14 CFR Part 25

Home , Erasure (heraldry), SAGE (game engine)

Pt. 25 14 CFR Ch. I (1–1–13 Edition)

TABLE I.—HIRF ENVIRONMENT I (5) From 400 MHz to 8 gigahertz (GHz), use radiated susceptibility tests at a minimum Field strength of 150 V/m peak with pulse modulation of 4 Frequency (volts/meter) percent duty cycle with a 1 kHz pulse repeti- Peak Average tion frequency. This signal must be switched on and off at a rate of 1 Hz with a duty cycle 10 kHz–2 MHz ...... 50 50 of 50 percent. 2 MHz–30 MHz ...... 100 100 (d) Equipment HIRF Test Level 2. Equipment 30 MHz–100 MHz ...... 50 50 HIRF test level 2 is HIRF environment II in 100 MHz–400 MHz ...... 100 100 table II of this appendix reduced by accept- 400 MHz–700 MHz ...... 700 50 able aircraft transfer function and attenu- 700 MHz–1 GHz ...... 700 100 ation curves. Testing must cover the fre- GHz–2 GHz ...... 2,000 200 2 GHz–6 GHz ...... 3,000 200 quency band of 10 kHz to 8 GHz. 6 GHz–8 GHz ...... 1,000 200 (e) Equipment HIRF Test Level 3. (1) From 10 8 GHz–12 GHz ...... 3,000 300 kHz to 400 MHz, use conducted susceptibility 12 GHz–18 GHz ...... 2,000 200 tests, starting at a minimum of 0.15 mA at 10 18 GHz–40 GHz ...... 600 200 kHz, increasing 20 dB per frequency decade In this table, the higher field strength applies at the fre- to a minimum of 7.5 mA at 500 kHz. quency band edges. (2) From 500 kHz to 40 MHz, use conducted susceptibility tests at a minimum of 7.5 mA. (b) HIRF environment II is specified in the (3) From 40 MHz to 400 MHz, use conducted following table: susceptibility tests, starting at a minimum of 7.5 mA at 40 MHz, decreasing 20 dB per fre- TABLE II.–HIRF ENVIRONMENT II quency decade to a minimum of 0.75 mA at 400 MHz. Field strength (volts/meter) (4) From 100 MHz to 8 GHz, use radiated Frequency susceptibility tests at a minimum of 5 V/m. Peak Average [Doc. No. FAA–2006–23657, 72 FR 44025, Aug. 6, 10 kHz–500 kHz ...... 20 20 2007] 500 kHz–2 MHz ...... 30 30 2 MHz–30 MHz ...... 100 100 30 MHz–100 MHz ...... 10 10 PART 25—AIRWORTHINESS STAND- 100 MHz–200 MHz ...... 30 10 ARDS: TRANSPORT CATEGORY 200 MHz–400 MHz ...... 10 10 400 MHz–1 GHz ...... 700 40 AIRPLANES 1 GHz–2 GHz ...... 1,300 160 2 GHz–4 GHz ...... 3,000 120 SPECIAL FEDERAL AVIATION REGULATION NO. 4 GHz–6 GHz ...... 3,000 160 6 GHz–8 GHz ...... 400 170 13 8 GHz–12 GHz ...... 1,230 230 SPECIAL FEDERAL AVIATION REGULATION NO. 12 GHz–18 GHz ...... 730 190 109 18 GHz–40 GHz ...... 600 150 In this table, the higher field strength applies at the fre- Subpart A—General quency band edges. Sec. (c) Equipment HIRF Test Level 1. (1) From 10 25.1 Applicability. kilohertz (kHz) to 400 megahertz (MHz), use 25.2 Special retroactive requirements. conducted susceptibility tests with contin- 25.3 Special provisions for ETOPS type de- uous wave (CW) and 1 kHz square wave mod- sign approvals. ulation with 90 percent depth or greater. The 25.5 Incorporations by reference. conducted susceptibility current must start at a minimum of 0.6 milliamperes (mA) at 10 Subpart B—Flight kHz, increasing 20 decibels (dB) per frequency decade to a minimum of 30 mA at 500 GENERAL kHz. 25.21 Proof of compliance. (2) From 500 kHz to 40 MHz, the conducted 25.23 Load distribution limits. susceptibility current must be at least 30 25.25 Weight limits. mA. 25.27 Center of gravity limits. (3) From 40 MHz to 400 MHz, use conducted 25.29 Empty weight and corresponding cen- susceptibility tests, starting at a minimum ter of gravity. of 30 mA at 40 MHz, decreasing 20 dB per fre- 25.31 Removable ballast. quency decade to a minimum of 3 mA at 400 25.33 Propeller speed and pitch limits. MHz. (4) From 100 MHz to 400 MHz, use radiated PERFORMANCE susceptibility tests at a minimum of 20 volts 25.101 General. per meter (V/m) peak with CW and 1 kHz 25.103 Stall speed. square wave modulation with 90 percent 25.105 Takeoff. depth or greater. 25.107 Takeoff speeds.

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25.109 Accelerate-stop distance. 25.349 Rolling conditions. 25.111 Takeoff path. 25.351 Yaw maneuver conditions. 25.113 Takeoff distance and takeoff run. 25.115 Takeoff flight path. SUPPLEMENTARY CONDITIONS 25.117 Climb: general. 25.361 Engine torque. 25.119 Landing climb: All-engines-operating. 25.363 Side load on engine and auxiliary 25.121 Climb: One-engine-inoperative. 25.123 En route flight paths. power unit mounts. 25.125 Landing. 25.365 Pressurized compartment loads. 25.367 Unsymmetrical loads due to engine CONTROLLABILITY AND MANEUVERABILITY failure. 25.371 Gyroscopic loads. 25.143 General. 25.373 Speed control devices. 25.145 Longitudinal control. 25.147 Directional and lateral control. CONTROL SURFACE AND SYSTEM LOADS 25.149 Minimum control speed. 25.391 Control surface loads: General. TRIM 25.393 Loads parallel to hinge line. 25.161 Trim. 25.395 Control system. 25.397 Control system loads. STABILITY 25.399 Dual control system. 25.171 General. 25.405 Secondary control system. 25.173 Static longitudinal stability. 25.407 Trim tab effects. 25.175 Demonstration of static longitudinal 25.409 Tabs. stability. 25.415 Ground gust conditions. 25.177 Static lateral-directional stability. 25.427 Unsymmetrical loads. 25.181 Dynamic stability. 25.445 Auxiliary aerodynamic surfaces. 25.457 Wing flaps. STALLS 25.459 Special devices. 25.201 Stall demonstration. GROUND LOADS 25.203 Stall characteristics. 25.207 Stall warning. 25.471 General. 25.473 Landing load conditions and assump- GROUND AND WATER HANDLING tions. CHARACTERISTICS 25.477 Landing gear arrangement. 25.231 Longitudinal stability and control. 25.479 Level landing conditions. 25.233 Directional stability and control. 25.481 Tail-down landing conditions. 25.235 Taxiing condition. 25.483 One-gear landing conditions. 25.237 Wind velocities. 25.485 Side load conditions. 25.239 Spray characteristics, control, and 25.487 Rebound landing condition. stability on water. 25.489 Ground handling conditions. 25.491 Taxi, takeoff and landing roll. MISCELLANEOUS FLIGHT REQUIREMENTS 25.493 Braked roll conditions. 25.251 Vibration and buffeting. 25.495 Turning. 25.253 High-speed characteristics. 25.497 Tail-wheel yawing. 25.255 Out-of-trim characteristics. 25.499 Nose-wheel yaw and steering. 25.503 Pivoting. Subpart C—Structure 25.507 Reversed braking. 25.509 Towing loads. GENERAL 25.511 Ground load: unsymmetrical loads on 25.301 Loads. multiple-wheel units. 25.303 Factor of safety. 25.519 Jacking and tie-down provisions. 25.305 Strength and deformation. 25.307 Proof of structure. WATER LOADS 25.521 General. FLIGHT LOADS 25.523 Design weights and center of gravity 25.321 General. positions. 25.525 Application of loads. FLIGHT MANEUVER AND GUST CONDITIONS 25.527 Hull and main float load factors. 25.331 Symmetric maneuvering conditions. 25.529 Hull and main float landing condi- 25.333 Flight maneuvering envelope. tions. 25.335 Design airspeeds. 25.531 Hull and main float takeoff condi- 25.337 Limit maneuvering load factors. tion. 25.341 Gust and turbulence loads. 25.533 Hull and main float bottom pressures. 25.343 Design fuel and oil loads. 25.535 Auxiliary float loads. 25.345 High lift devices. 25.537 Seawing loads.

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EMERGENCY LANDING CONDITIONS 25.753 Main float design. 25.755 Hulls. 25.561 General. 25.562 Emergency landing dynamic condi- PERSONNEL AND CARGO ACCOMMODATIONS tions. 25.563 Structural ditching provisions. 25.771 Pilot compartment. 25.772 Pilot compartment doors. FATIGUE EVALUATION 25.773 Pilot compartment view. 25.571 Damage—tolerance and fatigue eval- 25.775 Windshields and windows. uation of structure. 25.777 Cockpit controls. 25.779 Motion and effect of cockpit controls. LIGHTNING PROTECTION 25.781 Cockpit control knob shape. 25.581 Lightning protection. 25.783 Fuselage doors. 25.785 Seats, berths, safety belts, and har- Subpart D—Design and Construction nesses. 25.787 Stowage compartments. GENERAL 25.789 Retention of items of mass in pas- 25.601 General. senger and crew compartments and gal- 25.603 Materials. leys. 25.605 Fabrication methods. 25.791 Passenger information signs and plac- 25.607 Fasteners. ards. 25.609 Protection of structure. 25.793 Floor surfaces. 25.611 Accessibility provisions. 25.795 Security considerations. 25.613 Material strength properties and material design values. EMERGENCY PROVISIONS 25.619 Special factors. 25.801 Ditching. 25.621 Casting factors. 25.803 Emergency evacuation. 25.623 Bearing factors. 25.807 Emergency exits. 25.625 Fitting factors. 25.809 Emergency exit arrangement. 25.629 Aeroelastic stability requirements. 25.810 Emergency egress assist means and 25.631 Bird strike damage. escape routes. 25.811 Emergency exit marking. CONTROL SURFACES 25.812 Emergency lighting. 25.651 Proof of strength. 25.813 Emergency exit access. 25.655 Installation. 25.815 Width of aisle. 25.657 Hinges. 25.817 Maximum number of seats abreast. 25.819 Lower deck service compartments CONTROL SYSTEMS (including galleys). 25.671 General. 25.820 Lavatory doors. 25.672 Stability augmentation and automatic and power-operated systems. VENTILATION AND HEATING 25.675 Stops. 25.831 Ventilation. 25.677 Trim systems. 25.832 Cabin ozone concentration. 25.679 Control system gust locks. 25.833 Combustion heating systems. 25.681 Limit load static tests. 25.683 Operation tests. PRESSURIZATION 25.685 Control system details. 25.689 Cable systems. 25.841 Pressurized cabins. 25.693 Joints. 25.843 Tests for pressurized cabins. 25.697 Lift and drag devices, controls. 25.699 Lift and drag device indicator. FIRE PROTECTION 25.701 Flap and slat interconnection. 25.851 Fire extinguishers. 25.703 Takeoff warning system. 25.853 Compartment interiors. 25.854 Lavatory fire protection. LANDING GEAR 25.855 Cargo or baggage compartments. 25.721 General. 25.856 Thermal/Acoustic insulation mate- 25.723 Shock absorption tests. rials. 25.725–25.727 [Reserved] 25.857 Cargo compartment classification. 25.729 Retracting mechanism. 25.858 Cargo or baggage compartment 25.731 Wheels. smoke or fire detection systems. 25.733 Tires. 25.859 Combustion heater fire protection. 25.735 Brakes and braking systems. 25.863 Flammable fluid fire protection. 25.737 Skis. 25.865 Fire protection of flight controls, engine mounts, and other flight structure. FLOATS AND HULLS 25.867 Fire protection: other components. 25.751 Main float buoyancy. 25.869 Fire protection: systems.

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MISCELLANEOUS 25.1019 Oil strainer or filter. 25.1021 Oil system drains. 25.871 Leveling means. 25.875 Reinforcement near propellers. 25.1023 Oil radiators. 25.899 Electrical bonding and protection 25.1025 Oil valves. against static electricity. 25.1027 Propeller feathering system. COOLING Subpart E—Powerplant 25.1041 General. GENERAL 25.1043 Cooling tests. 25.901 Installation. 25.1045 Cooling test procedures. 25.903 Engines. 25.904 Automatic takeoff thrust control sys- INDUCTION SYSTEM tem (ATTCS). 25.1091 Air induction. 25.905 Propellers. 25.1093 Induction system icing protection. 25.907 Propeller vibration and fatigue. 25.1101 Carburetor air preheater design. 25.925 Propeller clearance. 25.1103 Induction system ducts and air duct 25.929 Propeller deicing. systems. 25.933 Reversing systems. 25.1105 Induction system screens. 25.934 Turbojet engine thrust reverser sys- 25.1107 Inter-coolers and after-coolers. tem tests. 25.937 Turbopropeller-drag limiting sys- EXHAUST SYSTEM tems. 25.1121 General. 25.939 Turbine engine operating characteristics. 25.1123 Exhaust piping. 25.941 Inlet, engine, and exhaust compat- 25.1125 Exhaust heat exchangers. ibility. 25.1127 Exhaust driven turbo-superchargers. 25.943 Negative acceleration. POWERPLANT CONTROLS AND ACCESSORIES 25.945 Thrust or power augmentation system. 25.1141 Powerplant controls: general. 25.1142 Auxiliary power unit controls. FUEL SYSTEM 25.1143 Engine controls. 25.951 General. 25.1145 Ignition switches. 25.952 Fuel system analysis and test. 25.1147 Mixture controls. 25.953 Fuel system independence. 25.1149 Propeller speed and pitch controls. 25.954 Fuel system lightning protection. 25.1153 Propeller feathering controls. 25.955 Fuel flow. 25.1155 Reverse thrust and propeller pitch 25.957 Flow between interconnected tanks. settings below the flight regime. 25.959 Unusable fuel supply. 25.1157 Carburetor air temperature controls. 25.961 Fuel system hot weather operation. 25.1159 Supercharger controls. 25.963 Fuel tanks: general. 25.1161 Fuel jettisoning system controls. 25.965 Fuel tank tests. 25.1163 Powerplant accessories. 25.967 Fuel tank installations. 25.1165 Engine ignition systems. 25.969 Fuel tank expansion space. 25.1167 Accessory gearboxes. 25.971 Fuel tank sump. 25.973 Fuel tank filler connection. POWERPLANT FIRE PROTECTION 25.975 Fuel tank vents and carburetor vapor 25.1181 Designated fire zones; regions in- vents. cluded. 25.977 Fuel tank outlet. 25.1182 Nacelle areas behind firewalls, and 25.979 Pressure fueling system. engine pod attaching structures con- 25.981 Fuel tank ignition prevention. taining flammable fluid lines. 25.1183 Flammable fluid-carrying compo- FUEL SYSTEM COMPONENTS nents. 25.991 Fuel pumps. 25.1185 Flammable fluids. 25.993 Fuel system lines and fittings. 25.1187 Drainage and ventilation of fire 25.994 Fuel system components. zones. 25.995 Fuel valves. 25.1189 Shutoff means. 25.997 Fuel strainer or filter. 25.1191 Firewalls. 25.999 Fuel system drains. 25.1192 Engine accessory section diaphragm. 25.1001 Fuel jettisoning system. 25.1193 Cowling and nacelle skin. 25.1195 Fire extinguishing systems. OIL SYSTEM 25.1197 Fire extinguishing agents. 25.1011 General. 25.1199 Extinguishing agent containers. 25.1013 Oil tanks. 25.1201 Fire extinguishing system materials. 25.1015 Oil tank tests. 25.1203 Fire detector system. 25.1017 Oil lines and fittings. 25.1207 Compliance.

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Subpart F—Equipment 25.1419 Ice protection. 25.1421 Megaphones. GENERAL 25.1423 Public address system.

25.1301 Function and installation. MISCELLANEOUS EQUIPMENT 25.1303 Flight and navigation instruments. 25.1305 Powerplant instruments. 25.1431 Electronic equipment. 25.1307 Miscellaneous equipment. 25.1433 Vacuum systems. 25.1309 Equipment, systems, and installa- 25.1435 Hydraulic systems. tions. 25.1438 Pressurization and pneumatic sys- 25.1310 Power source capacity and distribu- tems. tion. 25.1439 Protective breathing equipment. 25.1316 Electrical and electronic system 25.1441 Oxygen equipment and supply. lightning protection. 25.1443 Minimum mass flow of supplemental 25.1317 High-intensity Radiated Fields oxygen. (HIRF) Protection. 25.1445 Equipment standards for the oxygen distributing system. INSTRUMENTS: INSTALLATION 25.1447 Equipment standards for oxygen dispensing units. 25.1321 Arrangement and visibility. 25.1449 Means for determining use of oxy- 25.1322 Flightcrew alerting. gen. 25.1323 Airspeed indicating system. 25.1450 Chemical oxygen generators. 25.1325 Static pressure systems. 25.1453 Protection of oxygen equipment 25.1326 Pitot heat indication systems. from rupture. 25.1327 Magnetic direction indicator. 25.1455 Draining of fluids subject to freez- 25.1329 Flight guidance system. ing. 25.1331 Instruments using a power supply. 25.1457 Cockpit voice recorders. 25.1333 Instrument systems. 25.1459 Flight data recorders. 25.1337 Powerplant instruments. 25.1461 Equipment containing high energy ELECTRICAL SYSTEMS AND EQUIPMENT rotors. 25.1351 General. Subpart G—Operating Limitations and 25.1353 Electrical equipment and installa- Information tions. 25.1355 Distribution system. 25.1501 General. 25.1357 Circuit protective devices. 25.1360 Precautions against injury. OPERATING LIMITATIONS 25.1362 Electrical supplies for emergency 25.1503 Airspeed limitations: general. conditions. 25.1505 Maximum operating limit speed. 25.1363 Electrical system tests. 25.1507 Maneuvering speed. 25.1365 Electrical appliances, motors, and 25.1511 Flap extended speed. transformers. 25.1513 Minimum control speed. 25.1515 Landing gear speeds. LIGHTS 25.1516 Other speed limitations. 25.1381 Instrument lights. 25.1517 Rough air speed, VRA. 25.1383 Landing lights. 25.1519 Weight, center of gravity, and 25.1385 Position light system installation. weight distribution. 25.1387 Position light system dihedral an- 25.1521 Powerplant limitations. gles. 25.1522 Auxiliary power unit limitations. 25.1389 Position light distribution and in- 25.1523 Minimum flight crew. tensities. 25.1525 Kinds of operation. 25.1391 Minimum intensities in the hori- 25.1527 Ambient air temperature and oper- zontal plane of forward and rear position ating altitude. lights. 25.1529 Instructions for Continued Air- 25.1393 Minimum intensities in any vertical worthiness. plane of forward and rear position lights. 25.1531 Maneuvering flight load factors. 25.1395 Maximum intensities in overlapping 25.1533 Additional operating limitations. beams of forward and rear position 25.1535 ETOPS approval. lights. 25.1397 Color specifications. MARKINGS AND PLACARDS 25.1399 Riding light. 25.1541 General. 25.1401 Anticollision light system. 25.1543 Instrument markings: general. 25.1403 Wing icing detection lights. 25.1545 Airspeed limitation information. 25.1547 Magnetic direction indicator. SAFETY EQUIPMENT 25.1549 Powerplant and auxiliary power unit 25.1411 General. instruments. 25.1415 Ditching equipment. 25.1551 Oil quantity indication.

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25.1553 Fuel quantity indicator. APPENDIX N TO PART 25—FUEL TANK FLAM- 25.1555 Control markings. MABILITY EXPOSURE AND 25.1557 Miscellaneous markings and plac- RELIABILITYANALYSIS ards. AUTHORITY: 49 U.S.C. 106(g), 40113, 44701, 25.1561 Safety equipment. 44702 and 44704. 25.1563 Airspeed placard. SOURCE: Docket No. 5066, 29 FR 18291, Dec. AIRPLANE FLIGHT MANUAL 24, 1964, unless otherwise noted. 25.1581 General. 25.1583 Operating limitations. SPECIAL FEDERAL AVIATION REGULATION 25.1585 Operating procedures. NO. 13 25.1587 Performance information. 1. Applicability. Contrary provisions of the Subpart H—Electrical Wiring Civil Air Regulations regarding certification 1 Interconnection Systems (EWIS) notwithstanding, this regulation shall provide the basis for approval by the Adminis- 25.1701 Definition. trator of modifications of individual Douglas 25.1703 Function and installation: EWIS. DC–3 and Lockheed L–18 airplanes subse- 25.1705 Systems and functions: EWIS. quent to the effective date of this regulation. 25.1707 System separation: EWIS. 2. General modifications. Except as modified 25.1709 System safety: EWIS. in sections 3 and 4 of this regulation, an ap- 25.1711 Component identification: EWIS. plicant for approval of modifications to a 25.1713 Fire protection: EWIS. DC–3 or L–18 airplane which result in 25.1715 Electrical bonding and protection changes in design or in changes to approved against static electricity: EWIS. limitations shall show that the modifica- 25.1717 Circuit protective devices: EWIS. tions were accomplished in accordance with 25.1719 Accessibility provisions: EWIS. the rules of either Part 4a or Part 4b in ef- 25.1721 Protection of EWIS. fect on September 1, 1953, which are applica- 25.1723 Flammable fluid fire protection: ble to the modification being made: Provided, EWIS. That an applicant may elect to accomplish a 25.1725 Powerplants: EWIS. modification in accordance with the rules of 25.1727 Flammable fluid shutoff means: Part 4b in effect on the date of application EWIS. for the modification in lieu of Part 4a or 25.1729 Instructions for Continued Air- Part 4b as in effect on September 1, 1953: And worthiness: EWIS. provided further, That each specific modifica- 25.1731 Powerplant and APU fire detector tion must be accomplished in accordance system: EWIS. with all of the provisions contained in the 25.1733 Fire detector systems, general: elected rules relating to the particular modi- EWIS. fication. 3. Specific conditions for approval. An appli- Subpart I—Special Federal Aviation cant for any approval of the following spe- Regulations cific changes shall comply with section 2 of this regulation as modified by the applicable 25.1801 SFAR No. 111—Lavatory Oxygen provisions of this section. Systems. (a) Increase in take-off power limitation— APPENDIX A TO PART 25 1,200 to 1,350 horsepower. The engine take-off APPENDIX B TO PART 25 power limitation for the airplane may be in- APPENDIX C TO PART 25 creased to more than 1,200 horsepower but APPENDIX D TO PART 25 not to more than 1,350 horsepower per engine APPENDIX E TO PART 25 if the increase in power does not adversely APPENDIX F TO PART 25 affect the flight characteristics of the air- APPENDIX G TO PART 25—CONTINUOUS GUST plane. DESIGN CRITERIA (b) Increase in take-off power limitation to APPENDIX H TO PART 25—INSTRUCTIONS FOR more than 1,350 horsepower. The engine take- CONTINUED AIRWORTHINESS off power limitation for the airplane may be APPENDIX I TO PART 25—INSTALLATION OF AN increased to more than 1,350 horsepower per AUTOMATIC TAKEOFF THRUST CONTROL engine if compliance is shown with the flight SYSTEM (ATTCS) characteristics and ground handling require- APPENDIX J TO PART 25—EMERGENCY EVACU- ments of Part 4b. ATION (c) Installation of engines of not more than APPENDIX K TO PART 25—EXTENDED OPER- 1,830 cubic inches displacement and not having ATIONS (ETOPS) APPENDIX L TO PART 25—HIRF ENVIRON- 1 It is not intended to waive compliance MENTS AND EQUIPMENT HIRF TEST LEV- with such airworthiness requirements as are ELS included in the operating parts of the Civil APPENDIX M TO PART 25—FUEL TANK SYSTEM Air Regulations for specific types of oper- FLAMMABILITY REDUCTION MEANS ation.

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a certificated take-off rating of more than 1,350 (c) Airplane flight manual-performance oper- horsepower. Engines of not more than 1,830 ating information. An approved airplane flight cubic inches displacement and not having a manual shall be provided for each DC–3 and certificated take-off rating of more than L–18 airplane which has had new maximum 1,350 horsepower which necessitate a major certificated weights established under this modification of redesign of the engine instal- section. The airplane flight manual shall lation may be installed, if the engine fire contain the applicable performance informa- prevention and fire protection are equivalent tion prescribed in that part of the regula- to that on the prior engine installation. tions under which the new certificated (d) Installation of engines of more than 1,830 weights were established and such additional cubic inches displacement or having certificated information as may be necessary to enable take-off rating of more than 1,350 horsepower. the application of the take-off, en route, and Engines of more than 1,830 cubic inches dis- landing limitations prescribed for transport placement or having certificated take-off category airplanes in the operating parts of rating of more than 1,350 horsepower may be the Civil Air Regulations. installed if compliance is shown with the en- (d) Performance operating limitations. Each gine installation requirements of Part 4b: airplane for which new maximum certifi- Provided, That where literal compliance with cated weights are established in accordance the engine installation requirements of Part with paragraphs (a) or (b) of this section 4b is extremely difficult to accomplish and shall be considered a transport category air- would not contribute materially to the ob- plane for the purpose of complying with the jective sought, and the Administrator finds performance operating limitations applica- that the experience with the DC–3 or L–18 ble to the operations in which it is utilized. airplanes justifies it, he is authorized to ac- 5. Reference. Unless otherwise provided, all cept such measures of compliance as he finds references in this regulation to Part 4a and will effectively accomplish the basic objec- Part 4b are those parts of the Civil Air Regu- tive. lations in effect on September 1, 1953. 4. Establishment of new maximum certificated weights. An applicant for approval of new This regulation supersedes Special Civil maximum certificated weights shall apply Air Regulation SR–398 and shall remain ef- for an amendment of the airworthiness cer- fective until superseded or rescinded by the tificate of the airplane and shall show that Board. the weights sought have been established, [19 FR 5039, Aug. 11, 1954. Redesignated at 29 and the appropriate manual material ob- FR 19099, Dec. 30, 1964] tained, as provided in this section. NOTE: Transport category performance re- SPECIAL FEDERAL AVIATION REGULATION quirements result in the establishment of NO. 109 maximum certificated weights for various altitudes. 1. Applicability. Contrary provisions of 14 (a) Weights–25,200 to 26,900 for the DC–3 and CFR parts 21, 25, and 119 of this chapter not- 18,500 to 19,500 for the L–18. New maximum withstanding, an applicant is entitled to an certificated weights of more than 25,200 but amended type certificate or supplemental not more than 26,900 pounds for DC–3 and type certificate in the transport category, if more than 18,500 but not more than 19,500 the applicant complies with all applicable pounds for L–18 airplanes may be established provisions of this SFAR. in accordance with the transport category performance requirements of either Part 4a Operations or Part 4b, if the airplane at the new maximum weights can meet the structural re- 2. General. quirements of the elected part. (a) The passenger capacity may not exceed (b) Weights of more than 26,900 for the DC–3 60. If more than 60 passenger seats are in- and 19,500 for the L–18. New maximum certifi- stalled, then: cated weights of more than 26,900 pounds for (1) If the extra seats are not suitable for DC–3 and 19,500 pounds for L–18 airplanes occupancy during taxi, takeoff and landing, shall be established in accordance with the each extra seat must be clearly marked (e.g., structural performance, flight characteris- a placard on the top of an armrest, or a tics, and ground handling requirements of placard sewn into the top of the back cush- Part 4b: Provided, That where literal compli- ion) that the seat is not to be occupied dur- ance with the structural requirements of ing taxi, takeoff and landing. Part 4b is extremely difficult to accomplish (2) If the extra seats are suitable for occu- and would not contribute materially to the pancy during taxi, takeoff and landing (i.e., objective sought, and the Administrator meet all the strength and passenger injury finds that the experience with the DC–3 or L– criteria in part 25), then a note must be in- 18 airplanes justifies it, he is authorized to cluded in the Limitations Section of the Air- accept such measures of compliance as he plane Flight Manual that there are extra finds will effectively accomplish the basic seats installed but that the number of pas- objective. sengers on the airplane must not exceed 60.

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Additionally, there must be a placard in- in moderately rough air while occupying stalled adjacent to each door that can be aisles that are along the cabin sidewall, or used as a passenger boarding door that states where practicable, bordered by seats (seat that the maximum passenger capacity is 60. backs providing a 25-pound minimum break- The placard must be clearly legible to pas- away force are an acceptable means of com- sengers entering the airplane. pliance). (b) For airplanes outfitted with interior (b) Injury criteria for multiple occupancy doors under paragraph 10 of this SFAR, the side-facing seats. The following require- airplane flight manual (AFM) must include ments are only applicable to airplanes that an appropriate limitation that the airplane are subject to § 25.562. must be staffed with at least the following (1) Existing Criteria. All injury protection number of flight attendants who meet the re- criteria of § 25.562(c)(1) through (c)(6) apply to quirements of 14 CFR 91.533(b): (1) The number of flight attendants re- the occupants of side-facing seating. The quired by § 91.533(a)(1) and (2) of this chapter, Head Injury Criterion (HIC) assessments are and only required for head contact with the seat (2) At least one flight attendant if the air- and/or adjacent structures. plane model was originally certified for 75 (2) Body-to-Body Contact. Contact between passengers or more. the head, pelvis, torso or shoulder area of (c) The AFM must include appropriate lim- one Anthropomorphic Test Dummy (ATD) itation(s) to require a preflight passenger with the head, pelvis, torso or shoulder area briefing describing the appropriate functions of the ATD in the adjacent seat is not al- to be performed by the passengers and the lowed during the tests conducted in accord- relevant features of the airplane to ensure ance with § 25.562(b)(1) and (b)(2). Contact the safety of the passengers and crew. during rebound is allowed. (d) The airplane may not be offered for (3) Thoracic Trauma. If the torso of an ATD common carriage or operated for hire. The at the forward-most seat place impacts the operating limitations section of the AFM seat and/or adjacent structure during test- must be revised to prohibit any operations ing, compliance with the Thoracic Trauma involving the carriage of persons or property Index (TTI) injury criterion must be substan- for compensation or hire. The operators may tiated by dynamic test or by rational anal- receive remuneration to the extent con- ysis based on previous test(s) of a similar sistent with parts 125 and 91, subpart F, of seat installation. TTI data must be acquired this chapter. with a Side Impact Dummy (SID), as defined (e) A placard stating that ‘‘Operations in- by 49 CFR part 572, subpart F, or an equiva- volving the carriage of persons or property lent ATD or a more appropriate ATD and for compensation or hire are prohibited,’’ must be processed as defined in Federal must be located in the area of the Airworthi- Motor Vehicle Safety Standards (FMVSS) ness Certificate holder at the entrance to the part 571.214, section S6.13.5 (49 CFR 571.214). flightdeck. The TTI must be less than 85, as defined in (f) For passenger capacities of 45 to 60 pas- 49 CFR part 572, subpart F. Torso contact sengers, analysis must be submitted that during rebound is acceptable and need not be demonstrates that the airplane can be evacu- measured. ated in less than 90 seconds under the conditions specified in § 25.803 and appendix J to (4) Pelvis. If the pelvis of an ATD at any part 25. seat place impacts seat and/or adjacent (g) In order for any airplane certified under structure during testing, pelvic lateral accel- this SFAR to be placed in part 135 or part 121 eration injury criteria must be substantiated operations, the airplane must be brought by dynamic test or by rational analysis back into full compliance with the applica- based on previous test(s) of a similar seat in- ble operational part. stallation. Pelvic lateral acceleration may not exceed 130g. Pelvic acceleration data Equipment and Design must be processed as defined in FMVSS part 3. General. Unless otherwise noted, compli- 571.214, section S6.13.5 (49 CFR 571.214). ance is required with the applicable certifi- (5) Body-to-Wall/Furnishing Contact. If the cation basis for the airplane. Some provi- seat is installed aft of a structure—such as sions of this SFAR impose alternative re- an interior wall or furnishing that may con- quirements to certain airworthiness stand- tact the pelvis, upper arm, chest, or head of ards that do not apply to airplanes certifi- an occupant seated next to the structure— cated to earlier standards. Those airplanes the structure or a conservative representa- with an earlier certification basis are not re- tion of the structure and its stiffness must quired to comply with those alternative re- be included in the tests. It is recommended, quirements. but not required, that the contact surface of 4. Occupant Protection. the actual structure be covered with at least (a) Firm Handhold. In lieu of the require- two inches of energy absorbing protective ments of § 25.785(j), there must be means pro- padding (foam or equivalent) such as vided to enable persons to steady themselves Ensolite.

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(6) Shoulder Strap Loads. Where upper torso compartment, easily visible to all persons straps (shoulder straps) are used for sofa oc- entering the cabin in the immediate vicinity cupants, the tension loads in individual of each passenger entry door. straps may not exceed 1,750 pounds. If dual 7. Distance Between Exits. For an airplane straps are used for restraining the upper that is required to comply with § 25.807(f)(4), torso, the total strap tension loads may not in effect as of July 24, 1989, which has more exceed 2,000 pounds. than one passenger emergency exit on each (7) Occupant Retention. All side-facing seats side of the fuselage, no passenger emergency require end closures or other means to pre- exit may be more than 60 feet from any adja- vent the ATD’s pelvis from translating be- cent passenger emergency exit on the same yond the end of the seat at any time during side of the same deck of the fuselage, as testing. measured parallel to the airplane’s longitu- (8) Test Parameters. dinal axis between the nearest exit edges, (i) All seat positions need to be occupied by unless the following conditions are met: ATDs for the longitudinal tests. (a) Each passenger seat must be located (ii) A minimum of one longitudinal test, within 30 feet from the nearest exit on each conducted in accordance with the conditions side of the fuselage, as measured parallel to specified in § 25.562(b)(2), is required to assess the airplane’s longitudinal axis, between the the injury criteria as follows. Note that if a nearest exit edge and the front of the seat seat is installed aft of structure (such as an bottom cushion. interior wall or furnishing) that does not (b) The number of passenger seats located have a homogeneous surface, an additional between two adjacent pairs of emergency test or tests may be required to demonstrate exits (commonly referred to as a passenger that the injury criteria are met for the area zone) or between a pair of exits and a bulk- which an occupant could contact. For exam- head or a compartment door (commonly re- ple, different yaw angles could result in dif- ferred to as a ‘‘dead-end zone’’), may not ex- ferent injury considerations and may require ceed the following: separate tests to evaluate. (1) For zones between two pairs of exits, 50 (A) For configurations without structure percent of the combined rated capacity of (such as a wall or bulkhead) installed di- the two pairs of emergency exits. rectly forward of the forward seat place, Hy- (2) For zones between one pair of exits and brid II ATDs or equivalent must be in all a bulkhead, 40 percent of the rated capacity seat places. of the pair of emergency exits. (B) For configurations with structure (such (c) The total number of passenger seats in as a wall or bulkhead) installed directly for- the airplane may not exceed 33 percent of the ward of the forward seat place, a side impact maximum seating capacity for the airplane dummy or equivalent ATD or more appro- model using the exit ratings listed in priate ATD must be in the forward seat place § 25.807(g) for the original certified exits or and a Hybrid II ATD or equivalent must be the maximum allowable after modification in all other seat places. when exits are deactivated, whichever is less. (C) The test may be conducted with or (d) A distance of more than 60 feet between without deformed floor. adjacent passenger emergency exits on the (D) The test must be conducted with either same side of the same deck of the fuselage, no yaw or 10 degrees yaw for evaluating oc- as measured parallel to the airplane’s longi- cupant injury. Deviating from the no yaw tudinal axis between the nearest exit edges, condition may not result in the critical area is allowed only once on each side of the fuse- of contact not being evaluated. The upper lage. torso restraint straps, where installed, must 8. Emergency Exit Signs. In lieu of the re- remain on the occupant’s shoulder during quirements of § 25.811(d)(1) and (2) a single the impact condition of § 25.562(b)(2). sign at each exit may be installed provided: (c) For the vertical test, conducted in ac- (a) The sign can be read from the aisle cordance with the conditions specified in while directly facing the exit, and § 25.562(b)(1), Hybrid II ATDs or equivalent (b) The sign can be read from the aisle ad- must be used in all seat positions. jacent to the passenger seat that is farthest 5. Direct View. In lieu of the requirements from the exit and that does not have an in- of § 25.785(h)(2), to the extent practical with- tervening bulkhead/divider or exit. out compromising proximity to a required 9. Emergency Lighting. floor level emergency exit, the majority of (a) Exit Signs. In lieu of the requirements of installed flight attendant seats must be lo- § 25.812(b)(1), for airplanes that have a pas- cated to face the cabin area for which the senger seating configuration, excluding pilot flight attendant is responsible. seats, of 19 seats or less, the emergency exit 6. Passenger Information Signs. Compliance signs required by § 25.811(d)(1), (2), and (3) with § 25.791 is required except that for must have red letters at least 1-inch high on § 25.791(a), when smoking is to be prohibited, a white background at least 2 inches high. notification to the passengers may be pro- These signs may be internally electrically il- vided by a single placard so stating, to be luminated, or self illuminated by other than conspicuously located inside the passenger electrical means, with an initial brightness

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of at least 160 microlamberts. The color may 11. Width of Aisle. Compliance is required be reversed in the case of a sign that is self- with § 25.815, except that aisle width may be illuminated by other than electrical means. reduced to 0 inches between passenger seats (b) Floor Proximity Escape Path Marking. In during in-flight operations only, provided lieu of the requirements of § 25.812(e)(1), for that the applicant demonstrates that all cabin seating compartments that do not areas of the cabin are easily accessible by a have the main cabin aisle entering and crew member in the event of an emergency exiting the compartment, the following are (e.g., in-flight fire, decompression). Addition- applicable: ally, instructions must be provided at each (1) After a passenger leaves any passenger passenger seat for restoring the aisle width seat in the compartment, he/she must be required by § 25.815. Procedures must be es- able to exit the compartment to the main tablished and documented in the AFM to en- cabin aisle using only markings and visual sure that the required aisle widths are pro- features not more that 4 feet above the cabin vided during taxi, takeoff, and landing. floor, and 12. Materials for Compartment Interiors. (2) Proceed to the exits using the marking Compliance is required with the applicable system necessary to accomplish the actions provisions of § 25.853, except that compliance in § 25.812(e)(1) and (e)(2). with appendix F, parts IV and V, to part 25, (c) Transverse Separation of the Fuselage. In need not be demonstrated if it can be shown the event of a transverse separation of the by test or a combination of test and analysis fuselage, compliance must be shown with that the maximum time for evacuation of all § 25.812(l) except as follows: occupants does not exceed 45 seconds under (1) For each airplane type originally type the conditions specified in appendix J to part certificated with a maximum passenger seat- 25. ing capacity of 9 or less, not more than 50 13. Fire Detection. For airplanes with a type percent of all electrically illuminated emer- certificated passenger capacity of 20 or more, gency lights required by § 25.812 may be ren- there must be means that meet the require- dered inoperative in addition to the lights ments of § 25.858(a) through (d) to signal the that are directly damaged by the separation. flightcrew in the event of a fire in any iso- (2) For each airplane type originally type lated room not occupiable for taxi, takeoff certificated with a maximum passenger seat- and landing, which can be closed off from the ing capacity of 10 to 19, not more than 33 per- rest of the cabin by a door. The indication cent of all electrically illuminated emer- must identify the compartment where the gency lights required by § 25.812 may be ren- fire is located. This does not apply to lava- dered inoperative in addition to the lights tories, which continue to be governed by that are directly damaged by the separation. § 25.854. 10. Interior doors. In lieu of the requirements of § 25.813(e), interior doors may be in- 14. Cooktops. Each cooktop must be de- stalled between passenger seats and exits, signed and installed to minimize any poten- provided the following requirements are met. tial threat to the airplane, passengers, and (a) Each door between any passenger seat, crew. Compliance with this requirement occupiable for taxi, takeoff, and landing, and must be found in accordance with the fol- any emergency exit must have a means to lowing criteria: signal to the flightcrew, at the flightdeck, (a) Means, such as conspicuous burner-on that the door is in the open position for taxi, indicators, physical barriers, or handholds, takeoff and landing. must be installed to minimize the potential (b) Appropriate procedures/limitations for inadvertent personnel contact with hot must be established to ensure that any such surfaces of both the cooktop and cookware. door is in the open configuration for takeoff Conditions of turbulence must be considered. and landing. (b) Sufficient design means must be in- (c) Each door between any passenger seat cluded to restrain cookware while in place and any exit must have dual means to retain on the cooktop, as well as representative it in the open position, each of which is capa- contents, e.g., soup, sauces, etc., from the ef- ble of reacting the inertia loads specified in fects of flight loads and turbulence. Re- § 25.561. straints must be provided to preclude haz- (d) Doors installed across a longitudinal ardous movement of cookware and contents. aisle must translate laterally to open and These restraints must accommodate any close, e.g., pocket doors. cookware that is identified for use with the (e) Each door between any passenger seat cooktop. Restraints must be designed to be and any exit must be frangible in either di- easily utilized and effective in service. The rection. cookware restraint system should also be de- (f) Each door between any passenger seat signed so that it will not be easily disabled, and any exit must be operable from either thus rendering it unusable. Placarding must side, and if a locking mechanism is installed, be installed which prohibits the use of it must be capable of being unlocked from ei- cookware that cannot be accommodated by ther side without the use of special tools. the restraint system.

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(c) Placarding must be installed which pro- structural strength due to airplane corro- hibits the use of cooktops (i.e., power on any sion. burner) during taxi, takeoff, and landing. (h) Cooktop installations must provide (d) Means must be provided to address the adequate space for the user to immediately possibility of a fire occurring on or in the escape a hazardous cooktop condition. immediate vicinity of the cooktop. Two ac- (i) A means to shut off power to the ceptable means of complying with this re- cooktop must be provided at the galley con- quirement are as follows: taining the cooktop and in the cockpit. If ad- (1) Placarding must be installed that pro- ditional switches are introduced in the cock- hibits any burner from being powered when pit, revisions to smoke or fire emergency the cooktop is unattended. (NOTE: This procedures of the AFM will be required. would prohibit a single person from cooking (j) If the cooktop is required to have a lid on the cooktop and intermittently serving to enclose the cooktop there must be a food to passengers while any burner is pow- means to automatically shut off power to ered.) A fire detector must be installed in the the cooktop when the lid is closed. vicinity of the cooktop which provides an au- 15. Hand-Held Fire Extinguishers. dible warning in the passenger cabin, and a (a) For airplanes that were originally type fire extinguisher of appropriate size and ex- certificated with more than 60 passengers, tinguishing agent must be installed in the the number of hand-held fire extinguishers immediate vicinity of the cooktop. Access to must be the greater of— the extinguisher may not be blocked by a (1) That provided in accordance with the fire on or around the cooktop. requirements of § 25.851, or (2) An automatic, thermally activated fire (2) A number equal to the number of origi- suppression system must be installed to ex- nally type certificated exit pairs, regardless tinguish a fire at the cooktop and imme- of whether the exits are deactivated for the diately adjacent surfaces. The agent used in proposed configuration. the system must be an approved total flood- (b) Extinguishers must be evenly distrib- ing agent suitable for use in an occupied uted throughout the cabin. These extin- area. The fire suppression system must have guishers are in addition to those required by a manual override. The automatic activation paragraph 14 of this SFAR, unless it can be of the fire suppression system must also shown that the cooktop was installed in the automatically shut off power to the cooktop. immediate vicinity of the original exits. (e) The surfaces of the galley surrounding 16. Security. The requirements of § 25.795 are the cooktop which would be exposed to a fire not applicable to airplanes approved in ac- on the cooktop surface or in cookware on the cordance with this SFAR. cooktop must be constructed of materials [Doc. No. FAA–2007–28250, 74 FR 21541, May 8, that comply with the flammability require- 2009] ments of part III of appendix F to part 25. This requirement is in addition to the flammability requirements typically required of Subpart A—General the materials in these galley surfaces. Dur- ing the selection of these materials, consid- § 25.1 Applicability. eration must also be given to ensure that the flammability characteristics of the mate- (a) This part prescribes airworthiness rials will not be adversely affected by the use standards for the issue of type certifi- of cleaning agents and utensils used to re- cates, and changes to those certifi- move cooking stains. cates, for transport category airplanes. (f) The cooktop must be ventilated with a (b) Each person who applies under system independent of the airplane cabin and Part 21 for such a certificate or change cargo ventilation system. Procedures and must show compliance with the appli- time intervals must be established to inspect cable requirements in this part. and clean or replace the ventilation system to prevent a fire hazard from the accumula- § 25.2 Special retroactive require- tion of flammable oils and be included in the ments. instructions for continued airworthiness. The ventilation system ducting must be pro- The following special retroactive re- tected by a flame arrestor. [NOTE: The appli- quirements are applicable to an air- cant may find additional useful information plane for which the regulations ref- in Society of Automotive Engineers, Aero- erenced in the type certificate predate space Recommended Practice 85, Rev. E, en- the sections specified below— titled ‘‘Air Conditioning Systems for Sub- sonic Airplanes,’’ dated August 1, 1991.] (a) Irrespective of the date of applica- (g) Means must be provided to contain tion, each applicant for a supplemental spilled foods or fluids in a manner that will type certificate (or an amendment to a prevent the creation of a slipping hazard to type certificate) involving an increase occupants and will not lead to the loss of in passenger seating capacity to a total

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greater than that for which the air- on or after February 17, 2015, except plane has been type certificated must that, for an airplane configured for a show that the airplane concerned three person flight crew, the applicant meets the requirements of: need not comply with Appendix K, (1) Sections 25.721(d), 25.783(g), K25.1.4(a)(3), of this part, low fuel alert- 25.785(c), 25.803(c)(2) through (9), 25.803 ing. (d) and (e), 25.807 (a), (c), and (d), 25.809 (f) and (h), 25.811, 25.812, 25.813 (a), (b), [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, and (c), 25.815, 25.817, 25.853 (a) and (b), 2007] 25.855(a), 25.993(f), and 25.1359(c) in effect on October 24, 1967, and § 25.5 Incorporations by reference. (2) Sections 25.803(b) and 25.803(c)(1) (a) The materials listed in this sec- in effect on April 23, 1969. tion are incorporated by reference in (b) Irrespective of the date of applica- the corresponding sections noted. tion, each applicant for a supplemental These incorporations by reference were type certificate (or an amendment to a approved by the Director of the Federal type certificate) for an airplane manu- Register in accordance with 5 U.S.C. factured after October 16, 1987, must 552(a) and 1 CFR part 51. These mate- show that the airplane meets the re- rials are incorporated as they exist on quirements of § 25.807(c)(7) in effect on the date of the approval, and notice of July 24, 1989. any change in these materials will be (c) Compliance with subsequent revi- published in the FEDERAL REGISTER. sions to the sections specified in para- The materials are available for pur- graph (a) or (b) of this section may be elected or may be required in accord- chase at the corresponding addresses ance with § 21.101(a) of this chapter. noted below, and all are available for inspection at the National Archives [Amdt. 25–72, 55 FR 29773, July 20, 1990, as and Records Administration (NARA), amended by Amdt. 25–99, 65 FR 36266, June 7, and at FAA, Transport Airplane Direc- 2000] torate, Aircraft Certification Service, § 25.3 Special provisions for ETOPS 1601 Lind Avenue, SW., Renton, Wash- type design approvals. ington 98057–3356. For information on (a) Applicability. This section applies the availability of this material at to an applicant for ETOPS type design NARA, call 202–741–6030, or go to: http:// approval of an airplane: www.archives.gov/federallregister/ (1) That has an existing type certifi- codeloflfederallregulations/ cate on February 15, 2007; or ibrllocations.html. (2) For which an application for an (b) The following materials are avail- original type certificate was submitted able for purchase from the following before February 15, 2007. address: The National Technical Infor- (b) Airplanes with two engines. (1) For mation Services (NTIS), Springfield, ETOPS type design approval of an air- Virginia 22166. plane up to and including 180 minutes, (1) Fuel Tank Flammability Assess- an applicant must comply with ment Method User’s Manual, dated § 25.1535, except that it need not comply May 2008, document number DOT/FAA/ with the following provisions of Appen- AR–05/8, IBR approved for § 25.981 and dix K, K25.1.4, of this part: Appendix N. It can also be obtained at (i) K25.1.4(a), fuel system pressure and flow requirements; the following Web site: http:// (ii) K25.1.4(a)(3), low fuel alerting; www.fire.tc.faa.gov/systems/fueltank/ and FTFAM.stm. (iii) K25.1.4(c), engine oil tank design. (2) [Reserved] (2) For ETOPS type design approval [73 FR 42494, July 21, 2008] of an airplane beyond 180 minutes an applicant must comply with § 25.1535. (c) Airplanes with more than two engines. An applicant for ETOPS type design approval must comply with § 25.1535 for an airplane manufactured

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Subpart B—Flight pliance must be shown using the ice accretions defined in appendix C, assum- GENERAL ing normal operation of the airplane and its ice protection system in accord- § 25.21 Proof of compliance. ance with the operating limitations (a) Each requirement of this subpart and operating procedures established must be met at each appropriate com- by the applicant and provided in the bination of weight and center of grav- Airplane Flight Manual. ity within the range of loading condi- (2) No changes in the load distributions for which certification is re- tion limits of § 25.23, the weight limits quested. This must be shown— of § 25.25 (except where limited by per- (1) By tests upon an airplane of the formance requirements of this sub- type for which certification is re- part), and the center of gravity limits quested, or by calculations based on, of § 25.27, from those for non-icing con- and equal in accuracy to, the results of ditions, are allowed for flight in icing testing; and conditions or with ice accretion. (2) By systematic investigation of [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as each probable combination of weight amended by Amdt. 25–23, 35 FR 5671, Apr. 8, and center of gravity, if compliance 1970; Amdt. 25–42, 43 FR 2320, Jan. 16, 1978; cannot be reasonably inferred from Amdt. 25–72, 55 FR 29774, July 20, 1990; Amdt. combinations investigated. 25–121, 72 FR 44665, Aug. 8, 2007 Amdt. 25–135, (b) [Reserved] 76 FR 74654, Dec. 1, 2011] (c) The controllability, stability, trim, and stalling characteristics of § 25.23 Load distribution limits. the airplane must be shown for each al- (a) Ranges of weights and centers of titude up to the maximum expected in gravity within which the airplane may operation. be safely operated must be established. (d) Parameters critical for the test If a weight and center of gravity com- being conducted, such as weight, load- bination is allowable only within cer- ing (center of gravity and inertia), air- tain load distribution limits (such as speed, power, and wind, must be main- spanwise) that could be inadvertently tained within acceptable tolerances of exceeded, these limits and the cor- the critical values during flight test- responding weight and center of grav- ing. ity combinations must be established. (e) If compliance with the flight (b) The load distribution limits may characteristics requirements is depend- not exceed— ent upon a stability augmentation sys- (1) The selected limits; tem or upon any other automatic or (2) The limits at which the structure power-operated system, compliance is proven; or must be shown with §§ 25.671 and 25.672. (3) The limits at which compliance (f) In meeting the requirements of with each applicable flight require- §§ 25.105(d), 25.125, 25.233, and 25.237, the ment of this subpart is shown. wind velocity must be measured at a height of 10 meters above the surface, § 25.25 Weight limits. or corrected for the difference between (a) Maximum weights. Maximum the height at which the wind velocity weights corresponding to the airplane is measured and the 10-meter height. operating conditions (such as ramp, (g) The requirements of this subpart ground or water taxi, takeoff, en route, associated with icing conditions apply and landing), environmental conditions only if the applicant is seeking certifi- (such as altitude and temperature), and cation for flight in icing conditions. loading conditions (such as zero fuel (1) Each requirement of this subpart, weight, center of gravity position and except §§ 25.121(a), 25.123(c), 25.143(b)(1) weight distribution) must be estab- and (2), 25.149, 25.201(c)(2), 25.239, and lished so that they are not more than— 25.251(b) through (e), must be met in (1) The highest weight selected by icing conditions. Section 25.207(c) and the applicant for the particular condi- (d) must be met in the landing configu- tions; or ration in icing conditions, but need not (2) The highest weight at which com- be met for other configurations. Com- pliance with each applicable structural

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loading and flight requirement is water, and fluids intended for injection shown, except that for airplanes in the engine. equipped with standby power rocket (b) The condition of the airplane at engines the maximum weight must not the time of determining empty weight be more than the highest weight estab- must be one that is well defined and lished in accordance with appendix E of can be easily repeated. this part; or [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (3) The highest weight at which com- amended by Amdt. 25–42, 43 FR 2320, Jan. 16, pliance is shown with the certification 1978; Amdt. 25–72, 55 FR 29774, July 20, 1990] requirements of Part 36 of this chapter. (b) Minimum weight. The minimum § 25.31 Removable ballast. weight (the lowest weight at which Removable ballast may be used on compliance with each applicable re- showing compliance with the flight requirement of this part is shown) must quirements of this subpart. be established so that it is not less than— § 25.33 Propeller speed and pitch lim- (1) The lowest weight selected by the its. applicant; (a) The propeller speed and pitch (2) The design minimum weight (the must be limited to values that will en- lowest weight at which compliance sure— with each structural loading condition (1) Safe operation under normal oper- of this part is shown); or ating conditions; and (3) The lowest weight at which com- (2) Compliance with the performance pliance with each applicable flight re- requirements of §§ 25.101 through 25.125. quirement is shown. (b) There must be a propeller speed [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as limiting means at the governor. It amended by Amdt. 25–23, 35 FR 5671, Apr. 8, must limit the maximum possible gov- 1970; Amdt. 25–63, 53 FR 16365, May 6, 1988] erned engine speed to a value not exceeding the maximum allowable r.p.m. § 25.27 Center of gravity limits. (c) The means used to limit the low The extreme forward and the extreme pitch position of the propeller blades aft center of gravity limitations must must be set so that the engine does not be established for each practicably sep- exceed 103 percent of the maximum al- arable operating condition. No such lowable engine rpm or 99 percent of an limit may lie beyond— approved maximum overspeed, which- (a) The extremes selected by the ap- ever is greater, with— plicant; (1) The propeller blades at the low (b) The extremes within which the pitch limit and governor inoperative; structure is proven; or (2) The airplane stationary under (c) The extremes within which com- standard atmospheric conditions with pliance with each applicable flight re- no wind; and quirement is shown. (3) The engines operating at the takeoff manifold pressure limit for recipro- § 25.29 Empty weight and cor- cating engine powered airplanes or the responding center of gravity. maximum takeoff torque limit for tur- (a) The empty weight and cor- bopropeller engine-powered airplanes. responding center of gravity must be [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as determined by weighing the airplane amended by Amdt. 25–57, 49 FR 6848, Feb. 23, with— 1984; Amdt. 25–72, 55 FR 29774, July 20, 1990] (1) Fixed ballast; (2) Unusable fuel determined under PERFORMANCE § 25.959; and (3) Full operating fluids, including— § 25.101 General. (i) Oil; (a) Unless otherwise prescribed, air- (ii) Hydraulic fluid; and planes must meet the applicable per- (iii) Other fluids required for normal formance requirements of this subpart operation of airplane systems, except for ambient atmospheric conditions potable water, lavatory precharge and still air.

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(b) The performance, as affected by (f) Unless otherwise prescribed, in de- engine power or thrust, must be based termining the accelerate-stop dis- on the following relative humidities; tances, takeoff flight paths, takeoff (1) For turbine engine powered air- distances, and landing distances, planes, a relative humidity of— changes in the airplane’s configura- (i) 80 percent, at and below standard tion, speed, power, and thrust, must be temperatures; and made in accordance with procedures es- (ii) 34 percent, at and above standard tablished by the applicant for oper- temperatures plus 50 °F. ation in service. Between these two temperatures, the (g) Procedures for the execution of relative humidity must vary linearly. balked landings and missed approaches (2) For reciprocating engine powered associated with the conditions pre- airplanes, a relative humidity of 80 per- scribed in §§ 25.119 and 25.121(d) must be cent in a standard atmosphere. Engine established. power corrections for vapor pressure (h) The procedures established under must be made in accordance with the paragraphs (f) and (g) of this section following table: must— (1) Be able to be consistently exe- Vapor Specific humidity cuted in service by crews of average Altitude pressure e w (Lb. moisture Density ratio H (ft.) (In. Hg.) per lb. dry air) r / s=0.0023769 skill; (2) Use methods or devices that are 0 0.403 0.00849 0.99508 safe and reliable; and 1,000 .354 .00773 .96672 2,000 .311 .00703 .93895 (3) Include allowance for any time 3,000 .272 .00638 .91178 delays, in the execution of the proce- 4,000 .238 .00578 .88514 dures, that may reasonably be expected 5,000 .207 .00523 .85910 in service. 6,000 .1805 .00472 .83361 7,000 .1566 .00425 .80870 (i) The accelerate-stop and landing 8,000 .1356 .00382 .78434 distances prescribed in §§ 25.109 and 9,000 .1172 .00343 .76053 25.125, respectively, must be deter- 10,000 .1010 .00307 .73722 15,000 .0463 .001710 .62868 mined with all the airplane wheel 20,000 .01978 .000896 .53263 brake assemblies at the fully worn 25,000 .00778 .000436 .44806 limit of their allowable wear range.

(c) The performance must correspond [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as to the propulsive thrust available amended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; Amdt. 25–92, 63 FR 8318, Feb. 18, 1998] under the particular ambient atmospheric conditions, the particular flight § 25.103 Stall speed. condition, and the relative humidity specified in paragraph (b) of this sec- (a) The reference stall speed, VSR, is a tion. The available propulsive thrust calibrated airspeed defined by the ap- must correspond to engine power or plicant. VSR may not be less than a 1-g thrust, not exceeding the approved stall speed. VSR is expressed as: power or thrust less— V (1) Installation losses; and V ≥ CLMAX (2) The power or equivalent thrust SR n absorbed by the accessories and serv- ZW ices appropriate to the particular am- where: V = Calibrated airspeed obtained when bient atmospheric conditions and the CLMAX particular flight condition. the load factor-corrected lift coefficient (d) Unless otherwise prescribed, the ⎛ nW⎞ applicant must select the takeoff, en ⎜ ZW ⎟ route, approach, and landing configura- ⎝ qS ⎠ tions for the airplane. (e) The airplane configurations may is first a maximum during the maneuver vary with weight, altitude, and tem- prescribed in paragraph (c) of this section. In addition, when the maneuver is limited by a perature, to the extent they are com- device that abruptly pushes the nose down at patible with the operating procedures a selected angle of attack (e.g., a stick push- required by paragraph (f) of this sec- er), VCLMAX may not be less than the speed ex- tion. isting at the instant the device operates;

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nZW = Load factor normal to the flight path (1) In non-icing conditions; and at V CLMAX (2) In icing conditions, if in the con- W = Airplane gross weight; figuration of § 25.121(b) with the takeoff S = Aerodynamic reference wing area; and q = Dynamic pressure. ice accretion defined in appendix C: (i) The stall speed at maximum take- (b) VCLMAX is determined with: off weight exceeds that in non-icing (1) Engines idling, or, if that result- conditions by more than the greater of ant thrust causes an appreciable de- 3 knots CAS or 3 percent of V ; or crease in stall speed, not more than SR (ii) The degradation of the gradient zero thrust at the stall speed; (2) Propeller pitch controls (if appli- of climb determined in accordance with cable) in the takeoff position; § 25.121(b) is greater than one-half of (3) The airplane in other respects the applicable actual-to-net takeoff (such as flaps, landing gear, and ice ac- flight path gradient reduction defined cretions) in the condition existing in in § 25.115(b). the test or performance standard in (b) No takeoff made to determine the which V is being used; data required by this section may re- SR quire exceptional piloting skill or (4) The weight used when VSR is being used as a factor to determine compli- alertness. ance with a required performance (c) The takeoff data must be based standard; on— (5) The center of gravity position (1) In the case of land planes and am- that results in the highest value of ref- phibians: erence stall speed; and (i) Smooth, dry and wet, hard-sur- (6) The airplane trimmed for straight faced runways; and flight at a speed selected by the appli- (ii) At the option of the applicant, cant, but not less than 1.13VSR and not grooved or porous friction course wet, greater than 1.3VSR. hard-surfaced runways. (c) Starting from the stabilized trim (2) Smooth water, in the case of sea- condition, apply the longitudinal con- planes and amphibians; and trol to decelerate the airplane so that (3) Smooth, dry snow, in the case of the speed reduction does not exceed skiplanes. one knot per second. (d) The takeoff data must include, (d) In addition to the requirements of within the established operational lim- paragraph (a) of this section, when a its of the airplane, the following oper- device that abruptly pushes the nose ational correction factors: down at a selected angle of attack (e.g., (1) Not more than 50 percent of nomi- a stick pusher) is installed, the ref- nal wind components along the takeoff erence stall speed, VSR, may not be less path opposite to the direction of take- than 2 knots or 2 percent, whichever is off, and not less than 150 percent of greater, above the speed at which the nominal wind components along the device operates. takeoff path in the direction of takeoff. [Doc. No. 28404, 67 FR 70825, Nov. 26, 2002, as (2) Effective runway gradients. amended by Amdt. 25–121, 72 FR 44665, Aug. 8, 2007] [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–92, 63 FR 8318, Feb. 18, § 25.105 Takeoff. 1998; Amdt. 25–121, 72 FR 44665, Aug. 8, 2007] (a) The takeoff speeds prescribed by § 25.107 Takeoff speeds. § 25.107, the accelerate-stop distance prescribed by § 25.109, the takeoff path (a) V1 must be established in relation prescribed by § 25.111, the takeoff dis- to VEF as follows: tance and takeoff run prescribed by (1) VEF is the calibrated airspeed at § 25.113, and the net takeoff flight path which the critical engine is assumed to prescribed by § 25.115, must be deter- fail. VEF must be selected by the appli- mined in the selected configuration for cant, but may not be less than VMCG de- takeoff at each weight, altitude, and termined under § 25.149(e). ambient temperature within the oper- (2) V1, in terms of calibrated air- ational limits selected by the appli- speed, is selected by the applicant; cant— however, V1 may not be less than VEF 371

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plus the speed gained with critical en- (ii) 105 percent of VMC; gine inoperative during the time inter- (iii) The speed (determined in accord- val between the instant at which the ance with § 25.111(c)(2)) that allows critical engine is failed, and the in- reaching V2 before reaching a height of stant at which the pilot recognizes and 35 feet above the takeoff surface; or reacts to the engine failure, as indi- (iv) A speed that, if the airplane is cated by the pilot’s initiation of the rotated at its maximum practicable first action (e.g., applying brakes, re- rate, will result in a VLOF of not less ducing thrust, deploying speed brakes) than — to stop the airplane during accelerate- (A) 110 percent of VMU in the all-en- stop tests. gines-operating condition, and 105 per- (b) V2MIN, in terms of calibrated air- cent of VMU determined at the thrust- speed, may not be less than— to-weight ratio corresponding to the (1) 1.13 VSR for— one-engine-inoperative condition; or (i) Two-engine and three-engine tur- (B) If the V attitude is limited by bopropeller and reciprocating engine MU the geometry of the airplane (i.e., tail powered airplanes; and contact with the runway), 108 percent (ii) Turbojet powered airplanes with- of V in the all-engines-operating out provisions for obtaining a signifi- MU condition, and 104 percent of V deter- cant reduction in the one-engine-inop- MU mined at the thrust-to-weight ratio erative power-on stall speed; corresponding to the one-engine-inop- (2) 1.08 V for— SR erative condition. (i) Turbopropeller and reciprocating engine powered airplanes with more (2) For any given set of conditions than three engines; and (such as weight, configuration, and (ii) Turbojet powered airplanes with temperature), a single value of VR, ob- provisions for obtaining a significant tained in accordance with this para- reduction in the one-engine-inoper- graph, must be used to show compli- ative power-on stall speed; and ance with both the one-engine-inoperative and the all-engines-operating (3) 1.10 times VMC established under § 25.149. takeoff provisions. (3) It must be shown that the one-en- (c) V2, in terms of calibrated airspeed, must be selected by the appli- gine-inoperative takeoff distance, cant to provide at least the gradient of using a rotation speed of 5 knots less climb required by § 25.121(b) but may than VR established in accordance with not be less than— paragraphs (e)(1) and (2) of this section, (1) V ; does not exceed the corresponding one- 2MIN engine-inoperative takeoff distance (2) VR plus the speed increment attained (in accordance with § 25.111(c)(2)) using the established VR. The takeoff before reaching a height of 35 feet distances must be determined in ac- above the takeoff surface; and cordance with § 25.113(a)(1). (3) A speed that provides the maneu- (4) Reasonably expected variations in vering capability specified in § 25.143(h). service from the established takeoff procedures for the operation of the air- (d) VMU is the calibrated airspeed at and above which the airplane can safe- plane (such as over-rotation of the air- ly lift off the ground, and continue plane and out-of-trim conditions) may not result in unsafe flight characteris- the takeoff. VMU speeds must be selected by the applicant throughout the tics or in marked increases in the range of thrust-to-weight ratios to be scheduled takeoff distances established certificated. These speeds may be es- in accordance with § 25.113(a). tablished from free air data if these (f) VLOF is the calibrated airspeed at data are verified by ground takeoff which the airplane first becomes air- tests. borne. (e) VR, in terms of calibrated air- (g) VFTO, in terms of calibrated airspeed, must be selected in accordance speed, must be selected by the appli- with the conditions of paragraphs (e)(1) cant to provide at least the gradient of through (4) of this section: climb required by § 25.121(c), but may (1) VR may not be less than— not be less than— (i) V1; (1) 1.18 VSR; and 372

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(2) A speed that provides the maneu- pilot takes the first action to reject vering capability specified in § 25.143(h). the takeoff at the V1 for takeoff from a (h) In determining the takeoff speeds dry runway; and V1, VR, and V2 for flight in icing condi- (ii) With all engines still operating, tions, the values of VMCG, VMC, and VMU come to a full stop on dry runway from determined for non-icing conditions the speed reached as prescribed in para- may be used. graph (a)(2)(i) of this section; plus (iii) A distance equivalent to 2 sec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55466, Dec. 20, onds at the V1 for takeoff from a dry 1976; Amdt. 25–42, 43 FR 2320, Jan. 16, 1978; runway. Amdt. 25–92, 63 FR 8318, Feb. 18, 1998; Amdt. (b) The accelerate-stop distance on a 25–94, 63 FR 8848, Feb. 23, 1998; Amdt. 25–108, wet runway is the greater of the fol- 67 FR 70826, Nov. 26, 2002; Amdt. 25–121, 72 FR lowing distances: 44665, Aug. 8, 2007; Amdt. 25–135, 76 FR 74654, (1) The accelerate-stop distance on a Dec. 1, 2011] dry runway determined in accordance with paragraph (a) of this section; or § 25.109 Accelerate-stop distance. (2) The accelerate-stop distance de- (a) The accelerate-stop distance on a termined in accordance with paragraph dry runway is the greater of the fol- (a) of this section, except that the run- lowing distances: way is wet and the corresponding wet (1) The sum of the distances nec- runway values of VEF and V1 are used. essary to— In determining the wet runway accel- (i) Accelerate the airplane from a erate-stop distance, the stopping force standing start with all engines oper- from the wheel brakes may never ex- ating to VEF for takeoff from a dry run- ceed: way; (i) The wheel brakes stopping force (ii) Allow the airplane to accelerate determined in meeting the require- from VEF to the highest speed reached ments of § 25.101(i) and paragraph (a) of during the rejected takeoff, assuming this section; and the critical engine fails at VEF and the (ii) The force resulting from the wet pilot takes the first action to reject runway braking coefficient of friction the takeoff at the V1 for takeoff from a determined in accordance with para- dry runway; and graphs (c) or (d) of this section, as ap- (iii) Come to a full stop on a dry run- plicable, taking into account the dis- way from the speed reached as pre- tribution of the normal load between scribed in paragraph (a)(1)(ii) of this braked and unbraked wheels at the section; plus most adverse center-of-gravity position (iv) A distance equivalent to 2 sec- approved for takeoff. onds at the V1 for takeoff from a dry (c) The wet runway braking coeffi- runway. cient of friction for a smooth wet run- (2) The sum of the distances nec- way is defined as a curve of friction co- essary to— efficient versus ground speed and must (i) Accelerate the airplane from a be computed as follows: standing start with all engines oper- (1) The maximum tire-to-ground wet ating to the highest speed reached dur- runway braking coefficient of friction ing the rejected takeoff, assuming the is defined as:

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Where— Effi- Type of anti-skid system ciency Tire Pressure=maximum airplane operating value tire pressure (psi);

μt/gMAX=maximum tire-to-ground braking co- On-Off ...... 0.30 efficient; Quasi-Modulating ...... 0.50 Fully Modulating ...... 0.80 V=airplane true ground speed (knots); and Linear interpolation may be used for tire (d) At the option of the applicant, a pressures other than those listed. higher wet runway braking coefficient (2) The maximum tire-to-ground wet of friction may be used for runway sur- runway braking coefficient of friction faces that have been grooved or treated must be adjusted to take into account with a porous friction course material. the efficiency of the anti-skid system For grooved and porous friction course on a wet runway. Anti-skid system op- runways, the wet runway braking eration must be demonstrated by flight coefficent of friction is defined as ei- testing on a smooth wet runway, and ther: its efficiency must be determined. Un- (1) 70 percent of the dry runway brak- less a specific anti-skid system effi- ing coefficient of friction used to determine the dry runway accelerate-stop ciency is determined from a quan- distance; or titative analysis of the flight testing (2) The wet runway braking coeffi- on a smooth wet runway, the max- cient defined in paragraph (c) of this imum tire-to-ground wet runway brak- section, except that a specific anti-skid ing coefficient of friction determined system efficiency, if determined, is ap- in paragraph (c)(1) of this section must propriate for a grooved or porous fric- be multiplied by the efficiency value tion course wet runway, and the max- associated with the type of anti-skid imum tire-to-ground wet runway brak- system installed on the airplane: ing coefficient of friction is defined as:

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Where— teristics of the stopway and the vari- Tire Pressure=maximum airplane operating ations in these characteristics with tire pressure (psi); seasonal weather conditions (such as μ t/gMAX=maximum tire-to-ground braking co- temperature, rain, snow, and ice) with- efficient; in the established operational limits. V=airplane true ground speed (knots); and Linear interpolation may be used for tire (i) A flight test demonstration of the pressures other than those listed. maximum brake kinetic energy accelerate-stop distance must be conducted (e) Except as provided in paragraph with not more than 10 percent of the (f)(1) of this section, means other than allowable brake wear range remaining wheel brakes may be used to determine on each of the airplane wheel brakes. the accelerate-stop distance if that means— [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (1) Is safe and reliable; amended by Amdt. 25–42, 43 FR 2321, Jan. 16, (2) Is used so that consistent results 1978; Amdt. 25–92, 63 FR 8318, Feb. 18, 1998] can be expected under normal oper- § 25.111 Takeoff path. ating conditions; and (3) Is such that exceptional skill is (a) The takeoff path extends from a not required to control the airplane. standing start to a point in the takeoff (f) The effects of available reverse at which the airplane is 1,500 feet above thrust— the takeoff surface, or at which the (1) Shall not be included as an addi- transition from the takeoff to the en tional means of deceleration when de- route configuration is completed and termining the accelerate-stop distance VFTO is reached, whichever point is on a dry runway; and higher. In addition— (2) May be included as an additional (1) The takeoff path must be based on means of deceleration using rec- the procedures prescribed in § 25.101(f); ommended reverse thrust procedures (2) The airplane must be accelerated when determining the accelerate-stop on the ground to VEF, at which point distance on a wet runway, provided the the critical engine must be made inop- requirements of paragraph (e) of this erative and remain inoperative for the section are met. rest of the takeoff; and (g) The landing gear must remain ex- (3) After reaching VEF, the airplane tended throughout the accelerate-stop must be accelerated to V2. distance. (b) During the acceleration to speed (h) If the accelerate-stop distance in- V2, the nose gear may be raised off the cludes a stopway with surface charac- ground at a speed not less than VR. teristics substantially different from However, landing gear retraction may those of the runway, the takeoff data not be begun until the airplane is air- must include operational correction borne. factors for the accelerate-stop dis- (c) During the takeoff path deter- tance. The correction factors must ac- mination in accordance with para- count for the particular surface charac- graphs (a) and (b) of this section—

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(1) The slope of the airborne part of (3) The flight path must be based on the takeoff path must be positive at the airplane’s performance without each point; ground effect; and

(2) The airplane must reach V2 before (4) The takeoff path data must be it is 35 feet above the takeoff surface checked by continuous demonstrated and must continue at a speed as close takeoffs up to the point at which the as practical to, but not less than V2, airplane is out of ground effect and its until it is 400 feet above the takeoff speed is stabilized, to ensure that the surface; path is conservative relative to the (3) At each point along the takeoff continous path. path, starting at the point at which the The airplane is considered to be out of airplane reaches 400 feet above the the ground effect when it reaches a takeoff surface, the available gradient height equal to its wing span. of climb may not be less than— (e) For airplanes equipped with (i) 1.2 percent for two-engine air- standby power rocket engines, the planes; takeoff path may be determined in ac- (ii) 1.5 percent for three-engine air- cordance with section II of appendix E. planes; and [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (iii) 1.7 percent for four-engine air- amended by Amdt. 25–6, 30 FR 8468, July 2, planes. 1965; Amdt. 25–42, 43 FR 2321, Jan. 16, 1978; (4) The airplane configuration may Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; Amdt. 25–72, 55 FR 29774, July 20, 1990; Amdt. 25–94, not be changed, except for gear retrac- 63 FR 8848, Feb. 23, 1998; Amdt. 25–108, 67 FR tion and automatic propeller feath- 70826, Nov. 26, 2002; Amdt. 25–115, 69 FR 40527, ering, and no change in power or thrust July 2, 2004; Amdt. 25–121, 72 FR 44666; Aug. that requires action by the pilot may 8, 2007] be made until the airplane is 400 feet above the takeoff surface; and § 25.113 Takeoff distance and takeoff run. (5) If § 25.105(a)(2) requires the takeoff path to be determined for flight in (a) Takeoff distance on a dry runway icing conditions, the airborne part of is the greater of— the takeoff must be based on the air- (1) The horizontal distance along the plane drag: takeoff path from the start of the take- (i) With the takeoff ice accretion de- off to the point at which the airplane is fined in appendix C, from a height of 35 35 feet above the takeoff surface, deter- feet above the takeoff surface up to the mined under § 25.111 for a dry runway; point where the airplane is 400 feet or above the takeoff surface; and (2) 115 percent of the horizontal distance along the takeoff path, with all (ii) With the final takeoff ice accre- engines operating, from the start of the tion defined in appendix C, from the takeoff to the point at which the air- point where the airplane is 400 feet plane is 35 feet above the takeoff sur- above the takeoff surface to the end of face, as determined by a procedure con- the takeoff path. sistent with § 25.111. (d) The takeoff path must be deter- (b) Takeoff distance on a wet runway mined by a continuous demonstrated is the greater of— takeoff or by synthesis from segments. (1) The takeoff distance on a dry run- If the takeoff path is determined by the way determined in accordance with segmental method— paragraph (a) of this section; or (1) The segments must be clearly de- (2) The horizontal distance along the fined and must be related to the dis- takeoff path from the start of the take- tinct changes in the configuration, off to the point at which the airplane is power or thrust, and speed; 15 feet above the takeoff surface, (2) The weight of the airplane, the achieved in a manner consistent with configuration, and the power or thrust the achievement of V2 before reaching must be constant throughout each seg- 35 feet above the takeoff surface, deter- ment and must correspond to the most mined under § 25.111 for a wet runway. critical condition prevailing in the seg- (c) If the takeoff distance does not in- ment; clude a clearway, the takeoff run is

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equal to the takeoff distance. If the (1) 0.8 percent for two-engine air- takeoff distance includes a clearway— planes; (1) The takeoff run on a dry runway (2) 0.9 percent for three-engine air- is the greater of— planes; and (i) The horizontal distance along the (3) 1.0 percent for four-engine air- takeoff path from the start of the take- planes. off to a point equidistant between the (c) The prescribed reduction in climb point at which VLOF is reached and the gradient may be applied as an equiva- point at which the airplane is 35 feet lent reduction in acceleration along above the takeoff surface, as deter- that part of the takeoff flight path at mined under § 25.111 for a dry runway; which the airplane is accelerated in or level flight. (ii) 115 percent of the horizontal distance along the takeoff path, with all [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as engines operating, from the start of the amended by Amdt. 25–92, 63 FR 8320, Feb. 18, 1998] takeoff to a point equidistant between the point at which VLOF is reached and § 25.117 Climb: general. the point at which the airplane is 35 feet above the takeoff surface, deter- Compliance with the requirements of mined by a procedure consistent with §§ 25.119 and 25.121 must be shown at § 25.111. each weight, altitude, and ambient (2) The takeoff run on a wet runway temperature within the operational is the greater of— limits established for the airplane and (i) The horizontal distance along the with the most unfavorable center of takeoff path from the start of the take- gravity for each configuration. off to the point at which the airplane is § 25.119 Landing climb: All-engines-op- 15 feet above the takeoff surface, erating. achieved in a manner consistent with In the landing configuration, the the achievement of V2 before reaching 35 feet above the takeoff surface, as de- steady gradient of climb may not be termined under § 25.111 for a wet run- less than 3.2 percent, with the engines way; or at the power or thrust that is available (ii) 115 percent of the horizontal dis- 8 seconds after initiation of movement tance along the takeoff path, with all of the power or thrust controls from engines operating, from the start of the the minimum flight idle to the go- takeoff to a point equidistant between around power or thrust setting— (a) In non-icing conditions, with a the point at which VLOF is reached and the point at which the airplane is 35 climb speed of VREF determined in ac- feet above the takeoff surface, deter- cordance with § 25.125(b)(2)(i); and mined by a procedure consistent with (b) In icing conditions with the land- § 25.111. ing ice accretion defined in appendix C, and with a climb speed of VREF deter- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as mined in accordance with amended by Amdt. 25–23, 35 FR 5671, Apr. 8, § 25.125(b)(2)(ii). 1970; Amdt. 25–92, 63 FR 8320, Feb. 18, 1998] [Amdt. 25–121, 72 FR 44666; Aug. 8, 2007] § 25.115 Takeoff flight path. (a) The takeoff flight path shall be § 25.121 Climb: One-engine-inoper- considered to begin 35 feet above the ative. takeoff surface at the end of the take- (a) Takeoff; landing gear extended. In off distance determined in accordance the critical takeoff configuration exist- with § 25.113(a) or (b), as appropriate for ing along the flight path (between the the runway surface condition. points at which the airplane reaches (b) The net takeoff flight path data VLOF and at which the landing gear is must be determined so that they rep- fully retracted) and in the configura- resent the actual takeoff flight paths tion used in § 25.111 but without ground (determined in accordance with § 25.111 effect, the steady gradient of climb and with paragraph (a) of this section) must be positive for two-engine air- reduced at each point by a gradient of planes, and not less than 0.3 percent for climb equal to— three-engine airplanes or 0.5 percent

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for four-engine airplanes, at VLOF and (c) Final takeoff. In the en route con- with— figuration at the end of the takeoff (1) The critical engine inoperative path determined in accordance with and the remaining engines at the power § 25.111: or thrust available when retraction of (1) The steady gradient of climb may the landing gear is begun in accordance not be less than 1.2 percent for two-en- with § 25.111 unless there is a more crit- gine airplanes, 1.5 percent for three-en- ical power operating condition existing gine airplanes, and 1.7 percent for four- later along the flight path but before engine airplanes, at VFTO with— the point at which the landing gear is (i) The critical engine inoperative fully retracted; and and the remaining engines at the avail- (2) The weight equal to the weight able maximum continuous power or existing when retraction of the landing thrust; and gear is begun, determined under (ii) The weight equal to the weight § 25.111. existing at the end of the takeoff path, (b) Takeoff; landing gear retracted. In determined under § 25.111. the takeoff configuration existing at (2) The requirements of paragraph the point of the flight path at which (c)(1) of this section must be met: the landing gear is fully retracted, and (i) In non-icing conditions; and in the configuration used in § 25.111 but (ii) In icing conditions with the final without ground effect: takeoff ice accretion defined in appen- (1) The steady gradient of climb may dix C, if in the configuration of not be less than 2.4 percent for two-en- § 25.121(b) with the takeoff ice accre- gine airplanes, 2.7 percent for three-en- tion: gine airplanes, and 3.0 percent for four- (A) The stall speed at maximum engine airplanes, at V2 with: takeoff weight exceeds that in non- (i) The critical engine inoperative, icing conditions by more than the the remaining engines at the takeoff greater of 3 knots CAS or 3 percent of power or thrust available at the time VSR; or the landing gear is fully retracted, de- (B) The degradation of the gradient termined under § 25.111, unless there is of climb determined in accordance with a more critical power operating condi- § 25.121(b) is greater than one-half of tion existing later along the flight path the applicable actual-to-net takeoff but before the point where the airplane flight path gradient reduction defined reaches a height of 400 feet above the in § 25.115(b). takeoff surface; and (d) Approach. In a configuration cor- (ii) The weight equal to the weight responding to the normal all-engines- existing when the airplane’s landing operating procedure in which VSR for gear is fully retracted, determined this configuration does not exceed 110 under § 25.111. percent of the VSR for the related all- (2) The requirements of paragraph engines-operating landing configura- (b)(1) of this section must be met: tion: (i) In non-icing conditions; and (1) The steady gradient of climb may (ii) In icing conditions with the take- not be less than 2.1 percent for two-en- off ice accretion defined in appendix C, gine airplanes, 2.4 percent for three-en- if in the configuration of § 25.121(b) gine airplanes, and 2.7 percent for four- with the takeoff ice accretion: engine airplanes, with— (A) The stall speed at maximum (i) The critical engine inoperative, takeoff weight exceeds that in non- the remaining engines at the go-around icing conditions by more than the power or thrust setting; greater of 3 knots CAS or 3 percent of (ii) The maximum landing weight; VSR; or (iii) A climb speed established in con- (B) The degradation of the gradient nection with normal landing proce- of climb determined in accordance with dures, but not exceeding 1.4 VSR; and § 25.121(b) is greater than one-half of (iv) Landing gear retracted. the applicable actual-to-net takeoff (2) The requirements of paragraph flight path gradient reduction defined (d)(1) of this section must be met: in § 25.115(b). (i) In non-icing conditions; and

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(ii) In icing conditions with the ap- duction defined in paragraph (b) of this proach ice accretion defined in appen- section. dix C. The climb speed selected for non- (c) For three- or four-engine air- icing conditions may be used if the planes, the two-engine-inoperative net climb speed for icing conditions, com- flight path data must represent the ac- puted in accordance with paragraph tual climb performance diminished by (d)(1)(iii) of this section, does not ex- a gradient of climb of 0.3 percent for ceed that for non-icing conditions by three-engine airplanes and 0.5 percent more than the greater of 3 knots CAS for four-engine airplanes. or 3 percent. [Docket No. 5066, 29 FR 18291, Dec. 24, 1964, as [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–121, 72 FR 44666; Aug. 8, amended by Amdt. 25–84, 60 FR 30749, June 9, 2007] 1995; Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; Amdt. 25–121, 72 FR 44666; Aug. 8, 2007] § 25.125 Landing. (a) The horizontal distance necessary § 25.123 En route flight paths. to land and to come to a complete stop (a) For the en route configuration, (or to a speed of approximately 3 knots the flight paths prescribed in para- for water landings) from a point 50 feet graph (b) and (c) of this section must above the landing surface must be de- be determined at each weight, altitude, termined (for standard temperatures, and ambient temperature, within the at each weight, altitude, and wind operating limits established for the within the operational limits estab- airplane. The variation of weight along lished by the applicant for the air- the flight path, accounting for the pro- plane): gressive consumption of fuel and oil by (1) In non-icing conditions; and the operating engines, may be included (2) In icing conditions with the land- in the computation. The flight paths ing ice accretion defined in appendix C must be determined at a speed not less if VREF for icing conditions exceeds VREF than VFTO, with— for non-icing conditions by more than 5 (1) The most unfavorable center of knots CAS at the maximum landing gravity; weight. (2) The critical engines inoperative; (b) In determining the distance in (3) The remaining engines at the paragraph (a) of this section: available maximum continuous power (1) The airplane must be in the land- or thrust; and ing configuration. (4) The means for controlling the en- (2) A stabilized approach, with a cali- gine-cooling air supply in the position brated airspeed of not less than VREF, that provides adequate cooling in the must be maintained down to the 50-foot hot-day condition. height. (b) The one-engine-inoperative net (i) In non-icing conditions, VREF may flight path data must represent the ac- not be less than: tual climb performance diminished by (A) 1.23 VSR0; a gradient of climb of 1.1 percent for (B) VMCL established under § 25.149(f); two-engine airplanes, 1.4 percent for and three-engine airplanes, and 1.6 percent (C) A speed that provides the maneu- for four-engine airplanes— vering capability specified in § 25.143(h). (1) In non-icing conditions; and (ii) In icing conditions, VREF may not (2) In icing conditions with the en be less than: route ice accretion defined in appendix (A) The speed determined in para- C, if: graph (b)(2)(i) of this section; (i) A speed of 1.18 ‘‘VSR0 with the en (B) 1.23 VSR0 with the landing ice ac- route ice accretion exceeds the en cretion defined in appendix C if that route speed selected for non-icing con- speed exceeds VREF for non-icing condi- ditions by more than the greater of 3 tions by more than 5 knots CAS; and knots CAS or 3 percent of VSR; or (C) A speed that provides the maneu- (ii) The degradation of the gradient vering capability specified in § 25.143(h) of climb is greater than one-half of the with the landing ice accretion defined applicable actual-to-net flight path re- in appendix C.

