim. It must be possible to trim the ing ground running, symmetry checks, cyclic, collective, and directional control weighing and determining the center of grav- forces to zero at all approved IFR airspeeds, ity, lifting and shoring, and storage limita- power settings, and configurations appro- tions. priate to the type. (c) Diagrams of structural access plates IV. Static longitudinal stability. (a) General. and information needed to gain access for in- The helicopter must possess positive static spections when access plates are not pro- longitudinal control force stability at crit- vided. ical combinations of weight and center of (d) Details for the application of special in- gravity at the conditions specified in para- spection techniques including radiographic graphs IV (b) through (f) of this appendix. and ultrasonic testing where such processes The stick force must vary with speed so that are specified. any substantial speed change results in a (e) Information needed to apply protective stick force clearly perceptible to the pilot. treatments to the structure after inspection. The airspeed must return to within 10 per- (f) All data relative to structural fasteners cent of the trim speed when the control force such as identification, discard recommenda- is slowly released for each trim condition tions, and torque values. specified in paragraphs IV (b) through (f) of (g) A list of special tools needed. this appendix. (b) Climb. Stability must be shown in climb A29.4 AIRWORTHINESS LIMITATIONS SECTION thoughout the speed range 20 knots either The Instructions for Continued Airworthi- side of trim with— ness must contain a section titled Airworthi- (1) The helicopter trimmed at VYI; ness Limitations that is segregated and (2) Landing gear retracted (if retractable); clearly distinguishable from the rest of the and document. This section must set forth each (3) Power required for limit climb rate (at mandatory replacement time, structural in- least 1,000 fpm) at VYI or maximum contin- spection interval, and related structural in- uous power, whichever is less. spection procedure approved under § 29.571. If (c) Cruise. Stability must be shown the Instructions for Continued Airworthiness throughout the speed range from 0.7 to 1.1 VH consist of multiple documents, the section or VNEI, whichever is lower, not to exceed ±20 required by this paragraph must be included knots from trim with— in the principal manual. This section must (1) The helicopter trimmed and power ad- contain a legible statement in a prominent justed for level flight at 0.9 VH or 0.9 VNEI, location that reads: ‘‘The Airworthiness whichever is lower; and Limitations section is FAA approved and (2) Landing gear retracted (if retractable). specifies maintenance required under §§ 43.16 (d) Slow cruise. Stability must be shown and 91.403 of the Federal Aviation Regula- throughout the speed range from 0.9 V to tions unless an alternative program has been MINI 1.3 VMINI or 20 knots above trim speed, which- FAA approved.’’ ever is greater, with— [Amdt. 29–20, 45 FR 60178, Sept 11, 1980, as (1) The helicopter trimmed and power ad- amended by Amdt. 29–27, 54 FR 34330, Aug. 18, justed for level flight at 1.1 VMINI; and 1989] (2) Landing gear retracted (if retractable).