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(3) Changes in configuration, power CONTROLLABILITY AND or thrust, and speed, must be made in MANEUVERABILITY accordance with the established procedures for service operation. § 25.143 General. (4) The landing must be made with- (a) The airplane must be safely con- out excessive vertical acceleration, trollable and maneuverable during— tendency to bounce, nose over, ground (1) Takeoff; loop, porpoise, or water loop. (2) Climb; (5) The landings may not require ex- (3) Level flight; ceptional piloting skill or alertness. (4) Descent; and (c) For landplanes and amphibians, (5) Landing. the landing distance on land must be (b) It must be possible to make a determined on a level, smooth, dry, smooth transition from one flight con- hard-surfaced runway. In addition— dition to any other flight condition (1) The pressures on the wheel brak- without exceptional piloting skill, ing systems may not exceed those spec- alertness, or strength, and without ified by the brake manufacturer; danger of exceeding the airplane limit- (2) The brakes may not be used so as load factor under any probable oper- to cause excessive wear of brakes or ating conditions, including— tires; and (1) The sudden failure of the critical (3) Means other than wheel brakes engine; may be used if that means— (2) For airplanes with three or more engines, the sudden failure of the sec- (i) Is safe and reliable; ond critical engine when the airplane is (ii) Is used so that consistent results in the en route, approach, or landing can be expected in service; and configuration and is trimmed with the (iii) Is such that exceptional skill is critical engine inoperative; and not required to control the airplane. (3) Configuration changes, including (d) For seaplanes and amphibians, deployment or retraction of decelera- the landing distance on water must be tion devices. determined on smooth water. (c) The airplane must be shown to be (e) For skiplanes, the landing dis- safely controllable and maneuverable tance on snow must be determined on with the critical ice accretion appro- smooth, dry, snow. priate to the phase of flight defined in (f) The landing distance data must appendix C, and with the critical en- include correction factors for not more gine inoperative and its propeller (if than 50 percent of the nominal wind applicable) in the minimum drag posi- components along the landing path option: posite to the direction of landing, and (1) At the minimum V2 for takeoff; not less than 150 percent of the nomi- (2) During an approach and go- nal wind components along the landing around; and path in the direction of landing. (3) During an approach and landing. (g) If any device is used that depends (d) The following table prescribes, for on the operation of any engine, and if conventional wheel type controls, the the landing distance would be notice- maximum control forces permitted ably increased when a landing is made during the testing required by para- with that engine inoperative, the land- graph (a) through (c) of this section: ing distance must be determined with Force, in pounds, applied to the that engine inoperative unless the use control wheel or rudder pedals Pitch Roll Yaw of compensating means will result in a For short term application for landing distance not more than that pitch and roll control—two with each engine operating. hands available for control .... 75 50 For short term application for [Amdt. 25–121, 72 FR 44666; Aug. 8, 2007; 72 FR pitch and roll control—one 50467, Aug. 31, 2007] hand available for control ...... 50 25 For short term application for yaw control ...... 150 For long term application ...... 10 5 20

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(e) Approved operating procedures or of the stick force versus maneuvering conventional operating practices must load factor must lie within satisfactory be followed when demonstrating com- limits. The stick forces must not be so pliance with the control force limita- great as to make excessive demands on tions for short term application that the pilot’s strength when maneuvering are prescribed in paragraph (d) of this the airplane, and must not be so low section. The airplane must be in trim, that the airplane can easily be over- or as near to being in trim as practical, stressed inadvertently. Changes of gra- in the preceding steady flight condi- dient that occur with changes of load tion. For the takeoff condition, the air- factor must not cause undue difficulty plane must be trimmed according to in maintaining control of the airplane, the approved operating procedures. and local gradients must not be so low (f) When demonstrating compliance as to result in a danger of overcontrol- with the control force limitations for ling. long term application that are pre- (h) The maneuvering capabilities in a scribed in paragraph (d) of this section, constant speed coordinated turn at for- the airplane must be in trim, or as near ward center of gravity, as specified in to being in trim as practical. the following table, must be free of (g) When maneuvering at a constant stall warning or other characteristics airspeed or Mach number (up to VFC/ that might interfere with normal ma- MFC), the stick forces and the gradient neuvering:

Maneuvering Configuration Speed bank angle in a Thrust/power setting coordinated turn

1 Takeoff ...... V2 30° Asymmetric WAT-Limited. 2 3 Takeoff ...... V2 + XX 40° All-engines-operating climb. 1 En route ...... VFTO 40° Asymmetric WAT-Limited. Landing ...... VREF 40° Symmetric for ¥3° flight path angle. 1 A combination of weight, altitude, and temperature (WAT) such that the thrust or power setting produces the minimum climb gradient specified in § 25.121 for the flight condition. 2 Airspeed approved for all-engines-operating initial climb. 3 That thrust or power setting which, in the event of failure of the critical engine and without any crew action to adjust the thrust or power of the remaining engines, would result in the thrust or power specified for the takeoff condition at V2, or any lesser thrust or power setting that is used for all-engines-operating initial climb procedures.

(i) When demonstrating compliance (j) For flight in icing conditions be- with § 25.143 in icing conditions— fore the ice protection system has been (1) Controllability must be dem- activated and is performing its in- onstrated with the ice accretion de- tended function, it must be dem- fined in appendix C that is most crit- onstrated in flight with the ice accre- ical for the particular flight phase; tion defined in appendix C, part II(e) of (2) It must be shown that a push force this part that: is required throughout a pushover ma- (1) The airplane is controllable in a neuver down to a zero g load factor, or pull-up maneuver up to 1.5 g load fac- the lowest load factor obtainable if tor; and limited by elevator power or other de- (2) There is no pitch control force re- sign characteristic of the flight control versal during a pushover maneuver system. It must be possible to prompt- down to 0.5 g load factor. ly recover from the maneuver without [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as exceeding a pull control force of 50 amended by Amdt. 25–42, 43 FR 2321, Jan. 16, pounds; and 1978; Amdt. 25–84, 60 FR 30749, June 9, 1995; (3) Any changes in force that the Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; pilot must apply to the pitch control to Amdt. 25–121, 72 FR 44667, Aug. 8, 2007; Amdt. maintain speed with increasing sideslip 25–129, 74 FR 38339, Aug. 3, 2009] angle must be steadily increasing with no force reversals, unless the change in § 25.145 Longitudinal control. control force is gradual and easily con- (a) It must be possible, at any point trollable by the pilot without using ex- between the trim speed prescribed in ceptional piloting skill, alertness, or § 25.103(b)(6) and stall identification (as strength. defined in § 25.201(d)), to pitch the nose

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downward so that the acceleration to positions, and from the last gated posi- this selected trim speed is prompt with tion to the fully retracted position. (1) The airplane trimmed at the trim The requirements of paragraph (c) of speed prescribed in § 25.103(b)(6); this section also apply to retractions (2) The landing gear extended; from each approved landing position to (3) The wing flaps (i) retracted and the control position(s) associated with (ii) extended; and the high-lift device configuration(s) (4) Power (i) off and (ii) at maximum used to establish the go-around proce- continuous power on the engines. dure(s) from that landing position. In (b) With the landing gear extended, addition, the first gated control posi- no change in trim control, or exertion tion from the maximum landing posi- of more than 50 pounds control force tion must correspond with a configura- (representative of the maximum short tion of the high-lift devices used to es- term force that can be applied readily tablish a go-around procedure from a by one hand) may be required for the landing configuration. Each gated con- following maneuvers: trol position must require a separate (1) With power off, flaps retracted, and distinct motion of the control to and the airplane trimmed at 1.3 VSR1, pass through the gated position and extend the flaps as rapidly as possible must have features to prevent inad- while maintaining the airspeed at ap- vertent movement of the control proximately 30 percent above the ref- through the gated position. It must erence stall speed existing at each in- only be possible to make this separate stant throughout the maneuver. and distinct motion once the control (2) Repeat paragraph (b)(1) except ini- has reached the gated position. tially extend the flaps and then retract them as rapidly as possible. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (3) Repeat paragraph (b)(2), except at amended by Amdt. 25–23, 35 FR 5671, Apr. 8, the go-around power or thrust setting. 1970; Amdt. 25–72, 55 FR 29774, July 20, 1990; Amdt. 25–84, 60 FR 30749, June 9, 1995; Amdt. (4) With power off, flaps retracted, 25–98, 64 FR 6164, Feb. 8, 1999; 64 FR 10740, and the airplane trimmed at 1.3 VSR1, Mar. 5, 1999; Amdt. 25–108, 67 FR 70827, Nov. rapidly set go-around power or thrust 26, 2002] while maintaining the same airspeed. (5) Repeat paragraph (b)(4) except § 25.147 Directional and lateral con- with flaps extended. trol. (6) With power off, flaps extended, (a) Directional control; general. It must and the airplane trimmed at 1.3 VSR1, be possible, with the wings level, to obtain and maintain airspeeds between yaw into the operative engine and to VSW and either 1.6 VSR1 or VFE, which- safely make a reasonably sudden ever is lower. change in heading of up to 15 degrees in (c) It must be possible, without ex- the direction of the critical inoperative ceptional piloting skill, to prevent loss engine. This must be shown at 1.3 V R1 of altitude when complete retraction of S for heading changes up to 15 degrees the high lift devices from any position (except that the heading change at is begun during steady, straight, level which the rudder pedal force is 150 flight at 1.08 V for propeller powered SR1 pounds need not be exceeded), and airplanes, or 1.13 V for turbojet pow- SR1 with— ered airplanes, with— (1) The critical engine inoperative (1) Simultaneous movement of the and its propeller in the minimum drag power or thrust controls to the go- position; around power or thrust setting; (2) The landing gear extended; and (2) The power required for level flight (3) The critical combinations of land- at 1.3 VSR1, but not more than max- ing weights and altitudes. imum continuous power; (d) If gated high-lift device control (3) The most unfavorable center of positions are provided, paragraph (c) of gravity; this section applies to retractions of (4) Landing gear retracted; the high-lift devices from any position (5) Flaps in the approach position; from the maximum landing position to and the first gated position, between gated (6) Maximum landing weight.

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(b) Directional control; airplanes with for safety, without excessive control four or more engines. Airplanes with forces or travel. four or more engines must meet the re- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as quirements of paragraph (a) of this sec- amended by Amdt. 25–42, 43 FR 2321, Jan. 16, tion except that— 1978; Amdt. 25–72, 55 FR 29774, July 20, 1990; (1) The two critical engines must be Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; inoperative with their propellers (if ap- Amdt. 25–115, 69 FR 40527, July 2, 2004] plicable) in the minimum drag position; § 25.149 Minimum control speed. (2) [Reserved] (a) In establishing the minimum con- (3) The flaps must be in the most fa- trol speeds required by this section, the vorable climb position. method used to simulate critical en- (c) Lateral control; general. It must be gine failure must represent the most possible to make 20° banked turns, with critical mode of powerplant failure and against the inoperative engine, with respect to controllability ex- from steady flight at a speed equal to pected in service. 1.3 VSR1, with— (b) VMC is the calibrated airspeed at (1) The critical engine inoperative which, when the critical engine is sud- and its propeller (if applicable) in the denly made inoperative, it is possible minimum drag position; to maintain control of the airplane (2) The remaining engines at max- with that engine still inoperative and imum continuous power; maintain straight flight with an angle (3) The most unfavorable center of of bank of not more than 5 degrees. gravity; (c) VMC may not exceed 1.13 VSR (4) Landing gear (i) retracted and (ii) with— extended; (1) Maximum available takeoff power (5) Flaps in the most favorable climb or thrust on the engines; position; and (2) The most unfavorable center of (6) Maximum takeoff weight. gravity; (d) Lateral control; roll capability. With the critical engine inoperative, roll re- (3) The airplane trimmed for takeoff; sponse must allow normal maneuvers. (4) The maximum sea level takeoff Lateral control must be sufficient, at weight (or any lesser weight necessary the speeds likely to be used with one to show VMC); engine inoperative, to provide a roll (5) The airplane in the most critical rate necessary for safety without ex- takeoff configuration existing along cessive control forces or travel. the flight path after the airplane be- (e) Lateral control; airplanes with four comes airborne, except with the land- or more engines. Airplanes with four or ing gear retracted; more engines must be able to make 20° (6) The airplane airborne and the banked turns, with and against the in- ground effect negligible; and operative engines, from steady flight at (7) If applicable, the propeller of the inoperative engine— a speed equal to 1.3 VSR1, with maximum continuous power, and with the (i) Windmilling; airplane in the configuration pre- (ii) In the most probable position for scribed by paragraph (b) of this section. the specific design of the propeller con- (f) Lateral control; all engines oper- trol; or ating. With the engines operating, roll (iii) Feathered, if the airplane has an response must allow normal maneuvers automatic feathering device acceptable (such as recovery from upsets produced for showing compliance with the climb by gusts and the initiation of evasive requirements of § 25.121. maneuvers). There must be enough ex- (d) The rudder forces required to cess lateral control in sideslips (up to maintain control at VMC may not ex- sideslip angles that might be required ceed 150 pounds nor may it be nec- in normal operation), to allow a lim- essary to reduce power or thrust of the ited amount of maneuvering and to operative engines. During recovery, the correct for gusts. Lateral control must airplane may not assume any dan- be enough at any speed up to VFC/MFC gerous attitude or require exceptional to provide a peak roll rate necessary piloting skill, alertness, or strength to

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prevent a heading change of more than (4) The most favorable weight, or, at 20 degrees. the option of the applicant, as a func-

(e) VMCG, the minimum control speed tion of weight; on the ground, is the calibrated air- (5) For propeller airplanes, the pro- speed during the takeoff run at which, peller of the inoperative engine in the when the critical engine is suddenly position it achieves without pilot ac- made inoperative, it is possible to tion, assuming the engine fails while at maintain control of the airplane using the power or thrust necessary to main- the rudder control alone (without the tain a three degree approach path use of nosewheel steering), as limited angle; and by 150 pounds of force, and the lateral (6) Go-around power or thrust setting control to the extent of keeping the on the operating engine(s). wings level to enable the takeoff to be (g) For airplanes with three or more safely continued using normal piloting engines, VMCL-2, the minimum control skill. In the determination of VMCG, as- speed during approach and landing suming that the path of the airplane with one critical engine inoperative, is accelerating with all engines operating the calibrated airspeed at which, when is along the centerline of the runway, a second critical engine is suddenly its path from the point at which the made inoperative, it is possible to critical engine is made inoperative to maintain control of the airplane with the point at which recovery to a direc- both engines still inoperative, and tion parallel to the centerline is com- maintain straight flight with an angle pleted may not deviate more than 30 of bank of not more than 5 degrees. feet laterally from the centerline at VMCL-2 must be established with— (1) The airplane in the most critical any point. VMCG must be established with— configuration (or, at the option of the (1) The airplane in each takeoff con- applicant, each configuration) for ap- figuration or, at the option of the approach and landing with one critical plicant, in the most critical takeoff engine inoperative; configuration; (2) The most unfavorable center of (2) Maximum available takeoff power gravity; or thrust on the operating engines; (3) The airplane trimmed for approach with one critical engine inoper- (3) The most unfavorable center of ative; gravity; (4) The most unfavorable weight, or, (4) The airplane trimmed for takeoff; at the option of the applicant, as a and function of weight; (5) The most unfavorable weight in (5) For propeller airplanes, the pro- the range of takeoff weights. peller of the more critical inoperative (f) VMCL, the minimum control speed engine in the position it achieves with- during approach and landing with all out pilot action, assuming the engine engines operating, is the calibrated air- fails while at the power or thrust nec- speed at which, when the critical en- essary to maintain a three degree ap- gine is suddenly made inoperative, it is proach path angle, and the propeller of possible to maintain control of the air- the other inoperative engine feathered; plane with that engine still inoper- (6) The power or thrust on the operative, and maintain straight flight ating engine(s) necessary to maintain with an angle of bank of not more than an approach path angle of three de- 5 degrees. VMCL must be established grees when one critical engine is inop- with— erative; and (1) The airplane in the most critical (7) The power or thrust on the oper- configuration (or, at the option of the ating engine(s) rapidly changed, imme- applicant, each configuration) for ap- diately after the second critical engine proach and landing with all engines op- is made inoperative, from the power or erating; thrust prescribed in paragraph (g)(6) of (2) The most unfavorable center of this section to— gravity; (i) Minimum power or thrust; and (3) The airplane trimmed for ap- (ii) Go-around power or thrust set- proach with all engines operating; ting.

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(h) In demonstrations of VMCL and weight and configuration with power VMCL-2— settings corresponding to a 3 degree (1) The rudder force may not exceed glidepath, whichever is the most se- 150 pounds; vere, with the landing gear extended, (2) The airplane may not exhibit haz- the wing flaps (i) retracted and (ii) ex- ardous flight characteristics or require tended, and with the most unfavorable exceptional piloting skill, alertness, or combination of center of gravity posi- strength; tion and weight approved for landing; (3) Lateral control must be sufficient and to roll the airplane, from an initial (3) Level flight at any speed from 1.3 condition of steady flight, through an V , to V /M with the landing gear angle of 20 degrees in the direction nec- SR1 MO MO, and flaps retracted, and from 1.3 V to essary to initiate a turn away from the SR1 V with the landing gear extended. inoperative engine(s), in not more than LE 5 seconds; and (d) Longitudinal, directional, and lat- (4) For propeller airplanes, hazardous eral trim. The airplane must maintain flight characteristics must not be ex- longitudinal, directional, and lateral hibited due to any propeller position trim (and for the lateral trim, the achieved when the engine fails or dur- angle of bank may not exceed five de- ing any likely subsequent movements grees) at 1.3 VSR1 during climbing flight of the engine or propeller controls. with— (1) The critical engine inoperative; [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–42, 43 FR 2321, Jan. 16, (2) The remaining engines at max- 1978; Amdt. 25–72, 55 FR 29774, July 20, 1990; 55 imum continuous power; and FR 37607, Sept. 12, 1990; Amdt. 25–84, 60 FR (3) The landing gear and flaps re- 30749, June 9, 1995; Amdt. 25–108, 67 FR 70827, tracted. Nov. 26, 2002] (e) Airplanes with four or more engines. Each airplane with four or more TRIM engines must also maintain trim in § 25.161 Trim. rectilinear flight with the most unfa- (a) General. Each airplane must meet vorable center of gravity and at the the trim requirements of this section climb speed, configuration, and power after being trimmed, and without fur- required by § 25.123(a) for the purpose of ther pressure upon, or movement of, ei- establishing the en route flight paths ther the primary controls or their cor- with two engines inoperative. responding trim controls by the pilot [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as or the automatic pilot. amended by Amdt. 25–23, 35 FR 5671, Apr. 8, (b) Lateral and directional trim. The 1970; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; airplane must maintain lateral and di- Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; rectional trim with the most adverse Amdt. 25–115, 69 FR 40527, July 2, 2004] lateral displacement of the center of gravity within the relevant operating STABILITY limitations, during normally expected conditions of operation (including op- § 25.171 General. eration at any speed from 1.3 VSR1 to The airplane must be longitudinally, VMO/MMO). directionally, and laterally stable in (c) Longitudinal trim. The airplane accordance with the provisions of must maintain longitudinal trim dur- §§ 25.173 through 25.177. In addition, ing— suitable stability and control feel (1) A climb with maximum contin- (static stability) is required in any con- uous power at a speed not more than dition normally encountered in service, 1.3 V , with the landing gear re- SR1 if flight tests show it is necessary for tracted, and the flaps (i) retracted and safe operation. (ii) in the takeoff position; (2) Either a glide with power off at a [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as speed not more than 1.3 VSR1, or an ap- amended by Amdt. 25–7, 30 FR 13117, Oct. 15, proach within the normal range of ap- 1965] proach speeds appropriate to the

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§ 25.173 Static longitudinal stability. tation for use during climb for turbine Under the conditions specified in engines; and § 25.175, the characteristics of the eleva- (2) Is trimmed at the speed for best tor control forces (including friction) rate-of-climb except that the speed must be as follows: need not be less than 1.3 VSR1. (a) A pull must be required to obtain (b) Cruise. Static longitudinal sta- and maintain speeds below the speci- bility must be shown in the cruise con- fied trim speed, and a push must be re- dition as follows: quired to obtain and maintain speeds (1) With the landing gear retracted at above the specified trim speed. This high speed, the stick force curve must must be shown at any speed that can be have a stable slope at all speeds within obtained except speeds higher than the a range which is the greater of 15 per- landing gear or wing flap operating cent of the trim speed plus the result- limit speeds or VFC/MFC, whichever is ing free return speed range, or 50 knots appropriate, or lower than the min- plus the resulting free return speed imum speed for steady unstalled flight. range, above and below the trim speed (b) The airspeed must return to with- (except that the speed range need not in 10 percent of the original trim speed include speeds less than 1.3 VSR1, nor for the climb, approach, and landing speeds greater than VFC/MFC, nor speeds conditions specified in § 25.175 (a), (c), that require a stick force of more than and (d), and must return to within 7.5 50 pounds), with— percent of the original trim speed for (i) The wing flaps retracted; the cruising condition specified in (ii) The center of gravity in the most § 25.175(b), when the control force is adverse position (see § 25.27); slowly released from any speed within (iii) The most critical weight be- the range specified in paragraph (a) of tween the maximum takeoff and max- this section. imum landing weights; (c) The average gradient of the stable (iv) 75 percent of maximum contin- slope of the stick force versus speed uous power for reciprocating engines or curve may not be less than 1 pound for for turbine engines, the maximum each 6 knots. cruising power selected by the appli- (d) Within the free return speed range cant as an operating limitation (see specified in paragraph (b) of this sec- § 25.1521), except that the power need tion, it is permissible for the airplane, not exceed that required at VMO/MMO; without control forces, to stabilize on and speeds above or below the desired trim (v) The airplane trimmed for level speeds if exceptional attention on the flight with the power required in para- part of the pilot is not required to re- graph (b)(1)(iv) of this section. turn to and maintain the desired trim (2) With the landing gear retracted at speed and altitude. low speed, the stick force curve must [Amdt. 25–7, 30 FR 13117, Oct. 15, 1965] have a stable slope at all speeds within a range which is the greater of 15 per- § 25.175 Demonstration of static longi- cent of the trim speed plus the result- tudinal stability. ing free return speed range, or 50 knots Static longitudinal stability must be plus the resulting free return speed shown as follows: range, above and below the trim speed (a) Climb. The stick force curve must (except that the speed range need not have a stable slope at speeds between include speeds less than 1.3 VSR1, nor 85 and 115 percent of the speed at which speeds greater than the minimum the airplane— speed of the applicable speed range pre- (1) Is trimmed, with— scribed in paragraph (b)(1), nor speeds (i) Wing flaps retracted; that require a stick force of more than (ii) Landing gear retracted; 50 pounds), with— (iii) Maximum takeoff weight; and (i) Wing flaps, center of gravity posi- (iv) 75 percent of maximum contin- tion, and weight as specified in para- uous power for reciprocating engines or graph (b)(1) of this section; the maximum power or thrust selected (ii) Power required for level flight at by the applicant as an operating limi- a speed equal to (VMO + 1.3 VSR1)/2; and 386

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(iii) The airplane trimmed for level § 25.177 Static lateral-directional sta- flight with the power required in para- bility. graph (b)(2)(ii) of this section. (a) The static directional stability (3) With the landing gear extended, (as shown by the tendency to recover the stick force curve must have a sta- from a skid with the rudder free) must ble slope at all speeds within a range be positive for any landing gear and which is the greater of 15 percent of the flap position and symmetric power con- trim speed plus the resulting free re- dition, at speeds from 1.13 VSR1, up to turn speed range, or 50 knots plus the VFE, VLE, or VFC/MFC (as appropriate for resulting free return speed range, the airplane configuration). above and below the trim speed (except (b) The static lateral stability (as that the speed range need not include shown by the tendency to raise the low speeds less than 1.3 VSR1, nor speeds wing in a sideslip with the aileron con- greater than VLE, nor speeds that re- trols free) for any landing gear and flap quire a stick force of more than 50 position and symmetric power condi- pounds), with— tion, may not be negative at any air- (i) Wing flap, center of gravity posi- speed (except that speeds higher than tion, and weight as specified in para- VFE need not be considered for flaps ex- graph (b)(1) of this section; tended configurations nor speeds high- (ii) 75 percent of maximum contin- er than VLE for landing gear extended uous power for reciprocating engines configurations) in the following airspeed ranges: or, for turbine engines, the maximum cruising power selected by the appli- (1) From 1.13 VSR1 to VMO/MMO. cant as an operating limitation, except (2) From VMO/MMO to VFC/MFC, unless that the power need not exceed that re- the divergence is— (i) Gradual; quired for level flight at VLE; and (iii) The aircraft trimmed for level (ii) Easily recognizable by the pilot; and flight with the power required in paragraph (b)(3)(ii) of this section. (iii) Easily controllable by the pilot. (c) The following requirement must (c) Approach. The stick force curve be met for the configurations and speed must have a stable slope at speeds be- specified in paragraph (a) of this sec- tween V and 1.7 V , with— SW SR1 tion. In straight, steady sideslips over (1) Wing flaps in the approach posi- the range of sideslip angles appropriate tion; to the operation of the airplane, the ai- (2) Landing gear retracted; leron and rudder control movements (3) Maximum landing weight; and and forces must be substantially pro- (4) The airplane trimmed at 1.3 VSR1 portional to the angle of sideslip in a with enough power to maintain level stable sense. This factor of proportion- flight at this speed. ality must lie between limits found (d) Landing. The stick force curve necessary for safe operation. The range must have a stable slope, and the stick of sideslip angles evaluated must in- force may not exceed 80 pounds, at clude those sideslip angles resulting speeds between VSW and 1.7 VSR0 with— from the lesser of: (1) Wing flaps in the landing position; (1) One-half of the available rudder (2) Landing gear extended; control input; and (3) Maximum landing weight; (2) A rudder control force of 180 pounds. (4) The airplane trimmed at 1.3 VSR0 with— (d) For sideslip angles greater than those prescribed by paragraph (c) of (i) Power or thrust off, and this section, up to the angle at which (ii) Power or thrust for level flight. full rudder control is used or a rudder (5) The airplane trimmed at 1.3 VSR0 control force of 180 pounds is obtained, with power or thrust off. the rudder control forces may not re- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as verse, and increased rudder deflection amended by Amdt. 25–7, 30 FR 13117, Oct. 15, must be needed for increased angles of 1965; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; sideslip. Compliance with this require- Amdt. 25–115, 69 FR 40527, July 2, 2004] ment must be shown using straight,

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steady sideslips, unless full lateral con- (2) Representative weights within the trol input is achieved before reaching range for which certification is re- either full rudder control input or a quested; rudder control force of 180 pounds; a (3) The most adverse center of grav- straight, steady sideslip need not be ity for recovery; and maintained after achieving full lateral (4) The airplane trimmed for straight control input. This requirement must flight at the speed prescribed in be met at all approved landing gear and § 25.103(b)(6). flap positions for the range of oper- (c) The following procedures must be ating speeds and power conditions ap- used to show compliance with § 25.203; propriate to each landing gear and flap position with all engines operating. (1) Starting at a speed sufficiently above the stalling speed to ensure that [Amdt. 25–135, 76 FR 74654, Dec. 1, 2011] a steady rate of speed reduction can be established, apply the longitudinal § 25.181 Dynamic stability. control so that the speed reduction (a) Any short period oscillation, not does not exceed one knot per second including combined lateral-directional until the airplane is stalled. oscillations, occurring between 1.13 VSR (2) In addition, for turning flight and maximum allowable speed appro- stalls, apply the longitudinal control priate to the configuration of the air- to achieve airspeed deceleration rates plane must be heavily damped with the up to 3 knots per second. primary controls— (3) As soon as the airplane is stalled, (1) Free; and recover by normal recovery techniques. (2) In a fixed position. (d) The airplane is considered stalled (b) Any combined lateral-directional oscillations (‘‘Dutch roll’’) occurring when the behavior of the airplane gives the pilot a clear and distinctive indica- between 1.13 VSR and maximum allowable speed appropriate to the configu- tion of an acceptable nature that the ration of the airplane must be posi- airplane is stalled. Acceptable indica- tively damped with controls free, and tions of a stall, occurring either indi- must be controllable with normal use vidually or in combination, are— of the primary controls without requir- (1) A nose-down pitch that cannot be ing exceptional pilot skill. readily arrested; (2) Buffeting, of a magnitude and se- [Amdt. 25–42, 43 FR 2322, Jan. 16, 1978, as amended by Amdt. 25–72, 55 FR 29775, July 20, verity that is a strong and effective de- 1990; 55 FR 37607, Sept. 12, 1990; Amdt. 25–108, terrent to further speed reduction; or 67 FR 70827, Nov. 26, 2002] (3) The pitch control reaches the aft stop and no further increase in pitch STALLS attitude occurs when the control is held full aft for a short time before re- § 25.201 Stall demonstration. covery is initiated. (a) Stalls must be shown in straight flight and in 30 degree banked turns [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as with— amended by Amdt. 25–84, 60 FR 30750, June 9, 1995; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002] (1) Power off; and (2) The power necessary to maintain § 25.203 Stall characteristics. level flight at 1.5 VSR1 (where VSR1 corresponds to the reference stall speed at (a) It must be possible to produce and maximum landing weight with flaps in to correct roll and yaw by unreversed the approach position and the landing use of the aileron and rudder controls, gear retracted). up to the time the airplane is stalled. (b) In each condition required by No abnormal nose-up pitching may paragraph (a) of this section, it must occur. The longitudinal control force be possible to meet the applicable re- must be positive up to and throughout quirements of § 25.203 with— the stall. In addition, it must be pos- (1) Flaps, landing gear, and decelera- sible to promptly prevent stalling and tion devices in any likely combination to recover from a stall by normal use of positions approved for operation; of the controls.

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(b) For level wing stalls, the roll oc- identified in accordance with § 25.201(d) curring between the stall and the com- by not less than five knots or five per- pletion of the recovery may not exceed cent CAS, whichever is greater. Once approximately 20 degrees. initiated, stall warning must continue (c) For turning flight stalls, the ac- until the angle of attack is reduced to tion of the airplane after the stall may approximately that at which stall not be so violent or extreme as to warning began. make it difficult, with normal piloting (d) In addition to the requirement of skill, to effect a prompt recovery and paragraph (c) of this section, when the to regain control of the airplane. The speed is reduced at rates not exceeding maximum bank angle that occurs dur- one knot per second, in straight flight ing the recovery may not exceed— with engines idling and at the center- (1) Approximately 60 degrees in the of-gravity position specified in original direction of the turn, or 30 de- § 25.103(b)(5), VSW, in each normal con- grees in the opposite direction, for de- figuration, must exceed VSR by not less celeration rates up to 1 knot per sec- than three knots or three percent CAS, ond; and whichever is greater. (2) Approximately 90 degrees in the (e) In icing conditions, the stall original direction of the turn, or 60 de- warning margin in straight and turn- grees in the opposite direction, for de- ing flight must be sufficient to allow celeration rates in excess of 1 knot per the pilot to prevent stalling (as defined second. in § 25.201(d)) when the pilot starts a re- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as covery maneuver not less than three amended by Amdt. 25–84, 60 FR 30750, June 9, seconds after the onset of stall warn- 1995] ing. When demonstrating compliance with this paragraph, the pilot must § 25.207 Stall warning. perform the recovery maneuver in the (a) Stall warning with sufficient mar- same way as for the airplane in non- gin to prevent inadvertent stalling icing conditions. Compliance with this with the flaps and landing gear in any requirement must be demonstrated in normal position must be clear and dis- flight with the speed reduced at rates tinctive to the pilot in straight and not exceeding one knot per second, turning flight. with— (b) The warning must be furnished ei- (1) The more critical of the takeoff ther through the inherent aerodynamic ice and final takeoff ice accretions de- qualities of the airplane or by a device fined in appendix C for each configura- that will give clearly distinguishable tion used in the takeoff phase of flight; indications under expected conditions (2) The en route ice accretion defined of flight. However, a visual stall warn- in appendix C for the en route configu- ing device that requires the attention ration; of the crew within the cockpit is not (3) The holding ice accretion defined acceptable by itself. If a warning de- in appendix C for the holding configu- vice is used, it must provide a warning ration(s); in each of the airplane configurations (4) The approach ice accretion de- prescribed in paragraph (a) of this sec- fined in appendix C for the approach tion at the speed prescribed in para- configuration(s); and graphs (c) and (d) of this section. Ex- (5) The landing ice accretion defined cept for showing compliance with the in appendix C for the landing and go- stall warning margin prescribed in around configuration(s). paragraph (h)(3)(ii) of this section, stall (f) The stall warning margin must be warning for flight in icing conditions sufficient in both non-icing and icing must be provided by the same means as conditions to allow the pilot to prevent stall warning for flight in non-icing stalling when the pilot starts a recov- conditions. ery maneuver not less than one second (c) When the speed is reduced at rates after the onset of stall warning in slow- not exceeding one knot per second, down turns with at least 1.5 g load fac- stall warning must begin, in each nor- tor normal to the flight path and air- mal configuration, at a speed, VSW, ex- speed deceleration rates of at least 2 ceeding the speed at which the stall is knots per second. When demonstrating

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compliance with this paragraph for celeration rates of § 25.201(c)(2) need not icing conditions, the pilot must per- be demonstrated. form the recovery maneuver in the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as same way as for the airplane in non- amended by Amdt. 25–7, 30 FR 13118, Oct. 15, icing conditions. Compliance with this 1965; Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; requirement must be demonstrated in Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; flight with— Amdt. 25–121, 72 FR 44668, Aug. 8, 2007; Amdt. (1) The flaps and landing gear in any 25–129, 74 FR 38339, Aug. 3, 2009] normal position; GROUND AND WATER HANDLING (2) The airplane trimmed for straight CHARACTERISTICS flight at a speed of 1.3 VSR; and (3) The power or thrust necessary to § 25.231 Longitudinal stability and control. maintain level flight at 1.3 VSR. (g) Stall warning must also be pro- (a) Landplanes may have no uncon- vided in each abnormal configuration trollable tendency to nose over in any of the high lift devices that is likely to reasonably expected operating condi- be used in flight following system fail- tion or when rebound occurs during ures (including all configurations cov- landing or takeoff. In addition— ered by Airplane Flight Manual proce- (1) Wheel brakes must operate dures). smoothly and may not cause any undue (h) For flight in icing conditions be- tendency to nose over; and fore the ice protection system has been (2) If a tail-wheel landing gear is used, it must be possible, during the activated and is performing its in- takeoff ground run on concrete, to tended function, with the ice accretion maintain any attitude up to thrust line defined in appendix C, part II(e) of this level, at 75 percent of VSR1. part, the stall warning margin in (b) For seaplanes and amphibians, straight and turning flight must be suf- the most adverse water conditions safe ficient to allow the pilot to prevent for takeoff, taxiing, and landing, must stalling without encountering any ad- be established. verse flight characteristics when: (1) The speed is reduced at rates not [Docket No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–108, 67 FR 70828, Nov. exceeding one knot per second; 26, 2002] (2) The pilot performs the recovery maneuver in the same way as for flight § 25.233 Directional stability and con- in non-icing conditions; and trol. (3) The recovery maneuver is started (a) There may be no uncontrollable no earlier than: ground-looping tendency in 90° cross (i) One second after the onset of stall winds, up to a wind velocity of 20 knots warning if stall warning is provided by or 0.2 VSR0, whichever is greater, except the same means as for flight in non- that the wind velocity need not exceed icing conditions; or 25 knots at any speed at which the air- (ii) Three seconds after the onset of plane may be expected to be operated stall warning if stall warning is pro- on the ground. This may be shown ° vided by a different means than for while establishing the 90 cross compo- flight in non-icing conditions. nent of wind velocity required by § 25.237. (i) In showing compliance with para- (b) Landplanes must be satisfactorily graph (h) of this section, if stall warn- controllable, without exceptional piloting is provided by a different means in ing skill or alertness, in power-off land- icing conditions than for non-icing con- ings at normal landing speed, without ditions, compliance with § 25.203 must using brakes or engine power to main- be shown using the accretion defined in tain a straight path. This may be appendix C, part II(e) of this part. Com- shown during power-off landings made pliance with this requirement must be in conjunction with other tests. shown using the demonstration pre- (c) The airplane must have adequate scribed by § 25.201, except that the de- directional control during taxiing. This may be shown during taxiing prior to

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takeoffs made in conjunction with and in the conditions set forth in para- other tests. graph (b) of this section, there may be [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as no— amended by Amdt. 25–23, 35 FR 5671, Apr. 8, (1) Spray characteristics that would 1970; Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; impair the pilot’s view, cause damage, Amdt. 25–94, 63 FR 8848, Feb. 23, 1998; Amdt. or result in the taking in of an undue 25–108, 67 FR 70828, Nov. 26, 2002] quantity of water; (2) Dangerously uncontrollable § 25.235 Taxiing condition. porpoising, bounding, or swinging tend- The shock absorbing mechanism may ency; or not damage the structure of the air- (3) Immersion of auxiliary floats or plane when the airplane is taxied on sponsons, wing tips, propeller blades, the roughest ground that may reason- or other parts not designed to with- ably be expected in normal operation. stand the resulting water loads. § 25.237 Wind velocities. (b) Compliance with the requirements of paragraph (a) of this section (a) For land planes and amphibians, must be shown— the following applies: (1) A 90-degree cross component of (1) In water conditions, from smooth wind velocity, demonstrated to be safe to the most adverse condition estab- for takeoff and landing, must be established in accordance with § 25.231; lished for dry runways and must be at (2) In wind and cross-wind velocities, water currents, and associated waves least 20 knots or 0.2 VSR0, whichever is greater, except that it need not exceed and swells that may reasonably be ex- 25 knots. pected in operation on water; (2) The crosswind component for (3) At speeds that may reasonably be takeoff established without ice accre- expected in operation on water; tions is valid in icing conditions. (4) With sudden failure of the critical (3) The landing crosswind component engine at any time while on water; and must be established for: (5) At each weight and center of grav- (i) Non-icing conditions, and ity position, relevant to each operating (ii) Icing conditions with the landing condition, within the range of loading ice accretion defined in appendix C. conditions for which certification is re- (b) For seaplanes and amphibians, quested. the following applies: (c) In the water conditions of para- (1) A 90-degree cross component of graph (b) of this section, and in the wind velocity, up to which takeoff and corresponding wind conditions, the sea- landing is safe under all water condi- plane or amphibian must be able to tions that may reasonably be expected drift for five minutes with engines in- in normal operation, must be established and must be at least 20 knots or operative, aided, if necessary, by a sea anchor. 0.2 VSR0, whichever is greater, except that it need not exceed 25 knots. MISCELLANEOUS FLIGHT REQUIREMENTS (2) A wind velocity, for which taxiing is safe in any direction under all water § 25.251 Vibration and buffeting. conditions that may reasonably be expected in normal operation, must be es- (a) The airplane must be dem- tablished and must be at least 20 knots onstrated in flight to be free from any or 0.2 VSR0, whichever is greater, except vibration and buffeting that would pre- that it need not exceed 25 knots. vent continued safe flight in any likely operating condition. [Amdt. 25–42, 43 FR 2322, Jan. 16, 1978, as amended by Amdt. 25–108, 67 FR 70827, Nov. (b) Each part of the airplane must be 26, 2002; Amdt. 25–121, 72 FR 44668, Aug. 8, demonstrated in flight to be free from 2007] excessive vibration under any appropriate speed and power conditions up to § 25.239 Spray characteristics, control, VDF/MDF. The maximum speeds shown and stability on water. must be used in establishing the oper- (a) For seaplanes and amphibians, ating limitations of the airplane in ac- during takeoff, taxiing, and landing, cordance with § 25.1505.

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(c) Except as provided in paragraph normal attitude and its speed reduced (d) of this section, there may be no buf- to VMO/MMO, without— feting condition, in normal flight, in- (i) Exceptional piloting strength or cluding configuration changes during skill;

cruise, severe enough to interfere with (ii) Exceeding VD/MD, VDF/MDF, or the the control of the airplane, to cause ex- structural limitations; and cessive fatigue to the crew, or to cause (iii) Buffeting that would impair the structural damage. Stall warning buf- pilot’s ability to read the instruments feting within these limits is allowable. or control the airplane for recovery. (d) There may be no perceptible buf- (3) With the airplane trimmed at any feting condition in the cruise configu- speed up to VMO/MMO, there must be no ration in straight flight at any speed reversal of the response to control up to VMO/MMO, except that stall warn- input about any axis at any speed up to ing buffeting is allowable. VDF/MDF. Any tendency to pitch, roll, or (e) For an airplane with MD greater yaw must be mild and readily control- than .6 or with a maximum operating lable, using normal piloting tech- altitude greater than 25,000 feet, the niques. When the airplane is trimmed positive maneuvering load factors at at VMO/MMO, the slope of the elevator which the onset of perceptible buf- control force versus speed curve need feting occurs must be determined with not be stable at speeds greater than the airplane in the cruise configuration V /M , but there must be a push force for the ranges of airspeed or Mach FC FC at all speeds up to V /M and there number, weight, and altitude for which DF DF must be no sudden or excessive reduc- the airplane is to be certificated. The tion of elevator control force as V / envelopes of load factor, speed, alti- DF M is reached. tude, and weight must provide a suffi- DF cient range of speeds and load factors (4) Adequate roll capability to assure for normal operations. Probable inad- a prompt recovery from a lateral upset vertent excursions beyond the bound- condition must be available at any aries of the buffet onset envelopes may speed up to VDF/MDF. not result in unsafe conditions. (5) With the airplane trimmed at VMO/MMO, extension of the speedbrakes [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as over the available range of movements amended by Amdt. 25–23, 35 FR 5671, Apr. 8, of the pilot’s control, at all speeds 1970; Amdt. 25–72, 55 FR 29775, July 20, 1990; Amdt. 25–77, 57 FR 28949, June 29, 1992] above VMO/MMO, but not so high that VDF/MDF would be exceeded during the § 25.253 High-speed characteristics. maneuver, must not result in: (i) An excessive positive load factor (a) Speed increase and recovery charac- when the pilot does not take action to teristics. The following speed increase and recovery characteristics must be counteract the effects of extension; met: (ii) Buffeting that would impair the (1) Operating conditions and charac- pilot’s ability to read the instruments teristics likely to cause inadvertent or control the airplane for recovery; or speed increases (including upsets in (iii) A nose down pitching moment, pitch and roll) must be simulated with unless it is small. the airplane trimmed at any likely (b) Maximum speed for stability charac- cruise speed up to VMO/MMO. These con- teristics, VFC/MFC. VFC/MFC is the max- ditions and characteristics include gust imum speed at which the requirements upsets, inadvertent control move- of §§ 25.143(g), 25.147(f), 25.175(b)(1), ments, low stick force gradient in rela- 25.177(a) through (c), and 25.181 must be tion to control friction, passenger met with flaps and landing gear re- movement, leveling off from climb, and tracted. Except as noted in § 25.253(c), descent from Mach to airspeed limit al- VFC/MFC may not be less than a speed titudes. midway between VMO/MMO and VDF/MDF, (2) Allowing for pilot reaction time except that, for altitudes where Mach after effective inherent or artificial number is the limiting factor, MFC need speed warning occurs, it must be shown not exceed the Mach number at which that the airplane can be recovered to a effective speed warning occurs.

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(c) Maximum speed for stability charac- (c) Except as provided in paragraphs teristics in icing conditions. The max- (d) and (e) of this section, compliance imum speed for stability characteris- with the provisions of paragraph (a) of tics with the ice accretions defined in this section must be demonstrated in appendix C, at which the requirements flight over the acceleration range— of §§ 25.143(g), 25.147(f), 25.175(b)(1), (1) ¥1 g to +2.5 g; or 25.177(a) through (c), and 25.181 must be (2) 0 g to 2.0 g, and extrapolating by met, is the lower of: an acceptable method to ¥1 g and +2.5 (1) 300 knots CAS; g. (d) If the procedure set forth in para- (2) VFC; or (3) A speed at which it is dem- graph (c)(2) of this section is used to onstrated that the airframe will be free demonstrate compliance and marginal of ice accretion due to the effects of in- conditions exist during flight test with creased dynamic pressure. regard to reversal of primary longitudinal control force, flight tests must be [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as accomplished from the normal accel- amended by Amdt. 25–23, 35 FR 5671, Apr. 8, eration at which a marginal condition 1970; Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; is found to exist to the applicable limit Amdt. 25–72, 55 FR 29775, July 20, 1990; Amdt. specified in paragraph (b)(1) of this sec- 25–84, 60 FR 30750, June 9, 1995; Amdt. 25–121, 72 FR 44668, Aug. 8, 2007; Amdt. 25–135, 76 FR tion. 74654, Dec. 1, 2011] (e) During flight tests required by paragraph (a) of this section, the limit § 25.255 Out-of-trim characteristics. maneuvering load factors prescribed in §§ 25.333(b) and 25.337, and the maneu- (a) From an initial condition with vering load factors associated with the airplane trimmed at cruise speeds probable inadvertent excursions be- up to VMO/MMO, the airplane must have yond the boundaries of the buffet onset satisfactory maneuvering stability and envelopes determined under § 25.251(e), controllability with the degree of out- need not be exceeded. In addition, the of-trim in both the airplane nose-up entry speeds for flight test demonstra- and nose-down directions, which re- tions at normal acceleration values sults from the greater of— less than 1 g must be limited to the ex- (1) A three-second movement of the tent necessary to accomplish a recov- longitudinal trim system at its normal ery without exceeding VDF/MDF. rate for the particular flight condition (f) In the out-of-trim condition speci- with no aerodynamic load (or an equiv- fied in paragraph (a) of this section, it alent degree of trim for airplanes that must be possible from an overspeed do not have a power-operated trim sys- condition at VDF/MDF to produce at tem), except as limited by stops in the least 1.5 g for recovery by applying not trim system, including those required more than 125 pounds of longitudinal by § 25.655(b) for adjustable stabilizers; control force using either the primary or longitudinal control alone or the pri- (2) The maximum mistrim that can mary longitudinal control and the lon- be sustained by the autopilot while gitudinal trim system. If the longitu- maintaining level flight in the high dinal trim is used to assist in pro- speed cruising condition. ducing the required load factor, it must (b) In the out-of-trim condition speci- be shown at VDF/MDF that the longitu- fied in paragraph (a) of this section, dinal trim can be actuated in the air- when the normal acceleration is varied plane nose-up direction with the pri- from +1 g to the positive and negative mary surface loaded to correspond to values specified in paragraph (c) of this the least of the following airplane section— nose-up control forces: (1) The stick force vs. g curve must (1) The maximum control forces ex- have a positive slope at any speed up to pected in service as specified in §§ 25.301 and including VFC/MFC; and and 25.397. (2) At speeds between VFC/MFC and (2) The control force required to VDF/MDF the direction of the primary produce 1.5 g. longitudinal control force may not re- (3) The control force corresponding to verse. buffeting or other phenomena of such

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intensity that it is a strong deterrent proof of strength is shown by dynamic to further application of primary longi- tests simulating actual load condi- tudinal control force. tions, the 3-second limit does not [Amdt. 25–42, 43 FR 2322, Jan. 16, 1978] apply. Static tests conducted to ultimate load must include the ultimate deflections and ultimate deformation Subpart C—Structure induced by the loading. When analyt- GENERAL ical methods are used to show compliance with the ultimate load strength § 25.301 Loads. requirements, it must be shown that— (a) Strength requirements are speci- (1) The effects of deformation are not fied in terms of limit loads (the max- significant; imum loads to be expected in service) (2) The deformations involved are and ultimate loads (limit loads multi- fully accounted for in the analysis; or plied by prescribed factors of safety). (3) The methods and assumptions Unless otherwise provided, prescribed used are sufficient to cover the effects loads are limit loads. of these deformations. (b) Unless otherwise provided, the (c) Where structural flexibility is specified air, ground, and water loads such that any rate of load application must be placed in equilibrium with in- likely to occur in the operating condi- ertia forces, considering each item of tions might produce transient stresses mass in the airplane. These loads must appreciably higher than those cor- be distributed to conservatively ap- responding to static loads, the effects proximate or closely represent actual of this rate of application must be con- conditions. Methods used to determine sidered. load intensities and distribution must (d) [Reserved] be validated by flight load measure- (e) The airplane must be designed to ment unless the methods used for de- withstand any vibration and buffeting termining those loading conditions are that might occur in any likely oper- shown to be reliable. ating condition up to VD/MD, including (c) If deflections under load would stall and probable inadvertent excur- significantly change the distribution of sions beyond the boundaries of the buf- external or internal loads, this redis- fet onset envelope. This must be shown tribution must be taken into account. by analysis, flight tests, or other tests [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as found necessary by the Administrator. amended by Amdt. 25–23, 35 FR 5672, Apr. 8, (f) Unless shown to be extremely im- 1970] probable, the airplane must be designed to withstand any forced structural vi- § 25.303 Factor of safety. bration resulting from any failure, Unless otherwise specified, a factor of malfunction or adverse condition in safety of 1.5 must be applied to the pre- the flight control system. These must scribed limit load which are considered be considered limit loads and must be external loads on the structure. When a investigated at airspeeds up to VC/MC. loading condition is prescribed in terms of ultimate loads, a factor of [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5672, Apr. 8, safety need not be applied unless other- 1970; Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; wise specified. Amdt. 25–77, 57 FR 28949, June 29, 1992; Amdt. [Amdt. 25–23, 35 FR 5672, Apr. 8, 1970] 25–86, 61 FR 5220, Feb. 9, 1996]

§ 25.305 Strength and deformation. § 25.307 Proof of structure. (a) The structure must be able to (a) Compliance with the strength and support limit loads without detri- deformation requirements of this sub- mental permanent deformation. At any part must be shown for each critical load up to limit loads, the deformation loading condition. Structural analysis may not interfere with safe operation. may be used only if the structure con- (b) The structure must be able to forms to that for which experience has support ultimate loads without failure shown this method to be reliable. The for at least 3 seconds. However, when Administrator may require ultimate

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load tests in cases where limit load forces must be considered in equi- tests may be inadequate. librium with thrust and all aero- (b)–(c) [Reserved] dynamic moments, including moments (d) When static or dynamic tests are due to loads on components such as used to show compliance with the re- tail surfaces and nacelles. Critical quirements of § 25.305(b) for flight thrust values in the range from zero to structures, appropriate material cor- maximum continuous thrust must be rection factors must be applied to the considered. test results, unless the structure, or part thereof, being tested has features [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as such that a number of elements con- amended by Amdt. 25–23, 35 FR 5672, Apr. 8, tribute to the total strength of the 1970; Amdt. 25–86, 61 FR 5220, Feb. 9, 1996] structure and the failure of one ele- FLIGHT MANEUVER AND GUST ment results in the redistribution of CONDITIONS the load through alternate load paths. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.331 Symmetric maneuvering con- amended by Amdt. 25–23, 35 FR 5672, Apr. 8, ditions. 1970; Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; Amdt. 25–72, 55 FR 29775, July 20, 1990] (a) Procedure. For the analysis of the maneuvering flight conditions specified FLIGHT LOADS in paragraphs (b) and (c) of this section, the following provisions apply: § 25.321 General. (1) Where sudden displacement of a (a) Flight load factors represent the control is specified, the assumed rate ratio of the aerodynamic force compo- of control surface displacement may nent (acting normal to the assumed not be less than the rate that could be longitudinal axis of the airplane) to the applied by the pilot through the con- weight of the airplane. A positive load trol system. factor is one in which the aerodynamic (2) In determining elevator angles force acts upward with respect to the and chordwise load distribution in the airplane. maneuvering conditions of paragraphs (b) Considering compressibility ef- (b) and (c) of this section, the effect of fects at each speed, compliance with corresponding pitching velocities must the flight load requirements of this be taken into account. The in-trim and subpart must be shown— out-of-trim flight conditions specified (1) At each critical altitude within in § 25.255 must be considered. the range of altitudes selected by the (b) Maneuvering balanced conditions. applicant; Assuming the airplane to be in equi- (2) At each weight from the design librium with zero pitching accelera- minimum weight to the design max- tion, the maneuvering conditions A imum weight appropriate to each par- through I on the maneuvering envelope ticular flight load condition; and in § 25.333(b) must be investigated. (3) For each required altitude and weight, for any practicable distribution (c) Pitch maneuver conditions. The of disposable load within the operating conditions specified in paragraphs limitations recorded in the Airplane (c)(1) and (2) of this section must be in- Flight Manual. vestigated. The movement of the pitch (c) Enough points on and within the control surfaces may be adjusted to boundaries of the design envelope must take into account limitations imposed be investigated to ensure that the max- by the maximum pilot effort specified imum load for each part of the airplane by § 25.397(b), control system stops and structure is obtained. any indirect effect imposed by limita- (d) The significant forces acting on tions in the output side of the control the airplane must be placed in equi- system (for example, stalling torque or librium in a rational or conservative maximum rate obtainable by a power manner. The linear inertia forces must control system.) be considered in equilibrium with the (1) Maximum pitch control displacement thrust and all aerodynamic loads, at VA. The airplane is assumed to be while the angular (pitching) inertia flying in steady level flight (point A1, 395

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§ 25.333(b)) and the cockpit pitch con- (ii) A negative pitching acceleration trol is suddenly moved to obtain ex- (nose down) is assumed to be reached treme nose up pitching acceleration. In concurrently with the positive maneu- defining the tail load, the response of vering load factor (points A2 to D2, the airplane must be taken into ac- § 25.333(b)). This negative pitching account. Airplane loads that occur subse- celeration must be equal to at least quent to the time when normal acceleration at the c.g. exceeds the positive −26n limit maneuvering load factor (at point ()n− 1.5 ,() Radians/sec.2 A2 in § 25.333(b)), or the resulting v tailplane normal load reaches its max- where— imum, whichever occurs first, need not n is the positive load factor at the speed be considered. under consideration; and V is the air- (2) Specified control displacement. A plane equivalent speed in knots. checked maneuver, based on a rational pitching control motion vs. time pro- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5672, Apr. 8, file, must be established in which the 1970; Amdt. 25–46, 43 FR 50594, Oct. 30, 1978; 43 design limit load factor specified in FR 52495, Nov. 13, 1978; 43 FR 54082, Nov. 20, § 25.337 will not be exceeded. Unless 1978; Amdt. 25–72, 55 FR 29775, July 20, 1990; 55 lesser values cannot be exceeded, the FR 37607, Sept. 12, 1990; Amdt. 25–86, 61 FR airplane response must result in pitch- 5220, Feb. 9, 1996; Amdt. 25–91, 62 FR 40704, ing accelerations not less than the fol- July 29, 1997] lowing: (i) A positive pitching acceleration § 25.333 Flight maneuvering envelope. (nose up) is assumed to be reached con- (a) General. The strength require- currently with the airplane load factor ments must be met at each combina- of 1.0 (Points A1 to D1, § 25.333(b)). The positive acceleration must be equal to tion of airspeed and load factor on and at least within the boundaries of the representative maneuvering envelope (V-n dia- gram) of paragraph (b) of this section. 39n 2 ()n− 1.5 ,() Radians/sec. This envelope must also be used in de- v termining the airplane structural oper- where— ating limitations as specified in n is the positive load factor at the speed § 25.1501. under consideration, and V is the airplane equivalent speed in knots. (b) Maneuvering envelope.

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[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–86, 61 FR 5220, Feb. 9, 1996]

§ 25.335 Design airspeeds. (1) From an initial condition of stabilized flight at V /M the airplane is The selected design airspeeds are C C, upset, flown for 20 seconds along a equivalent airspeeds (EAS). Estimated flight path 7.5° below the initial path, values of V and V must be conserv- S0 S1 and then pulled up at a load factor of ative. 1.5g (0.5g acceleration increment). The (a) Design cruising speed, V For V C. C, speed increase occurring in this maneu- the following apply: ver may be calculated if reliable or (1) The minimum value of V must be C conservative aerodynamic data is used. sufficiently greater than V to provide B Power as specified in § 25.175(b)(1)(iv) is for inadvertent speed increases likely assumed until the pullup is initiated, to occur as a result of severe atmospheric turbulence. at which time power reduction and the use of pilot controlled drag devices (2) Except as provided in § 25.335(d)(2), may be assumed; VC may not be less than VB + 1.32 U REF (with U as specified in (2) The minimum speed margin must REF be enough to provide for atmospheric § 25.341(a)(5)(i)). However VC need not exceed the maximum speed in level variations (such as horizontal gusts, flight at maximum continuous power and penetration of jet streams and cold for the corresponding altitude. fronts) and for instrument errors and airframe production variations. These (3) At altitudes where VD is limited factors may be considered on a prob- by Mach number, VC may be limited to a selected Mach number. ability basis. The margin at altitude where MC is limited by compressibility (b) Design dive speed, VD. VD must be effects must not less than 0.07M unless selected so that VC/MC is not greater a lower margin is determined using a than 0.8 VD/MD, or so that the minimum rational analysis that includes the ef- speed margin between VC/MC and VD/MD is the greater of the following values: fects of any automatic systems. In any

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case, the margin may not be reduced to greater than the operating speed rec- less than 0.05M. ommended for the corresponding stage (c) Design maneuvering speed VA. For of flight (including balked landings) to VA, the following apply: allow for probable variations in control (1) VA may not be less than VS1 √n of airspeed and for transition from one where— flap position to another. (i) n is the limit positive maneu- (2) If an automatic flap positioning or vering load factor at VC; and load limiting device is used, the speeds (ii) VS1 is the stalling speed with flaps and corresponding flap positions pro- retracted. grammed or allowed by the device may (2) VA and VS must be evaluated at be used. the design weight and altitude under (3) VF may not be less than— consideration. (i) 1.6 VS1 with the flaps in takeoff po- (3) VA need not be more than VC or sition at maximum takeoff weight; the speed at which the positive C N max (ii) 1.8 V with the flaps in approach curve intersects the positive maneuver S1 position at maximum landing weight, load factor line, whichever is less. and (d) Design speed for maximum gust intensity, V . (iii) 1.8 VS0 with the flaps in landing B position at maximum landing weight. (1) VB may not be less than (f) Design drag device speeds, VDD. The selected design speed for each drag de- ⎡ KU Va⎤12 + g ref c vice must be sufficiently greater than VS1 ⎢1 ⎥ the speed recommended for the oper- ⎣ 498w ⎦ ation of the device to allow for prob- where— able variations in speed control. For VS1=the 1-g stalling speed based on CNAmax drag devices intended for use in high with the flaps retracted at the particular speed descents, VDD may not be less weight under consideration; than VD. When an automatic drag de- Vc=design cruise speed (knots equivalent air- vice positioning or load limiting means speed); is used, the speeds and corresponding Uref=the reference gust velocity (feet per second equivalent airspeed) from drag device positions programmed or § 25.341(a)(5)(i); allowed by the automatic means must w=average wing loading (pounds per square be used for design. foot) at the particular weight under consideration. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5672, Apr. 8, .88μ 1970; Amdt. 25–86, 61 FR 5220, Feb. 9, 1996; = Amdt. 25–91, 62 FR 40704, July 29, 1997] Kg 53. + μ § 25.337 Limit maneuvering load fac- 2w tors. μ = (a) Except where limited by max- ρcag imum (static) lift coefficients, the air- r=density of air (slugs/ft3); plane is assumed to be subjected to c=mean geometric chord of the wing (feet); symmetrical maneuvers resulting in g=acceleration due to gravity (ft/sec2); the limit maneuvering load factors pre- a=slope of the airplane normal force coeffi- scribed in this section. Pitching veloci- cient curve, CNA per radian; ties appropriate to the corresponding pull-up and steady turn maneuvers (2) At altitudes where VC is limited by Mach number— must be taken into account. (i) VB may be chosen to provide an (b) The positive limit maneuvering optimum margin between low and high load factor n for any speed up to Vn speed buffet boundaries; and, may not be less than 2.1+24,000/ (W (ii) VB need not be greater than VC. +10,000) except that n may not be less (e) Design flap speeds, VF. For VF, the than 2.5 and need not be greater than following apply: 3.8—where W is the design maximum (1) The design flap speed for each flap takeoff weight. position (established in accordance (c) The negative limit maneuvering with § 25.697(a)) must be sufficiently load factor—

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(1) May not be less than ¥1.0 at (i) At the airplane design speed VC: speeds up to VC; and Positive and negative gusts with ref- (2) Must vary linearly with speed erence gust velocities of 56.0 ft/sec EAS from the value at VC to zero at VD. must be considered at sea level. The (d) Maneuvering load factors lower reference gust velocity may be reduced than those specified in this section linearly from 56.0 ft/sec EAS at sea may be used if the airplane has design level to 44.0 ft/sec EAS at 15000 feet. features that make it impossible to ex- The reference gust velocity may be fur- ceed these values in flight. ther reduced linearly from 44.0 ft/sec [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as EAS at 15000 feet to 26.0 ft/sec EAS at amended by Amdt. 25–23, 35 FR 5672, Apr. 8, 50000 feet. 1970] (ii) At the airplane design speed VD: § 25.341 Gust and turbulence loads. The reference gust velocity must be 0.5 times the value obtained under (a) Discrete Gust Design Criteria. The § 25.341(a)(5)(i). airplane is assumed to be subjected to symmetrical vertical and lateral gusts (6) The flight profile alleviation fac- in level flight. Limit gust loads must tor, Fg, must be increased linearly from be determined in accordance with the the sea level value to a value of 1.0 at provisions: the maximum operating altitude de- (1) Loads on each part of the struc- fined in § 25.1527. At sea level, the flight ture must be determined by dynamic profile alleviation factor is determined analysis. The analysis must take into by the following equation: account unsteady aerodynamic characteristics and all significant structural FFF=+05. () degrees of freedom including rigid body ggzgm motions. (2) The shape of the gust must be: Where: ⎡ ⎛ π ⎞ ⎤ Z Uds s =− mo U = ⎢1- Cos⎜ ⎟ ⎥ Fgz 1 ; 2 ⎣ ⎝ H ⎠ ⎦ 250000 ≤ ≤ for 0 s 2H ⎛πR ⎞ where— = 1 Fgm R2 Tan ; s=distance penetrated into the gust (feet); ⎝ 4 ⎠ Uds=the design gust velocity in equivalent airspeed specified in paragraph (a)(4) of Maximum Landing Weight this section; and R = ; H=the gust gradient which is the distance 1 (feet) parallel to the airplane’s flight Maximum Take- off Weight path for the gust to reach its peak velocity. = Maximum Zero Fuel Weight R2 ; (3) A sufficient number of gust gra- Maximum Take- off Weight dient distances in the range 30 feet to 350 feet must be investigated to find Zmo=Maximum operating altitude defined in § 25.1527. the critical response for each load quantity. (7) When a stability augmentation (4) The design gust velocity must be: system is included in the analysis, the effect of any significant system non- 16 UUF= ()H linearities should be accounted for ds ref g 350 when deriving limit loads from limit where— gust conditions.

Uref=the reference gust velocity in equivalent (b) Continuous Gust Design Criteria. airspeed defined in paragraph (a)(5) of The dynamic response of the airplane this section. to vertical and lateral continuous tur- Fg=the flight profile alleviation factor defined in paragraph (a)(6) of this section. bulence must be taken into account. The continuous gust design criteria of (5) The following reference gust ve- appendix G of this part must be used to locities apply:

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establish the dynamic response unless be determined by rational analysis. more rational criteria are shown. The analysis must take into account the unsteady aerodynamic characteris- [Doc. No. 27902, 61 FR 5221, Feb. 9, 1996; 61 FR 9533, Mar. 8, 1996] tics and rigid body motions of the aircraft. The shape of the gust must be as § 25.343 Design fuel and oil loads. described in § 25.341(a)(2) except that—

(a) The disposable load combinations Uds=25 ft/sec EAS; must include each fuel and oil load in H=12.5 c; and the range from zero fuel and oil to the c=mean geometric chord of the wing (feet). selected maximum fuel and oil load. A (b) The airplane must be designed for structural reserve fuel condition, not the conditions prescribed in paragraph exceeding 45 minutes of fuel under the (a) of this section, except that the air- operating conditions in § 25.1001(e) and plane load factor need not exceed 1.0, (f), as applicable, may be selected. taking into account, as separate condi- (b) If a structural reserve fuel conditions, the effects of— tion is selected, it must be used as the (1) Propeller slipstream cor- minimum fuel weight condition for showing compliance with the flight responding to maximum continuous load requirements as prescribed in this power at the design flap speeds VF, and subpart. In addition— with takeoff power at not less than 1.4 (1) The structure must be designed times the stalling speed for the par- for a condition of zero fuel and oil in ticular flap position and associated the wing at limit loads corresponding maximum weight; and to— (2) A head-on gust of 25 feet per sec- (i) A maneuvering load factor of ond velocity (EAS). +2.25; and (c) If flaps or other high lift devices (ii) The gust conditions of § 25.341(a) are to be used in en route conditions, but assuming 85% of the design veloci- and with flaps in the appropriate posi- ties prescribed in § 25.341(a)(4). tion at speeds up to the flap design (2) Fatigue evaluation of the struc- speed chosen for these conditions, the ture must account for any increase in airplane is assumed to be subjected to operating stresses resulting from the symmetrical maneuvers and gusts design condition of paragraph (b)(1) of within the range determined by— this section; and (1) Maneuvering to a positive limit (3) The flutter, deformation, and vi- load factor as prescribed in § 25.337(b); bration requirements must also be met and with zero fuel. (2) The discrete vertical gust criteria in § 25.341(a). [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–18, 33 FR 12226, Aug. 30, (d) The airplane must be designed for 1968; Amdt. 25–72, 55 FR 37607, Sept. 12, 1990; a maneuvering load factor of 1.5 g at Amdt. 25–86, 61 FR 5221, Feb. 9, 1996] the maximum take-off weight with the wing-flaps and similar high lift devices § 25.345 High lift devices. in the landing configurations. (a) If wing flaps are to be used during [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as takeoff, approach, or landing, at the amended by Amdt. 25–46, 43 FR 50595, Oct. 30, design flap speeds established for these 1978; Amdt. 25–72, 55 FR 37607, Sept. 17, 1990; stages of flight under § 25.335(e) and Amdt. 25–86, 61 FR 5221, Feb. 9, 1996; Amdt. with the wing flaps in the cor- 25–91, 62 FR 40704, July 29, 1997] responding positions, the airplane is assumed to be subjected to symmet- § 25.349 Rolling conditions. rical maneuvers and gusts. The result- The airplane must be designed for ing limit loads must correspond to the loads resulting from the rolling condi- conditions determined as follows: tions specified in paragraphs (a) and (b) (1) Maneuvering to a positive limit of this section. Unbalanced aero- load factor of 2.0; and dynamic moments about the center of (2) Positive and negative gusts of 25 gravity must be reacted in a rational ft/sec EAS acting normal to the flight or conservative manner, considering path in level flight. Gust loads result- the principal masses furnishing the re- ing on each part of the structure must acting inertia forces.

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(a) Maneuvering. The following condi- gravity must be reacted in a rational tions, speeds, and aileron deflections or conservative manner considering the (except as the deflections may be lim- airplane inertia forces. In computing ited by pilot effort) must be considered the tail loads the yawing velocity may in combination with an airplane load be assumed to be zero. factor of zero and of two-thirds of the (a) With the airplane in unacceler- positive maneuvering factor used in de- ated flight at zero yaw, it is assumed sign. In determining the required aile- that the cockpit rudder control is sud- ron deflections, the torsional flexi- denly displaced to achieve the result- bility of the wing must be considered ing rudder deflection, as limited by: in accordance with § 25.301(b): (1) Conditions corresponding to (1) The control system on control steady rolling velocities must be inves- surface stops; or tigated. In addition, conditions cor- (2) A limit pilot force of 300 pounds responding to maximum angular accel- from VMC to VA and 200 pounds from VC/ eration must be investigated for air- MC to VD/MD, with a linear variation planes with engines or other weight between VA and VC/MC. concentrations outboard of the fuse- (b) With the cockpit rudder control lage. For the angular acceleration con- deflected so as always to maintain the ditions, zero rolling velocity may be maximum rudder deflection available assumed in the absence of a rational within the limitations specified in time history investigation of the ma- paragraph (a) of this section, it is as- neuver. sumed that the airplane yaws to the (2) At VA, a sudden deflection of the overswing sideslip angle. aileron to the stop is assumed. (c) With the airplane yawed to the (3) At VC, the aileron deflection must static equilibrium sideslip angle, it is be that required to produce a rate of assumed that the cockpit rudder con- roll not less than that obtained in trol is held so as to achieve the max- paragraph (a)(2) of this section. imum rudder deflection available with- (4) At VD, the aileron deflection must be that required to produce a rate of in the limitations specified in para- roll not less than one-third of that in graph (a) of this section. paragraph (a)(2) of this section. (d) With the airplane yawed to the (b) Unsymmetrical gusts. The airplane static equilibrium sideslip angle of is assumed to be subjected to unsym- paragraph (c) of this section, it is as- metrical vertical gusts in level flight. sumed that the cockpit rudder control The resulting limit loads must be de- is suddenly returned to neutral. termined from either the wing max- [Amdt. 25–91, 62 FR 40704, July 29, 1997] imum airload derived directly from § 25.341(a), or the wing maximum air- SUPPLEMENTARY CONDITIONS load derived indirectly from the vertical load factor calculated from § 25.361 Engine torque. § 25.341(a). It must be assumed that 100 percent of the wing air load acts on one (a) Each engine mount and its sup- side of the airplane and 80 percent of porting structure must be designed for the wing air load acts on the other the effects of— side. (1) A limit engine torque corresponding to takeoff power and pro- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as peller speed acting simultaneously amended by Amdt. 25–23, 35 FR 5672, Apr. 8, 1970; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; with 75 percent of the limit loads from Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] flight condition A of § 25.333(b); (2) A limit torque corresponding to § 25.351 Yaw maneuver conditions. the maximum continuous power and The airplane must be designed for propeller speed, acting simultaneously loads resulting from the yaw maneuver with the limit loads from flight condi- conditions specified in paragraphs (a) tion A of § 25.333(b); and through (d) of this section at speeds (3) For turbopropeller installations, from VMC to VD. Unbalanced aero- in addition to the conditions specified dynamic moments about the center of

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in paragraphs (a)(1) and (2) of this sec- § 25.365 Pressurized compartment tion, a limit engine torque cor- loads. responding to takeoff power and pro- For airplanes with one or more pres- peller speed, multiplied by a factor ac- surized compartments the following counting for propeller control system apply: malfunction, including quick feath- (a) The airplane structure must be ering, acting simultaneously with 1g strong enough to withstand the flight level flight loads. In the absence of a loads combined with pressure differen- rational analysis, a factor of 1.6 must tial loads from zero up to the max- be used. imum relief valve setting. (b) For turbine engine installations, (b) The external pressure distribution the engine mounts and supporting in flight, and stress concentrations and structure must be designed to with- fatigue effects must be accounted for. stand each of the following: (c) If landings may be made with the (1) A limit engine torque load im- compartment pressurized, landing posed by sudden engine stoppage due to loads must be combined with pressure malfunction or structural failure (such differential loads from zero up to the as compressor jamming). maximum allowed during landing. (2) A limit engine torque load im- (d) The airplane structure must be posed by the maximum acceleration of designed to be able to withstand the the engine. pressure differential loads cor- (c) The limit engine torque to be con- responding to the maximum relief sidered under paragraph (a) of this sec- valve setting multiplied by a factor of tion must be obtained by multiplying 1.33 for airplanes to be approved for op- mean torque for the specified power eration to 45,000 feet or by a factor of and speed by a factor of— 1.67 for airplanes to be approved for op- (1) 1.25 for turbopropeller installa- eration above 45,000 feet, omitting tions; other loads. (2) 1.33 for reciprocating engines with (e) Any structure, component or part, five or more cylinders; or inside or outside a pressurized com- (3) Two, three, or four, for engines partment, the failure of which could with four, three, or two cylinders, re- interfere with continued safe flight and spectively. landing, must be designed to withstand [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the effects of a sudden release of pres- amended by Amdt. 25–23, 35 FR 5672, Apr. 8, sure through an opening in any com- 1970; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; partment at any operating altitude re- Amdt. 25–72, 55 FR 29776, July 20, 1990] sulting from each of the following conditions: § 25.363 Side load on engine and auxil- (1) The penetration of the compart- iary power unit mounts. ment by a portion of an engine fol- (a) Each engine and auxiliary power lowing an engine disintegration; unit mount and its supporting struc- (2) Any opening in any pressurized ture must be designed for a limit load compartment up to the size Ho in factor in lateral direction, for the side square feet; however, small compart- load on the engine and auxiliary power ments may be combined with an adja- unit mount, at least equal to the max- cent pressurized compartment and both imum load factor obtained in the yaw- considered as a single compartment for ing conditions but not less than— openings that cannot reasonably be ex- (1) 1.33; or pected to be confined to the small com- (2) One-third of the limit load factor partment. The size Ho must be com- for flight condition A as prescribed in puted by the following formula: § 25.333(b). H =PA (b) The side load prescribed in para- o s graph (a) of this section may be as- where, sumed to be independent of other flight Ho=Maximum opening in square feet, need conditions. not exceed 20 square feet. P=(As/6240)+.024 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as As=Maximum cross-sectional area of the amended by Amdt. 25–23, 35 FR 5672, Apr. 8, pressurized shell normal to the longitu- 1970; Amdt. 25–91, 62 FR 40704, July 29, 1997] dinal axis, in square feet; and

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(3) The maximum opening caused by (3) The time history of the thrust airplane or equipment failures not decay and drag build-up occurring as a shown to be extremely improbable. result of the prescribed engine failures (f) In complying with paragraph (e) of must be substantiated by test or other this section, the fail-safe features of data applicable to the particular en- the design may be considered in deter- gine-propeller combination. mining the probability of failure or (4) The timing and magnitude of the penetration and probable size of open- probable pilot corrective action must ings, provided that possible improper be conservatively estimated, consid- operation of closure devices and inad- ering the characteristics of the par- vertent door openings are also consid- ticular engine-propeller-airplane com- ered. Furthermore, the resulting dif- bination. ferential pressure loads must be com- (b) Pilot corrective action may be as- bined in a rational and conservative sumed to be initiated at the time max- manner with 1–g level flight loads and imum yawing velocity is reached, but any loads arising from emergency de- not earlier than two seconds after the pressurization conditions. These loads engine failure. The magnitude of the may be considered as ultimate condi- corrective action may be based on the tions; however, any deformations asso- control forces specified in § 25.397(b) ex- ciated with these conditions must not cept that lower forces may be assumed interfere with continued safe flight and where it is shown by anaylsis or test landing. The pressure relief provided by that these forces can control the yaw intercompartment venting may also be and roll resulting from the prescribed considered. engine failure conditions. (g) Bulkheads, floors, and partitions § 25.371 Gyroscopic loads. in pressurized compartments for occupants must be designed to withstand The structure supporting any engine the conditions specified in paragraph or auxiliary power unit must be de- (e) of this section. In addition, reason- signed for the loads including the gyro- able design precautions must be taken scopic loads arising from the condi- to minimize the probability of parts tions specified in §§ 25.331, 25.341(a), becoming detached and injuring occu- 25.349, 25.351, 25.473, 25.479, and 25.481, pants while in their seats. with the engine or auxiliary power unit at the maximum rpm appropriate to [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the condition. For the purposes of com- amended by Amdt. 25–54, 45 FR 60172, Sept. pliance with this section, the pitch ma- 11, 1980; Amdt. 25–71, 55 FR 13477, Apr. 10, neuver in § 25.331(c)(1) must be carried 1990; Amdt. 25–72, 55 FR 29776, July 20, 1990; Amdt. 25–87, 61 FR 28695, June 5, 1996] out until the positive limit maneuvering load factor (point A2 in § 25.367 Unsymmetrical loads due to § 25.333(b)) is reached. engine failure. [Amdt. 25–91, 62 FR 40704, July 29, 1997] (a) The airplane must be designed for the unsymmetrical loads resulting § 25.373 Speed control devices. from the failure of the critical engine. If speed control devices (such as Turbopropeller airplanes must be de- spoilers and drag flaps) are installed signed for the following conditions in for use in en route conditions— combination with a single malfunction (a) The airplane must be designed for of the propeller drag limiting system, the symmetrical maneuvers prescribed considering the probable pilot correc- in § 25.333 and § 25.337, the yawing ma- tive action on the flight controls: neuvers prescribed in § 25.351, and the (1) At speeds between VMC and VD, the vertical and later gust conditions pre- loads resulting from power failure be- scribed in § 25.341(a), at each setting cause of fuel flow interruption are con- and the maximum speed associated sidered to be limit loads. with that setting; and (2) At speeds between VMC and VC, the (b) If the device has automatic oper- loads resulting from the disconnection ating or load limiting features, the air- of the engine compressor from the tur- plane must be designed for the maneu- bine or from loss of the turbine blades ver and gust conditions prescribed in are considered to be ultimate loads. paragraph (a) of this section, at the

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speeds and corresponding device posi- minimum forces prescribed in tions that the mechanism allows. § 25.397(c). [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–72, 55 FR 29776, July 20, amended by Amdt. 25–23, 35 FR 5672, Apr. 8, 1990; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996] 1970; Amdt. 25–72, 55 FR 29776, July 20, 1990]

CONTROL SURFACE AND SYSTEM LOADS § 25.397 Control system loads. (a) General. The maximum and min- § 25.391 Control surface loads: Gen- imum pilot forces, specified in para- eral. graph (c) of this section, are assumed The control surfaces must be de- to act at the appropriate control grips signed for the limit loads resulting or pads (in a manner simulating flight from the flight conditions in §§ 25.331, conditions) and to be reacted at the at- 25.341(a), 25.349 and 25.351 and the tachment of the control system to the ground gust conditions in § 25.415, con- control surface horn. sidering the requirements for— (b) Pilot effort effects. In the control surface flight loading condition, the air (a) Loads parallel to hinge line, in loads on movable surfaces and the cor- § 25.393; responding deflections need not exceed (b) Pilot effort effects, in § 25.397; those that would result in flight from (c) Trim tab effects, in § 25.407; the application of any pilot force with- (d) Unsymmetrical loads, in § 25.427; in the ranges specified in paragraph (c) and of this section. Two-thirds of the max- (e) Auxiliary aerodynamic surfaces, imum values specified for the aileron in § 25.445. and elevator may be used if control surface hinge moments are based on re- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–86, 61 FR 5222, Feb. 9, liable data. In applying this criterion, 1996] the effects of servo mechanisms, tabs, and automatic pilot systems, must be § 25.393 Loads parallel to hinge line. considered. (c) Limit pilot forces and torques. The (a) Control surfaces and supporting limit pilot forces and torques are as hinge brackets must be designed for in- follows: ertia loads acting parallel to the hinge line. Maximum Minimum Control forces or forces or (b) In the absence of more rational torques torques data, the inertia loads may be assumed Aileron: to be equal to KW, where— Stick ...... 100 lbs ...... 40 lbs. (1) K=24 for vertical surfaces; Wheel 1 ...... 80 D in.-lbs 2 ... 40 D in.-lbs. (2) K=12 for horizontal surfaces; and Elevator: Stick ...... 250 lbs ...... 100 lbs. (3) W=weight of the movable surfaces. Wheel (symmetrical) ..... 300 lbs ...... 100 lbs. Wheel (unsymmetrical) 3 ...... 100 lbs. § 25.395 Control system. Rudder ...... 300 lbs ...... 130 lbs. 1 The critical parts of the aileron control system must be de- (a) Longitudinal, lateral, directional, signed for a single tangential force with a limit value equal to and drag control system and their sup- 1.25 times the couple force determined from these criteria. 2 D=wheel diameter (inches). porting structures must be designed for 3 The unsymmetrical forces must be applied at one of the loads corresponding to 125 percent of normal handgrip points on the periphery of the control wheel. the computed hinge moments of the [Doc. 5066, 29 FR 18291, Dec. 24, 1964, as movable control surface in the condi- amended by Amdt. 25–38, 41 FR 55466, Dec. 20, tions prescribed in § 25.391. 1976; Amdt. 25–72, 55 FR 29776, July 20, 1990] (b) The system limit loads, except the loads resulting from ground gusts, § 25.399 Dual control system. need not exceed the loads that can be (a) Each dual control system must be produced by the pilot (or pilots) and by designed for the pilots operating in op- automatic or power devices operating position, using individual pilot forces the controls. not less than— (c) The loads must not be less than (1) 0.75 times those obtained under those resulting from application of the § 25.395; or

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(2) The minimum forces specified in fort forces up to those specified in § 25.397(c). § 25.397(b). (b) The control system must be de- (b) Balancing tabs. Balancing tabs signed for pilot forces applied in the must be designed for deflections con- same direction, using individual pilot sistent with the primary control sur- forces not less than 0.75 times those ob- face loading conditions. tained under § 25.395. (c) Servo tabs. Servo tabs must be designed for deflections consistent with § 25.405 Secondary control system. the primary control surface loading Secondary controls, such as wheel conditions obtainable within the pilot brake, spoiler, and tab controls, must maneuvering effort, considering pos- be designed for the maximum forces sible opposition from the trim tabs. that a pilot is likely to apply to those controls. The following values may be § 25.415 Ground gust conditions. used: (a) The control system must be designed as follows for control surface PILOT CONTROL FORCE LIMITS (SECONDARY loads due to ground gusts and taxiing CONTROLS) downwind: (1) The control system between the Control Limit pilot forces stops nearest the surfaces and the Miscellaneous: cockpit controls must be designed for *Crank, wheel, or lever .. ((1 + R) / 3) × 50 lbs., but not loads corresponding to the limit hinge less than 50 lbs. nor more moments H of paragraph (a)(2) of this than 150 lbs. (R=radius). (Ap- plicable to any angle within section. These loads need not exceed— 20° of plane of control). (i) The loads corresponding to the Twist ...... 133 in.–lbs. maximum pilot loads in § 25.397(c) for Push-pull ...... To be chosen by applicant. each pilot alone; or *Limited to flap, tab, stabilizer, spoiler, and landing gear op- (ii) 0.75 times these maximum loads eration controls. for each pilot when the pilot forces are § 25.407 Trim tab effects. applied in the same direction. (2) The control system stops nearest The effects of trim tabs on the con- the surfaces, the control system locks, trol surface design conditions must be and the parts of the systems (if any) accounted for only where the surface between these stops and locks and the loads are limited by maximum pilot ef- control surface horns, must be designed fort. In these cases, the tabs are con- for limit hinge moments H, in foot sidered to be deflected in the direction pounds, obtained from the formula, that would assist the pilot, and the deflections are— H=.0034KV2cS, where— (a) For elevator trim tabs, those re- V=65 (wind speed in knots) quired to trim the airplane at any K=limit hinge moment factor for ground point within the positive portion of the gusts derived in paragraph (b) of this sec- pertinent flight envelope in § 25.333(b), tion. except as limited by the stops; and c=mean chord of the control surface aft of (b) For aileron and rudder trim tabs, the hinge line (ft); those required to trim the airplane in S=area of the control surface aft of the hinge line (sq ft); the critical unsymmetrical power and loading conditions, with appropriate (b) The limit hinge moment factor K allowance for rigging tolerances. for ground gusts must be derived as follows: § 25.409 Tabs. Surface K Position of controls (a) Trim tabs. Trim tabs must be designed to withstand loads arising from (a) Aileron ...... 0.75 Control column locked all likely combinations of tab setting, or lashed in mid-position. primary control position, and airplane (b) ...... do ...... 1 1 Ailerons at full throw. speed (obtainable without exceeding ±0.50 the flight load conditions prescribed (c) Elevator ...... 1 1 (c) Elevator full down. ±0.75 for the airplane as a whole), when the (d) ...... do ...... 1 1 (d) Elevator full up. effect of the tab is opposed by pilot ef- ±0.75

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Surface K Position of controls (b) To provide for unsymmetrical loading when outboard fins extend (e) Rudder ...... 0.75 (e) Rudder in neutral. (f) ...... do ...... 0.75 (f) Rudder at full throw. above and below the horizontal surface, the critical vertical surface loading 1 A positive value of K indicates a moment tending to de- press the surface, while a negative value of K indicates a mo- (load per unit area) determined under ment tending to raise the surface. § 25.391 must also be applied as follows: (1) 100 percent to the area of the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as vertical surfaces above (or below) the amended by Amdt. 25–72, 55 FR 29776, July 20, horizontal surface. 1990; Amdt. 25–91, 62 FR 40705, July 29, 1997] (2) 80 percent to the area below (or § 25.427 Unsymmetrical loads. above) the horizontal surface. (a) In designing the airplane for lat- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as eral gust, yaw maneuver and roll ma- amended by Amdt. 25–86, 61 FR 5222, Feb. 9, neuver conditions, account must be 1996] taken of unsymmetrical loads on the § 25.457 Wing flaps. empennage arising from effects such as slipstream and aerodynamic inter- Wing flaps, their operating mecha- ference with the wing, vertical fin and nisms, and their supporting structures other aerodynamic surfaces. must be designed for critical loads occurring in the conditions prescribed in (b) The horizontal tail must be as- § 25.345, accounting for the loads occur- sumed to be subjected to unsymmet- ring during transition from one flap po- rical loading conditions determined as sition and airspeed to another. follows: (1) 100 percent of the maximum load- § 25.459 Special devices. ing from the symmetrical maneuver conditions of § 25.331 and the vertical The loading for special devices using gust conditions of § 25.341(a) acting sep- aerodynamic surfaces (such as slots, arately on the surface on one side of slats and spoilers) must be determined from test data. the plane of symmetry; and (2) 80 percent of these loadings acting [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as on the other side. amended by Amdt. 25–72, 55 FR 29776, July 20, (c) For empennage arrangements 1990] where the horizontal tail surfaces have GROUND LOADS dihedral angles greater than plus or minus 10 degrees, or are supported by § 25.471 General. the vertical tail surfaces, the surfaces (a) For limit and the supporting structure must be Loads and equilibrium. ground loads— designed for gust velocities specified in (1) Limit ground loads obtained § 25.341(a) acting in any orientation at under this subpart are considered to be right angles to the flight path. external forces applied to the airplane (d) Unsymmetrical loading on the structure; and empennage arising from buffet condi- (2) In each specified ground load con- tions of § 25.305(e) must be taken into dition, the external loads must be account. placed in equilibrium with the linear [Doc. No. 27902, 61 FR 5222, Feb. 9, 1996] and angular inertia loads in a rational or conservative manner. § 25.445 Auxiliary aerodynamic sur- (b) Critical centers of gravity. The crit- faces. ical centers of gravity within the range (a) When significant, the aero- for which certification is requested dynamic influence between auxiliary must be selected so that the maximum aerodynamic surfaces, such as out- design loads are obtained in each land- board fins and winglets, and their sup- ing gear element. Fore and aft, porting aerodynamic surfaces, must be vertical, and lateral airplane centers of taken into account for all loading con- gravity must be considered. Lateral ditions including pitch, roll, and yaw displacements of the c.g. from the air- maneuvers, and gusts as specified in plane centerline which would result in § 25.341(a) acting at any orientation at main gear loads not greater than 103 right angles to the flight path. percent of the critical design load for

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symmetrical loading conditions may be (d) The landing gear dynamic charac- selected without considering the ef- teristics must be validated by tests as fects of these lateral c.g. displacements defined in § 25.723(a). on the loading of the main gear ele- (e) The coefficient of friction between ments, or on the airplane structure the tires and the ground may be estab- provided— lished by considering the effects of (1) The lateral displacement of the skidding velocity and tire pressure. c.g. results from random passenger or However, this coefficient of friction cargo disposition within the fuselage or need not be more than 0.8. from random unsymmetrical fuel load- [Amdt. 25–91, 62 FR 40705, July 29, 1997; Amdt. ing or fuel usage; and 25–91, 62 FR 45481, Aug. 27, 1997; Amdt 25–103, (2) Appropriate loading instructions 66 FR 27394, May 16, 2001] for random disposable loads are included under the provisions of § 25.477 Landing gear arrangement. § 25.1583(c)(1) to ensure that the lateral Sections 25.479 through 25.485 apply displacement of the center of gravity is to airplanes with conventional ar- maintained within these limits. rangements of main and nose gears, or (c) Landing gear dimension data. Fig- main and tail gears, when normal oper- ure 1 of appendix A contains the basic ating techniques are used. landing gear dimension data. § 25.479 Level landing conditions. [Amdt. 25–23, 35 FR 5673, Apr. 8, 1970] (a) In the level attitude, the airplane § 25.473 Landing load conditions and is assumed to contact the ground at assumptions. forward velocity components, ranging (a) For the landing conditions speci- from VL1 to 1.25 VL2 parallel to the fied in § 25.479 to § 25.485 the airplane is ground under the conditions prescribed assumed to contact the ground— in § 25.473 with— (1) V equal to V (TAS) at the ap- (1) In the attitudes defined in § 25.479 L1 S0 propriate landing weight and in stand- and § 25.481; ard sea level conditions; and (2) With a limit descent velocity of 10 (2) V equal to V (TAS) at the ap- fps at the design landing weight (the L2 S0 propriate landing weight and altitudes maximum weight for landing condi- in a hot day temperature of 41 degrees tions at maximum descent velocity); F. above standard. and (3) The effects of increased contact (3) With a limit descent velocity of 6 speed must be investigated if approval fps at the design take-off weight (the of downwind landings exceeding 10 maximum weight for landing condi- knots is requested. tions at a reduced descent velocity). (b) For the level landing attitude for (4) The prescribed descent velocities airplanes with tail wheels, the condi- may be modified if it is shown that the tions specified in this section must be airplane has design features that make investigated with the airplane hori- it impossible to develop these veloci- zontal reference line horizontal in ac- ties. cordance with Figure 2 of Appendix A (b) Airplane lift, not exceeding air- of this part. plane weight, may be assumed unless (c) For the level landing attitude for the presence of systems or procedures airplanes with nose wheels, shown in significantly affects the lift. Figure 2 of Appendix A of this part, the (c) The method of analysis of air- conditions specified in this section plane and landing gear loads must take must be investigated assuming the fol- into account at least the following ele- lowing attitudes: ments: (1) An attitude in which the main (1) Landing gear dynamic character- wheels are assumed to contact the istics. ground with the nose wheel just clear (2) Spin-up and springback. of the ground; and (3) Rigid body response. (2) If reasonably attainable at the (4) Structural dynamic response of specified descent and forward veloci- the airframe, if significant. ties, an attitude in which the nose and

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main wheels are assumed to contact accordance with figure 3 of appendix A. the ground simultaneously. Ground reaction conditions on the tail (d) In addition to the loading condi- wheel are assumed to act— tions prescribed in paragraph (a) of this (1) Vertically; and section, but with maximum vertical (2) Up and aft through the axle at 45 ground reactions calculated from para- degrees to the ground line. graph (a), the following apply: (c) For the tail-down landing condi- (1) The landing gear and directly af- tion for airplanes with nose wheels, the fected attaching structure must be designed for the maximum vertical airplane is assumed to be at an atti- ground reaction combined with an aft tude corresponding to either the stall- acting drag component of not less than ing angle or the maximum angle allow- 25% of this maximum vertical ground ing clearance with the ground by each reaction. part of the airplane other than the (2) The most severe combination of main wheels, in accordance with figure loads that are likely to arise during a 3 of appendix A, whichever is less. lateral drift landing must be taken [Docket No. 5066, 29 FR 18291, Dec. 24, 1964, as into account. In absence of a more ra- amended by Amdt. 25–91, 62 FR 40705, July 29, tional analysis of this condition, the 1997; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] following must be investigated: (i) A vertical load equal to 75% of the § 25.483 One-gear landing conditions. maximum ground reaction of § 25.473 must be considered in combination For the one-gear landing conditions, with a drag and side load of 40% and the airplane is assumed to be in the 25% respectively of that vertical load. level attitude and to contact the (ii) The shock absorber and tire de- ground on one main landing gear, in flections must be assumed to be 75% of accordance with Figure 4 of Appendix the deflection corresponding to the A of this part. In this attitude— maximum ground reaction of (a) The ground reactions must be the § 25.473(a)(2). This load case need not be same as those obtained on that side considered in combination with flat under § 25.479(d)(1), and tires. (b) Each unbalanced external load (3) The combination of vertical and must be reacted by airplane inertia in drag components is considered to be a rational or conservative manner. acting at the wheel axle centerline. [Docket No. 5066, 29 FR 18291, Dec. 24, 1964, as [Amdt. 25–91, 62 FR 40705, July 29, 1997; Amdt. amended by Amdt. 25–91, 62 FR 40705, July 29, 25–91, 62 FR 45481, Aug. 27, 1997] 1997] § 25.481 Tail-down landing conditions. § 25.485 Side load conditions. (a) In the tail-down attitude, the airplane is assumed to contact the ground In addition to § 25.479(d)(2) the fol- at forward velocity components, rang- lowing conditions must be considered: (a) For the side load condition, the ing from VL1 to VL2 parallel to the ground under the conditions prescribed airplane is assumed to be in the level in § 25.473 with— attitude with only the main wheels (1) VL1 equal to VS0 (TAS) at the ap- contacting the ground, in accordance propriate landing weight and in stand- with figure 5 of appendix A. ard sea level conditions; and (b) Side loads of 0.8 of the vertical re- (2) VL2 equal to VS0 (TAS) at the ap- action (on one side) acting inward and propriate landing weight and altitudes 0.6 of the vertical reaction (on the in a hot day temperature of 41 degrees other side) acting outward must be F. above standard. combined with one-half of the max- (3) The combination of vertical and imum vertical ground reactions ob- drag components considered to be act- tained in the level landing conditions. ing at the main wheel axle centerline. These loads are assumed to be applied (b) For the tail-down landing condi- at the ground contact point and to be tion for airplanes with tail wheels, the resisted by the inertia of the airplane. main and tail wheels are assumed to contact the ground simultaneously, in

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The drag loads may be assumed to be 0.8, must be combined with the vertical zero. ground reaction and applied at the ground contact point. [Docket No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–91, 62 FR 40705, July 29, (b) For an airplane with a nose wheel 1997] the limit vertical load factor is 1.2 at the design landing weight, and 1.0 at § 25.487 Rebound landing condition. the design ramp weight. A drag reac- (a) The landing gear and its sup- tion equal to the vertical reaction, porting structure must be investigated multiplied by a coefficient of friction for the loads occurring during rebound of 0.8, must be combined with the of the airplane from the landing sur- vertical reaction and applied at the face. ground contact point of each wheel (b) With the landing gear fully ex- with brakes. The following two atti- tended and not in contact with the tudes, in accordance with figure 6 of ground, a load factor of 20.0 must act appendix A, must be considered: on the unsprung weights of the landing (1) The level attitude with the wheels gear. This load factor must act in the contacting the ground and the loads direction of motion of the unsprung distributed between the main and nose weights as they reach their limiting gear. Zero pitching acceleration is as- positions in extending with relation to sumed. the sprung parts of the landing gear. (2) The level attitude with only the main gear contacting the ground and § 25.489 Ground handling conditions. with the pitching moment resisted by Unless otherwise prescribed, the angular acceleration. landing gear and airplane structure (c) A drag reaction lower than that must be investigated for the conditions prescribed in this section may be used in §§ 25.491 through 25.509 with the air- if it is substantiated that an effective plane at the design ramp weight (the drag force of 0.8 times the vertical re- maximum weight for ground handling action cannot be attained under any conditions). No wing lift may be con- likely loading condition. sidered. The shock absorbers and tires (d) An airplane equipped with a nose may be assumed to be in their static gear must be designed to withstand the position. loads arising from the dynamic pitch- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ing motion of the airplane due to sud- amended by Amdt. 25–23, 35 FR 5673, Apr. 8, den application of maximum braking 1970] force. The airplane is considered to be at design takeoff weight with the nose § 25.491 Taxi, takeoff and landing roll. and main gears in contact with the Within the range of appropriate ground, and with a steady-state ground speeds and approved weights, vertical load factor of 1.0. The steady- the airplane structure and landing gear state nose gear reaction must be com- are assumed to be subjected to loads bined with the maximum incremental not less than those obtained when the nose gear vertical reaction caused by aircraft is operating over the roughest the sudden application of maximum ground that may reasonably be ex- braking force as described in para- pected in normal operation. graphs (b) and (c) of this section. [Amdt. 25–91, 62 FR 40705, July 29, 1997] (e) In the absence of a more rational analysis, the nose gear vertical reac- § 25.493 Braked roll conditions. tion prescribed in paragraph (d) of this section must be calculated according (a) An airplane with a tail wheel is to the following formula: assumed to be in the level attitude with the load on the main wheels, in accordance with figure 6 of appendix A. W ⎡ fAEμ ⎤ V = T ⎢B + ⎥ The limit vertical load factor is 1.2 at N AB+ ⎣ AB++μ E⎦ the design landing weight and 1.0 at the design ramp weight. A drag reac- Where:

tion equal to the vertical reaction mul- VN=Nose gear vertical reaction. tiplied by a coefficient of friction of WT=Design takeoff weight. 409

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A=Horizontal distance between the c.g. of ground reaction at that point are as- the airplane and the nose wheel. sumed. B=Horizontal distance between the c.g. of (b) With the airplane assumed to be the airplane and the line joining the centers of the main wheels. in static equilibrium with the loads re- E=Vertical height of the c.g. of the airplane sulting from the use of brakes on one above the ground in the 1.0 g static con- side of the main landing gear, the nose dition. gear, its attaching structure, and the μ=Coefficient of friction of 0.80. fuselage structure forward of the cen- f=Dynamic response factor; 2.0 is to be used ter of gravity must be designed for the unless a lower factor is substantiated. In following loads: the absence of other information, the dynamic response factor f may be defined (1) A vertical load factor at the cen- by the equation: ter of gravity of 1.0. (2) A forward acting load at the air- ⎛ −πξ ⎞ plane center of gravity of 0.8 times the =+ ⎜ ⎟ vertical load on one main gear. f 1 exp⎜ ⎟ ⎝ 1− ξ 2 ⎠ (3) Side and vertical loads at the ground contact point on the nose gear Where: that are required for static equi- x is the effective critical damping ratio of librium. the rigid body pitching mode about the (4) A side load factor at the airplane main landing gear effective ground con- center of gravity of zero. tact point. (c) If the loads prescribed in para- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as graph (b) of this section result in a amended by Amdt. 25–23, 35 FR 5673, Apr. 8, nose gear side load higher than 0.8 1970; Amdt. 25–97, 63 FR 29072, May 27, 1998] times the vertical nose gear load, the § 25.495 Turning. design nose gear side load may be limited to 0.8 times the vertical load, with In the static position, in accordance unbalanced yawing moments assumed with figure 7 of appendix A, the air- to be resisted by airplane inertia plane is assumed to execute a steady forces. turn by nose gear steering, or by appli- (d) For other than the nose gear, its cation of sufficient differential power, attaching structure, and the forward so that the limit load factors applied at fuselage structure, the loading condi- the center of gravity are 1.0 vertically tions are those prescribed in paragraph and 0.5 laterally. The side ground reaction of each wheel must be 0.5 of the (b) of this section, except that— vertical reaction. (1) A lower drag reaction may be used if an effective drag force of 0.8 times § 25.497 Tail-wheel yawing. the vertical reaction cannot be reached (a) A vertical ground reaction equal under any likely loading condition; and to the static load on the tail wheel, in (2) The forward acting load at the combination with a side component of center of gravity need not exceed the equal magnitude, is assumed. maximum drag reaction on one main (b) If there is a swivel, the tail wheel gear, determined in accordance with is assumed to be swiveled 90° to the air- § 25.493(b). plane longitudinal axis with the result- (e) With the airplane at design ramp ant load passing through the axle. weight, and the nose gear in any steer- (c) If there is a lock, steering device, able position, the combined application or shimmy damper the tail wheel is of full normal steering torque and also assumed to be in the trailing posi- vertical force equal to 1.33 times the tion with the side load acting at the maximum static reaction on the nose ground contact point. gear must be considered in designing the nose gear, its attaching structure, § 25.499 Nose-wheel yaw and steering. and the forward fuselage structure. (a) A vertical load factor of 1.0 at the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as airplane center of gravity, and a side amended by Amdt. 25–23, 35 FR 5673, Apr. 8, component at the nose wheel ground 1970; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; contact equal to 0.8 of the vertical Amdt. 25–91, 62 FR 40705, July 29, 1997]

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§ 25.503 Pivoting. (2) The shock struts and tires must be in their static positions; and (a) The airplane is assumed to pivot (3) With W as the design ramp about one side of the main gear with T weight, the towing load, F is— the brakes on that side locked. The TOW, (i) 0.3 W for W less than 30,000 limit vertical load factor must be 1.0 T T pounds; and the coefficient of friction 0.8. (ii) (6W + 450,000)/70 for W between (b) The airplane is assumed to be in T T 30,000 and 100,000 pounds; and static equilibrium, with the loads being (iii) 0.15 W for W over 100,000 applied at the ground contact points, T T pounds. in accordance with figure 8 of appendix (b) For towing points not on the A. landing gear but near the plane of sym- § 25.507 Reversed braking. metry of the airplane, the drag and side tow load components specified for (a) The airplane must be in a three the auxiliary gear apply. For towing point static ground attitude. Hori- points located outboard of the main zontal reactions parallel to the ground gear, the drag and side tow load compo- and directed forward must be applied nents specified for the main gear apply. at the ground contact point of each Where the specified angle of swivel wheel with brakes. The limit loads cannot be reached, the maximum ob- must be equal to 0.55 times the vertical tainable angle must be used. load at each wheel or to the load devel- (c) The towing loads specified in oped by 1.2 times the nominal max- paragraph (d) of this section must be imum static brake torque, whichever is reacted as follows: less. (1) The side component of the towing (b) For airplanes with nose wheels, load at the main gear must be reacted the pitching moment must be balanced by a side force at the static ground line by rotational inertia. of the wheel to which the load is ap- (c) For airplanes with tail wheels, the plied. resultant of the ground reactions must (2) The towing loads at the auxiliary pass through the center of gravity of gear and the drag components of the the airplane. towing loads at the main gear must be reacted as follows: § 25.509 Towing loads. (i) A reaction with a maximum value (a) The towing loads specified in equal to the vertical reaction must be paragraph (d) of this section must be applied at the axle of the wheel to considered separately. These loads which the load is applied. Enough air- must be applied at the towing fittings plane inertia to achieve equilibrium and must act parallel to the ground. In must be applied. addition— (ii) The loads must be reacted by air- (1) A vertical load factor equal to 1.0 plane inertia. must be considered acting at the center (d) The prescribed towing loads are as of gravity; follows:

Load Tow point Position Magnitude No. Direction

Main gear ...... 0.75 FTOW per main 1 Forward, parallel to drag axis. gear unit. 2 Forward, at 30° to drag axis. 3 Aft, parallel to drag axis. 4 Aft, at 30° to drag axis. Auxiliary gear ...... Swiveled forward ...... 1.0 FTOW ...... 5 Forward. 6 Aft. Swiveled aft ...... do ...... 7 Forward. 8 Aft. Swiveled 45° from forward ..... 0.5 FTOW ...... 9 Forward, in plane of wheel. 10 Aft, in plane of wheel. Swiveled 45° from aft ...... do ...... 11 Forward, in plane of wheel. 12 Aft, in plane of wheel.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5673, Apr. 8, 1970]

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§ 25.511 Ground load: unsymmetrical account the physical arrangement of loads on multiple-wheel units. the gear components. In addition— (a) General. Multiple-wheel landing (1) The deflation of any one tire for gear units are assumed to be subjected each multiple wheel landing gear unit, to the limit ground loads prescribed in and the deflation of any two critical this subpart under paragraphs (b) tires for each landing gear unit using through (f) of this section. In addi- four or more wheels per unit, must be tion— considered; and (1) A tandem strut gear arrangement (2) The ground reactions must be ap- is a multiple-wheel unit; and plied to the wheels with inflated tires (2) In determining the total load on a except that, for multiple-wheel gear gear unit with respect to the provisions units with more than one shock strut, of paragraphs (b) through (f) of this a rational distribution of the ground section, the transverse shift in the load reactions between the deflated and in- centroid, due to unsymmetrical load flated tires, accounting for the dif- distribution on the wheels, may be ne- glected. ferences in shock strut extensions re- (b) Distribution of limit loads to wheels; sulting from a deflated tire, may be tires inflated. The distribution of the used. limit loads among the wheels of the (d) Landing conditions. For one and landing gear must be established for for two deflated tires, the applied load each landing, taxiing, and ground han- to each gear unit is assumed to be 60 dling condition, taking into account percent and 50 percent, respectively, of the effects of the following factors: the limit load applied to each gear for (1) The number of wheels and their each of the prescribed landing condi- physical arrangements. For truck type tions. However, for the drift landing landing gear units, the effects of any condition of § 25.485, 100 percent of the seesaw motion of the truck during the vertical load must be applied. landing impact must be considered in (e) Taxiing and ground handling condi- determining the maximum design loads tions. For one and for two deflated for the fore and aft wheel pairs. tires— (2) Any differentials in tire diameters (1) The applied side or drag load fac- resulting from a combination of manu- tor, or both factors, at the center of facturing tolerances, tire growth, and gravity must be the most critical value tire wear. A maximum tire-diameter up to 50 percent and 40 percent, respec- differential equal to 2⁄3 of the most unfavorable combination of diameter tively, of the limit side or drag load variations that is obtained when tak- factors, or both factors, corresponding ing into account manufacturing toler- to the most severe condition resulting ances, tire growth, and tire wear, may from consideration of the prescribed be assumed. taxiing and ground handling condi- (3) Any unequal tire inflation pres- tions; sure, assuming the maximum variation (2) For the braked roll conditions of to be ±5 percent of the nominal tire in- § 25.493 (a) and (b)(2), the drag loads on flation pressure. each inflated tire may not be less than (4) A runway crown of zero and a run- those at each tire for the symmetrical way crown having a convex upward load distribution with no deflated tires; shape that may be approximated by a (3) The vertical load factor at the slope of 11⁄2 percent with the hori- center of gravity must be 60 percent zontal. Runway crown effects must be and 50 percent, respectively, of the fac- considered with the nose gear unit on tor with no deflated tires, except that either slope of the crown. it may not be less than 1g; and (5) The airplane attitude. (4) Pivoting need not be considered. (6) Any structural deflections. (c) Deflated tires. The effect of de- (f) Towing conditions. For one and for flated tires on the structure must be two deflated tires, the towing load, considered with respect to the loading FTOW, must be 60 percent and 50 percent, conditions specified in paragraphs (d) respectively, of the load prescribed. through (f) of this section, taking into

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§ 25.519 Jacking and tie-down provi- most severe sea conditions likely to be sions. encountered. (a) General. The airplane must be de- (b) Unless a more rational analysis of signed to withstand the limit load con- the water loads is made, or the stand- ditions resulting from the static ards in ANC–3 are used, §§ 25.523 ground load conditions of paragraph (b) through 25.537 apply. of this section and, if applicable, para- (c) The requirements of this section graph (c) of this section at the most and §§ 25.523 through 25.537 apply also to critical combinations of airplane amphibians. weight and center of gravity. The max- § 25.523 Design weights and center of imum allowable load at each jack pad gravity positions. must be specified. (b) Jacking. The airplane must have (a) Design weights. The water load re- provisions for jacking and must with- quirements must be met at each oper- stand the following limit loads when ating weight up to the design landing the airplane is supported on jacks— weight except that, for the takeoff con- (1) For jacking by the landing gear at dition prescribed in § 25.531, the design the maximum ramp weight of the air- water takeoff weight (the maximum plane, the airplane structure must be weight for water taxi and takeoff run) designed for a vertical load of 1.33 must be used. times the vertical static reaction at (b) Center of gravity positions. The each jacking point acting singly and in critical centers of gravity within the combination with a horizontal load of limits for which certification is re- 0.33 times the vertical static reaction quested must be considered to reach applied in any direction. maximum design loads for each part of (2) For jacking by other airplane the seaplane structure. structure at maximum approved jack- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ing weight: amended by Amdt. 25–23, 35 FR 5673, Apr. 8, (i) The airplane structure must be de- 1970] signed for a vertical load of 1.33 times the vertical reaction at each jacking § 25.525 Application of loads. point acting singly and in combination (a) Unless otherwise prescribed, the with a horizontal load of 0.33 times the seaplane as a whole is assumed to be vertical static reaction applied in any subjected to the loads corresponding to direction. the load factors specified in § 25.527. (ii) The jacking pads and local struc- (b) In applying the loads resulting ture must be designed for a vertical from the load factors prescribed in load of 2.0 times the vertical static re- § 25.527, the loads may be distributed action at each jacking point, acting over the hull or main float bottom (in singly and in combination with a hori- order to avoid excessive local shear zontal load of 0.33 times the vertical loads and bending moments at the lo- static reaction applied in any direc- cation of water load application) using tion. pressures not less than those pre- (c) Tie-down. If tie-down points are scribed in § 25.533(b). provided, the main tie-down points and (c) For twin float seaplanes, each local structure must withstand the float must be treated as an equivalent limit loads resulting from a 65-knot hull on a fictitious seaplane with a horizontal wind from any direction. weight equal to one-half the weight of [Doc. No. 26129, 59 FR 22102, Apr. 28, 1994] the twin float seaplane. (d) Except in the takeoff condition of WATER LOADS § 25.531, the aerodynamic lift on the seaplane during the impact is assumed § 25.521 General. to be 2⁄3 of the weight of the seaplane. (a) Seaplanes must be designed for the water loads developed during take- § 25.527 Hull and main float load fac- off and landing, with the seaplane in tors. any attitude likely to occur in normal (a) Water reaction load factors nW operation, and at the appropriate for- must be computed in the following ward and sinking velocities under the manner:

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(1) For the step landing case § 25.529 Hull and main float landing conditions. CV1 2 n = S0 (a) Symmetrical step, bow, and stern w ⎛ 2 ⎞ 1 landing. For symmetrical step, bow, 3 β 3 ⎝Tan ⎠ W and stern landings, the limit water reaction load factors are those computed (2) For the bow and stern landing under § 25.527. In addition— cases (1) For symmetrical step landings, the resultant water load must be ap- CV = 1 S02 × K1 plied at the keel, through the center of nw 1 2 gravity, and must be directed per- ⎛ 2 ⎞ 3 3 Tan 3 β W ()1+ r 2 pendicularly to the keel line; ⎝ ⎠ x (2) For symmetrical bow landings, (b) The following values are used: the resultant water load must be applied at the keel, one-fifth of the longi- (1) nW=water reaction load factor (that is, the water reaction divided by tudinal distance from the bow to the seaplane weight). step, and must be directed perpendicularly to the keel line; and (2) C1=empirical seaplane operations factor equal to 0.012 (except that this (3) For symmetrical stern landings, factor may not be less than that nec- the resultant water load must be ap- essary to obtain the minimum value of plied at the keel, at a point 85 percent step load factor of 2.33). of the longitudinal distance from the step to the stern post, and must be di- (3) VS0=seaplane stalling speed in knots with flaps extended in the appro- rected perpendicularly to the keel line. priate landing position and with no (b) Unsymmetrical landing for hull and slipstream effect. single float seaplanes. Unsymmetrical (4) b=angle of dead rise at the longi- step, bow, and stern landing conditions tudinal station at which the load fac- must be investigated. In addition— tor is being determined in accordance (1) The loading for each condition with figure 1 of appendix B. consists of an upward component and a (5) W= seaplane design landing side component equal, respectively, to weight in pounds. 0.75 and 0.25 tan b times the resultant load in the corresponding symmetrical (6) K1=empirical hull station weighing factor, in accordance with figure 2 landing condition; and of appendix B. (2) The point of application and di-

(7) rx=ratio of distance, measured par- rection of the upward component of the allel to hull reference axis, from the load is the same as that in the sym- center of gravity of the seaplane to the metrical condition, and the point of ap- hull longitudinal station at which the plication of the side component is at load factor is being computed to the ra- the same longitudinal station as the dius of gyration in pitch of the sea- upward component but is directed in- plane, the hull reference axis being a ward perpendicularly to the plane of straight line, in the plane of sym- symmetry at a point midway between metry, tangential to the keel at the the keel and chine lines. main step. (c) Unsymmetrical landing; twin float (c) For a twin float seaplane, because seaplanes. The unsymmetrical loading of the effect of flexibility of the attach- consists of an upward load at the step ment of the floats to the seaplane, the of each float of 0.75 and a side load of factor K1 may be reduced at the bow 0.25 tan b at one float times the step and stern to 0.8 of the value shown in landing load reached under § 25.527. The figure 2 of appendix B. This reduction side load is directed inboard, per- applies only to the design of the carry- pendicularly to the plane of symmetry through and seaplane structure. midway between the keel and chine [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as lines of the float, at the same longitu- amended by Amdt. 25–23, 35 FR 5673, Apr. 8, dinal station as the upward load. 1970]

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§ 25.531 Hull and main float takeoff tended in the appropriate takeoff posi- condition. tion; and bK=angle of dead rise at keel, in accordance For the wing and its attachment to with figure 1 of appendix B. the hull or main float— (2) For a flared bottom, the pressure (a) The aerodynamic wing lift is as- at the beginning of the flare is the sumed to be zero; and same as that for an unflared bottom, (b) A downward inertia load, cor- and the pressure between the chine and responding to a load factor computed the beginning of the flare varies lin- from the following formula, must be early, in accordance with figure 3 of ap- applied: pendix B. The pressure distribution is the same as that prescribed in para- CV2 n = TO S1 graph (b)(1) of this section for an ⎛ 2 ⎞ 1 unflared bottom except that the pres- 3 β 3 ⎝tan ⎠ W sure at the chine is computed as follows: where— n=inertia load factor; KV =×2 S12 CTO=empirical seaplane operations factor PCch 3 equal to 0.004; tan β VS1=seaplane stalling speed (knots) at the de- where— sign takeoff weight with the flaps ex- P =pressure (p.s.i.) at the chine; tended in the appropriate takeoff posi- ch C =0.0016; tion; 3 K2=hull station weighing factor, in accord- b=angle of dead rise at the main step (de- ance with figure 2 of appendix B; grees); and VS1=seaplane stalling speed at the design W=design water takeoff weight in pounds. water takeoff weight with flaps extended [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as in the appropriate takeoff position; and amended by Amdt. 25–23, 35 FR 5673, Apr. 8, b=angle of dead rise at appropriate station. 1970] The area over which these pressures § 25.533 Hull and main float bottom are applied must simulate pressures oc- pressures. curring during high localized impacts on the hull or float, but need not ex- (a) General. The hull and main float structure, including frames and bulk- tend over an area that would induce heads, stringers, and bottom plating, critical stresses in the frames or in the must be designed under this section. overall structure. (c) Distributed pressures. For the de- (b) Local pressures. For the design of sign of the frames, keel, and chine the bottom plating and stringers and structure, the following pressure dis- their attachments to the supporting tributions apply: structure, the following pressure dis- (1) Symmetrical pressures are com- tributions must be applied: puted as follows: (1) For an unflared bottom, the pressure at the chine is 0.75 times the pres- KV2 2 sure at the keel, and the pressures be- PC=× S0 tween the keel and chine vary linearly, 4 tan β in accordance with figure 3 of appendix where— B. The pressure at the keel (psi) is P=pressure (p.s.i.); computed as follows: C4=0.078 C1 (with C1 computed under § 25.527); K2=hull station weighing factor, determined KV2 2 in accordance with figure 2 of appendix PC=× S1 B; k 2 β tan k VS0=seaplane stalling speed (Knots) with landing flaps extended in the appropriate where— position and with no slipstream effect; Pk=pressure (p.s.i.) at the keel; and C2=0.00213; VS0=seaplane stalling speed with landing K2=hull station weighing factor, in accord- flaps extended in the appropriate posi- ance with figure 2 of appendix B; tion and with no slipstream effect; and VS1=seaplane stalling speed (Knots) at the de- b=angle of dead rise at appropriate sta- sign water takeoff weight with flaps ex- tion.

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(2) The unsymmetrical pressure dis- metry of the float to the radius of gyra- tribution consists of the pressures pre- tion in roll. scribed in paragraph (c)(1) of this sec- (c) Bow loading. The resultant limit tion on one side of the hull or main load must be applied in the plane of float centerline and one-half of that symmetry of the float at a point one- pressure on the other side of the hull or fourth of the distance from the bow to main float centerline, in accordance the step and must be perpendicular to with figure 3 of appendix B. the tangent to the keel line at that These pressures are uniform and must point. The magnitude of the resultant be applied simultaneously over the en- load is that specified in paragraph (b) tire hull or main float bottom. The of this section. loads obtained must be carried into the (d) Unsymmetrical step loading. The re- sidewall structure of the hull proper, sultant water load consists of a compo- but need not be transmitted in a fore nent equal to 0.75 times the load speci- and aft direction as shear and bending fied in paragraph (a) of this section and loads. a side component equal to 3.25 tan b [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as times the load specified in paragraph amended by Amdt. 25–23, 35 FR 5673, Apr. 8, (b) of this section. The side load must 1970] be applied perpendicularly to the plane of symmetry of the float at a point § 25.535 Auxiliary float loads. midway between the keel and the (a) General. Auxiliary floats and their chine. attachments and supporting structures (e) Unsymmetrical bow loading. The re- must be designed for the conditions sultant water load consists of a compo- prescribed in this section. In the cases nent equal to 0.75 times the load speci- specified in paragraphs (b) through (e) fied in paragraph (b) of this section and of this section, the prescribed water a side component equal to 0.25 tan b loads may be distributed over the float times the load specified in paragraph bottom to avoid excessive local loads, (c) of this section. The side load must using bottom pressures not less than be applied perpendicularly to the plane those prescribed in paragraph (g) of of symmetry at a point midway be- this section. tween the keel and the chine. (b) Step loading. The resultant water (f) Immersed float condition. The re- load must be applied in the plane of sultant load must be applied at the symmetry of the float at a point three- centroid of the cross section of the fourths of the distance from the bow to float at a point one-third of the dis- the step and must be perpendicular to tance from the bow to the step. The the keel. The resultant limit load is limit load components are as follows: computed as follows, except that the = value of L need not exceed three times vertical ρgV the weight of the displaced water when 2 2 the float is completely submerged: ρ 3 ⎛ ⎞ aft = CV KV x 2 ⎝ S ⎠ 2 0 3 CV5 2 W 2 2 = S0 ρ 3 ⎛ ⎞ L 2 side = CV KV 2 y ⎝ S ⎠ 3 2 3 2 0 tan β sy()1+ r where— where— r=mass density of water (slugs/ft.2); L=limit load (lbs.); V=volume of float (ft.2); C5=0.0053; Cx=coefficient of drag force, equal to 0.133; VS0=seaplane stalling speed (knots) with Cy=coefficient of side force, equal to 0.106; landing flaps extended in the appropriate K=0.8, except that lower values may be used position and with no slipstream effect; if it is shown that the floats are incapa- W=seaplane design landing weight in pounds; ble of submerging at a speed of 0.8 VS0 in 3 bS=angle of dead rise at a station ⁄4 of the normal operations; distance from the bow to the step, but VS0=seaplane stalling speed (knots) with need not be less than 15 degrees; and landing flaps extended in the appropriate ry=ratio of the lateral distance between the position and with no slipstream effect; center of gravity and the plane of sym- and

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g=acceleration due to gravity (ft./sec.2). (2) When such positioning is not prac- (g) Float bottom pressures. The float tical (e.g. fuselage mounted engines or bottom pressures must be established auxiliary power units) each such item under § 25.533, except that the value of of mass shall be restrained under all loads up to those specified in paragraph K2 in the formulae may be taken as 1.0. The angle of dead rise to be used in de- (b)(3) of this section. The local attach- termining the float bottom pressures is ments for these items should be de- set forth in paragraph (b) of this sec- signed to withstand 1.33 times the spection. ified loads if these items are subject to severe wear and tear through frequent [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as removal (e.g. quick change interior amended by Amdt. 25–23, 35 FR 5673, Apr. 8, 1970] items). (d) Seats and items of mass (and § 25.537 Seawing loads. their supporting structure) must not Seawing design loads must be based deform under any loads up to those on applicable test data. specified in paragraph (b)(3) of this section in any manner that would impede EMERGENCY LANDING CONDITIONS subsequent rapid evacuation of occupants. § 25.561 General. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) The airplane, although it may be amended by Amdt. 25–23, 35 FR 5673, Apr. 8, damaged in emergency landing condi- 1970; Amdt. 25–64, 53 FR 17646, May 17, 1988; tions on land or water, must be de- Amdt. 25–91, 62 FR 40706, July 29, 1997] signed as prescribed in this section to protect each occupant under those con- § 25.562 Emergency landing dynamic ditions. conditions. (b) The structure must be designed to (a) The seat and restraint system in give each occupant every reasonable the airplane must be designed as pre- chance of escaping serious injury in a scribed in this section to protect each minor crash landing when— occupant during an emergency landing (1) Proper use is made of seats, belts, condition when— and all other safety design provisions; (1) Proper use is made of seats, safety (2) The wheels are retracted (where belts, and shoulder harnesses provided applicable); and for in the design; and (3) The occupant experiences the following ultimate inertia forces acting (2) The occupant is exposed to loads separately relative to the surrounding resulting from the conditions pre- structure: scribed in this section. (i) Upward, 3.0g (b) Each seat type design approved (ii) Forward, 9.0g for crew or passenger occupancy during (iii) Sideward, 3.0g on the airframe; takeoff and landing must successfully and 4.0g on the seats and their attach- complete dynamic tests or be dem- ments. onstrated by rational analysis based on (iv) Downward, 6.0g dynamic tests of a similar type seat, in (v) Rearward, 1.5g accordance with each of the following (c) For equipment, cargo in the pas- emergency landing conditions. The senger compartments and any other tests must be conducted with an occu- large masses, the following apply: pant simulated by a 170-pound (1) Except as provided in paragraph anthropomorphic test dummy, as de- (c)(2) of this section, these items must fined by 49 CFR Part 572, Subpart B, or be positioned so that if they break its equivalent, sitting in the normal loose they will be unlikely to: upright position. (i) Cause direct injury to occupants; (1) A change in downward vertical ve- (ii) Penetrate fuel tanks or lines or locity (D v) of not less than 35 feet per cause fire or explosion hazard by dam- second, with the airplane’s longitu- age to adjacent systems; or dinal axis canted downward 30 degrees (iii) Nullify any of the escape facili- with respect to the horizontal plane ties provided for use after an emer- and with the wings level. Peak floor de- gency landing. celeration must occur in not more than

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0.08 seconds after impact and must (1) Where upper torso straps are used reach a minimum of 14g. for crewmembers, tension loads in indi- (2) A change in forward longitudinal vidual straps must not exceed 1,750 velocity (D v) of not less than 44 feet pounds. If dual straps are used for re- per second, with the airplane’s longitu- straining the upper torso, the total dinal axis horizontal and yawed 10 de- strap tension loads must not exceed grees either right or left, whichever 2,000 pounds. would cause the greatest likelihood of (2) The maximum compressive load the upper torso restraint system measured between the pelvis and the (where installed) moving off the occu- lumbar column of the anthropomorphic pant’s shoulder, and with the wings dummy must not exceed 1,500 pounds. level. Peak floor deceleration must (3) The upper torso restraint straps occur in not more than 0.09 seconds (where installed) must remain on the after impact and must reach a min- occupant’s shoulder during the impact. imum of 16g. Where floor rails or floor (4) The lap safety belt must remain fittings are used to attach the seating on the occupant’s pelvis during the im- devices to the test fixture, the rails or pact. fittings must be misaligned with re- (5) Each occupant must be protected spect to the adjacent set of rails or fit- from serious head injury under the con- tings by at least 10 degrees vertically ditions prescribed in paragraph (b) of (i.e., out of Parallel) with one rolled 10 this section. Where head contact with degrees. seats or other structure can occur, pro- (c) The following performance meas- tection must be provided so that the ures must not be exceeded during the head impact does not exceed a Head In- dynamic tests conducted in accordance jury Criterion (HIC) of 1,000 units. The with paragraph (b) of this section: level of HIC is defined by the equation:

⎧ ⎡ ⎤25. ⎫ ⎪ 1 t2 ⎪ HIC=−⎨() t t ⎢ ∫ atdt() ⎥ ⎬ 21 − t ⎪ ⎣⎢()tt211 ⎦⎥ ⎪ ⎩ ⎭max

Where: § 25.563 Structural ditching provi- t1 is the initial integration time, sions. t2 is the final integration time, and Structural strength considerations of a(t) is the total acceleration vs. time curve ditching provisions must be in accord- for the head strike, and where ance with § 25.801(e). (t) is in seconds, and (a) is in units of gravity (g). FATIGUE EVALUATION

(6) Where leg injuries may result § 25.571 Damage—tolerance and fa- from contact with seats or other struc- tigue evaluation of structure. ture, protection must be provided to (a) General. An evaluation of the prevent axially compressive loads ex- strength, detail design, and fabrication ceeding 2,250 pounds in each femur. must show that catastrophic failure (7) The seat must remain attached at due to fatigue, corrosion, manufac- all points of attachment, although the turing defects, or accidental damage, structure may have yielded. will be avoided throughout the oper- (8) Seats must not yield under the ational life of the airplane. This eval- tests specified in paragraphs (b)(1) and uation must be conducted in accord- (b)(2) of this section to the extent they ance with the provisions of paragraphs would impede rapid evacuation of the (b) and (e) of this section, except as airplane occupants. specified in paragraph (c) of this section, for each part of the structure that [Amdt. 25–64, 53 FR 17646, May 17, 1988] could contribute to a catastrophic failure (such as wing, empennage, control

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surfaces and their systems, the fuse- onstrated that load path failure, par- lage, engine mounting, landing gear, tial failure, or crack arrest will be de- and their related primary attach- tected and repaired during normal ments). For turbojet powered air- maintenance, inspection, or operation planes, those parts that could con- of an airplane prior to failure of the re- tribute to a catastrophic failure must maining structure. also be evaluated under paragraph (d) (b) Damage-tolerance evaluation. The of this section. In addition, the fol- evaluation must include a determina- lowing apply: tion of the probable locations and (1) Each evaluation required by this modes of damage due to fatigue, corro- section must include— sion, or accidental damage. Repeated (i) The typical loading spectra, tem- load and static analyses supported by peratures, and humidities expected in test evidence and (if available) service service; experience must also be incorporated (ii) The identification of principal in the evaluation. Special consider- structural elements and detail design ation for widespread fatigue damage points, the failure of which could cause must be included where the design is catastrophic failure of the airplane; such that this type of damage could and occur. An LOV must be established (iii) An analysis, supported by test that corresponds to the period of time, evidence, of the principal structural stated as a number of total accumu- elements and detail design points iden- lated flight cycles or flight hours or tified in paragraph (a)(1)(ii) of this sec- both, during which it is demonstrated tion. that widespread fatigue damage will (2) The service history of airplanes of not occur in the airplane structure. similar structural design, taking due This demonstration must be by full- account of differences in operating con- scale fatigue test evidence. The type ditions and procedures, may be used in certificate may be issued prior to com- the evaluations required by this sec- pletion of full-scale fatigue testing, tion. provided the Administrator has ap- (3) Based on the evaluations required proved a plan for completing the re- by this section, inspections or other quired tests. In that case, the Air- procedures must be established, as nec- worthiness Limitations section of the essary, to prevent catastrophic failure, Instructions for Continued Airworthi- and must be included in the Airworthi- ness required by § 25.1529 must specify ness Limitations section of the In- that no airplane may be operated be- structions for Continued Airworthiness yond a number of cycles equal to 1⁄2 the required by § 25.1529. The limit of valid- number of cycles accumulated on the ity of the engineering data that sup- fatigue test article, until such testing ports the structural maintenance pro- is completed. The extent of damage for gram (hereafter referred to as LOV), residual strength evaluation at any stated as a number of total accumu- time within the operational life of the lated flight cycles or flight hours or airplane must be consistent with the both, established by this section must initial detectability and subsequent also be included in the Airworthiness growth under repeated loads. The resid- Limitations section of the Instructions ual strength evaluation must show for Continued Airworthiness required that the remaining structure is able to by § 25.1529. Inspection thresholds for withstand loads (considered as static the following types of structure must ultimate loads) corresponding to the be established based on crack growth following conditions: analyses and/or tests, assuming the (1) The limit symmetrical maneu- structure contains an initial flaw of vering conditions specified in § 25.337 at the maximum probable size that could all speeds up to Vc and in § 25.345. exist as a result of manufacturing or (2) The limit gust conditions speci- service-induced damage: fied in § 25.341 at the specified speeds up (i) Single load path structure, and to VC and in § 25.345. (ii) Multiple load path ‘‘fail-safe’’ (3) The limit rolling conditions speci- structure and crack arrest ‘‘fail-safe’’ fied in § 25.349 and the limit unsymmet- structure, where it cannot be dem- rical conditions specified in §§ 25.367

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and 25.427 (a) through (c), at speeds up during which likely structural damage to VC. occurs as a result of— (4) The limit yaw maneuvering condi- (1) Impact with a 4-pound bird when tions specified in § 25.351(a) at the spec- the velocity of the airplane relative to ified speeds up to VC. the bird along the airplane’s flight (5) For pressurized cabins, the fol- path is equal to Vc at sea level or 0.85Vc lowing conditions: at 8,000 feet, whichever is more critical; (i) The normal operating differential (2) Uncontained fan blade impact; pressure combined with the expected (3) Uncontained engine failure; or external aerodynamic pressures applied simultaneously with the flight loading (4) Uncontained high energy rotating conditions specified in paragraphs machinery failure. (b)(1) through (4) of this section, if they The damaged structure must be able to have a significant effect. withstand the static loads (considered (ii) The maximum value of normal as ultimate loads) which are reason- operating differential pressure (includ- ably expected to occur on the flight. ing the expected external aerodynamic Dynamic effects on these static loads pressures during 1 g level flight) multi- need not be considered. Corrective ac- plied by a factor of 1.15, omitting other tion to be taken by the pilot following loads. the incident, such as limiting maneu- (6) For landing gear and directly-af- vers, avoiding turbulence, and reducing fected airframe structure, the limit speed, must be considered. If signifi- ground loading conditions specified in cant changes in structural stiffness or §§ 25.473, 25.491, and 25.493. geometry, or both, follow from a struc- If significant changes in structural tural failure or partial failure, the ef- stiffness or geometry, or both, follow fect on damage tolerance must be fur- from a structural failure, or partial ther investigated. failure, the effect on damage tolerance must be further investigated. [Amdt. 25–45, 43 FR 46242, Oct. 5, 1978, as (c) Fatigue (safe-life) evaluation. Com- amended by Amdt. 25–54, 45 FR 60173, Sept. pliance with the damage-tolerance re- 11, 1980; Amdt. 25–72, 55 FR 29776, July 20, quirements of paragraph (b) of this sec- 1990; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; Amdt. 25–96, 63 FR 15714, Mar. 31, 1998; 63 FR tion is not required if the applicant es- 23338, Apr. 28, 1998; Amdt. 25–132, 75 FR 69781, tablishes that their application for par- Nov. 15, 2010] ticular structure is impractical. This structure must be shown by analysis, LIGHTNING PROTECTION supported by test evidence, to be able to withstand the repeated loads of vari- § 25.581 Lightning protection. able magnitude expected during its (a) The airplane must be protected service life without detectable cracks. against catastrophic effects from light- Appropriate safe-life scatter factors must be applied. ning. (d) Sonic fatigue strength. It must be (b) For metallic components, compli- shown by analysis, supported by test ance with paragraph (a) of this section evidence, or by the service history of may be shown by— airplanes of similar structural design (1) Bonding the components properly and sonic excitation environment, to the airframe; or that— (2) Designing the components so that (1) Sonic fatigue cracks are not prob- a strike will not endanger the airplane. able in any part of the flight structure (c) For nonmetallic components, subject to sonic excitation; or compliance with paragraph (a) of this (2) Catastrophic failure caused by section may be shown by— sonic cracks is not probable assuming (1) Designing the components to min- that the loads prescribed in paragraph imize the effect of a strike; or (b) of this section are applied to all (2) Incorporating acceptable means of areas affected by those cracks. diverting the resulting electrical cur- (e) Damage-tolerance (discrete source) rent so as not to endanger the airplane. evaluation. The airplane must be capable of successfully completing a flight [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970]

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Subpart D—Design and limitations of the airplane using nor- Construction mal pilot skill and strength; or (2) Its loss could result in reduction GENERAL in pitch, yaw, or roll control capability or response below that required by § 25.601 General. Subpart B of this chapter. The airplane may not have design (b) The fasteners specified in para- features or details that experience has graph (a) of this section and their lock- shown to be hazardous or unreliable. ing devices may not be adversely af- The suitability of each questionable fected by the environmental conditions design detail and part must be estab- associated with the particular installa- lished by tests. tion. (c) No self-locking nut may be used § 25.603 Materials. on any bolt subject to rotation in oper- The suitability and durability of ma- ation unless a nonfriction locking de- terials used for parts, the failure of vice is used in addition to the self-lock- which could adversely affect safety, ing device. must— (a) Be established on the basis of ex- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] perience or tests; § 25.609 Protection of structure. (b) Conform to approved specifications (such as industry or military Each part of the structure must— specifications, or Technical Standard (a) Be suitably protected against de- Orders) that ensure their having the terioration or loss of strength in serv- strength and other properties assumed ice due to any cause, including— in the design data; and (1) Weathering; (c) Take into account the effects of (2) Corrosion; and environmental conditions, such as tem- (3) Abrasion; and perature and humidity, expected in (b) Have provisions for ventilation service. and drainage where necessary for pro- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tection. amended by Amdt. 25–38, 41 FR 55466, Dec. 20 1976; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978] § 25.611 Accessibility provisions. (a)Means must be provided to allow § 25.605 Fabrication methods. inspection (including inspection of (a) The methods of fabrication used principal structural elements and con- must produce a consistently sound trol systems), replacement of parts structure. If a fabrication process (such normally requiring replacement, ad- as gluing, spot welding, or heat treat- justment, and lubrication as necessary ing) requires close control to reach this for continued airworthiness. The in- objective, the process must be per- spection means for each item must be formed under an approved process spec- practicable for the inspection interval ification. for the item. Nondestructive inspection (b) Each new aircraft fabrication method must be substantiated by a aids may be used to inspect structural test program. elements where it is impracticable to provide means for direct visual inspec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tion if it is shown that the inspection amended by Amdt. 25–46, 43 FR 50595, Oct. 30, is effective and the inspection proce- 1978] dures are specified in the maintenance § 25.607 Fasteners. manual required by § 25.1529. (b) EWIS must meet the accessibility (a) Each removable bolt, screw, nut, requirements of § 25.1719. pin, or other removable fastener must incorporate two separate locking de- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970, as vices if— amended by Amdt. 25–123, 72 FR 63404, Nov. 8, (1) Its loss could preclude continued 2007] flight and landing within the design

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§ 25.613 Material strength properties (a) Uncertain; and material design values. (b) Likely to deteriorate in service (a) Material strength properties must before normal replacement; or be based on enough tests of material (c) Subject to appreciable variability meeting approved specifications to es- because of uncertainties in manufac- tablish design values on a statistical turing processes or inspection methods. basis. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (b) Material design values must be amended by Amdt. 25–23, 35 FR 5674, Apr. 8, chosen to minimize the probability of 1970] structural failures due to material variability. Except as provided in para- § 25.621 Casting factors. graphs (e) and (f) of this section, com- (a) General. The factors, tests, and in- pliance must be shown by selecting ma- spections specified in paragraphs (b) terial design values which assure mate- through (d) of this section must be ap- rial strength with the following prob- plied in addition to those necessary to ability: establish foundry quality control. The (1) Where applied loads are eventu- inspections must meet approved speci- ally distributed through a single mem- fications. Paragraphs (c) and (d) of this ber within an assembly, the failure of section apply to any structural cast- which would result in loss of structural ings except castings that are pressure integrity of the component, 99 percent tested as parts of hydraulic or other probability with 95 percent confidence. fluid systems and do not support struc- (2) For redundant structure, in which tural loads. the failure of individual elements (b) Bearing stresses and surfaces. The would result in applied loads being casting factors specified in paragraphs safely distributed to other load car- (c) and (d) of this section— rying members, 90 percent probability (1) Need not exceed 1.25 with respect with 95 percent confidence. to bearing stresses regardless of the (c) The effects of environmental con- method of inspection used; and ditions, such as temperature and mois- (2) Need not be used with respect to ture, on material design values used in the bearing surfaces of a part whose an essential component or structure bearing factor is larger than the appli- must be considered where these effects cable casting factor. are significant within the airplane op- (c) Critical castings. For each casting erating envelope. whose failure would preclude continued (d) [Reserved] safe flight and landing of the airplane (e) Greater material design values or result in serious injury to occu- may be used if a ‘‘premium selection’’ pants, the following apply: of the material is made in which a (1) Each critical casting must— specimen of each individual item is (i) Have a casting factor of not less tested before use to determine that the than 1.25; and actual strength properties of that par- (ii) Receive 100 percent inspection by ticular item will equal or exceed those visual, radiographic, and magnetic par- used in design. ticle or penetrant inspection methods (f) Other material design values may or approved equivalent nondestructive be used if approved by the Adminis- inspection methods. trator. (2) For each critical casting with a [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as casting factor less than 1.50, three sam- amended by Amdt. 25–46, 43 FR 50595, Oct. 30, ple castings must be static tested and 1978; Amdt. 25–72, 55 FR 29776, July 20, 1990; shown to meet— Amdt. 25–112, 68 FR 46431, Aug. 5, 2003] (i) The strength requirements of § 25.305 at an ultimate load cor- § 25.619 Special factors. responding to a casting factor of 1.25; The factor of safety prescribed in and § 25.303 must be multiplied by the high- (ii) The deformation requirements of est pertinent special factor of safety § 25.305 at a load of 1.15 times the limit prescribed in §§ 25.621 through 25.625 for load. each part of the structure whose (3) Examples of these castings are strength is— structural attachment fittings, parts of

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flight control systems, control surface § 25.625 Fitting factors. hinges and balance weight attach- For each fitting (a part or terminal ments, seat, berth, safety belt, and fuel used to join one structural member to and oil tank supports and attachments, another), the following apply: and cabin pressure valves. (a) For each fitting whose strength is (d) Noncritical castings. For each cast- not proven by limit and ultimate load ing other than those specified in para- tests in which actual stress conditions graph (c) of this section, the following are simulated in the fitting and sur- apply: rounding structures, a fitting factor of (1) Except as provided in paragraphs at least 1.15 must be applied to each (d)(2) and (3) of this section, the casting part of— factors and corresponding inspections (1) The fitting; must meet the following table: (2) The means of attachment; and

Casting factor Inspection (3) The bearing on the joined members. 2.0 or more ...... 100 percent visual. (b) No fitting factor need be used— Less than 2.0 but more 100 percent visual, and magnetic (1) For joints made under approved than 1.5. particle or penetrant or equivalent nondestructive inspection practices and based on comprehensive methods. test data (such as continuous joints in 1.25 through 1.50 ...... 100 percent visual, magnetic par- metal plating, welded joints, and scarf ticle or penetrant, and radiographic, or approved equivalent joints in wood); or nondestructive inspection meth- (2) With respect to any bearing sur- ods. face for which a larger special factor is used. (2) The percentage of castings in- (c) For each integral fitting, the part spected by nonvisual methods may be must be treated as a fitting up to the reduced below that specified in para- point at which the section properties graph (d)(1) of this section when an ap- become typical of the member. proved quality control procedure is es- (d) For each seat, berth, safety belt, tablished. and harness, the fitting factor specified (3) For castings procured to a speci- in § 25.785(f)(3) applies. fication that guarantees the mechanical properties of the material in the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5674, Apr. 8, casting and provides for demonstration 1970; Amdt. 25–72, 55 FR 29776, July 20, 1990] of these properties by test of coupons cut from the castings on a sampling § 25.629 Aeroelastic stability require- basis— ments. (i) A casting factor of 1.0 may be (a) General. The aeroelastic stability used; and evaluations required under this section (ii) The castings must be inspected as include flutter, divergence, control re- provided in paragraph (d)(1) of this sec- versal and any undue loss of stability tion for casting factors of ‘‘1.25 through and control as a result of structural de- 1.50’’ and tested under paragraph (c)(2) formation. The aeroelastic evaluation of this section. must include whirl modes associated with any propeller or rotating device § 25.623 Bearing factors. that contributes significant dynamic (a) Except as provided in paragraph forces. Compliance with this section (b) of this section, each part that has must be shown by analyses, wind tun- nel tests, ground vibration tests, flight clearance (free fit), and that is subject tests, or other means found necessary to pounding or vibration, must have a by the Administrator. bearing factor large enough to provide (b) Aeroelastic stability envelopes. The for the effects of normal relative mo- airplane must be designed to be free tion. from aeroelastic instability for all con- (b) No bearing factor need be used for figurations and design conditions with- a part for which any larger special fac- in the aeroelastic stability envelopes tor is prescribed. as follows:

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(1) For normal conditions without ternally mounted aerodynamic body failures, malfunctions, or adverse con- (such as an external fuel tank). ditions, all combinations of altitudes (5) For airplanes with engines that and speeds encompassed by the VD/MD have propellers or large rotating de- versus altitude envelope enlarged at all vices capable of significant dynamic points by an increase of 15 percent in forces, any single failure of the engine equivalent airspeed at both constant structure that would reduce the rigid- Mach number and constant altitude. In ity of the rotational axis. addition, a proper margin of stability (6) The absence of aerodynamic or gy- must exist at all speeds up to VD/MD roscopic forces resulting from the most and, there must be no large and rapid adverse combination of feathered pro- reduction in stability as VD/MD is ap- pellers or other rotating devices capa- proached. The enlarged envelope may ble of significant dynamic forces. In be limited to Mach 1.0 when MD is less addition, the effect of a single feath- than 1.0 at all design altitudes, and ered propeller or rotating device must (2) For the conditions described in be coupled with the failures of para- § 25.629(d) below, for all approved alti- graphs (d)(4) and (d)(5) of this section. tudes, any airspeed up to the greater (7) Any single propeller or rotating airspeed defined by; device capable of significant dynamic (i) The VD/MD envelope determined by forces rotating at the highest likely § 25.335(b); or, overspeed. (ii) An altitude-airspeed envelope de- (8) Any damage or failure condition, fined by a 15 percent increase in equiv- required or selected for investigation alent airspeed above VC at constant al- by § 25.571. The single structural fail- titude, from sea level to the altitude of ures described in paragraphs (d)(4) and the intersection of 1.15 VC with the ex- (d)(5) of this section need not be consid- tension of the constant cruise Mach ered in showing compliance with this number line, MC, then a linear vari- section if; ation in equivalent airspeed to MC+.05 (i) The structural element could not at the altitude of the lowest VC/MC fail due to discrete source damage re- intersection; then, at higher altitudes, sulting from the conditions described up to the maximum flight altitude, the in § 25.571(e), and boundary defined by a .05 Mach in- (ii) A damage tolerance investigation crease in MC at constant altitude. in accordance with § 25.571(b) shows (c) Balance weights. If concentrated that the maximum extent of damage balance weights are used, their effec- assumed for the purpose of residual tiveness and strength, including sup- strength evaluation does not involve porting structure, must be substan- complete failure of the structural ele- tiated. ment. (d) Failures, malfunctions, and adverse (9) Any damage, failure, or malfunc- conditions. The failures, malfunctions, tion considered under §§ 25.631, 25.671, and adverse conditions which must be 25.672, and 25.1309. considered in showing compliance with (10) Any other combination of fail- this section are: ures, malfunctions, or adverse condi- (1) Any critical fuel loading conditions not shown to be extremely im- tions, not shown to be extremely improbable. probable, which may result from mis- (e) Flight flutter testing. Full scale management of fuel. flight flutter tests at speeds up to VDF/ (2) Any single failure in any flutter MDF must be conducted for new type damper system. designs and for modifications to a type (3) For airplanes not approved for op- design unless the modifications have eration in icing conditions, the max- been shown to have an insignificant ef- imum likely ice accumulation expected fect on the aeroelastic stability. These as a result of an inadvertent encounter. tests must demonstrate that the air- (4) Failure of any single element of plane has a proper margin of damping the structure supporting any engine, at all speeds up to VDF/MDF, and that independently mounted propeller shaft, there is no large and rapid reduction in large auxiliary power unit, or large ex- damping as VDF/MDF, is approached. If a 424

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failure, malfunction, or adverse condi- which the airplane is shown to meet tion is simulated during flight test in the trim requirements of § 25.161. showing compliance with paragraph (d) of this section, the maximum speed in- § 25.657 Hinges. vestigated need not exceed VFC/MFC if it (a) For control surface hinges, in- is shown, by correlation of the flight cluding ball, roller, and self-lubricated test data with other test data or anal- bearing hinges, the approved rating of yses, that the airplane is free from any the bearing may not be exceeded. For aeroelastic instability at all speeds nonstandard bearing hinge configura- within the altitude-airspeed envelope tions, the rating must be established described in paragraph (b)(2) of this on the basis of experience or tests and, section. in the absence of a rational investiga- [Doc. No. 26007, 57 FR 28949, June 29, 1992] tion, a factor of safety of not less than 6.67 must be used with respect to the § 25.631 Bird strike damage. ultimate bearing strength of the softest material used as a bearing. The empennage structure must be de- (b) Hinges must have enough signed to assure capability of contin- strength and rigidity for loads parallel ued safe flight and landing of the air- to the hinge line. plane after impact with an 8-pound bird when the velocity of the airplane (rel- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] ative to the bird along the airplane’s CONTROL SYSTEMS flight path) is equal to VC at sea level, selected under § 25.335(a). Compliance with this section by provision of redun- § 25.671 General. dant structure and protected location (a) Each control and control system of control system elements or protec- must operate with the ease, smooth- tive devices such as splitter plates or ness, and positiveness appropriate to energy absorbing material is accept- its function. able. Where compliance is shown by (b) Each element of each flight con- analysis, tests, or both, use of data on trol system must be designed, or dis- airplanes having similar structural de- tinctively and permanently marked, to sign is acceptable. minimize the probability of incorrect assembly that could result in the mal- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] functioning of the system. CONTROL SURFACES (c) The airplane must be shown by analysis, tests, or both, to be capable § 25.651 Proof of strength. of continued safe flight and landing after any of the following failures or (a) Limit load tests of control sur- jamming in the flight control system faces are required. These tests must in- and surfaces (including trim, lift, drag, clude the horn or fitting to which the and feel systems), within the normal control system is attached. flight envelope, without requiring ex- (b) Compliance with the special fac- ceptional piloting skill or strength. tors requirements of §§ 25.619 through Probable malfunctions must have only 25.625 and 25.657 for control surface minor effects on control system oper- hinges must be shown by analysis or ation and must be capable of being individual load tests. readily counteracted by the pilot. (1) Any single failure, excluding jam- § 25.655 Installation. ming (for example, disconnection or (a) Movable tail surfaces must be in- failure of mechanical elements, or stalled so that there is no interference structural failure of hydraulic compo- between any surfaces when one is held nents, such as actuators, control spool in its extreme position and the others housing, and valves). are operated through their full angular (2) Any combination of failures not movement. shown to be extremely improbable, ex- (b) If an adjustable stabilizer is used, cluding jamming (for example, dual it must have stops that will limit its electrical or hydraulic system failures, range of travel to the maximum for or any single failure in combination

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with any probable hydraulic or elec- at any speed or altitude within the ap- trical failure). proved operating limitations that is (3) Any jam in a control position nor- critical for the type of failure being mally encountered during takeoff, considered; climb, cruise, normal turns, descent, (2) The controllability and maneuver- and landing unless the jam is shown to ability requirements of this part are be extremely improbable, or can be al- met within a practical operational leviated. A runaway of a flight control flight envelope (for example, speed, al- to an adverse position and jam must be titude, normal acceleration, and air- accounted for if such runaway and sub- plane configurations) which is de- sequent jamming is not extremely im- scribed in the Airplane Flight Manual; probable. and (d) The airplane must be designed so (3) The trim, stability, and stall char- that it is controllable if all engines acteristics are not impaired below a fail. Compliance with this requirement level needed to permit continued safe may be shown by analysis where that flight and landing. method has been shown to be reliable. [Amdt. 25–23, 35 FR 5675 Apr. 8, 1970] [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5674, Apr. 8, § 25.675 Stops. 1970] (a) Each control system must have § 25.672 Stability augmentation and stops that positively limit the range of automatic and power-operated sys- motion of each movable aerodynamic tems. surface controlled by the system. If the functioning of stability aug- (b) Each stop must be located so that mentation or other automatic or wear, slackness, or take-up adjust- power-operated systems is necessary to ments will not adversely affect the show compliance with the flight char- control characteristics of the airplane acteristics requirements of this part, because of a change in the range of sur- such systems must comply with § 25.671 face travel. and the following: (c) Each stop must be able to with- (a) A warning which is clearly distin- stand any loads corresponding to the guishable to the pilot under expected design conditions for the control sys- flight conditions without requiring his tem. attention must be provided for any [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as failure in the stability augmentation amended by Amdt. 25–38, 41 FR 55466, Dec. 20, system or in any other automatic or 1976] power-operated system which could result in an unsafe condition if the pilot § 25.677 Trim systems. were not aware of the failure. Warning (a) Trim controls must be designed to systems must not activate the control prevent inadvertent or abrupt oper- systems. ation and to operate in the plane, and (b) The design of the stability aug- with the sense of motion, of the air- mentation system or of any other auto- plane. matic or power-operated system must (b) There must be means adjacent to permit initial counteraction of failures the trim control to indicate the direc- of the type specified in § 25.671(c) with- tion of the control movement relative out requiring exceptional pilot skill or to the airplane motion. In addition, strength, by either the deactivation of there must be clearly visible means to the system, or a failed portion thereof, indicate the position of the trim device or by overriding the failure by move- with respect to the range of adjustment of the flight controls in the nor- ment. The indicator must be clearly mal sense. marked with the range within which it (c) It must be shown that after any has been demonstrated that takeoff is single failure of the stability aug- safe for all center of gravity positions mentation system or any other auto- approved for takeoff. matic or power-operated system— (c) Trim control systems must be de- (1) The airplane is safely controllable signed to prevent creeping in flight. when the failure or malfunction occurs Trim tab controls must be irreversible

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unless the tab is appropriately bal- (c) Excessive deflection. anced and shown to be free from flutter. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5675, Apr. 8, (d) If an irreversible tab control sys- 1970] tem is used, the part from the tab to the attachment of the irreversible unit § 25.685 Control system details. to the airplane structure must consist of a rigid connection. (a) Each detail of each control system must be designed and installed to [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as prevent jamming, chafing, and inter- amended by Amdt. 25–23, 35 FR 5675, Apr. 8, ference from cargo, passengers, loose 1970; Amdt. 25–115, 69 FR 40527, July 2, 2004] objects, or the freezing of moisture. § 25.679 Control system gust locks. (b) There must be means in the cockpit to prevent the entry of foreign ob- (a) There must be a device to prevent jects into places where they would jam damage to the control surfaces (including tabs), and to the control system, the system. from gusts striking the airplane while (c) There must be means to prevent it is on the ground or water. If the de- the slapping of cables or tubes against vice, when engaged, prevents normal other parts. operation of the control surfaces by the (d) Sections 25.689 and 25.693 apply to pilot, it must— cable systems and joints. (1) Automatically disengage when the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as pilot operates the primary flight con- amended by Amdt. 25–38, 41 FR 55466, Dec. 20, trols in a normal manner; or 1976] (2) Limit the operation of the airplane so that the pilot receives unmis- § 25.689 Cable systems. takable warning at the start of takeoff. (a) Each cable, cable fitting, turn- (b) The device must have means to buckle, splice, and pulley must be ap- preclude the possibility of it becoming proved. In addition— inadvertently engaged in flight. (1) No cable smaller than 1⁄8 inch in § 25.681 Limit load static tests. diameter may be used in the aileron, elevator, or rudder systems; and (a) Compliance with the limit load requirements of this Part must be (2) Each cable system must be de- shown by tests in which— signed so that there will be no haz- (1) The direction of the test loads ardous change in cable tension produces the most severe loading in the throughout the range of travel under control system; and operating conditions and temperature (2) Each fitting, pulley, and bracket variations. used in attaching the system to the (b) Each kind and size of pulley must main structure is included. correspond to the cable with which it is (b) Compliance must be shown (by used. Pulleys and sprockets must have analyses or individual load tests) with closely fitted guards to prevent the ca- the special factor requirements for bles and chains from being displaced or control system joints subject to angu- fouled. Each pulley must lie in the lar motion. plane passing through the cable so that the cable does not rub against the pul- § 25.683 Operation tests. ley flange. It must be shown by operation tests (c) Fairleads must be installed so that when portions of the control sys- that they do not cause a change in tem subject to pilot effort loads are cable direction of more than three de- loaded to 80 percent of the limit load grees. specified for the system and the pow- (d) Clevis pins subject to load or mo- ered portions of the control system are tion and retained only by cotter pins loaded to the maximum load expected may not be used in the control system. in normal operation, the system is free (e) Turnbuckles must be attached to from— parts having angular motion in a man- (a) Jamming; ner that will positively prevent binding (b) Excessive friction; and throughout the range of travel.

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(f) There must be provisions for vis- § 25.699 Lift and drag device indicator. ual inspection of fairleads, pulleys, ter- (a) There must be means to indicate minals, and turnbuckles. to the pilots the position of each lift or drag device having a separate control § 25.693 Joints. in the cockpit to adjust its position. In Control system joints (in push-pull addition, an indication of unsymmet- systems) that are subject to angular rical operation or other malfunction in motion, except those in ball and roller the lift or drag device systems must be bearing systems, must have a special provided when such indication is nec- factor of safety of not less than 3.33 essary to enable the pilots to prevent with respect to the ultimate bearing or counteract an unsafe flight or strength of the softest material used as ground condition, considering the ef- a bearing. This factor may be reduced fects on flight characteristics and per- to 2.0 for joints in cable control sys- formance. tems. For ball or roller bearings, the (b) There must be means to indicate approved ratings may not be exceeded. to the pilots the takeoff, en route, approach, and landing lift device posi- [Amdt. 25–72, 55 FR 29777, July 20, 1990] tions. (c) If any extension of the lift and § 25.697 Lift and drag devices, con- drag devices beyond the landing posi- trols. tion is possible, the controls must be (a) Each lift device control must be clearly marked to identify this range designed so that the pilots can place of extension. the device in any takeoff, en route, ap- [Amdt. 25–23, 35 FR 5675, Apr. 8, 1970] proach, or landing position established under § 25.101(d). Lift and drag devices § 25.701 Flap and slat interconnection. must maintain the selected positions, (a) Unless the airplane has safe flight except for movement produced by an characteristics with the flaps or slats automatic positioning or load limiting retracted on one side and extended on device, without further attention by the other, the motion of flaps or slats the pilots. on opposite sides of the plane of sym- (b) Each lift and drag device control metry must be synchronized by a me- must be designed and located to make chanical interconnection or approved inadvertent operation improbable. Lift equivalent means. and drag devices intended for ground (b) If a wing flap or slat interconnec- operation only must have means to tion or equivalent means is used, it prevent the inadvertant operation of must be designed to account for the ap- their controls in flight if that oper- plicable unsymmetrical loads, includ- ation could be hazardous. ing those resulting from flight with the (c) The rate of motion of the surfaces engines on one side of the plane of sym- in response to the operation of the con- metry inoperative and the remaining trol and the characteristics of the engines at takeoff power. automatic positioning or load limiting (c) For airplanes with flaps or slats device must give satisfactory flight that are not subjected to slipstream and performance characteristics under conditions, the structure must be de- steady or changing conditions of air- signed for the loads imposed when the speed, engine power, and airplane atti- wing flaps or slats on one side are car- tude. rying the most severe load occurring in the prescribed symmetrical conditions (d) The lift device control must be and those on the other side are car- designed to retract the surfaces from rying not more than 80 percent of that the fully extended position, during load. steady flight at maximum continuous (d) The interconnection must be de- engine power at any speed below VF signed for the loads resulting when +9.0 (knots). interconnected flap or slat surfaces on [Amdt. 25–23, 35 FR 5675, Apr. 8, 1970, as one side of the plane of symmetry are amended by Amdt. 25–46, 43 FR 50595, Oct. 30, jammed and immovable while the sur- 1978; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] faces on the other side are free to move

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and the full power of the surface actu- pilots seats, of 10 seats or more, the ating system is applied. spillage of enough fuel from any part of [Amdt. 25–72, 55 FR 29777, July 20, 1990] the fuel system to constitute a fire hazard. § 25.703 Takeoff warning system. (b) Each airplane that has a pas- A takeoff warning system must be in- senger seating configuration excluding stalled and must meet the following re- pilots seats, of 10 seats or more must quirements: be designed so that with the airplane (a) The system must provide to the under control it can be landed on a pilots an aural warning that is auto- paved runway with any one or more matically activated during the initial landing gear legs not extended without portion of the takeoff roll if the air- sustaining a structural component fail- plane is in a configuration, including ure that is likely to cause the spillage any of the following, that would not of enough fuel to constitute a fire haz- allow a safe takeoff: ard. (1) The wing flaps or leading edge de- (c) Compliance with the provisions of vices are not within the approved range this section may be shown by analysis of takeoff positions. or tests, or both. (2) Wing spoilers (except lateral control spoilers meeting the requirements [Amdt. 25–32, 37 FR 3969, Feb. 24, 1972] of § 25.671), speed brakes, or longitudinal trim devices are in a position § 25.723 Shock absorption tests. that would not allow a safe takeoff. (a) The analytical representation of (b) The warning required by para- the landing gear dynamic characteris- graph (a) of this section must continue tics that is used in determining the until— landing loads must be validated by en- (1) The configuration is changed to ergy absorption tests. A range of tests allow a safe takeoff; must be conducted to ensure that the (2) Action is taken by the pilot to analytical representation is valid for terminate the takeoff roll; the design conditions specified in (3) The airplane is rotated for takeoff; § 25.473. or (1) The configurations subjected to (4) The warning is manually deacti- energy absorption tests at limit design vated by the pilot. conditions must include at least the (c) The means used to activate the design landing weight or the design system must function properly takeoff weight, whichever produces the throughout the ranges of takeoff greater value of landing impact energy. weights, altitudes, and temperatures for which certification is requested. (2) The test attitude of the landing gear unit and the application of appro- [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978] priate drag loads during the test must simulate the airplane landing condi- LANDING GEAR tions in a manner consistent with the § 25.721 General. development of rational or conservative limit loads. (a) The main landing gear system (b) The landing gear may not fail in must be designed so that if it fails due to overloads during takeoff and landing a test, demonstrating its reserve en- (assuming the overloads to act in the ergy absorption capacity, simulating a upward and aft directions), the failure descent velocity of 12 f.p.s. at design mode is not likely to cause— landing weight, assuming airplane lift (1) For airplanes that have passenger not greater than airplane weight act- seating configuration, excluding pilots ing during the landing impact. seats, of nine seats or less, the spillage (c) In lieu of the tests prescribed in of enough fuel from any fuel system in this section, changes in previously ap- the fuselage to constitute a fire hazard; proved design weights and minor and changes in design may be substantiated (2) For airplanes that have a pas- by analyses based on previous tests senger seating configuration, excluding conducted on the same basic landing

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gear system that has similar energy (c) Emergency operation. There must absorption characteristics. be an emergency means for extending the landing gear in the event of— [Doc. No. 1999–5835, 66 FR 27394, May 16, 2001] (1) Any reasonably probable failure in §§ 25.725–25.727 [Reserved] the normal retraction system; or (2) The failure of any single source of § 25.729 Retracting mechanism. hydraulic, electric, or equivalent en- (a) General. For airplanes with re- ergy supply. tractable landing gear, the following (d) Operation test. The proper func- apply: tioning of the retracting mechanism (1) The landing gear retracting mech- must be shown by operation tests. anism, wheel well doors, and sup- (e) Position indicator and warning de- porting structure, must be designed vice. If a retractable landing gear is for— used, there must be a landing gear po- (i) The loads occurring in the flight sition indicator easily visible to the conditions when the gear is in the re- pilot or to the appropriate crew mem- tracted position, bers (as well as necessary devices to ac- (ii) The combination of friction tuate the indicator) to indicate with- loads, inertia loads, brake torque loads, out ambiguity that the retractable air loads, and gyroscopic loads result- units and their associated doors are se- ing from the wheels rotating at a pe- cured in the extended (or retracted) po- ripheral speed equal to 1.23VSR (with sition. The means must be designed as the wing-flaps in take-off position at follows: design take-off weight), occurring dur- (1) If switches are used, they must be ing retraction and extension at any located and coupled to the landing gear airspeed up to 1.5 VSR1 (with the wing- mechanical systems in a manner that flaps in the approach position at design prevents an erroneous indication of landing weight), and ‘‘down and locked’’ if the landing gear (iii) Any load factor up to those spec- is not in a fully extended position, or of ified in § 25.345(a) for the wing-flaps ex- ‘‘up and locked’’ if the landing gear is tended condition. not in the fully retracted position. The (2) Unless there are other means to switches may be located where they decelerate the airplane in flight at this are operated by the actual landing gear speed, the landing gear, the retracting locking latch or device. mechanism, and the airplane structure (2) The flightcrew must be given an (including wheel well doors) must be aural warning that functions continu- designed to withstand the flight loads ously, or is periodically repeated, if a occurring with the landing gear in the landing is attempted when the landing extended position at any speed up to gear is not locked down. 0.67 V C. (3) The warning must be given in suf- (3) Landing gear doors, their oper- ficient time to allow the landing gear ating mechanism, and their supporting to be locked down or a go-around to be structures must be designed for the made. yawing maneuvers prescribed for the airplane in addition to the conditions (4) There must not be a manual shut- of airspeed and load factor prescribed off means readily available to the in paragraphs (a)(1) and (2) of this sec- flightcrew for the warning required by tion. paragraph (e)(2) of this section such (b) Landing gear lock. There must be that it could be operated instinctively, positive means to keep the landing inadvertently, or by habitual reflexive gear extended in flight and on the action. ground. There must be positive means (5) The system used to generate the to keep the landing gear and doors in aural warning must be designed to the correct retracted position in flight, minimize false or inappropriate alerts. unless it can be shown that lowering of (6) Failures of systems used to in- the landing gear or doors, or flight hibit the landing gear aural warning, with the landing gear or doors ex- that would prevent the warning system tended, at any speed, is not hazardous. from operating, must be improbable.

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(7) A flightcrew alert must be pro- by the Administrator that is not ex- vided whenever the landing gear posi- ceeded under— tion is not consistent with the landing (1) The loads on the main wheel tire, gear selector lever position. corresponding to the most critical (f) Protection of equipment on landing combination of airplane weight (up to gear and in wheel wells. Equipment that maximum weight) and center of grav- is essential to the safe operation of the ity position, and airplane and that is located on the (2) The loads corresponding to the landing gear and in wheel wells must ground reactions in paragraph (b) of be protected from the damaging effects this section, on the nose wheel tire, ex- of— cept as provided in paragraphs (b)(2) (1) A bursting tire; and (b)(3) of this section. (2) A loose tire tread, unless it is (b) The applicable ground reactions shown that a loose tire tread cannot for nose wheel tires are as follows: cause damage. (1) The static ground reaction for the (3) Possible wheel brake tempera- tire corresponding to the most critical tures. combination of airplane weight (up to maximum ramp weight) and center of [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as gravity position with a force of 1.0g amended by Amdt. 25–23, 35 FR 5676, Apr. 8, acting downward at the center of grav- 1970; Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; Amdt. 25–72, 55 FR 29777, July 20, 1990; Amdt. ity. This load may not exceed the load 25–75, 56 FR 63762, Dec. 5, 1991; Amdt. 25–136, rating of the tire. 77 FR 1617, Jan. 11, 2012] (2) The ground reaction of the tire corresponding to the most critical § 25.731 Wheels. combination of airplane weight (up to (a) Each main and nose wheel must maximum landing weight) and center be approved. of gravity position combined with (b) The maximum static load rating forces of 1.0g downward and 0.31g for- of each wheel may not be less than the ward acting at the center of gravity. corresponding static ground reaction The reactions in this case must be dis- with— tributed to the nose and main wheels (1) Design maximum weight; and by the principles of statics with a drag (2) Critical center of gravity. reaction equal to 0.31 times the vertical load at each wheel with brakes (c) The maximum limit load rating of capable of producing this ground reac- each wheel must equal or exceed the tion. This nose tire load may not ex- maximum radial limit load determined ceed 1.5 times the load rating of the under the applicable ground load re- tire. quirements of this part. (3) The ground reaction of the tire (d) Overpressure burst prevention. corresponding to the most critical Means must be provided in each wheel combination of airplane weight (up to to prevent wheel failure and tire burst maximum ramp weight) and center of that may result from excessive pressur- gravity position combined with forces ization of the wheel and tire assembly. of 1.0g downward and 0.20g forward act- (e) Braked wheels. Each braked wheel ing at the center of gravity. The reac- must meet the applicable requirements tions in this case must be distributed of § 25.735. to the nose and main wheels by the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as principles of statics with a drag reac- amended by Amdt. 25–72, 55 FR 29777, July 20, tion equal to 0.20 times the vertical 1990; Amdt. 25–107, 67 FR 20420, Apr. 24, 2002] load at each wheel with brakes capable of producing this ground reaction. This § 25.733 Tires. nose tire load may not exceed 1.5 times (a) When a landing gear axle is fitted the load rating of the tire. with a single wheel and tire assembly, (c) When a landing gear axle is fitted the wheel must be fitted with a suit- with more than one wheel and tire as- able tire of proper fit with a speed rat- sembly, such as dual or dual-tandem, ing approved by the Administrator each wheel must be fitted with a suit- that is not exceeded under critical con- able tire of proper fit with a speed rat- ditions and with a load rating approved ing approved by the Administrator

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that is not exceeded under critical con- cause or support a hazardous fire on ditions, and with a load rating ap- the ground or in flight. proved by the Administrator that is (c) Brake controls. The brake controls not exceeded by— must be designed and constructed so (1) The loads on each main wheel that: tire, corresponding to the most critical (1) Excessive control force is not re- combination of airplane weight (up to quired for their operation. maximum weight) and center of grav- (2) If an automatic braking system is ity position, when multiplied by a fac- installed, means are provided to: tor of 1.07; and (i) Arm and disarm the system, and (2) Loads specified in paragraphs (a)(2), (b)(1), (b)(2), and (b)(3) of this (ii) Allow the pilot(s) to override the section on each nose wheel tire. system by use of manual braking. (d) Each tire installed on a retract- (d) Parking brake. The airplane must able landing gear system must, at the have a parking brake control that, maximum size of the tire type expected when selected on, will, without further in service, have a clearance to sur- attention, prevent the airplane from rounding structure and systems that is rolling on a dry and level paved runway adequate to prevent unintended con- when the most adverse combination of tact between the tire and any part of maximum thrust on one engine and up the structure or systems. to maximum ground idle thrust on any, (e) For an airplane with a maximum or all, other engine(s) is applied. The certificated takeoff weight of more control must be suitably located or be than 75,000 pounds, tires mounted on adequately protected to prevent inad- braked wheels must be inflated with vertent operation. There must be indi- dry nitrogen or other gases shown to be cation in the cockpit when the parking inert so that the gas mixture in the brake is not fully released. tire does not contain oxygen in excess (e) Antiskid system. If an antiskid sys- of 5 percent by volume, unless it can be tem is installed: shown that the tire liner material will (1) It must operate satisfactorily over not produce a volatile gas when heated the range of expected runway condi- or that means are provided to prevent tions, without external adjustment. tire temperatures from reaching unsafe (2) It must, at all times, have pri- levels. ority over the automatic braking sys- [Amdt. 25–48, 44 FR 68752, Nov. 29, 1979; Amdt. tem, if installed. 25–72, 55 FR 29777, July 20, 1990, as amended (f) Kinetic energy capacity—(1) Design by Amdt. 25–78, 58 FR 11781, Feb. 26, 1993] landing stop. The design landing stop is an operational landing stop at max- § 25.735 Brakes and braking systems. imum landing weight. The design land- (a) Approval. Each assembly con- ing stop brake kinetic energy absorp- sisting of a wheel(s) and brake(s) must tion requirement of each wheel, brake, be approved. and tire assembly must be determined. (b) Brake system capability. The brake It must be substantiated by dynamom- system, associated systems and compo- eter testing that the wheel, brake and nents must be designed and con- tire assembly is capable of absorbing structed so that: not less than this level of kinetic en- (1) If any electrical, pneumatic, hy- ergy throughout the defined wear draulic, or mechanical connecting or range of the brake. The energy absorp- transmitting element fails, or if any tion rate derived from the airplane single source of hydraulic or other manufacturer’s braking requirements brake operating energy supply is lost, must be achieved. The mean decelera- it is possible to bring the airplane to tion must not be less than 10 fps 2. rest with a braked roll stopping dis- (2) Maximum kinetic energy accelerate- tance of not more than two times that stop. The maximum kinetic energy ac- obtained in determining the landing celerate-stop is a rejected takeoff for distance as prescribed in § 25.125. the most critical combination of air- (2) Fluid lost from a brake hydraulic plane takeoff weight and speed. The ac- system following a failure in, or in the celerate-stop brake kinetic energy ab- vicinity of, the brakes is insufficient to sorption requirement of each wheel,

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brake, and tire assembly must be de- (i) Brake wear indicators. Means must termined. It must be substantiated by be provided for each brake assembly to dynamometer testing that the wheel, indicate when the heat sink is worn to brake, and tire assembly is capable of the permissible limit. The means must absorbing not less than this level of ki- be reliable and readily visible. netic energy throughout the defined (j) Overtemperature burst prevention. wear range of the brake. The energy Means must be provided in each braked absorption rate derived from the air- wheel to prevent a wheel failure, a tire plane manufacturer’s braking require- burst, or both, that may result from ments must be achieved. The mean de- elevated brake temperatures. Addition- celeration must not be less than 6 fps2. ally, all wheels must meet the require- (3) Most severe landing stop. The most ments of § 25.731(d). severe landing stop is a stop at the (k) Compatibility. Compatibility of most critical combination of airplane the wheel and brake assemblies with landing weight and speed. The most se- the airplane and its systems must be vere landing stop brake kinetic energy substantiated. absorption requirement of each wheel, [Doc. No. FAA–1999–6063, 67 FR 20420, Apr. 24, brake, and tire assembly must be de- 2002, as amended by Amdt. 25–108, 67 FR termined. It must be substantiated by 70827, Nov. 26, 2002; 68 FR 1955, Jan. 15, 2003] dynamometer testing that, at the de- clared fully worn limit(s) of the brake § 25.737 Skis. heat sink, the wheel, brake and tire as- Each ski must be approved. The max- sembly is capable of absorbing not less imum limit load rating of each ski than this level of kinetic energy. The must equal or exceed the maximum most severe landing stop need not be limit load determined under the appli- considered for extremely improbable cable ground load requirements of this failure conditions or if the maximum part. kinetic energy accelerate-stop energy is more severe. FLOATS AND HULLS (g) Brake condition after high kinetic energy dynamometer stop(s). Following § 25.751 Main float buoyancy. the high kinetic energy stop dem- Each main float must have— onstration(s) required by paragraph (f) (a) A buoyancy of 80 percent in excess of this section, with the parking brake of that required to support the max- promptly and fully applied for at least imum weight of the seaplane or am- 3 minutes, it must be demonstrated phibian in fresh water; and that for at least 5 minutes from appli- (b) Not less than five watertight com- cation of the parking brake, no condi- partments approximately equal in vol- tion occurs (or has occurred during the ume. stop), including fire associated with the tire or wheel and brake assembly, § 25.753 Main float design. that could prejudice the safe and com- Each main float must be approved plete evacuation of the airplane. and must meet the requirements of (h) Stored energy systems. An indica- § 25.521. tion to the flightcrew of the usable stored energy must be provided if a § 25.755 Hulls. stored energy system is used to show (a) Each hull must have enough wa- compliance with paragraph (b)(1) of tertight compartments so that, with this section. The available stored en- any two adjacent compartments flood- ergy must be sufficient for: ed, the buoyancy of the hull and auxil- (1) At least 6 full applications of the iary floats (and wheel tires, if used) brakes when an antiskid system is not provides a margin of positive stability operating; and great enough to minimize the prob- (2) Bringing the airplane to a com- ability of capsizing in rough, fresh plete stop when an antiskid system is water. operating, under all runway surface (b) Bulkheads with watertight doors conditions for which the airplane is may be used for communication be- certificated. tween compartments.

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PERSONNEL AND CARGO enter the pilot compartment in the ACCOMMODATIONS event that the flightcrew becomes incapacitated. § 25.771 Pilot compartment. [Doc. No. 24344, 55 FR 29777, July 20, 1990, as (a) Each pilot compartment and its amended by Amdt. 25–106, 67 FR 2127, Jan. 15, equipment must allow the minimum 2002] flight crew (established under § 25.1523) to perform their duties without unrea- § 25.773 Pilot compartment view. sonable concentration or fatigue. (a) Nonprecipitation conditions. For (b) The primary controls listed in nonprecipitation conditions, the fol- § 25.779(a), excluding cables and control lowing apply: rods, must be located with respect to the propellers so that no member of the (1) Each pilot compartment must be minimum flight crew (established arranged to give the pilots a suffi- under § 25.1523), or part of the controls, ciently extensive, clear, and undis- lies in the region between the plane of torted view, to enable them to safely rotation of any inboard propeller and perform any maneuvers within the op- the surface generated by a line passing erating limitations of the airplane, in- through the center of the propeller hub cluding taxiing takeoff, approach, and making an angle of five degrees for- landing. ward or aft of the plane of rotation of (2) Each pilot compartment must be the propeller. free of glare and reflection that could (c) If provision is made for a second interfere with the normal duties of the pilot, the airplane must be controllable minimum flight crew (established with equal safety from either pilot under § 25.1523). This must be shown in seat. day and night flight tests under non- (d) The pilot compartment must be precipitation conditions. constructed so that, when flying in (b) Precipitation conditions. For pre- rain or snow, it will not leak in a man- cipitation conditions, the following ner that will distract the crew or harm apply: the structure. (1) The airplane must have a means (e) Vibration and noise characteris- to maintain a clear portion of the tics of cockpit equipment may not windshield, during precipitation condi- interfere with safe operation of the air- tions, sufficient for both pilots to have plane. a sufficiently extensive view along the flight path in normal flight attitudes [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–4, 30 FR 6113, Apr. 30, of the airplane. This means must be de- 1965] signed to function, without continuous attention on the part of the crew, in— § 25.772 Pilot compartment doors. (i) Heavy rain at speeds up to 1.5 VSR1 For an airplane that has a lockable with lift and drag devices retracted; door installed between the pilot com- and partment and the passenger compart- (ii) The icing conditions specified in ment: § 25.1419 if certification for flight in (a) For airplanes with a maximum icing conditions is requested. passenger seating configuration of (2) No single failure of the systems more than 20 seats, the emergency exit used to provide the view required by configuration must be designed so that paragraph (b)(1) of this section must neither crewmembers nor passengers cause the loss of that view by both pi- require use of the flightdeck door in lots in the specified precipitation con- order to reach the emergency exits pro- ditions. vided for them; and (3) The first pilot must have a win- (b) Means must be provided to enable dow that— flight crewmembers to directly enter (i) Is openable under the conditions the passenger compartment from the prescribed in paragraph (b)(1) of this pilot compartment if the cockpit door section when the cabin is not pressur- becomes jammed. ized; (c) There must be an emergency (ii) Provides the view specified in means to enable a flight attendant to paragraph (b)(1) of this section; and

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(iii) Provides sufficient protection (c) Unless it can be shown by analysis from the elements against impairment or tests that the probability of occur- of the pilot’s vision. rence of a critical windshield frag- (4) The openable window specified in mentation condition is of a low order, paragraph (b)(3) of this section need the airplane must have a means to not be provided if it is shown that an minimize the danger to the pilots from area of the transparent surface will re- flying windshield fragments due to bird main clear sufficient for at least one impact. This must be shown for each pilot to land the airplane safely in the transparent pane in the cockpit that— event of— (1) Appears in the front view of the (i) Any system failure or combina- airplane; tion of failures which is not extremely (2) Is inclined 15 degrees or more to improbable, in accordance with the longitudinal axis of the airplane; § 25.1309, under the precipitation condi- and tions specified in paragraph (b)(1) of (3) Has any part of the pane located this section. where its fragmentation will constitute (ii) An encounter with severe hail, a hazard to the pilots. birds, or insects. (d) The design of windshields and (c) Internal windshield and window windows in pressurized airplanes must fogging. The airplane must have a be based on factors peculiar to high al- means to prevent fogging of the inter- titude operation, including the effects nal portions of the windshield and win- of continuous and cyclic pressurization dow panels over an area which would loadings, the inherent characteristics provide the visibility specified in para- of the material used, and the effects of graph (a) of this section under all in- temperatures and temperature dif- ternal and external ambient condi- ferentials. The windshield and window tions, including precipitation condi- panels must be capable of withstanding tions, in which the airplane is intended the maximum cabin pressure differen- to be operated. tial loads combined with critical aero- (d) Fixed markers or other guides dynamic pressure and temperature ef- must be installed at each pilot station fects after any single failure in the in- to enable the pilots to position them- stallation or associated systems. It selves in their seats for an optimum may be assumed that, after a single combination of outside visibility and failure that is obvious to the flight instrument scan. If lighted markers or crew (established under § 25.1523), the guides are used they must comply with cabin pressure differential is reduced the requirements specified in § 25.1381. from the maximum, in accordance with [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as appropriate operating limitations, to amended by Amdt. 25–23, 35 FR 5676, Apr. 8, allow continued safe flight of the air- 1970; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; plane with a cabin pressure altitude of Amdt. 25–72, 55 FR 29778, July 20, 1990; Amdt. not more than 15,000 feet. 25–108, 67 FR 70827, Nov. 26, 2002; Amdt. 25– (e) The windshield panels in front of 121, 72 FR 44669, Aug. 8, 2007; Amdt. 25–136, 77 FR 1618, Jan. 11, 2012] the pilots must be arranged so that, assuming the loss of vision through any § 25.775 Windshields and windows. one panel, one or more panels remain available for use by a pilot seated at a (a) Internal panes must be made of pilot station to permit continued safe nonsplintering material. flight and landing. (b) Windshield panes directly in front of the pilots in the normal conduct of [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as their duties, and the supporting struc- amended by Amdt. 25–23, 35 FR 5676, Apr. 8, tures for these panes, must withstand, 1970; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976] without penetration, the impact of a four-pound bird when the velocity of § 25.777 Cockpit controls. the airplane (relative to the bird along (a) Each cockpit control must be lo- the airplane’s flight path) is equal to cated to provide convenient operation the value of VC, at sea level, selected and to prevent confusion and inad- under § 25.335(a). vertent operation.