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(e) Descent. Stability must be shown flight without undue pilot effort. Additional throughout the speed range 20 knots either unrelated probable failures affecting the con- side of trim with— trol system must be considered; and (1) The helicopter trimmed at 0.8 VH or 0.8 (2) The flight characteristics requirements VNEI (or 0.8 VLE for the landing gear extended in Subpart B of Part 29 must be met through- case), whichever is lower; out a practical flight envelope. (2) Power required for 1,000 fpm descent at (b) The SAS must be designed so that it trim speed; and cannot create a hazardous deviation in flight (3) Landing gear extended and retracted, if path or produce hazardous loads on the heli- applicable. copter during normal operation or in the (f) Approach. Stability must be shown event of malfunction or failure, assuming throughout the speed range from 0.7 times corrective action begins within an appro- the minimum recommended approach speed priate period of time. Where multiple sys- to 20 knots above the maximum rec- tems are installed, subsequent malfunction ommended approach speed with— conditions must be considered in sequence (1) The helicopter trimmed at the rec- unless their occurrence is shown to be im- ommended approach speed or speeds; probable. (2) Landing gear extended and retracted, if VIII. Equipment, systems, and installation. applicable; and The basic equipment and installation must (3) Power required to maintain a 3° glide comply with Subpart F of Part 29 through path and power required to maintain the Amendment 29–14, with the following excep- steepest approach gradient for which ap- tions and additions: proval is requested. (a) Flight and navigation instruments. (1) A V. Static lateral-directional stability. (a) magnetic gyro-stabilized direction indicator Static directional stability must be positive instead of the gyroscopic direction indicator throughout the approved ranges of airspeed, required by § 29.1303(h); and power, and vertical speed. In straight, steady (2) A standby attitude indicator which sideslips up to ±10° from trim, directional meets the requirements of §§ 29.1303(g)(1) control position must increase in approxi- through (7), instead of a rate-of-turn indi- mately constant proportion to angle of side- cator required by § 29.1303(g). If standby bat- slip. At greater angles up to the maximum teries are provided, they may be charged sideslip angle appropriate to the type, in- from the aircraft electrical system if ade- creased directional control position must quate isolation is incorporated. The system produce increased angle of sideslip. must be designed so that the standby bat- (b) During sideslips up to ±10° from trim teries may not be used for engine starting. throughout the approved ranges of airspeed, (b) Miscellaneous requirements. (1) Instru- power, and vertical speed there must be no ment systems and other systems essential negative dihedral stability perceptible to the for IFR flight that could be adversely af- pilot through lateral control motion or fected by icing must be provided with ade- force. Longitudinal cycle movement with quate ice protection whether or not the sideslip must not be excessive. rotorcraft is certificated for operation in VI. Dynamic stability. (a) Any oscillation icing conditions. having a period of less than 5 seconds must (2) There must be means in the generating damp to 1/2 amplitude in not more than one system to automatically de-energize and dis- cycle. connect from the main bus any power source (b) Any oscillation having a period of 5 sec- developing hazardous overvoltage. onds or more but less than 10 seconds must (3) Each required flight instrument using a damp to 1/2 amplitude in not more than two power supply (electric, vacuum, etc.) must cycles. have a visual means integral with the instru- (c) Any oscillation having a period of 10 ment to indicate the adequacy of the power seconds or more but less than 20 seconds being supplied. must be damped. (4) When multiple systems performing like (d) Any oscillation having a period of 20 functions are required, each system must be seconds or more may not achieve double am- grouped, routed, and spaced so that physical plitude in less than 20 seconds. separation between systems is provided to (e) Any aperiodic response may not achieve ensure that a single malfunction will not ad- double amplitude in less than 9 seconds. versely affect more than one system. VII. Stability augmentation system (SAS). (a) (5) For systems that operate the required If a SAS is used, the reliability of the SAS flight instruments at each pilot’s station— must be related to the effects of its failure. (i) Only the required flight instruments for The occurrence of any failure condition the first pilot may be connected to that op- which would prevent continued safe flight erating system; and landing must be extremely improbable. (ii) Additional instruments, systems, or For any failure condition of the SAS which equipment may not be connected to an oper- is not shown to be extremely improbable— ating system for a second pilot unless provi- (1) The helicopter must be safely control- sions are made to ensure the continued nor- lable and capable of prolonged instrument mal functioning of the required instruments

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in the event of any malfunction of the addi- weight, altitude, and temperature for which tional instruments, systems, or equipment approval is requested. which is not shown to be extremely improb- [Amdt. 29–21, 48 FR 4392, Jan. 31, 1983, as able; amended by Amdt. 29–31, 55 FR 38967, Sept. (iii) The equipment, systems, and installa- 21, 1990; 55 FR 41309, Oct. 10, 1990; Amdt. 29– tions must be designed so that one display of 40, 61 FR 21908, May 10, 1996] the information essential to the safety of flight which is provided by the instruments APPENDIX C TO PART 29—ICING will remain available to a pilot, without ad- CERTIFICATION ditional crew-member action, after any sin- gle failure or combination of failures that is (a) Continuous maximum icing. The max- not shown to be extremely improbable; and imum continuous intensity of atmospheric (iv) For single-pilot configurations, instru- icing conditions (continuous maximum ments which require a static source must be icing) is defined by the variables of the cloud provided with a means of selecting an alter- liquid water content, the mean effective di- ameter of the cloud droplets, the ambient air nate source and that source must be cali- temperature, and the interrelationship of brated. these three variables as shown in Figure 1 of (6) In determining compliance with the re- this appendix. The limiting icing envelope in quirements of § 29.1351(d)(2), the supply of terms of altitude and temperature is given in electrical power to all systems necessary for Figure 2 of this appendix. The interrelation- flight under IFR must be included in the ship of cloud liquid water content with drop evaluation. diameter and altitude is determined from (c) Thunderstorm lights. In addition to the Figures 1 and 2. The cloud liquid water con- instrument lights required by § 29.1381(a), tent for continuous maximum icing condi- thunderstorm lights which provide high in- tions of a horizontal extent, other than 17.4 tensity white flood lighting to the basic nautical miles, is determined by the value of flight instruments must be provided. The liquid water content of Figure 1, multiplied thunderstorm lights must be installed to by the appropriate factor from Figure 3 of meet the requirements of § 29.1381(b). this appendix. IX. Rotorcraft Flight Manual. A Rotorcraft (b) Intermittent maximum icing. The inter- Flight Manual or Rotorcraft Flight Manual mittent maximum intensity of atmospheric IFR Supplement must be provided and must icing conditions (intermittent maximum icing) is defined by the variables of the cloud contain— liquid water content, the mean effective di- (a) Limitations. The approved IFR flight en- ameter of the cloud droplets, the ambient air velope, the IFR flightcrew composition, the temperature, and the interrelationship of revised kinds of operation, and the steepest these three variables as shown in Figure 4 of IFR precision approach gradient for which this appendix. The limiting icing envelope in the helicopter is approved; terms of altitude and temperature is given in (b) Procedures. Required information for Figure 5 of this appendix. The interrelation- proper operation of IFR systems and the rec- ship of cloud liquid water content with drop ommended procedures in the event of sta- diameter and altitude is determined from bility augmentation or electrical system Figures 4 and 5. The cloud liquid water con- failures; and tent for intermittent maximum icing condi-

(c) Performance. If VYI differs from VY, tions of a horizontal extent, other than 2.6 nautical miles, is determined by the value of climb performance at VYI and with maximum continuous power throughout the ranges of cloud liquid water content of Figure 4 multi- plied by the appropriate factor in Figure 6 of this appendix.