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(b) The direction of movement of tions efficiently and without inter- cockpit controls must meet the re- fering with each other. quirements of § 25.779. Wherever prac- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ticable, the sense of motion involved in amended by Amdt. 25–46, 43 FR 50596, Oct. 30, the operation of other controls must 1978] correspond to the sense of the effect of the operation upon the airplane or § 25.779 Motion and effect of cockpit upon the part operated. Controls of a controls. variable nature using a rotary motion Cockpit controls must be designed so must move clockwise from the off posi- that they operate in accordance with tion, through an increasing range, to the following movement and actuation: the full on position. (a) Aerodynamic controls: (c) The controls must be located and (1) Primary. arranged, with respect to the pilots’ seats, so that there is full and unre- Controls Motion and effect stricted movement of each control without interference from the cockpit Aileron ...... Right (clockwise) for right wing down. structure or the clothing of the min- Elevator ...... Rearward for nose up. imum flight crew (established under Rudder ...... Right pedal forward for nose right. § 25.1523) when any member of this flight crew, from 5′2″ to 6′3″ in height, (2) Secondary. is seated with the seat belt and shoulder harness (if provided) fastened. Controls Motion and effect (d) Identical powerplant controls for Flaps (or auxiliary lift Forward for flaps up; rearward for each engine must be located to prevent devices). flaps down. confusion as to the engines they con- Trim tabs (or equiva- Rotate to produce similar rotation of trol. lent). the airplane about an axis parallel to the axis of the control. (e) Wing flap controls and other auxiliary lift device controls must be lo- (b) Powerplant and auxiliary con- cated on top of the pedestal, aft of the trols: throttles, centrally or to the right of (1) Powerplant. the pedestal centerline, and not less than 10 inches aft of the landing gear Controls Motion and effect control. Power or thrust ...... Forward to increase forward thrust (f) The landing gear control must be and rearward to increase rear- located forward of the throttles and ward thrust. must be operable by each pilot when Propellers ...... Forward to increase rpm. seated with seat belt and shoulder har- Mixture ...... Forward or upward for rich. ness (if provided) fastened. Carburetor air heat ...... Forward or upward for cold. Supercharger ...... Forward or upward for low blower. (g) Control knobs must be shaped in For turbosuperchargers, forward, accordance with § 25.781. In addition, upward, or clockwise, to increase the knobs must be of the same color, pressure. and this color must contrast with the color of control knobs for other pur- (2) Auxiliary. poses and the surrounding cockpit. Controls Motion and effect (h) If a flight engineer is required as part of the minimum flight crew (es- Landing gear ...... Down to extend. tablished under § 25.1523), the airplane must have a flight engineer station lo- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as cated and arranged so that the flight amended by Amdt. 25–72, 55 FR 29778, July 20, crewmembers can perform their func- 1990]

§ 25.781 Cockpit control knob shape. Cockpit control knobs must conform to the general shapes (but not necessarily the exact sizes or specific proportions) in the following figure:

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[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–72, 55 FR 29779, July 20, 1990]

§ 25.783 Fuselage doors. tion as a secondary bulkhead under the prescribed failure conditions of part 25. (a) General. This section applies to fu- These doors must meet the require- selage doors, which includes all doors, ments of this section, taking into ac- hatches, openable windows, access pan- count both pressurized and unpres- els, covers, etc., on the exterior of the surized flight, and must be designed as fuselage that do not require the use of follows: tools to open or close. This also applies (1) Each door must have means to to each door or hatch through a pres- safeguard against opening in flight as a sure bulkhead, including any bulkhead result of mechanical failure, or failure that is specifically designed to func- of any single structural element.

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(2) Each door that could be a hazard cated pressurization prevention means if it unlatches must be designed so that if, from every possible position of the unlatching during pressurized and un- door, it will remain open to the extent pressurized flight from the fully closed, that it prevents pressurization or safe- latched, and locked condition is ex- ly close and latch as pressurization tremely improbable. This must be takes place. This must also be shown shown by safety analysis. with any single failure and malfunc- (3) Each element of each door oper- tion, except that— ating system must be designed or, (i) With failures or malfunctions in where impracticable, distinctively and the latching mechanism, it need not permanently marked, to minimize the latch after closing; and probability of incorrect assembly and (ii) With jamming as a result of me- adjustment that could result in a mal- chanical failure or blocking debris, the function. door need not close and latch if it can (4) All sources of power that could be shown that the pressurization loads initiate unlocking or unlatching of any on the jammed door or mechanism door must be automatically isolated would not result in an unsafe condi- from the latching and locking systems tion. prior to flight and it must not be pos- (d) Latching and locking. The latching sible to restore power to the door dur- and locking mechanisms must be de- ing flight. signed as follows: (5) Each removable bolt, screw, nut, (1) There must be a provision to latch pin, or other removable fastener must each door. meet the locking requirements of (2) The latches and their operating § 25.607. mechanism must be designed so that, (6) Certain doors, as specified by under all airplane flight and ground § 25.807(h), must also meet the applica- loading conditions, with the door ble requirements of §§ 25.809 through latched, there is no force or torque 25.812 for emergency exits. tending to unlatch the latches. In addi- (b) Opening by persons. There must be tion, the latching system must include a means to safeguard each door against a means to secure the latches in the opening during flight due to inad- latched position. This means must be vertent action by persons. In addition, independent of the locking system. design precautions must be taken to minimize the possibility for a person to (3) Each door subject to pressuriza- open a door intentionally during flight. tion, and for which the initial opening If these precautions include the use of movement is not inward, must— auxiliary devices, those devices and (i) Have an individual lock for each their controlling systems must be de- latch; signed so that— (ii) Have the lock located as close as (1) No single failure will prevent practicable to the latch; and more than one exit from being opened; (iii) Be designed so that, during pres- and surized flight, no single failure in the (2) Failures that would prevent open- locking system would prevent the ing of the exit after landing are im- locks from restraining the latches nec- probable. essary to secure the door. (c) Pressurization prevention means. (4) Each door for which the initial There must be a provision to prevent opening movement is inward, and pressurization of the airplane to an un- unlatching of the door could result in a safe level if any door subject to pres- hazard, must have a locking means to surization is not fully closed, latched, prevent the latches from becoming dis- and locked. engaged. The locking means must en- (1) The provision must be designed to sure sufficient latching to prevent function after any single failure, or opening of the door even with a single after any combination of failures not failure of the latching mechanism. shown to be extremely improbable. (5) It must not be possible to position (2) Doors that meet the conditions the lock in the locked position if the described in paragraph (h) of this sec- latch and the latching mechanism are tion are not required to have a dedi- not in the latched position.

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(6) It must not be possible to unlatch lighting conditions, or by means of a the latches with the locks in the flashlight or equivalent light source. locked position. Locks must be de- (g) Certain maintenance doors, remov- signed to withstand the limit loads re- able emergency exits, and access panels. sulting from— Some doors not normally opened ex- (i) The maximum operator effort cept for maintenance purposes or emer- when the latches are operated manu- gency evacuation and some access pan- ally; els need not comply with certain para- (ii) The powered latch actuators, if graphs of this section as follows: installed; and (1) Access panels that are not subject (iii) The relative motion between the to cabin pressurization and would not latch and the structural counterpart. be a hazard if open during flight need (7) Each door for which unlatching not comply with paragraphs (a) would not result in a hazard is not re- through (f) of this section, but must quired to have a locking mechanism have a means to prevent inadvertent meeting the requirements of para- opening during flight. graphs (d)(3) through (d)(6) of this sec- (2) Inward-opening removable emer- tion. gency exits that are not normally re- (e) Warning, caution, and advisory in- moved, except for maintenance pur- dications. Doors must be provided with poses or emergency evacuation, and the following indications: flight deck-openable windows need not (1) There must be a positive means to comply with paragraphs (c) and (f) of indicate at each door operator’s station this section. that all required operations to close, (3) Maintenance doors that meet the latch, and lock the door(s) have been conditions of paragraph (h) of this sec- completed. tion, and for which a placard is pro- (2) There must be a positive means vided limiting use to maintenance ac- clearly visible from each operator sta- cess, need not comply with paragraphs tion for any door that could be a haz- (c) and (f) of this section. ard if unlatched to indicate if the door (h) Doors that are not a hazard. For is not fully closed, latched, and locked. the purposes of this section, a door is (3) There must be a visual means on considered not to be a hazard in the un- the flight deck to signal the pilots if latched condition during flight, pro- any door is not fully closed, latched, vided it can be shown to meet all of the and locked. The means must be de- following conditions: signed such that any failure or com- (1) Doors in pressurized compart- bination of failures that would result ments would remain in the fully closed in an erroneous closed, latched, and position if not restrained by the locked indication is improbable for— latches when subject to a pressure (i) Each door that is subject to pres- greater than 1⁄2 psi. Opening by persons, surization and for which the initial either inadvertently or intentionally, opening movement is not inward; or need not be considered in making this (ii) Each door that could be a hazard determination. if unlatched. (2) The door would remain inside the (4) There must be an aural warning airplane or remain attached to the air- to the pilots prior to or during the ini- plane if it opens either in pressurized tial portion of takeoff roll if any door or unpressurized portions of the flight. is not fully closed, latched, and locked, This determination must include the and its opening would prevent a safe consideration of inadvertent and inten- takeoff and return to landing. tional opening by persons during either (f) Visual inspection provision. Each pressurized or unpressurized portions door for which unlatching of the door of the flight. could be a hazard must have a provi- (3) The disengagement of the latches sion for direct visual inspection to de- during flight would not allow depres- termine, without ambiguity, if the surization of the cabin to an unsafe door is fully closed, latched, and level. This safety assessment must in- locked. The provision must be perma- clude the physiological effects on the nent and discernible under operational occupants.

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(4) The open door during flight would (e) Each berth must be designed so not create aerodynamic interference that the forward part has a padded end that could preclude safe flight and board, canvas diaphragm, or equivalent landing. means, that can withstand the static (5) The airplane would meet the load reaction of the occupant when structural design requirements with subjected to the forward inertia force the door open. This assessment must specified in § 25.561. Berths must be free include the aeroelastic stability re- from corners and protuberances likely quirements of § 25.629, as well as the to cause injury to a person occupying strength requirements of subpart C of the berth during emergency conditions. this part. (f) Each seat or berth, and its sup- (6) The unlatching or opening of the porting structure, and each safety belt door must not preclude safe flight and or harness and its anchorage must be landing as a result of interaction with designed for an occupant weight of 170 other systems or structures. pounds, considering the maximum load [Doc. No. 2003–14193, 69 FR 24501, May 3, 2004] factors, inertia forces, and reactions among the occupant, seat, safety belt, § 25.785 Seats, berths, safety belts, and and harness for each relevant flight harnesses. and ground load condition (including (a) A seat (or berth for a nonambu- the emergency landing conditions pre- lant person) must be provided for each scribed in § 25.561). In addition— occupant who has reached his or her (1) The structural analysis and test- second birthday. ing of the seats, berths, and their sup- (b) Each seat, berth, safety belt, har- porting structures may be determined ness, and adjacent part of the airplane by assuming that the critical load in at each station designated as occupi- the forward, sideward, downward, up- able during takeoff and landing must ward, and rearward directions (as de- be designed so that a person making termined from the prescribed flight, proper use of these facilities will not ground, and emergency landing condi- suffer serious injury in an emergency tions) acts separately or using selected landing as a result of the inertia forces combinations of loads if the required specified in §§ 25.561 and 25.562. strength in each specified direction is (c) Each seat or berth must be ap- substantiated. The forward load factor proved. need not be applied to safety belts for (d) Each occupant of a seat that berths. makes more than an 18-degree angle (2) Each pilot seat must be designed with the vertical plane containing the for the reactions resulting from the ap- airplane centerline must be protected plication of the pilot forces prescribed from head injury by a safety belt and in § 25.395. an energy absorbing rest that will sup- (3) The inertia forces specified in port the arms, shoulders, head, and § 25.561 must be multiplied by a factor spine, or by a safety belt and shoulder of 1.33 (instead of the fitting factor pre- harness that will prevent the head scribed in § 25.625) in determining the from contacting any injurious object. strength of the attachment of each Each occupant of any other seat must seat to the structure and each belt or be protected from head injury by a harness to the seat or structure. safety belt and, as appropriate to the (g) Each seat at a flight deck station type, location, and angle of facing of must have a restraint system con- each seat, by one or more of the fol- sisting of a combined safety belt and lowing: shoulder harness with a single-point re- (1) A shoulder harness that will pre- lease that permits the flight deck occu- vent the head from contacting any in- pant, when seated with the restraint jurious object. system fastened, to perform all of the (2) The elimination of any injurious occupant’s necessary flight deck func- object within striking radius of the tions. There must be a means to secure head. each combined restraint system when (3) An energy absorbing rest that will not in use to prevent interference with support the arms, shoulders, head, and the operation of the airplane and with spine. rapid egress in an emergency.

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(h) Each seat located in the pas- ducting the necessary enroute inspec- senger compartment and designated for tion. use during takeoff and landing by a [Amdt. 25–72, 55 FR 29780, July 20, 1990, as flight attendant required by the oper- amended by Amdt. 25–88, 61 FR 57956, Nov. 8, ating rules of this chapter must be: 1996] (1) Near a required floor level emergency exit, except that another loca- § 25.787 Stowage compartments. tion is acceptable if the emergency (a) Each compartment for the stow- egress of passengers would be enhanced age of cargo, baggage, carry-on arti- with that location. A flight attendant cles, and equipment (such as life rafts), seat must be located adjacent to each and any other stowage compartment Type A or B emergency exit. Other must be designed for its placarded max- flight attendant seats must be evenly imum weight of contents and for the critical load distribution at the appro- distributed among the required floor- priate maximum load factors cor- level emergency exits to the extent responding to the specified flight and feasible. ground load conditions, and to the (2) To the extent possible, without emergency landing conditions of compromising proximity to a required § 25.561(b), except that the forces speci- floor level emergency exit, located to fied in the emergency landing condi- provide a direct view of the cabin area tions need not be applied to compart- for which the flight attendant is re- ments located below, or forward, of all sponsible. occupants in the airplane. If the air- (3) Positioned so that the seat will plane has a passenger seating configu- not interfere with the use of a passage- ration, excluding pilots seats, of 10 way or exit when the seat is not in use. seats or more, each stowage compart- (4) Located to minimize the prob- ment in the passenger cabin, except for ability that occupants would suffer in- underseat and overhead compartments jury by being struck by items dislodged for passenger convenience, must be from service areas, stowage compart- completely enclosed. ments, or service equipment. (b) There must be a means to prevent the contents in the compartments from (5) Either forward or rearward facing becoming a hazard by shifting, under with an energy absorbing rest that is the loads specified in paragraph (a) of designed to support the arms, shoul- this section. For stowage compart- ders, head, and spine. ments in the passenger and crew cabin, (6) Equipped with a restraint system if the means used is a latched door, the consisting of a combined safety belt design must take into consideration and shoulder harness unit with a single the wear and deterioration expected in point release. There must be means to service. secure each restraint system when not (c) If cargo compartment lamps are in use to prevent interference with installed, each lamp must be installed rapid egress in an emergency. so as to prevent contact between lamp (i) Each safety belt must be equipped bulb and cargo. with a metal to metal latching device. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (j) If the seat backs do not provide a amended by Amdt. 25–32, 37 FR 3969, Feb. 24, firm handhold, there must be a hand- 1972; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; grip or rail along each aisle to enable Amdt. 25–51, 45 FR 7755, Feb. 4, 1980] persons to steady themselves while § 25.789 Retention of items of mass in using the aisles in moderately rough passenger and crew compartments air. and galleys. (k) Each projecting object that would (a) Means must be provided to pre- injure persons seated or moving about vent each item of mass (that is part of the airplane in normal flight must be the airplane type design) in a passenger padded. or crew compartment or galley from (l) Each forward observer’s seat re- becoming a hazard by shifting under quired by the operating rules must be the appropriate maximum load factors shown to be suitable for use in con- corresponding to the specified flight

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and ground load conditions, and to the § 25.795 Security considerations. emergency landing conditions of (a) Protection of flightcrew compart- § 25.561(b). ment. If a flightdeck door is required by (b) Each interphone restraint system operating rules: must be designed so that when sub- (1) The bulkhead, door, and any other jected to the load factors specified in accessible boundary separating the § 25.561(b)(3), the interphone will re- flightcrew compartment from occupied main in its stowed position. areas must be designed to resist forc- [Amdt. 25–32, 37 FR 3969, Feb. 24, 1972, as ible intrusion by unauthorized persons amended by Amdt. 25–46, 43 FR 50596, Oct. 30, and be capable of withstanding impacts 1978] of 300 joules (221.3 foot pounds). (2) The bulkhead, door, and any other § 25.791 Passenger information signs accessible boundary separating the and placards. flightcrew compartment from occupied (a) If smoking is to be prohibited, areas must be designed to resist a con- there must be at least one placard so stant 250 pound (1,113 Newtons) tensile stating that is legible to each person load on accessible handholds, including seated in the cabin. If smoking is to be the doorknob or handle. allowed, and if the crew compartment (3) The bulkhead, door, and any other is separated from the passenger com- boundary separating the flightcrew partment, there must be at least one compartment from any occupied areas sign notifying when smoking is prohib- must be designed to resist penetration ited. Signs which notify when smoking by small arms fire and fragmentation is prohibited must be operable by a devices to a level equivalent to level member of the flightcrew and, when il- IIIa of the National Institute of Justice luminated, must be legible under all (NIJ) Standard 0101.04. probable conditions of cabin illumina- (b) Airplanes with a maximum cer- tion to each person seated in the cabin. tificated passenger seating capacity of more than 60 persons or a maximum (b) Signs that notify when seat belts certificated takeoff gross weight of should be fastened and that are in- over 100,000 pounds (45,359 Kilograms) stalled to comply with the operating must be designed to limit the effects of rules of this chapter must be operable an explosive or incendiary device as by a member of the flightcrew and, follows: when illuminated, must be legible (1) Flightdeck smoke protection. Means under all probable conditions of cabin must be provided to limit entry of illumination to each person seated in smoke, fumes, and noxious gases into the cabin. the flightdeck. (c) A placard must be located on or (2) Passenger cabin smoke protection. adjacent to the door of each receptacle Means must be provided to prevent pas- used for the disposal of flammable senger incapacitation in the cabin re- waste materials to indicate that use of sulting from smoke, fumes, and nox- the receptacle for disposal of ciga- ious gases as represented by the initial rettes, etc., is prohibited. combined volumetric concentrations of (d) Lavatories must have ‘‘No Smok- 0.59% carbon monoxide and 1.23% car- ing’’ or ‘‘No Smoking in Lavatory’’ bon dioxide. placards conspicuously located on or (3) Cargo compartment fire suppression. adjacent to each side of the entry door. An extinguishing agent must be capa- (e) Symbols that clearly express the ble of suppressing a fire. All cargo- intent of the sign or placard may be compartment fire suppression systems used in lieu of letters. must be designed to withstand the fol- [Amdt. 25–72, 55 FR 29780, July 20, 1990] lowing effects, including support structure displacements or adjacent mate- § 25.793 Floor surfaces. rials displacing against the distribution system: The floor surface of all areas which (i) Impact or damage from a 0.5-inch are likely to become wet in service diameter aluminum sphere traveling at must have slip resistant properties. 430 feet per second (131.1 meters per [Amdt. 25–51, 45 FR 7755, Feb. 4, 1980] second);

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(ii) A 15-pound per square-inch (103.4 other objects from a simple inspection kPa) pressure load if the projected sur- in the following areas of the airplane face area of the component is greater cabin: than 4 square feet. Any single dimen- (i) Areas above the overhead bins sion greater than 4 feet (1.22 meters) must be designed to prevent objects may be assumed to be 4 feet (1.22 me- from being hidden from view in a sim- ters) in length; and ple search from the aisle. Designs that (iii) A 6-inch (0.152 meters) displace- prevent concealment of objects with ment, except where limited by the fu- volumes 20 cubic inches and greater selage contour, from a single point force applied anywhere along the dis- satisfy this requirement. tribution system where relative move- (ii) Toilets must be designed to pre- ment between the system and its at- vent the passage of solid objects great- tachment can occur. er than 2.0 inches in diameter. (iv) Paragraphs (b)(3)(i) through (iii) (iii) Life preservers or their storage of this section do not apply to compo- locations must be designed so that nents that are redundant and separated tampering is evident. in accordance with paragraph (c)(2) of (d) Exceptions. Airplanes used solely this section or are installed remotely to transport cargo only need to meet from the cargo compartment. the requirements of paragraphs (b)(1), (c) An airplane with a maximum cer- (b)(3), and (c)(2) of this section. tificated passenger seating capacity of (e) Material Incorporated by Reference. more than 60 persons or a maximum You must use National Institute of certificated takeoff gross weight of Justice (NIJ) Standard 0101.04, Ballistic over 100,000 pounds (45,359 Kilograms) Resistance of Personal Body Armor, must comply with the following: (1) Least risk bomb location. An air- June 2001, Revision A, to establish bal- plane must be designed with a des- listic resistance as required by para- ignated location where a bomb or other graph (a)(3) of this section. explosive device could be placed to best (1) The Director of the Federal Reg- protect flight-critical structures and ister approved the incorporation by ref- systems from damage in the case of erence of this document under 5 U.S.C. detonation. 552(a) and 1 CFR part 51. (2) Survivability of systems. (i) Except (2) You may review copies of NIJ where impracticable, redundant air- Standard 0101.04 at the: plane systems necessary for continued (i) FAA Transport Airplane Direc- safe flight and landing must be phys- torate, 1601 Lind Avenue, SW., Renton, ically separated, at a minimum, by an Washington 98055; amount equal to a sphere of diameter (ii) National Institute of Justice (NIJ), http://www.ojp.usdoj.gov/nij, tele- = ()π DH2 0 / phone (202) 307–2942; or (iii) National Archives and Records (where H0 is defined under § 25.365(e)(2) Administration (NARA). For informa- of this part and D need not exceed 5.05 tion on the availability of this mate- feet (1.54 meters)). The sphere is ap- rial at NARA go to http:// plied everywhere within the fuselage— www.archives.gov/federallregister/ limited by the forward bulkhead and codeloflfederallregulations/ the aft bulkhead of the passenger cabin ibrllocations.html or call (202) 741–6030. and cargo compartment beyond which (3) You may obtain copies of NIJ only one-half the sphere is applied. Standard 0101.04 from the National (ii) Where compliance with paragraph (c)(2)(i) of this section is impracticable, Criminal Justice Reference Service, other design precautions must be taken P.O. Box 6000, Rockville, MD 20849–6000, to maximize the survivability of those telephone (800) 851–3420. systems. [Amdt. Nos. 25–127; 121–341, 73 FR 63879, Oct. (3) Interior design to facilitate searches. 28, 2008, as amended at 74 FR 22819, May 15, Design features must be incorporated 2009] that will deter concealment or promote discovery of weapons, explosives, or

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EMERGENCY PROVISIONS as with the landing gear retracted, considering the possibility of the airplane § 25.801 Ditching. being on fire. (a) If certification with ditching pro- (b) [Reserved] visions is requested, the airplane must (c) For airplanes having a seating ca- meet the requirements of this section pacity of more than 44 passengers, it and §§ 25.807(e), 25.1411, and 25.1415(a). must be shown that the maximum (b) Each practicable design measure, seating capacity, including the number compatible with the general character- of crewmembers required by the oper- istics of the airplane, must be taken to ating rules for which certification is minimize the probability that in an requested, can be evacuated from the emergency landing on water, the be- airplane to the ground under simulated havior of the airplane would cause im- emergency conditions within 90 sec- mediate injury to the occupants or onds. Compliance with this require- would make it impossible for them to ment must be shown by actual dem- escape. onstration using the test criteria out- (c) The probable behavior of the air- lined in appendix J of this part unless plane in a water landing must be inves- the Administrator finds that a com- tigated by model tests or by compari- bination of analysis and testing will son with airplanes of similar configura- provide data equivalent to that which tion for which the ditching characteris- would be obtained by actual dem- tics are known. Scoops, flaps, projec- onstration. tions, and any other factor likely to af- (d)–(e) [Reserved] fect the hydrodynamic characteristics [Doc. No. 24344, 55 FR 29781, July 20, 1990] of the airplane, must be considered. (d) It must be shown that, under rea- § 25.807 Emergency exits. sonably probable water conditions, the (a) Type. For the purpose of this part, flotation time and trim of the airplane the types of exits are defined as fol- will allow the occupants to leave the lows: airplane and enter the liferafts re- (1) Type I. This type is a floor-level quired by § 25.1415. If compliance with exit with a rectangular opening of not this provision is shown by buoyancy less than 24 inches wide by 48 inches and trim computations, appropriate al- high, with corner radii not greater lowances must be made for probable than eight inches. structural damage and leakage. If the (2) Type II. This type is a rectangular airplane has fuel tanks (with fuel jetti- opening of not less than 20 inches wide soning provisions) that can reasonably by 44 inches high, with corner radii not be expected to withstand a ditching greater than seven inches. Type II exits without leakage, the jettisonable vol- must be floor-level exits unless located ume of fuel may be considered as buoy- over the wing, in which case they must ancy volume. not have a step-up inside the airplane (e) Unless the effects of the collapse of more than 10 inches nor a step-down of external doors and windows are ac- outside the airplane of more than 17 counted for in the investigation of the inches. probable behavior of the airplane in a (3) Type III. This type is a rectan- water landing (as prescribed in para- gular opening of not less than 20 inches graphs (c) and (d) of this section), the wide by 36 inches high with corner external doors and windows must be radii not greater than seven inches, designed to withstand the probable and with a step-up inside the airplane maximum local pressures. of not more than 20 inches. If the exit [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as is located over the wing, the step-down amended by Amdt. 25–72, 55 FR 29781, July 20, outside the airplane may not exceed 27 1990] inches. (4) Type IV. This type is a rectan- § 25.803 Emergency evacuation. gular opening of not less than 19 inches (a) Each crew and passenger area wide by 26 inches high, with corner must have emergency means to allow radii not greater than 6.3 inches, lo- rapid evacuation in crash landings, cated over the wing, with a step-up in- with the landing gear extended as well side the airplane of not more than 29

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inches and a step-down outside the air- (e) Uniformity. Exits must be distrib- plane of not more than 36 inches. uted as uniformly as practical, taking (5) Ventral. This type is an exit from into account passenger seat distribu- the passenger compartment through tion. the pressure shell and the bottom fuse- (f) Location. (1) Each required pas- lage skin. The dimensions and physical senger emergency exit must be acces- configuration of this type of exit must sible to the passengers and located allow at least the same rate of egress where it will afford the most effective as a Type I exit with the airplane in means of passenger evacuation. the normal ground attitude, with land- (2) If only one floor-level exit per side ing gear extended. is prescribed, and the airplane does not (6) Tailcone. This type is an aft exit have a tailcone or ventral emergency from the passenger compartment exit, the floor-level exits must be in through the pressure shell and through the rearward part of the passenger an openable cone of the fuselage aft of compartment unless another location the pressure shell. The means of open- affords a more effective means of pas- ing the tailcone must be simple and ob- senger evacuation. vious and must employ a single oper- (3) If more than one floor-level exit ation. per side is prescribed, and the airplane (7) Type A. This type is a floor-level does not have a combination cargo and exit with a rectangular opening of not passenger configuration, at least one less than 42 inches wide by 72 inches floor-level exit must be located in each high, with corner radii not greater side near each end of the cabin. than seven inches. (4) For an airplane that is required to have more than one passenger emer- (8) Type B. This type is a floor-level gency exit for each side of the fuselage, exit with a rectangular opening of not no passenger emergency exit shall be less than 32 inches wide by 72 inches more than 60 feet from any adjacent high, with corner radii not greater passenger emergency exit on the same than six inches. side of the same deck of the fuselage, (9) Type C. This type is a floor-level as measured parallel to the airplane’s exit with a rectangular opening of not longitudinal axis between the nearest less than 30 inches wide by 48 inches exit edges. high, with corner radii not greater (g) Type and number required. The than 10 inches. maximum number of passenger seats (b) Step down distance. Step down dis- permitted depends on the type and tance, as used in this section, means number of exits installed in each side the actual distance between the bot- of the fuselage. Except as further re- tom of the required opening and a usa- stricted in paragraphs (g)(1) through ble foot hold, extending out from the (g)(9) of this section, the maximum fuselage, that is large enough to be ef- number of passenger seats permitted fective without searching by sight or for each exit of a specific type installed feel. in each side of the fuselage is as fol- (c) Over-sized exits. Openings larger lows: than those specified in this section, Type A 110 whether or not of rectangular shape, Type B 75 may be used if the specified rectan- Type C 55 gular opening can be inscribed within Type I 45 the opening and the base of the in- Type II 40 scribed rectangular opening meets the Type III 35 specified step-up and step-down Type IV 9 heights. (1) For a passenger seating configura- (d) Asymmetry. Exits of an exit pair tion of 1 to 9 seats, there must be at need not be diametrically opposite least one Type IV or larger overwing each other nor of the same size; how- exit in each side of the fuselage or, if ever, the number of passenger seats overwing exits are not provided, at permitted under paragraph (g) of this least one exit in each side that meets section is based on the smaller of the the minimum dimensions of a Type III two exits. exit.

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(2) For a passenger seating configura- less than 56 inches from the passenger tion of more than 9 seats, each exit compartment floor, 15 additional pas- must be a Type III or larger exit. senger seats. (3) For a passenger seating configura- (h) Other exits. The following exits tion of 10 to 19 seats, there must be at also must meet the applicable emer- least one Type III or larger exit in each gency exit requirements of §§ 25.809 side of the fuselage. through 25.812, and must be readily ac- (4) For a passenger seating configura- cessible: tion of 20 to 40 seats, there must be at (1) Each emergency exit in the pas- least two exits, one of which must be a senger compartment in excess of the Type II or larger exit, in each side of minimum number of required emer- the fuselage. gency exits. (5) For a passenger seating configura- (2) Any other floor-level door or exit tion of 41 to 110 seats, there must be at that is accessible from the passenger least two exits, one of which must be a compartment and is as large or larger Type I or larger exit, in each side of than a Type II exit, but less than 46 the fuselage. inches wide. (6) For a passenger seating configura- (3) Any other ventral or tail cone tion of more than 110 seats, the emer- passenger exit. gency exits in each side of the fuselage (i) Ditching emergency exits for pas- must include at least two Type I or sengers. Whether or not ditching cer- larger exits. tification is requested, ditching emer- (7) The combined maximum number gency exits must be provided in accord- of passenger seats permitted for all ance with the following requirements, Type III exits is 70, and the combined unless the emergency exits required by maximum number of passenger seats paragraph (g) of this section already permitted for two Type III exits in meet them: each side of the fuselage that are sepa- (1) For airplanes that have a pas- rated by fewer than three passenger senger seating configuration of nine or seat rows is 65. fewer seats, excluding pilot seats, one (8) If a Type A, Type B, or Type C exit above the waterline in each side of exit is installed, there must be at least the airplane, meeting at least the di- two Type C or larger exits in each side mensions of a Type IV exit. of the fuselage. (2) For airplanes that have a pas- (9) If a passenger ventral or tailcone senger seating configuration of 10 of exit is installed and that exit provides more seats, excluding pilot seats, one at least the same rate of egress as a exit above the waterline in a side of the Type III exit with the airplane in the airplane, meeting at least the dimen- most adverse exit opening condition sions of a Type III exit for each unit (or that would result from the collapse of part of a unit) of 35 passenger seats, one or more legs of the landing gear, an but no less than two such exits in the increase in the passenger seating con- passenger cabin, with one on each side figuration is permitted as follows: of the airplane. The passenger seat/ (i) For a ventral exit, 12 additional exit ratio may be increased through passenger seats. the use of larger exits, or other means, (ii) For a tailcone exit incorporating provided it is shown that the evacu- a floor level opening of not less than 20 ation capability during ditching has inches wide by 60 inches high, with cor- been improved accordingly. ner radii not greater than seven inches, (3) If it is impractical to locate side in the pressure shell and incorporating exits above the waterline, the side an approved assist means in accordance exits must be replaced by an equal with § 25.810(a), 25 additional passenger number of readily accessible overhead seats. hatches of not less than the dimensions (iii) For a tailcone exit incorporating of a Type III exit, except that for air- an opening in the pressure shell which planes with a passenger configuration is at least equivalent to a Type III of 35 or fewer seats, excluding pilot emergency exit with respect to dimen- seats, the two required Type III side sions, step-up and step-down distance, exits need be replaced by only one and with the top of the opening not overhead hatch.

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(j) Flightcrew emergency exits. For air- tudes corresponding to collapse of one planes in which the proximity of pas- or more legs of the landing gear; and senger emergency exits to the (2) Within 10 seconds measured from flightcrew area does not offer a conven- the time when the opening means is ac- ient and readily accessible means of tuated to the time when the exit is evacuation of the flightcrew, and for fully opened. all airplanes having a passenger seat- (3) Even though persons may be ing capacity greater than 20, flightcrew crowded against the door on the inside exits shall be located in the flightcrew of the airplane. area. Such exits shall be of sufficient (c) The means of opening emergency size and so located as to permit rapid exits must be simple and obvious; may evacuation by the crew. One exit shall not require exceptional effort; and be provided on each side of the air- must be arranged and marked so that plane; or, alternatively, a top hatch it can be readily located and operated, shall be provided. Each exit must en- even in darkness. Internal exit-opening compass an unobstructed rectangular means involving sequence operations opening of at least 19 by 20 inches un- (such as operation of two handles or less satisfactory exit utility can be latches, or the release of safety demonstrated by a typical crew- catches) may be used for flightcrew member. emergency exits if it can be reasonably [Amdt. 25–72, 55 FR 29781, July 20, 1990, as established that these means are sim- amended by Amdt. 25–88, 61 FR 57956, Nov. 8, ple and obvious to crewmembers 1996; 62 FR 1817, Jan. 13, 1997; Amdt. 25–94, 63 trained in their use. FR 8848, Feb. 23, 1998; 63 FR 12862, Mar. 16, (d) If a single power-boost or single 1998; Amdt. 25–114, 69 FR 24502, May 3, 2004] power-operated system is the primary § 25.809 Emergency exit arrangement. system for operating more than one exit in an emergency, each exit must (a) Each emergency exit, including be capable of meeting the requirements each flightcrew emergency exit, must of paragraph (b) of this section in the be a moveable door or hatch in the ex- event of failure of the primary system. ternal walls of the fuselage, allowing Manual operation of the exit (after an unobstructed opening to the out- failure of the primary system) is ac- side. In addition, each emergency exit ceptable. must have means to permit viewing of (e) Each emergency exit must be the conditions outside the exit when shown by tests, or by a combination of the exit is closed. The viewing means analysis and tests, to meet the require- may be on or adjacent to the exit pro- ments of paragraphs (b) and (c) of this vided no obstructions exist between the section. exit and the viewing means. Means must also be provided to permit view- (f) Each door must be located where ing of the likely areas of evacuee persons using them will not be endan- ground contact. The likely areas of gered by the propellers when appro- evacuee ground contact must be priate operating procedures are used. viewable during all lighting conditions (g) There must be provisions to mini- with the landing gear extended as well mize the probability of jamming of the as in all conditions of landing gear col- emergency exits resulting from fuse- lapse. lage deformation in a minor crash (b) Each emergency exit must be landing. openable from the inside and the out- (h) When required by the operating side except that sliding window emer- rules for any large passenger-carrying gency exits in the flight crew area need turbojet-powered airplane, each ven- not be openable from the outside if tral exit and tailcone exit must be— other approved exits are convenient (1) Designed and constructed so that and readily accessible to the flight it cannot be opened during flight; and crew area. Each emergency exit must (2) Marked with a placard readable be capable of being opened, when there from a distance of 30 inches and in- is no fuselage deformation— stalled at a conspicuous location near (1) With the airplane in the normal the means of opening the exit, stating ground attitude and in each of the atti- that the exit has been designed and

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constructed so that it cannot be opened (iii) It must be of such length after during flight. full deployment that the lower end is (i) Each emergency exit must have a self-supporting on the ground and pro- means to retain the exit in the open vides safe evacuation of occupants to position, once the exit is opened in an the ground after collapse of one or emergency. The means must not re- more legs of the landing gear. quire separate action to engage when (iv) It must have the capability, in the exit is opened, and must require 25-knot winds directed from the most positive action to disengage. critical angle, to deploy and, with the assistance of only one person, to re- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as main usable after full deployment to amended by Amdt. 25–15, 32 FR 13264, Sept. 20, 1967; Amdt. 25–32, 37 FR 3970, Feb. 24, 1972; evacuate occupants safely to the Amdt. 25–34, 37 FR 25355, Nov. 30, 1972; Amdt. ground. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. 25–47, (v) For each system installation 44 FR 61325, Oct. 25, 1979; Amdt. 25–72, 55 FR (mockup or airplane installed), five 29782, July 20, 1990; Amdt. 25–114, 69 FR 24502, consecutive deployment and inflation May 3, 2004; Amdt. 25–116, 69 FR 62788, Oct. 27, tests must be conducted (per exit) 2004] without failure, and at least three tests of each such five-test series must be § 25.810 Emergency egress assist conducted using a single representative means and escape routes. sample of the device. The sample de- (a) Each non over-wing Type A, Type vices must be deployed and inflated by B or Type C exit, and any other non the system’s primary means after over-wing landplane emergency exit being subjected to the inertia forces more than 6 feet from the ground with specified in § 25.561(b). If any part of the the airplane on the ground and the system fails or does not function prop- landing gear extended, must have an erly during the required tests, the approved means to assist the occupants cause of the failure or malfunction in descending to the ground. must be corrected by positive means (1) The assisting means for each pas- and after that, the full series of five senger emergency exit must be a self- consecutive deployment and inflation supporting slide or equivalent; and, in tests must be conducted without fail- the case of Type A or Type B exits, it ure. must be capable of carrying simulta- (2) The assisting means for flightcrew neously two parallel lines of evacuees. emergency exits may be a rope or any In addition, the assisting means must other means demonstrated to be suit- be designed to meet the following re- able for the purpose. If the assisting quirements— means is a rope, or an approved device (i) It must be automatically deployed equivalent to a rope, it must be— and deployment must begin during the (i) Attached to the fuselage structure interval between the time the exit at or above the top of the emergency opening means is actuated from inside exit opening, or, for a device at a pi- the airplane and the time the exit is lot’s emergency exit window, at an- fully opened. However, each passenger other approved location if the stowed emergency exit which is also a pas- device, or its attachment, would reduce senger entrance door or a service door the pilot’s view in flight; must be provided with means to pre- (ii) Able (with its attachment) to vent deployment of the assisting means withstand a 400-pound static load. when it is opened from either the in- (b) Assist means from the cabin to side or the outside under non- the wing are required for each type A emergency conditions for normal use. or Type B exit located above the wing (ii) Except for assisting means in- and having a stepdown unless the exit stalled at Type C exits, it must be without an assist-means can be shown automatically erected within 6 seconds to have a rate of passenger egress at after deployment is begun. Assisting least equal to that of the same type of means installed at Type C exits must non over-wing exit. If an assist means be automatically erected within 10 sec- is required, it must be automatically onds from the time the opening means deployed and automatically erected of the exit is actuated. concurrent with the opening of the

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exit. In the case of assist means in- within 10 seconds from the time the stalled at Type C exits, it must be self- opening means of the exit is actuated, supporting within 10 seconds from the and that provided for the escape route time the opening means of the exits is leading from any other exit type must actuated. For all other exit types, it be automatically erected within 10 sec- must be self-supporting 6 seconds after onds after actuation of the erection deployment is begun. system. (c) An escape route must be estab- (e) If an integral stair is installed in lished from each overwing emergency a passenger entry door that is qualified exit, and (except for flap surfaces suit- as a passenger emergency exit, the able as slides) covered with a slip re- stair must be designed so that, under sistant surface. Except where a means the following conditions, the effective- for channeling the flow of evacuees is ness of passenger emergency egress will provided— not be impaired: (1) The escape route from each Type (1) The door, integral stair, and oper- A or Type B passenger emergency exit, ating mechanism have been subjected or any common escape route from two to the inertia forces specified in Type III passenger emergency exits, § 25.561(b)(3), acting separately relative must be at least 42 inches wide; that to the surrounding structure. from any other passenger emergency (2) The airplane is in the normal exit must be at least 24 inches wide; ground attitude and in each of the atti- and tudes corresponding to collapse of one (2) The escape route surface must or more legs of the landing gear. have a reflectance of at least 80 percent, and must be defined by markings [Amdt. 25–72, 55 FR 29782, July 20, 1990, as with a surface-to-marking contrast amended by Amdt. 25–88, 61 FR 57958, Nov. 8, ratio of at least 5:1. 1996; 62 FR 1817, Jan. 13, 1997; Amdt. 25–114, 69 FR 24502, May 3, 2004] (d) Means must be provided to assist evacuees to reach the ground for all § 25.811 Emergency exit marking. Type C exits located over the wing and, if the place on the airplane structure (a) Each passenger emergency exit, at which the escape route required in its means of access, and its means of paragraph (c) of this section termi- opening must be conspicuously nates is more than 6 feet from the marked. ground with the airplane on the ground (b) The identity and location of each and the landing gear extended, for all passenger emergency exit must be rec- other exit types. ognizable from a distance equal to the (1) If the escape route is over the width of the cabin. flap, the height of the terminal edge (c) Means must be provided to assist must be measured with the flap in the the occupants in locating the exits in takeoff or landing position, whichever conditions of dense smoke. is higher from the ground. (d) The location of each passenger (2) The assisting means must be usa- emergency exit must be indicated by a ble and self-supporting with one or sign visible to occupants approaching more landing gear legs collapsed and along the main passenger aisle (or under a 25-knot wind directed from the aisles). There must be— most critical angle. (1) A passenger emergency exit loca- (3) The assisting means provided for tor sign above the aisle (or aisles) near each escape route leading from a Type each passenger emergency exit, or at A or B emergency exit must be capable another overhead location if it is more of carrying simultaneously two par- practical because of low headroom, ex- allel lines of evacuees; and, the assist- cept that one sign may serve more ing means leading from any other exit than one exit if each exit can be seen type must be capable of carrying as readily from the sign; many parallel lines of evacuees as (2) A passenger emergency exit mark- there are required escape routes. ing sign next to each passenger emer- (4) The assisting means provided for gency exit, except that one sign may each escape route leading from a Type serve two such exits if they both can be C exit must be automatically erected seen readily from the sign; and

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(3) A sign on each bulkhead or divider the darker color is 15 percent or less, that prevents fore and aft vision along the reflectance of the lighter color the passenger cabin to indicate emer- must be at least 45 percent. ‘‘Reflec- gency exits beyond and obscured by the tance’’ is the ratio of the luminous flux bulkhead or divider, except that if this reflected by a body to the luminous is not possible the sign may be placed flux it receives. When the reflectance at another appropriate location. of the darker color is greater than 15 (e) The location of the operating han- percent, at least a 30-percent difference dle and instructions for opening exits between its reflectance and the reflec- from the inside of the airplane must be tance of the lighter color must be pro- shown in the following manner: vided. (1) Each passenger emergency exit (3) In the case of exists other than must have, on or near the exit, a mark- those in the side of the fuselage, such ing that is readable from a distance of as ventral or tailcone exists, the exter- 30 inches. nal means of opening, including in- (2) Each Type A, Type B, Type C or structions if applicable, must be con- Type I passenger emergency exit oper- spicuously marked in red, or bright ating handle must— chrome yellow if the background color (i) Be self-illuminated with an initial is such that red is inconspicuous. When brightness of at least 160 micro- the opening means is located on only lamberts; or one side of the fuselage, a conspicuous (ii) Be conspicuously located and well marking to that effect must be pro- illuminated by the emergency lighting vided on the other side. even in conditions of occupant crowd- (g) Each sign required by paragraph ing at the exit. (d) of this section may use the word (3) [Reserved] ‘‘exit’’ in its legend in place of the (4) Each Type A, Type B, Type C, term ‘‘emergency exit’’. Type I, or Type II passenger emergency [Amdt. 25–15, 32 FR 13264, Sept. 20, 1967, as exit with a locking mechanism re- amended by Amdt. 25–32, 37 FR 3970, Feb. 24, leased by rotary motion of the handle 1972; Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; 43 must be marked— FR 52495, Nov. 13, 1978; Amdt. 25–79, 58 FR (i) With a red arrow, with a shaft at 45229, Aug. 26, 1993; Amdt. 25–88, 61 FR 57958, least three-fourths of an inch wide and Nov. 8, 1996] a head twice the width of the shaft, extending along at least 70 degrees of arc § 25.812 Emergency lighting. at a radius approximately equal to (a) An emergency lighting system, three-fourths of the handle length. independent of the main lighting sys- (ii) So that the centerline of the exit tem, must be installed. However, the handle is within ±1 inch of the pro- sources of general cabin illumination jected point of the arrow when the han- may be common to both the emergency dle has reached full travel and has re- and the main lighting systems if the leased the locking mechanism, and power supply to the emergency light- (iii) With the word ‘‘open’’ in red let- ing system is independent of the power ters 1 inch high, placed horizontally supply to the main lighting system. near the head of the arrow. The emergency lighting system must (f) Each emergency exit that is re- include: quired to be openable from the outside, (1) Illuminated emergency exit mark- and its means of opening, must be ing and locating signs, sources of gen- marked on the outside of the airplane. eral cabin illumination, interior light- In addition, the following apply: ing in emergency exit areas, and floor (1) The outside marking for each pas- proximity escape path marking. senger emergency exit in the side of (2) Exterior emergency lighting. the fuselage must include a 2-inch col- (b) Emergency exit signs— ored band outlining the exit. (1) For airplanes that have a pas- (2) Each outside marking including senger seating configuration, excluding the band, must have color contrast to pilot seats, of 10 seats or more must be readily distinguishable from the sur- meet the following requirements: rounding fuselage surface. The contrast (i) Each passenger emergency exit lo- must be such that if the reflectance of cator sign required by § 25.811(d)(1) and

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each passenger emergency exit mark- (d) The floor of the passageway leading sign required by § 25.811(d)(2) must ing to each floor-level passenger emer- have red letters at least 11⁄2 inches high gency exit, between the main aisles on an illuminated white background, and the exit openings, must be pro- and must have an area of at least 21 vided with illumination that is not less square inches excluding the letters. than 0.02 foot-candle measured along a The lighted background-to-letter con- line that is within 6 inches of and par- trast must be at least 10 : 1. The letter allel to the floor and is centered on the height to stroke-width ratio may not passenger evacuation path. be more than 7 : 1 nor less than 6 : 1. (e) Floor proximity emergency es- These signs must be internally elec- cape path marking must provide emer- trically illuminated with a background gency evacuation guidance for pas- brightness of at least 25 foot-lamberts sengers when all sources of illumina- and a high-to-low background contrast tion more than 4 feet above the cabin no greater than 3 : 1. aisle floor are totally obscured. In the (ii) Each passenger emergency exit dark of the night, the floor proximity sign required by § 25.811(d)(3) must have emergency escape path marking must red letters at least 11⁄2 inches high on a enable each passenger to— white background having an area of at (1) After leaving the passenger seat, least 21 square inches excluding the visually identify the emergency escape letters. These signs must be internally path along the cabin aisle floor to the electrically illuminated or self-illumi- first exits or pair of exits forward and nated by other than electrical means aft of the seat; and and must have an initial brightness of (2) Readily identify each exit from at least 400 microlamberts. The colors the emergency escape path by ref- may be reversed in the case of a sign erence only to markings and visual fea- that is self-illuminated by other than tures not more than 4 feet above the electrical means. cabin floor. (2) For airplanes that have a passenger seating configuration, excluding (f) Except for subsystems provided in pilot seats, of nine seats or less, that accordance with paragraph (h) of this are required by § 25.811(d)(1), (2), and (3) section that serve no more than one as- must have red letters at least 1 inch sist means, are independent of the air- high on a white background at least 2 plane’s main emergency lighting sys- inches high. These signs may be inter- tem, and are automatically activated nally electrically illuminated, or self- when the assist means is erected, the illuminated by other than electrical emergency lighting system must be de- means, with an initial brightness of at signed as follows. least 160 microlamberts. The colors (1) The lights must be operable may be reversed in the case of a sign manually from the flight crew station that is self-illuminated by other than and from a point in the passenger com- electrical means. partment that is readily accessible to a (c) General illumination in the pas- normal flight attendant seat. senger cabin must be provided so that (2) There must be a flight crew warn- when measured along the centerline of ing light which illuminates when power main passenger aisle(s), and cross is on in the airplane and the emergency aisle(s) between main aisles, at seat lighting control device is not armed. arm-rest height and at 40-inch inter- (3) The cockpit control device must vals, the average illumination is not have an ‘‘on,’’ ‘‘off,’’ and ‘‘armed’’ posi- less than 0.05 foot-candle and the illu- tion so that when armed in the cockpit mination at each 40-inch interval is not or turned on at either the cockpit or less than 0.01 foot-candle. A main pas- flight attendant station the lights will senger aisle(s) is considered to extend either light or remain lighted upon along the fuselage from the most for- interruption (except an interruption ward passenger emergency exit or caused by a transverse vertical separa- cabin occupant seat, whichever is far- tion of the fuselage during crash land- ther forward, to the most rearward pas- ing) of the airplane’s normal electric senger emergency exit or cabin occu- power. There must be a means to safe- pant seat, whichever is farther aft. guard against inadvertent operation of

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the control device from the ‘‘armed’’ or gency lighting system, and is auto- ‘‘on’’ positions. matically activated when the assist (g) Exterior emergency lighting must means is erected, the lighting provi- be provided as follows: sions— (1) At each overwing emergency exit (i) May not be adversely affected by the illumination must be— stowage; and (i) Not less than 0.03 foot-candle (ii) Must provide illumination of not (measured normal to the direction of less than 0.03 foot-candle (measured the incident light) on a 2-square-foot normal to the direction of incident area where an evacuee is likely to light) at the ground and of the erected make his first step outside the cabin; assist means where an evacuee would (ii) Not less than 0.05 foot-candle (measured normal to the direction of normally make first contact with the the incident light) for a minimum ground, with the airplane in each of width of 42 inches for a Type A the attitudes corresponding to the col- overwing emergency exit and two feet lapse of one or more legs of the landing for all other overwing emergency exits gear. along the 30 percent of the slip-resist- (i) The energy supply to each emer- ant portion of the escape route re- gency lighting unit must provide the quired in § 25.810(c) that is farthest required level of illumination for at from the exit; and least 10 minutes at the critical ambient (iii) Not less than 0.03 foot-candle on conditions after emergency landing. the ground surface with the landing (j) If storage batteries are used as the gear extended (measured normal to the energy supply for the emergency light- direction of the incident light) where ing system, they may be recharged an evacuee using the established escape from the airplane’s main electric power route would normally make first con- system: Provided, That, the charging tact with the ground. circuit is designed to preclude inad- (2) At each non-overwing emergency vertent battery discharge into charg- exit not required by § 25.810(a) to have ing circuit faults. descent assist means the illumination (k) Components of the emergency must be not less than 0.03 foot-candle lighting system, including batteries, (measured normal to the direction of the incident light) on the ground sur- wiring relays, lamps, and switches face with the landing gear extended must be capable of normal operation where an evacuee is likely to make after having been subjected to the iner- first contact with the ground outside tia forces listed in § 25.561(b). the cabin. (l) The emergency lighting system (h) The means required in must be designed so that after any sin- §§ 25.810(a)(1) and (d) to assist the occu- gle transverse vertical separation of pants in descending to the ground must the fuselage during crash landing— be illuminated so that the erected as- (1) Not more than 25 percent of all sist means is visible from the airplane. electrically illuminated emergency (1) If the assist means is illuminated lights required by this section are ren- by exterior emergency lighting, it dered inoperative, in addition to the must provide illumination of not less lights that are directly damaged by the than 0.03 foot-candle (measured normal separation; to the direction of the incident light) (2) Each electrically illuminated exit at the ground end of the erected assist sign required under § 25.811(d)(2) re- means where an evacuee using the es- mains operative exclusive of those that tablished escape route would normally are directly damaged by the separa- make first contact with the ground, tion; and with the airplane in each of the attitudes corresponding to the collapse of (3) At least one required exterior one or more legs of the landing gear. emergency light for each side of the (2) If the emergency lighting sub- airplane remains operative exclusive of system illuminating the assist means serves no other assist means, is independent of the airplane’s main emer-

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those that are directly damaged by the est main aisle and a Type A or B exit; separation. and (2) A cross-aisle which leads to the [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967, as amended by Amdt. 25–28, 36 FR 16899, Aug. 26, immediate vicinity of each passageway 1971; Amdt. 25–32, 37 FR 3971, Feb. 24, 1972; between the nearest main aisle and a Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. Type 1, Type II, or Type III exit; except 25–58, 49 FR 43186, Oct. 26, 1984; Amdt. 25–88, that when two Type III exits are lo- 61 FR 57958, Nov. 8, 1996; Amdt. 25–116, 69 FR cated within three passenger rows of 62788, Oct. 27, 2004; Amdt. 25–128, 74 FR 25645, each other, a single cross-aisle may be May 29, 2009] used if it leads to the vicinity between the passageways from the nearest main § 25.813 Emergency exit access. aisle to each exit. Each required emergency exit must (b) Adequate space to allow crew- be accessible to the passengers and lo- member(s) to assist in the evacuation cated where it will afford an effective of passengers must be provided as fol- means of evacuation. Emergency exit lows: distribution must be as uniform as (1) Each assist space must be a rec- practical, taking passenger distribu- tangle on the floor, of sufficient size to tion into account; however, the size enable a crewmember, standing erect, and location of exits on both sides of to effectively assist evacuees. The as- the cabin need not be symmetrical. If sist space must not reduce the unob- only one floor level exit per side is pre- structed width of the passageway below scribed, and the airplane does not have that required for the exit. a tailcone or ventral emergency exit, (2) For each Type A or B exit, assist the floor level exit must be in the rear- space must be provided at each side of ward part of the passenger compart- the exit regardless of whether an assist ment, unless another location affords a means is required by § 25.810(a). more effective means of passenger (3) For each Type C, I or II exit in- evacuation. Where more than one floor stalled in an airplane with seating for level exit per side is prescribed, at more than 80 passengers, an assist least one floor level exit per side must space must be provided at one side of be located near each end of the cabin, the passageway regardless of whether except that this provision does not an assist means is required by apply to combination cargo/passenger § 25.810(a). configurations. In addition— (4) For each Type C, I or II exit, an (a) There must be a passageway lead- assist space must be provided at one ing from the nearest main aisle to each side of the passageway if an assist Type A, Type B, Type C, Type I, or means is required by § 25.810(a). Type II emergency exit and between in- (5) For any tailcone exit that quali- dividual passenger areas. Each passage- fies for 25 additional passenger seats way leading to a Type A or Type B exit under the provisions of § 25.807(g)(9)(ii), must be unobstructed and at least 36 an assist space must be provided, if an inches wide. Passageways between indi- assist means is required by § 25.810(a). vidual passenger areas and those lead- (6) There must be a handle, or han- ing to Type I, Type II, or Type C emer- dles, at each assist space, located to gency exits must be unobstructed and enable the crewmember to steady him- at least 20 inches wide. Unless there self or herself: are two or more main aisles, each Type (i) While manually activating the as- A or B exit must be located so that sist means (where applicable) and, there is passenger flow along the main (ii) While assisting passengers during aisle to that exit from both the forward an evacuation. and aft directions. If two or more main (c) The following must be provided aisles are provided, there must be un- for each Type III or Type IV exit—(1) obstructed cross-aisles at least 20 There must be access from the nearest inches wide between main aisles. There aisle to each exit. In addition, for each must be— Type III exit in an airplane that has a (1) A cross-aisle which leads directly passenger seating configuration of 60 or to each passageway between the near- more—

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(i) Except as provided in paragraph plane in which it is installed, there (c)(1)(ii), the access must be provided must be placards that— by an unobstructed passageway that is (i) Are readable by all persons seated at least 10 inches in width for interior adjacent to and facing a passageway to arrangements in which the adjacent the exit; seat rows on the exit side of the aisle (ii) Accurately state or illustrate the contain no more than two seats, or 20 proper method of opening the exit, in- inches in width for interior arrange- cluding the use of handholds; and ments in which those rows contain (iii) If the exit is a removable hatch, three seats. The width of the passage- state the weight of the hatch and indi- way must be measured with adjacent cate an appropriate location to place seats adjusted to their most adverse the hatch after removal. position. The centerline of the required (d) If it is necessary to pass through passageway width must not be dis- a passageway between passenger com- placed more than 5 inches horizontally partments to reach any required emer- from that of the exit. gency exit from any seat in the pas- (ii) In lieu of one 10- or 20-inch passenger cabin, the passageway must be sageway, there may be two passage- unobstructed. However, curtains may ways, between seat rows only, that be used if they allow free entry must be at least 6 inches in width and through the passageway. lead to an unobstructed space adjacent (e) No door may be installed between to each exit. (Adjacent exits must not any passenger seat that is occupiable share a common passageway.) The for takeoff and landing and any pas- width of the passageways must be senger emergency exit, such that the measured with adjacent seats adjusted door crosses any egress path (including to their most adverse position. The un- aisles, crossaisles and passageways). obstructed space adjacent to the exit (f) If it is necessary to pass through a must extend vertically from the floor doorway separating any crewmember to the ceiling (or bottom of sidewall seat (except those seats on the stowage bins), inboard from the exit for flightdeck), occupiable for takeoff and a distance not less than the width of landing, from any emergency exit, the the narrowest passenger seat installed door must have a means to latch it in on the airplane, and from the forward the open position. The latching means edge of the forward passageway to the must be able to withstand the loads aft edge of the aft passageway. The exit imposed upon it when the door is sub- opening must be totally within the fore jected to the ultimate inertia forces, and aft bounds of the unobstructed relative to the surrounding structure, listed in § 25.561(b). space. (2) In addition to the access— [Amdt. 25–1, 30 FR 3204, Mar. 9, 1965, as (i) For airplanes that have a pas- amended by Amdt. 25–15, 32 FR 13265, Sept. senger seating configuration of 20 or 20, 1967; Amdt. 25–32, 37 FR 3971, Feb. 24, 1972; Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. more, the projected opening of the exit 25–72, 55 FR 29783, July 20, 1990; Amdt. 25–76, provided must not be obstructed and 57 FR 19244, May 4, 1992; Amdt. 25–76, 57 FR there must be no interference in open- 29120, June 30, 1992; Amdt. 25–88, 61 FR 57958, ing the exit by seats, berths, or other Nov. 8, 1996; Amdt. 25–116, 69 FR 62788, Oct. protrusions (including any seatback in 27, 2004; Amdt. 25–128, 74 FR 25645, May 29, the most adverse position) for a dis- 2009] tance from that exit not less than the § 25.815 Width of aisle. width of the narrowest passenger seat installed on the airplane. The passenger aisle width at any (ii) For airplanes that have a pas- point between seats must equal or ex- senger seating configuration of 19 or ceed the values in the following table: fewer, there may be minor obstructions Minimum passenger in this region, if there are compen- aisle width (inches) sating factors to maintain the effec- Passenger seating capacity Less than 25 in. and tiveness of the exit. 25 in. from more from (3) For each Type III exit, regardless floor floor of the passenger capacity of the air- 10 or less ...... 1 12 15

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Minimum passenger (c) There must be an aural emer- aisle width (inches) gency alarm system, audible during Passenger seating capacity Less than 25 in. and normal and emergency conditions, to 25 in. from more from floor floor enable crewmembers on the flight deck and at each required floor level emer- 11 through 19 ...... 12 20 gency exit to alert occupants of each 20 or more ...... 15 20 lower deck service compartment of an 1 A narrower width not less than 9 inches may be approved emergency situation. when substantiated by tests found necessary by the Administrator. (d) There must be a means, readily detectable by occupants of each lower [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967, as deck service compartment, that indi- amended by Amdt. 25–38, 41 FR 55466, Dec. 20, cates when seat belts should be fas- 1976] tened. § 25.817 Maximum number of seats (e) If a public address system is in- abreast. stalled in the airplane, speakers must On airplanes having only one pas- be provided in each lower deck service senger aisle, no more than three seats compartment. abreast may be placed on each side of (f) For each occupant permitted in a the aisle in any one row. lower deck service compartment, there must be a forward or aft facing seat [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967] which meets the requirements of § 25.819 Lower deck service compart- § 25.785(d), and must be able to with- ments (including galleys). stand maximum flight loads when occupied. For airplanes with a service compartment located below the main deck, (g) For each powered lift system in- which may be occupied during taxi or stalled between a lower deck service flight but not during takeoff or land- compartment and the main deck for ing, the following apply: the carriage of persons or equipment, (a) There must be at least two emer- or both, the system must meet the fol- gency evacuation routes, one at each lowing requirements: end of each lower deck service com- (1) Each lift control switch outside partment or two having sufficient sepa- the lift, except emergency stop but- ration within each compartment, tons, must be designed to prevent the which could be used by each occupant activation of the life if the lift door, or of the lower deck service compartment the hatch required by paragraph (g)(3) to rapidly evacuate to the main deck of this section, or both are open. under normal and emergency lighting (2) An emergency stop button, that conditions. The routes must provide for when activated will immediately stop the evacuation of incapacitated per- the lift, must be installed within the sons, with assistance. The use of the lift and at each entrance to the lift. evacuation routes may not be depend- (3) There must be a hatch capable of ent on any powered device. The routes being used for evacuating persons from must be designed to minimize the pos- the lift that is openable from inside sibility of blockage which might result and outside the lift without tools, with from fire, mechanical or structural the lift in any position. failure, or persons standing on top of or [Amdt. 25–53, 45 FR 41593, June 19, 1980; 45 FR against the escape routes. In the event 43154, June 26, 1980; Amdt. 25–110; 68 FR 36883, the airplane’s main power system or June 19, 2003] compartment main lighting system should fail, emergency illumination for § 25.820 Lavatory doors. each lower deck service compartment All lavatory doors must be designed must be automatically provided. to preclude anyone from becoming (b) There must be a means for two- trapped inside the lavatory. If a lock- way voice communication between the ing mechanism is installed, it must be flight deck and each lower deck service capable of being unlocked from the compartment, which remains available outside without the aid of special tools. following loss of normal electrical power generating system. [Doc. No. 2003–14193, 69 FR 24502, May 3, 2004]

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VENTILATION AND HEATING without depressurizing beyond safe limits. § 25.831 Ventilation. (e) Except as provided in paragraph (a) Under normal operating condi- (f) of this section, means must be pro- tions and in the event of any probable vided to enable the occupants of the failure conditions of any system which following compartments and areas to would adversely affect the ventilating control the temperature and quantity air, the ventilation system must be de- of ventilating air supplied to their signed to provide a sufficient amount compartment or area independently of of uncontaminated air to enable the the temperature and quantity of air crewmembers to perform their duties supplied to other compartments and without undue discomfort or fatigue areas: and to provide reasonable passenger (1) The flight crew compartment. comfort. For normal operating condi- (2) Crewmember compartments and tions, the ventilation system must be areas other than the flight crew com- designed to provide each occupant with partment unless the crewmember com- an airflow containing at least 0.55 partment or area is ventilated by air pounds of fresh air per minute. interchange with other compartments (b) Crew and passenger compartment or areas under all operating conditions. air must be free from harmful or haz- (f) Means to enable the flight crew to ardous concentrations of gases or va- control the temperature and quantity pors. In meeting this requirement, the of ventilating air supplied to the flight following apply: crew compartment independently of (1) Carbon monoxide concentrations the temperature and quantity of ven- in excess of 1 part in 20,000 parts of air tilating air supplied to other compart- are considered hazardous. For test pur- ments are not required if all of the fol- poses, any acceptable carbon monoxide lowing conditions are met: detection method may be used. (2) Carbon dioxide concentration dur- (1) The total volume of the flight ing flight must be shown not to exceed crew and passenger compartments is 0.5 percent by volume (sea level equiva- 800 cubic feet or less. lent) in compartments normally occu- (2) The air inlets and passages for air pied by passengers or crewmembers. to flow between flight crew and pas- (c) There must be provisions made to senger compartments are arranged to ensure that the conditions prescribed provide compartment temperatures in paragraph (b) of this section are met within 5 degrees F. of each other and after reasonably probable failures or adequate ventilation to occupants in malfunctioning of the ventilating, both compartments. heating, pressurization, or other sys- (3) The temperature and ventilation tems and equipment. controls are accessible to the flight (d) If accumulation of hazardous crew. quantities of smoke in the cockpit area (g) The exposure time at any given is reasonably probable, smoke evacu- temperature must not exceed the val- ation must be readily accomplished, ues shown in the following graph after starting with full pressurization and any improbable failure condition.

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[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. 25–87, 61 FR 28695, June 5, 1996; Amdt. 25–89, 61 FR 63956, Dec. 2, 1996]

§ 25.832 Cabin ozone concentration. § 25.833 Combustion heating systems. (a) The airplane cabin ozone con- Combustion heaters must be ap- centration during flight must be shown proved. not to exceed— [Amdt. 25–72, 55 FR 29783, July 20, 1990] (1) 0.25 parts per million by volume, sea level equivalent, at any time above PRESSURIZATION flight level 320; and (2) 0.1 parts per million by volume, § 25.841 Pressurized cabins. sea level equivalent, time-weighted average during any 3-hour interval above (a) Pressurized cabins and compart- flight level 270. ments to be occupied must be equipped (b) For the purpose of this section, to provide a cabin pressure altitude of ‘‘sea level equivalent’’ refers to condi- not more than 8,000 feet at the max- tions of 25 °C and 760 millimeters of imum operating altitude of the air- mercury pressure. plane under normal operating condi- (c) Compliance with this section tions. must be shown by analysis or tests (1) If certification for operation based on airplane operational proce- above 25,000 feet is requested, the air- dures and performance limitations, plane must be designed so that occu- that demonstrate that either— pants will not be exposed to cabin pres- (1) The airplane cannot be operated sure altitudes in excess of 15,000 feet at an altitude which would result in after any probable failure condition in cabin ozone concentrations exceeding the pressurization system. the limits prescribed by paragraph (a) (2) The airplane must be designed so of this section; or that occupants will not be exposed to a (2) The airplane ventilation system, cabin pressure altitude that exceeds including any ozone control equipment, the following after decompression from will maintain cabin ozone concentra- any failure condition not shown to be tions at or below the limits prescribed extremely improbable: by paragraph (a) of this section. (i) Twenty-five thousand (25,000) feet [Amdt. 25–50, 45 FR 3883, Jan. 1, 1980, as for more than 2 minutes; or amended by Amdt. 25–56, 47 FR 58489, Dec. 30, (ii) Forty thousand (40,000) feet for 1982; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] any duration.

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(3) Fuselage structure, engine and (8) The pressure sensors necessary to system failures are to be considered in meet the requirements of paragraphs evaluating the cabin decompression. (b)(5) and (b)(6) of this section and (b) Pressurized cabins must have at § 25.1447(c), must be located and the least the following valves, controls, sensing system designed so that, in the and indicators for controlling cabin event of loss of cabin pressure in any pressure: passenger or crew compartment (in- (1) Two pressure relief valves to auto- cluding upper and lower lobe galleys), matically limit the positive pressure the warning and automatic presen- differential to a predetermined value tation devices, required by those provi- at the maximum rate of flow delivered sions, will be actuated without any by the pressure source. The combined delay that would significantly increase capacity of the relief valves must be the hazards resulting from decompres- large enough so that the failure of any sion. one valve would not cause an appre- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ciable rise in the pressure differential. amended by Amdt. 25–38, 41 FR 55466, Dec. 20, The pressure differential is positive 1976; Amdt. 25–87, 61 FR 28696, June 5, 1996] when the internal pressure is greater than the external. § 25.843 Tests for pressurized cabins. (2) Two reverse pressure differential (a) Strength test. The complete pres- relief valves (or their equivalents) to surized cabin, including doors, win- automatically prevent a negative pres- dows, and valves, must be tested as a sure differential that would damage pressure vessel for the pressure dif- the structure. One valve is enough, ferential specified in § 25.365(d). however, if it is of a design that reasonably precludes its malfunctioning. (b) Functional tests. The following functional tests must be performed: (3) A means by which the pressure differential can be rapidly equalized. (1) Tests of the functioning and capacity of the positive and negative (4) An automatic or manual regulator pressure differential valves, and of the for controlling the intake or exhaust airflow, or both, for maintaining the emergency release valve, to stimulate required internal pressures and airflow the effects of closed regulator valves. rates. (2) Tests of the pressurization system (5) Instruments at the pilot or flight to show proper functioning under each engineer station to show the pressure possible condition of pressure, tem- differential, the cabin pressure alti- perature, and moisture, up to the max- tude, and the rate of change of the imum altitude for which certification cabin pressure altitude. is requested. (6) Warning indication at the pilot or (3) Flight tests, to show the perform- flight engineer station to indicate ance of the pressure supply, pressure when the safe or preset pressure dif- and flow regulators, indicators, and ferential and cabin pressure altitude warning signals, in steady and stepped limits are exceeded. Appropriate warn- climbs and descents at rates cor- ing markings on the cabin pressure dif- responding to the maximum attainable ferential indicator meet the warning within the operating limitations of the requirement for pressure differential airplane, up to the maximum altitude limits and an aural or visual signal (in for which certification is requested. addition to cabin altitude indicating (4) Tests of each door and emergency means) meets the warning requirement exit, to show that they operate prop- for cabin pressure altitude limits if it erly after being subjected to the flight warns the flight crew when the cabin tests prescribed in paragraph (b)(3) of pressure altitude exceeds 10,000 feet. this section. (7) A warning placard at the pilot or flight engineer station if the structure FIRE PROTECTION is not designed for pressure differentials up to the maximum relief valve § 25.851 Fire extinguishers. setting in combination with landing (a) Hand fire extinguishers. (1) The fol- loads. lowing minimum number of hand fire

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extinguishers must be conveniently lo- (i) No extinguishing agent likely to cated and evenly distributed in pas- enter personnel compartments will be senger compartments: hazardous to the occupants; and (ii) No discharge of the extinguisher Passenger capacity No. of extinguishers can cause structural damage. 7 through 30 ...... 1 (2) The capacity of each required 31 through 60 ...... 2 built-in fire extinguishing system must 61 through 200 ...... 3 be adequate for any fire likely to occur 201 through 300 ...... 4 in the compartment where used, con- 301 through 400 ...... 5 401 through 500 ...... 6 sidering the volume of the compart- 501 through 600 ...... 7 ment and the ventilation rate. 601 through 700 ...... 8 [Amdt. 25–74, 56 FR 15456, Apr. 16, 1991]

(2) At least one hand fire extin- § 25.853 Compartment interiors. guisher must be conveniently located in the pilot compartment. For each compartment occupied by (3) At least one readily accessible the crew or passengers, the following apply: hand fire extinguisher must be avail- (a) Materials (including finishes or able for use in each Class A or Class B decorative surfaces applied to the ma- cargo or baggage compartment and in terials) must meet the applicable test each Class E cargo or baggage compart- criteria prescribed in part I of appendix ment that is accessible to crew- F of this part, or other approved equiv- members in flight. alent methods, regardless of the pas- (4) At least one hand fire extin- senger capacity of the airplane. guisher must be located in, or readily (b) [Reserved] accessible for use in, each galley lo- (c) In addition to meeting the re- cated above or below the passenger quirements of paragraph (a) of this sec- compartment. tion, seat cushions, except those on (5) Each hand fire extinguisher must flight crewmember seats, must meet be approved. the test requirements of part II of ap- (6) At least one of the required fire pendix F of this part, or other equiva- extinguishers located in the passenger lent methods, regardless of the pas- compartment of an airplane with a passenger capacity of the airplane. senger capacity of at least 31 and not (d) Except as provided in paragraph more than 60, and at least two of the (e) of this section, the following inte- fire extinguishers located in the pas- rior components of airplanes with passenger compartment of an airplane senger capacities of 20 or more must with a passenger capacity of 61 or more also meet the test requirements of must contain Halon 1211 parts IV and V of appendix F of this (bromochlorodifluoromethane CBrC1 part, or other approved equivalent F2), or equivalent, as the extinguishing method, in addition to the flamma- agent. The type of extinguishing agent bility requirements prescribed in para- used in any other extinguisher required graph (a) of this section: by this section must be appropriate for (1) Interior ceiling and wall panels, the kinds of fires likely to occur where other than lighting lenses and win- used. dows; (7) The quantity of extinguishing (2) Partitions, other than transparent agent used in each extinguisher re- panels needed to enhance cabin safety; quired by this section must be appro- (3) Galley structure, including ex- priate for the kinds of fires likely to posed surfaces of stowed carts and occur where used. standard containers and the cavity (8) Each extinguisher intended for walls that are exposed when a full com- use in a personnel compartment must plement of such carts or containers is be designed to minimize the hazard of not carried; and toxic gas concentration. (4) Large cabinets and cabin stowage (b) Built-in fire extinguishers. If a compartments, other than underseat built-in fire extinguisher is provided— stowage compartments for stowing (1) Each built-in fire extinguishing small items such as magazines and system must be installed so that— maps.

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(e) The interiors of compartments, each disposal receptacle upon occur- such as pilot compartments, galleys, rence of a fire in that receptacle. lavatories, crew rest quarters, cabinets [Amdt. 25–74, 56 FR 15456, Apr. 16, 1991] and stowage compartments, need not meet the standards of paragraph (d) of § 25.855 Cargo or baggage compart- this section, provided the interiors of ments. such compartments are isolated from For each cargo or baggage compart- the main passenger cabin by doors or ment, the following apply: equivalent means that would normally (a) The compartment must meet one be closed during an emergency landing of the class requirements of § 25.857. condition. (b) Class B through Class E cargo or (f) Smoking is not allowed in lava- baggage compartments, as defined in tories. If smoking is allowed in any § 25.857, must have a liner, and the liner area occupied by the crew or pas- must be separate from (but may be at- sengers, an adequate number of self- tached to) the airplane structure. contained, removable ashtrays must be (c) Ceiling and sidewall liner panels provided in designated smoking sec- of Class C compartments must meet tions for all seated occupants. the test requirements of part III of ap- (g) Regardless of whether smoking is pendix F of this part or other approved allowed in any other part of the air- equivalent methods. plane, lavatories must have self-con- (d) All other materials used in the tained, removable ashtrays located construction of the cargo or baggage conspicuously on or near the entry side compartment must meet the applicable of each lavatory door, except that one test criteria prescribed in part I of ap- ashtray may serve more than one lava- pendix F of this part or other approved tory door if the ashtray can be seen equivalent methods. readily from the cabin side of each lav- (e) No compartment may contain any atory served. controls, lines, equipment, or acces- (h) Each receptacle used for the dis- sories whose damage or failure would posal of flammable waste material affect safe operation, unless those must be fully enclosed, constructed of items are protected so that— at least fire resistant materials, and (1) They cannot be damaged by the must contain fires likely to occur in it movement of cargo in the compart- under normal use. The capability of the ment, and receptacle to contain those fires under (2) Their breakage or failure will not all probable conditions of wear, mis- create a fire hazard. alignment, and ventilation expected in (f) There must be means to prevent service must be demonstrated by test. cargo or baggage from interfering with the functioning of the fire protective [Amdt. 25–83, 60 FR 6623, Feb. 2, 1995, as features of the compartment. amended by Amdt. 25–116, 69 FR 62788, Oct. (g) Sources of heat within the com- 27, 2004] partment must be shielded and insu- § 25.854 Lavatory fire protection. lated to prevent igniting the cargo or baggage. For airplanes with a passenger capac- (h) Flight tests must be conducted to ity of 20 or more: show compliance with the provisions of (a) Each lavatory must be equipped § 25.857 concerning— with a smoke detector system or equiv- (1) Compartment accessibility, alent that provides a warning light in (2) The entries of hazardous quan- the cockpit, or provides a warning tities of smoke or extinguishing agent light or audible warning in the pas- into compartments occupied by the senger cabin that would be readily de- crew or passengers, and tected by a flight attendant; and (3) The dissipation of the extin- (b) Each lavatory must be equipped guishing agent in Class C compart- with a built-in fire extinguisher for ments. each disposal receptacle for towels, (i) During the above tests, it must be paper, or waste, located within the lav- shown that no inadvertent operation of atory. The extinguisher must be de- smoke or fire detectors in any com- signed to discharge automatically into partment would occur as a result of

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fire contained in any other compart- smoke, flames, or extinguishing agent, ment, either during or after extin- will enter any compartment occupied guishment, unless the extinguishing by the crew or passengers; system floods each such compartment (3) There is a separate approved simultaneously. smoke detector or fire detector system (j) Cargo or baggage compartment to give warning at the pilot or flight electrical wiring interconnection sys- engineer station. tem components must meet the re- (c) Class C. A Class C cargo or bag- quirements of § 25.1721. gage compartment is one not meeting [Amdt. 25–72, 55 FR 29784, July 20, 1990, as the requirements for either a Class A amended by Amdt. 25–93, 63 FR 8048, Feb. 17, or B compartment but in which— 1998; Amdt. 25–116, 69 FR 62788, Oct. 27, 2004; (1) There is a separate approved Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] smoke detector or fire detector system to give warning at the pilot or flight § 25.856 Thermal/Acoustic insulation engineer station; materials. (2) There is an approved built-in fire (a) Thermal/acoustic insulation ma- extinguishing or suppression system terial installed in the fuselage must controllable from the cockpit. meet the flame propagation test re- (3) There are means to exclude haz- quirements of part VI of Appendix F to ardous quantities of smoke, flames, or this part, or other approved equivalent extinguishing agent, from any com- test requirements. This requirement partment occupied by the crew or pas- does not apply to ‘‘small parts,’’ as de- sengers; fined in part I of Appendix F of this (4) There are means to control ven- part. tilation and drafts within the compart- (b) For airplanes with a passenger ca- ment so that the extinguishing agent pacity of 20 or greater, thermal/acous- used can control any fire that may tic insulation materials (including the start within the compartment. means of fastening the materials to the (d) [Reserved] fuselage) installed in the lower half of (e) Class E. A Class E cargo compart- the airplane fuselage must meet the ment is one on airplanes used only for flame penetration resistance test re- the carriage of cargo and in which— quirements of part VII of Appendix F (1) [Reserved] to this part, or other approved equiva- (2) There is a separate approved lent test requirements. This require- smoke or fire detector system to give ment does not apply to thermal/acous- warning at the pilot or flight engineer tic insulation installations that the station; FAA finds would not contribute to fire (3) There are means to shut off the penetration resistance. ventilating airflow to, or within, the [Amdt. 25–111, 68 FR 45059, July 31, 2003] compartment, and the controls for these means are accessible to the flight § 25.857 Cargo compartment classifica- crew in the crew compartment; tion. (4) There are means to exclude haz- (a) Class A; A Class A cargo or bag- ardous quantities of smoke, flames, or gage compartment is one in which— noxious gases, from the flight crew (1) The presence of a fire would be compartment; and easily discovered by a crewmember (5) The required crew emergency while at his station; and exits are accessible under any cargo (2) Each part of the compartment is loading condition. easily accessible in flight. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (b) Class B. A Class B cargo or bag- amended by Amdt. 25–32, 37 FR 3972, Feb. 24, gage compartment is one in which— 1972; Amdt. 25–60, 51 FR 18243, May 16, 1986; (1) There is sufficient access in flight Amdt. 25–93, 63 FR 8048, Feb. 17, 1998] to enable a crewmember to effectively reach any part of the compartment § 25.858 Cargo or baggage compart- with the contents of a hand fire extin- ment smoke or fire detection sys- guisher; tems. (2) When the access provisions are If certification with cargo or baggage being used, no hazardous quantity of compartment smoke or fire detection

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provisions is requested, the following that any fire originating in the heater must be met for each cargo or baggage can be contained in the duct; and compartment with those provisions: (2) Each part of any ventilating duct (a) The detection system must pro- passing through any region having a vide a visual indication to the flight flammable fluid system must be con- crew within one minute after the start structed or isolated from that system of a fire. so that the malfunctioning of any com- (b) The system must be capable of de- ponent of that system cannot intro- tecting a fire at a temperature signifi- duce flammable fluids or vapors into cantly below that at which the struc- the ventilating airstream. tural integrity of the airplane is sub- (c) Combustion air ducts. Each com- stantially decreased. bustion air duct must be fireproof for a (c) There must be means to allow the distance great enough to prevent dam- crew to check in flight, the functioning age from backfiring or reverse flame of each fire detector circuit. propagation. In addition— (d) The effectiveness of the detection (1) No combustion air duct may have system must be shown for all approved a common opening with the ventilating operating configurations and condi- airstream unless flames from backfires tions. or reverse burning cannot enter the ventilating airstream under any oper- [Amdt. 25–54, 45 FR 60173, Sept. 11, 1980, as ating condition, including reverse flow amended by Amdt. 25–93, 63 FR 8048, Feb. 17, or malfunctioning of the heater or its 1998] associated components; and § 25.859 Combustion heater fire pro- (2) No combustion air duct may re- tection. strict the prompt relief of any backfire that, if so restricted, could cause heat- (a) Combustion heater fire zones. The er failure. following combustion heater fire zones (d) Heater controls; general. Provision must be protected from fire in accord- must be made to prevent the hazardous ance with the applicable provisions of accumulation of water or ice on or in §§ 25.1181 through 25.1191 and §§ 25.1195 any heater control component, control through 25.1203; system tubing, or safety control. (1) The region surrounding the heat- (e) Heater safety controls. For each er, if this region contains any flam- combustion heater there must be the mable fluid system components (ex- following safety control means: 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 to that (2) The region surrounding the heat- heater at a point remote from that er, if the heater fuel system has fit- heater when any of the following oc- tings that, if they leaked, would allow curs: fuel or vapors to enter this region. (i) The heat exchanger temperature (3) The part of the ventilating air exceeds safe limits. passage that surrounds the combustion (ii) The ventilating air temperature chamber. However, no fire extinguish- exceeds safe limits. ment is required in cabin ventilating (iii) The combustion airflow becomes air passages. inadequate for safe operation. (b) Ventilating air ducts. Each ven- (iv) The ventilating airflow becomes tilating air duct passing through any inadequate for safe operation. fire zone must be fireproof. In addi- (2) The means of complying with tion— paragraph (e)(1) of this section for any (1) Unless isolation is provided by individual heater must— fireproof valves or by equally effective (i) Be independent of components means, the ventilating air duct down- serving any other heater whose heat stream of each heater must be fireproof output is essential for safe operation; for a distance great enough to ensure and

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(ii) Keep the heater off until re- § 25.863 Flammable fluid fire protec- started by the crew. tion. (3) There must be means to warn the (a) In each area where flammable crew when any heater whose heat out- fluids or vapors might escape by leak- put is essential for safe operation has age of a fluid system, there must be been shut off by the automatic means means to minimize the probability of prescribed in paragraph (e)(1) of this ignition of the fluids and vapors, and section. the resultant hazards if ignition does (f) Air intakes. Each combustion and occur. ventilating air intake must be located (b) Compliance with paragraph (a) of so that no flammable fluids or vapors this section must be shown by analysis can enter the heater system under any or tests, and the following factors must operating condition— be considered: (1) During normal operation; or (1) Possible sources and paths of fluid (2) As a result of the malfunctioning leakage, and means of detecting leak- of any other component. age. (g) Heater exhaust. Heater exhaust (2) Flammability characteristics of systems must meet the provisions of fluids, including effects of any combus- §§ 25.1121 and 25.1123. In addition, there tible or absorbing materials. must be provisions in the design of the (3) Possible ignition sources, includ- heater exhaust system to safely expel ing electrical faults, overheating of the products of combustion to prevent equipment, and malfunctioning of pro- the occurrence of— tective devices. (1) Fuel leakage from the exhaust to (4) Means available for controlling or surrounding compartments; extinguishing a fire, such as stopping (2) Exhaust gas impingement on sur- flow of fluids, shutting down equip- rounding equipment or structure; ment, fireproof containment, or use of (3) Ignition of flammable fluids by extinguishing agents. the exhaust, if the exhaust is in a com- (5) Ability of airplane components partment containing flammable fluid that are critical to safety of flight to lines; and withstand fire and heat. (4) Restriction by the exhaust of the (c) If action by the flight crew is re- prompt relief of backfires that, if so required to prevent or counteract a fluid stricted, could cause heater failure. fire (e.g., equipment shutdown or actu- (h) Heater fuel systems. Each heater ation of a fire extinguisher) quick act- fuel system must meet each power- ing means must be provided to alert plant fuel system requirement affect- the crew. ing safe heater operation. Each heater (d) Each area where flammable fluids fuel system component within the ven- or vapors might escape by leakage of a tilating airstream must be protected fluid system must be identified and de- by shrouds so that no leakage from fined. those components can enter the ven- [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970, as tilating airstream. amended by Amdt. 25–46, 43 FR 50597, Oct. 30, (i) Drains. There must be means to 1978] safely drain fuel that might accumulate within the combustion chamber or § 25.865 Fire protection of flight con- the heat exchanger. In addition— trols, engine mounts, and other (1) Each part of any drain that oper- flight structure. ates at high temperatures must be pro- Essential flight controls, engine tected in the same manner as heater mounts, and other flight structures lo- exhausts; and cated in designated fire zones or in ad- (2) Each drain must be protected jacent areas which would be subjected from hazardous ice accumulation under to the effects of fire in the fire zone any operating condition. must be constructed of fireproof material or shielded so that they are capa- [Doc. No. 5066, 29 FR 18291, Dec. 24 1964, as ble of withstanding the effects of fire. amended by Amdt. 25–11, 32 FR 6912, May 5, 1967; Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970]

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§ 25.867 Fire protection: other compo- MISCELLANEOUS nents. § 25.871 Leveling means. (a) Surfaces to the rear of the nacelles, within one nacelle diameter of There must be means for determining when the airplane is in a level position the nacelle centerline, must be at least on the ground. fire-resistant. (b) Paragraph (a) of this section does [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] not apply to tail surfaces to the rear of the nacelles that could not be readily § 25.875 Reinforcement near propellers. affected by heat, flames, or sparks coming from a designated fire zone or (a) Each part of the airplane near the engine compartment of any nacelle. propeller tips must be strong and stiff enough to withstand the effects of the [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] induced vibration and of ice thrown from the propeller. § 25.869 Fire protection: systems. (b) No window may be near the pro- (a) Electrical system components: peller tips unless it can withstand the (1) Components of the electrical sys- most severe ice impact likely to occur. tem must meet the applicable fire and § 25.899 Electrical bonding and protec- smoke protection requirements of tion against static electricity. §§ 25.831(c) and 25.863. (a) Electrical bonding and protection (2) Equipment that is located in des- against static electricity must be de- ignated fire zones and is used during signed to minimize accumulation of emergency procedures must be at least electrostatic charge that would cause— fire resistant. (1) Human injury from electrical (3) EWIS components must meet the shock, requirements of § 25.1713. (2) Ignition of flammable vapors, or (b) Each vacuum air system line and (3) Interference with installed elec- fitting on the discharge side of the trical/electronic equipment. pump that might contain flammable (b) Compliance with paragraph (a) of vapors or fluids must meet the require- this section may be shown by— ments of § 25.1183 if the line or fitting is (1) Bonding the components properly in a designated fire zone. Other vacuum to the airframe; or air systems components in designated (2) Incorporating other acceptable fire zones must be at least fire resist- means to dissipate the static charge so ant. as not to endanger the airplane, per- (c) Oxygen equipment and lines sonnel, or operation of the installed must— electrical/electronic systems. (1) Not be located in any designated [Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] fire zone, (2) Be protected from heat that may Subpart E—Powerplant be generated in, or escape from, any designated fire zone, and GENERAL (3) Be installed so that escaping oxy- § 25.901 Installation. gen cannot cause ignition of grease, fluid, or vapor accumulations that are (a) For the purpose of this part, the present in normal operation or as a re- airplane powerplant installation includes each component that— sult of failure or malfunction of any (1) Is necessary for propulsion; system. (2) Affects the control of the major [Amdt. 25–72, 55 FR 29784, July 20, 1990, as propulsive units; or amended by Amdt. 25–113, 69 FR 12530, Mar. (3) Affects the safety of the major 16, 2004; Amdt. 25–123, 72 FR 63405, Nov. 8, propulsive units between normal in- 2007] spections or overhauls. (b) For each powerplant— (1) The installation must comply with—

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(i) The installation instructions pro- NOTE: § 33.77 of this chapter in effect on Oc- vided under §§ 33.5 and 35.3 of this chap- tober 31, 1974, was published in 14 CFR parts ter; and 1 to 59, Revised as of January 1, 1975. See 39 FR 35467, October 1, 1974. (ii) The applicable provisions of this subpart; (b) Engine isolation. The powerplants (2) The components of the installa- must be arranged and isolated from tion must be constructed, arranged, each other to allow operation, in at and installed so as to ensure their con- least one configuration, so that the tinued safe operation between normal failure or malfunction of any engine, or inspections or overhauls; of any system that can affect the en- (3) The installation must be acces- gine, will not— sible for necessary inspections and (1) Prevent the continued safe oper- maintenance; and ation of the remaining engines; or (4) The major components of the in- (2) Require immediate action by any stallation must be electrically bonded crewmember for continued safe oper- to the other parts of the airplane. ation. (c) For each powerplant and auxiliary (c) Control of engine rotation. There power unit installation, it must be es- must be means for stopping the rota- tablished that no single failure or mal- tion of any engine individually in function or probable combination of flight, except that, for turbine engine failures will jeopardize the safe oper- installations, the means for stopping ation of the airplane except that the the rotation of any engine need be pro- failure of structural elements need not vided only where continued rotation be considered if the probability of such could jeopardize the safety of the air- failure is extremely remote. plane. Each component of the stopping (d) Each auxiliary power unit instal- system on the engine side of the fire- lation must meet the applicable provi- wall that might be exposed to fire must sions of this subpart. be at least fire-resistant. If hydraulic propeller feathering systems are used [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as for this purpose, the feathering lines amended by Amdt. 25–23, 35 FR 5676, Apr. 8, must be at least fire resistant under 1970; Amdt. 25–40, 42 FR 15042, Mar. 17, 1977; the operating conditions that may be Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. expected to exist during feathering. 25–126, 73 FR 63345, Oct. 24, 2008] (d) Turbine engine installations. For § 25.903 Engines. turbine engine installations— (1) Design precautions must be taken (a) Engine type certificate. (1) Each en- to minimize the hazards to the airplane gine must have a type certificate and in the event of an engine rotor failure must meet the applicable requirements or of a fire originating within the en- of part 34 of this chapter. gine which burns through the engine (2) Each turbine engine must comply case. with one of the following: (2) The powerplant systems associ- (i) Sections 33.76, 33.77 and 33.78 of ated with engine control devices, sys- this chapter in effect on December 13, tems, and instrumentation, must be de- 2000, or as subsequently amended; or signed to give reasonable assurance (ii) Sections 33.77 and 33.78 of this that those engine operating limitations chapter in effect on April 30, 1998, or as that adversely affect turbine rotor subsequently amended before Decem- structural integrity will not be exceed- ber 13, 2000; or ed in service. (iii) Comply with § 33.77 of this chap- (e) Restart capability. (1) Means to re- ter in effect on October 31, 1974, or as start any engine in flight must be pro- subsequently amended prior to April vided. 30, 1998, unless that engine’s foreign ob- (2) An altitude and airspeed envelope ject ingestion service history has re- must be established for in-flight engine sulted in an unsafe condition; or restarting, and each engine must have (iv) Be shown to have a foreign object a restart capability within that enve- ingestion service history in similar in- lope. stallation locations which has not re- (3) For turbine engine powered air- sulted in any unsafe condition. planes, if the minimum windmilling

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speed of the engines, following the § 25.907 Propeller vibration and fa- inflight shutdown of all engines, is in- tigue. sufficient to provide the necessary This section does not apply to fixed- electrical power for engine ignition, a pitch wood propellers of conventional power source independent of the en- design. gine-driven electrical power generating (a) The applicant must determine the system must be provided to permit in- magnitude of the propeller vibration flight engine ignition for restarting. stresses or loads, including any stress (f) Auxiliary Power Unit. Each auxil- peaks and resonant conditions, iary power unit must be approved or throughout the operational envelope of meet the requirements of the category the airplane by either: for its intended use. (1) Measurement of stresses or loads [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as through direct testing or analysis amended by Amdt. 25–23, 35 FR 5676, Apr. 8, based on direct testing of the propeller 1970; Amdt. 25–40, 42 FR 15042, Mar. 17, 1977; on the airplane and engine installation Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; Amdt. for which approval is sought; or 25–72, 55 FR 29784, July 20, 1990; Amdt. 25–73, (2) Comparison of the propeller to 55 FR 32861, Aug. 10, 1990; Amdt. 25–94, 63 FR similar propellers installed on similar 8848, Feb. 23, 1998; Amdt. 25–95, 63 FR 14798, airplane installations for which these Mar. 26, 1998; Amdt. 25–100, 65 FR 55854, Sept. 14, 2000] measurements have been made. (b) The applicant must demonstrate § 25.904 Automatic takeoff thrust con- by tests, analysis based on tests, or trol system (ATTCS). previous experience on similar designs that the propeller does not experience Each applicant seeking approval for harmful effects of flutter throughout installation of an engine power control the operational envelope of the air- system that automatically resets the plane. power or thrust on the operating en- (c) The applicant must perform an gine(s) when any engine fails during evaluation of the propeller to show the takeoff must comply with the re- that failure due to fatigue will be quirements of appendix I of this part. avoided throughout the operational life [Amdt. 25–62, 52 FR 43156, Nov. 9, 1987] of the propeller using the fatigue and structural data obtained in accordance § 25.905 Propellers. with part 35 of this chapter and the vi- (a) Each propeller must have a type bration data obtained from compliance certificate. with paragraph (a) of this section. For (b) Engine power and propeller shaft the purpose of this paragraph, the pro- rotational speed may not exceed the peller includes the hub, blades, blade limits for which the propeller is certifi- retention component and any other cated. propeller component whose failure due (c) The propeller blade pitch control to fatigue could be catastrophic to the system must meet the requirements of airplane. This evaluation must include: §§ 35.21, 35.23, 35.42 and 35.43 of this (1) The intended loading spectra in- chapter. cluding all reasonably foreseeable propeller vibration and cyclic load pat- (d) Design precautions must be taken terns, identified emergency conditions, to minimize the hazards to the airplane allowable overspeeds and overtorques, in the event a propeller blade fails or is and the effects of temperatures and hu- released by a hub failure. The hazards midity expected in service. which must be considered include dam- (2) The effects of airplane and pro- age to structure and vital systems due peller operating and airworthiness lim- to impact of a failed or released blade itations. and the unbalance created by such failure or release. [Amdt. 25–126, 73 FR 63345, Oct. 24, 2008]

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.925 Propeller clearance. amended by Amdt. 25–54, 45 FR 60173, Sept. 11, 1980; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; Unless smaller clearances are sub- Amdt. 25–72, 55 FR 29784, July 20, 1990; Amdt. stantiated, propeller clearances with 25–126, 73 FR 63345, Oct. 24, 2008] the airplane at maximum weight, with

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the most adverse center of gravity, and that during any reversal in flight the with the propeller in the most adverse engine will produce no more than flight pitch position, may not be less than idle thrust. In addition, it must be the following: shown by analysis or test, or both, (a) Ground clearance. There must be a that— clearance of at least seven inches (for (i) Each operable reverser can be re- each airplane with nose wheel landing stored to the forward thrust position; gear) or nine inches (for each airplane and with tail wheel landing gear) between (ii) The airplane is capable of contin- each propeller and the ground with the ued safe flight and landing under any landing gear statically deflected and in possible position of the thrust reverser. the level takeoff, or taxiing attitude, (2) Each system intended for inflight whichever is most critical. In addition, use must be designed so that no unsafe there must be positive clearance be- condition will result during normal op- tween the propeller and the ground eration of the system, or from any fail- when in the level takeoff attitude with ure (or reasonably likely combination the critical tire(s) completely deflated of failures) of the reversing system, and the corresponding landing gear under any anticipated condition of op- strut bottomed. eration of the airplane including (b) Water clearance. There must be a ground operation. Failure of structural clearance of at least 18 inches between elements need not be considered if the each propeller and the water, unless probability of this kind of failure is ex- compliance with § 25.239(a) can be tremely remote. shown with a lesser clearance. (3) Each system must have means to (c) Structural clearance. There must prevent the engine from producing be— more than idle thrust when the revers- (1) At least one inch radial clearance ing system malfunctions, except that it between the blade tips and the airplane may produce any greater forward structure, plus any additional radial thrust that is shown to allow direc- clearance necessary to prevent harmful tional control to be maintained, with vibration; (2) At least one-half inch longitudinal aerodynamic means alone, under the clearance between the propeller blades most critical reversing condition ex- or cuffs and stationary parts of the air- pected in operation. plane; and (b) For propeller reversing systems— (3) Positive clearance between other (1) Each system intended for ground rotating parts of the propeller or spin- operation only must be designed so ner and stationary parts of the air- that no single failure (or reasonably plane. likely combination of failures) or malfunction of the system will result in [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as unwanted reverse thrust under any ex- amended by Amdt. 25–72, 55 FR 29784, July 20, pected operating condition. Failure of 1990] structural elements need not be consid- § 25.929 Propeller deicing. ered if this kind of failure is extremely remote. (a) For airplanes intended for use (2) Compliance with this section may where icing may be expected, there be shown by failure analysis or testing, must be a means to prevent or remove or both, for propeller systems that hazardous ice accumulation on propel- allow propeller blades to move from lers or on accessories where ice accu- the flight low-pitch position to a posi- mulation would jeopardize engine per- tion that is substantially less than formance. that at the normal flight low-pitch po- (b) If combustible fluid is used for sition. The analysis may include or be propeller deicing, §§ 25.1181 through supported by the analysis made to 25.1185 and 25.1189 apply. show compliance with the require- § 25.933 Reversing systems. ments of § 35.21 of this chapter for the propeller and associated installation (a) For turbojet reversing systems— components. (1) Each system intended for ground operation only must be designed so [Amdt. 25–72, 55 FR 29784, July 20, 1990]

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§ 25.934 Turbojet engine thrust re- odynamic control of the airplane may verser system tests. not result in any condition that would Thrust reversers installed on tur- require exceptional skill, alertness, or bojet engines must meet the require- strength on the part of the pilot to ments of § 33.97 of this chapter. avoid exceeding an operational or structural limitation of the airplane; [Amdt. 25–23, 35 FR 5677, Apr. 8, 1970] and (c) In showing compliance with para- § 25.937 Turbopropeller-drag limiting graph (b) of this section, the pilot systems. strength required may not exceed the Turbopropeller power airplane pro- limits set forth in § 25.143(d), subject to peller-drag limiting systems must be the conditions set forth in paragraphs designed so that no single failure or (e) and (f) of § 25.143. malfunction of any of the systems during normal or emergency operation re- [Amdt. 25–38, 41 FR 55467, Dec. 20, 1976, as sults in propeller drag in excess of that amended by Amdt. 25–121, 72 FR 44669, Aug. 8, 2007] for which the airplane was designed under § 25.367. Failure of structural ele- § 25.943 Negative acceleration. ments of the drag limiting systems need not be considered if the prob- No hazardous malfunction of an en- ability of this kind of failure is ex- gine, an auxiliary power unit approved tremely remote. for use in flight, or any component or system associated with the powerplant § 25.939 Turbine engine operating or auxiliary power unit may occur characteristics. when the airplane is operated at the (a) Turbine engine operating charac- negative accelerations within the teristics must be investigated in flight flight envelopes prescribed in § 25.333. to determine that no adverse charac- This must be shown for the greatest teristics (such as stall, surge, or flame- duration expected for the acceleration. out) are present, to a hazardous degree, [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] during normal and emergency operation within the range of operating § 25.945 Thrust or power augmentation limitations of the airplane and of the system. engine. (a) General. Each fluid injection sys- (b) [Reserved] tem must provide a flow of fluid at the (c) The turbine engine air inlet sys- rate and pressure established for proper tem may not, as a result of air flow dis- engine functioning under each intended tortion during normal operation, cause operating condition. If the fluid can vibration harmful to the engine. freeze, fluid freezing may not damage [Amdt. 25–11, 32 FR 6912, May 5, 1967, as the airplane or adversely affect air- amended by Amdt. 25–40, 42 FR 15043, Mar. 17, plane performance. 1977] (b) Fluid tanks. Each augmentation system fluid tank must meet the fol- § 25.941 Inlet, engine, and exhaust lowing requirements: compatibility. (1) Each tank must be able to with- For airplanes using variable inlet or stand without failure the vibration, in- exhaust system geometry, or both— ertia, fluid, and structural loads that it (a) The system comprised of the may be subject to in operation. inlet, engine (including thrust aug- (2) The tanks as mounted in the air- mentation systems, if incorporated), plane must be able to withstand with- and exhaust must be shown to function out failure or leakage an internal pres- properly under all operating conditions sure 1.5 times the maximum operating for which approval is sought, including pressure. all engine rotating speeds and power (3) If a vent is provided, the venting settings, and engine inlet and exhaust must be effective under all normal configurations; flight conditions. (b) The dynamic effects of the oper- (4) [Reserved] ation of these (including consideration (5) Each tank must have an expan- of probable malfunctions) upon the aer- sion space of not less than 2 percent of

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the tank capacity. It must be impos- (d) Each fuel system for a turbine en- sible to fill the expansion space inad- gine powered airplane must meet the vertently with the airplane in the nor- applicable fuel venting requirements of mal ground attitude. part 34 of this chapter. (c) Augmentation system drains [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as must be designed and located in ac- amended by Amdt. 25–23, 35 FR 5677, Apr. 8, cordance with § 25.1455 if— 1970; Amdt. 25–36, 39 FR 35460, Oct. 1, 1974; (1) The augmentation system fluid is Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. subject to freezing; and 25–73, 55 FR 32861, Aug. 10, 1990] (2) The fluid may be drained in flight or during ground operation. § 25.952 Fuel system analysis and test. (d) The augmentation liquid tank ca- (a) Proper fuel system functioning pacity available for the use of each en- under all probable operating conditions gine must be large enough to allow op- must be shown by analysis and those eration of the airplane under the ap- tests found necessary by the Adminis- proved procedures for the use of liquid- trator. Tests, if required, must be made augmented power. The computation of using the airplane fuel system or a test liquid consumption must be based on article that reproduces the operating the maximum approved rate appro- characteristics of the portion of the priate for the desired engine output fuel system to be tested. and must include the effect of tempera- (b) The likely failure of any heat ex- ture on engine performance as well as changer using fuel as one of its fluids any other factors that might vary the may not result in a hazardous condi- amount of liquid required. tion. (e) This section does not apply to fuel injection systems. [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977, as § 25.953 Fuel system independence. amended by Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–115, 69 FR 40527, July 2, 2004] Each fuel system must meet the requirements of § 25.903(b) by— FUEL SYSTEM (a) Allowing the supply of fuel to each engine through a system inde- § 25.951 General. pendent of each part of the system sup- (a) Each fuel system must be con- plying fuel to any other engine; or structed and arranged to ensure a flow (b) Any other acceptable method. of fuel at a rate and pressure established for proper engine and auxiliary § 25.954 Fuel system lightning protec- power unit functioning under each tion. likely operating condition, including The fuel system must be designed any maneuver for which certification is and arranged to prevent the ignition of requested and during which the engine fuel vapor within the system by— or auxiliary power unit is permitted to (a) Direct lightning strikes to areas be in operation. having a high probability of stroke at- (b) Each fuel system must be ar- tachment; ranged so that any air which is intro- (b) Swept lightning strokes to areas duced into the system will not result where swept strokes are highly prob- in— able; and (1) Power interruption for more than (c) Corona and streamering at fuel 20 seconds for reciprocating engines; or vent outlets. (2) Flameout for turbine engines. (c) Each fuel system for a turbine en- [Amdt. 25–14, 32 FR 11629, Aug. 11, 1967] gine must be capable of sustained operation throughout its flow and pressure § 25.955 Fuel flow. range with fuel initially saturated with (a) Each fuel system must provide at water at 80 °F and having 0.75cc of free least 100 percent of the fuel flow re- water per gallon added and cooled to quired under each intended operating the most critical condition for icing condition and maneuver. Compliance likely to be encountered in operation. must be shown as follows:

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(1) Fuel must be delivered to each en- than the quantity at which the first gine at a pressure within the limits evidence of engine malfunction occurs specified in the engine type certificate. under the most adverse fuel feed condi- (2) The quantity of fuel in the tank tion for all intended operations and may not exceed the amount established flight maneuvers involving fuel feeding as the unusable fuel supply for that from that tank. Fuel system compo- tank under the requirements of § 25.959 nent failures need not be considered. plus that necessary to show compliance with this section. [Amdt. 25–23, 35 FR 5677, Apr. 8, 1970, as amended by Amdt. 25–40, 42 FR 15043, Mar. 17, (3) Each main pump must be used 1977] that is necessary for each operating condition and attitude for which com- § 25.961 Fuel system hot weather oper- pliance with this section is shown, and ation. the appropriate emergency pump must (a) The fuel system must perform be substituted for each main pump so satisfactorily in hot weather operation. used. This must be shown by showing that (4) If there is a fuel flowmeter, it the fuel system from the tank outlets must be blocked and the fuel must flow to each engine is pressurized, under all through the meter or its bypass. intended operations, so as to prevent (b) If an engine can be supplied with vapor formation, or must be shown by fuel from more than one tank, the fuel climbing from the altitude of the air- system must— port elected by the applicant to the (1) For each reciprocating engine, maximum altitude established as an supply the full fuel pressure to that en- operating limitation under § 25.1527. If a gine in not more than 20 seconds after climb test is elected, there may be no switching to any other fuel tank con- evidence of vapor lock or other mal- taining usable fuel when engine malfunctioning during the climb test con- functioning becomes apparent due to ducted under the following conditions: the depletion of the fuel supply in any (1) For reciprocating engine powered tank from which the engine can be fed; airplanes, the engines must operate at and maximum continuous power, except (2) For each turbine engine, in addi- that takeoff power must be used for the tion to having appropriate manual altitudes from 1,000 feet below the crit- switching capability, be designed to ical altitude through the critical alti- prevent interruption of fuel flow to tude. The time interval during which that engine, without attention by the takeoff power is used may not be less flight crew, when any tank supplying than the takeoff time limitation. fuel to that engine is depleted of usable (2) For turbine engine powered air- fuel during normal operation, and any planes, the engines must operate at other tank, that normally supplies fuel takeoff power for the time interval se- to that engine alone, contains usable lected for showing the takeoff flight fuel. path, and at maximum continuous [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as power for the rest of the climb. amended by Amdt. 25–11, 32 FR 6912, May 5, (3) The weight of the airplane must 1967] be the weight with full fuel tanks, minimum crew, and the ballast necessary § 25.957 Flow between interconnected to maintain the center of gravity with- tanks. in allowable limits. If fuel can be pumped from one tank (4) The climb airspeed may not ex- to another in flight, the fuel tank ceed— vents and the fuel transfer system (i) For reciprocating engine powered must be designed so that no structural airplanes, the maximum airspeed es- damage to the tanks can occur because tablished for climbing from takeoff to of overfilling. the maximum operating altitude with the airplane in the following configura- § 25.959 Unusable fuel supply. tion: The unusable fuel quantity for each (A) Landing gear retracted. fuel tank and its fuel system compo- (B) Wing flaps in the most favorable nents must be established at not less position.

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(C) Cowl flaps (or other means of con- ments, low energy engine debris, or trolling the engine cooling supply) in other likely debris. the position that provides adequate (2) All covers must be fire resistant cooling in the hot-day condition. as defined in part 1 of this chapter. (D) Engine operating within the max- (f) For pressurized fuel tanks, a imum continuous power limitations. means with fail-safe features must be (E) Maximum takeoff weight; and provided to prevent the buildup of an (ii) For turbine engine powered air- excessive pressure difference between planes, the maximum airspeed estab- the inside and the outside of the tank. lished for climbing from takeoff to the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as maximum operating altitude. amended by Amdt. 25–40, 42 FR 15043, Mar. 17, (5) The fuel temperature must be at 1977; Amdt. 25–69, 54 FR 40354, Sept. 29, 1989] least 110 °F. (b) The test prescribed in paragraph § 25.965 Fuel tank tests. (a) of this section may be performed in (a) It must be shown by tests that the flight or on the ground under closely fuel tanks, as mounted in the airplane, simulated flight conditions. If a flight can withstand, without failure or leak- test is performed in weather cold age, the more critical of the pressures enough to interfere with the proper resulting from the conditions specified conduct of the test, the fuel tank sur- in paragraphs (a)(1) and (2) of this sec- faces, fuel lines, and other fuel system tion. In addition, it must be shown by parts subject to cold air must be insu- either analysis or tests, that tank sur- lated to simulate, insofar as prac- faces subjected to more critical pres- ticable, flight in hot weather. sures resulting from the condition of [Amdt. 25–11, 32 FR 6912, May 5, 1967, as paragraphs (a)(3) and (4) of this section, amended by Amdt. 25–57, 49 FR 6848, Feb. 23, are able to withstand the following 1984] pressures: (1) An internal pressure of 3.5 psi. § 25.963 Fuel tanks: general. (2) 125 percent of the maximum air (a) Each fuel tank must be able to pressure developed in the tank from withstand, without failure, the vibra- ram effect. tion, inertia, fluid, and structural loads (3) Fluid pressures developed during that it may be subjected to in oper- maximum limit accelerations, and de- ation. flections, of the airplane with a full (b) Flexible fuel tank liners must be tank. approved or must be shown to be suit- (4) Fluid pressures developed during able for the particular application. the most adverse combination of air- (c) Integral fuel tanks must have fa- plane roll and fuel load. cilities for interior inspection and re- (b) Each metallic tank with large un- pair. supported or unstiffened flat surfaces, (d) Fuel tanks within the fuselage whose failure or deformation could contour must be able to resist rupture cause fuel leakage, must be able to and to retain fuel, under the inertia withstand the following test, or its forces prescribed for the emergency equivalent, without leakage or exces- landing conditions in § 25.561. In addi- sive deformation of the tank walls: tion, these tanks must be in a pro- (1) Each complete tank assembly and tected position so that exposure of the its supports must be vibration tested tanks to scraping action with the while mounted to simulate the actual ground is unlikely. installation. (e) Fuel tank access covers must (2) Except as specified in paragraph comply with the following criteria in (b)(4) of this section, the tank assembly order to avoid loss of hazardous quan- must be vibrated for 25 hours at an am- tities of fuel: plitude of not less than 1⁄32 of an inch (1) All covers located in an area (unless another amplitude is substan- where experience or analysis indicates tiated) while 2⁄3 filled with water or a strike is likely must be shown by other suitable test fluid. analysis or tests to minimize penetra- (3) The test frequency of vibration tion and deformation by tire frag- must be as follows:

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(i) If no frequency of vibration result- (1) There must be pads, if necessary, ing from any r.p.m. within the normal to prevent chafing between the tank operating range of engine speeds is and its supports; critical, the test frequency of vibration (2) Padding must be nonabsorbent or must be 2,000 cycles per minute. treated to prevent the absorption of (ii) If only one frequency of vibration fluids; resulting from any r.p.m. within the (3) If a flexible tank liner is used, it normal operating range of engine must be supported so that it is not re- speeds is critical, that frequency of vi- quired to withstand fluid loads; and bration must be the test frequency. (4) Each interior surface of the tank compartment must be smooth and free (iii) If more than one frequency of vi- of projections that could cause wear of bration resulting from any r.p.m. with- the liner unless— in the normal operating range of en- (i) Provisions are made for protection gine speeds is critical, the most crit- of the liner at these points; or ical of these frequencies must be the (ii) The construction of the liner test frequency. itself provides that protection. (4) Under paragraphs (b)(3)(ii) and (b) Spaces adjacent to tank surfaces (iii) of this section, the time of test must be ventilated to avoid fume accu- must be adjusted to accomplish the mulation due to minor leakage. If the same number of vibration cycles that tank is in a sealed compartment, ven- would be accomplished in 25 hours at tilation may be limited to drain holes the frequency specified in paragraph large enough to prevent excessive pres- (b)(3)(i) of this section. sure resulting from altitude changes. (5) During the test, the tank assem- (c) The location of each tank must bly must be rocked at the rate of 16 to meet the requirements of § 25.1185(a). 20 complete cycles per minute, through (d) No engine nacelle skin imme- an angle of 15° on both sides of the hor- diately behind a major air outlet from izontal (30° total), about the most crit- the engine compartment may act as ical axis, for 25 hours. If motion about the wall of an integral tank. more than one axis is likely to be crit- (e) Each fuel tank must be isolated ical, the tank must be rocked about from personnel compartments by a fumeproof and fuelproof enclosure. each critical axis for 121⁄2 hours. (c) Except where satisfactory oper- § 25.969 Fuel tank expansion space. ating experience with a similar tank in Each fuel tank must have an expan- a similar installation is shown, non- sion space of not less than 2 percent of metallic tanks must withstand the test the tank capacity. It must be impos- specified in paragraph (b)(5) of this sec- sible to fill the expansion space inad- tion, with fuel at a temperature of 110 vertently with the airplane in the nor- ° F. During this test, a representative mal ground attitude. For pressure fuel- specimen of the tank must be installed ing systems, compliance with this sec- in a supporting structure simulating tion may be shown with the means pro- the installation in the airplane. vided to comply with § 25.979(b). (d) For pressurized fuel tanks, it must be shown by analysis or tests [Amdt. 25–11, 32 FR 6913, May 5, 1967] that the fuel tanks can withstand the § 25.971 Fuel tank sump. maximum pressure likely to occur on the ground or in flight. (a) Each fuel tank must have a sump with an effective capacity, in the nor- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as mal ground attitude, of not less than amended by Amdt. 25–11, 32 FR 6913, May 5, the greater of 0.10 percent of the tank 1967; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] capacity or one-sixteenth of a gallon unless operating limitations are estab- § 25.967 Fuel tank installations. lished to ensure that the accumulation (a) Each fuel tank must be supported of water in service will not exceed the so that tank loads (resulting from the sump capacity. weight of the fuel in the tanks) are not (b) Each fuel tank must allow drain- concentrated on unsupported tank sur- age of any hazardous quantity of water faces. In addition— from any part of the tank to its sump

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with the airplane in the ground atti- (5) There may be no point in any vent tude. line where moisture can accumulate (c) Each fuel tank sump must have with the airplane in the ground atti- an accessible drain that— tude or the level flight attitude, unless (1) Allows complete drainage of the drainage is provided; and sump on the ground; (6) No vent or drainage provision may (2) Discharges clear of each part of end at any point— the airplane; and (i) Where the discharge of fuel from (3) Has manual or automatic means the vent outlet would constitute a fire for positive locking in the closed posi- hazard; or tion. (ii) From which fumes could enter personnel compartments. § 25.973 Fuel tank filler connection. (b) Carburetor vapor vents. Each car- Each fuel tank filler connection must buretor with vapor elimination connec- prevent the entrance of fuel into any tions must have a vent line to lead va- part of the airplane other than the pors back to one of the fuel tanks. In tank itself. In addition— addition— (a) [Reserved] (1) Each vent system must have (b) Each recessed filler connection means to avoid stoppage by ice; and that can retain any appreciable quan- (2) If there is more than one fuel tity of fuel must have a drain that dis- tank, and it is necessary to use the charges clear of each part of the air- tanks in a definite sequence, each plane; vapor vent return line must lead back (c) Each filler cap must provide a to the fuel tank used for takeoff and fuel-tight seal; and landing. (d) Each fuel filling point must have a provision for electrically bonding the § 25.977 Fuel tank outlet. airplane to ground fueling equipment. (a) There must be a fuel strainer for [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the fuel tank outlet or for the booster amended by Amdt. 25–40, 42 FR 15043, Mar. 17, pump. This strainer must— 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990; (1) For reciprocating engine powered Amdt. 25–115, 69 FR 40527, July 2, 2004] airplanes, have 8 to 16 meshes per inch; and § 25.975 Fuel tank vents and carburetor vapor vents. (2) For turbine engine powered airplanes, prevent the passage of any ob- (a) Fuel tank vents. Each fuel tank ject that could restrict fuel flow or must be vented from the top part of the damage any fuel system component. expansion space so that venting is ef- (b) [Reserved] fective under any normal flight condi- (c) The clear area of each fuel tank tion. In addition— outlet strainer must be at least five (1) Each vent must be arranged to times the area of the outlet line. avoid stoppage by dirt or ice forma- (d) The diameter of each strainer tion; must be at least that of the fuel tank (2) The vent arrangement must pre- outlet. vent siphoning of fuel during normal (e) Each finger strainer must be ac- operation; cessible for inspection and cleaning. (3) The venting capacity and vent pressure levels must maintain accept- [Amdt. 25–11, 32 FR 6913, May 5, 1967, as able differences of pressure between amended by Amdt. 25–36, 39 FR 35460, Oct. 1, the interior and exterior of the tank, 1974] during— (i) Normal flight operation; § 25.979 Pressure fueling system. (ii) Maximum rate of ascent and de- For pressure fueling systems, the fol- scent; and lowing apply: (iii) Refueling and defueling (where (a) Each pressure fueling system fuel applicable); manifold connection must have means (4) Airspaces of tanks with inter- to prevent the escape of hazardous connected outlets must be inter- quantities of fuel from the system if connected; the fuel entry valve fails.

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(b) An automatic shutoff means must ation, failure, or malfunction could in- be provided to prevent the quantity of crease the temperature inside the tank. fuel in each tank from exceeding the (3) Demonstrating that an ignition maximum quantity approved for that source could not result from each sin- tank. This means must— gle failure, from each single failure in (1) Allow checking for proper shutoff combination with each latent failure operation before each fueling of the condition not shown to be extremely tank; and remote, and from all combinations of (2) Provide indication at each fueling failures not shown to be extremely im- station of failure of the shutoff means probable. The effects of manufacturing to stop the fuel flow at the maximum variability, aging, wear, corrosion, and quantity approved for that tank. likely damage must be considered. (c) A means must be provided to pre- (b) Except as provided in paragraphs vent damage to the fuel system in the (b)(2) and (c) of this section, no fuel event of failure of the automatic shut- tank Fleet Average Flammability Ex- off means prescribed in paragraph (b) posure on an airplane may exceed three of this section. percent of the Flammability Exposure (d) The airplane pressure fueling sys- Evaluation Time (FEET) as defined in tem (not including fuel tanks and fuel Appendix N of this part, or that of a tank vents) must withstand an ulti- fuel tank within the wing of the air- mate load that is 2.0 times the load plane model being evaluated, which- arising from the maximum pressures, ever is greater. If the wing is not a con- including surge, that is likely to occur ventional unheated aluminum wing, during fueling. The maximum surge the analysis must be based on an as- pressure must be established with any sumed Equivalent Conventional combination of tank valves being ei- Unheated Aluminum Wing Tank. ther intentionally or inadvertently (1) Fleet Average Flammability Ex- closed. posure is determined in accordance (e) The airplane defueling system with Appendix N of this part. The as- (not including fuel tanks and fuel tank sessment must be done in accordance vents) must withstand an ultimate with the methods and procedures set load that is 2.0 times the load arising forth in the Fuel Tank Flammability from the maximum permissible Assessment Method User’s Manual, defueling pressure (positive or nega- dated May 2008, document number tive) at the airplane fueling connec- DOT/FAA/AR–05/8 (incorporated by ref- tion. erence, see § 25.5). (2) Any fuel tank other than a main [Amdt. 25–11, 32 FR 6913, May 5, 1967, as fuel tank on an airplane must meet the amended by Amdt. 25–38, 41 FR 55467, Dec. 20, flammability exposure criteria of Ap- 1976; Amdt. 25–72, 55 FR 29785, July 20, 1990] pendix M to this part if any portion of the tank is located within the fuselage § 25.981 Fuel tank ignition prevention. contour. (a) No ignition source may be present (3) As used in this paragraph, at each point in the fuel tank or fuel (i) Equivalent Conventional Unheated tank system where catastrophic failure Aluminum Wing Tank is an integral could occur due to ignition of fuel or tank in an unheated semi-monocoque vapors. This must be shown by: aluminum wing of a subsonic airplane (1) Determining the highest tempera- that is equivalent in aerodynamic per- ture allowing a safe margin below the formance, structural capability, fuel lowest expected autoignition tempera- tank capacity and tank configuration ture of the fuel in the fuel tanks. to the designed wing. (2) Demonstrating that no tempera- (ii) Fleet Average Flammability Expo- ture at each place inside each fuel tank sure is defined in Appendix N to this where fuel ignition is possible will ex- part and means the percentage of time ceed the temperature determined under each fuel tank ullage is flammable for paragraph (a)(1) of this section. This a fleet of an airplane type operating must be verified under all probable op- over the range of flight lengths. erating, failure, and malfunction con- (iii) Main Fuel Tank means a fuel ditions of each component whose oper- tank that feeds fuel directly into one

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or more engines and holds required fuel (b) Emergency pumps. There must be reserves continually throughout each emergency pumps or another main flight. pump to feed each engine immediately (c) Paragraph (b) of this section does after failure of any main pump (other not apply to a fuel tank if means are than a fuel injection pump approved as provided to mitigate the effects of an part of the engine). ignition of fuel vapors within that fuel tank such that no damage caused by an § 25.993 Fuel system lines and fittings. ignition will prevent continued safe (a) Each fuel line must be installed flight and landing. (d) Critical design configuration con- and supported to prevent excessive vi- trol limitations (CDCCL), inspections, bration and to withstand loads due to or other procedures must be estab- fuel pressure and accelerated flight lished, as necessary, to prevent devel- conditions. opment of ignition sources within the (b) Each fuel line connected to com- fuel tank system pursuant to para- ponents of the airplane between which graph (a) of this section, to prevent in- relative motion could exist must have creasing the flammability exposure of provisions for flexibility. the tanks above that permitted under (c) Each flexible connection in fuel paragraph (b) of this section, and to lines that may be under pressure and prevent degradation of the perform- subjected to axial loading must use ance and reliability of any means pro- flexible hose assemblies. vided according to paragraphs (a) or (c) (d) Flexible hose must be approved or of this section. These CDCCL, inspec- must be shown to be suitable for the tions, and procedures must be included particular application. in the Airworthiness Limitations sec- (e) No flexible hose that might be ad- tion of the instructions for continued versely affected by exposure to high airworthiness required by § 25.1529. temperatures may be used where exces- Visible means of identifying critical features of the design must be placed in sive temperatures will exist during op- areas of the airplane where foreseeable eration or after engine shut-down. maintenance actions, repairs, or alter- (f) Each fuel line within the fuselage ations may compromise the critical de- must be designed and installed to allow sign configuration control limitations a reasonable degree of deformation and (e.g., color-coding of wire to identify stretching without leakage. separation limitation). These visible [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as means must also be identified as amended by Amdt. 25–15, 32 FR 13266, Sept. CDCCL. 20, 1967] [Doc. No. 1999–6411, 66 FR 23129, May 7, 2001, as amended at Doc. No. FAA–2005–22997, 73 § 25.994 Fuel system components. FR 42494, July 21, 2008] Fuel system components in an engine nacelle or in the fuselage must be pro- FUEL SYSTEM COMPONENTS tected from damage which could result § 25.991 Fuel pumps. in spillage of enough fuel to constitute (a) Main pumps. Each fuel pump re- a fire hazard as a result of a wheels-up quired for proper engine operation, or landing on a paved runway. required to meet the fuel system re- [Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] quirements of this subpart (other than those in paragraph (b) of this section, § 25.995 Fuel valves. is a main pump. For each main pump, In addition to the requirements of provision must be made to allow the bypass of each positive displacement § 25.1189 for shutoff means, each fuel fuel pump other than a fuel injection valve must— pump (a pump that supplies the proper (a) [Reserved] flow and pressure for fuel injection (b) Be supported so that no loads re- when the injection is not accomplished sulting from their operation or from in a carburetor) approved as part of the engine.

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accelerated flight conditions are trans- event of a landing with landing gear re- mitted to the lines attached to the tracted. valve. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55467, Dec. 20, amended by Amdt. 25–40, 42 FR 15043, Mar. 17, 1976] 1977] § 25.1001 Fuel jettisoning system. § 25.997 Fuel strainer or filter. (a) A fuel jettisoning system must be There must be a fuel strainer or filter installed on each airplane unless it is between the fuel tank outlet and the shown that the airplane meets the inlet of either the fuel metering device climb requirements of §§ 25.119 and or an engine driven positive displace- 25.121(d) at maximum takeoff weight, ment pump, whichever is nearer the less the actual or computed weight of fuel tank outlet. This fuel strainer or fuel necessary for a 15-minute flight filter must— comprised of a takeoff, go-around, and (a) Be accessible for draining and landing at the airport of departure cleaning and must incorporate a screen with the airplane configuration, speed, or element which is easily removable; power, and thrust the same as that (b) Have a sediment trap and drain used in meeting the applicable takeoff, except that it need not have a drain if approach, and landing climb perform- the strainer or filter is easily remov- ance requirements of this part. able for drain purposes; (b) If a fuel jettisoning system is re- (c) Be mounted so that its weight is quired it must be capable of jettisoning not supported by the connecting lines enough fuel within 15 minutes, starting or by the inlet or outlet connections of with the weight given in paragraph (a) the strainer or filter itself, unless ade- of this section, to enable the airplane quate strength margins under all load- to meet the climb requirements of ing conditions are provided in the lines §§ 25.119 and 25.121(d), assuming that the and connections; and fuel is jettisoned under the conditions, (d) Have the capacity (with respect to except weight, found least favorable operating limitations established for during the flight tests prescribed in the engine) to ensure that engine fuel paragraph (c) of this section. system functioning is not impaired, (c) Fuel jettisoning must be dem- with the fuel contaminated to a degree onstrated beginning at maximum take- (with respect to particle size and den- off weight with flaps and landing gear sity) that is greater than that estab- up and in— lished for the engine in Part 33 of this (1) A power-off glide at 1.3 VSR1; chapter. (2) A climb at the one-engine inoper- [Amdt. 25–36, 39 FR 35460, Oct. 1, 1974, as ative best rate-of-climb speed, with the amended by Amdt. 25–57, 49 FR 6848, Feb. 23, critical engine inoperative and the re- 1984] maining engines at maximum continuous power; and § 25.999 Fuel system drains. (3) Level flight at 1.3 V SR1; if the re- (a) Drainage of the fuel system must sults of the tests in the conditions be accomplished by the use of fuel specified in paragraphs (c)(1) and (2) of strainer and fuel tank sump drains. this section show that this condition (b) Each drain required by paragraph could be critical. (a) of this section must— (d) During the flight tests prescribed (1) Discharge clear of all parts of the in paragraph (c) of this section, it must airplane; be shown that— (2) Have manual or automatic means (1) The fuel jettisoning system and for positive locking in the closed posi- its operation are free from fire hazard; tion; and (2) The fuel discharges clear of any (3) Have a drain valve— part of the airplane; (i) That is readily accessible and (3) Fuel or fumes do not enter any which can be easily opened and closed; parts of the airplane; and and (4) The jettisoning operation does not (ii) That is either located or pro- adversely affect the controllability of tected to prevent fuel spillage in the the airplane.

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(e) For reciprocating engine powered (b) The usable oil capacity may not airplanes, means must be provided to be less than the product of the endur- prevent jettisoning the fuel in the ance of the airplane under critical op- tanks used for takeoff and landing erating conditions and the approved below the level allowing 45 minutes maximum allowable oil consumption of flight at 75 percent maximum contin- the engine under the same conditions, uous power. However, if there is an plus a suitable margin to ensure sys- auxiliary control independent of the tem circulation. Instead of a rational main jettisoning control, the system analysis of airplane range for the pur- may be designed to jettison the re- pose of computing oil requirements for maining fuel by means of the auxiliary reciprocating engine powered air- jettisoning control. planes, the following fuel/oil ratios (f) For turbine engine powered air- may be used: planes, means must be provided to pre- (1) For airplanes without a reserve vent jettisoning the fuel in the tanks oil or oil transfer system, a fuel/oil used for takeoff and landing below the ratio of 30:1 by volume. level allowing climb from sea level to (2) For airplanes with either a re- 10,000 feet and thereafter allowing 45 serve oil or oil transfer system, a fuel/ minutes cruise at a speed for maximum oil ratio of 40:1 by volume. range. However, if there is an auxiliary (c) Fuel/oil ratios higher than those control independent of the main jetti- prescribed in paragraphs (b)(1) and (2) soning control, the system may be de- of this section may be used if substan- signed to jettison the remaining fuel tiated by data on actual engine oil con- by means of the auxiliary jettisoning sumption. control. § 25.1013 Oil tanks. (g) The fuel jettisoning valve must be designed to allow flight personnel to (a) Installation. Each oil tank instal- close the valve during any part of the lation must meet the requirements of jettisoning operation. § 25.967. (b) Expansion space. Oil tank expan- (h) Unless it is shown that using any sion space must be provided as follows: means (including flaps, slots, and slats) (1) Each oil tank used with a recipro- for changing the airflow across or cating engine must have an expansion around the wings does not adversely af- space of not less than the greater of 10 fect fuel jettisoning, there must be a percent of the tank capacity or 0.5 gal- placard, adjacent to the jettisoning lon, and each oil tank used with a tur- control, to warn flight crewmembers bine engine must have an expansion against jettisoning fuel while the space of not less than 10 percent of the means that change the airflow are tank capacity. being used. (2) Each reserve oil tank not directly (i) The fuel jettisoning system must connected to any engine may have an be designed so that any reasonably expansion space of not less than two probable single malfunction in the sys- percent of the tank capacity. tem will not result in a hazardous con- (3) It must be impossible to fill the dition due to unsymmetrical jetti- expansion space inadvertently with the soning of, or inability to jettison, fuel. airplane in the normal ground attitude. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) Filler connection. Each recessed oil amended by Amdt. 25–18, 33 FR 12226, Aug. 30, tank filler connection that can retain 1968; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; any appreciable quantity of oil must Amdt. 25–108, 67 FR 70827, Nov. 26, 2002] have a drain that discharges clear of each part of the airplane. In addition, OIL SYSTEM each oil tank filler cap must provide an oil-tight seal. § 25.1011 General. (d) Vent. Oil tanks must be vented as (a) Each engine must have an inde- follows: pendent oil system that can supply it (1) Each oil tank must be vented with an appropriate quantity of oil at a from the top part of the expansion temperature not above that safe for space so that venting is effective under continuous operation. any normal flight condition.

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(2) Oil tank vents must be arranged (b) Breather lines must be arranged so that condensed water vapor that so that— might freeze and obstruct the line can- (1) Condensed water vapor that might not accumulate at any point. freeze and obstruct the line cannot ac- (e) Outlet. There must be means to cumulate at any point; prevent entrance into the tank itself, (2) The breather discharge does not or into the tank outlet, of any object constitute a fire hazard if foaming oc- that might obstruct the flow of oil curs or causes emitted oil to strike the through the system. No oil tank outlet pilot’s windshield; and may be enclosed by any screen or guard (3) The breather does not discharge that would reduce the flow of oil below into the engine air induction system. a safe value at any operating temperature. There must be a shutoff valve at § 25.1019 Oil strainer or filter. the outlet of each oil tank used with a turbine engine, unless the external por- (a) Each turbine engine installation tion of the oil system (including the oil must incorporate an oil strainer or fil- tank supports) is fireproof. ter through which all of the engine oil (f) Flexible oil tank liners. Each flexi- flows and which meets the following re- ble oil tank liner must be approved or quirements: must be shown to be suitable for the (1) Each oil strainer or filter that has particular application. a bypass must be constructed and installed so that oil will flow at the nor- [Doc. No. 5066, 29 FR 18291, Dec. 24, as amended by Amdt. 25–19, 33 FR 15410, Oct. 17, 1968; mal rate through the rest of the sys- Amdt. 25–23, 35 FR 5677, Apr. 8, 1970; Amdt. tem with the strainer or filter com- 25–36, 39 FR 35460, Oct. 1, 1974; Amdt. 25–57, 49 pletely blocked. FR 6848, Feb. 23, 1984; Amdt. 25–72, 55 FR (2) The oil strainer or filter must 29785, July 20, 1990] have the capacity (with respect to operating limitations established for the § 25.1015 Oil tank tests. engine) to ensure that engine oil sys- Each oil tank must be designed and tem functioning is not impaired when installed so that— the oil is contaminated to a degree (a) It can withstand, without failure, (with respect to particle size and den- each vibration, inertia, and fluid load sity) that is greater than that estab- that it may be subjected to in oper- lished for the engine under Part 33 of ation; and this chapter. (b) It meets the provisions of § 25.965, except— (3) The oil strainer or filter, unless it (1) The test pressure— is installed at an oil tank outlet, must (i) For pressurized tanks used with a incorporate an indicator that will indi- turbine engine, may not be less than 5 cate contamination before it reaches p.s.i. plus the maximum operating the capacity established in accordance pressure of the tank instead of the with paragraph (a)(2) of this section. pressure specified in § 25.965(a); and (4) The bypass of a strainer or filter (ii) For all other tanks may not be must be constructed and installed so less than 5 p.s.i. instead of the pressure that the release of collected contami- specified in § 25.965(a); and nants is minimized by appropriate lo- (2) The test fluid must be oil at 250 cation of the bypass to ensure that col- °F. instead of the fluid specified in lected contaminants are not in the by- § 25.965(c). pass flow path. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (5) An oil strainer or filter that has amended by Amdt. 25–36, 39 FR 35461, Oct. 1, no bypass, except one that is installed 1974] at an oil tank outlet, must have a means to connect it to the warning § 25.1017 Oil lines and fittings. system required in § 25.1305(c)(7). (a) Each oil line must meet the re- (b) Each oil strainer or filter in a quirements of § 25.993 and each oil line powerplant installation using recipro- and fitting in any designated fire zone cating engines must be constructed and must meet the requirements of installed so that oil will flow at the § 25.1183. normal rate through the rest of the

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system with the strainer or filter ele- source of oil for lubricating the engine ment completely blocked. during operation. (d) Provision must be made to pre- [Amdt. 25–36, 39 FR 35461, Oct. 1, 1974, as vent sludge or other foreign matter amended by Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] from affecting the safe operation of the propeller feathering system. § 25.1021 Oil system drains. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as A drain (or drains) must be provided amended by Amdt. 25–38, 41 FR 55467, Dec. 20, to allow safe drainage of the oil sys- 1976] tem. Each drain must— COOLING (a) Be accessible; and (b) Have manual or automatic means § 25.1041 General. for positive locking in the closed posi- The powerplant and auxiliary power tion. unit cooling provisions must be able to [Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] maintain the temperatures of powerplant components, engine fluids, and § 25.1023 Oil radiators. auxiliary power unit components and fluids within the temperature limits (a) Each oil radiator must be able to established for these components and withstand, without failure, any vibra- fluids, under ground, water, and flight tion, inertia, and oil pressure load to operating conditions, and after normal which it would be subjected in oper- engine or auxiliary power unit shut- ation. down, or both. (b) Each oil radiator air duct must be located so that, in case of fire, flames [Amdt. 25–38, 41 FR 55467, Dec. 20, 1976] coming from normal openings of the engine nacelle cannot impinge directly § 25.1043 Cooling tests. upon the radiator. (a) General. Compliance with § 25.1041 must be shown by tests, under critical § 25.1025 Oil valves. ground, water, and flight operating (a) Each oil shutoff must meet the re- conditions. For these tests, the fol- quirements of § 25.1189. lowing apply: (b) The closing of oil shutoff means (1) If the tests are conducted under may not prevent propeller feathering. conditions deviating from the maximum ambient atmospheric tempera- (c) Each oil valve must have positive ture, the recorded powerplant tempera- stops or suitable index provisions in tures must be corrected under para- the ‘‘on’’ and ‘‘off’’ positions and must graphs (c) and (d) of this section. be supported so that no loads resulting (2) No corrected temperatures deter- from its operation or from accelerated mined under paragraph (a)(1) of this flight conditions are transmitted to section may exceed established limits. the lines attached to the valve. (3) For reciprocating engines, the fuel used during the cooling tests must be § 25.1027 Propeller feathering system. the minimum grade approved for the (a) If the propeller feathering system engines, and the mixture settings must depends on engine oil, there must be be those normally used in the flight means to trap an amount of oil in the stages for which the cooling tests are tank if the supply becomes depleted conducted. The test procedures must be due to failure of any part of the lubri- as prescribed in § 25.1045. cating system other than the tank (b) Maximum ambient atmospheric tem- itself. perature. A maximum ambient atmos- (b) The amount of trapped oil must pheric temperature corresponding to be enough to accomplish the feathering sea level conditions of at least 100 de- operation and must be available only grees F must be established. The as- to the feathering pump. sumed temperature lapse rate is 3.6 de- (c) The ability of the system to ac- grees F per thousand feet of altitude complish feathering with the trapped above sea level until a temperature of oil must be shown. This may be done ¥69.7 degrees F is reached, above which on the ground using an auxiliary altitude the temperature is considered

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constant at ¥69.7 degrees F. However, the time of entry). The takeoff cooling for winterization installations, the ap- test must be preceded by a period dur- plicant may select a maximum ambi- ing which the powerplant component ent atmospheric temperature cor- and engine fluid temperatures are sta- responding to sea level conditions of bilized with the engines at ground idle. less than 100 degrees F. (c) Cooling tests for each stage of (c) Correction factor (except cylinder flight must be continued until— barrels). Unless a more rational correc- (1) The component and engine fluid tion applies, temperatures of engine temperatures stabilize; fluids and powerplant components (ex- (2) The stage of flight is completed; cept cylinder barrels) for which tem- or perature limits are established, must be corrected by adding to them the dif- (3) An operating limitation is ference between the maximum ambient reached. atmospheric temperature and the tem- (d) For reciprocating engine powered perature of the ambient air at the time airplanes, it may be assumed, for cool- of the first occurrence of the maximum ing test purposes, that the takeoff component or fluid temperature re- stage of flight is complete when the corded during the cooling test. airplane reaches an altitude of 1,500 (d) Correction factor for cylinder barrel feet above the takeoff surface or temperatures. Unless a more rational reaches a point in the takeoff where correction applies, cylinder barrel tem- the transition from the takeoff to the peratures must be corrected by adding en route configuration is completed to them 0.7 times the difference be- and a speed is reached at which compli- tween the maximum ambient atmos- ance with § 25.121(c) is shown, which- pheric temperature and the tempera- ever point is at a higher altitude. The ture of the ambient air at the time of airplane must be in the following con- the first occurrence of the maximum figuration: cylinder barrel temperature recorded (1) Landing gear retracted. during the cooling test. (2) Wing flaps in the most favorable [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as position. amended by Amdt. 25–42, 43 FR 2323, Jan. 16, (3) Cowl flaps (or other means of con- 1978] trolling the engine cooling supply) in the position that provides adequate § 25.1045 Cooling test procedures. cooling in the hot-day condition. (a) Compliance with § 25.1041 must be (4) Critical engine inoperative and its shown for the takeoff, climb, en route, propeller stopped. and landing stages of flight that cor- (5) Remaining engines at the max- respond to the applicable performance imum continuous power available for requirements. The cooling tests must the altitude. be conducted with the airplane in the (e) For hull seaplanes and amphib- configuration, and operating under the ians, cooling must be shown during conditions, that are critical relative to taxiing downwind for 10 minutes, at cooling during each stage of flight. For five knots above step speed. the cooling tests, a temperature is ‘‘stabilized’’ when its rate of change is [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as less than two degrees F. per minute. amended by Amdt. 25–57, 49 FR 6848, Feb. 23, (b) Temperatures must be stabilized 1984] under the conditions from which entry is made into each stage of flight being INDUCTION SYSTEM investigated, unless the entry condi- § 25.1091 Air induction. tion normally is not one during which component and the engine fluid tem- (a) The air induction system for each peratures would stabilize (in which engine and auxiliary power unit must case, operation through the full entry supply— condition must be conducted before (1) The air required by that engine entry into the stage of flight being in- and auxiliary power unit under each vestigated in order to allow tempera- operating condition for which certifi- tures to reach their natural levels at cation is requested; and

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(2) The air for proper fuel metering ture of 30 F., each airplane with alti- and mixture distribution with the in- tude engines using— duction system valves in any position. (1) Conventional venturi carburetors (b) Each reciprocating engine must have a preheater that can provide a have an alternate air source that pre- heat rise of 120 F. with the engine at 60 vents the entry of rain, ice, or any percent of maximum continuous power; other foreign matter. or (c) Air intakes may not open within (2) Carburetors tending to reduce the the cowling, unless— probability of ice formation has a pre- (1) That part of the cowling is iso- heater that can provide a heat rise of lated from the engine accessory section 100 °F. with the engine at 60 percent of by means of a fireproof diaphragm; or (2) For reciprocating engines, there maximum continuous power. are means to prevent the emergence of (b) Turbine engines. (1) Each turbine backfire flames. engine must operate throughout the (d) For turbine engine powered air- flight power range of the engine (in- planes and airplanes incorporating aux- cluding idling), without the accumula- iliary power units— tion of ice on the engine, inlet system (1) There must be means to prevent components, or airframe components hazardous quantities of fuel leakage or that would adversely affect engine op- overflow from drains, vents, or other eration or cause a serious loss of power components of flammable fluid systems or thrust— from entering the engine or auxiliary (i) Under the icing conditions speci- power unit intake system; and fied in appendix C, and (2) The airplane must be designed to (ii) In falling and blowing snow with- prevent water or slush on the runway, in the limitations established for the taxiway, or other airport operating airplane for such operation. surfaces from being directed into the (2) Each turbine engine must idle for engine or auxiliary power unit air inlet 30 minutes on the ground, with the air ducts in hazardous quantities, and the bleed available for engine icing protec- air inlet ducts must be located or pro- tion at its critical condition, without tected so as to minimize the ingestion of foreign matter during takeoff, land- adverse effect, in an atmosphere that is ° ° ing, and taxiing. at a temperature between 15 and 30 F ¥ ° ¥ ° (e) If the engine induction system (between 9 and 1 C) and has a liq- contains parts or components that uid water content not less than 0.3 could be damaged by foreign objects grams per cubic meter in the form of entering the air inlet, it must be shown drops having a mean effective diameter by tests or, if appropriate, by analysis not less than 20 microns, followed by that the induction system design can momentary operation at takeoff power withstand the foreign object ingestion or thrust. During the 30 minutes of idle test conditions of §§ 33.76, 33.77 and operation, the engine may be run up 33.78(a)(1) of this chapter without fail- periodically to a moderate power or ure of parts or components that could thrust setting in a manner acceptable create a hazard. to the Administrator. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) Supercharged reciprocating engines. amended by Amdt. 25–38, 41 FR 55467, Dec. 20, For each engine having a supercharger 1976; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977; to pressurize the air before it enters Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. the carburetor, the heat rise in the air 25–100, 65 FR 55854, Sept. 14, 2000] caused by that supercharging at any § 25.1093 Induction system icing pro- altitude may be utilized in determining tection. compliance with paragraph (a) of this section if the heat rise utilized is that (a) Reciprocating engines. Each reciprocating engine air induction system which will be available, automatically, must have means to prevent and eliminate icing. Unless this is done by other means, it must be shown that, in air free of visible moisture at a tempera-

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for the applicable altitude and oper- and entering any other compartment ating condition because of super- or area of the airplane in which a haz- charging. ard would be created resulting from the entry of hot gases. The materials used [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55467, Dec. 20, to form the remainder of the induction 1976; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977; system duct and plenum chamber of Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. the auxiliary power unit must be capa- 25–72, 55 FR 29785, July 20, 1990] ble of resisting the maximum heat conditions likely to occur. § 25.1101 Carburetor air preheater de- (f) Each auxiliary power unit induc- sign. tion system duct must be constructed Each carburetor air preheater must of materials that will not absorb or be designed and constructed to— trap hazardous quantities of flammable (a) Ensure ventilation of the pre- fluids that could be ignited in the heater when the engine is operated in event of a surge or reverse flow condi- cold air; tion. (b) Allow inspection of the exhaust [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as manifold parts that it surrounds; and amended by Amdt. 25–46, 43 FR 50597, Oct. 30, (c) Allow inspection of critical parts 1978] of the preheater itself. § 25.1105 Induction system screens. § 25.1103 Induction system ducts and If induction system screens are air duct systems. used— (a) Each induction system duct up- (a) Each screen must be upstream of stream of the first stage of the engine the carburetor; supercharger and of the auxiliary (b) No screen may be in any part of power unit compressor must have a the induction system that is the only drain to prevent the hazardous accu- passage through which air can reach mulation of fuel and moisture in the the engine, unless it can be deiced by ground attitude. No drain may dis- heated air; charge where it might cause a fire haz- (c) No screen may be deiced by alco- ard. hol alone; and (b) Each induction system duct must (d) It must be impossible for fuel to be— strike any screen. (1) Strong enough to prevent induction system failures resulting from § 25.1107 Inter-coolers and after-cool- normal backfire conditions; and ers. (2) Fire-resistant if it is in any fire Each inter-cooler and after-cooler zone for which a fire-extinguishing sys- must be able to withstand any vibra- tem is required, except that ducts for tion, inertia, and air pressure load to auxiliary power units must be fireproof which it would be subjected in oper- within the auxiliary power unit fire ation. zone. (c) Each duct connected to compo- EXHAUST SYSTEM nents between which relative motion could exist must have means for flexi- § 25.1121 General. bility. For powerplant and auxiliary power (d) For turbine engine and auxiliary unit installations the following apply: power unit bleed air duct systems, no (a) Each exhaust system must ensure hazard may result if a duct failure oc- safe disposal of exhaust gases without curs at any point between the air duct fire hazard or carbon monoxide con- source and the airplane unit served by tamination in any personnel compart- the air. ment. For test purposes, any accept- (e) Each auxiliary power unit induc- able carbon monoxide detection meth- tion system duct must be fireproof for od may be used to show the absence of a sufficient distance upstream of the carbon monoxide. auxiliary power unit compartment to (b) Each exhaust system part with a prevent hot gas reverse flow from burn- surface hot enough to ignite flammable ing through auxiliary power unit ducts fluids or vapors must be located or

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shielded so that leakage from any sys- other load to which it would be sub- tem carrying flammable fluids or va- jected in operation. In addition— pors will not result in a fire caused by (1) Each exchanger must be suitable impingement of the fluids or vapors on for continued operation at high tem- any part of the exhaust system includ- peratures and resistant to corrosion ing shields for the exhaust system. from exhaust gases; (c) Each component that hot exhaust (2) There must be means for the in- gases could strike, or that could be spection of the critical parts of each subjected to high temperatures from exchanger; exhaust system parts, must be fire- (3) Each exchanger must have cooling proof. All exhaust system components provisions wherever it is subject to must be separated by fireproof shields contact with exhaust gases; and from adjacent parts of the airplane (4) No exhaust heat exchanger or that are outside the engine and auxil- muff may have any stagnant areas or iary power unit compartments. liquid traps that would increase the (d) No exhaust gases may discharge probability of ignition of flammable so as to cause a fire hazard with respect to any flammable fluid vent or fluids or vapors that might be present drain. in case of the failure or malfunction of (e) No exhaust gases may discharge components carrying flammable fluids. where they will cause a glare seriously (b) If an exhaust heat exchanger is affecting pilot vision at night. used for heating ventilating air— (f) Each exhaust system component (1) There must be a secondary heat must be ventilated to prevent points of exchanger between the primary ex- excessively high temperature. haust gas heat exchanger and the ven- (g) Each exhaust shroud must be ventilating air system; or tilated or insulated to avoid, during (2) Other means must be used to pre- normal operation, a temperature high clude the harmful contamination of the enough to ignite any flammable fluids ventilating air. or vapors external to the shroud. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55467, Dec. 20, amended by Amdt. 25–40, 42 FR 15043, Mar. 17, 1976] 1977] § 25.1127 Exhaust driven turbo-super- § 25.1123 Exhaust piping. chargers. For powerplant and auxiliary power (a) Each exhaust driven turbo-super- unit installations, the following apply: charger must be approved or shown to (a) Exhaust piping must be heat and be suitable for the particular applica- corrosion resistant, and must have pro- tion. It must be installed and sup- visions to prevent failure due to expan- ported to ensure safe operation be- sion by operating temperatures. tween normal inspections and over- (b) Piping must be supported to with- hauls. In addition, there must be provi- stand any vibration and inertia loads sions for expansion and flexibility be- to which it would be subjected in oper- tween exhaust conduits and the tur- ation; and bine. (c) Piping connected to components (b) There must be provisions for lu- between which relative motion could bricating the turbine and for cooling exist must have means for flexibility. turbine parts where temperatures are [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as critical. amended by Amdt. 25–40, 42 FR 15044, Mar. 17, (c) If the normal turbo-supercharger 1977] control system malfunctions, the turbine speed may not exceed its max- § 25.1125 Exhaust heat exchangers. imum allowable value. Except for the For reciprocating engine powered waste gate operating components, the airplanes, the following apply: components provided for meeting this (a) Each exhaust heat exchanger requirement must be independent of must be constructed and installed to the normal turbo-supercharger con- withstand each vibration, inertia, and trols.

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POWERPLANT CONTROLS AND (b) Power and thrust controls must ACCESSORIES be arranged to allow— (1) Separate control of each engine; § 25.1141 Powerplant controls: general. and Each powerplant control must be lo- (2) Simultaneous control of all en- cated, arranged, and designed under gines. §§ 25.777 through 25.781 and marked (c) Each power and thrust control under § 25.1555. In addition, it must must provide a positive and imme- meet the following requirements: diately responsive means of controlling (a) Each control must be located so its engine. that it cannot be inadvertently oper- (d) For each fluid injection (other ated by persons entering, leaving, or than fuel) system and its controls not moving normally in, the cockpit. provided and approved as part of the (b) Each flexible control must be ap- engine, the applicant must show that proved or must be shown to be suitable the flow of the injection fluid is ade- for the particular application. quately controlled. (c) Each control must have sufficient (e) If a power or thrust control incor- strength and rigidity to withstand op- porates a fuel shutoff feature, the con- erating loads without failure and with- trol must have a means to prevent the out excessive deflection. inadvertent movement of the control (d) Each control must be able to into the shutoff position. The means maintain any set position without con- must— stant attention by flight crewmembers (1) Have a positive lock or stop at the and without creep due to control loads idle position; and or vibration. (2) Require a separate and distinct (e) The portion of each powerplant operation to place the control in the control located in a designated fire shutoff position. zone that is required to be operated in [Amdt. 25–23, 35 FR 5677, Apr. 8, 1970, as the event of fire must be at least fire amended by Amdt. 25–38, 41 FR 55467, Dec. 20, resistant. 1976; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984] (f) For powerplant valve controls located in the flight deck there must be § 25.1145 Ignition switches. a means: (a) Ignition switches must control (1) For the flightcrew to select each each engine ignition circuit on each intended position or function of the engine. valve; and (b) There must be means to quickly (2) To indicate to the flightcrew: shut off all ignition by the grouping of (i) The selected position or function switches or by a master ignition con- of the valve; and trol. (ii) When the valve has not responded (c) Each group of ignition switches, as intended to the selected position or except ignition switches for turbine en- function. gines for which continuous ignition is [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as not required, and each master ignition amended by Amdt. 25–40, 42 FR 15044, Mar. 17, control must have a means to prevent 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990; its inadvertent operation. Amdt. 25–115, 69 FR 40527, July 2, 2004] [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.1142 Auxiliary power unit con- amended by Amdt. 25–40, 42 FR 15044 Mar. 17, trols. 1977] Means must be provided on the flight § 25.1147 Mixture controls. deck for starting, stopping, and emergency shutdown of each installed auxil- (a) If there are mixture controls, iary power unit. each engine must have a separate control. The controls must be grouped and [Amdt. 25–46, 43 FR 50598, Oct. 30, 1978] arranged to allow— (1) Separate control of each engine; § 25.1143 Engine controls. and (a) There must be a separate power or (2) Simultaneous control of all en- thrust control for each engine. gines.

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(b) Each intermediate position of the thrust regime for turbojet powered air- mixture controls that corresponds to a planes). normal operating setting must be iden- [Amdt. 25–11, 32 FR 6913, May 5, 1967] tifiable by feel and sight. (c) The mixture controls must be ac- § 25.1157 Carburetor air temperature cessible to both pilots. However, if controls. there is a separate flight engineer sta- There must be a separate carburetor tion with a control panel, the controls air temperature control for each en- need be accessible only to the flight engine. gineer. § 25.1159 Supercharger controls. § 25.1149 Propeller speed and pitch controls. Each supercharger control must be accessible to the pilots or, if there is a (a) There must be a separate pro- separate flight engineer station with a peller speed and pitch control for each control panel, to the flight engineer. propeller. (b) The controls must be grouped and § 25.1161 Fuel jettisoning system con- arranged to allow— trols. (1) Separate control of each pro- Each fuel jettisoning system control peller; and must have guards to prevent inad- (2) Simultaneous control of all pro- vertent operation. No control may be pellers. near any fire extinguisher control or (c) The controls must allow synchro- other control used to combat fire. nization of all propellers. (d) The propeller speed and pitch con- § 25.1163 Powerplant accessories. trols must be to the right of, and at (a) Each engine mounted accessory least one inch below, the pilot’s throt- must— tle controls. (1) Be approved for mounting on the engine involved; § 25.1153 Propeller feathering controls. (2) Use the provisions on the engine (a) There must be a separate pro- for mounting; and peller feathering control for each pro- (3) Be sealed to prevent contamina- peller. The control must have means to tion of the engine oil system and the prevent its inadvertent operation. accessory system. (b) If feathering is accomplished by (b) Electrical equipment subject to movement of the propeller pitch or arcing or sparking must be installed to speed control lever, there must be minimize the probability of contact means to prevent the inadvertent with any flammable fluids or vapors movement of this lever to the feath- that might be present in a free state. ering position during normal oper- (c) If continued rotation of an engine- ation. driven cabin supercharger or of any remote accessory driven by the engine is [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as hazardous if malfunctioning occurs, amended by Amdt. 25–11, 32 FR 6913, May 5, 1967] there must be means to prevent rotation without interfering with the con- § 25.1155 Reverse thrust and propeller tinued operation of the engine. pitch settings below the flight re- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as gime. amended by Amdt. 25–57, 49 FR 6849, Feb. 23, Each control for reverse thrust and 1984] for propeller pitch settings below the flight regime must have means to pre- § 25.1165 Engine ignition systems. vent its inadvertent operation. The (a) Each battery ignition system means must have a positive lock or must be supplemented by a generator stop at the flight idle position and that is automatically available as an must require a separate and distinct alternate source of electrical energy to operation by the crew to displace the allow continued engine operation if control from the flight regime (forward any battery becomes depleted.

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(b) The capacity of batteries and gen- (c) Possible misalignments and tor- erators must be large enough to meet sional loadings of the gearbox, trans- the simultaneous demands of the en- mission, and shaft system, expected to gine ignition system and the greatest result under normal operating condi- demands of any electrical system com- tions must be evaluated. ponents that draw electrical energy from the same source. [Amdt. 25–38, 41 FR 55467, Dec. 20, 1976] (c) The design of the engine ignition POWERPLANT FIRE PROTECTION system must account for— (1) The condition of an inoperative § 25.1181 Designated fire zones; re- generator; gions included. (2) The condition of a completely depleted battery with the generator run- (a) Designated fire zones are— ning at its normal operating speed; and (1) The engine power section; (3) The condition of a completely de- (2) The engine accessory section; pleted battery with the generator oper- (3) Except for reciprocating engines, ating at idling speed, if there is only any complete powerplant compartment one battery. in which no isolation is provided be- (d) Magneto ground wiring (for sepa- tween the engine power section and the rate ignition circuits) that lies on the engine accessory section; engine side of the fire wall, must be in- (4) Any auxiliary power unit com- stalled, located, or protected, to mini- partment; mize the probability of simultaneous (5) Any fuel-burning heater and other failure of two or more wires as a result combustion equipment installation de- of mechanical damage, electrical scribed in § 25.859; faults, or other cause. (6) The compressor and accessory sec- (e) No ground wire for any engine tions of turbine engines; and may be routed through a fire zone of (7) Combustor, turbine, and tailpipe another engine unless each part of that sections of turbine engine installations wire within that zone is fireproof. that contain lines or components car- (f) Each ignition system must be rying flammable fluids or gases. independent of any electrical circuit, (b) Each designated fire zone must not used for assisting, controlling, or meet the requirements of §§ 25.863, analyzing the operation of that system. 25.865, 25.867, 25.869, and 25.1185 through (g) There must be means to warn ap- 25.1203. propriate flight crewmembers if the malfunctioning of any part of the elec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as trical system is causing the continuous amended by Amdt. 25–11, 32 FR 6913, May 5, 1967; Amdt. 25–23, 35 FR 5677, Apr. 8, 1970; discharge of any battery necessary for Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. engine ignition. 25–115, 69 FR 40527, July 2, 2004] (h) Each engine ignition system of a turbine powered airplane must be con- § 25.1182 Nacelle areas behind fire- sidered an essential electrical load. walls, and engine pod attaching structures containing flammable [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as fluid lines. amended by Amdt. 25–23, 35 FR 5677, Apr. 8, 1970; Amdt. 25–72, 55 FR 29785, July 20, 1990] (a) Each nacelle area immediately behind the firewall, and each portion of § 25.1167 Accessory gearboxes. any engine pod attaching structure For airplanes equipped with an acces- containing flammable fluid lines, must sory gearbox that is not certificated as meet each requirement of §§ 25.1103(b), part of an engine— 25.1165 (d) and (e), 25.1183, 25.1185(c), (a) The engine with gearbox and con- 25.1187, 25.1189, and 25.1195 through necting transmissions and shafts at- 25.1203, including those concerning des- tached must be subjected to the tests ignated fire zones. However, engine pod specified in § 33.49 or § 33.87 of this chap- attaching structures need not contain ter, as applicable; fire detection or extinguishing means. (b) The accessory gearbox must meet (b) For each area covered by para- the requirements of §§ 33.25 and 33.53 or graph (a) of this section that contains 33.91 of this chapter, as applicable; and a retractable landing gear, compliance

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with that paragraph need only be tions, lines, and control provide a de- shown with the landing gear retracted. gree of safety equal to that which would exist if the tank or reservoir [Amdt. 25–11, 32 FR 6913, May 5, 1967] were outside such a zone. § 25.1183 Flammable fluid-carrying (b) There must be at least one-half components. inch of clear airspace between each (a) Except as provided in paragraph tank or reservoir and each firewall or (b) of this section, each line, fitting, shroud isolating a designated fire zone. and other component carrying flam- (c) Absorbent materials close to mable fluid in any area subject to en- flammable fluid system components gine fire conditions, and each compo- that might leak must be covered or nent which conveys or contains flam- treated to prevent the absorption of mable fluid in a designated fire zone hazardous quantities of fluids. must be fire resistant, except that [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964 as flammable fluid tanks and supports in amended by Amdt. 25–19, 33 FR 15410, Oct. 17, a designated fire zone must be fireproof 1968; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] or be enclosed by a fireproof shield unless damage by fire to any non-fire- § 25.1187 Drainage and ventilation of proof part will not cause leakage or fire zones. spillage of flammable fluid. Compo- (a) There must be complete drainage nents must be shielded or located to of each part of each designated fire safeguard against the ignition of leak- zone to minimize the hazards resulting ing flammable fluid. An integral oil from failure or malfunctioning of any sump of less than 25-quart capacity on component containing flammable a reciprocating engine need not be fire- fluids. The drainage means must be— proof nor be enclosed by a fireproof (1) Effective under conditions ex- shield. pected to prevail when drainage is (b) Paragraph (a) of this section does needed; and not apply to— (2) Arranged so that no discharged (1) Lines, fittings, and components fluid will cause an additional fire haz- which are already approved as part of a type certificated engine; and ard. (2) Vent and drain lines, and their fit- (b) Each designated fire zone must be tings, whose failure will not result in, ventilated to prevent the accumulation or add to, a fire hazard. of flammable vapors. (c) All components, including ducts, (c) No ventilation opening may be within a designated fire zone must be where it would allow the entry of flam- fireproof if, when exposed to or dam- mable fluids, vapors, or flame from aged by fire, they could— other zones. (1) Result in fire spreading to other (d) Each ventilation means must be regions of the airplane; or arranged so that no discharged vapors (2) Cause unintentional operation of, will cause an additional fire hazard. or inability to operate, essential serv- (e) Unless the extinguishing agent ca- ices or equipment. pacity and rate of discharge are based on maximum air flow through a zone, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as there must be means to allow the crew amended by Amdt. 25–11, 32 FR 6913, May 5, 1967; Amdt. 25–36, 39 FR 35461, Oct. 1, 1974; to shut off sources of forced ventilation Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. to any fire zone except the engine 25–101, 65 FR 79710, Dec. 19, 2000] power section of the nacelle and the combustion heater ventilating air § 25.1185 Flammable fluids. ducts. (a) Except for the integral oil sumps specified in § 25.1183(a), no tank or res- § 25.1189 Shutoff means. ervoir that is a part of a system con- (a) Each engine installation and each taining flammable fluids or gases may fire zone specified in § 25.1181(a)(4) and be in a designated fire zone unless the (5) must have a means to shut off or fluid contained, the design of the sys- otherwise prevent hazardous quantities tem, the materials used in the tank, of fuel, oil, deicer, and other flammable the shut-off means, and all connec- fluids, from flowing into, within, or

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through any designated fire zone, ex- (2) Constructed so that no hazardous cept that shutoff means are not re- quantity of air, fluid, or flame can pass quired for— from the compartment to other parts (1) Lines, fittings, and components of the airplane; forming an integral part of an engine; (3) Constructed so that each opening and is sealed with close fitting fireproof (2) Oil systems for turbine engine in- grommets, bushings, or firewall fit- stallations in which all components of tings; and the system in a designated fire zone, (4) Protected against corrosion. including oil tanks, are fireproof or located in areas not subject to engine § 25.1192 Engine accessory section dia- fire conditions. phragm. (b) The closing of any fuel shutoff For reciprocating engines, the engine valve for any engine may not make power section and all portions of the fuel unavailable to the remaining en- exhaust system must be isolated from gines. the engine accessory compartment by a (c) Operation of any shutoff may not diaphragm that complies with the fire- interfere with the later emergency op- wall requirements of § 25.1191. eration of other equipment, such as the [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970] means for feathering the propeller. (d) Each flammable fluid shutoff § 25.1193 Cowling and nacelle skin. means and control must be fireproof or (a) Each cowling must be constructed must be located and protected so that and supported so that it can resist any any fire in a fire zone will not affect its vibration, inertia, and air load to operation. which it may be subjected in operation. (e) No hazardous quantity of flam- (b) Cowling must meet the drainage mable fluid may drain into any des- and ventilation requirements of ignated fire zone after shutoff. § 25.1187. (f) There must be means to guard (c) On airplanes with a diaphragm against inadvertent operation of the isolating the engine power section from shutoff means and to make it possible the engine accessory section, each part for the crew to reopen the shutoff of the accessory section cowling sub- means in flight after it has been closed. ject to flame in case of fire in the en- (g) Each tank-to-engine shutoff valve gine power section of the powerplant must be located so that the operation must— of the valve will not be affected by (1) Be fireproof; and powerplant or engine mount structural (2) Meet the requirements of § 25.1191. failure. (d) Each part of the cowling subject (h) Each shutoff valve must have a to high temperatures due to its near- means to relieve excessive pressure ac- ness to exhaust system parts or ex- cumulation unless a means for pressure haust gas impingement must be fire- relief is otherwise provided in the sys- proof. tem. (e) Each airplane must— [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (1) Be designed and constructed so amended by Amdt. 25–23, 35 FR 5677, Apr. 8, that no fire originating in any fire zone 1970; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984] can enter, either through openings or by burning through external skin, any § 25.1191 Firewalls. other zone or region where it would (a) Each engine, auxiliary power create additional hazards; unit, fuel-burning heater, other com- (2) Meet paragraph (e)(1) of this sec- bustion equipment intended for oper- tion with the landing gear retracted (if ation in flight, and the combustion, applicable); and turbine, and tailpipe sections of tur- (3) Have fireproof skin in areas sub- bine engines, must be isolated from the ject to flame if a fire starts in the en- rest of the airplane by firewalls, gine power or accessory sections. shrouds, or equivalent means. (b) Each firewall and shroud must § 25.1195 Fire extinguishing systems. be— (a) Except for combustor, turbine, (1) Fireproof; and tail pipe sections of turbine engine

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installations that contain lines or com- bon dioxide fuselage compartment fire ponents carrying flammable fluids or extinguishing systems for which— gases for which it is shown that a fire (1) Five pounds or less of carbon diox- originating in these sections can be ide will be discharged, under estab- controlled, there must be a fire extin- lished fire control procedures, into any guisher system serving each designated fuselage compartment; or fire zone. (2) There is protective breathing (b) The fire extinguishing system, the equipment for each flight crewmember quantity of the extinguishing agent, on flight deck duty. the rate of discharge, and the discharge [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as distribution must be adequate to extin- amended by Amdt. 25–38, 41 FR 55467, Dec. 20, guish fires. It must be shown by either 1976; Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] actual or simulated flights tests that under critical airflow conditions in § 25.1199 Extinguishing agent con- flight the discharge of the extin- tainers. guishing agent in each designated fire (a) Each extinguishing agent con- zone specified in paragraph (a) of this tainer must have a pressure relief to section will provide an agent con- prevent bursting of the container by centration capable of extinguishing excessive internal pressures. fires in that zone and of minimizing (b) The discharge end of each dis- the probability of reignition. An indi- charge line from a pressure relief con- vidual ‘‘one-shot’’ system may be used nection must be located so that dis- for auxiliary power units, fuel burning charge of the fire extinguishing agent heaters, and other combustion equip- would not damage the airplane. The ment. For each other designated fire line must also be located or protected zone, two discharges must be provided to prevent clogging caused by ice or each of which produces adequate agent other foreign matter. concentration. (c) There must be a means for each (c) The fire extinguishing system for fire extinguishing agent container to a nacelle must be able to simulta- indicate that the container has dis- neously protect each zone of the na- charged or that the charging pressure celle for which protection is provided. is below the established minimum nec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as essary for proper functioning. amended by Amdt. 25–46, 43 FR 50598, Oct. 30, (d) The temperature of each con- 1978] tainer must be maintained, under intended operating conditions, to prevent § 25.1197 Fire extinguishing agents. the pressure in the container from— (a) Fire extinguishing agents must— (1) Falling below that necessary to (1) Be capable of extinguishing provide an adequate rate of discharge; flames emanating from any burning of or fluids or other combustible materials (2) Rising high enough to cause pre- in the area protected by the fire extin- mature discharge. guishing system; and (e) If a pyrotechnic capsule is used to (2) Have thermal stability over the discharge the extinguishing agent, temperature range likely to be experi- each container must be installed so enced in the compartment in which that temperature conditions will not they are stored. cause hazardous deterioration of the (b) If any toxic extinguishing agent is pyrotechnic capsule. used, provisions must be made to pre- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as vent harmful concentrations of fluid or amended by Amdt. 25–23, 35 FR 5678, Apr. 8, fluid vapors (from leakage during nor- 1970; Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] mal operation of the airplane or as a result of discharging the fire extin- § 25.1201 Fire extinguishing system guisher on the ground or in flight) from materials. entering any personnel compartment, (a) No material in any fire extin- even though a defect may exist in the guishing system may react chemically extinguishing system. This must be with any extinguishing agent so as to shown by test except for built-in car- create a hazard.

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(b) Each system component in an en- (h) EWIS for each fire or overheat de- gine compartment must be fireproof. tector system in a fire zone must meet the requirements of § 25.1731. § 25.1203 Fire detector system. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) There must be approved, quick amended by Amdt. 25–23, 35 FR 5678, Apr. 8, acting fire or overheat detectors in 1970; Amdt. 25–26, 36 FR 5493, Mar. 24, 1971; each designated fire zone, and in the Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] combustion, turbine, and tailpipe sections of turbine engine installations, in § 25.1207 Compliance. numbers and locations ensuring Unless otherwise specified, compli- prompt detection of fire in those zones. ance with the requirements of §§ 25.1181 (b) Each fire detector system must be through 25.1203 must be shown by a full constructed and installed so that— scale fire test or by one or more of the (1) It will withstand the vibration, in- following methods: ertia, and other loads to which it may (a) Tests of similar powerplant con- be subjected in operation; figurations; (2) There is a means to warn the crew (b) Tests of components; in the event that the sensor or associ- (c) Service experience of aircraft ated wiring within a designated fire with similar powerplant configura- zone is severed at one point, unless the tions; system continues to function as a sat- (d) Analysis. isfactory detection system after the severing; and [Amdt. 25–46, 43 FR 50598, Oct. 30, 1978] (3) There is a means to warn the crew in the event of a short circuit in the Subpart F—Equipment sensor or associated wiring within a designated fire zone, unless the system GENERAL continues to function as a satisfactory § 25.1301 Function and installation. detection system after the short circuit. (a) Each item of installed equipment (c) No fire or overheat detector may must— be affected by any oil, water, other (1) Be of a kind and design appro- fluids or fumes that might be present. priate to its intended function; (d) There must be means to allow the (2) Be labeled as to its identification, crew to check, in flight, the func- function, or operating limitations, or tioning of each fire or overheat detec- any applicable combination of these tor electric circuit. factors; (e) Components of each fire or over- (3) Be installed according to limita- heat detector system in a fire zone tions specified for that equipment; and must be fire-resistant. (4) Function properly when installed. (f) No fire or overheat detector sys- (b) EWIS must meet the require- tem component for any fire zone may ments of subpart H of this part. pass through another fire zone, un- [Dockt. No. 5066, Amdt. 1–6, 29 FR 18333, Dec. less— 24, 1964, as amended by Amdt. 25–123, 72 FR (1) It is protected against the possi- 63405, Nov. 8, 2007] bility of false warnings resulting from fires in zones through which it passes; § 25.1303 Flight and navigation instru- or ments. (2) Each zone involved is simulta- (a) The following flight and naviga- neously protected by the same detector tion instruments must be installed so and extinguishing system. that the instrument is visible from (g) Each fire detector system must be each pilot station: constructed so that when it is in the (1) A free air temperature indicator configuration for installation it will or an air-temperature indicator which not exceed the alarm activation time provides indications that are convert- approved for the detectors using the re- ible to free-air temperature. sponse time criteria specified in the ap- (2) A clock displaying hours, min- propriate Technical Standard Order for utes, and seconds with a sweep-second the detector. pointer or digital presentation.

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(3) A direction indicator (non- (a) For all airplanes. (1) A fuel pres- stabilized magnetic compass). sure warning means for each engine, or (b) The following flight and naviga- a master warning means for all engines tion instruments must be installed at with provision for isolating the indi- each pilot station: vidual warning means from the master (1) An airspeed indicator. If airspeed warning means. limitations vary with altitude, the in- (2) A fuel quantity indicator for each dicator must have a maximum allow- fuel tank. able airspeed indicator showing the (3) An oil quantity indicator for each variation of VMO with altitude. oil tank. (2) An altimeter (sensitive). (4) An oil pressure indicator for each (3) A rate-of-climb indicator (vertical independent pressure oil system of speed). each engine. (4) A gyroscopic rate-of-turn indi- (5) An oil pressure warning means for cator combined with an integral slip- each engine, or a master warning skid indicator (turn-and-bank indi- means for all engines with provision cator) except that only a slip-skid indi- for isolating the individual warning cator is required on large airplanes means from the master warning means. with a third attitude instrument sys- (6) An oil temperature indicator for tem useable through flight attitudes of each engine. 360° of pitch and roll and installed in (7) Fire-warning devices that provide accordance with § 121.305(k) of this visual and audible warning. title. (8) An augmentation liquid quantity (5) A bank and pitch indicator (gyro- indicator (appropriate for the manner scopically stabilized). in which the liquid is to be used in op- (6) A direction indicator (gyroscop- eration) for each tank. ically stabilized, magnetic or non- (b) For reciprocating engine-powered magnetic). airplanes. In addition to the powerplant instruments required by paragraph (a) (c) The following flight and naviga- of this section, the following power- tion instruments are required as pre- plant instruments are required: scribed in this paragraph: (1) A carburetor air temperature indi- (1) A speed warning device is required cator for each engine. for turbine engine powered airplanes (2) A cylinder head temperature indi- and for airplanes with V /M greater MO MO cator for each air-cooled engine. than 0.8 V /M or 0.8 V /M The DF DF D D. (3) A manifold pressure indicator for speed warning device must give effec- each engine. tive aural warning (differing distinc- (4) A fuel pressure indicator (to indi- tively from aural warnings used for cate the pressure at which the fuel is other purposes) to the pilots, whenever supplied) for each engine. the speed exceeds V plus 6 knots or MO (5) A fuel flowmeter, or fuel mixture M +0.01. The upper limit of the pro- MO indicator, for each engine without an duction tolerance for the warning de- automatic altitude mixture control. vice may not exceed the prescribed (6) A tachometer for each engine. warning speed. (7) A device that indicates, to the (2) A machmeter is required at each flight crew (during flight), any change pilot station for airplanes with com- in the power output, for each engine pressibility limitations not otherwise with— indicated to the pilot by the airspeed (i) An automatic propeller feathering indicating system required under para- system, whose operation is initiated by graph (b)(1) of this section. a power output measuring system; or [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as (ii) A total engine piston displace- amended by Amdt. 25–24, 35 FR 7108, May 6, ment of 2,000 cubic inches or more. 1970; Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; (8) A means to indicate to the pilot Amdt. 25–90, 62 FR 13253, Mar. 19, 1997] when the propeller is in reverse pitch, for each reversing propeller. § 25.1305 Powerplant instruments. (c) For turbine engine-powered air- The following are required power- planes. In addition to the powerplant plant instruments: instruments required by paragraph (a)

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of this section, the following power- (e) For turbopropeller-powered air- plant instruments are required: planes. In addition to the powerplant (1) A gas temperature indicator for instruments required by paragraphs (a) each engine. and (c) of this section, the following (2) A fuel flowmeter indicator for powerplant instruments are required: each engine. (1) A torque indicator for each en- (3) A tachometer (to indicate the gine. speed of the rotors with established (2) Position indicating means to indi- limiting speeds) for each engine. cate to the flight crew when the pro- (4) A means to indicate, to the flight peller blade angle is below the flight crew, the operation of each engine low pitch position, for each propeller. starter that can be operated continu- (f) For airplanes equipped with fluid ously but that is neither designed for systems (other than fuel) for thrust or continuous operation nor designed to power augmentation, an approved prevent hazard if it failed. means must be provided to indicate the (5) An indicator to indicate the func- proper functioning of that system to tioning of the powerplant ice protec- the flight crew. tion system for each engine. (6) An indicator for the fuel strainer [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as or filter required by § 25.997 to indicate amended by Amdt. 25–35, 39 FR 1831, Jan. 15, the occurrence of contamination of the 1974; Amdt. 25–36, 39 FR 35461, Oct. 1, 1974; strainer or filter before it reaches the Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. capacity established in accordance 25–54, 45 FR 60173, Sept. 11, 1980; Amdt. 25–72, with § 25.997(d). 55 FR 29785, July 20, 1990; Amdt. 25–115, 69 FR 40527, July 2, 2004] (7) A warning means for the oil strainer or filter required by § 25.1019, if § 25.1307 Miscellaneous equipment. it has no bypass, to warn the pilot of the occurrence of contamination of the The following is required miscella- strainer or filter screen before it neous equipment: reaches the capacity established in ac- (a) [Reserved] cordance with § 25.1019(a)(2). (b) Two or more independent sources (8) An indicator to indicate the prop- of electrical energy. er functioning of any heater used to (c) Electrical protective devices, as prevent ice clogging of fuel system prescribed in this part. components. (d) Two systems for two-way radio (d) For turbojet engine powered air- communications, with controls for planes. In addition to the powerplant each accessible from each pilot station, instruments required by paragraphs (a) designed and installed so that failure of and (c) of this section, the following one system will not preclude operation powerplant instruments are required: of the other system. The use of a com- (1) An indicator to indicate thrust, or mon antenna system is acceptable if a parameter that is directly related to adequate reliability is shown. thrust, to the pilot. The indication (e) Two systems for radio navigation, must be based on the direct measure- with controls for each accessible from ment of thrust or of parameters that each pilot station, designed and in- are directly related to thrust. The indi- stalled so that failure of one system cator must indicate a change in thrust resulting from any engine malfunction, will not preclude operation of the other damage, or deterioration. system. The use of a common antenna (2) A position indicating means to in- system is acceptable if adequate reli- dicate to the flightcrew when the ability is shown. thrust reversing device— [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as (i) Is not in the selected position, and amended by Amdt. 25–46, 43 FR 50598, Oct. 30, (ii) Is in the reverse thrust position, 1978; Amdt. 25–54, 45 FR 60173, Sept. 11, 1980; for each engine using a thrust revers- Amdt. 25–72, 55 FR 29785, July 20, 1990] ing device. (3) An indicator to indicate rotor system unbalance.