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[Amdt. 29–21, 48 FR 4393, Jan. 31, 1983]

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APPENDIX D TO PART 29—CRITERIA FOR craft in the normal course of their duties DEMONSTRATION OF EMERGENCY may not be used as passengers. EVACUATION PROCEDURES UNDER (i) No passenger may be assigned a specific § 29.803 seat except as the Administrator may re- quire. Except as required by paragraph (1) of (a) The demonstration must be conducted this appendix, no employee of the applicant either during the dark of the night or during may be seated next to an emergency exit, ex- daylight with the dark of night simulated. If cept as allowed by the Administrator. the demonstration is conducted indoors dur- (j) Seat belts and shoulder harnesses (as re- ing daylight hours, it must be conducted in- quired) must be fastened. side a darkened hangar having doors and (k) Before the start of the demonstration, windows covered. In addition, the doors and approximately one-half of the total average windows of the rotorcraft must be covered if amount of carry-on baggage, blankets, pil- the hangar illumination exceeds that of a lows, and other similar articles must be dis- moonless night. Illumination on the floor or tributed at several locations in the aisles ground may be used, but it must be kept low and emergency exit access ways to create and shielded against shining into the minor obstructions. rotorcraft’s windows or doors. (l) No prior indication may be given to any (b) The rotorcraft must be in a normal at- crewmember or passenger of the particular titude with landing gear extended. exits to be used in the demonstration. (c) Safety equipment such as mats or in- (m) The applicant may not practice, re- verted liferafts may be placed on the floor or hearse, or describe the demonstration for the ground to protect participants. No other equipment that is not part of the rotorcraft’s participants nor may any participant have emergency evacuation equipment may be taken part in this type of demonstration used to aid the participants in reaching the within the preceding 6 months. ground. (n) A pretakeoff passenger briefing may be (d) Except as provided in paragraph (a) of given. The passengers may also be advised to this appendix, only the rotorcraft’s emer- follow directions of crewmembers, but not be gency lighting system may provide illumina- instructed on the procedures to be followed tion. in the demonstration. (e) All emergency equipment required for (o) If safety equipment, as allowed by para- the planned operation of the rotorcraft must graph (c) of this appendix, is provided, either be installed. all passenger and cockpit windows must be (f) Each external door and exit and each in- blacked out or all emergency exits must ternal door or curtain must be in the takeoff have safety equipment to prevent disclosure configuration. of the available emergency exits. (g) Each crewmember must be seated in (p) Not more than 50 percent of the emer- the normally assigned seat for takeoff and gency exits in the sides of the fuselage of a must remain in that seat until receiving the rotorcraft that meet all of the requirements signal for commencement of the demonstra- applicable to the required emergency exits tion. For compliance with this section, each for that rotorcraft may be used for dem- crewmember must be— onstration. Exits that are not to be used for (1) A member of a regularly scheduled line the demonstration must have the exit handle crew; or deactivated or must be indicated by red (2) A person having knowledge of the oper- lights, red tape, or other acceptable means ation of exits and emergency equipment. placed outside the exits to indicate fire or (h) A representative passenger load of per- other reasons why they are unusable. The sons in normal health must be used as fol- exits to be used must be representative of all lows: the emergency exits on the rotorcraft and (1) At least 25 percent must be over 50 must be designated by the applicant, subject years of age, with at least 40 percent of these to approval by the Administrator. If in- being females. stalled, at least one floor level exit (Type I; (2) The remaining, 75 percent or less, must § 29.807(a)(1)) must be used as required by be 50 years of age or younger, with at least § 29.807(c). 30 percent of these being females. (q) All evacuees must leave the rotorcraft (3) Three life-size dolls, not included as by a means provided as part of the part of the total passenger load, must be car- rotorcraft’s equipment. ried by passengers to simulate live infants 2 (r) Approved procedures must be fully uti- years old or younger, except for a total pas- lized during the demonstration. senger load of fewer than 44 but more than (s) The evacuation time period is com- 19, one doll must be carried. A doll is not re- pleted when the last occupant has evacuated quired for a 19 or fewer passenger load. the rotorcraft and is on the ground. (4) Crewmembers, mechanics, and training personnel who maintain or operate the rotor- [Amdt. 27–26, 55 FR 8005, Mar. 6, 1990]

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