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§ 25.1309 Equipment, systems, and in- procedures, the ability to provide con- stallations. tinuous, safe service under foreseeable (a) The equipment, systems, and in- environmental conditions may be stallations whose functioning is re- shown by environmental tests, design quired by this subchapter, must be de- analysis, or reference to previous com- signed to ensure that they perform parable service experience on other air- their intended functions under any craft. foreseeable operating condition. (f) EWIS must be assessed in accord- (b) The airplane systems and associ- ance with the requirements of § 25.1709. ated components, considered separately and in relation to other systems, [Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as must be designed so that— amended by Amdt. 25–38, 41 FR 55467, Dec. 20, (1) The occurrence of any failure con- 1976; Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] dition which would prevent the continued safe flight and landing of the air- § 25.1310 Power source capacity and plane is extremely improbable, and distribution. (2) The occurrence of any other failure conditions which would reduce the (a) Each installation whose func- capability of the airplane or the ability tioning is required for type certifi- of the crew to cope with adverse oper- cation or under operating rules and ating conditions is improbable. that requires a power supply is an ‘‘es- (c) Warning information must be pro- sential load’’ on the power supply. The vided to alert the crew to unsafe sys- power sources and the system must be tem operating conditions, and to enable to supply the following power able them to take appropriate correc- loads in probable operating combina- tive action. Systems, controls, and as- tions and for probable durations: sociated monitoring and warning (1) Loads connected to the system means must be designed to minimize with the system functioning normally. crew errors which could create addi- (2) Essential loads, after failure of tional hazards. any one prime mover, power converter, (d) Compliance with the require- or energy storage device. ments of paragraph (b) of this section (3) Essential loads after failure of— must be shown by analysis, and where (i) Any one engine on two-engine air- necessary, by appropriate ground, planes; and flight, or simulator tests. The analysis must consider— (ii) Any two engines on airplanes (1) Possible modes of failure, includ- with three or more engines. ing malfunctions and damage from ex- (4) Essential loads for which an alter- ternal sources. nate source of power is required, after (2) The probability of multiple fail- any failure or malfunction in any one ures and undetected failures. power supply system, distribution sys- (3) The resulting effects on the air- tem, or other utilization system. plane and occupants, considering the (b) In determining compliance with stage of flight and operating condi- paragraphs (a)(2) and (3) of this section, tions, and the power loads may be assumed to be (4) The crew warning cues, corrective reduced under a monitoring procedure action required, and the capability of consistent with safety in the kinds of detecting faults. operation authorized. Loads not re- (e) In showing compliance with para- quired in controlled flight need not be graphs (a) and (b) of this section with considered for the two-engine-inoper- regard to the electrical system and ative condition on airplanes with three equipment design and installation, or more engines. critical environmental conditions must be considered. For electrical genera- [Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] tion, distribution, and utilization equipment required by or used in com- § 25.1316 Electrical and electronic sys- plying with this chapter, except equip- tem lightning protection. ment covered by Technical Standard (a) Each electrical and electronic Orders containing environmental test system that performs a function, for

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which failure would prevent the contin- ment providing these functions is ex- ued safe flight and landing of the air- posed to equipment HIRF test level 1 plane, must be designed and installed or 2, as described in appendix L to this so that— part. (1) The function is not adversely af- (c) Each electrical and electronic sys- fected during and after the time the tem that performs a function whose airplane is exposed to lightning; and failure would reduce the capability of (2) The system automatically recov- the airplane or the ability of the ers normal operation of that function flightcrew to respond to an adverse op- in a timely manner after the airplane erating condition must be designed and is exposed to lightning. installed so the system is not adversely (b) Each electrical and electronic affected when the equipment providing system that performs a function, for the function is exposed to equipment which failure would reduce the capa- HIRF test level 3, as described in ap- bility of the airplane or the ability of pendix L to this part. the flightcrew to respond to an adverse (d) Before December 1, 2012, an elec- operating condition, must be designed trical or electronic system that per- and installed so that the function re- forms a function whose failure would covers normal operation in a timely prevent the continued safe flight and manner after the airplane is exposed to landing of an airplane may be designed lightning. and installed without meeting the pro- [Doc. No. FAA–2010–0224, Amdt. 25–134, 76 FR visions of paragraph (a) provided— 33135, June 8, 2011] (1) The system has previously been shown to comply with special condi- § 25.1317 High-intensity Radiated tions for HIRF, prescribed under § 21.16, Fields (HIRF) Protection. issued before December 1, 2007; (a) Except as provided in paragraph (2) The HIRF immunity characteris- (d) of this section, each electrical and tics of the system have not changed electronic system that performs a func- since compliance with the special con- tion whose failure would prevent the ditions was demonstrated; and continued safe flight and landing of the (3) The data used to demonstrate airplane must be designed and installed compliance with the special conditions so that— is provided. (1) The function is not adversely af- [Doc. No. FAA–2006–23657, 72 FR 44025, Aug. 6, fected during and after the time the 2007] airplane is exposed to HIRF environment I, as described in appendix L to INSTRUMENTS: INSTALLATION this part; (2) The system automatically recov- § 25.1321 Arrangement and visibility. ers normal operation of that function, (a) Each flight, navigation, and pow- in a timely manner, after the airplane erplant instrument for use by any pilot is exposed to HIRF environment I, as must be plainly visible to him from his described in appendix L to this part, station with the minimum practicable unless the system’s recovery conflicts deviation from his normal position and with other operational or functional line of vision when he is looking for- requirements of the system; and ward along the flight path. (3) The system is not adversely af- (b) The flight instruments required fected during and after the time the by § 25.1303 must be grouped on the in- airplane is exposed to HIRF environ- strument panel and centered as nearly ment II, as described in appendix L to as practicable about the vertical plane this part. of the pilot’s forward vision. In addi- (b) Each electrical and electronic tion— system that performs a function whose (1) The instrument that most effec- failure would significantly reduce the tively indicates attitude must be on capability of the airplane or the ability the panel in the top center position; of the flightcrew to respond to an ad- (2) The instrument that most effec- verse operating condition must be de- tively indicates airspeed must be adja- signed and installed so the system is cent to and directly to the left of the not adversely affected when the equip- instrument in the top center position:

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(3) The instrument that most effec- (3) Advisory: For conditions that re- tively indicates altitude must be adja- quire flightcrew awareness and may re- cent to and directly to the right of the quire subsequent flightcrew response. instrument in the top center position; (c) Warning and caution alerts must: and (1) Be prioritized within each cat- (4) The instrument that most effec- egory, when necessary. tively indicates direction of flight (2) Provide timely attention-getting must be adjacent to and directly below cues through at least two different the instrument in the top center posi- senses by a combination of aural, vis- tion. ual, or tactile indications. (c) Required powerplant instruments (3) Permit each occurrence of the at- must be closely grouped on the instru- tention-getting cues required by para- ment panel. In addition— graph (c)(2) of this section to be ac- (1) The location of identical power- knowledged and suppressed, unless plant instruments for the engines must they are required to be continuous. prevent confusion as to which engine (d) The alert function must be de- each instrument relates; and signed to minimize the effects of false (2) Powerplant instruments vital to and nuisance alerts. In particular, it the safe operation of the airplane must must be designed to: be plainly visible to the appropriate (1) Prevent the presentation of an crewmembers. alert that is inappropriate or unneces- (d) Instrument panel vibration may sary. not damage or impair the accuracy of (2) Provide a means to suppress an any instrument. attention-getting component of an (e) If a visual indicator is provided to alert caused by a failure of the alerting indicate malfunction of an instrument, function that interferes with the it must be effective under all probable flightcrew’s ability to safely operate cockpit lighting conditions. the airplane. This means must not be readily available to the flightcrew so [Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as that it could be operated inadvertently amended by Amdt. 25–41, 42 FR 36970, July 18, or by habitual reflexive action. When 1977] an alert is suppressed, there must be a clear and unmistakable annunciation § 25.1322 Flightcrew alerting. to the flightcrew that the alert has (a) Flightcrew alerts must: been suppressed. (1) Provide the flightcrew with the (e) Visual alert indications must: information needed to: (1) Conform to the following color (i) Identify non-normal operation or convention: airplane system conditions, and (i) Red for warning alert indications. (ii) Determine the appropriate ac- (ii) Amber or yellow for caution alert tions, if any. indications. (2) Be readily and easily detectable (iii) Any color except red or green for and intelligible by the flightcrew under advisory alert indications. all foreseeable operating conditions, (2) Use visual coding techniques, to- including conditions where multiple gether with other alerting function ele- alerts are provided. ments on the flight deck, to distin- (3) Be removed when the alerting guish between warning, caution, and condition no longer exists. advisory alert indications, if they are (b) Alerts must conform to the fol- presented on monochromatic displays lowing prioritization hierarchy based that are not capable of conforming to on the urgency of flightcrew awareness the color convention in paragraph (e)(1) and response. of this section. (1) Warning: For conditions that re- (f) Use of the colors red, amber, and quire immediate flightcrew awareness yellow on the flight deck for functions and immediate flightcrew response. other than flightcrew alerting must be limited and must not adversely affect (2) Caution: For conditions that re- flightcrew alerting. quire immediate flightcrew awareness and subsequent flightcrew response. [Amdt. 25–131, 75 FR 67209, Nov. 2, 2010]

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§ 25.1323 Airspeed indicating system. (h) Each system must be arranged, so For each airspeed indicating system, far as practicable, to prevent malfunc- the following apply: tion or serious error due to the entry of (a) Each airspeed indicating instru- moisture, dirt, or other substances. ment must be approved and must be (i) Each system must have a heated calibrated to indicate true airspeed (at pitot tube or an equivalent means of sea level with a standard atmosphere) preventing malfunction due to icing. with a minimum practicable instru- (j) Where duplicate airspeed indica- ment calibration error when the cor- tors are required, their respective pitot responding pitot and static pressures tubes must be far enough apart to are applied. avoid damage to both tubes in a colli- (b) Each system must be calibrated sion with a bird. to determine the system error (that is, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the relation between IAS and CAS) in amended by Amdt. 25–57, 49 FR 6849, Feb. 23, flight and during the accelerated take- 1984; Amdt. 25–108, 67 FR 70828, Nov. 26, 2002; off ground run. The ground run calibra- Amdt. 25–109, 67 FR 76656, Dec. 12, 2002] tion must be determined— (1) From 0.8 of the minimum value of § 25.1325 Static pressure systems. V1 to the maximum value of V2, consid- (a) Each instrument with static air ering the approved ranges of altitude case connections must be vented to the and weight; and outside atmosphere through an appro- (2) With the flaps and power settings priate piping system. corresponding to the values determined (b) Each static port must be designed in the establishment of the takeoff and located in such manner that the path under § 25.111 assuming that the static pressure system performance is critical engine fails at the minimum least affected by airflow variation, or value of V1. by moisture or other foreign matter, (c) The airspeed error of the installa- and that the correlation between air tion, excluding the airspeed indicator pressure in the static pressure system instrument calibration error, may not and true ambient atmospheric static exceed three percent or five knots, pressure is not changed when the air- whichever is greater, throughout the plane is exposed to the continuous and speed range, from— intermittent maximum icing condi- (1) VMO to 1.23 VSR1, with flaps re- tions defined in appendix C of this part. tracted; and (c) The design and installation of the (2) 1.23 VSR0 to VFE with flaps in the static pressure system must be such landing position. that— (d) From 1.23 VSR to the speed at (1) Positive drainage of moisture is which stall warning begins, the IAS provided; chafing of the tubing and ex- must change perceptibly with CAS and cessive distortion or restriction at in the same sense, and at speeds below bends in the tubing is avoided; and the stall warning speed the IAS must not materials used are durable, suitable for change in an incorrect sense. the purpose intended, and protected 2 (e) From VMO to VMO + ⁄3 (VDF ¥ against corrosion; and VMO), the IAS must change perceptibly (2) It is airtight except for the port with CAS and in the same sense, and at into the atmosphere. A proof test must higher speeds up to VDF the IAS must be conducted to demonstrate the integ- not change in an incorrect sense. rity of the static pressure system in (f) There must be no indication of the following manner: airspeed that would cause undue dif- (i) Unpressurized airplanes. Evacuate ficulty to the pilot during the takeoff the static pressure system to a pres- between the initiation of rotation and sure differential of approximately 1 the achievement of a steady climbing inch of mercury or to a reading on the condition. altimeter, 1,000 feet above the airplane (g) The effects of airspeed indicating elevation at the time of the test. With- system lag may not introduce signifi- out additional pumping for a period of cant takeoff indicated airspeed bias, or 1 minute, the loss of indicated altitude significant errors in takeoff or accel- must not exceed 100 feet on the altim- erate-stop distances. eter.

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(ii) Pressurized airplanes. Evacuate static pressure source being open or the static pressure system until a pres- blocked. sure differential equivalent to the max- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as imum cabin pressure differential for amended by Amdt. 25–5, 30 FR 8261, June 29, which the airplane is type certificated 1965; Amdt. 25–12, 32 FR 7587, May 24, 1967; is achieved. Without additional pump- Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. ing for a period of 1 minute, the loss of 25–108, 67 FR 70828, Nov. 26, 2002] indicated altitude must not exceed 2 percent of the equivalent altitude of § 25.1326 Pitot heat indication systems. the maximum cabin differential pres- If a flight instrument pitot heating sure or 100 feet, whichever is greater. system is installed, an indication sys- (d) Each pressure altimeter must be tem must be provided to indicate to approved and must be calibrated to in- the flight crew when that pitot heating dicate pressure altitude in a standard system is not operating. The indication atmosphere, with a minimum prac- system must comply with the following ticable calibration error when the cor- requirements: responding static pressures are applied. (a) The indication provided must in- (e) Each system must be designed and corporate an amber light that is in installed so that the error in indicated clear view of a flight crewmember. pressure altitude, at sea level, with a (b) The indication provided must be standard atmosphere, excluding instru- designed to alert the flight crew if ei- ment calibration error, does not result ther of the following conditions exist: in an error of more than ±30 feet per 100 (1) The pitot heating system is knots speed for the appropriate con- switched ‘‘off’’. figuration in the speed range between (2) The pitot heating system is 1.23 VSR0 with flaps extended and 1.7 switched ‘‘on’’ and any pitot tube heat- VSR1 with flaps retracted. However, the ing element is inoperative. error need not be less than ±30 feet. (f) If an altimeter system is fitted [Amdt. 25–43, 43 FR 10339, Mar. 13, 1978] with a device that provides corrections § 25.1327 Magnetic direction indicator. to the altimeter indication, the device must be designed and installed in such (a) Each magnetic direction indicator manner that it can be bypassed when it must be installed so that its accuracy malfunctions, unless an alternate al- is not excessively affected by the air- timeter system is provided. Each cor- plane’s vibration or magnetic fields. rection device must be fitted with a (b) The compensated installation means for indicating the occurrence of may not have a deviation, in level reasonably probable malfunctions, in- flight, greater than 10 degrees on any cluding power failure, to the flight heading. crew. The indicating means must be effective for any cockpit lighting condi- § 25.1329 Flight guidance system. tion likely to occur. (a) Quick disengagement controls for (g) Except as provided in paragraph the autopilot and autothrust functions (h) of this section, if the static pressure must be provided for each pilot. The system incorporates both a primary autopilot quick disengagement con- and an alternate static pressure source, trols must be located on both control the means for selecting one or the wheels (or equivalent). The autothrust other source must be designed so quick disengagement controls must be that— located on the thrust control levers. (1) When either source is selected, the Quick disengagement controls must be other is blocked off; and readily accessible to each pilot while (2) Both sources cannot be blocked operating the control wheel (or equiva- off simultaneously. lent) and thrust control levers. (h) For unpressurized airplanes, para- (b) The effects of a failure of the sys- graph (g)(1) of this section does not tem to disengage the autopilot or apply if it can be demonstrated that autothrust functions when manually the static pressure system calibration, commanded by the pilot must be as- when either static pressure source is sessed in accordance with the require- selected, is not changed by the other ments of § 25.1309.

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(c) Engagement or switching of the consistent manner. The indications flight guidance system, a mode, or a must be visible to each pilot under all sensor may not cause a transient re- expected lighting conditions. sponse of the airplane’s flight path any (j) Following disengagement of the greater than a minor transient, as de- autopilot, a warning (visual and audi- fined in paragraph (n)(1) of this section. tory) must be provided to each pilot (d) Under normal conditions, the dis- and be timely and distinct from all engagement of any automatic control other cockpit warnings. function of a flight guidance system (k) Following disengagement of the may not cause a transient response of autothrust function, a caution must be the airplane’s flight path any greater provided to each pilot. than a minor transient. (l) The autopilot may not create a (e) Under rare normal and non-nor- potential hazard when the flightcrew mal conditions, disengagement of any applies an override force to the flight automatic control function of a flight controls. guidance system may not result in a transient any greater than a signifi- (m) During autothrust operation, it cant transient, as defined in paragraph must be possible for the flightcrew to (n)(2) of this section. move the thrust levers without requir- (f) The function and direction of mo- ing excessive force. The autothrust tion of each command reference con- may not create a potential hazard trol, such as heading select or vertical when the flightcrew applies an override speed, must be plainly indicated on, or force to the thrust levers. adjacent to, each control if necessary (n) For purposes of this section, a to prevent inappropriate use or confu- transient is a disturbance in the con- sion. trol or flight path of the airplane that (g) Under any condition of flight ap- is not consistent with response to propriate to its use, the flight guidance flightcrew inputs or environmental system may not produce hazardous conditions. loads on the airplane, nor create haz- (1) A minor transient would not sig- ardous deviations in the flight path. nificantly reduce safety margins and This applies to both fault-free oper- would involve flightcrew actions that ation and in the event of a malfunc- are well within their capabilities. A tion, and assumes that the pilot begins minor transient may involve a slight corrective action within a reasonable increase in flightcrew workload or period of time. some physical discomfort to passengers (h) When the flight guidance system or cabin crew. is in use, a means must be provided to (2) A significant transient may lead avoid excursions beyond an acceptable to a significant reduction in safety margin from the speed range of the margins, an increase in flightcrew normal flight envelope. If the airplane workload, discomfort to the flightcrew, experiences an excursion outside this or physical distress to the passengers range, a means must be provided to or cabin crew, possibly including non- prevent the flight guidance system fatal injuries. Significant transients do from providing guidance or control to not require, in order to remain within an unsafe speed. or recover to the normal flight enve- (i) The flight guidance system func- lope, any of the following: tions, controls, indications, and alerts (i) Exceptional piloting skill, alert- must be designed to minimize ness, or strength. flightcrew errors and confusion con- (ii) Forces applied by the pilot which cerning the behavior and operation of are greater than those specified in the flight guidance system. Means § 25.143(c). must be provided to indicate the cur- (iii) Accelerations or attitudes in the rent mode of operation, including any airplane that might result in further armed modes, transitions, and rever- sions. Selector switch position is not hazard to secured or non-secured occu- an acceptable means of indication. The pants. controls and indications must be [Doc. No. FAA–2004–18775, 71 FR 18191, Apr. grouped and presented in a logical and 11, 2006]

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§ 25.1331 Instruments using a power vided by the instruments, including at- supply. titude, direction, airspeed, and altitude (a) For each instrument required by will remain available to the pilots, § 25.1303(b) that uses a power supply, without additional crewmember ac- the following apply: tion, after any single failure or com- (1) Each instrument must have a vis- bination of failures that is not shown ual means integral with, the instru- to be extremely improbable; and ment, to indicate when power adequate (c) Additional instruments, systems, to sustain proper instrument perform- or equipment may not be connected to ance is not being supplied. The power the operating systems for the required must be measured at or near the point instruments, unless provisions are where it enters the instruments. For made to ensure the continued normal electric instruments, the power is con- functioning of the required instru- sidered to be adequate when the volt- ments in the event of any malfunction age is within approved limits. of the additional instruments, systems, (2) Each instrument must, in the or equipment which is not shown to be event of the failure of one power extremely improbable. source, be supplied by another power [Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as source. This may be accomplished amended by Amdt. 25–41, 42 FR 36970, July 18, automatically or by manual means. 1977] (3) If an instrument presenting navigation data receives information from § 25.1337 Powerplant instruments. sources external to that instrument (a) Instruments and instrument lines. and loss of that information would (1) Each powerplant and auxiliary render the presented data unreliable, power unit instrument line must meet the instrument must incorporate a visual means to warn the crew, when such the requirements of §§ 25.993 and 25.1183. loss of information occurs, that the (2) Each line carrying flammable presented data should not be relied fluids under pressure must— upon. (i) Have restricting orifices or other (b) As used in this section, ‘‘instru- safety devices at the source of pressure ment’’ includes devices that are phys- to prevent the escape of excessive fluid ically contained in one unit, and de- if the line fails; and vices that are composed of two or more (ii) Be installed and located so that physically separate units or compo- the escape of fluids would not create a nents connected together (such as a re- hazard. mote indicating gyroscopic direction (3) Each powerplant and auxiliary indicator that includes a magnetic power unit instrument that utilizes sensing element, a gyroscopic unit, an flammable fluids must be installed and amplifier and an indicator connected located so that the escape of fluid together). would not create a hazard. (b) Fuel quantity indicator. There [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–41, 42 FR 36970, July 18, must be means to indicate to the flight 1977] crewmembers, the quantity, in gallons or equivalent units, of usable fuel in § 25.1333 Instrument systems. each tank during flight. In addition— For systems that operate the instru- (1) Each fuel quantity indicator must ments required by § 25.1303(b) which are be calibrated to read ‘‘zero’’ during located at each pilot’s station— level flight when the quantity of fuel (a) Means must be provided to con- remaining in the tank is equal to the nect the required instruments at the unusable fuel supply determined under first pilot’s station to operating sys- § 25.959; tems which are independent of the op- (2) Tanks with interconnected outlets erating systems at other flight crew and airspaces may be treated as one stations, or other equipment; tank and need not have separate indi- (b) The equipment, systems, and in- cators; and stallations must be designed so that (3) Each exposed sight gauge, used as one display of the information essen- a fuel quantity indicator, must be pro- tial to the safety of flight which is protected against damage.

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(c) Fuel flowmeter system. If a fuel (1) Power sources function properly flowmeter system is installed, each when independent and when connected metering component must have a in combination; means for bypassing the fuel supply if (2) No failure or malfunction of any malfunction of that component se- power source can create a hazard or verely restricts fuel flow. impair the ability of remaining sources (d) Oil quantity indicator. There must to supply essential loads; be a stick gauge or equivalent means (3) The system voltage and frequency to indicate the quantity of oil in each (as applicable) at the terminals of all tank. If an oil transfer or reserve oil essential load equipment can be main- supply system is installed, there must tained within the limits for which the be a means to indicate to the flight equipment is designed, during any crew, in flight, the quantity of oil in probable operating condition; and each tank. (4) System transients due to switch- (e) Turbopropeller blade position indi- ing, fault clearing, or other causes do cator. Required turbopropeller blade not make essential loads inoperative, position indicators must begin indi- and do not cause a smoke or fire haz- cating before the blade moves more ard. than eight degrees below the flight low (5) There are means accessible, in pitch stop. The source of indication flight, to appropriate crewmembers for must directly sense the blade position. the individual and collective dis- (f) Fuel pressure indicator. There must connection of the electrical power be means to measure fuel pressure, in sources from the system. each system supplying reciprocating (6) There are means to indicate to ap- engines, at a point downstream of any propriate crewmembers the generating fuel pump except fuel injection pumps. system quantities essential for the safe In addition— operation of the system, such as the (1) If necessary for the maintenance voltage and current supplied by each of proper fuel delivery pressure, there generator. must be a connection to transmit the carburetor air intake static pressure to (c) External power. If provisions are the proper pump relief valve connec- made for connecting external power to tion; and the airplane, and that external power (2) If a connection is required under can be electrically connected to equip- paragraph (f)(1) of this section, the ment other than that used for engine gauge balance lines must be independ- starting, means must be provided to ently connected to the carburetor inlet ensure that no external power supply pressure to avoid erroneous readings. having a reverse polarity, or a reverse phase sequence, can supply power to [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the airplane’s electrical system. amended by Amdt. 25–40, 42 FR 15044, Mar. 17, (d) Operation without normal electrical 1977] power. It must be shown by analysis, ELECTRICAL SYSTEMS AND EQUIPMENT tests, or both, that the airplane can be operated safely in VFR conditions, for § 25.1351 General. a period of not less than five minutes, (a) Electrical system capacity. The re- with the normal electrical power (elec- quired generating capacity, and num- trical power sources excluding the bat- ber and kinds of power sources must— tery) inoperative, with critical type (1) Be determined by an electrical fuel (from the standpoint of flameout load analysis; and and restart capability), and with the (2) Meet the requirements of § 25.1309. airplane initially at the maximum cer- (b) Generating system. The generating tificated altitude. Parts of the elec- system includes electrical power trical system may remain on if— sources, main power busses, trans- (1) 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 both the part must be designed so that— turned off and the part turned on; and

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(2) The parts turned on are elec- (6) Nickel cadmium battery installa- trically and mechanically isolated tions must have— from the parts turned off. (i) A system to control the charging rate of the battery automatically so as [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–41, 42 FR 36970, July 18, to prevent battery overheating; 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990] (ii) A battery temperature sensing and over-temperature warning system § 25.1353 Electrical equipment and in- with a means for disconnecting the stallations. battery from its charging source in the (a) Electrical equipment and controls event of an over-temperature condi- must be installed so that operation of tion; or any one unit or system of units will (iii) A battery failure sensing and not adversely affect the simultaneous warning system with a means for dis- operation of any other electrical unit connecting the battery from its charg- or system essential to safe operation. ing source in the event of battery fail- Any electrical interference likely to be ure. present in the airplane must not result (c) Electrical bonding must provide in hazardous effects on the airplane or an adequate electrical return path its systems. under both normal and fault condi- (b) Storage batteries must be de- tions, on airplanes having grounded signed and installed as follows: electrical systems. (1) Safe cell temperatures and pres- [Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] sures must be maintained during any probable charging or discharging con- § 25.1355 Distribution system. dition. No uncontrolled increase in cell (a) The distribution system includes temperature may result when the bat- the distribution busses, their associ- tery is recharged (after previous com- ated feeders, and each control and pro- plete discharge)— tective device. (i) At maximum regulated voltage or power; (b) [Reserved] (ii) During a flight of maximum dura- (c) If two independent sources of elec- tion; and trical power for particular equipment (iii) Under the most adverse cooling or systems are required by this chap- condition likely to occur in service. ter, in the event of the failure of one (2) Compliance with paragraph (b)(1) power source for such equipment or of this section must be shown by test system, another power source (includ- unless experience with similar bat- ing its separate feeder) must be auto- teries and installations has shown that matically provided or be manually se- maintaining safe cell temperatures and lectable to maintain equipment or sys- pressures presents no problem. tem operation. (3) No explosive or toxic gases emit- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ted by any battery in normal oper- amended by Amdt. 25–23, 35 FR 5679, Apr. 8, ation, or as the result of any probable 1970; Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] malfunction in the charging system or battery installation, may accumulate § 25.1357 Circuit protective devices. in hazardous quantities within the air- (a) Automatic protective devices plane. must be used to minimize distress to (4) No corrosive fluids or gases that the electrical system and hazard to the may escape from the battery may dam- airplane in the event of wiring faults or age surrounding airplane structures or serious malfunction of the system or adjacent essential equipment. connected equipment. (5) Each nickel cadmium battery in- (b) The protective and control de- stallation must have provisions to pre- vices in the generating system must be vent any hazardous effect on structure designed to de-energize and disconnect or essential systems that may be faulty power sources and power trans- caused by the maximum amount of mission equipment from their associ- heat the battery can generate during a ated busses with sufficient rapidity to short circuit of the battery or of indi- provide protection from hazardous vidual cells. over-voltage and other malfunctioning.

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(c) Each resettable circuit protective gency landing or ditching. The circuits device must be designed so that, when for these services must be designed, an overload or circuit fault exists, it protected, and installed so that the will open the circuit irrespective of the risk of the services being rendered inef- position of the operating control. fective under these emergency condi- (d) If the ability to reset a circuit tions is minimized. breaker or replace a fuse is essential to [Amdt. 25–123, 72 FR 63406, Nov. 8, 2007] safety in flight, that circuit breaker or fuse must be located and identified so § 25.1363 Electrical system tests. that it can be readily reset or replaced (a) When laboratory tests of the elec- in flight. Where fuses are used, there trical system are conducted— must be spare fuses for use in flight (1) The tests must be performed on a equal to at least 50% of the number of mock-up using the same generating fuses of each rating required for com- equipment used in the airplane; plete circuit protection. (2) The equipment must simulate the (e) Each circuit for essential loads electrical characteristics of the dis- must have individual circuit protec- tribution wiring and connected loads to tion. However, individual protection the extent necessary for valid test re- for each circuit in an essential load sults; and system (such as each position light cir- (3) Laboratory generator drives must cuit in a system) is not required. simulate the actual prime movers on (f) For airplane systems for which the airplane with respect to their reac- the ability to remove or reset power tion to generator loading, including during normal operations is necessary, loading due to faults. the system must be designed so that (b) For each flight condition that circuit breakers are not the primary cannot be simulated adequately in the means to remove or reset system power laboratory or by ground tests on the unless specifically designed for use as a airplane, flight tests must be made. switch. (g) Automatic reset circuit breakers § 25.1365 Electrical appliances, motors, may be used as integral protectors for and transformers. electrical equipment (such as thermal (a) Domestic appliances must be de- cut-outs) if there is circuit protection signed and installed so that in the to protect the cable to the equipment. event of failures of the electrical supply or control system, the require- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–123, 72 FR 63405, Nov. 8, ments of § 25.1309(b), (c), and (d) will be 2007] satisfied. Domestic appliances are items such as cooktops, ovens, coffee § 25.1360 Precautions against injury. makers, water heaters, refrigerators, (a) Shock. The electrical system and toilet flush systems that are must be designed to minimize risk of placed on the airplane to provide serv- electric shock to crew, passengers, and ice amenities to passengers. (b) Galleys and cooking appliances servicing personnel and to mainte- must be installed in a way that mini- nance personnel using normal pre- mizes risk of overheat or fire. cautions. (c) Domestic appliances, particularly (b) Burns. The temperature of any those in galley areas, must be installed part that may be handled by a crew- or protected so as to prevent damage or member during normal operations contamination of other equipment or must not cause dangerous inadvertent systems from fluids or vapors which movement by the crewmember or in- may be present during normal oper- jury to the crewmember. ation or as a result of spillage, if such [Amdt. 25–123, 72 FR 63406, Nov. 8, 2007] damage or contamination could create a hazardous condition. § 25.1362 Electrical supplies for emer- (d) Unless compliance with § 25.1309(b) gency conditions. is provided by the circuit protective A suitable electrical supply must be device required by § 25.1357(a), electric provided to those services required for motors and transformers, including emergency procedures after an emer- those installed in domestic systems,

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must have a suitable thermal protec- far apart as practicable and installed tion device to prevent overheating forward on the airplane so that, with under normal operation and failure the airplane in the normal flying posi- conditions, if overheating could create tion, the red light is on the left side a smoke or fire hazard. and the green light is on the right side. [Amdt. 25–123, 72 FR 63406, Nov. 8, 2007] Each light must be approved. (c) Rear position light. The rear posi- LIGHTS tion light must be a white light mounted as far aft as practicable on the tail § 25.1381 Instrument lights. or on each wing tip, and must be ap- (a) The instrument lights must— proved. (1) Provide sufficient illumination to (d) Light covers and color filters. Each make each instrument, switch and light cover or color filter must be at other device necessary for safe oper- least flame resistant and may not ation easily readable unless sufficient change color or shape or lose any ap- illumination is available from another preciable light transmission during source; and normal use. (2) Be installed so that— [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (i) Their direct rays are shielded from amended by Amdt. 25–38, 41 FR 55468, Dec. 20, the pilot’s eyes; and 1976] (ii) No objectionable reflections are visible to the pilot. § 25.1387 Position light system dihe- (b) Unless undimmed instrument dral angles. lights are satisfactory under each ex- (a) Except as provided in paragraph pected flight condition, there must be a (e) of this section, each forward and means to control the intensity of illu- rear position light must, as installed, mination. show unbroken light within the dihe- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as dral angles described in this section. amended by Amdt. 25–72, 55 FR 29785, July 20, (b) Dihedral angle L (left) is formed 1990] by two intersecting vertical planes, the first parallel to the longitudinal axis of § 25.1383 Landing lights. the airplane, and the other at 110 de- (a) Each landing light must be ap- grees to the left of the first, as viewed proved, and must be installed so that— when looking forward along the longi- (1) No objectionable glare is visible tudinal axis. to the pilot; (c) Dihedral angle R (right) is formed (2) The pilot is not adversely affected by two intersecting vertical planes, the by halation; and first parallel to the longitudinal axis of (3) It provides enough light for night the airplane, and the other at 110 de- landing. grees to the right of the first, as viewed (b) Except when one switch is used when looking forward along the longi- for the lights of a multiple light instal- tudinal axis. lation at one location, there must be a (d) Dihedral angle A (aft) is formed separate switch for each light. by two intersecting vertical planes (c) There must be a means to indicate making angles of 70 degrees to the to the pilots when the landing lights right and to the left, respectively, to a are extended. vertical plane passing through the longitudinal axis, as viewed when looking § 25.1385 Position light system installa- aft along the longitudinal axis. tion. (e) If the rear position light, when (a) General. Each part of each posi- mounted as far aft as practicable in action light system must meet the appli- cordance with § 25.1385(c), cannot show cable requirements of this section and unbroken light within dihedral angle A each system as a whole must meet the (as defined in paragraph (d) of this sec- requirements of §§ 25.1387 through tion), a solid angle or angles of ob- 25.1397. structed visibility totaling not more (b) Forward position lights. Forward than 0.04 steradians is allowable within position lights must consist of a red that dihedral angle, if such solid angle and a green light spaced laterally as is within a cone whose apex is at the

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rear position light and whose elements more than 100 candles, the maximum make an angle of 30° with a vertical overlap intensities between them may line passing through the rear position exceed the values given in § 25.1395 if light. the overlap intensity in Area A is not [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as more than 10 percent of peak position amended by Amdt. 25–30, 36 FR 21278, Nov. 5, light intensity and the overlap inten- 1971] sity in Area B is not greater than 2.5 percent of peak position light inten- § 25.1389 Position light distribution sity. and intensities. (a) General. The intensities prescribed § 25.1391 Minimum intensities in the in this section must be provided by new horizontal plane of forward and rear position lights. equipment with light covers and color filters in place. Intensities must be de- Each position light intensity must termined with the light source oper- equal or exceed the applicable values in ating at a steady value equal to the av- the following table: erage luminous output of the source at Angle from right the normal operating voltage of the or left of longitu- Dihedral angle (light in- dinal axis, meas- Intensity airplane. The light distribution and included) ured from dead (candles) tensity of each position light must ahead meet the requirements of paragraph (b) L and R (forward red and 0° to 10° ...... 40 of this section. green). 10° to 20° ...... 30 (b) Forward and rear position lights. 20° to 110° ...... 5 The light distribution and intensities A (rear white) ...... 110° to 180° ...... 20 of forward and rear position lights must be expressed in terms of min- § 25.1393 Minimum intensities in any imum intensities in the horizontal vertical plane of forward and rear plane, minimum intensities in any position lights. vertical plane, and maximum inten- Each position light intensity must sities in overlapping beams, within di- equal or exceed the applicable values in hedral angles L, R, and A, and must the following table: meet the following requirements: (1) Intensities in the horizontal plane. Angle above or below the horizontal plane Intensity, l Each intensity in the horizontal plane 0° ...... 1.00 (the plane containing the longitudinal 0° to 5° ...... 0.90 axis of the airplane and perpendicular 5° to 10° ...... 0.80 10° to 15° ...... 0.70 to the plane of symmetry of the air- 15° to 20° ...... 0.50 plane) must equal or exceed the values 20° to 30° ...... 0.30 in § 25.1391. 30° to 40° ...... 0.10 ° ° (2) Intensities in any vertical plane. 40 to 90 ...... 0.05 Each intensity in any vertical plane (the plane perpendicular to the hori- § 25.1395 Maximum intensities in over- zontal plane) must equal or exceed the lapping beams of forward and rear appropriate value in § 25.1393, where I is position lights. the minimum intensity prescribed in No position light intensity may ex- § 25.1391 for the corresponding angles in ceed the applicable values in the fol- the horizontal plane. lowing table, except as provided in (3) Intensities in overlaps between adja- § 25.1389(b)(3). cent signals. No intensity in any overlap between adjacent signals may ex- Maximum intensity ceed the values given in § 25.1395, except Overlaps Area A Area B that higher intensities in overlaps may (candles) (candles) be used with main beam intensities Green in dihedral angle L ...... 10 1 substantially greater than the minima Red in dihedral angle R ...... 10 1 Green in dihedral angle A ...... 5 1 specified in §§ 25.1391 and 25.1393 if the Red in dihedral angle A ...... 5 1 overlap intensities in relation to the Rear white in dihedral angle L ...... 5 1 main beam intensities do not adversely Rear white in dihedral angle R ..... 5 1 affect signal clarity. When the peak intensity of the forward position lights is Where—

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(a) Area A includes all directions in (2) Meets the requirements of para- the adjacent dihedral angle that pass graphs (b) through (f) of this section. through the light source and intersect (b) Field of coverage. The system must the common boundary plane at more consist of enough lights to illuminate than 10 degrees but less than 20 de- the vital areas around the airplane grees; and considering the physical configuration (b) Area B includes all directions in and flight characteristics of the air- the adjacent dihedral angle that pass plane. The field of coverage must ex- through the light source and intersect tend in each direction within at least the common boundary plane at more 75 degrees above and 75 degrees below than 20 degrees. the horizontal plane of the airplane, except that a solid angle or angles of § 25.1397 Color specifications. obstructed visibility totaling not more Each position light color must have than 0.03 steradians is allowable within the applicable International Commis- a solid angle equal to 0.15 steradians sion on Illumination chromaticity co- centered about the longitudinal axis in ordinates as follows: the rearward direction. (a) Aviation red— (c) Flashing characteristics. The ar- y is not greater than 0.335; and rangement of the system, that is, the z is not greater than 0.002. number of light sources, beam width, speed of rotation, and other character- (b) Aviation green— istics, must give an effective flash fre- x is not greater than 0.440¥0.320y ; quency of not less than 40, nor more x is not greater than y¥0.170; and than 100 cycles per minute. The effec- y is not less than 0.390¥0.170x. tive flash frequency is the frequency at (c) Aviation white— which the airplane’s complete anticollision light system is observed from x is not less than 0.300 and not greater than a distance, and applies to each sector 0.540; of light including any overlaps that y is not less than x¥0.040; or y0¥0.010, whichever is the smaller; and exist when the system consists of more y is not greater than x+0.020 nor 0.636¥0.400x; than one light source. In overlaps, Where y0 is the y coordinate of the Planckian flash frequencies may exceed 100, but radiator for the value of x considered. not 180 cycles per minute. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (d) Color. Each anticollision light amended by Amdt. 25–27, 36 FR 12972, July 10, must be either aviation red or aviation 1971] white and must meet the applicable requirements of § 25.1397. § 25.1399 Riding light. (e) Light intensity. The minimum (a) Each riding (anchor) light re- light intensities in all vertical planes, quired for a seaplane or amphibian measured with the red filter (if used) must be installed so that it can— and expressed in terms of ‘‘effective’’ (1) Show a white light for at least 2 intensities, must meet the require- nautical miles at night under clear at- ments of paragraph (f) of this section. mospheric conditions; and The following relation must be as- (2) Show the maximum unbroken sumed: light practicable when the airplane is moored or drifting on the water. t2 (b) Externally hung lights may be ∫ I(t)dt I = t1 used. e +− 02. ()tt2 1 § 25.1401 Anticollision light system. where: (a) General. The airplane must have Ie=effective intensity (candles). an anticollision light system that— I(t)=instantaneous intensity as a function of (1) Consists of one or more approved time. anticollision lights located so that t2—t1=flash time interval (seconds). their light will not impair the crew’s vision or detract from the conspicuity Normally, the maximum value of effec- of the position lights; and tive intensity is obtained when t2 and t1 505

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are chosen so that the effective inten- pants for which certification for ditch- sity is equal to the instantaneous in- ing is requested. tensity at t2 and t1. (2) Liferafts must be stowed near (f) Minimum effective intensities for exits through which the rafts can be anticollision lights. Each anticollision launched during an unplanned ditch- light effective intensity must equal or ing. exceed the applicable values in the fol- (3) Rafts automatically or remotely lowing table. released outside the airplane must be attached to the airplane by means of Effective Angle above or below the horizontal plane intensity the static line prescribed in § 25.1415. (candles) (4) The stowage provisions for each 0° to 5° ...... 400 portable liferaft must allow rapid de- 5° to 10° ...... 240 tachment and removal of the raft for 10° to 20° ...... 80 use at other than the intended exits. 20° to 30° ...... 40 30° to 75° ...... 20 (e) Long-range signaling device. The stowage provisions for the long-range [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as signaling device required by § 25.1415 amended by Amdt. 25–27, 36 FR 12972, July 10, must be near an exit available during 1971; Amdt. 25–41, 42 FR 36970, July 18, 1977] an unplanned ditching. (f) Life preserver stowage provisions. § 25.1403 Wing icing detection lights. The stowage provisions for life pre- Unless operations at night in known servers described in § 25.1415 must ac- or forecast icing conditions are prohib- commodate one life preserver for each ited by an operating limitation, a occupant for which certification for means must be provided for illu- ditching is requested. Each life pre- minating or otherwise determining the server must be within easy reach of formation of ice on the parts of the each seated occupant. wings that are critical from the stand- (g) Life line stowage provisions. If cer- point of ice accumulation. Any illu- tification for ditching under § 25.801 is mination that is used must be of a type requested, there must be provisions to that will not cause glare or reflection store life lines. These provisions that would handicap crewmembers in must— the performance of their duties. (1) Allow one life line to be attached to each side of the fuselage; and [Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] (2) Be arranged to allow the life lines SAFETY EQUIPMENT to be used to enable the occupants to stay on the wing after ditching. § 25.1411 General. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) Accessibility. Required safety amended by Amdt. 25–32, 37 FR 3972, Feb. 24, equipment to be used by the crew in an 1972; Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; emergency must be readily accessible. Amdt. 25–53, 45 FR 41593, June 19, 1980; Amdt. (b) Stowage provisions. Stowage provi- 25–70, 54 FR 43925, Oct. 27, 1989; Amdt. 25–79, sions for required emergency equip- 58 FR 45229, Aug. 26, 1993; Amdt. 25–116, 69 FR ment must be furnished and must— 62789, Oct. 27, 2004] (1) Be arranged so that the equip- § 25.1415 Ditching equipment. ment is directly accessible and its location is obvious; and (a) Ditching equipment used in air- (2) Protect the safety equipment planes to be certificated for ditching from inadvertent damage. under § 25.801, and required by the oper- (c) Emergency exit descent device. The ating rules of this chapter, must meet stowage provisions for the emergency the requirements of this section. exit descent devices required by (b) Each liferaft and each life pre- § 25.810(a) must be at each exit for server must be approved. In addition— which they are intended. (1) Unless excess rafts of enough ca- (d) Liferafts. (1) The stowage provi- pacity are provided, the buoyancy and sions for the liferafts described in seating capacity beyond the rated ca- § 25.1415 must accommodate enough pacity of the rafts must accommodate rafts for the maximum number of occu- all occupants of the airplane in the

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event of a loss of one raft of the largest (c) Caution information, such as an rated capacity; and amber caution light or equivalent, (2) Each raft must have a trailing must be provided to alert the line, and must have a static line de- flightcrew when the anti-ice or de-ice signed to hold the raft near the air- system is not functioning normally. plane but to release it if the airplane (d) For turbine engine powered air- becomes totally submerged. planes, the ice protection provisions of (c) Approved survival equipment this section are considered to be appli- must be attached to each liferaft. cable primarily to the airframe. For (d) There must be an approved sur- the powerplant installation, certain ad- vival type emergency locator trans- ditional provisions of subpart E of this mitter for use in one life raft. (e) For airplanes not certificated for part may be found applicable. (e) One ditching under § 25.801 and not having of the following methods of icing detec- approved life preservers, there must be tion and activation of the airframe ice an approved flotation means for each protection system must be provided: occupant. This means must be within (1) A primary ice detection system easy reach of each seated occupant and that automatically activates or alerts must be readily removable from the the flightcrew to activate the airframe airplane. ice protection system; (2) A definition of visual cues for rec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–29, 36 FR 18722, Sept. ognition of the first sign of ice accre- 21, 1971; Amdt 25–50, 45 FR 38348, June 9, 1980; tion on a specified surface combined Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. with an advisory ice detection system 25–82, 59 FR 32057, June 21, 1994] that alerts the flightcrew to activate the airframe ice protection system; or § 25.1419 Ice protection. (3) Identification of conditions con- If the applicant seeks certification ducive to airframe icing as defined by for flight in icing conditions, the air- an appropriate static or total air tem- plane must be able to safely operate in perature and visible moisture for use the continuous maximum and inter- by the flightcrew to activate the air- mittent maximum icing conditions of frame ice protection system. appendix C. To establish this— (a) An analysis must be performed to (f) Unless the applicant shows that establish that the ice protection for the airframe ice protection system the various components of the airplane need not be operated during specific is adequate, taking into account the phases of flight, the requirements of various airplane operational configura- paragraph (e) of this section are appli- tions; and cable to all phases of flight. (b) To verify the ice protection anal- (g) After the initial activation of the ysis, to check for icing anomalies, and airframe ice protection system— to demonstrate that the ice protection (1) The ice protection system must be system and its components are effec- designed to operate continuously; tive, the airplane or its components (2) The airplane must be equipped must be flight tested in the various with a system that automatically cy- operational configurations, in meas- cles the ice protection system; or ured natural atmospheric icing condi- (3) An ice detection system must be tions and, as found necessary, by one provided to alert the flightcrew each or more of the following means: time the ice protection system must be (1) Laboratory dry air or simulated cycled. icing tests, or a combination of both, of (h) Procedures for operation of the the components or models of the com- ice protection system, including acti- ponents. vation and deactivation, must be estab- (2) Flight dry air tests of the ice protection system as a whole, or of its in- lished and documented in the Airplane dividual components. Flight Manual. (3) Flight tests of the airplane or its [Amdt. 25–72, 55 FR 29785, July 20, 1990, as components in measured simulated amended by Amdt. 25–121, 72 FR 44669, Aug. 8, icing conditions. 2007; Amdt. 25–129, 74 FR 38339, Aug. 3, 2009]

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§ 25.1421 Megaphones. communication between seated flight attendants. If a megaphone is installed, a restraining means must be provided that [Doc. No. 26003, 58 FR 45229, Aug. 26, 1993, as is capable of restraining the mega- amended by Amdt. 25–115, 69 FR 40527, July 2, phone when it is subjected to the ulti- 2004] mate inertia forces specified in MISCELLANEOUS EQUIPMENT § 25.561(b)(3). [Amdt. 25–41, 42 FR 36970, July 18, 1977] § 25.1431 Electronic equipment. (a) In showing compliance with § 25.1423 Public address system. § 25.1309 (a) and (b) with respect to A public address system required by radio and electronic equipment and this chapter must— their installations, critical environ- (a) Be powerable when the aircraft is mental conditions must be considered. in flight or stopped on the ground, (b) Radio and electronic equipment after the shutdown or failure of all en- must be supplied with power under the gines and auxiliary power units, or the requirements of § 25.1355(c). disconnection or failure of all power (c) Radio and electronic equipment, sources dependent on their continued controls, and wiring must be installed operation, for— so that operation of any one unit or (1) A time duration of at least 10 min- system of units will not adversely af- utes, including an aggregate time dura- fect the simultaneous operation of any tion of at least 5 minutes of announce- other radio or electronic unit, or sys- ments made by flight and cabin crew- tem of units, required by this chapter. members, considering all other loads (d) Electronic equipment must be de- which may remain powered by the signed and installed such that it does same source when all other power not cause essential loads to become in- sources are inoperative; and operative as a result of electrical (2) An additional time duration in its power supply transients or transients standby state appropriate or required from other causes. for any other loads that are powered by [Docket No. 5066, 29 FR 18291, Dec. 24, 1964, as the same source and that are essential amended by Amdt. 25–113, 69 FR 12530, Mar. to safety of flight or required during 16, 2004] emergency conditions. (b) Be capable of operation within 3 § 25.1433 Vacuum systems. seconds from the time a microphone is There must be means, in addition to removed from its stowage. the normal pressure relief, to auto- (c) Be intelligible at all passenger matically relieve the pressure in the seats, lavatories, and flight attendant discharge lines from the vacuum air seats and work stations. pump when the delivery temperature of (d) Be designed so that no unused, the air becomes unsafe. unstowed microphone will render the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as system inoperative. amended by Amdt. 25–72, 55 FR 29785, July 20, (e) Be capable of functioning inde- 1990] pendently of any required crewmember interphone system. § 25.1435 Hydraulic systems. (f) Be accessible for immediate use (a) Element design. Each element of from each of two flight crewmember the hydraulic system must be designed stations in the pilot compartment. to: (g) For each required floor-level pas- (1) Withstand the proof pressure senger emergency exit which has an ad- without permanent deformation that jacent flight attendant seat, have a would prevent it from performing its microphone which is readily accessible intended functions, and the ultimate to the seated flight attendant, except pressure without rupture. The proof that one microphone may serve more and ultimate pressures are defined in than one exit, provided the proximity terms of the design operating pressure of the exits allows unassisted verbal (DOP) as follows:

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plane manufacturer, which must be Element Proof Ultimate (xDOP) (xDOP) identified by appropriate markings as 1. Tubes and fittings...... 1.5 3.0 required by § 25.1541. 2. Pressure vessels containing gas: (c) Tests. Tests must be conducted on High pressure (e.g., accumulators) 3.0 4.0 the hydraulic system(s), and/or sub- Low pressure (e.g., reservoirs) ...... 1.5 3.0 system(s) and elements, except that 3. Hoses ...... 2.0 4.0 4. All other elements ...... 1.5 2.0 analysis may be used in place of or to supplement testing, where the analysis (2) Withstand, without deformation is shown to be reliable and appropriate. that would prevent it from performing All internal and external influences its intended function, the design oper- must be taken into account to an ex- ating pressure in combination with tent necessary to evaluate their ef- limit structural loads that may be im- fects, and to assure reliable system and posed; element functioning and integration. (3) Withstand, without rupture, the Failure or unacceptable deficiency of design operating pressure multiplied by an element or system must be cor- a factor of 1.5 in combination with ulti- rected and be sufficiently retested, mate structural load that can reason- where necessary. ably occur simultaneously; (1) The system(s), subsystem(s), or (4) Withstand the fatigue effects of element(s) must be subjected to per- all cyclic pressures, including tran- formance, fatigue, and endurance tests sients, and associated externally in- representative of airplane ground and duced loads, taking into account the flight operations. consequences of element failure; and (2) The complete system must be (5) Perform as intended under all en- tested to determine proper functional vironmental conditions for which the performance and relation to the other airplane is certificated. systems, including simulation of rel- (b) System design. Each hydraulic sys- evant failure conditions, and to sup- tem must: port or validate element design. (1) Have means located at a (3) The complete hydraulic system(s) flightcrew station to indicate appro- must be functionally tested on the air- priate system parameters, if plane in normal operation over the (i) It performs a function necessary range of motion of all associated user for continued safe flight and landing; systems. The test must be conducted at or the system relief pressure or 1.25 times (ii) In the event of hydraulic system the DOP if a system pressure relief de- malfunction, corrective action by the vice is not part of the system design. crew to ensure continued safe flight Clearances between hydraulic system and landing is necessary; elements and other systems or struc- (2) Have means to ensure that system tural elements must remain adequate pressures, including transient pres- and there must be no detrimental ef- sures and pressures from fluid volu- fects. metric changes in elements that are [Doc. No. 28617, 66 FR 27402, May 16, 2001] likely to remain closed long enough for such changes to occur, are within the § 25.1438 Pressurization and pneu- design capabilities of each element, matic systems. such that they meet the requirements (a) Pressurization system elements defined in § 25.1435(a)(1) through (a)(5); must be burst pressure tested to 2.0 (3) Have means to minimize the re- times, and proof pressure tested to 1.5 lease of harmful or hazardous con- times, the maximum normal operating centrations of hydraulic fluid or vapors pressure. into the crew and passenger compart- (b) Pneumatic system elements must ments during flight; be burst pressure tested to 3.0 times, (4) Meet the applicable requirements and proof pressure tested to 1.5 times, of §§ 25.863, 25.1183, 25.1185, and 25.1189 if the maximum normal operating pres- a flammable hydraulic fluid is used; sure. and (c) An analysis, or a combination of (5) Be designed to use any suitable analysis and test, may be substituted hydraulic fluid specified by the air- for any test required by paragraph (a)

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or (b) of this section if the Adminis- to the inside of the device and prevent trator finds it equivalent to the re- any outward leakage causing signifi- quired test. cant increase in the oxygen content of the local ambient atmosphere. If a de- [Amdt. 25–41, 42 FR 36971, July 18, 1977] mand oxygen system is used, a supply § 25.1439 Protective breathing equip- of 300 liters of free oxygen at 70 °F. and ment. 760 mm. Hg. pressure is considered to (a) Fixed (stationary, or built in) pro- be of 15-minute duration at the pre- tective breathing equipment must be scribed altitude and minute volume. If installed for the use of the flightcrew, a continuous flow open circuit protec- and at least one portable protective tive breathing system is used, a flow breathing equipment shall be located rate of 60 liters per minute at 8,000 feet at or near the flight deck for use by a (45 liters per minute at sea level) and a flight crewmember. In addition, port- supply of 600 liters of free oxygen at 70 ° able protective breathing equipment F. and 760 mm. Hg. pressure is consid- must be installed for the use of appro- ered to be of 15-minute duration at the priate crewmembers for fighting fires prescribed altitude and minute volume. in compartments accessible in flight Continuous flow systems must not in- other than the flight deck. This increase the ambient oxygen content of cludes isolated compartments and the local atmosphere above that of de- upper and lower lobe galleys, in which mand systems. BTPD refers to body ° crewmember occupancy is permitted temperature conditions (that is, 37 C., during flight. Equipment must be in- at ambient pressure, dry). stalled for the maximum number of (6) The equipment must meet the re- crewmembers expected to be in the quirements of § 25.1441. area during any operation. [Doc. No. FAA–2002–13859, 69 FR 40528, July 2, (b) For protective breathing equip- 2004] ment required by paragraph (a) of this section or by the applicable Operating § 25.1441 Oxygen equipment and sup- Regulations: ply. (1) The equipment must be designed (a) If certification with supplemental to protect the appropriate crewmember oxygen equipment is requested, the from smoke, carbon dioxide, and other equipment must meet the requirements harmful gases while on flight deck of this section and §§ 25.1443 through duty or while combating fires. 25.1453. (2) The equipment must include— (b) The oxygen system must be free (i) Masks covering the eyes, nose and from hazards in itself, in its method of mouth, or operation, and in its effect upon other (ii) Masks covering the nose and components. mouth, plus accessory equipment to (c) There must be a means to allow cover the eyes. the crew to readily determine, during (3) Equipment, including portable flight, the quantity of oxygen available equipment, must allow communication in each source of supply. with other crewmembers while in use. (d) The oxygen flow rate and the oxy- Equipment available at flightcrew as- gen equipment for airplanes for which signed duty stations must also enable certification for operation above 40,000 the flightcrew to use radio equipment. feet is requested must be approved. (4) The part of the equipment pro- tecting the eyes shall not cause any ap- § 25.1443 Minimum mass flow of sup- preciable adverse effect on vision and plemental oxygen. must allow corrective glasses to be (a) If continuous flow equipment is worn. installed for use by flight crew- (5) The equipment must supply pro- members, the minimum mass flow of tective oxygen of 15 minutes duration supplemental oxygen required for each per crewmember at a pressure altitude crewmember may not be less than the of 8,000 feet with a respiratory minute flow required to maintain, during in- volume of 30 liters per minute BTPD. spiration, a mean tracheal oxygen par- The equipment and system must be de- tial pressure of 149 mm. Hg. when signed to prevent any inward leakage breathing 15 liters per minute, BTPS,

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and with a maximum tidal volume of graph (a) or (b) of this section, which- 700 cc. with a constant time interval ever is applicable. between respirations. (b) If demand equipment is installed § 25.1445 Equipment standards for the for use by flight crewmembers, the oxygen distributing system. minimum mass flow of supplemental (a) When oxygen is supplied to both oxygen required for each crewmember crew and passengers, the distribution may not be less than the flow required system must be designed for either— to maintain, during inspiration, a (1) A source of supply for the flight mean tracheal oxygen partial pressure crew on duty and a separate source for of 122 mm. Hg., up to and including a the passengers and other crewmembers; cabin pressure altitude of 35,000 feet, or and 95 percent oxygen between cabin (2) A common source of supply with pressure altitudes of 35,000 and 40,000 means to separately reserve the min- feet, when breathing 20 liters per imum supply required by the flight minute BTPS. In addition, there must crew on duty. be means to allow the crew to use undi- (b) Portable walk-around oxygen luted oxygen at their discretion. units of the continuous flow, diluter- (c) For passengers and cabin attend- demand, and straight demand kinds ants, the minimum mass flow of sup- may be used to meet the crew or pas- plemental oxygen required for each senger breathing requirements. person at various cabin pressure altitudes may not be less than the flow re- § 25.1447 Equipment standards for ox- quired to maintain, during inspiration ygen dispensing units. and while using the oxygen equipment (including masks) provided, the fol- If oxygen dispensing units are in- lowing mean tracheal oxygen partial stalled, the following apply: pressures: (a) There must be an individual dis- (1) At cabin pressure altitudes above pensing unit for each occupant for 10,000 feet up to and including 18,500 whom supplemental oxygen is to be feet, a mean tracheal oxygen partial supplied. Units must be designed to pressure of 100 mm. Hg. when breathing cover the nose and mouth and must be 15 liters per minute, BTPS, and with a equipped with a suitable means to re- tidal volume of 700 cc. with a constant tain the unit in position on the face. time interval between respirations. Flight crew masks for supplemental (2) At cabin pressure altitudes above oxygen must have provisions for the 18,500 feet up to and including 40,000 use of communication equipment. feet, a mean tracheal oxygen partial (b) If certification for operation up to pressure of 83.8 mm. Hg. when breath- and including 25,000 feet is requested, ing 30 liters per minute, BTPS, and an oxygen supply terminal and unit of with a tidal volume of 1,100 cc. with a oxygen dispensing equipment for the constant time interval between res- immediate use of oxygen by each crew- pirations. member must be within easy reach of (d) If first-aid oxygen equipment is that crewmember. For any other occu- installed, the minimum mass flow of pants, the supply terminals and dis- oxygen to each user may not be less pensing equipment must be located to than four liters per minute, STPD. allow the use of oxygen as required by However, there may be a means to de- the operating rules in this chapter. crease this flow to not less than two li- (c) If certification for operation ters per minute, STPD, at any cabin al- above 25,000 feet is requested, there titude. The quantity of oxygen re- must be oxygen dispensing equipment quired is based upon an average flow meeting the following requirements: rate of three liters per minute per per- (1) There must be an oxygen dis- son for whom first-aid oxygen is re- pensing unit connected to oxygen sup- quired. ply terminals immediately available to (e) If portable oxygen equipment is each occupant, wherever seated, and at installed for use by crewmembers, the least two oxygen dispensing units con- minimum mass flow of supplemental nected to oxygen terminals in each lav- oxygen is the same as specified in para- atory. The total number of dispensing

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units and outlets in the cabin must ex- pensing unit connected to the portable ceed the number of seats by at least 10 oxygen supply. percent. The extra units must be as [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as uniformly distributed throughout the amended by Amdt. 25–41, 42 FR 36971, July 18, cabin as practicable. If certification for 1977; Amdt. 25–87, 61 FR 28696, June 5, 1996; operation above 30,000 feet is requested, Amdt. 25–116, 69 FR 62789, Oct. 27, 2004] the dispensing units providing the required oxygen flow must be automati- § 25.1449 Means for determining use of oxygen. cally presented to the occupants before the cabin pressure altitude exceeds There must be a means to allow the 15,000 feet. The crew must be provided crew to determine whether oxygen is with a manual means of making the being delivered to the dispensing equipment. dispensing units immediately available in the event of failure of the automatic § 25.1450 Chemical oxygen generators. system. (a) For the purpose of this section, a (2) Each flight crewmember on flight chemical oxygen generator is defined deck duty must be provided with a as a device which produces oxygen by quick-donning type oxygen dispensing chemical reaction. unit connected to an oxygen supply (b) Each chemical oxygen generator terminal. This dispensing unit must be must be designed and installed in ac- immediately available to the flight cordance with the following require- crewmember when seated at his sta- ments: tion, and installed so that it: (1) Surface temperature developed by (i) Can be placed on the face from its the generator during operation may ready position, properly secured, not create a hazard to the airplane or sealed, and supplying oxygen upon de- to its occupants. mand, with one hand, within five sec- (2) Means must be provided to relieve onds and without disturbing eyeglasses any internal pressure that may be haz- or causing delay in proceeding with ardous. emergency duties; and (c) In addition to meeting the re- (ii) Allows, while in place, the per- quirements in paragraph (b) of this sec- formance of normal communication tion, each portable chemical oxygen functions. generator that is capable of sustained operation by successive replacement of (3) The oxygen dispensing equipment a generator element must be placarded for the flight crewmembers must be: to show— (i) The diluter demand or pressure de- (1) The rate of oxygen flow, in liters mand (pressure demand mask with a per minute; diluter demand pressure breathing reg- (2) The duration of oxygen flow, in ulator) type, or other approved oxygen minutes, for the replaceable generator equipment shown to provide the same element; and degree of protection, for airplanes to be (3) A warning that the replaceable operated above 25,000 feet. generator element may be hot, unless (ii) The pressure demand (pressure the element construction is such that demand mask with a diluter demand the surface temperature cannot exceed pressure breathing regulator) type with 100 degrees F. mask-mounted regulator, or other ap- [Amdt. 25–41, 42 FR 36971, July 18, 1977] proved oxygen equipment shown to provide the same degree of protection, § 25.1453 Protection of oxygen equip- for airplanes operated at altitudes ment from rupture. where decompressions that are not ex- Oxygen pressure tanks, and lines be- tremely improbable may expose the tween tanks and the shutoff means, flightcrew to cabin pressure altitudes must be— in excess of 34,000 feet. (a) Protected from unsafe tempera- (4) Portable oxygen equipment must tures; and be immediately available for each (b) Located where the probability and cabin attendant. The portable oxygen hazards of rupture in a crash landing equipment must have the oxygen dis- are minimized.

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§ 25.1455 Draining of fluids subject to cockpit noise conditions and played freezing. back. Repeated aural or visual play- If fluids subject to freezing may be back of the record may be used in eval- drained overboard in flight or during uating intelligibility. ground operation, the drains must be (c) Each cockpit voice recorder must designed and located to prevent the be installed so that the part of the formation of hazardous quantities of communication or audio signals speci- ice on the airplane as a result of the fied in paragraph (a) of this section ob- drainage. tained from each of the following sources is recorded on a separate chan- [Amdt. 25–23, 35 FR 5680, Apr. 8, 1970] nel: (1) For the first channel, from each § 25.1457 Cockpit voice recorders. boom, mask, or hand-held microphone, (a) Each cockpit voice recorder re- headset, or speaker used at the first quired by the operating rules of this pilot station. chapter must be approved and must be (2) For the second channel from each installed so that it will record the fol- boom, mask, or hand-held microphone, lowing: headset, or speaker used at the second (1) Voice communications trans- pilot station. mitted from or received in the airplane (3) For the third channel—from the by radio. cockpit-mounted area microphone. (2) Voice communications of flight (4) For the fourth channel, from— crewmembers on the flight deck. (3) Voice communications of flight (i) Each boom, mask, or hand-held crewmembers on the flight deck, using microphone, headset, or speaker used the airplane’s interphone system. at the station for the third and fourth (4) Voice or audio signals identifying crew members; or navigation or approach aids introduced (ii) If the stations specified in para- into a headset or speaker. graph (c)(4)(i) of this section are not re- (5) Voice communications of flight quired or if the signal at such a station crewmembers using the passenger loud- is picked up by another channel, each speaker system, if there is such a sys- microphone on the flight deck that is tem and if the fourth channel is avail- used with the passenger loudspeaker able in accordance with the require- system, if its signals are not picked up ments of paragraph (c)(4)(ii) of this sec- by another channel. tion. (5) As far as is practicable all sounds (6) If datalink communication equip- received by the microphone listed in ment is installed, all datalink commu- paragraphs (c)(1), (2), and (4) of this nications, using an approved data mes- section must be recorded without sage set. Datalink messages must be interruption irrespective of the posi- recorded as the output signal from the tion of the interphone-transmitter key communications unit that translates switch. The design shall ensure that the signal into usable data. sidetone for the flight crew is produced (b) The recording requirements of only when the interphone, public ad- paragraph (a)(2) of this section must be dress system, or radio transmitters are met by installing a cockpit-mounted in use. area microphone, located in the best (d) Each cockpit voice recorder must position for recording voice commu- be installed so that— nications originating at the first and (1)(i) It receives its electrical power second pilot stations and voice commu- from the bus that provides the max- nications of other crewmembers on the imum reliability for operation of the flight deck when directed to those sta- cockpit voice recorder without jeopard- tions. The microphone must be so lo- izing service to essential or emergency cated and, if necessary, the pre- loads. amplifiers and filters of the recorder (ii) It remains powered for as long as must be so adjusted or supplemented, possible without jeopardizing emer- that the intelligibility of the recorded gency operation of the airplane. communications is as high as prac- (2) There is an automatic means to ticable when recorded under flight simultaneously stop the recorder and

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prevent each erasure feature from func- ability of inadvertent operation and ac- tioning, within 10 minutes after crash tuation of the device during crash im- impact; pact. (3) There is an aural or visual means (g) Each recorder container must— for preflight checking of the recorder (1) Be either bright orange or bright for proper operation; yellow; (4) Any single electrical failure exter- (2) Have reflective tape affixed to its nal to the recorder does not disable external surface to facilitate its loca- both the cockpit voice recorder and the tion under water; and flight data recorder; (3) Have an underwater locating de- (5) It has an independent power vice, when required by the operating source— rules of this chapter, on or adjacent to (i) That provides 10 ± 1 minutes of the container which is secured in such electrical power to operate both the manner that they are not likely to be cockpit voice recorder and cockpit- separated during crash impact. mounted area microphone; [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (ii) That is located as close as prac- amended by Amdt. 25–2, 30 FR 3932, Mar. 26, ticable to the cockpit voice recorder; 1965; Amdt. 25–16, 32 FR 13914, Oct. 6, 1967; and Amdt. 25–41, 42 FR 36971, July 18, 1977; Amdt. (iii) To which the cockpit voice re- 25–65, 53 FR 26143, July 11, 1988; Amdt. 25–124, corder and cockpit-mounted area 73 FR 12563, Mar. 7, 2008; 74 FR 32800, July 9, microphone are switched automati- 2009] cally in the event that all other power to the cockpit voice recorder is inter- § 25.1459 Flight data recorders. rupted either by normal shutdown or (a) Each flight recorder required by by any other loss of power to the elec- the operating rules of this chapter trical power bus; and must be installed so that— (6) It is in a separate container from (1) It is supplied with airspeed, alti- the flight data recorder when both are tude, and directional data obtained required. If used to comply with only from sources that meet the accuracy the cockpit voice recorder require- requirements of §§ 25.1323, 25.1325, and ments, a combination unit may be in- 25.1327, as appropriate; stalled. (2) The vertical acceleration sensor is (e) The recorder container must be rigidly attached, and located longitu- located and mounted to minimize the dinally either within the approved cen- probability of rupture of the container ter of gravity limits of the airplane, or as a result of crash impact and con- at a distance forward or aft of these sequent heat damage to the recorder limits that does not exceed 25 percent from fire. of the airplane’s mean aerodynamic (1) Except as provided in paragraph chord; (e)(2) of this section, the recorder con- (3)(i) It receives its electrical power tainer must be located as far aft as from the bus that provides the max- practicable, but need not be outside of imum reliability for operation of the the pressurized compartment, and may flight data recorder without jeopard- not be located where aft-mounted en- izing service to essential or emergency gines may crush the container during loads. impact. (ii) It remains powered for as long as (2) If two separate combination dig- possible without jeopardizing emer- ital flight data recorder and cockpit gency operation of the airplane. voice recorder units are installed in- (4) There is an aural or visual means stead of one cockpit voice recorder and for preflight checking of the recorder one digital flight data recorder, the for proper recording of data in the stor- combination unit that is installed to age medium; comply with the cockpit voice recorder (5) Except for recorders powered sole- requirements may be located near the ly by the engine-driven electrical gen- cockpit. erator system, there is an automatic (f) If the cockpit voice recorder has a means to simultaneously stop a re- bulk erasure device, the installation corder that has a data erasure feature must be designed to minimize the prob- and prevent each erasure feature from

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functioning, within 10 minutes after any dedicated parameters must be re- crash impact; corded on flight recorders in addition (6) There is a means to record data to or in place of existing requirements. from which the time of each radio [Amdt. 25–8, 31 FR 127, Jan. 6, 1966, as amend- transmission either to or from ATC can ed by Amdt. 25–25, 35 FR 13192, Aug. 19, 1970; be determined; Amdt. 25–37, 40 FR 2577, Jan. 14, 1975; Amdt. (7) Any single electrical failure exter- 25–41, 42 FR 36971, July 18, 1977; Amdt. 25–65, nal to the recorder does not disable 53 FR 26144, July 11, 1988; Amdt. 25–124, 73 FR both the cockpit voice recorder and the 12563, Mar. 7, 2008; 74 FR 32800, July 9, 2009] flight data recorder; and (8) It is in a separate container from § 25.1461 Equipment containing high the cockpit voice recorder when both energy rotors. are required. If used to comply with (a) Equipment containing high en- only the flight data recorder require- ergy rotors must meet paragraph (b), ments, a combination unit may be in- (c), or (d) of this section. stalled. If a combination unit is in- (b) High energy rotors contained in stalled as a cockpit voice recorder to equipment must be able to withstand comply with § 25.1457(e)(2), a combina- damage caused by malfunctions, vibration unit must be used to comply with tion, abnormal speeds, and abnormal this flight data recorder requirement. temperatures. In addition— (b) Each nonejectable record con- (1) Auxiliary rotor cases must be able tainer must be located and mounted so to contain damage caused by the fail- as to minimize the probability of con- ure of high energy rotor blades; and tainer rupture resulting from crash im- (2) Equipment control devices, sys- pact and subsequent damage to the tems, and instrumentation must rea- record from fire. In meeting this re- sonably ensure that no operating limi- quirement the record container must tations affecting the integrity of high be located as far aft as practicable, but energy rotors will be exceeded in serv- need not be aft of the pressurized com- ice. partment, and may not be where aft- (c) It must be shown by test that mounted engines may crush the con- equipment containing high energy ro- tainer upon impact. tors can contain any failure of a high (c) A correlation must be established energy rotor that occurs at the highest between the flight recorder readings of speed obtainable with the normal speed airspeed, altitude, and heading and the control devices inoperative. corresponding readings (taking into ac- (d) Equipment containing high en- count correction factors) of the first pi- ergy rotors must be located where lot’s instruments. The correlation rotor failure will neither endanger the must cover the airspeed range over occupants nor adversely affect contin- which the airplane is to be operated, ued safe flight. the range of altitude to which the airplane is limited, and 360 degrees of [Amdt. 25–41, 42 FR 36971, July 18, 1977] heading. Correlation may be established on the ground as appropriate. Subpart G—Operating Limitations (d) Each recorder container must— and Information (1) Be either bright orange or bright yellow; § 25.1501 General. (2) Have reflective tape affixed to its (a) Each operating limitation speci- external surface to facilitate its loca- fied in §§ 25.1503 through 25.1533 and tion under water; and other limitations and information nec- (3) Have an underwater locating de- essary for safe operation must be es- vice, when required by the operating tablished. rules of this chapter, on or adjacent to (b) The operating limitations and the container which is secured in such other information necessary for safe a manner that they are not likely to be operation must be made available to separated during crash impact. the crewmembers as prescribed in (e) Any novel or unique design or §§ 25.1541 through 25.1587. operational characteristics of the aircraft shall be evaluated to determine if [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978]

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OPERATING LIMITATIONS gear, as determined under § 25.729 or by flight characteristics. If the extension § 25.1503 Airspeed limitations: general. speed is not the same as the retraction When airspeed limitations are a func- speed, the two speeds must be des- tion of weight, weight distribution, al- ignated as VLO(EXT) and VLO(RET), respec- titude, or Mach number, limitations tively. corresponding to each critical com- (b) The established landing gear ex- bination of these factors must be estab- tended speed VLE may not exceed the lished. speed at which it is safe to fly with the landing gear secured in the fully ex- § 25.1505 Maximum operating limit tended position, and that determined speed. under § 25.729. The maximum operating limit speed [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (VMO/MMO airspeed or Mach Number, amended by Amdt. 25–38, 41 FR 55468, Dec. 20, whichever is critical at a particular al- 1976] titude) is a speed that may not be de- liberately exceeded in any regime of § 25.1516 Other speed limitations. flight (climb, cruise, or descent), unless Any other limitation associated with a higher speed is authorized for flight speed must be established. test or pilot training operations. V / MO [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] MMO must be established so that it is not greater than the design cruising § 25.1517 Rough air speed, VRA. speed V and so that it is sufficiently C A rough air speed, V , for use as the below V /M or V /M to make it RA D D DF DF, recommended turbulence penetration highly improbable that the latter airspeed in § 25.1585(a)(8), must be es- speeds will be inadvertently exceeded tablished, which— in operations. The speed margin be- (1) Is not greater than the design air- tween V /M and V /M or V M/ MO MO D D DF DF speed for maximum gust intensity, se- may not be less than that determined lected for V ; and under § 25.335(b) or found necessary dur- B (2) Is not less than the minimum ing the flight tests conducted under value of V specified in § 25.335(d); and § 25.253. B (3) Is sufficiently less than VMO to en- [Amdt. 25–23, 35 FR 5680, Apr. 8, 1970] sure that likely speed variation during rough air encounters will not cause the § 25.1507 Maneuvering speed. overspeed warning to operate too fre- The maneuvering speed must be es- quently. In the absence of a rational tablished so that it does not exceed the investigation substantiating the use of design maneuvering speed VA deter- other values, VRA must be less than mined under § 25.335(c). VMO—35 knots (TAS). [Doc. No. 27902, 61 FR 5222, Feb. 9, 1996] § 25.1511 Flap extended speed. The established flap extended speed § 25.1519 Weight, center of gravity, and VFE must be established so that it does weight distribution. not exceed the design flap speed VF The airplane weight, center of grav- chosen under §§ 25.335(e) and 25.345, for ity, and weight distribution limita- the corresponding flap positions and tions determined under §§ 25.23 through engine powers. 25.27 must be established as operating limitations. § 25.1513 Minimum control speed.

The minimum control speed VMC de- § 25.1521 Powerplant limitations. termined under § 25.149 must be estab- (a) General. The powerplant limita- lished as an operating limitation. tions prescribed in this section must be established so that they do not exceed § 25.1515 Landing gear speeds. the corresponding limits for which the (a) The established landing gear oper- engines or propellers are type certifi- ating speed or speeds, VLO, may not ex- cated and do not exceed the values on ceed the speed at which it is safe both which compliance with any other re- to extend and to retract the landing quirement of this part is based.

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(b) Reciprocating engine installations. § 25.1522 Auxiliary power unit limita- Operating limitations relating to the tions. following must be established for recip- If an auxiliary power unit is installed rocating engine installations: in the airplane, limitations established (1) Horsepower or torque, r.p.m., for the auxiliary power unit, including manifold pressure, and time at critical categories of operation, must be speci- pressure altitude and sea level pressure fied as operating limitations for the altitude for— airplane. (i) Maximum continuous power (relating to unsupercharged operation or [Amdt. 25–72, 55 FR 29786, July 20, 1990] to operation in each supercharger mode § 25.1523 Minimum flight crew. as applicable); and (ii) Takeoff power (relating to unsu- The minimum flight crew must be es- percharged operation or to operation in tablished so that it is sufficient for safe each supercharger mode as applicable). operation, considering— (2) Fuel grade or specification. (a) The workload on individual crew- (3) Cylinder head and oil tempera- members; tures. (b) The accessibility and ease of oper- (4) Any other parameter for which a ation of necessary controls by the ap- limitation has been established as part propriate crewmember; and of the engine type certificate except (c) The kind of operation authorized that a limitation need not be estab- under § 25.1525. lished for a parameter that cannot be The criteria used in making the deter- exceeded during normal operation due minations required by this section are to the design of the installation or to set forth in appendix D. another established limitation. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) Turbine engine installations. Oper- amended by Amdt. 25–3, 30 FR 6067, Apr. 29, ating limitations relating to the fol- 1965] lowing must be established for turbine engine installations: § 25.1525 Kinds of operation. (1) Horsepower, torque or thrust, The kinds of operation to which the r.p.m., gas temperature, and time for— airplane is limited are established by (i) Maximum continuous power or the category in which it is eligible for thrust (relating to augmented or un- certification and by the installed augmented operation as applicable). equipment. (ii) Takeoff power or thrust (relating to augmented or unaugmented oper- § 25.1527 Ambient air temperature and operating altitude. ation as applicable). (2) Fuel designation or specification. The extremes of the ambient air tem- (3) Any other parameter for which a perature and operating altitude for limitation has been established as part which operation is allowed, as limited of the engine type certificate except by flight, structural, powerplant, func- that a limitation need not be estab- tional, or equipment characteristics, lished for a parameter that cannot be must be established. exceeded during normal operation due [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] to the design of the installation or to another established limitation. § 25.1529 Instructions for Continued (d) Ambient temperature. An ambient Airworthiness. temperature limitation (including lim- The applicant must prepare Instruc- itations for winterization installations, tions for Continued Airworthiness in if applicable) must be established as accordance with appendix H to this the maximum ambient atmospheric part that are acceptable to the Admin- temperature established in accordance istrator. The instructions may be in- with § 25.1043(b). complete at type certification if a program exists to ensure their completion [Amdt. 25–72, 55 FR 29786, July 20, 1990] prior to delivery of the first airplane or

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issuance of a standard certificate of § 25.1535 ETOPS approval. airworthiness, whichever occurs later. Except as provided in § 25.3, each ap- [Amdt. 25–54, 45 FR 60173, Sept. 11, 1980] plicant seeking ETOPS type design approval must comply with the provi- § 25.1531 Maneuvering flight load fac- sions of Appendix K of this part. tors. [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, Load factor limitations, not exceed- 2007] ing the positive limit load factors determined from the maneuvering dia- MARKINGS AND PLACARDS gram in § 25.333(b), must be established. § 25.1541 General. § 25.1533 Additional operating limita- (a) The airplane must contain— tions. (1) The specified markings and plac- (a) Additional operating limitations ards; and must be established as follows: (2) Any additional information, in- (1) The maximum takeoff weights strument markings, and placards re- must be established as the weights at quired for the safe operation if there which compliance is shown with the are unusual design, operating, or han- applicable provisions of this part (in- dling characteristics. cluding the takeoff climb provisions of (b) Each marking and placard pre- § 25.121(a) through (c), for altitudes and scribed in paragraph (a) of this sec- ambient temperatures). tion— (2) The maximum landing weights (1) Must be displayed in a con- must be established as the weights at spicuous place; and which compliance is shown with the (2) May not be easily erased, dis- applicable provisions of this part (in- figured, or obscured. cluding the landing and approach climb provisions of §§ 25.119 and 25.121(d) for § 25.1543 Instrument markings: gen- altitudes and ambient temperatures). eral. (3) The minimum takeoff distances For each instrument— must be established as the distances at (a) When markings are on the cover which compliance is shown with the glass of the instrument, there must be applicable provisions of this part (in- means to maintain the correct align- cluding the provisions of §§ 25.109 and ment of the glass cover with the face of 25.113, for weights, altitudes, tempera- the dial; and tures, wind components, runway surface conditions (dry and wet), and run- (b) Each instrument marking must way gradients) for smooth, hard-sur- be clearly visible to the appropriate faced runways. Additionally, at the op- crewmember. tion of the applicant, wet runway take- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as off distances may be established for amended by Amdt. 25–72, 55 FR 29786, July 20, runway surfaces that have been 1990] grooved or treated with a porous friction course, and may be approved for § 25.1545 Airspeed limitation informa- use on runways where such surfaces tion. have been designed constructed, and The airspeed limitations required by maintained in a manner acceptable to § 25.1583 (a) must be easily read and un- the Administrator. derstood by the flight crew. (b) The extremes for variable factors (such as altitude, temperature, wind, § 25.1547 Magnetic direction indicator. and runway gradients) are those at (a) A placard meeting the require- which compliance with the applicable ments of this section must be installed provisions of this part is shown. on, or near, the magnetic direction in- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as dicator. amended by Amdt. 25–38, 41 FR 55468, Dec. 20, (b) The placard must show the cali- 1976; Amdt. 25–72, 55 FR 29786, July 20, 1990; bration of the instrument in level Amdt. 25–92, 63 FR 8321, Feb. 18, 1998] flight with the engines operating.

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(c) The placard must state whether (2) If safe operation requires the use the calibration was made with radio re- of any tanks in a specific sequence, ceivers on or off. that sequence must be marked on, or (d) Each calibration reading must be adjacent to, the selector for those in terms of magnetic heading in not tanks; and more than 45 degree increments. (3) Each valve control for each engine must be marked to indicate the posi- § 25.1549 Powerplant and auxiliary tion corresponding to each engine con- power unit instruments. trolled. For each required powerplant and (d) For accessory, auxiliary, and auxiliary power unit instrument, as ap- emergency controls— propriate to the type of instrument— (1) Each emergency control (includ- (a) Each maximum and, if applicable, ing each fuel jettisoning and fluid shut- minimum safe operating limit must be off must be colored red; and marked with a red radial or a red line; (2) Each visual indicator required by (b) Each normal operating range § 25.729(e) must be marked so that the must be marked with a green arc or pilot can determine at any time when green line, not extending beyond the the wheels are locked in either extreme maximum and minimum safe limits; position, if retractable landing gear is (c) Each takeoff and precautionary used. range must be marked with a yellow arc or a yellow line; and § 25.1557 Miscellaneous markings and (d) Each engine, auxiliary power placards. unit, or propeller speed range that is (a) Baggage and cargo compartments restricted because of excessive vibra- and ballast location. Each baggage and tion stresses must be marked with red cargo compartment, and each ballast arcs or red lines. location must have a placard stating [Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] any limitations on contents, including weight, that are necessary under the § 25.1551 Oil quantity indication. loading requirements. However, Each oil quantity indicating means underseat compartments designed for must be marked to indicate the quan- the storage of carry-on articles weigh- tity of oil readily and accurately. ing not more than 20 pounds need not have a loading limitation placard. [Amdt. 25–72, 55 FR 29786, July 20, 1990] (b) Powerplant fluid filler openings. The following apply: § 25.1553 Fuel quantity indicator. (1) Fuel filler openings must be If the unusable fuel supply for any marked at or near the filler cover tank exceeds one gallon, or five per- with— cent of the tank capacity, whichever is (i) The word ‘‘fuel’’; greater, a red arc must be marked on (ii) For reciprocating engine powered its indicator extending from the cali- airplanes, the minimum fuel grade; brated zero reading to the lowest read- (iii) For turbine engine powered air- ing obtainable in level flight. planes, the permissible fuel designa- tions; and § 25.1555 Control markings. (iv) For pressure fueling systems, the (a) Each cockpit control, other than maximum permissible fueling supply primary flight controls and controls pressure and the maximum permissible whose function is obvious, must be defueling pressure. plainly marked as to its function and (2) Oil filler openings must be method of operation. marked at or near the filler cover with (b) Each aerodynamic control must the word ‘‘oil’’. be marked under the requirements of (3) Augmentation fluid filler open- §§ 25.677 and 25.699. ings must be marked at or near the (c) For powerplant fuel controls— filler cover to identify the required (1) Each fuel tank selector control fluid. must be marked to indicate the posi- (c) Emergency exit placards. Each tion corresponding to each tank and to emergency exit placard must meet the each existing cross feed position; requirements of § 25.811.

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(d) Doors. Each door that must be ble noise standards of part 36 of this used in order to reach any required chapter. emergency exit must have a suitable (b) Approved information. Each part of placard stating that the door is to be the manual listed in §§ 25.1583 through latched in the open position during 25.1587, that is appropriate to the air- takeoff and landing. plane, must be furnished, verified, and approved, and must be segregated, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–32, 37 FR 3972, Feb. 24, identified, and clearly distinguished 1972; Amdt. 25–38, 41 FR 55468, Dec. 20, 1976; from each unapproved part of that Amdt. 25–72, 55 FR 29786, July 20, 1990] manual. (c) [Reserved] § 25.1561 Safety equipment. (d) Each Airplane Flight Manual (a) Each safety equipment control to must include a table of contents if the be operated by the crew in emergency, complexity of the manual indicates a such as controls for automatic liferaft need for it. releases, must be plainly marked as to [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978, as its method of operation. amended by Amdt. 25–72, 55 FR 29786, July 20, (b) Each location, such as a locker or 1990] compartment, that carries any fire extinguishing, signaling, or other life § 25.1583 Operating limitations. saving equipment must be marked ac- (a) Airspeed limitations. The following cordingly. airspeed limitations and any other air- (c) Stowage provisions for required speed limitations necessary for safe op- emergency equipment must be con- eration must be furnished: spicuously marked to identify the con- (1) The maximum operating limit tents and facilitate the easy removal of speed VMO/MMO and a statement that the equipment. this speed limit may not be delib- (d) Each liferaft must have obviously erately exceeded in any regime of marked operating instructions. flight (climb, cruise, or descent) unless (e) Approved survival equipment a higher speed is authorized for flight must be marked for identification and test or pilot training. method of operation. (2) If an airspeed limitation is based [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as upon compressibility effects, a state- amended by Amdt. 25–46, 43 FR 50598, Oct. 30, ment to this effect and information as 1978] to any symptoms, the probable behavior of the airplane, and the rec- § 25.1563 Airspeed placard. ommended recovery procedures. A placard showing the maximum air- (3) The maneuvering speed estab- speeds for flap extension for the take- lished under § 25.1507 and statements, off, approach, and landing positions as applicable to the particular design, must be installed in clear view of each explaining that: pilot. (i) Full application of pitch, roll, or yaw controls should be confined to AIRPLANE FLIGHT MANUAL speeds below the maneuvering speed; and § 25.1581 General. (ii) Rapid and large alternating con- (a) Furnishing information. An Air- trol inputs, especially in combination plane Flight Manual must be furnished with large changes in pitch, roll, or with each airplane, and it must contain yaw, and full control inputs in more the following: than one axis at the same time, should (1) Information required by §§ 25.1583 be avoided as they may result in struc- through 25.1587. tural failures at any speed, including (2) Other information that is nec- below the maneuvering speed. essary for safe operation because of de- (4) The flap extended speed VFE and sign, operating, or handling character- the pertinent flap positions and engine istics. powers. (3) Any limitation, procedure, or (5) The landing gear operating speed other information established as a con- or speeds, and a statement explaining dition of compliance with the applica- the speeds as defined in § 25.1515(a).

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(6) The landing gear extended speed described in terms of accelerations, VLE, if greater than VLO, and a state- must be furnished. ment that this is the maximum speed at which the airplane can be safely [Doc. No. 5066, 29 FR 1891, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55468, Dec, 20, flown with the landing gear extended. 1976; Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; (b) Powerplant limitations. The fol- Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; Amdt. lowing information must be furnished: 25–72, 55 FR 29787, July 20, 1990; Amdt. 25–105, (1) Limitations required by § 25.1521 66 FR 34024, June 26, 2001; 75 FR 49818, Aug. and § 25.1522. 16, 2010] (2) Explanation of the limitations, when appropriate. § 25.1585 Operating procedures. (3) Information necessary for mark- (a) Operating procedures must be fur- ing the instruments required by nished for— §§ 25.1549 through 25.1553. (c) Weight and loading distribution. (1) Normal procedures peculiar to the The weight and center of gravity limi- particular type or model encountered tations established under § 25.1519 must in connection with routine operations; be furnished in the Airplane Flight (2) Non-normal procedures for mal- Manual. All of the following informa- function cases and failure conditions tion, including the weight distribution involving the use of special systems or limitations established under § 25.1519, the alternative use of regular systems; must be presented either in the Air- and plane Flight Manual or in a separate (3) Emergency procedures for foresee- weight and balance control and loading able but unusual situations in which document that is incorporated by ref- immediate and precise action by the erence in the Airplane Flight Manual: crew may be expected to substantially (1) The condition of the airplane and reduce the risk of catastrophe. the items included in the empty weight (b) Information or procedures not di- as defined in accordance with § 25.29. rectly related to airworthiness or not (2) Loading instructions necessary to under the control of the crew, must not ensure loading of the airplane within be included, nor must any procedure the weight and center of gravity limits, that is accepted as basic airmanship. and to maintain the loading within (c) Information identifying each op- these limits in flight. (3) If certification for more than one erating condition in which the fuel sys- center of gravity range is requested, tem independence prescribed in § 25.953 the appropriate limitations, with re- is necessary for safety must be fur- gard to weight and loading procedures, nished, together with instructions for for each separate center of gravity placing the fuel system in a configura- range. tion used to show compliance with that (d) Flight crew. The number and func- section. tions of the minimum flight crew de- (d) The buffet onset envelopes, deter- termined under § 25.1523 must be fur- mined under § 25.251 must be furnished. nished. The buffet onset envelopes presented (e) Kinds of operation. The kinds of may reflect the center of gravity at operation approved under § 25.1525 must which the airplane is normally loaded be furnished. during cruise if corrections for the ef- (f) Ambient air temperatures and oper- fect of different center of gravity loca- ating altitudes. The extremes of the am- tions are furnished. bient air temperatures and operating (e) Information must be furnished altitudes established under § 25.1527 that indicates that when the fuel quan- must be furnished. tity indicator reads ‘‘zero’’ in level (g) [Reserved] flight, any fuel remaining in the fuel (h) Additional operating limitations. tank cannot be used safely in flight. The operating limitations established (f) Information on the total quantity under § 25.1533 must be furnished. of usable fuel for each fuel tank must (i) Maneuvering flight load factors. The be furnished. positive maneuvering limit load factors for which the structure is proven, [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001]

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§ 25.1587 Performance information. Subpart H—Electrical Wiring (a) Each Airplane Flight Manual Interconnection Systems (EWIS) must contain information to permit conversion of the indicated tempera- SOURCE: Docket No. FAA–2004–18379, 72 FR ture to free air temperature if other 63406, Nov. 8, 2007, unless otherwise noted. than a free air temperature indicator is used to comply with the requirements § 25.1701 Definition. of § 25.1303(a)(1). (a) As used in this chapter, electrical (b) Each Airplane Flight Manual wiring interconnection system (EWIS) must contain the performance informa- means any wire, wiring device, or com- tion computed under the applicable bination of these, including termi- provisions of this part (including nation devices, installed in any area of §§ 25.115, 25.123, and 25.125 for the the airplane for the purpose of trans- weights, altitudes, temperatures, wind mitting electrical energy, including components, and runway gradients, as data and signals, between two or more applicable) within the operational lim- intended termination points. This in- its of the airplane, and must contain cludes: the following: (1) Wires and cables. (2) Bus bars. (1) In each case, the conditions of (3) The termination point on elec- power, configuration, and speeds, and trical devices, including those on re- the procedures for handling the air- lays, interrupters, switches, plane and any system having a signifi- contactors, terminal blocks and circuit cant effect on the performance infor- breakers, and other circuit protection mation. devices. (2) VSR determined in accordance (4) Connectors, including feed- with § 25.103. through connectors. (3) The following performance infor- (5) Connector accessories. mation (determined by extrapolation (6) Electrical grounding and bonding and computed for the range of weights devices and their associated connec- between the maximum landing weight tions. and the maximum takeoff weight): (7) Electrical splices. (i) Climb in the landing configura- (8) Materials used to provide addition. tional protection for wires, including (ii) Climb in the approach configura- wire insulation, wire sleeving, and con- tion. duits that have electrical termination (iii) Landing distance. for the purpose of bonding. (4) Procedures established under (9) Shields or braids. § 25.101(f) and (g) that are related to the (10) Clamps and other devices used to limitations and information required route and support the wire bundle. by § 25.1533 and by this paragraph (b) in (11) Cable tie devices. the form of guidance material, includ- (12) Labels or other means of identi- ing any relevant limitations or infor- fication. mation. (13) Pressure seals. (5) An explanation of significant or (14) EWIS components inside shelves, unusual flight or ground handling char- panels, racks, junction boxes, distribu- acteristics of the airplane. tion panels, and back-planes of equipment racks, including, but not limited (6) Corrections to indicated values of to, circuit board back-planes, wire in- airspeed, altitude, and outside air tem- tegration units, and external wiring of perature. equipment. (7) An explanation of operational (b) Except for the equipment indi- landing runway length factors included cated in paragraph (a)(14) of this sec- in the presentation of the landing dis- tion, EWIS components inside the fol- tance, if appropriate. lowing equipment, and the external [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001, connectors that are part of that equip- as amended by Amdt. 25–108, 67 FR 70828, ment, are excluded from the definition Nov. 26, 2002] in paragraph (a) of this section:

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(1) Electrical equipment or avionics the applicable requirements for that that are qualified to environmental system. conditions and testing procedures when (1) § 25.773(b)(2) Pilot compartment those conditions and procedures are— view. (i) Appropriate for the intended func- (2) § 25.981 Fuel tank ignition pre- tion and operating environment, and vention. (ii) Acceptable to the FAA. (3) § 25.1165 Engine ignition systems. (2) Portable electrical devices that (4) § 25.1310 Power source capacity are not part of the type design of the and distribution. airplane. This includes personal enter- (5) § 25.1316 System lightning protec- tainment devices and laptop com- tion. puters. (6) § 25.1331(a)(2) Instruments using a (3) Fiber optics. power supply. (7) § 25.1351 General. § 25.1703 Function and installation: (8) § 25.1355 Distribution system. EWIS. (9) § 25.1360 Precautions against in- (a) Each EWIS component installed jury. in any area of the aircraft must: (10) § 25.1362 Electrical supplies for (1) Be of a kind and design appro- emergency conditions. priate to its intended function. (11) § 25.1365 Electrical appliances, (2) Be installed according to limita- motors, and transformers. tions specified for the EWIS compo- (12) § 25.1431(c) and (d) Electronic nents. equipment. (3) Perform the function for which it was intended without degrading the § 25.1707 System separation: EWIS. airworthiness of the airplane. (a) Each EWIS must be designed and (4) Be designed and installed in a way installed with adequate physical sepa- that will minimize mechanical strain. ration from other EWIS and airplane (b) Selection of wires must take into systems so that an EWIS component account known characteristics of the failure will not create a hazardous con- wire in relation to each installation dition. Unless otherwise stated, for the and application to minimize the risk of purposes of this section, adequate wire damage, including any arc track- physical separation must be achieved ing phenomena. by separation distance or by a barrier (c) The design and installation of the that provides protection equivalent to main power cables (including generator that separation distance. cables) in the fuselage must allow for a (b) Each EWIS must be designed and reasonable degree of deformation and installed so that any electrical inter- stretching without failure. ference likely to be present in the air- (d) EWIS components located in plane will not result in hazardous ef- areas of known moisture accumulation fects upon the airplane or its systems. must be protected to minimize any (c) Wires and cables carrying heavy hazardous effects due to moisture. current, and their associated EWIS components, must be designed and in- § 25.1705 Systems and functions: EWIS. stalled to ensure adequate physical (a) EWIS associated with any system separation and electrical isolation so required for type certification or by op- that damage to circuits associated erating rules must be considered an in- with essential functions will be mini- tegral part of that system and must be mized under fault conditions. considered in showing compliance with (d) Each EWIS associated with inde- the applicable requirements for that pendent airplane power sources or system. power sources connected in combina- (b) For systems to which the fol- tion must be designed and installed to lowing rules apply, the components of ensure adequate physical separation EWIS associated with those systems and electrical isolation so that a fault must be considered an integral part of in any one airplane power source EWIS that system or systems and must be will not adversely affect any other considered in showing compliance with independent power sources. In addition:

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(1) Airplane independent electrical (1) Chafing, jamming, or other inter- power sources must not share a com- ference are prevented. mon ground terminating location. (2) An EWIS component failure will (2) Airplane system static grounds not create a hazardous condition. must not share a common ground ter- (3) Failure of any flight or other me- minating location with any of the air- chanical control systems cables or sys- plane’s independent electrical power tems components will not damage the sources. EWIS and create a hazardous condi- (e) Except to the extent necessary to tion. provide electrical connection to the (j) EWIS must be designed and in- fuel systems components, the EWIS stalled with adequate physical separa- must be designed and installed with tion between the EWIS components adequate physical separation from fuel and heated equipment, hot air ducts, lines and other fuel system compo- and lines, so that: nents, so that: (1) An EWIS component failure will (1) An EWIS component failure will not create a hazardous condition. not create a hazardous condition. (2) Any hot air leakage or heat gen- (2) Any fuel leakage onto EWIS com- erated onto EWIS components will not ponents will not create a hazardous create a hazardous condition. condition. (k) For systems for which redun- (f) Except to the extent necessary to dancy is required, by certification provide electrical connection to the rules, by operating rules, or as a result hydraulic systems components, EWIS of the assessment required by § 25.1709, must be designed and installed with EWIS components associated with adequate physical separation from hy- those systems must be designed and in- draulic lines and other hydraulic sys- stalled with adequate physical separa- tem components, so that: tion. (1) An EWIS component failure will (l) Each EWIS must be designed and not create a hazardous condition. installed so there is adequate physical (2) Any hydraulic fluid leakage onto separation between it and other air- EWIS components will not create a craft components and aircraft struc- hazardous condition. ture, and so that the EWIS is protected (g) Except to the extent necessary to from sharp edges and corners, to mini- provide electrical connection to the ox- mize potential for abrasion/chafing, vi- ygen systems components, EWIS must bration damage, and other types of me- be designed and installed with ade- chanical damage. quate physical separation from oxygen lines and other oxygen system compo- § 25.1709 System safety: EWIS. nents, so that an EWIS component failure will not create a hazardous condi- Each EWIS must be designed and in- tion. stalled so that: (h) Except to the extent necessary to (a) Each catastrophic failure condi- provide electrical connection to the tion— water/waste systems components, (1) Is extremely improbable; and EWIS must be designed and installed (2) Does not result from a single fail- with adequate physical separation from ure. water/waste lines and other water/ (b) Each hazardous failure condition waste system components, so that: is extremely remote. (1) An EWIS component failure will not create a hazardous condition. § 25.1711 Component identification: (2) Any water/waste leakage onto EWIS. EWIS components will not create a (a) EWIS components must be labeled hazardous condition. or otherwise identified using a con- (i) EWIS must be designed and in- sistent method that facilitates identi- stalled with adequate physical separa- fication of the EWIS component, its tion between the EWIS and flight or function, and its design limitations, if other mechanical control systems ca- any. bles and associated system compo- (b) For systems for which redundancy nents, so that: is required, by certification rules, by

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operating rules, or as a result of the as- provided by EWIS components must sessment required by § 25.1709, EWIS provide an electrical return path capa- components associated with those sys- ble of carrying both normal and fault tems must be specifically identified currents without creating a shock haz- with component part number, function, ard or damage to the EWIS compo- and separation requirement for bun- nents, other airplane system compo- dles. nents, or airplane structure. (1) The identification must be placed along the wire, cable, or wire bundle at § 25.1717 Circuit protective devices: appropriate intervals and in areas of EWIS. the airplane where it is readily visible Electrical wires and cables must be to maintenance, repair, or alteration designed and installed so they are com- personnel. patible with the circuit protection de- (2) If an EWIS component cannot be vices required by § 25.1357, so that a fire marked physically, then other means or smoke hazard cannot be created of identification must be provided. under temporary or continuous fault (c) The identifying markings re- conditions. quired by paragraphs (a) and (b) of this section must remain legible through- § 25.1719 Accessibility provisions: out the expected service life of the EWIS. EWIS component. Access must be provided to allow in- (d) The means used for identifying spection and replacement of any EWIS each EWIS component as required by component as necessary for continued this section must not have an adverse airworthiness. effect on the performance of that component throughout its expected service § 25.1721 Protection of EWIS. life. (a) No cargo or baggage compartment (e) Identification for EWIS modifica- may contain any EWIS whose damage tions to the type design must be con- or failure may affect safe operation, sistent with the identification scheme unless the EWIS is protected so that: of the original type design. (1) It cannot be damaged by movement of cargo or baggage in the com- § 25.1713 Fire protection: EWIS. partment. (a) All EWIS components must meet (2) Its breakage or failure will not the applicable fire and smoke protec- create a fire hazard. tion requirements of § 25.831(c) of this (b) EWIS must be designed and in- part. stalled to minimize damage and risk of (b) EWIS components that are lo- damage to EWIS by movement of peo- cated in designated fire zones and are ple in the airplane during all phases of used during emergency procedures flight, maintenance, and servicing. must be fire resistant. (c) EWIS must be designed and in- (c) Insulation on electrical wire and stalled to minimize damage and risk of electrical cable, and materials used to damage to EWIS by items carried onto provide additional protection for the the aircraft by passengers or cabin wire and cable, installed in any area of crew. the airplane, must be self-extinguishing when tested in accordance § 25.1723 Flammable fluid fire protection: EWIS. with the applicable portions of Appen- dix F, part I, of 14 CFR part 25. EWIS components located in each area where flammable fluid or vapors § 25.1715 Electrical bonding and pro- might escape by leakage of a fluid sys- tection against static electricity: tem must be considered a potential ig- EWIS. nition source and must meet the re- (a) EWIS components used for elec- quirements of § 25.863. trical bonding and protection against static electricity must meet the re- § 25.1725 Powerplants: EWIS. quirements of § 25.899. (a) EWIS associated with any power- (b) On airplanes having grounded plant must be designed and installed so electrical systems, electrical bonding that the failure of an EWIS component

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will not prevent the continued safe op- (b) No EWIS component of any fire or eration of the remaining powerplants overheat detector system for any fire or require immediate action by any zone may pass through another fire crewmember for continued safe oper- zone, unless: ation, in accordance with the require- (1) It is protected against the possi- ments of § 25.903(b). bility of false warnings resulting from (b) Design precautions must be taken fires in zones through which it passes; to minimize hazards to the airplane or due to EWIS damage in the event of a (2) Each zone involved is simulta- powerplant rotor failure or a fire originating within the powerplant that neously protected by the same detector burns through the powerplant case, in and extinguishing system. accordance with the requirements of (c) EWIS that are part of each fire or § 25.903(d)(1). overheat detector system in a fire zone must meet the requirements of § 25.1727 Flammable fluid shutoff § 25.1203. means: EWIS. EWIS associated with each flam- § 25.1733 Fire detector systems, gen- mable fluid shutoff means and control eral: EWIS. must be fireproof or must be located EWIS associated with any installed and protected so that any fire in a fire fire protection system, including those zone will not affect operation of the required by §§ 25.854 and 25.858, must be flammable fluid shutoff means, in ac- considered an integral part of the sys- cordance with the requirements of tem in showing compliance with the § 25.1189. applicable requirements for that sys- § 25.1729 Instructions for Continued tem. Airworthiness: EWIS. The applicant must prepare Instruc- Subpart I—Special Federal tions for Continued Airworthiness ap- Aviation Regulations plicable to EWIS in accordance with Appendix H sections H25.4 and H25.5 to SOURCE: Docket No. FAA–2011–0186, Amdt. this part that are approved by the 25–133, 76 FR 12555, Mar. 8, 2011, unless other- FAA. wise noted.

§ 25.1731 Powerplant and APU fire de- § 25.1801 SFAR No. 111—Lavatory Oxy- tector system: EWIS. gen Systems. (a) EWIS that are part of each fire or The requirements of § 121.1500 of this overheat detector system in a fire zone must be fire-resistant. chapter also apply to this part.

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APPENDIX A TO PART 25

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APPENDIX B TO PART 25

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APPENDIX C TO PART 25 by the appropriate factor from figure 3 of this appendix. Part I—Atmospheric Icing Conditions (b) Intermittent maximum icing. The intermittent maximum intensity of atmospheric (a) Continuous maximum icing. The max- icing conditions (intermittent maximum imum continuous intensity of atmospheric icing) is defined by the variables of the cloud icing conditions (continuous maximum liquid water content, the mean effective di- icing) is defined by the variables of the cloud ameter of the cloud droplets, the ambient air liquid water content, the mean effective di- temperature, and the interrelationship of ameter of the cloud droplets, the ambient air these three variables as shown in figure 4 of temperature, and the interrelationship of this appendix. The limiting icing envelope in these three variables as shown in figure 1 of terms of altitude and temperature is given in this appendix. The limiting icing envelope in figure 5 of this appendix. The inter-relation- terms of altitude and temperature is given in ship of cloud liquid water content with drop figure 2 of this appendix. The inter-relation- diameter and altitude is determined from ship of cloud liquid water content with drop figures 4 and 5. The cloud liquid water con- diameter and altitude is determined from tent for intermittent maximum icing condi- figures 1 and 2. The cloud liquid water con- tions of a horizontal extent, other than 2.6 tent for continuous maximum icing condi- nautical miles, is determined by the value of tions of a horizontal extent, other than 17.4 cloud liquid water content of figure 4 multi- nautical miles, is determined by the value of plied by the appropriate factor in figure 6 of liquid water content of figure 1, multiplied this appendix.

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FIGURE 2

t~ ~. :~:: :~:..: :.:...:: ,~- .. CONTINUOUS MAXIMUM (STRATIFORM CLOUDS)

ATMOSPH ERIC ICING CONDITIONS ". :::::::. ::..:-::-: _ i'i ~::i i: "; ii; AMBIENT TEMPERATURE VS PRESSURE ALTITUDE

.... ,:i:;:~;: :... , :':1:;;", ';:1 :.::1 .... ':i •.•• :.;"...... _ .... :+~r= . "' ··b ~~~ ~T ~::~ i?~~ :~~:' :~;:: .... "::; :~J:~~: ~~: ~::: ~;~ ::~~~ ::;; :::; ::::: ~;:;: ::J~~: ...... ;:::.~ ..'"

...... ;::1.' :•• :, •• ··:.4 •••••• :1····.... •..• •••... •••••••.•.

il ; ••• ;...:[ 1/1/ 17171/1/ 17 [2[717171/1;;12[1 I •• :. I ...... '.'.-.-'~ ... ~ -- . .. ---_ .. '''~''.Ej;-0 '::;"i ......

•• ;; ::.'Eii, :::: :::: ... fi'd':-:' ::" ~" .. ... :::c ''''1'-':-...', .".,,~ •• " ". ".:

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FIGURE 5

..•••.•••..•••.•••• ".' INTERMITTENT MAXIMUM (CUMULI FORM CLOUDS):i ..•. .•••• .. .'c •••• •••.•.•••. •••• ••••• ATMOSPHERIC ICING CONDITIONS t .•••.••.... :•••••. ~" •• '.. •... :1' ..... , •••• ...... •••• ...... AMBIENT TEMPERA l'UR E VS PRESSURE ALTITUDE .. +ii-4•• ".,.j,...... -4 •.• 4+"'8 ...... :. .. ._..+ ·.~rl •.••. •...••.•• ·..TI1111 1 .•.•.•

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(c) Takeoff maximum icing. The maximum temperature at ground level of minus 9 de- intensity of atmospheric icing conditions for grees Celsius (¥9 °C). The takeoff maximum takeoff (takeoff maximum icing) is defined icing conditions extend from ground level to by the cloud liquid water content of 0.35 g/ a height of 1,500 feet above the level of the m3, the mean effective diameter of the cloud takeoff surface. droplets of 20 microns, and the ambient air

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Part II—Airframe Ice Accretions for Showing ments of § 25.21(g), any of the ice accretions Compliance With Subpart B. defined in paragraph (a) of this section may be used for any other flight phase if it is (a) Ice accretions—General. The most crit- shown to be more critical than the specific ical ice accretion in terms of airplane per- ice accretion defined for that flight phase. formance and handling qualities for each Configuration differences and their effects flight phase must be used to show compli- on ice accretions must be taken into ac- ance with the applicable airplane perform- count. ance and handling requirements in icing con- (c) The ice accretion that has the most ad- ditions of subpart B of this part. Applicants verse effect on handling qualities may be must demonstrate that the full range of at- used for airplane performance tests provided mospheric icing conditions specified in part I any difference in performance is conserv- of this appendix have been considered, in- atively taken into account. cluding the mean effective drop diameter, (d) For both unprotected and protected liquid water content, and temperature appro- parts, the ice accretion for the takeoff phase priate to the flight conditions (for example, may be determined by calculation, assuming configuration, speed, angle-of-attack, and al- the takeoff maximum icing conditions de- titude). The ice accretions for each flight fined in appendix C, and assuming that: phase are defined as follows: (1) Takeoff ice is the most critical ice accre- (1) Airfoils, control surfaces and, if appli- tion on unprotected surfaces and any ice ac- cable, propellers are free from frost, snow, or cretion on the protected surfaces appropriate ice at the start of the takeoff; to normal ice protection system operation, (2) The ice accretion starts at liftoff; occurring between liftoff and 400 feet above (3) The critical ratio of thrust/power-to- the takeoff surface, assuming accretion weight; starts at liftoff in the takeoff maximum (4) Failure of the critical engine occurs at icing conditions of part I, paragraph (c) of VEF; and this appendix. (5) Crew activation of the ice protection (2) Final takeoff ice is the most critical ice system is in accordance with a normal oper- accretion on unprotected surfaces, and any ating procedure provided in the Airplane ice accretion on the protected surfaces ap- Flight Manual, except that after beginning propriate to normal ice protection system the takeoff roll, it must be assumed that the operation, between 400 feet and either 1,500 crew takes no action to activate the ice pro- feet above the takeoff surface, or the height tection system until the airplane is at least at which the transition from the takeoff to 400 feet above the takeoff surface. the en route configuration is completed and (e) The ice accretion before the ice protection system has been activated and is per- VFTO is reached, whichever is higher. Ice accretion is assumed to start at liftoff in the forming its intended function is the critical takeoff maximum icing conditions of part I, ice accretion formed on the unprotected and paragraph (c) of this appendix. normally protected surfaces before activa- (3) En route ice is the critical ice accretion tion and effective operation of the ice pro- on the unprotected surfaces, and any ice ac- tection system in continuous maximum at- cretion on the protected surfaces appropriate mospheric icing conditions. This ice accre- to normal ice protection system operation, tion only applies in showing compliance to during the en route phase. §§ 25.143(j) and 25.207(h), and 25.207(i). (4) Holding ice is the critical ice accretion [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as on the unprotected surfaces, and any ice ac- amended by Amdt. 25–121, 72 FR 44669, Aug. 8, cretion on the protected surfaces appropriate 2007; 72 FR 50467, Aug. 31, 2007; Amdt. 25–129, to normal ice protection system operation, 74 FR 38340, Aug. 3, 2009] during the holding flight phase. (5) Approach ice is the critical ice accretion APPENDIX D TO PART 25 on the unprotected surfaces, and any ice accretion on the protected surfaces appropriate Criteria for determining minimum flight crew. to normal ice protection system operation The following are considered by the Agency following exit from the holding flight phase in determining the minimum flight crew and transition to the most critical approach under § 25.1523: configuration. (a) Basic workload functions. The following (6) Landing ice is the critical ice accretion basic workload functions are considered: on the unprotected surfaces, and any ice ac- (1) Flight path control. cretion on the protected surfaces appropriate (2) Collision avoidance. to normal ice protection system operation (3) Navigation. following exit from the approach flight phase (4) Communications. and transition to the final landing configura- (5) Operation and monitoring of aircraft tion. engines and systems. (b) In order to reduce the number of ice ac- (6) Command decisions. cretions to be considered when dem- (b) Workload factors. The following work- onstrating compliance with the require- load factors are considered significant when

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analyzing and demonstrating workload for APPENDIX E TO PART 25 minimum flight crew determination: (1) The accessibility, ease, and simplicity I—Limited Weight Credit For Airplanes of operation of all necessary flight, power, Equipped With Standby Power and equipment controls, including emer- (a) Each applicant for an increase in the gency fuel shutoff valves, electrical controls, maximum certificated takeoff and landing electronic controls, pressurization system weights of an airplane equipped with a type- controls, and engine controls. certificated standby power rocket engine (2) The accessibility and conspicuity of all may obtain an increase as specified in para- necessary instruments and failure warning graph (b) if— devices such as fire warning, electrical sys- (1) The installation of the rocket engine tem malfunction, and other failure or cau- has been approved and it has been estab- tion indicators. The extent to which such in- lished by flight test that the rocket engine struments or devices direct the proper cor- and its controls can be operated safely and rective action is also considered. reliably at the increase in maximum weight; and (3) The number, urgency, and complexity (2) The Airplane Flight Manual, or the of operating procedures with particular con- placard, markings or manuals required in sideration given to the specific fuel manage- place thereof, set forth in addition to any ment schedule imposed by center of gravity, other operating limitations the Adminis- structural or other considerations of an air- trator may require, the increased weight ap- worthiness nature, and to the ability of each proved under this regulation and a prohibi- engine to operate at all times from a single tion against the operation of the airplane at tank or source which is automatically re- the approved increased weight when— plenished if fuel is also stored in other tanks. (i) The installed standby power rocket en- (4) The degree and duration of con- gines have been stored or installed in excess centrated mental and physical effort in- of the time limit established by the manu- volved in normal operation and in diagnosing facturer of the rocket engine (usually sten- and coping with malfunctions and emer- ciled on the engine casing); or gencies. (ii) The rocket engine fuel has been ex- (5) The extent of required monitoring of pended or discharged. the fuel, hydraulic, pressurization, elec- (b) The currently approved maximum take- trical, electronic, deicing, and other systems off and landing weights at which an airplane while en route. is certificated without a standby power rock- (6) The actions requiring a crewmember to et engine installation may be increased by an amount that does not exceed any of the be unavailable at his assigned duty station, following: including: observation of systems, emer- (1) An amount equal in pounds to 0.014 IN, gency operation of any control, and emer- where I is the maximum usable impulse in gencies in any compartment. pounds-seconds available from each standby (7) The degree of automation provided in power rocket engine and N is the number of the aircraft systems to afford (after failures rocket engines installed. or malfunctions) automatic crossover or iso- (2) An amount equal to 5 percent of the lation of difficulties to minimize the need for maximum certificated weight approved in flight crew action to guard against loss of accordance with the applicable airworthiness hydraulic or electric power to flight controls regulations without standby power rocket or to other essential systems. engines installed. (8) The communications and navigation (3) An amount equal to the weight of the workload. rocket engine installation. (9) The possibility of increased workload (4) An amount that, together with the cur- associated with any emergency that may rently approved maximum weight, would lead to other emergencies. equal the maximum structural weight estab- (10) Incapacitation of a flight crewmember lished for the airplane without standby rock- whenever the applicable operating rule re- et engines installed. quires a minimum flight crew of at least two II—Performance Credit for Transport Category pilots. Airplanes Equipped With Standby Power (c) Kind of operation authorized. The determination of the kind of operation authorized The Administrator may grant performance requires consideration of the operating rules credit for the use of standby power on trans- under which the airplane will be operated. port category airplanes. However, the per- Unless an applicant desires approval for a formance credit applies only to the max- more limited kind of operation. It is assumed imum certificated takeoff and landing that each airplane certificated under this weights, the takeoff distance, and the take- Part will operate under IFR conditions. off paths, and may not exceed that found by the Administrator to result in an overall [Amdt. 25–3, 30 FR 6067, Apr. 29, 1965] level of safety in the takeoff, approach, and

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landing regimes of flight equivalent to that feet above the takeoff surface for turbine- prescribed in the regulations under which powered airplanes. the airplane was originally certificated with- (4) Maximum certificated takeoff weights. The out standby power. For the purposes of this maximum certificated takeoff weights must appendix, ‘‘standby power’’ is power or be determined at all altitudes, and at ambi- thrust, or both, obtained from rocket en- ent temperatures, if applicable, at which per- gines for a relatively short period and actu- formance credit is to be applied and may not ated only in cases of emergency. The fol- exceed the weights established in compliance lowing provisions apply: with paragraphs (a) and (b) of this section. (1) Takeoff; general. The takeoff data pre- (a) The conditions of paragraphs (2)(b) scribed in paragraphs (2) and (3) of this ap- through (d) must be met at the maximum pendix must be determined at all weights certificated takeoff weight. and altitudes, and at ambient temperatures (b) Without the use of standby power, the if applicable, at which performance credit is airplane must meet all of the en route re- to be applied. quirements of the applicable airworthiness (2) Takeoff path. regulations under which the airplane was (a) The one-engine-inoperative takeoff originally certificated. In addition, turbine- path with standby power in use must be de- powered airplanes without the use of standby termined in accordance with the perform- power must meet the final takeoff climb re- ance requirements of the applicable air- quirements prescribed in the applicable airworthiness regulations. worthiness regulations. (b) The one-engine-inoperative takeoff (5) Maximum certificated landing weights. path (excluding that part where the airplane (a) The maximum certificated landing is on or just above the takeoff surface) deter- weights (one-engine-inoperative approach mined in accordance with paragraph (a) of and all-engine-operating landing climb) must this section must lie above the one-engine- be determined at all altitudes, and at ambi- inoperative takeoff path without standby ent temperatures if applicable, at which per- power at the maximum takeoff weight at formance credit is to be applied and must which all of the applicable air-worthiness re- not exceed that established in compliance quirements are met. For the purpose of this with paragraph (b) of this section. comparison, the flight path is considered to (b) The flight path, with the engines oper- extend to at least a height of 400 feet above ating at the power or thrust, or both, appro- the takeoff surface. priate to the airplane configuration and with (c) The takeoff path with all engines oper- standby power in use, must lie above the ating, but without the use of standby power, flight path without standby power in use at must reflect a conservatively greater overall the maximum weight at which all of the ap- level of performance than the one-engine-in- plicable airworthiness requirements are met. operative takeoff path established in accord- In addition, the flight paths must comply ance with paragraph (a) of this section. The with subparagraphs (i) and (ii) of this para- margin must be established by the Adminis- graph. trator to insure safe day-to-day operations, (i) The flight paths must be established but in no case may it be less than 15 percent. without changing the appropriate airplane The all-engines-operating takeoff path must configuration. be determined by a procedure consistent (ii) The flight paths must be carried out for with that established in complying with a minimum height of 400 feet above the point paragraph (a) of this section. where standby power is actuated. (d) For reciprocating-engine-powered air- (6) Airplane configuration, speed, and power planes, the takeoff path to be scheduled in and thrust; general. Any change in the air- the Airplane Flight Manual must represent plane’s configuration, speed, and power or the one-engine-operative takeoff path deter- thrust, or both, must be made in accordance mined in accordance with paragraph (a) of with the procedures established by the appli- this section and modified to reflect the pro- cant for the operation of the airplane in cedure (see paragraph (6)) established by the service and must comply with paragraphs (a) applicant for flap retraction and attainment through (c) of this section. In addition, proof the en route speed. The scheduled takeoff cedures must be established for the execu- path must have a positive slope at all points tion of balked landings and missed ap- of the airborne portion and at no point must proaches. it lie above the takeoff path specified in (a) The Administrator must find that the paragraph (a) of this section. procedure can be consistently executed in (3) Takeoff distance. The takeoff distance service by crews of average skill. must be the horizontal distance along the (b) The procedure may not involve methods one-engine-inoperative take off path deter- or the use of devices which have not been mined in accordance with paragraph (2)(a) proven to be safe and reliable. from the start of the takeoff to the point (c) Allowances must be made for such time where the airplane attains a height of 50 feet delays in the execution of the procedures as above the takeoff surface for reciprocating- may be reasonably expected to occur during engine-powered airplanes and a height of 35 service.

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(7) Installation and operation; standby power. (iv) Clear plastic windows and signs, parts The standby power unit and its installation constructed in whole or in part of elas- must comply with paragraphs (a) and (b) of tomeric materials, edge lighted instrument this section. assemblies consisting of two or more instru- (a) The standby power unit and its instal- ments in a common housing, seat belts, lation must not adversely affect the safety of shoulder harnesses, and cargo and baggage the airplane. tiedown equipment, including containers, (b) The operation of the standby power bins, pallets, etc., used in passenger or crew unit and its control must have proven to be compartments, may not have an average safe and reliable. burn rate greater than 2.5 inches per minute when tested horizontally in accordance with [Amdt. 25–6, 30 FR 8468, July 2, 1965] the applicable portions of this appendix. (v) Except for small parts (such as knobs, APPENDIX F TO PART 25 handles, rollers, fasteners, clips, grommets, rub strips, pulleys, and small electrical Part I—Test Criteria and Procedures for parts) that would not contribute signifi- Showing Compliance with § 25.853, or § 25.855. cantly to the propagation of a fire and for (a) Material test criteria—(1) Interior com- electrical wire and cable insulation, mate- partments occupied by crew or passengers. (i) rials in items not specified in paragraphs Interior ceiling panels, interior wall panels, (a)(1)(i), (ii), (iii), or (iv) of part I of this ap- partitions, galley structure, large cabinet pendix may not have a burn rate greater walls, structural flooring, and materials used than 4.0 inches per minute when tested hori- in the construction of stowage compart- zontally in accordance with the applicable ments (other than underseat stowage com- portions of this appendix. partments and compartments for stowing (2) Cargo and baggage compartments not oc- small items such as magazines and maps) cupied by crew or passengers. must be self-extinguishing when tested (i) [Reserved] vertically in accordance with the applicable (ii) A cargo or baggage compartment de- portions of part I of this appendix. The aver- fined in § 25.857 as Class B or E must have a age burn length may not exceed 6 inches and liner constructed of materials that meet the the average flame time after removal of the requirements of paragraph (a)(1)(ii) of part I flame source may not exceed 15 seconds. of this appendix and separated from the air- Drippings from the test specimen may not plane structure (except for attachments). In continue to flame for more than an average addition, such liners must be subjected to of 3 seconds after falling. the 45 degree angle test. The flame may not (ii) Floor covering, textiles (including penetrate (pass through) the material during draperies and upholstery), seat cushions, application of the flame or subsequent to its padding, decorative and nondecorative coat- removal. The average flame time after re- ed fabrics, leather, trays and galley fur- moval of the flame source may not exceed 15 nishings, electrical conduit, air ducting, seconds, and the average glow time may not joint and edge covering, liners of Class B and exceed 10 seconds. E cargo or baggage compartments, floor pan- (iii) A cargo or baggage compartment de- els of Class B, C, D, or E cargo or baggage fined in § 25.857 as Class B, C, D, or E must compartments, cargo covers and trans- have floor panels constructed of materials parencies, molded and thermoformed parts, which meet the requirements of paragraph air ducting joints, and trim strips (decora- (a)(1)(ii) of part I of this appendix and which tive and chafing), that are constructed of are separated from the airplane structure materials not covered in subparagraph (iv) (except for attachments). Such panels must below, must be self-extinguishing when test- be subjected to the 45 degree angle test. The ed vertically in accordance with the applica- flame may not penetrate (pass through) the ble portions of part I of this appendix or material during application of the flame or other approved equivalent means. The aver- subsequent to its removal. The average age burn length may not exceed 8 inches, and flame time after removal of the flame source the average flame time after removal of the may not exceed 15 seconds, and the average flame source may not exceed 15 seconds. glow time may not exceed 10 seconds. Drippings from the test specimen may not (iv) Insulation blankets and covers used to continue to flame for more than an average protect cargo must be constructed of mate- of 5 seconds after falling. rials that meet the requirements of para- (iii) Motion picture film must be safety graph (a)(1)(ii) of part I of this appendix. Tie- film meeting the Standard Specifications for down equipment (including containers, bins, Safety Photographic Film PHI.25 (available and pallets) used in each cargo and baggage from the American National Standards Insti- compartment must be constructed of mate- tute, 1430 Broadway, New York, NY 10018). If rials that meet the requirements of para- the film travels through ducts, the ducts graph (a)(1)(v) of part I of this appendix. must meet the requirements of subparagraph (3) Electrical system components. Insulation (ii) of this paragraph. on electrical wire or cable installed in any

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area of the fuselage must be self-extin- vertical test, or Method 5906 for horizontal guishing when subjected to the 60 degree test test (available from the General Services Ad- specified in part I of this appendix. The aver- ministration, Business Service Center, Re- age burn length may not exceed 3 inches, and gion 3, Seventh & D Streets SW., Wash- the average flame time after removal of the ington, DC 20407). Specimens which are too flame source may not exceed 30 seconds. large for the cabinet must be tested in simi- Drippings from the test specimen may not lar draft-free conditions. continue to flame for more than an average (4) Vertical test. A minimum of three speci- of 3 seconds after falling. mens must be tested and results averaged. (b) Test Procedures—(1) Conditioning. Speci- For fabrics, the direction of weave cor- mens must be conditioned to 70 ±5 F., and at responding to the most critical flammability 50 percent ±5 percent relative humidity until conditions must be parallel to the longest di- moisture equilibrium is reached or for 24 mension. Each specimen must be supported hours. Each specimen must remain in the vertically. The specimen must be exposed to conditioning environment until it is sub- a Bunsen or Tirrill burner with a nominal 3⁄8- jected to the flame. inch I.D. tube adjusted to give a flame of 11⁄2 (2) Specimen configuration. Except for small inches in height. The minimum flame tem- parts and electrical wire and cable insula- perature measured by a calibrated thermo- tion, materials must be tested either as sec- couple pyrometer in the center of the flame tion cut from a fabricated part as installed must be 1550 °F. The lower edge of the speci- in the airplane or as a specimen simulating men must be 3⁄4-inch above the top edge of a cut section, such as a specimen cut from a the burner. The flame must be applied to the flat sheet of the material or a model of the center line of the lower edge of the specimen. fabricated part. The specimen may be cut For materials covered by paragraph (a)(1)(i) from any location in a fabricated part; how- of part I of this appendix, the flame must be ever, fabricated units, such as sandwich pan- applied for 60 seconds and then removed. For els, may not be separated for test. Except as materials covered by paragraph (a)(1)(ii) of noted below, the specimen thickness must be part I of this appendix, the flame must be ap- no thicker than the minimum thickness to plied for 12 seconds and then removed. Flame be qualified for use in the airplane. Test time, burn length, and flaming time of drip- specimens of thick foam parts, such as seat pings, if any, may be recorded. The burn cushions, must be 1⁄2-inch in thickness. Test length determined in accordance with sub- specimens of materials that must meet the paragraph (7) of this paragraph must be requirements of paragraph (a)(1)(v) of part I measured to the nearest tenth of an inch. of this appendix must be no more than 1⁄8- (5) Horizontal test. A minimum of three inch in thickness. Electrical wire and cable specimens must be tested and the results specimens must be the same size as used in averaged. Each specimen must be supported the airplane. In the case of fabrics, both the horizontally. The exposed surface, when in- warp and fill direction of the weave must be stalled in the aircraft, must be face down for tested to determine the most critical flam- the test. The specimen must be exposed to a mability condition. Specimens must be Bunsen or Tirrill burner with a nominal 3⁄8- mounted in a metal frame so that the two inch I.D. tube adjusted to give a flame of 11⁄2 long edges and the upper edge are held se- inches in height. The minimum flame tem- curely during the vertical test prescribed in perature measured by a calibrated thermo- subparagraph (4) of this paragraph and the couple pyrometer in the center of the flame two long edges and the edge away from the must be 1550 °F. The specimen must be posi- flame are held securely during the horizontal tioned so that the edge being tested is cen- test prescribed in subparagraph (5) of this tered 3⁄4-inch above the top of the burner. paragraph. The exposed area of the specimen The flame must be applied for 15 seconds and must be at least 2 inches wide and 12 inches then removed. A minimum of 10 inches of long, unless the actual size used in the air- specimen must be used for timing purposes, plane is smaller. The edge to which the burn- approximately 11⁄2 inches must burn before er flame is applied must not consist of the the burning front reaches the timing zone, finished or protected edge of the specimen and the average burn rate must be recorded. but must be representative of the actual (6) Forty-five degree test. A minimum of cross-section of the material or part as in- three specimens must be tested and the re- stalled in the airplane. The specimen must sults averaged. The specimens must be sup- be mounted in a metal frame so that all four ported at an angle of 45° to a horizontal sur- edges are held securely and the exposed area face. The exposed surface when installed in of the specimen is at least 8 inches by 8 the aircraft must be face down for the test. inches during the 45° test prescribed in sub- The specimens must be exposed to a Bunsen paragraph (6) of this paragraph. or Tirrill burner with a nominal 3⁄8-inch I.D. (3) Apparatus. Except as provided in sub- tube adjusted to give a flame of 11⁄2 inches in paragraph (7) of this paragraph, tests must height. The minimum flame temperature be conducted in a draft-free cabinet in ac- measured by a calibrated thermocouple py- cordance with Federal Test Method Standard rometer in the center of the flame must be 191 Model 5903 (revised Method 5902) for the 1550 °F. Suitable precautions must be taken

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to avoid drafts. The flame must be applied (1) At least three sets of seat bottom and for 30 seconds with one-third contacting the seat back cushion specimens must be tested. material at the center of the specimen and (2) If the cushion is constructed with a fire then removed. Flame time, glow time, and blocking material, the fire blocking material whether the flame penetrates (passes must completely enclose the cushion foam through) the specimen must be recorded. core material. (7) Sixty degree test. A minimum of three (3) Each specimen tested must be fab- specimens of each wire specification (make ricated using the principal components (i.e., and size) must be tested. The specimen of foam core, flotation material, fire blocking wire or cable (including insulation) must be material, if used, and dress covering) and as- placed at an angle of 60° with the horizontal sembly processes (representative seams and in the cabinet specified in subparagraph (3) closures) intended for use in the production of this paragraph with the cabinet door open articles. If a different material combination during the test, or must be placed within a is used for the back cushion than for the bot- chamber approximately 2 feet high by 1 foot tom cushion, both material combinations by 1 foot, open at the top and at one vertical must be tested as complete specimen sets, side (front), and which allows sufficient flow each set consisting of a back cushion speci- of air for complete combustion, but which is men and a bottom cushion specimen. If a free from drafts. The specimen must be par- cushion, including outer dress covering, is allel to and approximately 6 inches from the demonstrated to meet the requirements of front of the chamber. The lower end of the this appendix using the oil burner test, the specimen must be held rigidly clamped. The dress covering of that cushion may be re- upper end of the specimen must pass over a placed with a similar dress covering provided pulley or rod and must have an appropriate the burn length of the replacement covering, weight attached to it so that the specimen is as determined by the test specified in held tautly throughout the flammability § 25.853(c), does not exceed the corresponding test. The test specimen span between lower burn length of the dress covering used on the clamp and upper pulley or rod must be 24 cushion subjected to the oil burner test. inches and must be marked 8 inches from the (4) For at least two-thirds of the total lower end to indicate the central point for number of specimen sets tested, the burn length from the burner must not reach the flame application. A flame from a Bunsen or side of the cushion opposite the burner. The Tirrill burner must be applied for 30 seconds burn length must not exceed 17 inches. Burn at the test mark. The burner must be mount- length is the perpendicular distance from the ed underneath the test mark on the speci- inside edge of the seat frame closest to the men, perpendicular to the specimen and at burner to the farthest evidence of damage to an angle of 30° to the vertical plane of the the test specimen due to flame impingement, specimen. The burner must have a nominal including areas of partial or complete con- bore of 3⁄8-inch and be adjusted to provide a sumption, charring, or embrittlement, but 3-inch high flame with an inner cone ap- not including areas sooted, stained, warped, proximately one-third of the flame height. or discolored, or areas where material has The minimum temperature of the hottest shrunk or melted away from the heat source. portion of the flame, as measured with a (5) The average percentage weight loss calibrated thermocouple pyrometer, may not must not exceed 10 percent. Also, at least ° be less than 1750 F. The burner must be posi- two-thirds of the total number of specimen tioned so that the hottest portion of the sets tested must not exceed 10 percent flame is applied to the test mark on the weight loss. All droppings falling from the wire. Flame time, burn length, and flaming cushions and mounting stand are to be dis- time of drippings, if any, must be recorded. carded before the after-test weight is deter- The burn length determined in accordance mined. The percentage weight loss for a spec- with paragraph (8) of this paragraph must be imen set is the weight of the specimen set measured to the nearest tenth of an inch. before testing less the weight of the speci- Breaking of the wire specimens is not consid- men set after testing expressed as the per- ered a failure. centage of the weight before testing. (8) Burn length. Burn length is the distance (b) Test Conditions. Vertical air velocity from the original edge to the farthest evi- should average 25 fpm±10 fpm at the top of dence of damage to the test specimen due to the back seat cushion. Horizontal air veloc- flame impingement, including areas of par- ity should be below 10 fpm just above the tial or complete consumption, charring, or bottom seat cushion. Air velocities should be embrittlement, but not including areas soot- measured with the ventilation hood oper- ed, stained, warped, or discolored, nor areas ating and the burner motor off. where material has shrunk or melted away (c) Test Specimens. (1) For each test, one set from the heat source. of cushion specimens representing a seat bottom and seat back cushion must be used. Part II—Flammability of Seat Cushions (2) The seat bottom cushion specimen must (a) Criteria for Acceptance. Each seat cush- be 18 ±1⁄8 inches (457 ±3 mm) wide by 20 ±1⁄8 ion must meet the following criteria: inches (508 ±3 mm) deep by 4 ±1⁄8 inches (102

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±3 mm) thick, exclusive of fabric closures plane of the insulating board in a plane par- and seam overlap. allel to the exit of the test burner cone. (3) The seat back cushion specimen must (4) Thermocouples. The seven thermocouples 1 1 be 18 ± ⁄8 inches (432 ±3 mm) wide by 25 ± ⁄8 to be used for testing must be 1⁄16- to 1⁄8-inch inches (635 ±3 mm) high by 2 ±1⁄8 inches (51 ±3 metal sheathed, ceramic packed, type K, mm) thick, exclusive of fabric closures and grounded thermocouples with a nominal 22 seam overlap. to 30 American wire gage (AWG)-size con- (4) The specimens must be conditioned at ductor. The seven thermocouples must be at- 70 ±5 °F (21 ±2 °C) 55%±10% relative humidity tached to a steel angle bracket to form a for at least 24 hours before testing. thermocouple rake for placement in the test (d) Test Apparatus. The arrangement of the stand during burner calibration, as shown in test apparatus is shown in Figures 1 through Figure 5. 5 and must include the components described (5) Apparatus Arrangement. The test burner in this section. Minor details of the appa- must be mounted on a suitable stand to posi- ratus may vary, depending on the model tion the exit of the burner cone a distance of burner used. 4 ±1⁄8 inches (102 ±3 mm) from one side of the (1) Specimen Mounting Stand. The mounting specimen mounting stand. The burner stand stand for the test specimens consists of steel should have the capability of allowing the angles, as shown in Figure 1. The length of burner to be swung away from the specimen ±1 the mounting stand legs is 12 ⁄8 inches (305 mounting stand during warmup periods. ± 3 mm). The mounting stand must be used (6) Data Recording. A recording potentiom- for mounting the test specimen seat bottom eter or other suitable calibrated instrument and seat back, as shown in Figure 2. The with an appropriate range must be used to mounting stand should also include a suit- measure and record the outputs of the calo- able drip pan lined with aluminum foil, dull rimeter and the thermocouples. side up. (7) Weight Scale. Weighing Device—A device (2) Test Burner. The burner to be used in must be used that with proper procedures testing must— (i) Be a modified gun type; may determine the before and after test (ii) Have an 80-degree spray angle nozzle weights of each set of seat cushion specimens nominally rated for 2.25 gallons/hour at 100 within 0.02 pound (9 grams). A continuous psi; weighing system is preferred. (iii) Have a 12-inch (305 mm) burner cone (8) Timing Device. A stopwatch or other de- installed at the end of the draft tube, with vice (calibrated to ±1 second) must be used to an opening 6 inches (152 mm) high and 11 measure the time of application of the burn- inches (280 mm) wide, as shown in Figure 3; er flame and self-extinguishing time or test and duration. (iv) Have a burner fuel pressure regulator (e) Preparation of Apparatus. Before calibra- that is adjusted to deliver a nominal 2.0 gal- tion, all equipment must be turned on and lon/hour of # 2 Grade kerosene or equivalent the burner fuel must be adjusted as specified required for the test. in paragraph (d)(2). Burner models which have been used success- (f) Calibration. To ensure the proper ther- fully in testing are the Lennox Model OB–32, mal output of the burner, the following test Carlin Model 200 CRD, and Park Model DPL must be made: 3400. FAA published reports pertinent to this (1) Place the calorimeter on the test stand type of burner are: (1) Powerplant as shown in Figure 4 at a distance of 4 ±1⁄8 Enginering Report No. 3A, Standard Fire inches (102 ±3 mm) from the exit of the burn- Test Apparatus and Procedure for Flexible er cone. Hose Assemblies, dated March 1978; and (2) (2) Turn on the burner, allow it to run for Report No. DOT/FAA/RD/76/213, Reevaluation 2 minutes for warmup, and adjust the burner of Burner Characteristics for Fire Resistance air intake damper to produce a reading of Tests, dated January 1977. 10.5 ±0.5 BTU/ft2-sec. (11.9 ±0.6 w/cm2) on the (3) Calorimeter. calorimeter to ensure steady state condi- (i) The calorimeter to be used in testing tions have been achieved. Turn off the burn- must be a (0–15.0 BTU/ft2-sec. 0–17.0 W/cm2) er. calorimeter, accurate ±3%, mounted in a 6- (3) Replace the calorimeter with the ther- inch by 12-inch (152 by 305 mm) by 3⁄4-inch (19 mocouple rake (Figure 5). mm) thick calcium silicate insulating board (4) Turn on the burner and ensure that the which is attached to a steel angle bracket for thermocouples are reading 1900 ±100 °F (1038 placement in the test stand during burner ±38 °C) to ensure steady state conditions calibration, as shown in Figure 4. have been achieved. (ii) Because crumbling of the insulating (5) If the calorimeter and thermocouples do board with service can result in misalign- not read within range, repeat steps in para- ment of the calorimeter, the calorimeter graphs 1 through 4 and adjust the burner air must be monitored and the mounting intake damper until the proper readings are shimmed, as necessary, to ensure that the obtained. The thermocouple rake and the calorimeter face is flush with the exposed calorimeter should be used frequently to

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maintain and record calibrated test param- (5) To begin the test, swing the burner into eters. Until the specific apparatus has dem- the test position and simultaneously start onstrated consistency, each test should be the timing device. calibrated. After consistency has been con- (6) Expose the seat bottom cushion speci- firmed, several tests may be conducted with men to the burner flame for 2 minutes and the pre-test calibration before and a calibra- then turn off the burner. Immediately swing tion check after the series. the burner away from the test position. Ter- (g) Test Procedure. The flammability of minate test 7 minutes after initiating cush- each set of specimens must be tested as fol- ion exposure to the flame by use of a gaseous lows: extinguishing agent (i.e., Halon or CO2). (1) Record the weight of each set of seat (7) Determine the weight of the remains of bottom and seat back cushion specimens to the seat cushion specimen set left on the be tested to the nearest 0.02 pound (9 grams). mounting stand to the nearest 0.02 pound (9 (2) Mount the seat bottom and seat back grams) excluding all droppings. cushion test specimens on the test stand as (h) Test Report. With respect to all speci- shown in Figure 2, securing the seat back men sets tested for a particular seat cushion cushion specimen to the test stand at the for which testing of compliance is performed, top. the following information must be recorded: (3) Swing the burner into position and en- (1) An identification and description of the sure that the distance from the exit of the specimens being tested. burner cone to the side of the seat bottom (2) The number of specimen sets tested. cushion specimen is 4 ±1⁄8 inches (102 ±3 mm). (3) The initial weight and residual weight (4) Swing the burner away from the test of each set, the calculated percentage weight position. Turn on the burner and allow it to loss of each set, and the calculated average run for 2 minutes to provide adequate percentage weight loss for the total number warmup of the burner cone and flame sta- of sets tested. bilization. (4) The burn length for each set tested.

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Part III—Test Method To Determine Flame Pen- such as joints, lamp assemblies, etc., the etration Resistance of Cargo Compartment failure of which would affect the capability Liners. of the liner to safely contain a fire. (3) There must be no flame penetration of (a) Criteria for Acceptance. (1) At least three any specimen within 5 minutes after applica- specimens of cargo compartment sidewall or tion of the flame source, and the peak tem- ceiling liner panels must be tested. perature measured at 4 inches above the (2) Each specimen tested must simulate upper surface of the horizontal test sample the cargo compartment sidewall or ceiling must not exceed 400 °F. liner panel, including any design features,

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(b) Summary of Method. This method pro- orimeter face is parallel to the exit plane of vides a laboratory test procedure for meas- the test burner cone. uring the capability of cargo compartment (4) Thermocouples. The seven thermocouples lining materials to resist flame penetration to be used for testing must be 1⁄16 inch ce- with a 2 gallon per hour (GPH) #2 Grade ker- ramic sheathed, type K, grounded osene or equivalent burner fire source. Ceil- thermocouples with a nominal 30 American ing and sidewall liner panels may be tested wire gage (AWG) size conductor. The seven individually provided a baffle is used to sim- thermocouples must be attached to a steel ulate the missing panel. Any specimen that angle bracket to form a thermocouple rake passes the test as a ceiling liner panel may for placement in the test stand during burn- be used as a sidewall liner panel. er calibration as shown in Figure 3 of this (c) Test Specimens. (1) The specimen to be part of this appendix. tested must measure 16 ±1⁄8 inches (406 ±3 (5) Apparatus Arrangement. The test burner mm) by 24+1⁄8 inches (610 ±3 mm). must be mounted on a suitable stand to posi- (2) The specimens must be conditioned at tion the exit of the burner cone a distance of 70 °F.±5 °F. (21 °C. ±2 °C.) and 55%±5% humid- 8 inches from the ceiling liner panel and 2 ity for at least 24 hours before testing. inches from the sidewall liner panel. The (d) Test Apparatus. The arrangement of the burner stand should have the capability of test apparatus, which is shown in Figure 3 of allowing the burner to be swung away from Part II and Figures 1 through 3 of this part the test specimen during warm-up periods. of appendix F, must include the components (6) Instrumentation. A recording potentiom- described in this section. Minor details of the eter or other suitable instrument with an ap- apparatus may vary, depending on the model propriate range must be used to measure and of the burner used. record the outputs of the calorimeter and (1) Specimen Mounting Stand. The mounting the thermocouples. stand for the test specimens consists of steel (7) Timing Device. A stopwatch or other de- angles as shown in Figure 1. vice must be used to measure the time of (2) Test Burner. The burner to be used in flame application and the time of flame pen- tesing must— etration, if it occurs. (i) Be a modified gun type. (e) Preparation of Apparatus. Before calibra- (ii) Use a suitable nozzle and maintain fuel tion, all equipment must be turned on and pressure to yield a 2 GPH fuel flow. For ex- allowed to stabilize, and the burner fuel flow ample: an 80 degree nozzle nominally rated must be adjusted as specified in paragraph at 2.25 GPH and operated at 85 pounds per (d)(2). square inch (PSI) gage to deliver 2.03 GPH. (f) Calibration. To ensure the proper ther- (iii) Have a 12 inch (305 mm) burner exten- mal output of the burner the following test sion installed at the end of the draft tube must be made: with an opening 6 inches (152 mm) high and (1) Remove the burner extension from the 11 inches (280 mm) wide as shown in Figure end of the draft tube. Turn on the blower 3 of Part II of this appendix. portion of the burner without turning the (iv) Have a burner fuel pressure regulator fuel or igniters on. Measure the air velocity that is adjusted to deliver a nominal 2.0 GPH using a hot wire anemometer in the center of of #2 Grade kerosene or equivalent. the draft tube across the face of the opening. Burner models which have been used success- Adjust the damper such that the air velocity fully in testing are the Lenox Model OB–32, is in the range of 1550 to 1800 ft./min. If tabs Carlin Model 200 CRD and Park Model DPL. are being used at the exit of the draft tube, The basic burner is described in FAA Power- they must be removed prior to this measure- plant Engineering Report No. 3A, Standard ment. Reinstall the draft tube extension Fire Test Apparatus and Procedure for Flexi- cone. ble Hose Assemblies, dated March 1978; how- (2) Place the calorimeter on the test stand ever, the test settings specified in this ap- as shown in Figure 2 at a distance of 8 inches pendix differ in some instances from those (203 mm) from the exit of the burner cone to specified in the report. simulate the position of the horizontal test (3) Calorimeter. (i) The calorimeter to be specimen. used in testing must be a total heat flux Foil (3) Turn on the burner, allow it to run for Type Gardon Gage of an appropriate range 2 minutes for warm-up, and adjust the damp- (approximately 0 to 15.0 British thermal unit er to produce a calorimeter reading of 8.0 ±0.5 (BTU) per ft.2 sec., 0–17.0 watts/cm2). The cal- BTU per ft.2 sec. (9.1 ±0.6 Watts/cm2). orimeter must be mounted in a 6 inch by 12 (4) Replace the calorimeter with the ther- inch (152 by 305 mm) by 3⁄4 inch (19 mm) thick mocouple rake (see Figure 3). insulating block which is attached to a steel (5) Turn on the burner and ensure that angle bracket for placement in the test stand each of the seven thermocouples reads 1700 during burner calibration as shown in Figure °F. ±100 °F. (927 °C. ±38 °C.) to ensure steady 2 of this part of this appendix. state conditions have been achieved. If the (ii) The insulating block must be mon- temperature is out of this range, repeat steps itored for deterioration and the mounting 2 through 5 until proper readings are ob- shimmed as necessary to ensure that the cal- tained.

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(6) Turn off the burner and remove the The test may be terminated earlier if flame thermocouple rake. penetration is observed. (7) Repeat (1) to ensure that the burner is (5) When testing ceiling liner panels, in the correct range. record the peak temperature measured 4 (g) Test Procedure. (1) Mount a thermo- inches above the sample. couple of the same type as that used for cali- (6) Record the time at which flame pene- bration at a distance of 4 inches (102 mm) tration occurs if applicable. above the horizontal (ceiling) test specimen. The thermocouple should be centered over (h) Test Report. The test report must in- the burner cone. clude the following: (2) Mount the test specimen on the test (1) A complete description of the materials stand shown in Figure 1 in either the hori- tested including type, manufacturer, thick- zontal or vertical position. Mount the insu- ness, and other appropriate data. lating material in the other position. (2) Observations of the behavior of the test (3) Position the burner so that flames will specimens during flame exposure such as not impinge on the specimen, turn the burn- delamination, resin ignition, smoke, ect., iner on, and allow it to run for 2 minutes. Ro- cluding the time of such occurrence. tate the burner to apply the flame to the (3) The time at which flame penetration specimen and simultaneously start the tim- occurs, if applicable, for each of the three ing device. specimens tested. (4) Expose the test specimen to the flame for 5 minutes and then turn off the burner. (4) Panel orientation (ceiling or sidewall).

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Part IV—Test Method To Determine the Heat this part IV. The truncated diamond-shaped Release Rate From Cabin Materials Exposed mask of .042 ± .002 inch (1.07 ± .05mm) stainless to Radiant Heat. steel must be added to provide uniform heat flux density over the area occupied by the (a) Summary of Method. Three or more vertical sample. specimens representing the completed air- (4) Air Distribution System. The air entering craft component are tested. Each test speci- the environmental chamber must be distrib- men is injected into an environmental cham- uted by a .25 inch (6.3 mm) thick aluminum ber through which a constant flow of air plate having eight No. 4 drill-holes, located 2 passes. The specimen’s exposure is deter- inches (51 mm) from sides on 4 inch (102 mm) mined by a radiant heat source adjusted to centers, mounted at the base of the environ- produce, on the specimen, the desired total mental chamber. A second plate of 18 guage 2 heat flux of 3.5 W/cm . The specimen is tested stainless steel having 120, evenly spaced, No. with the exposed surface vertical. Combus- 28 drill holes must be mounted 6 inches (152 tion is initiated by piloted ignition. The mm) above the aluminum plate. A well-regu- combustion products leaving the chamber lated air supply is required. The air-supply are monitored in order to calculate the re- manifold at the base of the pyramidal sec- lease rate of heat. tion must have 48, evenly spaced, No. 26 drill (b) Apparatus. The Ohio State University holes located .38 inch (10 mm) from the inner (OSU) rate of heat release apparatus, as de- edge of the manifold, resulting in an airflow scribed below, is used. This is a modified split of approximately three to one within version of the rate of heat release apparatus the apparatus. standardized by the American Society of (5) Exhaust Stack. An exhaust stack, Testing and Materials (ASTM), ASTM E–906. 5.25 × 2.75 inches (133 × 70 mm) in cross section, (1) This apparatus is shown in Figures 1A and 10 inches (254 mm) long, fabricated from and 1B of this part IV. All exterior surfaces 28 guage stainless steel must be mounted on of the apparatus, except the holding cham- the outlet of the pyramidal section. A. ber, must be insulated with 1 inch (25 mm) 1.0 × 3.0 inch (25 × 76 mm) baffle plate of thick, low density, high temperature, fiber- .018 ± .002 inch (.50 ± .05 mm) stainless steel glass board insulation. A gasketed door, must be centered inside the stack, perpen- through which the sample injection rod dicular to the air flow, 3 inches (76 mm) slides, must be used to form an airtight clo- above the base of the stack. sure on the specimen hold chamber. (6) Specimen Holders. (i) The specimen must (2) Thermopile. The temperature difference be tested in a vertical orientation. The speci- between the air entering the environmental men holder (Figure 3 of this part IV) must chamber and that leaving must be monitored incorporate a frame that touches the speci- by a thermopile having five hot, and five men (which is wrapped with aluminum foil cold, 24-guage Chromel-Alumel junctions. as required by paragraph (d)(3) of this Part) The hot junctions must be spaced across the along only the .25 inch (6 mm) perimeter. A top of the exhaust stack, .38 inches (10 mm) ‘‘V’’ shaped spring is used to hold the assem- below the top of the chimney. The bly together. A detachable .50 × .50 × 5.91 inch thermocouples must have a .050 ±.010 inch (12 × 12 × 150 mm) drip pan and two .020 inch (.5 (1.3 ±.3mm) diameter, ball-type, welded tip. mm) stainless steel wires (as shown in Fig- One thermocouple must be located in the ure 3 of this part IV) must be used for testing geometric center, with the other four located materials prone to melting and dripping. The 1.18 inch (30 mm) from the center along the positioning of the spring and frame may be diagonal toward each of the corners (Figure changed to accommodate different specimen 5 of this part IV). The cold junctions must be thicknesses by inserting the retaining rod in located in the pan below the lower air dis- different holes on the specimen holder. tribution plate (see paragraph (b)(4) of this (ii) Since the radiation shield described in part IV). Thermopile hot junctions must be ASTM E–906 is not used, a guide pin must be cleared of soot deposits as needed to main- added to the injection mechanism. This fits tain the calibrated sensitivity. into a slotted metal plate on the injection (3) Radiation Source. A radiant heat source mechanism outside of the holding chamber. incorporating four Type LL silicon carbide It can be used to provide accurate posi- elements, 20 inches (508 mm) long by .63 inch tioning of the specimen face after injection. (16 mm) O.D., must be used, as shown in Fig- The front surface of the specimen must be 3.9 ures 2A and 2B of this part IV. The heat inches (100 mm) from the closed radiation source must have a nominal resistance of 1.4 doors after injection. ohms and be capable of generating a flux up (iii) The specimen holder clips onto the to 100 kW/m2. The silicone carbide elements mounted bracket (Figure 3 of this part IV). must be mounted in the stainless steel panel The mounting bracket must be attached to box by inserting them through .63 inch (16 the injection rod by three screws that pass mm) holes in .03 inch (1 mm) thick ceramic through a wide-area washer welded onto a 1⁄2- fiber or calcium-silicate millboard. Loca- inch (13 mm) nut. The end of the injection tions of the holes in the pads and stainless rod must be threaded to screw into the nut, steel cover plates are shown in Figure 2B of and a .020 inch (5.1 mm) thick wide area

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washer must be held between two 1⁄2-inch (13 ports, all radiating in the same direction. mm) nuts that are adjusted to tightly cover The first hole must be .50 inch (13 mm) from the hole in the radiation doors through the closed end of the tubing. The tube must which the injection rod or calibration calo- be positioned above the specimen holder so rimeter pass. that the holes are placed above the specimen (7) Calorimeter. A total-flux type calo- as shown in Figure 1B of this part IV. The rimeter must be mounted in the center of a fuel supplied to the burner must be methane 1⁄2-inch Kaowool ‘‘M’’ board inserted in the mixed with air in a ratio of approximately sample holder to measure the total heat flux. 50/50 by volume. The total gas flow must be The calorimeter must have a view angle of adjusted to produce flame lengths of 1 inch 180 degrees and be calibrated for incident (25 mm). When the gas/air ratio and the flow flux. The calorimeter calibration must be ac- rate are properly adjusted, approximately .25 ceptable to the Administrator. inch (6 mm) of the flame length appears yel- (8) Pilot-Flame Positions. Pilot ignition of low in color. the specimen must be accomplished by si- (c) Calibration of Equipment—(1) Heat Re- multaneously exposing the specimen to a lease Rate. A calibration burner, as shown in lower pilot burner and an upper pilot burner, Figure 4, must be placed over the end of the as described in paragraph (b)(8)(i) and lower pilot flame tubing using a gas tight (b)(8)(ii) or (b)(8)(iii) of this part IV, respec- connection. The flow of gas to the pilot tively. Since intermittent pilot flame extin- flame must be at least 99 percent methane guishment for more than 3 seconds would in- and must be accurately metered. Prior to validate the test results, a spark ignitor may usage, the wet test meter must be properly be installed to ensure that the lower pilot leveled and filled with distilled water to the burner remains lighted. tip of the internal pointer while no gas is (i) Lower Pilot Burner. The pilot-flame tub- flowing. Ambient temperature and pressure ing must be .25 inch (6.3 mm) O.D., .03 inch of the water are based on the internal wet (0.8mm) wall, stainless steel tubing. A mix- test meter temperature. A baseline flow rate ture of 120 cm3/min. of methane and 850 cm3/ of approximately 1 liter/min. must be set and min. of air must be fed to the lower pilot increased to higher preset flows of 4, 6, 8, 6 flame burner. The normal position of the end and 4 liters/min. Immediately prior to re- of the pilot burner tubing is .40 inch (10 mm) cording methane flow rates, a flow rate of 8 from and perpendicular to the exposed liters/min. must be used for 2 minutes to pre- vertical surface of the specimen. The center- condition the chamber. This is not recorded line at the outlet of the burner tubing must as part of calibration. The rate must be de- intersect the vertical centerline of the sam- termined by using a stopwatch to time a ple at a point .20 inch (5 mm) above the lower complete revolution of the wet test meter for exposed edge of the specimen. both the baseline and higher flow, with the (ii) Standard Three-Hole Upper Pilot Burner. flow returned to baseline before changing to The pilot burner must be a straight length of the next higher flow. The thermopile base- .25 inch (6.3 mm) O.D., .03 inch (0.8 mm) wall, line voltage must be measured. The gas flow stainless steel tubing that is 14 inches (360 to the burner must be increased to the high- mm) long. One end of the tubing must be er preset flow and allowed to burn for 2.0 closed, and three No. 40 drill holes must be minutes, and the thermopile voltage must be drilled into the tubing, 2.38 inch (60 mm) measured. The sequence must be repeated apart, for gas ports, all radiating in the same until all five values have been determined. direction. The first hole must be .19 inch (5 The average of the five values must be used mm) from the closed end of the tubing. The as the calibration factor. The procedure tube must be positioned .75 inch (19 mm) must be repeated if the percent relative above and .75 inch (19 mm) behind the ex- standard deviation is greater than 5 percent. posed upper edge of the specimen. The mid- Calculations are shown in paragraph (f) of dle hole must be in the vertical plane perpen- this part IV. dicular to the exposed surface of the speci- (2) Flux Uniformity. Uniformity of flux over men which passes through its vertical cen- the specimen must be checked periodically terline and must be pointed toward the radi- and after each heating element change to de- ation source. The gas supplied to the burner termine if it is within acceptable limits of must be methane and must be adjusted to plus or minus 5 percent. produce flame lengths of 1 inch (25 mm). (3) As noted in paragraph (b)(2) of this part (iii) Optional Fourteen-Hole Upper Pilot IV, thermopile hot junctions must be cleared Burner. This burner may be used in lieu of of soot deposits as needed to maintain the the standard three-hole burner described in calibrated sensitivity. paragraph (b)(8)(ii) of this part IV. The pilot (d) Preparation of Test Specimens. (1) The burner must be a straight length of .25 inch test specimens must be representative of the (6.3 mm) O.D., .03 inch (0.8 mm) wall, stain- aircraft component in regard to materials less steel tubing that is 15.75 inches (400 mm) and construction methods. The standard size long. One end of the tubing must be closed, for the test specimens is 5.91 ±.03 × 5.91 ±.03 and 14 No. 59 drill holes must be drilled into inches (149 ±1 × 149 ±1 mm). The thickness of the tubing, .50 inch (13 mm) apart, for gas the specimen must be the same as that of the

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aircraft component it represents up to a (4) The specimen must be placed in the maximum thickness of 1.75 inches (45 mm). hold chamber with the radiation doors Test specimens representing thicker compo- closed. The airtight outer door must be se- nents must be 1.75 inches (45 mm). cured, and the recording devices must be (2) Conditioning. Specimens must be condi- started. The specimen must be retained in tioned as described in Part 1 of this appen- the hold chamber for 60 seconds, plus or dix. minus 10 seconds, before injection. The ther- (3) Mounting. Each test specimen must be mopile ‘‘zero’’ value must be determined dur- wrapped tightly on all sides of the specimen, ing the last 20 seconds of the hold period. except for the one surface that is exposed The sample must not be injected before com- with a single layer of .001 inch (.025 mm) alu- pletion of the ‘‘zero’’ value determination. minum foil. (5) When the specimen is to be injected, the (e) Procedure. (1) The power supply to the radiation doors must be opened. After the radiant panel must be set to produce a radi- specimen is injected into the environmental ant flux of 3.5 ±.05 W/cm2, as measured at the chamber, the radiation doors must be closed point the center of the specimen surface will behind the specimen. occupy when positioned for the test. The ra- (6) [Reserved] diant flux must be measured after the air (7) Injection of the specimen and closure of flow through the equipment is adjusted to the inner door marks time zero. A record of the desired rate. the thermopile output with at least one data (2) After the pilot flames are lighted, their point per second must be made during the position must be checked as described in time the specimen is in the environmental paragraph (b)(8) of this part IV. chamber. (3) Air flow through the apparatus must be (8) The test duration is five minutes. The controlled by a circular plate orifice located lower pilot burner and the upper pilot burner in a 1.5 inch (38.1 mm) I.D. pipe with two must remain lighted for the entire duration pressure measuring points, located 1.5 inches of the test, except that there may be inter- (38 mm) upstream and .75 inches (19 mm) mittent flame extinguishment for periods downstream of the orifice plate. The pipe that do not exceed 3 seconds. Furthermore, if must be connected to a manometer set at a the optional three-hole upper burner is used, pressure differential of 7.87 inches (200 mm) at least two flamelets must remain lighted of Hg. (See Figure 1B of this part IV.) The for the entire duration of the test, except total air flow to the equipment is approxi- that there may be intermittent flame extin- mately .04 m3/seconds. The stop on the guishment of all three flamelets for periods vertical specimen holder rod must be ad- that do not exceed 3 seconds. justed so that the exposed surface of the (9) A minimum of three specimens must be specimen is positioned 3.9 inches (100 mm) tested. from the entrance when injected into the en- (f) Calculations. (1) The calibration factor is vironmental chamber. calculated as follows:

(FF− ) (210. 8− 22)k 273 P− P mole CH4 STP WATT min kw K = 1 O × cal ×× v ××× h − (VV1 O ) mole Ta 760 22.41 . 01433kcal 1000w

F0=flow of methane at baseline (1pm) Kh=calibration factor (kw/mv) F =higher preset flow of methane (1pm) 1 (3) The integral of the heat release rate is V =thermopile voltage at baseline (mv) 0 the total heat release as a function of time V =thermopile voltage at higher flow (mv) 1 and is calculated by multiplying the rate by Ta=Ambient temperature (K) the data sampling frequency in minutes and P=Ambient pressure (mm Hg) summing the time from zero to two minutes. Pv=Water vapor pressure (mm Hg) (g) Criteria. The total positive heat release (2) Heat release rates may be calculated over the first two minutes of exposure for from the reading of the thermopile output each of the three or more samples tested voltage at any instant of time as: must be averaged, and the peak heat release rate for each of the samples must be aver- ()VVK− aged. The average total heat release must HRR = mbn not exceed 65 kilowatt-minutes per square 2 meter, and the average peak heat release .02323m rate must not exceed 65 kilowatts per square HRR=heat release rate (kw/m2) meter. Vb=baseline voltage (mv) (h) Report. The test report must include Vm=measured thermopile voltage (mv) the following for each specimen tested: 561

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(1) Description of the specimen. (4) If melting, sagging, delaminating, or (2) Radiant heat flux to the specimen, ex- other behavior that affects the exposed sur- pressed in W/cm2. face area or the mode of burning occurs, (3) Data giving release rates of heat (in kW/ these behaviors must be reported, together m2 ) as a function of time, either graphically with the time at which such behaviors were or tabulated at intervals no greater than 10 observed. seconds. The calibration factor (k ) must be (5) The peak heat release and the 2-minute n integrated heat release rate must be re- recorded. ported.

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FIGURES TO PART IV OF APPENDIX F

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Part V. Test Method To Determine the Smoke Part VI—Test Method To Determine the Flam- Emission Characteristics of Cabin Materials mability and Flame Propagation Characteris- tics of Thermal/Acoustic Insulation Mate- (a) Summary of Method. The specimens rials must be constructed, conditioned, and tested in the flaming mode in accordance with Use this test method to evaluate the flam- American Society of Testing and Materials mability and flame propagation characteris- (ASTM) Standard Test Method ASTM F814– tics of thermal/acoustic insulation when exposed to both a radiant heat source and a 83. flame. (b) Acceptance Criteria. The specific optical (a) Definitions. smoke density (Ds), which is obtained by ‘‘Flame propagation’’ means the furthest averaging the reading obtained after 4 min- distance of the propagation of visible flame utes with each of the three specimens, shall towards the far end of the test specimen, not exceed 200. measured from the midpoint of the ignition source flame. Measure this distance after initially applying the ignition source and before all flame on the test specimen is extin- guished. The measurement is not a determination of burn length made after the test.

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‘‘Radiant heat source’’ means an electric material encapsulated by a film covering and or air propane panel. foams. ‘‘Thermal/acoustic insulation’’ means a ‘‘Zero point’’ means the point of applica- material or system of materials used to pro- tion of the pilot burner to the test specimen. vide thermal and/or acoustic protection. Ex- (b) Test apparatus. amples include fiberglass or other batting

(1) Radiant panel test chamber. Conduct window to provide access to the movable tests in a radiant panel test chamber (see specimen platform holder. The bottom of the figure 1 above). Place the test chamber under test chamber must be a sliding steel plat- an exhaust hood to facilitate clearing the form that has provision for securing the test chamber of smoke after each test. The radi- specimen holder in a fixed and level position. ant panel test chamber must be an enclosure The chamber must have an internal chimney 55 inches (1397 mm) long by 19.5 (495 mm) with exterior dimensions of 5.1 inches (129 deep by 28 (710 mm) to 30 inches (maximum) mm) wide, by 16.2 inches (411 mm) deep by 13 (762 mm) above the test specimen. Insulate inches (330 mm) high at the opposite end of the sides, ends, and top with a fibrous ce- the chamber from the radiant energy source. ramic insulation, such as Kaowool MTM board. On the front side, provide a 52 by 12- The interior dimensions must be 4.5 inches inch (1321 by 305 mm) draft-free, high-tem- (114 mm) wide by 15.6 inches (395 mm) deep. perature, glass window for viewing the sam- The chimney must extend to the top of the ple during testing. Place a door below the chamber (see figure 2).

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(2) Radiant heat source. Mount the radiant temperatures up to 1300 °F (704 °C). An air heat energy source in a cast iron frame or propane panel must be made of a porous re- equivalent. An electric panel must have six, fractory material and have a radiation sur- 3-inch wide emitter strips. The emitter strips face of 12 by 18 inches (305 by 457 mm). The must be perpendicular to the length of the panel must be capable of operating at tem- panel. The panel must have a radiation sur- peratures up to 1,500 °F (816 °C). See figures 3a 7 1 face of 12 ⁄8 by 18 ⁄2 inches (327 by 470 mm). and 3b. The panel must be capable of operating at

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(i) Electric radiant panel. The radiant panel (i) The sliding platform serves as the hous- must be 3-phase and operate at 208 volts. A ing for test specimen placement. Brackets single-phase, 240 volt panel is also accept- may be attached (via wing nuts) to the top able. Use a solid-state power controller and lip of the platform in order to accommodate microprocessor-based controller to set the various thicknesses of test specimens. Place electric panel operating parameters. the test specimens on a sheet of Kaowool (ii) Gas radiant panel. Use propane (liquid MTM board or 1260 Standard Board (manufac- petroleum gas—2.1 UN 1075) for the radiant tured by Thermal Ceramics and available in panel fuel. The panel fuel system must con- Europe), or equivalent, either resting on the sist of a venturi-type aspirator for mixing bottom lip of the sliding platform or on the gas and air at approximately atmospheric base of the brackets. It may be necessary to pressure. Provide suitable instrumentation for monitoring and controlling the flow of use multiple sheets of material based on the fuel and air to the panel. Include an air flow thickness of the test specimen (to meet the gauge, an air flow regulator, and a gas pres- sample height requirement). Typically, these sure gauge. non-combustible sheets of material are (iii) Radiant panel placement. Mount the available in 1⁄4 inch (6 mm) thicknesses. See panel in the chamber at 30° to the horizontal figure 4. A sliding platform that is deeper specimen plane, and 71⁄2 inches above the zero than the 2-inch (50.8mm) platform shown in point of the specimen. figure 4 is also acceptable as long as the sam- (3) Specimen holding system. ple height requirement is met.

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(ii) Attach a 1⁄2 inch (13 mm) piece of (iii) Place the test specimen horizontally Kaowool MTM board or other high tempera- on the non-combustible board(s). Place a ture material measuring 411⁄2 by 81⁄4 inches steel retaining/securing frame fabricated of (1054 by 210 mm) to the back of the platform. mild steel, having a thickness of 1⁄8 inch (3.2 This board serves as a heat retainer and pro- mm) and overall dimensions of 23 by 131⁄8 tects the test specimen from excessive inches (584 by 333 mm) with a specimen open- preheating. The height of this board must ing of 19 by 103⁄4 inches (483 by 273 mm) over not impede the sliding platform movement the test specimen. The front, back, and right (in and out of the test chamber). If the plat- portions of the top flange of the frame must form has been fabricated such that the back rest on the top of the sliding platform, and side of the platform is high enough to pre- the bottom flanges must pinch all 4 sides of vent excess preheating of the specimen when the test specimen. The right bottom flange the sliding platform is out, a retainer board must be flush with the sliding platform. See is not necessary. figure 5.

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(4) Pilot Burner. The pilot burner used to (19 mm). A 3⁄4 inch (19 mm) guide (such as a ignite the specimen must be a thin strip of metal) may be soldered to the BernzomaticTM commercial propane venturi top of the burner to aid in setting the flame torch with an axially symmetric burner tip height. The overall flame length must be ap- and a propane supply tube with an orifice di- proximately 5 inches long (127 mm). Provide ameter of 0.006 inches (0.15 mm). The length a way to move the burner out of the ignition of the burner tube must be 27⁄8 inches (71 position so that the flame is horizontal and mm). The propane flow must be adjusted via at least 2 inches (50 mm) above the specimen gas pressure through an in-line regulator to plane. See figure 6. produce a blue inner cone length of 3⁄4 inch

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(5) Thermocouples. Install a 24 American (E) The graphite plate must be electrically Wire Gauge (AWG) Type K (Chromel-Alumel) heated, have a clear surface area on each thermocouple in the test chamber for tem- side of the plate of at least 2 by 2 inches (51 perature monitoring. Insert it into the by 51 mm), and be 1⁄8 inch ±1⁄16 inch thick (3.2 chamber through a small hole drilled ±1.6 mm). through the back of the chamber. Place the (F) Center the 2 transducers on opposite thermocouple so that it extends 11 inches sides of the plates at equal distances from (279 mm) out from the back of the chamber the plate. wall, 111⁄2 inches (292 mm) from the right side (G) The distance of the calorimeter to the of the chamber wall, and is 2 inches (51 mm) plate must be no less than 0.0625 inches (1.6 below the radiant panel. The use of other mm), nor greater than 0.375 inches (9.5 mm). thermocouples is optional. (H) The range used in calibration must be (6) Calorimeter. The calorimeter must be a at least 0–3.5 BTUs/ft2 second (0–3.9 Watts/ one-inch cylindrical water-cooled, total heat cm2) and no greater than 0–5.7 BTUs/ft2 sec- flux density, foil type Gardon Gage that has ond (0–6.4 Watts/cm2). a range of 0 to 5 BTU/ft2-second (0 to 5.7 (I) The recording device used must record Watts/cm2). the 2 transducers simultaneously or at least (7) Calorimeter calibration specification and within 1⁄10 of each other. procedure. (8) Calorimeter fixture. With the sliding plat- (i) Calorimeter specification. form pulled out of the chamber, install the (A) Foil diameter must be 0.25 ±0.005 inches calorimeter holding frame and place a sheet (6.35 ±0.13 mm). of non-combustible material in the bottom of the sliding platform adjacent to the hold- (B) Foil thickness must be 0.0005 ±0.0001 ing frame. This will prevent heat losses dur- inches (0.013 ±0.0025 mm). ing calibration. The frame must be 131⁄8 (C) Foil material must be thermocouple inches (333 mm) deep (front to back) by 8 grade Constantan. inches (203 mm) wide and must rest on the (D) Temperature measurement must be a top of the sliding platform. It must be fab- Copper Constantan thermocouple. ricated of 1⁄8 inch (3.2 mm) flat stock steel (E) The copper center wire diameter must and have an opening that accommodates a 1⁄2 be 0.0005 inches (0.013 mm). inch (12.7 mm) thick piece of refractory (F) The entire face of the calorimeter must board, which is level with the top of the slid- be lightly coated with ‘‘Black Velvet’’ paint ing platform. The board must have three 1- having an emissivity of 96 or greater. inch (25.4 mm) diameter holes drilled (ii) Calorimeter calibration. (A) The calibra- through the board for calorimeter insertion. tion method must be by comparison to a like The distance to the radiant panel surface standardized transducer. from the centerline of the first hole (‘‘zero’’ (B) The standardized transducer must meet position) must be 71⁄2 ±1⁄8 inches (191 ±3 mm). the specifications given in paragraph VI(b)(6) The distance between the centerline of the of this appendix. first hole to the centerline of the second hole (C) Calibrate the standard transducer must be 2 inches (51 mm). It must also be the against a primary standard traceable to the same distance from the centerline of the sec- National Institute of Standards and Tech- ond hole to the centerline of the third hole. nology (NIST). See figure 7. A calorimeter holding frame (D) The method of transfer must be a heat- that differs in construction is acceptable as ed graphite plate. long as the height from the centerline of the

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first hole to the radiant panel and the distance between holes is the same as described in this paragraph.

(9) Instrumentation. Provide a calibrated recover material is over-cut enough to be cording device with an appropriate range or drawn down the sides without compressing a computerized data acquisition system to the core material. The fastening means measure and record the outputs of the calo- should be as continuous as possible along the rimeter and the thermocouple. The data ac- length of the seams. The specimen thickness quisition system must be capable of record- must be of the same thickness as installed in ing the calorimeter output every second dur- the airplane. ing calibration. (3) Specimen Dimensions. To facilitate prop- (10) Timing device. Provide a stopwatch or er placement of specimens in the sliding other device, accurate to ±1 second/hour, to platform housing, cut non-rigid core mate- measure the time of application of the pilot rials, such as fiberglass, 121⁄2 inches (318mm) burner flame. wide by 23 inches (584mm) long. Cut rigid (c) Test specimens. (1) Specimen preparation. materials, such as foam, 111⁄2 ±1⁄4 inches (292 Prepare and test a minimum of three test mm ±6mm) wide by 23 inches (584mm) long in specimens. If an oriented film cover material order to fit properly in the sliding platform is used, prepare and test both the warp and housing and provide a flat, exposed surface fill directions. equal to the opening in the housing. (2) Construction. Test specimens must in- (d) Specimen conditioning. Condition the clude all materials used in construction of test specimens at 70 ±5 °F (21 ±2 °C) and 55% the insulation (including batting, film, ±10% relative humidity, for a minimum of 24 scrim, tape etc.). Cut a piece of core material hours prior to testing. such as foam or fiberglass, and cut a piece of (e) Apparatus Calibration. (1) With the slid- film cover material (if used) large enough to ing platform out of the chamber, install the cover the core material. Heat sealing is the calorimeter holding frame. Push the plat- preferred method of preparing fiberglass form back into the chamber and insert the samples, since they can be made without calorimeter into the first hole (‘‘zero’’ posi- compressing the fiberglass (‘‘box sample’’). tion). See figure 7. Close the bottom door lo- Cover materials that are not heat sealable cated below the sliding platform. The dis- may be stapled, sewn, or taped as long as the tance from the centerline of the calorimeter

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to the radiant panel surface at this point up to 1 hour). The pilot burner must be off must be 7.1⁄2 inches ±1⁄8 (191 mm ±3). Prior to and in the down position during this time. igniting the radiant panel, ensure that the (3) After steady-state conditions have been calorimeter face is clean and that there is reached, move the calorimeter 2 inches (51 water running through the calorimeter. mm) from the ‘‘zero’’ position (first hole) to (2) Ignite the panel. Adjust the fuel/air position 1 and record the heat flux. Move the mixture to achieve 1.5 BTUs/ft2-second ±5% calorimeter to position 2 and record the heat (1.7 Watts/cm2 ±5%) at the ‘‘zero’’ position. If flux. Allow enough time at each position for using an electric panel, set the power con- the calorimeter to stabilize. Table 1 depicts troller to achieve the proper heat flux. Allow typical calibration values at the three posi- the unit to reach steady state (this may take tions.

TABLE 1—CALIBRATION TABLE

Position BTU’s/ft2sec Watts/cm2

‘‘Zero’’ Position ...... 1.5 1.7 Position 1 ...... 1.51–1.50–1.49 1.71–1.70–1.69 Position 2 ...... 1.43–1.44 1.62–1.63

(4) Open the bottom door, remove the calo- make a slit in the film cover to purge any air rimeter and holder fixture. Use caution as inside. This allows the operator to maintain the fixture is very hot. the proper test specimen position (level with (f) Test Procedure. (1) Ignite the pilot burn- the top of the platform) and to allow ventila- er. Ensure that it is at least 2 inches (51 mm) tion of gases during testing. A longitudinal above the top of the platform. The burner slit, approximately 2 inches (51mm) in must not contact the specimen until the test length, must be centered 3 inches ±1⁄2 inch begins. (76mm±13mm) from the left flange of the se- (2) Place the test specimen in the sliding curing frame. A utility knife is acceptable platform holder. Ensure that the test sample for slitting the film cover. surface is level with the top of the platform. (4) Immediately push the sliding platform At ‘‘zero’’ point, the specimen surface must into the chamber and close the bottom door. be 71⁄2 inches ±1⁄8 inch (191 mm ±3) below the (5) Bring the pilot burner flame into con- radiant panel. tact with the center of the specimen at the (3) Place the retaining/securing frame over ‘‘zero’’ point and simultaneously start the the test specimen. It may be necessary (due timer. The pilot burner must be at a 27° to compression) to adjust the sample (up or angle with the sample and be approximately down) in order to maintain the distance from 1⁄2 inch (12 mm) above the sample. See figure the sample to the radiant panel (71⁄2 inches 7. A stop, as shown in figure 8, allows the op- ±1⁄8 inch (191 mm±3) at ‘‘zero’’ position). With erator to position the burner correctly each film/fiberglass assemblies, it is critical to time.

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(6) Leave the burner in position for 15 sec- to the left of the centerline of the pilot onds and then remove to a position at least flame application. 2 inches (51 mm) above the specimen. (2) The flame time after removal of the (g) Report. (1) Identify and describe the pilot burner may not exceed 3 seconds on any test specimen. specimen. (2) Report any shrinkage or melting of the Part VII—Test Method To Determine the test specimen. Burnthrough Resistance of Thermal/Acoustic (3) Report the flame propagation distance. Insulation Materials If this distance is less than 2 inches, report this as a pass (no measurement required). Use the following test method to evaluate the burnthrough resistance characteristics (4) Report the after-flame time. of aircraft thermal/acoustic insulation mate- (h) Requirements. (1) There must be no rials when exposed to a high intensity open flame propagation beyond 2 inches (51 mm) flame.

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(a) Definitions. the test rig, at an angle of 30° with respect to Burnthrough time means the time, in sec- vertical. onds, for the burner flame to penetrate the Specimen set means two insulation blanket test specimen, and/or the time required for specimens. Both specimens must represent the heat flux to reach 2.0 Btu/ft2sec (2.27 W/ the same production insulation blanket con- 2 cm ) on the inboard side, at a distance of 12 struction and materials, proportioned to cor- inches (30.5 cm) from the front surface of the respond to the specimen size. insulation blanket test frame, whichever is (b) Apparatus. (1) The arrangement of the sooner. The burnthrough time is measured at test apparatus is shown in figures 1 and 2 and the inboard side of each of the insulation blanket specimens. must include the capability of swinging the Insulation blanket specimen means one of burner away from the test specimen during two specimens positioned in either side of warm-up.

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(2) Test burner. The test burner must be a eters such as fuel pressure, nozzle depth, sta- modified gun-type such as the Park Model tor position, and intake airflow must be DPL 3400. Flame characteristics are highly properly adjusted to achieve the correct dependent on actual burner setup. Param- flame output.

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(i) Nozzle. A nozzle must maintain the fuel (ii) Fuel Rail. The fuel rail must be ad- pressure to yield a nominal 6.0 gal/hr (0.378 L/ justed to position the fuel nozzle at a depth min) fuel flow. A Monarch-manufactured 80° of 0.3125 inch (8 mm) from the end plane of PL (hollow cone) nozzle nominally rated at the exit stator, which must be mounted in 6.0 gal/hr at 100 lb/in2 (0.71 MPa) delivers a the end of the draft tube. proper spray pattern. (iii) Internal Stator. The internal stator, located in the middle of the draft tube, must

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be positioned at a depth of 3.75 inches (95 (vi) Fuel. Use JP–8, Jet A, or their inter- mm) from the tip of the fuel nozzle. The sta- national equivalent, at a flow rate of 6.0 ±0.2 tor must also be positioned such that the in- gal/hr (0.378 ±0.0126 L/min). If this fuel is un- tegral igniters are located at an angle mid- available, ASTM K2 fuel (Number 2 grade way between the 10 and 11 o’clock position, kerosene) or ASTM D2 fuel (Number 2 grade when viewed looking into the draft tube. fuel oil or Number 2 diesel fuel) are accept- Minor deviations to the igniter angle are ac- able if the nominal fuel flow rate, tempera- ceptable if the temperature and heat flux re- ture, and heat flux measurements conform to quirements conform to the requirements of the requirements of paragraph VII(e) of this paragraph VII(e) of this appendix. (iv) Blower Fan. The cylindrical blower fan appendix. used to pump air through the burner must (vii) Fuel pressure regulator. Provide a fuel measure 5.25 inches (133 mm) in diameter by pressure regulator, adjusted to deliver a 3.5 inches (89 mm) in width. nominal 6.0 gal/hr (0.378 L/min) flow rate. An (v) Burner cone. Install a 12 +0.125-inch (305 operating fuel pressure of 100 lb/in2 (0.71 ±3 mm) burner extension cone at the end of MPa) for a nominally rated 6.0 gal/hr 80° the draft tube. The cone must have an open- spray angle nozzle (such as a PL type) deliv- ing 6 ±0.125-inch (152 ±3 mm) high and 11 ers 6.0 ±0.2 gal/hr (0.378 ±0.0126 L/min). ±0.125-inch (280 ±3 mm) wide (see figure 3).

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(3) Calibration rig and equipment. (i) Con- perature. Position the calibration rigs to struct individual calibration rigs to incor- allow movement of the burner from the test porate a calorimeter and thermocouple rake rig position to either the heat flux or tem- for the measurement of heat flux and temperature position with minimal difficulty.

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(ii) Calorimeter. The calorimeter must be a ±3 mm) by 0.75 ±0.125 inch (19 mm ±3 mm) total heat flux, foil type Gardon Gage of an thick insulating block which is attached to appropriate range such as 0–20 Btu/ft 2-sec (0– the heat flux calibration rig during calibra- 22.7 W/cm 2), accurate to ±3% of the indicated tion (figure 4). Monitor the insulating block reading. The heat flux calibration method for deterioration and replace it when nec- must be in accordance with paragraph essary. Adjust the mounting as necessary to VI(b)(7) of this appendix. ensure that the calorimeter face is parallel (iii) Calorimeter mounting. Mount the calo- to the exit plane of the test burner cone. rimeter in a 6- by 12- ±0.125 inch (152- by 305-

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(iv) Thermocouples. Provide seven 1⁄8-inch Wire Gauge (AWG) size conductor for cali- (3.2 mm) ceramic packed, metal sheathed, bration. Attach the thermocouples to a steel type K (Chromel-alumel), grounded junction angle bracket to form a thermocouple rake thermocouples with a nominal 24 American

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for placement in the calibration rig during inch (3.2 mm) thick steel as shown in figure burner calibration (figure 5). 1, except for the center vertical former, (v) Air velocity meter. Use a vane-type air which should be 1⁄4-inch (6.4 mm) thick to velocity meter to calibrate the velocity of minimize warpage. The specimen mounting air entering the burner. An Omega Engineer- frame stringers (horizontal) should be bolted ing Model HH30A is satisfactory. Use a suit- to the test frame formers (vertical) such that able adapter to attach the measuring device the expansion of the stringers will not cause to the inlet side of the burner to prevent air the entire structure to warp. Use the mount- from entering the burner other than through ing frame for mounting the two insulation the measuring device, which would produce erroneously low readings. Use a flexible duct, blanket test specimens as shown in figure 2. measuring 4 inches wide (102 mm) by 20 feet (5) Backface calorimeters. Mount two total long (6.1 meters), to supply fresh air to the heat flux Gardon type calorimeters behind burner intake to prevent damage to the air the insulation test specimens on the back velocity meter from ingested soot. An op- side (cold) area of the test specimen mount- tional airbox permanently mounted to the ing frame as shown in figure 6. Position the burner intake area can effectively house the calorimeters along the same plane as the air velocity meter and provide a mounting burner cone centerline, at a distance of 4 port for the flexible intake duct. inches (102 mm) from the vertical centerline (4) Test specimen mounting frame. Make the of the test frame. mounting frame for the test specimens of 1⁄8-

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(i) The calorimeters must be a total heat The heat flux calibration method must com- flux, foil type Gardon Gage of an appropriate ply with paragraph VI(b)(7) of this appendix. range such as 0–5 Btu/ft2-sec (0–5.7 W/cm2), accurate to ±3% of the indicated reading.

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(6) Instrumentation. Provide a recording po- (i) Fire barrier material. If the insulation tentiometer or other suitable calibrated in- blanket is constructed with a fire barrier strument with an appropriate range to meas- material, place the fire barrier material in a ure and record the outputs of the calo- manner reflective of the installed arrange- rimeter and the thermocouples. ment For example, if the material will be (7) Timing device. Provide a stopwatch or placed on the outboard side of the insulation other device, accurate to ±1%, to measure material, inside the moisture film, place it the time of application of the burner flame the same way in the test specimen. and burnthrough time. (ii) Insulation material. Blankets that uti- (8) Test chamber. Perform tests in a suitable lize more than one variety of insulation chamber to reduce or eliminate the possi- (composition, density, etc.) must have speci- bility of test fluctuation due to air move- men sets constructed that reflect the insula- ment. The chamber must have a minimum tion combination used. If, however, several floor area of 10 by 10 feet (305 by 305 cm). blanket types use similar insulation com- (i) Ventilation hood. Provide the test cham- binations, it is not necessary to test each ber with an exhaust system capable of re- combination if it is possible to bracket the moving the products of combustion expelled various combinations. during tests. (iii) Moisture barrier film. If a production (c) Test Specimens. (1) Specimen preparation. blanket construction utilizes more than one Prepare a minimum of three specimen sets of type of moisture barrier film, perform sepa- the same construction and configuration for rate tests on each combination. For example, testing. if a polyimide film is used in conjunction (2) Insulation blanket test specimen. with an insulation in order to enhance the (i) For batt-type materials such as fiber- burnthrough capabilities, also test the same glass, the constructed, finished blanket spec- insulation when used with a polyvinyl fluo- imen assemblies must be 32 inches wide by 36 ride film. inches long (81.3 by 91.4 cm), exclusive of (iv) Installation on test frame. Attach the heat sealed film edges. blanket test specimens to the test frame (ii) For rigid and other non-conforming using 12 steel spring type clamps as shown in types of insulation materials, the finished figure 7. Use the clamps to hold the blankets test specimens must fit into the test rig in in place in both of the outer vertical such a manner as to replicate the actual in- formers, as well as the center vertical former service installation. (4 clamps per former). The clamp surfaces (3) Construction. Make each of the speci- should measure 1 inch by 2 inches (25 by 51 mens tested using the principal components mm). Place the top and bottom clamps 6 (i.e., insulation, fire barrier material if used, inches (15.2 cm) from the top and bottom of and moisture barrier film) and assembly the test frame, respectively. Place the mid- processes (representative seams and clo- dle clamps 8 inches (20.3 cm) from the top sures). and bottom clamps.

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(Note: For blanket materials that cannot blower. Measure the airflow of the test be installed in accordance with figure 7 chamber using a vane anemometer or equiv- above, the blankets must be installed in a alent measuring device. The vertical air ve- manner approved by the FAA.) locity just behind the top of the upper insu- (v) Conditioning. Condition the specimens lation blanket test specimen must be 100 ±50 at 70° ±5 °F (21° ±2 °C) and 55% ±10% relative ft/min (0.51 ±0.25 m/s). The horizontal air ve- humidity for a minimum of 24 hours prior to locity at this point must be less than 50 ft/ testing. min (0.25 m/s). (d) Preparation of apparatus. (1) Level and (3) If a calibrated flow meter is not avail- center the frame assembly to ensure align- able, measure the fuel flow rate using a grad- ment of the calorimeter and/or thermocouple uated cylinder of appropriate size. Turn on rake with the burner cone. the burner motor/fuel pump, after insuring (2) Turn on the ventilation hood for the that the igniter system is turned off. Collect test chamber. Do not turn on the burner the fuel via a plastic or rubber tube into the

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graduated cylinder for a 2-minute period. De- rimeter face. Ensure that the horizontal cen- termine the flow rate in gallons per hour. terline of the burner cone is offset 1 inch The fuel flow rate must be 6.0 ±0.2 gallons per below the horizontal centerline of the calo- hour (0.378 ±0.0126 L/min). rimeter (figure 8). Without disturbing the (e) Calibration. (1) Position the burner in calorimeter position, rotate the burner in front of the calorimeter so that it is centered front of the thermocouple rake, such that and the vertical plane of the burner cone exit the middle thermocouple (number 4 of 7) is is 4 ±0.125 inches (102 ±3 mm) from the calo- centered on the burner cone.

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Ensure that the horizontal centerline of the average temperature of each thermo- the burner cone is also offset 1 inch below couple over this 30-second period and record. the horizontal centerline of the thermo- The average temperature of each of the 7 couple tips. Re-check measurements by ro- thermocouples should be 1900 °F ±100 °F (1038 tating the burner to each position to ensure ±56 °C). proper alignment between the cone and the (6) If either the heat flux or the tempera- calorimeter and thermocouple rake. (Note: tures are not within the specified range, ad- The test burner mounting system must in- just the burner intake air velocity and re- corporate ‘‘detents’’ that ensure proper cen- tering of the burner cone with respect to peat the procedures of paragraphs (4) and (5) both the calorimeter and the thermocouple above to obtain the proper values. Ensure rakes, so that rapid positioning of the burner that the inlet air velocity is within the can be achieved during the calibration proce- range of 2150 ft/min ±50 ft/min (10.92 ±0.25 m/ dure.) s). (2) Position the air velocity meter in the (7) Calibrate prior to each test until con- adapter or airbox, making certain that no sistency has been demonstrated. After con- gaps exist where air could leak around the sistency has been confirmed, several tests air velocity measuring device. Turn on the may be conducted with calibration con- blower/motor while ensuring that the fuel so- ducted before and after a series of tests. lenoid and igniters are off. Adjust the air in- (f) Test procedure. (1) Secure the two insula- take velocity to a level of 2150 ft/min, (10.92 tion blanket test specimens to the test m/s) then turn off the blower/motor. (Note: frame. The insulation blankets should be at- The Omega HH30 air velocity meter meas- tached to the test rig center vertical former ures 2.625 inches in diameter. To calculate using four spring clamps positioned as shown the intake airflow, multiply the cross-sectional area (0.03758 ft2) by the air velocity in figure 7 (according to the criteria of para- (2150 ft/min) to obtain 80.80 ft3/min. An air graph paragraph (c)(3)(iv) of this part of this velocity meter other than the HH30 unit can appendix). be used, provided the calculated airflow of (2) Ensure that the vertical plane of the 80.80 ft3/min (2.29 m3/min) is equivalent.) burner cone is at a distance of 4 ±0.125 inch (3) Rotate the burner from the test posi- (102 ±3 mm) from the outer surface of the tion to the warm-up position. Prior to light- horizontal stringers of the test specimen ing the burner, ensure that the calorimeter frame, and that the burner and test frame face is clean of soot deposits, and there is are both situated at a 30° angle with respect water running through the calorimeter. Ex- to vertical. amine and clean the burner cone of any evi- (3) When ready to begin the test, direct the dence of buildup of products of combustion, burner away from the test position to the soot, etc. Soot buildup inside the burner warm-up position so that the flame will not cone may affect the flame characteristics impinge on the specimens prematurely. Turn and cause calibration difficulties. Since the on and light the burner and allow it to sta- burner cone may distort with time, dimen- bilize for 2 minutes. sions should be checked periodically. (4) While the burner is still rotated to the (4) To begin the test, rotate the burner into warm-up position, turn on the blower/motor, the test position and simultaneously start igniters and fuel flow, and light the burner. the timing device. Allow it to warm up for a period of 2 min- (5) Expose the test specimens to the burner utes. Move the burner into the calibration flame for 4 minutes and then turn off the position and allow 1 minute for calorimeter burner. Immediately rotate the burner out of stabilization, then record the heat flux once the test position. every second for a period of 30 seconds. Turn (6) Determine (where applicable) the off burner, rotate out of position, and allow burnthrough time, or the point at which the to cool. Calculate the average heat flux over heat flux exceeds 2.0 Btu/ft2-sec (2.27 W/cm2). this 30-second duration. The average heat (g) Report. (1) Identify and describe the flux should be 16.0 ±0.8 Btu/ft2 sec (18.2 ±0.9 W/ specimen being tested. 2 cm ). (2) Report the number of insulation blan- (5) Position the burner in front of the ther- ket specimens tested. mocouple rake. After checking for proper (3) Report the burnthrough time (if any), alignment, rotate the burner to the warm-up position, turn on the blower/motor, igniters and the maximum heat flux on the back face and fuel flow, and light the burner. Allow it of the insulation blanket test specimen, and to warm up for a period of 2 minutes. Move the time at which the maximum occurred. the burner into the calibration position and (h) Requirements. (1) Each of the two insula- allow 1 minute for thermocouple stabilization blanket test specimens must not allow tion, then record the temperature of each of fire or flame penetration in less than 4 min- the 7 thermocouples once every second for a utes. period of 30 seconds. Turn off burner, rotate (2) Each of the two insulation blanket test out of position, and allow to cool. Calculate specimens must not allow more than 2.0 Btu/

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ft2-sec (2.27 W/cm2) on the cold side of the in- similar design with extensive satisfactory sulation specimens at a point 12 inches (30.5 service experience, it will be acceptable to cm) from the face of the test rig. select Us at Vc less than 85 fps, but not less than 75 fps, with linear decrease from that [Amdt. 25–32, 37 FR 3972, Feb. 24, 1972] value at 20,000 feet to 30 fps at 80,000 feet. EDITORIAL NOTE: For FEDERAL REGISTER ci- The following factors will be taken into ac- tations affecting appendix F to Part 25, see count when assessing comparability to a the List of CFR Sections Affected, which ap- similar design: pears in the Finding Aids section of the (1) The transfer function of the new design printed volume and at www.fdsys.gov. should exhibit no unusual characteristics as compared to the similar design which will APPENDIX G TO PART 25—CONTINUOUS significantly affect response to turbulence; GUST DESIGN CRITERIA e.g., coalescence of modal response in the frequency regime which can result in a sig- The continuous gust design criteria in this nificant increase of loads. appendix must be used in establishing the (2) The typical mission of the new airplane dynamic response of the airplane to vertical is substantially equivalent to that of the and lateral continuous turbulence unless a similar design. more rational criteria is used. The following (3) The similar design should demonstrate gust load requirements apply to mission the adequacy of the Us selected. analysis and design envelope analysis: (ii) At speed VB: Us is equal to 1.32 times (a) The limit gust loads utilizing the con- the values obtained under paragraph (b)(3)(i) tinuous turbulence concept must be deter- of this appendix. 1 mined in accordance with the provisions of (iii) At speed VD: Us is equal to ⁄2 the val- either paragraph (b) or paragraphs (c) and (d) ues obtained under paragraph (b)(3)(i) of this of this appendix. appendix. (b) Design envelope analysis. The limit loads (iv) At speeds between VB and Vc and be- must be determined in accordance with the tween Vc and VD: Us is equal to a value ob- following: tained by linear interpolation. (1) All critical altitudes, weights, and (4) When a stability augmentation system weight distributions, as specified in is included in the analysis, the effect of sys- § 25.321(b), and all critical speeds within the tem nonlinearities on loads at the limit load ranges indicated in paragraph (b)(3) of this level must be realistically or conservatively appendix must be considered. accounted for. (2) Values of A¯ (ratio of root-mean-square (c) Mission analysis. Limit loads must be incremental load root-mean-square gust ve- determined in accordance with the following: locity) must be determined by dynamic anal- (1) The expected utilization of the airplane ysis. The power spectral density of the at- must be represented by one or more flight mospheric turbulence must be as given by profiles in which the load distribution and the equation— the variation with time of speed, altitude, gross weight, and center of gravity position + 8 ()Ω 2 are defined. These profiles must be divided 1 1339. L into mission segments or blocks, for anal- φσπ()Ω = 2 3 ysis, and average or effective values of the L / 11 + ()Ω 2 6 pertinent parameters defined for each seg- []1 1339. L ment. (2) For each of the mission segments de- where: fined under paragraph (c)(1) of this appendix, =power-spectral density (ft./sec.) 2/rad./ft. ¯ j values of A and No must be determined by s=root-mean-square gust velocity, ft./sec. analysis. A¯ is defined as the ratio of root- W=reduced frequency, radians per foot. mean-square incremental load to root-mean- L=2,500 ft. square gust velocity and No is the radius of (3) The limit loads must be obtained by gyration of the load power spectral density multiplying the A¯ values determined by the function about zero frequency. The power dynamic analysis by the following values of spectral density of the atmospheric turbu- the gust velocity Us™ lence must be given by the equation set forth (i) At speed Vc: Us=85 fps true gust velocity in paragraph (b)(2) of this appendix. in the interval 0 to 30,000 ft. altitude and is (3) For each of the load and stress quan- linearly decreased to 30 fps true gust veloc- tities selected, the frequency of exceedance ity at 80,000 ft. altitude. Where the Adminis- must be determined as a function of load trator finds that a design is comparable to a level by means of the equation—

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⎡ ⎛ − ⎞ ⎛ − ⎞⎤ ⎢ yyone−−gg yy one ⎥ NN()γ = ∑ t P exp⎜ − ⎟ +−P exp ⎜ ⎟ o ⎢ 1 ⎜ bA ⎟ 2 ⎜ bA ⎟⎥ ⎣ ⎝ 1 ⎠ ⎝ 2 ⎠⎦

where— ation must be given to the fraction of flight t=selected time interval. time that the system may be inoperative. y=net value of the load or stress. The flight profiles of paragraph (c)(1) of this Yone=g=value of the load or stress in one-g appendix must include flight with the sys- level flight. tem inoperative for this fraction of the flight N(y)=average number of exceedances of the time. When a stability augmentation system indicated value of the load or stress in is included in the analysis, the effect of sys- unit time. tem nonlinearities on loads at the limit load è=symbol denoting summation over all mis- level must be conservatively accounted for. sion segments. (d) Supplementary design envelope analysis. ¯ No, A=parameters determined by dynamic In addition to the limit loads defined by analysis as defined in paragraph (c)(2) of paragraph (c) of this appendix, limit loads this appendix. must also be determined in accordance with P1, P2, b1, b2=parameters defining the prob- paragraph (b) of this appendix, except that— ability distributions of root-mean-square (1) In paragraph (b)(3)(i) of this appendix, gust velocity, to be read from Figures 1 the value of Us=85 fps true gust velocity is and 2 of this appendix. replaced by Us=60 fps true gust velocity on The limit gust loads must be read from the the interval 0 to 30,000 ft. altitude, and is lin- frequency of exceedance curves at a fre- early decreased to 25 fps true gust velocity quency of exceedance of 2×10¥5 exceedances at 80,000 ft. altitude; and per hour. Both positive and negative load di- (2) In paragraph (b) of this appendix, the rections must be considered in determining reference to paragraphs (b)(3)(i) through the limit loads. (b)(3)(iii) of this appendix is to be understood (4) If a stability augmentation system is as referring to the paragraph as modified by utilized to reduce the gust loads, consider- paragraph (d)(1).

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[Amdt. 25–54, 45 FR 60173, Sept. 11, 1980]

APPENDIX H TO PART 25—INSTRUCTIONS the Instructions for Continued Airworthiness FOR CONTINUED AIRWORTHINESS for each engine and propeller (hereinafter designated ‘‘products’’), for each appliance H25.1 General. required by this chapter, and any required (a) This appendix specifies requirements information relating to the interface of for preparation of Instructions for Continued those appliances and products with the air- Airworthiness as required by §§ 25.1529, plane. If Instructions for Continued Air- 25.1729, and applicable provisions of parts 21 worthiness are not supplied by the manufac- and 26 of this chapter. turer of an appliance or product installed in (b) The Instructions for Continued Air- the airplane, the Instructions for Continued worthiness for each airplane must include Airworthiness for the airplane must include

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the information essential to the continued essary to provide for the continued air- airworthiness of the airplane. worthiness of the airplane. (c) The applicant must submit to the FAA (2) Troubleshooting information describing a program to show how changes to the In- probable malfunctions, how to recognize structions for Continued Airworthiness made those malfunctions, and the remedial action by the applicant or by the manufacturers or for those malfunctions. products and appliances installed in the air- (3) Information describing the order and plane will be distributed. method of removing and replacing products H25.2 Format. and parts with any necessary precautions to (a) The Instructions for Continued Air- be taken. worthiness must be in the form of a manual (4) Other general procedural instructions or manuals as appropriate for the quantity including procedures for system testing dur- of data to be provided. ing ground running, symmetry checks, (b) The format of the manual or manuals weighing and determining the center of grav- must provide for a practical arrangement. ity, lifting and shoring, and storage limitations. H25.3 Content. (c) Diagrams of structural access plates The contents of the manual or manuals and information needed to gain access for in- must be prepared in the English language. spections when access plates are not pro- The Instructions for Continued Airworthi- vided. ness must contain the following manuals or sections, as appropriate, and information: (d) Details for the application of special inspection techniques including radiographic (a) Airplane maintenance manual or section. and ultrasonic testing where such processes (1) Introduction information that includes an are specified. explanation of the airplane’s features and data to the extent necessary for mainte- (e) Information needed to apply protective nance or preventive maintenance. treatments to the structure after inspection. (2) A description of the airplane and its (f) All data relative to structural fasteners systems and installations including its en- such as identification, discard recommenda- gines, propellers, and appliances. tions, and torque values. (3) Basic control and operation information (g) A list of special tools needed. describing how the airplane components and H25.4 Airworthiness Limitations section. systems are controlled and how they oper- (a) The Instructions for Continued Air- ate, including any special procedures and worthiness must contain a section titled Air- limitations that apply. worthiness Limitations that is segregated (4) Servicing information that covers de- and clearly distinguishable from the rest of tails regarding servicing points, capacities of the document. This section must set forth— tanks, reservoirs, types of fluids to be used, pressures applicable to the various systems, (1) Each mandatory modification time, re- location of access panels for inspection and placement time, structural inspection inter- servicing, locations of lubrication points, lu- val, and related structural inspection proce- bricants to be used, equipment required for dure approved under § 25.571. servicing, tow instructions and limitations, (2) Each mandatory replacement time, in- mooring, jacking, and leveling information. spection interval, related inspection proce- (b) Maintenance instructions. (1) Scheduling dure, and all critical design configuration information for each part of the airplane and control limitations approved under § 25.981 its engines, auxiliary power units, propellers, for the fuel tank system. accessories, instruments, and equipment (3) Any mandatory replacement time of that provides the recommended periods at EWIS components as defined in section which they should be cleaned, inspected, ad- 25.1701. justed, tested, and lubricated, and the degree (4) A limit of validity of the engineering of inspection, the applicable wear tolerances, data that supports the structural mainte- and work recommended at these periods. nance program (LOV), stated as a total num- However, the applicant may refer to an ac- ber of accumulated flight cycles or flight cessory, instrument, or equipment manufac- hours or both, approved under § 25.571. Until turer as the source of this information if the the full-scale fatigue testing is completed applicant shows that the item has an excep- and the FAA has approved the LOV, the tionally high degree of complexity requiring number of cycles accumulated by the air- specialized maintenance techniques, test plane cannot be greater than 1⁄2 the number equipment, or expertise. The recommended of cycles accumulated on the fatigue test ar- overhaul periods and necessary cross ref- ticle. erences to the Airworthiness Limitations (b) If the Instructions for Continued Air- section of the manual must also be included. worthiness consist of multiple documents, In addition, the applicant must include an the section required by this paragraph must inspection program that includes the fre- be included in the principal manual. This quency and extent of the inspections nec- section must contain a legible statement in

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a prominent location that reads: ‘‘The Air- in the form of a document appropriate for worthiness Limitations section is FAA-ap- the information to be provided, and they proved and specifies maintenance required must be easily recognizable as EWIS ICA. under §§ 43.16 and 91.403 of the Federal Avia- This document must either contain the re- tion Regulations, unless an alternative pro- quired EWIS ICA or specifically reference gram has been FAA approved.’’ other portions of the ICA that contain this H25.5 Electrical Wiring Interconnection Sys- information. tem (EWIS) Instructions for Continued Air- worthiness. [Amdt. 25–54, 45 FR 60177, Sept. 11, 1980, as (a) The applicant must prepare Instruc- amended by Amdt. 25–68, 54 FR 34329, Aug. 18, tions for Continued Airworthiness (ICA) ap- 1989; Amdt. 25–102, 66 FR 23130, May 7, 2001; plicable to EWIS as defined by § 25.1701 that Amdt. 25–123, 72 FR 63408, Nov. 8, 2007; Amdt. are approved by the FAA and include the fol- 25–132, 75 FR 69782, Nov. 15, 2010] lowing: (1) Maintenance and inspection require- APPENDIX I TO PART 25—INSTALLATION ments for the EWIS developed with the use OF AN AUTOMATIC TAKEOFF THRUST of an enhanced zonal analysis procedure that includes: CONTROL SYSTEM (ATTCS) (i) Identification of each zone of the air- I25.1 General. plane. (ii) Identification of each zone that con- (a) This appendix specifies additional re- tains EWIS. quirements for installation of an engine (iii) Identification of each zone containing power control system that automatically EWIS that also contains combustible mate- resets thrust or power on operating engine(s) rials. in the event of any one engine failure during (iv) Identification of each zone in which takeoff. EWIS is in close proximity to both primary (b) With the ATTCS and associated sys- and back-up hydraulic, mechanical, or elec- tems functioning normally as designed, all trical flight controls and lines. applicable requirements of Part 25, except as (v) Identification of— provided in this appendix, must be met with- (A) Tasks, and the intervals for performing out requiring any action by the crew to in- those tasks, that will reduce the likelihood crease thrust or power. of ignition sources and accumulation of combustible material, and I25.2 Definitions. (B) Procedures, and the intervals for per- (a) Automatic Takeoff Thrust Control System forming those procedures, that will effec- (ATTCS). An ATTCS is defined as the entire tively clean the EWIS components of com- automatic system used on takeoff, including bustible material if there is not an effective all devices, both mechanical and electrical, task to reduce the likelihood of combustible that sense engine failure, transmit signals, material accumulation. actuate fuel controls or power levers or in- (vi) Instructions for protections and cau- crease engine power by other means on oper- tion information that will minimize con- ating engines to achieve scheduled thrust or tamination and accidental damage to EWIS, power increases, and furnish cockpit infor- as applicable, during performance of mainte- mation on system operation. nance, alteration, or repairs. (2) Acceptable EWIS maintenance prac- (b) Critical Time Interval. When conducting tices in a standard format. an ATTCS takeoff, the critical time interval (3) Wire separation requirements as deter- is between V1 minus 1 second and a point on mined under § 25.1707. the minimum performance, all-engine flight (4) Information explaining the EWIS iden- path where, assuming a simultaneous occur- tification method and requirements for iden- rence of an engine and ATTCS failure, the tifying any changes to EWIS under § 25.1711. resulting minimum flight path thereafter (5) Electrical load data and instructions for intersects the Part 25 required actual flight updating that data. path at no less than 400 feet above the take- (b) The EWIS ICA developed in accordance off surface. This time interval is shown in with the requirements of H25.5(a)(1) must be the following illustration:

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I25.3 Performance and System Reliability Re- during takeoff with the ATTCS system func- quirements. tioning. The applicant must comply with the per- I25.4 Thrust Setting. formance and ATTCS reliability require- The initial takeoff thrust or power setting ments as follows: on each engine at the beginning of the take- (a) An ATTCS failure or a combination of off roll may not be less than any of the fol- failures in the ATTCS during the critical lowing: time interval: (a) Ninety (90) percent of the thrust or (1) Shall not prevent the insertion of the power set by the ATTCS (the maximum maximum approved takeoff thrust or power, or takeoff thrust or power approved for the air- must be shown to be an improbable event. plane under existing ambient conditions); (2) Shall not result in a significant loss or (b) That required to permit normal oper- reduction in thrust or power, or must be ation of all safety-related systems and equip- shown to be an extremely improbable event. ment dependent upon engine thrust or power (b) The concurrent existence of an ATTCS lever position; or failure and an engine failure during the crit- (c) That shown to be free of hazardous en- ical time interval must be shown to be ex- gine response characteristics when thrust or tremely improbable. power is advanced from the initial takeoff (c) All applicable performance require- thrust or power to the maximum approved ments of Part 25 must be met with an engine takeoff thrust or power. failure occurring at the most critical point I25.5 Powerplant Controls.

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(a) In addition to the requirements of (b) The airplane must be in a normal atti- § 25.1141, no single failure or malfunction, or tude with landing gear extended. probable combination thereof, of the ATTCS, (c) Unless the airplane is equipped with an including associated systems, may cause the off-wing descent means, stands or ramps may failure of any powerplant function necessary be used for descent from the wing to the for safety. ground. Safety equipment such as mats or (b) The ATTCS must be designed to: inverted life rafts may be placed on the floor (1) Apply thrust or power on the operating or ground to protect participants. No other engine(s), following any one engine failure equipment that is not part of the emergency during takeoff, to achieve the maximum ap- evacuation equipment of the airplane may be proved takeoff thrust or power without ex- used to aid the participants in reaching the ceeding engine operating limits; ground. (2) Permit manual decrease or increase in (d) Except as provided in paragraph (a) of thrust or power up to the maximum takeoff this appendix, only the airplane’s emergency thrust or power approved for the airplane lighting system may provide illumination. under existing conditions through the use of (e) All emergency equipment required for the power lever. For airplanes equipped with the planned operation of the airplane must limiters that automatically prevent engine be installed. operating limits from being exceeded under (f) Each internal door or curtain must be existing ambient conditions, other means in the takeoff configuration. may be used to increase the thrust or power (g) Each crewmember must be seated in in the event of an ATTCS failure provided the normally assigned seat for takeoff and the means is located on or forward of the must remain in the seat until receiving the power levers; is easily identified and oper- signal for commencement of the demonstra- ated under all operating conditions by a sin- tion. Each crewmember must be a person gle action of either pilot with the hand that having knowledge of the operation of exits is normally used to actuate the power levers; and emergency equipment and, if compliance and meets the requirements of § 25.777 (a), with § 121.291 is also being demonstrated, (b), and (c); each flight attendant must be a member of a (3) Provide a means to verify to the regularly scheduled line crew. flightcrew before takeoff that the ATTCS is (h) A representative passenger load of per- in a condition to operate; and sons in normal health must be used as fol- (4) Provide a means for the flightcrew to lows: deactivate the automatic function. This (1) At least 40 percent of the passenger load means must be designed to prevent inad- must be female. vertent deactivation. (2) At least 35 percent of the passenger load I25.6 Powerplant Instruments. must be over 50 years of age. (3) At least 15 percent of the passenger load In addition to the requirements of § 25.1305: must be female and over 50 years of age. (a) A means must be provided to indicate (4) Three life-size dolls, not included as when the ATTCS is in the armed or ready part of the total passenger load, must be car- condition; and ried by passengers to simulate live infants 2 (b) If the inherent flight characteristics of years old or younger. the airplane do not provide adequate warn- (5) Crewmembers, mechanics, and training ing that an engine has failed, a warning sys- personnel, who maintain or operate the air- tem that is independent of the ATTCS must plane in the normal course of their duties, be provided to give the pilot a clear warning may not be used as passengers. of any engine failure during takeoff. (i) No passenger may be assigned a specific [Amdt. 25–62, 52 FR 43156, Nov. 9, 1987] seat except as the Administrator may require. Except as required by subparagraph APPENDIX J TO PART 25—EMERGENCY (g) of this paragraph, no employee of the ap- EVACUATION plicant may be seated next to an emergency exit. The following test criteria and procedures (j) Seat belts and shoulder harnesses (as re- must be used for showing compliance with quired) must be fastened. § 25.803: (k) Before the start of the demonstration, (a) The emergency evacuation must be con- approximately one-half of the total average ducted with exterior ambient light levels of amount of carry-on baggage, blankets, pil- no greater than 0.3 foot-candles prior to the lows, and other similar articles must be dis- activation of the airplane emergency light- tributed at several locations in aisles and ing system. The source(s) of the initial exte- emergency exit access ways to create minor rior ambient light level may remain active obstructions. or illuminated during the actual demonstra- (l) No prior indication may be given to any tion. There must, however, be no increase in crewmember or passenger of the particular the exterior ambient light level except for exits to be used in the demonstration. that due to activation of the airplane emer- (m) The applicant may not practice, re- gency lighting system. hearse, or describe the demonstration for the

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participants nor may any participant have APPENDIX K TO PART 25—EXTENDED taken part in this type of demonstration OPERATIONS (ETOPS) within the preceding 6 months. (n) Prior to entering the demonstration This appendix specifies airworthiness re- aircraft, the passengers may also be advised quirements for the approval of an airplane- to follow directions of crewmembers but may engine combination for extended operations not be instructed on the procedures to be fol- (ETOPS). For two-engine airplanes, the applicant must comply with sections K25.1 and lowed in the demonstration, except with re- K25.2 of this appendix. For airplanes with spect to safety procedures in place for the more than two engines, the applicant must demonstration or which have to do with the comply with sections K25.1 and K25.3 of this demonstration site. Prior to the start of the appendix. demonstration, the pre-takeoff passenger K25.1 Design requirements. briefing required by § 121.571 may be given. K25.1.1 Part 25 compliance. Flight attendants may assign demonstration The airplane-engine combination must subjects to assist persons from the bottom of comply with the requirements of part 25 con- a slide, consistent with their approved train- sidering the maximum flight time and the ing program. longest diversion time for which the appli- (o) The airplane must be configured to pre- cant seeks approval. vent disclosure of the active emergency exits K25.1.2 Human factors. to demonstration participants in the air- An applicant must consider crew workload, plane until the start of the demonstration. operational implications, and the crew’s and passengers’ physiological needs during con- (p) Exits used in the demonstration must tinued operation with failure effects for the consist of one exit from each exit pair. The longest diversion time for which it seeks ap- demonstration may be conducted with the proval. escape slides, if provided, inflated and the K25.1.3 Airplane systems. exits open at the beginning of the dem- (a) Operation in icing conditions. onstration. In this case, all exits must be (1) The airplane must be certificated for configured such that the active exits are not operation in icing conditions in accordance disclosed to the occupants. If this method is with § 25.1419. used, the exit preparation time for each exit (2) The airplane must be able to safely con- utilized must be accounted for, and exits duct an ETOPS diversion with the most crit- that are not to be used in the demonstration ical ice accretion resulting from: must not be indicated before the demonstra- (i) Icing conditions encountered at an alti- tion has started. The exits to be used must tude that the airplane would have to fly fol- be representative of all of the emergency lowing an engine failure or cabin decompres- exits on the airplane and must be designated sion. by the applicant, subject to approval by the (ii) A 15-minute hold in the continuous maximum icing conditions specified in Ap- Administrator. At least one floor level exit pendix C of this part with a liquid water con- must be used. tent factor of 1.0. (q) Except as provided in paragraph (c) of (iii) Ice accumulated during approach and this section, all evacuees must leave the air- landing in the icing conditions specified in plane by a means provided as part of the air- Appendix C of this part. plane’s equipment. (b) Electrical power supply. The airplane (r) The applicant’s approved procedures must be equipped with at least three inde- must be fully utilized, except the flightcrew pendent sources of electrical power. must take no active role in assisting others (c) Time limited systems. The applicant must inside the cabin during the demonstration. define the system time capability of each (s) The evacuation time period is com- ETOPS significant system that is time-lim- pleted when the last occupant has evacuated ited. the airplane and is on the ground. Provided K25.1.4 Propulsion systems. that the acceptance rate of the stand or (a) Fuel system design. Fuel necessary to complete an ETOPS flight (including a diver- ramp is no greater than the acceptance rate sion for the longest time for which the appli- of the means available on the airplane for de- cant seeks approval) must be available to the scent from the wing during an actual crash operating engines at the pressure and fuel- situation, evacuees using stands or ramps al- flow required by § 25.955 under any airplane lowed by paragraph (c) of this appendix are failure condition not shown to be extremely considered to be on the ground when they are improbable. Types of failures that must be on the stand or ramp. considered include, but are not limited to: [Amdt. 25–72, 55 FR 29788, July 20, 1990, as crossfeed valve failures, automatic fuel man- amended by Amdt. 25–79, Aug. 26, 1993; Amdt. agement system failures, and normal electrical power generation failures. 25–117, 69 FR 67499, Nov. 17, 2004] (1) If the engine has been certified for limited operation with negative engine-fuel-

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pump-inlet pressures, the following require- plane up to the maximum diversion time ments apply: being approved. (i) Airplane demonstration-testing must (b) Required markings or placards. cover worst case cruise and diversion condi- (c) The airborne equipment required for ex- tions involving: tended operations and flightcrew operating (A) Fuel grade and temperature. procedures for this equipment. (B) Thrust or power variations. (d) The system time capability for the fol- (C) Turbulence and negative G. lowing: (D) Fuel system components degraded within their approved maintenance limits. (1) The most limiting fire suppression sys- (ii) Unusable-fuel quantity in the suction tem for Class C cargo or baggage compart- feed configuration must be determined in ac- ments. cordance with § 25.959. (2) The most limiting ETOPS significant (2) For two-engine airplanes to be certifi- system other than fire suppression systems cated for ETOPS beyond 180 minutes, one for Class C cargo or baggage compartments. fuel boost pump in each main tank and at (e) This statement: ‘‘The type-design reli- least one crossfeed valve, or other means for ability and performance of this airplane-en- transferring fuel, must be powered by an gine combination has been evaluated under independent electrical power source other 14 CFR 25.1535 and found suitable for (iden- than the three power sources required to tify maximum approved diversion time) ex- comply with section K25.1.3(b) of this appen- tended operations (ETOPS) when the con- dix. This requirement does not apply if the figuration, maintenance, and procedures normal fuel boost pressure, crossfeed valve standard contained in (identify the CMP doc- actuation, or fuel transfer capability is not ument) are met. The actual maximum ap- provided by electrical power. proved diversion time for this airplane may (3) An alert must be displayed to the be less based on its most limiting system flightcrew when the quantity of fuel avail- time capability. This finding does not con- able to the engines falls below the level re- stitute operational approval to conduct quired to fly to the destination. The alert ETOPS.’’ must be given when there is enough fuel re- K25.2. Two-engine airplanes. maining to safely complete a diversion. This An applicant for ETOPS type design ap- alert must account for abnormal fuel man- proval of a two-engine airplane must use one agement or transfer between tanks, and pos- of the methods described in section K25.2.1, sible loss of fuel. This paragraph does not K25.2.2, or K25.2.3 of this appendix. apply to airplanes with a required flight en- K25.2.1 Service experience method. gineer. An applicant for ETOPS type design ap- (b) APU design. If an APU is needed to com- proval using the service experience method ply with this appendix, the applicant must must comply with sections K25.2.1(a) and demonstrate that: K25.2.1(b) of this appendix before conducting (1) The reliability of the APU is adequate the assessments specified in sections to meet those requirements; and K25.2.1(c) and K25.2.1(d) of this appendix, and (2) If it is necessary that the APU be able the flight test specified in section K25.2.1(e) to start in flight, it is able to start at any al- of this appendix. titude up to the maximum operating altitude of the airplane, or 45,000 feet, whichever is (a) Service experience. The world fleet for lower, and run for the remainder of any the airplane-engine combination must accu- flight . mulate a minimum of 250,000 engine-hours. (c) Engine oil tank design. The engine oil The FAA may reduce this number of hours if tank filler cap must comply with § 33.71(c)(4) the applicant identifies compensating fac- of this chapter. tors that are acceptable to the FAA. The K25.1.5 Engine-condition monitoring. compensating factors may include experi- Procedures for engine-condition moni- ence on another airplane, but experience on toring must be specified and validated in ac- the candidate airplane must make up a sig- cordance with Part 33, Appendix A, para- nificant portion of the total service experi- graph A33.3(c) of this chapter. ence. K25.1.6 Configuration, maintenance, and (b) In-flight shutdown (IFSD) rates. The procedures. demonstrated 12-month rolling average IFSD The applicant must list any configuration, rate for the world fleet of the airplane-en- operating and maintenance requirements, gine combination must be commensurate hardware life limits, MMEL constraints, and with the level of ETOPS approval being ETOPS approval in a CMP document. sought. K25.1.7 Airplane flight manual. (1) For type design approval up to and in- The airplane flight manual must contain cluding 120 minutes: An IFSD rate of 0.05 or the following information applicable to the less per 1,000 world-fleet engine-hours, unless ETOPS type design approval: otherwise approved by the FAA. Unless the (a) Special limitations, including any limi- IFSD rate is 0.02 or less per 1,000 world-fleet tation associated with operation of the air- engine-hours, the applicant must provide a

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list of corrective actions in the CMP docu- number of cycles, and the distribution of ment specified in section K25.1.6 of this ap- hours and cycles. pendix, that, when taken, would result in an (v) The mean time between failures IFSD rate of 0.02 or less per 1,000 fleet en- (MTBF) of propulsion system components gine-hours. that affect reliability. (2) For type design approval up to and in- (vi) A history of the IFSD rates since including 180 minutes: An IFSD rate of 0.02 or troduction into service using a 12-month less per 1,000 world-fleet engine-hours, unless rolling average. otherwise approved by the FAA. If the air- (2) The cause or potential cause of each plane-engine combination does not meet this item listed in K25.2.1(c)(1)(i) must have a cor- rate by compliance with an existing 120- rective action or actions that are shown to minute CMP document, then new or addi- be effective in preventing future occur- tional CMP requirements that the applicant rences. Each corrective action must be iden- has demonstrated would achieve this IFSD tified in the CMP document specified in sec- rate must be added to the CMP document. tion K25.1.6. A corrective action is not re- (3) For type design approval beyond 180 quired: minutes: An IFSD rate of 0.01 or less per 1,000 (i) For an item where the manufacturer is fleet engine-hours unless otherwise approved unable to determine a cause or potential by the FAA. If the airplane-engine combina- cause. tion does not meet this rate by compliance (ii) For an event where it is technically with an existing 120-minute or 180-minute unfeasible to develop a corrective action. CMP document, then new or additional CMP (iii) If the world-fleet IFSD rate— requirements that the applicant has dem- (A) Is at or below 0.02 per 1,000 world-fleet engine-hours for approval up to and includ- onstrated would achieve this IFSD rate must ing 180-minute ETOPS; or be added to the CMP document. (B) Is at or below 0.01 per 1,000 world-fleet (c) Propulsion system assessment. (1) The ap- engine-hours for approval greater than 180- plicant must conduct a propulsion system minute ETOPS. assessment based on the following data col- (d) Airplane systems assessment. The appli- lected from the world-fleet of the airplane- cant must conduct an airplane systems as- engine combination: sessment. The applicant must show that the (i) A list of all IFSD’s, unplanned ground airplane systems comply with § 25.1309(b) engine shutdowns, and occurrences (both using available in-service reliability data for ground and in-flight) when an engine was not ETOPS significant systems on the candidate shut down, but engine control or the desired airplane-engine combination. Each cause or thrust or power level was not achieved, in- potential cause of a relevant design, manu- cluding engine flameouts. Planned IFSD’s facturing, operational, and maintenance performed during flight training need not be problem occurring in service must have a included. For each item, the applicant must corrective action or actions that are shown provide— to be effective in preventing future occur- (A) Each airplane and engine make, model, rences. Each corrective action must be iden- and serial number; tified in the CMP document specified in sec- (B) Engine configuration, and major alter- tion K25.1.6 of this appendix. A corrective ac- ation history; tion is not required if the problem would not (C) Engine position; significantly impact the safety or reliability (D) Circumstances leading up to the engine of the airplane system involved. A relevant shutdown or occurrence; problem is a problem with an ETOPS group (E) Phase of flight or ground operation; 1 significant system that has or could result (F) Weather and other environmental con- in, an IFSD or diversion. The applicant must ditions; and include in this assessment relevant problems (G) Cause of engine shutdown or occur- with similar or identical equipment installed rence. on other types of airplanes to the extent (ii) A history of unscheduled engine re- such information is reasonably available. moval rates since introduction into service (e) Airplane flight test. The applicant must (using 6- and 12-month rolling averages), conduct a flight test to validate the with a summary of the major causes for the flightcrew’s ability to safely conduct an removals. ETOPS diversion with an inoperative engine (iii) A list of all propulsion system events and worst-case ETOPS Significant System (whether or not caused by maintenance or failures and malfunctions that could occur in flightcrew error), including dispatch delays, service. The flight test must validate the air- cancellations, aborted takeoffs, turnbacks, plane’s flying qualities and performance diversions, and flights that continue to des- with the demonstrated failures and malfunc- tination after the event. tions. (iv) The total number of engine hours and K25.2.2 Early ETOPS method. cycles, the number of hours for the engine An applicant for ETOPS type design ap- with the highest number of hours, the num- proval using the Early ETOPS method must ber of cycles for the engine with the highest comply with the following requirements:

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(a) Assessment of relevant experience with air- (ii) Completely disassembled and the pro- planes previously certificated under part 25. pulsion system hardware inspected to deter- The applicant must identify specific correc- mine whether it meets the service limits tive actions taken on the candidate airplane specified in the Instructions for Continued to prevent relevant design, manufacturing, Airworthiness submitted in compliance with operational, and maintenance problems ex- § 25.1529. perienced on airplanes previously certifi- (2) The applicant must identify, track, and cated under part 25 manufactured by the ap- resolve each cause or potential cause of plicant. Specific corrective actions are not IFSD, loss of thrust control, or other power required if the nature of a problem is such loss encountered during this inspection in that the problem would not significantly im- accordance with the problem tracking and pact the safety or reliability of the airplane resolution system specified in section K25.2.2 system involved. A relevant problem is a (h) of this appendix. problem with an ETOPS group 1 significant (e) New technology testing. Technology new system that has or could result in an IFSD to the applicant, including substantially new or diversion. The applicant must include in manufacturing techniques, must be tested to this assessment relevant problems of sup- substantiate its suitability for the airplane plier-provided ETOPS group 1 significant design. systems and similar or identical equipment (f) APU validation test. If an APU is needed used on airplanes built by other manufactur- to comply with this appendix, one APU of ers to the extent such information is reason- the type to be certified with the airplane ably available. must be tested for 3,000 equivalent airplane (b) Propulsion system design. (1) The engine operational cycles. Following completion of used in the applicant’s airplane design must the test, the APU must be disassembled and be approved as eligible for Early ETOPS in inspected. The applicant must identify, accordance with § 33.201 of this chapter. track, and resolve each cause or potential (2) The applicant must design the propul- cause of an inability to start or operate the sion system to preclude failures or malfunc- APU in flight as intended in accordance with tions that could result in an IFSD. The ap- the problem tracking and resolution system plicant must show compliance with this re- specified in section K25.2.2(h) of this appen- quirement by analysis, test, in-service expe- dix. rience on other airplanes, or other means ac- (g) Airplane demonstration. For each air- ceptable to the FAA. If analysis is used, the plane-engine combination to be approved for applicant must show that the propulsion sys- ETOPS, the applicant must flight test at tem design will minimize failures and mal- least one airplane to demonstrate that the functions with the objective of achieving the airplane, and its components and equipment following IFSD rates: are capable of functioning properly during (i) An IFSD rate of 0.02 or less per 1,000 ETOPS flights and diversions of the longest world-fleet engine-hours for type design ap- duration for which the applicant seeks approval up to and including 180 minutes. proval. This flight testing may be performed (ii) An IFSD rate of 0.01 or less per 1,000 in conjunction with, but may not substitute world-fleet engine-hours for type design ap- for the flight testing required by § 21.35(b)(2) proval beyond 180 minutes. of this chapter. (c) Maintenance and operational procedures. (1) The airplane demonstration flight test The applicant must validate all maintenance program must include: and operational procedures for ETOPS sig- (i) Flights simulating actual ETOPS, in- nificant systems. The applicant must iden- cluding flight at normal cruise altitude, step tify, track, and resolve any problems found climbs, and, if applicable, APU operation. during the validation in accordance with the (ii) Maximum duration flights with max- problem tracking and resolution system imum duration diversions. specified in section K25.2.2(h) of this appen- (iii) Maximum duration engine-inoperative dix. diversions distributed among the engines in- (d) Propulsion system validation test. (1) The stalled on the airplanes used for the airplane installed engine configuration for which ap- demonstration flight test program. At least proval is being sought must comply with two one-engine-inoperative diversions must § 33.201(c) of this chapter. The test engine be conducted at maximum continuous thrust must be configured with a complete airplane or power using the same engine. nacelle package, including engine-mounted (iv) Flights under non-normal conditions equipment, except for any configuration dif- to demonstrate the flightcrew’s ability to ferences necessary to accommodate test safely conduct an ETOPS diversion with stand interfaces with the engine nacelle worst-case ETOPS significant system fail- package. At the conclusion of the test, the ures or malfunctions that could occur in propulsion system must be— service. (i) Visually inspected according to the ap- (v) Diversions to airports that represent plicant’s on-wing inspection recommenda- airports of the types used for ETOPS diver- tions and limits; and sions.

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(vi) Repeated exposure to humid and in- system specified in section K25.2.2(h) of this clement weather on the ground followed by a appendix. long-duration flight at normal cruise alti- (h) Problem tracking and resolution system. tude. (1) The applicant must establish and main- (2) The airplane demonstration flight test tain a problem tracking and resolution sys- program must validate the adequacy of the tem. The system must: airplane’s flying qualities and performance, (i) Contain a process for prompt reporting and the flightcrew’s ability to safely conduct to the responsible FAA aircraft certification an ETOPS diversion under the conditions office of each occurrence reportable under specified in section K25.2.2(g)(1) of this ap- § 21.4(a)(6) encountered during the phases of pendix. airplane and engine development used to as- (3) During the airplane demonstration sess Early ETOPS eligibility. flight test program, each test airplane must be operated and maintained using the appli- (ii) Contain a process for notifying the re- cant’s recommended operating and mainte- sponsible FAA aircraft certification office of nance procedures. each proposed corrective action that the ap- (4) At the completion of the airplane dem- plicant determines necessary for each prob- onstration flight test program, each ETOPS lem identified from the occurrences reported significant system must undergo an on-wing under section K25.2.2. (h)(1)(i) of this appen- inspection or test in accordance with the dix. The timing of the notification must per- tasks defined in the proposed Instructions mit appropriate FAA review before taking for Continued Airworthiness to establish its the proposed corrective action. condition for continued safe operation. Each (2) If the applicant is seeking ETOPS type engine must also undergo a gas path inspec- design approval of a change to an airplane- tion. These inspections must be conducted in engine combination previously approved for a manner to identify abnormal conditions ETOPS, the problem tracking and resolution that could result in an IFSD or diversion. system need only address those problems The applicant must identify, track and re- specified in the following table, provided the solve any abnormal conditions in accordance applicant obtains prior authorization from with the problem tracking and resolution the FAA:

If the change does not require a new airplane type certificiate Then the Problem Tracking and Resolution System must ad- and . . . dress . . .

(i) Requires a new engine type certificate ...... All problems applicable to the new engine installation, and for the remainder of the airplane, problems in changed systems only. (ii) Does not require a new engine type certificate ...... Problems in changed systems only.

(i) Acceptance criteria. The type and fre- An applicant for ETOPS type design ap- quency of failures and malfunctions on proval of an airplane with more than two en- ETOPS significant systems that occur dur- gines must use one of the methods described ing the airplane flight test program and the in section K25.3.1, K25.3.2, or K25.3.3 of this airplane demonstration flight test program appendix. specified in section K25.2.2(g) of this appen- K25.3.1 Service experience method. dix must be consistent with the type and fre- An applicant for ETOPS type design ap- quency of failures and malfunctions that proval using the service experience method would be expected to occur on currently cer- must comply with section K25.3.1(a) of this tificated airplanes approved for ETOPS. appendix before conducting the airplane sys- K25.2.3. Combined service experience and tems assessment specified in K25.3.1(b), and Early ETOPS method. the flight test specified in section K25.3.1(c) An applicant for ETOPS type design ap- of this appendix. proval using the combined service experience (a) Service experience. The world fleet for and Early ETOPS method must comply with the airplane-engine combination must accu- the following requirements. mulate a minimum of 250,000 engine-hours. (a) A service experience requirement of not The FAA may reduce this number of hours if less than 15,000 engine-hours for the world the applicant identifies compensating fac- fleet of the candidate airplane-engine com- tors that are acceptable to the FAA. The bination. compensating factors may include experi- (b) The Early ETOPS requirements of ence on another airplane, but experience on K25.2.2, except for the airplane demonstra- the candidate airplane must make up a sig- tion specified in section K25.2.2(g) of this ap- nificant portion of the total required service pendix; and experience. (c) The flight test requirement of section (b) Airplane systems assessment. The appli- K25.2.1(e) of this appendix. cant must conduct an airplane systems as- K25.3. Airplanes with more than two engines. sessment. The applicant must show that the

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airplane systems comply with the § 25.1309(b) airplane, and its components and equipment using available in-service reliability data for are capable of functioning properly during ETOPS significant systems on the candidate ETOPS flights and diversions of the longest airplane-engine combination. Each cause or duration for which the applicant seeks ap- potential cause of a relevant design, manu- proval. This flight testing may be performed facturing, operational or maintenance prob- in conjunction with, but may not substitute lem occurring in service must have a correc- for the flight testing required by § 21.35(b)(2). tive action or actions that are shown to be (1) The airplane demonstration flight test effective in preventing future occurrences. program must include: Each corrective action must be identified in (i) Flights simulating actual ETOPS in- the CMP document specified in section cluding flight at normal cruise altitude, step K25.1.6 of this appendix. A corrective action climbs, and, if applicable, APU operation. is not required if the problem would not sig- (ii) Maximum duration flights with max- nificantly impact the safety or reliability of imum duration diversions. the airplane system involved. A relevant (iii) Maximum duration engine-inoperative problem is a problem with an ETOPS group diversions distributed among the engines in- 1 significant system that has or could result stalled on the airplanes used for the airplane in an IFSD or diversion. The applicant must demonstration flight test program. At least include in this assessment relevant problems two one engine-inoperative diversions must with similar or identical equipment installed be conducted at maximum continuous thrust on other types of airplanes to the extent or power using the same engine. such information is reasonably available. (iv) Flights under non-normal conditions (c) Airplane flight test. The applicant must to validate the flightcrew’s ability to safely conduct a flight test to validate the conduct an ETOPS diversion with worst-case flightcrew’s ability to safely conduct an ETOPS significant system failures or mal- ETOPS diversion with an inoperative engine functions that could occur in service. and worst-case ETOPS significant system (v) Diversions to airports that represent failures and malfunctions that could occur in airports of the types used for ETOPS diver- service. The flight test must validate the air- sions. plane’s flying qualities and performance (vi) Repeated exposure to humid and in- with the demonstrated failures and malfunc- clement weather on the ground followed by a tions. long duration flight at normal cruise alti- K25.3.2 Early ETOPS method. tude. An applicant for ETOPS type design ap- (2) The airplane demonstration flight test proval using the Early ETOPS method must program must validate the adequacy of the comply with the following requirements: airplane’s flying qualities and performance, (a) Maintenance and operational procedures. and the flightcrew’s ability to safely conduct The applicant must validate all maintenance an ETOPS diversion under the conditions and operational procedures for ETOPS sig- specified in section K25.3.2(d)(1) of this ap- nificant systems. The applicant must iden- pendix. tify, track and resolve any problems found (3) During the airplane demonstration during the validation in accordance with the flight test program, each test airplane must problem tracking and resolution system be operated and maintained using the appli- specified in section K25.3.2(e) of this appen- cant’s recommended operating and mainte- dix. nance procedures. (b) New technology testing. Technology new (4) At the completion of the airplane dem- to the applicant, including substantially new onstration, each ETOPS significant system manufacturing techniques, must be tested to must undergo an on-wing inspection or test substantiate its suitability for the airplane in accordance with the tasks defined in the design. proposed Instructions for Continued Air- (c) APU validation test. If an APU is needed worthiness to establish its condition for con- to comply with this appendix, one APU of tinued safe operation. Each engine must also the type to be certified with the airplane undergo a gas path inspection. These inspec- must be tested for 3,000 equivalent airplane tions must be conducted in a manner to iden- operational cycles. Following completion of tify abnormal conditions that could result in the test, the APU must be disassembled and an IFSD or diversion. The applicant must inspected. The applicant must identify, identify, track and resolve any abnormal track, and resolve each cause or potential conditions in accordance with the problem cause of an inability to start or operate the tracking and resolution system specified in APU in flight as intended in accordance with section K25.3.2(e) of this appendix. the problem tracking and resolution system (e) Problem tracking and resolution system. specified in section K25.3.2(e) of this appen- (1) The applicant must establish and main- dix. tain a problem tracking and resolution sys- (d) Airplane demonstration. For each air- tem. The system must: plane-engine combination to be approved for (i) Contain a process for prompt reporting ETOPS, the applicant must flight test at to the responsible FAA aircraft certification least one airplane to demonstrate that the office of each occurrence reportable under

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§ 21.4(a)(6) encountered during the phases of appropriate FAA review before taking the airplane and engine development used to as- proposed corrective action. sess Early ETOPS eligibility. (2) If the applicant is seeking ETOPS type (ii) Contain a process for notifying the redesign approval of a change to an airplane- sponsible FAA aircraft certification office of engine combination previously approved for each proposed corrective action that the ap- ETOPS, the problem tracking and resolution plicant determines necessary for each prob- system need only address those problems lem identified from the occurrences reported specified in the following table, provided the under section K25.3.2(h)(1)(i) of this appendix. applicant obtains prior authorization from The timing of the notification must permit the FAA:

If the change does not require a new airplane type certificate Then the Problem Tracking and Resolution System must ad- and . . . dress . . .

(f) Acceptance criteria. The type and fre- TABLE I.—HIRF ENVIRONMENT I quency of failures and malfunctions on ETOPS significant systems that occur dur- Field strength ing the airplane flight test program and the Frequency (volts/meter) airplane demonstration flight test program Peak Average specified in section K25.3.2(d) of this appendix must be consistent with the type and fre- 10 kHz–2 MHz ...... 50 50 2 MHz–30 MHz ...... 100 100 quency of failures and malfunctions that 30 MHz–100 MHz ...... 50 50 would be expected to occur on currently cer- 100 MHz–400 MHz ...... 100 100 tificated airplanes approved for ETOPS. 400 MHz–700 MHz ...... 700 50 K25.3.3 Combined service experience and 700 MHz–1 GHz ...... 700 100 Early ETOPS method. 1 GHz–2 GHz ...... 2,000 200 2 GHz–6 GHz ...... 3,000 200 An applicant for ETOPS type design ap- 6 GHz–8 GHz ...... 1,000 200 proval using the Early ETOPS method must 8 GHz–12 GHz ...... 3,000 300 comply with the following requirements: 12 GHz–18 GHz ...... 2,000 200 (a) A service experience requirement of 18 GHz–40 GHz ...... 600 200 less than 15,000 engine-hours for the world In this table, the higher field strength applies at the fre- fleet of the candidate airplane-engine com- quency band edges. bination; (b) HIRF environment II is specified in the (b) The Early ETOPS requirements of sec- following table: tion K25.3.2 of this appendix, except for the airplane demonstration specified in section TABLE II.–HIRF ENVIRONMENT II K25.3.2(d) of this appendix; and Field strength (c) The flight test requirement of section (volts/meter) K25.3.1(c) of this appendix. Frequency Peak Average [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, 2007] 10 kHz–500 kHz ...... 20 20 500 kHz–2 MHz ...... 30 30 2 MHz–30 MHz ...... 100 100 APPENDIX L TO PART 25—HIRF ENVI- 30 MHz–100 MHz ...... 10 10 RONMENTS AND EQUIPMENT HIRF 100 MHz–200 MHz ...... 30 10 TEST LEVELS 200 MHz–400 MHz ...... 10 10 400 MHz–1 GHz ...... 700 40 This appendix specifies the HIRF environ- 1 GHz–2 GHz ...... 1,300 160 2 GHz–4 GHz ...... 3,000 120 ments and equipment HIRF test levels for 4 GHz–6 GHz ...... 3,000 160 electrical and electronic systems under 6 GHz–8 GHz ...... 400 170 § 25.1317. The field strength values for the 8 GHz–12 GHz ...... 1,230 230 HIRF environments and equipment HIRF 12 GHz–18 GHz ...... 730 190 test levels are expressed in root-mean-square 18 GHz–40 GHz ...... 600 150 units measured during the peak of the modu- In this table, the higher field strength applies at the fre- lation cycle. quency band edges. (a) HIRF environment I is specified in the (c) Equipment HIRF Test Level 1. following table: (1) From 10 kilohertz (kHz) to 400 megahertz (MHz), use conducted susceptibility tests with continuous wave (CW) and 1 kHz

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square wave modulation with 90 percent (1) When any FRM is operational but the depth or greater. The conducted suscepti- fuel tank is not inert and the tank is flam- bility current must start at a minimum of mable; and 0.6 milliamperes (mA) at 10 kHz, increasing (2) When any FRM is inoperative and the 20 decibels (dB) per frequency decade to a tank is flammable. minimum of 30 mA at 500 kHz. (b) The Fleet Average Flammability Expo- (2) From 500 kHz to 40 MHz, the conducted sure, as defined in Appendix N of this part, of susceptibility current must be at least 30 each fuel tank may not exceed 3 percent of mA. the portion of the FEET occurring during ei- (3) From 40 MHz to 400 MHz, use conducted ther ground or takeoff/climb phases of flight susceptibility tests, starting at a minimum during warm days. The analysis must con- of 30 mA at 40 MHz, decreasing 20 dB per fre- sider the following conditions. quency decade to a minimum of 3 mA at 400 (1) The analysis must use the subset of MHz. (4) From 100 MHz to 400 MHz, use radiated those flights that begin with a sea level ° susceptibility tests at a minimum of 20 volts ground ambient temperature of 80 F (stand- ° per meter (V/m) peak with CW and 1 kHz ard day plus 21 F atmosphere) or above, square wave modulation with 90 percent from the flammability exposure analysis depth or greater. done for overall performance. (5) From 400 MHz to 8 gigahertz (GHz), use (2) For the ground and takeoff/climb phases radiated susceptibility tests at a minimum of flight, the average flammability exposure of 150 V/m peak with pulse modulation of 4 must be calculated by dividing the time dur- percent duty cycle with a 1 kHz pulse repeti- ing the specific flight phase the fuel tank is tion frequency. This signal must be switched flammable by the total time of the specific on and off at a rate of 1 Hz with a duty cycle flight phase. of 50 percent. (3) Compliance with this paragraph may be (d) Equipment HIRF Test Level 2. Equipment shown using only those flights for which the HIRF test level 2 is HIRF environment II in airplane is dispatched with the flammability table II of this appendix reduced by accept- reduction means operational. able aircraft transfer function and attenu- M25.2 Showing compliance. ation curves. Testing must cover the fre- (a) The applicant must provide data from quency band of 10 kHz to 8 GHz. analysis, ground testing, and flight testing, (e) Equipment HIRF Test Level 3. or any combination of these, that: (1) From 10 kHz to 400 MHz, use conducted (1) Validate the parameters used in the susceptibility tests, starting at a minimum analysis required by paragraph M25.1 of this of 0.15 mA at 10 kHz, increasing 20 dB per fre- appendix; quency decade to a minimum of 7.5 mA at 500 kHz. (2) Substantiate that the FRM is effective (2) From 500 kHz to 40 MHz, use conducted at limiting flammability exposure in all susceptibility tests at a minimum of 7.5 mA. compartments of each tank for which the (3) From 40 MHz to 400 MHz, use conducted FRM is used to show compliance with para- susceptibility tests, starting at a minimum graph M25.1 of this appendix; and of 7.5 mA at 40 MHz, decreasing 20 dB per fre- (3) Describe the circumstances under which quency decade to a minimum of 0.75 mA at the FRM would not be operated during each 400 MHz. phase of flight. (4) From 100 MHz to 8 GHz, use radiated (b) The applicant must validate that the susceptibility tests at a minimum of 5 V/m. FRM meets the requirements of paragraph M25.1 of this appendix with any airplane or [Doc. No. FAA–2006–23657, 72 FR 44026, Aug. 6, engine configuration affecting the perform- 2007] ance of the FRM for which approval is sought. APPENDIX M TO PART 25—FUEL TANK M25.3 Reliability indications and mainte- SYSTEM FLAMMABILITY REDUCTION nance access. MEANS (a) Reliability indications must be pro- M25.1 Fuel tank flammability exposure re- vided to identify failures of the FRM that quirements. would otherwise be latent and whose identi- (a) The Fleet Average Flammability Expo- fication is necessary to ensure the fuel tank sure of each fuel tank, as determined in ac- with an FRM meets the fleet average flam- cordance with Appendix N of this part, may mability exposure requirements listed in not exceed 3 percent of the Flammability Ex- paragraph M25.1 of this appendix, including posure Evaluation Time (FEET), as defined when the FRM is inoperative. in Appendix N of this part. As a portion of (b) Sufficient accessibility to FRM reli- this 3 percent, if flammability reduction ability indications must be provided for means (FRM) are used, each of the following maintenance personnel or the flightcrew. time periods may not exceed 1.8 percent of (c) The access doors and panels to the fuel the FEET: tanks with FRMs (including any tanks that

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communicate with a tank via a vent sys- substantiates that the tank is a conven- tem), and to any other confined spaces or en- tional unheated wing tank. closed areas that could contain hazardous at- (b) This appendix defines parameters af- mosphere under normal conditions or failure fecting fuel tank flammability that must be conditions, must be permanently stenciled, used in performing the analysis. These in- marked, or placarded to warn maintenance clude parameters that affect all airplanes personnel of the possible presence of a poten- within the fleet, such as a statistical dis- tially hazardous atmosphere. tribution of ambient temperature, fuel flash M25.4 Airworthiness limitations and proce- point, flight lengths, and airplane descent dures. rate. Demonstration of compliance also re- (a) If FRM is used to comply with para- quires application of factors specific to the graph M25.1 of this appendix, Airworthiness airplane model being evaluated. Factors that Limitations must be identified for all main- need to be included are maximum range, tenance or inspection tasks required to iden- cruise mach number, typical altitude where tify failures of components within the FRM the airplane begins initial cruise phase of that are needed to meet paragraph M25.1 of flight, fuel temperature during both ground this appendix. and flight times, and the performance of a (b) Maintenance procedures must be devel- flammability reduction means (FRM) if in- oped to identify any hazards to be considered stalled. during maintenance of the FRM. These pro- (c) The following definitions, input vari- cedures must be included in the instructions ables, and data tables must be used in the for continued airworthiness (ICA). program to determine fleet average flamma- M25.5 Reliability reporting. bility exposure for a specific airplane model. The effects of airplane component failures N25.2 Definitions. on FRM reliability must be assessed on an (a) Bulk Average Fuel Temperature means on-going basis. The applicant/holder must do the average fuel temperature within the fuel the following: tank or different sections of the tank if the (a) Demonstrate effective means to ensure tank is subdivided by baffles or compart- collection of FRM reliability data. The ments. means must provide data affecting FRM reli- (b) Flammability Exposure Evaluation Time ability, such as component failures. (FEET). The time from the start of preparing (b) Unless alternative reporting procedures the airplane for flight, through the flight are approved by the FAA Oversight Office, as and landing, until all payload is unloaded, defined in part 26 of this subchapter, provide and all passengers and crew have dis- a report to the FAA every six months for the embarked. In the Monte Carlo program, the first five years after service introduction. flight time is randomly selected from the After that period, continued reporting every Flight Length Distribution (Table 2), the pre-flight times are provided as a function of six months may be replaced with other reli- the flight time, and the post-flight time is a ability tracking methods found acceptable to constant 30 minutes. the FAA or eliminated if it is established (c) Flammable. With respect to a fluid or that the reliability of the FRM meets, and gas, flammable means susceptible to igniting will continue to meet, the exposure require- readily or to exploding (14 CFR Part 1, Defi- ments of paragraph M25.1 of this appendix. nitions). A non-flammable ullage is one (c) Develop service instructions or revise where the fuel-air vapor is too lean or too the applicable airplane manual, according to rich to burn or is inert as defined below. For a schedule approved by the FAA Oversight the purposes of this appendix, a fuel tank Office, as defined in part 26 of this sub- that is not inert is considered flammable chapter, to correct any failures of the FRM when the bulk average fuel temperature that occur in service that could increase any within the tank is within the flammable fuel tank’s Fleet Average Flammability Ex- range for the fuel type being used. For any posure to more than that required by para- fuel tank that is subdivided into sections by graph M25.1 of this appendix. baffles or compartments, the tank is consid- [Doc. No. FAA–2005–22997, 73 FR 42494, July ered flammable when the bulk average fuel 21, 2008] temperature within any section of the tank, that is not inert, is within the flammable APPENDIX N TO PART 25—FUEL TANK range for the fuel type being used. FLAMMABILITY EXPOSURE AND RELI- (d) Flash Point. The flash point of a flam- ABILITY ANALYSIS mable fluid means the lowest temperature at which the application of a flame to a heated N25.1 General. sample causes the vapor to ignite momen- (a) This appendix specifies the require- tarily, or ‘‘flash.’’ Table 1 of this appendix ments for conducting fuel tank fleet average provides the flash point for the standard fuel flammability exposure analyses required to to be used in the analysis. meet § 25.981(b) and Appendix M of this part. (e) Fleet average flammability exposure is the For fuel tanks installed in aluminum wings, percentage of the flammability exposure a qualitative assessment is sufficient if it evaluation time (FEET) each fuel tank

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ullage is flammable for a fleet of an airplane not allowed in the analysis. The analysis type operating over the range of flight must be done in accordance with the meth- lengths in a world-wide range of environ- ods and procedures set forth in the Fuel mental conditions and fuel properties as de- Tank Flammability Assessment Method fined in this appendix. User’s Manual, dated May 2008, document (f) Gaussian Distribution is another name number DOT/FAA/AR–05/8 (incorporated by for the normal distribution, a symmetrical reference, see § 25.5). The parameters speci- frequency distribution having a precise fied in sections N25.3(b) and (c) of this appen- mathematical formula relating the mean dix must be used in the fuel tank flamma- and standard deviation of the samples. bility exposure ‘‘Monte Carlo’’ analysis. Gaussian distributions yield bell-shaped fre- (b) The following parameters are defined in quency curves having a preponderance of val- the Monte Carlo analysis and provided in ues around the mean with progressively paragraph N25.4 of this appendix: fewer observations as the curve extends out- (1) Cruise Ambient Temperature, as de- ward. fined in this appendix. (g) Hazardous atmosphere. An atmosphere (2) Ground Ambient Temperature, as de- that may expose maintenance personnel, fined in this appendix. passengers or flight crew to the risk of death, incapacitation, impairment of ability (3) Fuel Flash Point, as defined in this ap- to self-rescue (that is, escape unaided from a pendix. confined space), injury, or acute illness. (4) Flight Length Distribution, as defined (h) Inert. For the purpose of this appendix, in Table 2 of this appendix. the tank is considered inert when the bulk (5) Airplane Climb and Descent Profiles, as average oxygen concentration within each defined in the Fuel Tank Flammability As- compartment of the tank is 12 percent or less sessment Method User’s Manual, dated May from sea level up to 10,000 feet altitude, then 2008, document number DOT/FAA/AR–05/8 linearly increasing from 12 percent at 10,000 (incorporated by reference in § 25.5). feet to 14.5 percent at 40,000 feet altitude, (c) Parameters that are specific to the par- and extrapolated linearly above that alti- ticular airplane model under evaluation that tude. must be provided as inputs to the Monte (i) Inerting. A process where a noncombus- Carlo analysis are: tible gas is introduced into the ullage of a (1) Airplane cruise altitude. fuel tank so that the ullage becomes non- (2) Fuel tank quantities. If fuel quantity flammable. affects fuel tank flammability, inputs to the (j) Monte Carlo Analysis. The analytical Monte Carlo analysis must be provided that method that is specified in this appendix as represent the actual fuel quantity within the the compliance means for assessing the fleet fuel tank or compartment of the fuel tank average flammability exposure time for a throughout each of the flights being evalu- fuel tank. ated. Input values for this data must be ob- (k) Oxygen evolution occurs when oxygen tained from ground and flight test data or dissolved in the fuel is released into the the approved FAA fuel management proce- ullage as the pressure and temperature in dures. the fuel tank are reduced. (3) Airplane cruise mach number. (l) Standard deviation is a statistical measure of the dispersion or variation in a dis- (4) Airplane maximum range. tribution, equal to the square root of the (5) Fuel tank thermal characteristics. If arithmetic mean of the squares of the devi- fuel temperature affects fuel tank flamma- ations from the arithmetic means. bility, inputs to the Monte Carlo analysis (m) Transport Effects. For purposes of this must be provided that represent the actual appendix, transport effects are the change in bulk average fuel temperature within the fuel vapor concentration in a fuel tank fuel tank at each point in time throughout caused by low fuel conditions and fuel con- each of the flights being evaluated. For fuel densation and vaporization. tanks that are subdivided by baffles or com- (n) Ullage. The volume within the fuel tank partments, bulk average fuel temperature not occupied by liquid fuel. inputs must be provided for each section of N25.3 Fuel tank flammability exposure anal- the tank. Input values for these data must be ysis. obtained from ground and flight test data or (a) A flammability exposure analysis must a thermal model of the tank that has been be conducted for the fuel tank under evalua- validated by ground and flight test data. tion to determine fleet average flammability (6) Maximum airplane operating tempera- exposure for the airplane and fuel types ture limit, as defined by any limitations in under evaluation. For fuel tanks that are the airplane flight manual. subdivided by baffles or compartments, an (7) Airplane Utilization. The applicant analysis must be performed either for each must provide data supporting the number of section of the tank, or for the section of the flights per day and the number of hours per tank having the highest flammability expo- flight for the specific airplane model under sure. Consideration of transport effects is evaluation. If there is no existing airplane

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fleet data to support the airplane being eval- fect flammability exposure, substantiating uated, the applicant must provide substan- data, and any airworthiness limitations and tiation that the number of flights per day other conditions assumed in the analysis. and the number of hours per flight for that N25.4 Variables and data tables. airplane model is consistent with the exist- The following data must be used when con- ing fleet data they propose to use. ducting a flammability exposure analysis to (d) Fuel Tank FRM Model. If FRM is used, determine the fleet average flammability ex- an FAA approved Monte Carlo program must posure. Variables used to calculate fleet be used to show compliance with the flam- flammability exposure must include atmos- mability requirements of § 25.981 and Appen- pheric ambient temperatures, flight length, dix M of this part. The program must deter- flammability exposure evaluation time, fuel mine the time periods during each flight flash point, thermal characteristics of the phase when the fuel tank or compartment fuel tank, overnight temperature drop, and with the FRM would be flammable. The fol- oxygen evolution from the fuel into the lowing factors must be considered in estab- ullage. lishing these time periods: (a) Atmospheric Ambient Temperatures (1) Any time periods throughout the flam- and Fuel Properties. mability exposure evaluation time and under (1) In order to predict flammability expo- the full range of expected operating condi- sure during a given flight, the variation of tions, when the FRM is operating properly ground ambient temperatures, cruise ambi- but fails to maintain a non-flammable fuel ent temperatures, and a method to compute tank because of the effects of the fuel tank the transition from ground to cruise and vent system or other causes, back again must be used. The variation of (2) If dispatch with the system inoperative the ground and cruise ambient temperatures under the Master Minimum Equipment List and the flash point of the fuel is defined by (MMEL) is requested, the time period as- a Gaussian curve, given by the 50 percent sumed in the reliability analysis (60 flight value and a ±1-standard deviation value. hours must be used for a 10-day MMEL dis- (2) Ambient Temperature: Under the pro- patch limit unless an alternative period has gram, the ground and cruise ambient tem- been approved by the Administrator), peratures are linked by a set of assumptions (3) Frequency and duration of time periods on the atmosphere. The temperature varies of FRM inoperability, substantiated by test with altitude following the International or analysis acceptable to the FAA, caused by Standard Atmosphere (ISA) rate of change latent or known failures, including airplane from the ground ambient temperature until system shut-downs and failures that could the cruise temperature for the flight is cause the FRM to shut down or become inop- reached. Above this altitude, the ambient erative. temperature is fixed at the cruise ambient (4) Effects of failures of the FRM that temperature. This results in a variation in could increase the flammability exposure of the upper atmospheric temperature. For cold the fuel tank. days, an inversion is applied up to 10,000 feet, (5) If an FRM is used that is affected by ox- and then the ISA rate of change is used. ygen concentrations in the fuel tank, the (3) Fuel properties: time periods when oxygen evolution from the (i) For Jet A fuel, the variation of flash fuel results in the fuel tank or compartment point of the fuel is defined by a Gaussian exceeding the inert level. The applicant curve, given by the 50 percent value and a ±1- must include any times when oxygen evo- standard deviation, as shown in Table 1 of lution from the fuel in the tank or compart- this appendix. ment under evaluation would result in a (ii) The flammability envelope of the fuel flammable fuel tank. The oxygen evolution that must be used for the flammability expo- rate that must be used is defined in the Fuel sure analysis is a function of the flash point Tank Flammability Assessment Method of the fuel selected by the Monte Carlo for a User’s Manual, dated May 2008, document given flight. The flammability envelope for number DOT/FAA/AR–05/8 (incorporated by the fuel is defined by the upper flammability reference in § 25.5). limit (UFL) and lower flammability limit (6) If an inerting system FRM is used, the (LFL) as follows: effects of any air that may enter the fuel (A) LFL at sea level = flash point tempera- tank following the last flight of the day due ture of the fuel at sea level minus 10 °F. LFL to changes in ambient temperature, as de- decreases from sea level value with increas- fined in Table 4, during a 12-hour overnight ing altitude at a rate of 1 °F per 808 feet. period. (B) UFL at sea level = flash point tempera- (e) The applicant must submit to the FAA ture of the fuel at sea level plus 63.5 °F. UFL Oversight Office for approval the fuel tank decreases from the sea level value with in- flammability analysis, including the air- creasing altitude at a rate of 1 °F per 512 plane-specific parameters identified under feet. paragraph N25.3(c) of this appendix and any (4) For each flight analyzed, a separate deviations from the parameters identified in random number must be generated for each paragraph N25.3(b) of this appendix that af- of the three parameters (ground ambient

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temperature, cruise ambient temperature, and fuel flash point) using the Gaussian distribution defined in Table 1 of this appendix.

TABLE 1.—GAUSSIAN DISTRIBUTION FOR GROUND AMBIENT TEMPERATURE, CRUISE AMBIENT TEMPERATURE, AND FUEL FLASH POINT

Temperature in deg F Parameter Ground ambient Cruise ambient Fuel flash point temperature temperature (FP)

Mean Temp ...... 59.95 ¥70 120 Neg 1 std dev ...... 20.14 8 8 Pos 1 std dev ...... 17.28 8 8

(b) The Flight Length Distribution defined in Table 2 must be used in the Monte Carlo analysis.

TABLE 2.—FLIGHT LENGTH DISTRIBUTION

Flight length (NM) Airplane maximum range—nautical miles (NM) From To 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

Distribution of flight lengths (percentage of total)

0 200 11.7 7.5 6.2 5.5 4.7 4.0 3.4 3.0 2.6 2.3 200 400 27.3 19.9 17.0 15.2 13.2 11.4 9.7 8.5 7.5 6.7 400 600 46.3 40.0 35.7 32.6 28.5 24.9 21.2 18.7 16.4 14.8 600 800 10.3 11.6 11.0 10.2 9.1 8.0 6.9 6.1 5.4 4.8 800 1000 4.4 8.5 8.6 8.2 7.4 6.6 5.7 5.0 4.5 4.0 1000 1200 0.0 4.8 5.3 5.3 4.8 4.3 3.8 3.3 3.0 2.7 1200 1400 0.0 3.6 4.4 4.5 4.2 3.8 3.3 3.0 2.7 2.4 1400 1600 0.0 2.2 3.3 3.5 3.3 3.1 2.7 2.4 2.2 2.0 1600 1800 0.0 1.2 2.3 2.6 2.5 2.4 2.1 1.9 1.7 1.6 1800 2000 0.0 0.7 2.2 2.6 2.6 2.5 2.2 2.0 1.8 1.7 2000 2200 0.0 0.0 1.6 2.1 2.2 2.1 1.9 1.7 1.6 1.4 2200 2400 0.0 0.0 1.1 1.6 1.7 1.7 1.6 1.4 1.3 1.2 2400 2600 0.0 0.0 0.7 1.2 1.4 1.4 1.3 1.2 1.1 1.0 2600 2800 0.0 0.0 0.4 0.9 1.0 1.1 1.0 0.9 0.9 0.8 2800 3000 0.0 0.0 0.2 0.6 0.7 0.8 0.7 0.7 0.6 0.6 3000 3200 0.0 0.0 0.0 0.6 0.8 0.8 0.8 0.8 0.7 0.7 3200 3400 0.0 0.0 0.0 0.7 1.1 1.2 1.2 1.1 1.1 1.0 3400 3600 0.0 0.0 0.0 0.7 1.3 1.6 1.6 1.5 1.5 1.4 3600 3800 0.0 0.0 0.0 0.9 2.2 2.7 2.8 2.7 2.6 2.5 3800 4000 0.0 0.0 0.0 0.5 2.0 2.6 2.8 2.8 2.7 2.6 4000 4200 0.0 0.0 0.0 0.0 2.1 3.0 3.2 3.3 3.2 3.1 4200 4400 0.0 0.0 0.0 0.0 1.4 2.2 2.5 2.6 2.6 2.5 4400 4600 0.0 0.0 0.0 0.0 1.0 2.0 2.3 2.5 2.5 2.4 4600 4800 0.0 0.0 0.0 0.0 0.6 1.5 1.8 2.0 2.0 2.0 4800 5000 0.0 0.0 0.0 0.0 0.2 1.0 1.4 1.5 1.6 1.5 5000 5200 0.0 0.0 0.0 0.0 0.0 0.8 1.1 1.3 1.3 1.3 5200 5400 0.0 0.0 0.0 0.0 0.0 0.8 1.2 1.5 1.6 1.6 5400 5600 0.0 0.0 0.0 0.0 0.0 0.9 1.7 2.1 2.2 2.3 5600 5800 0.0 0.0 0.0 0.0 0.0 0.6 1.6 2.2 2.4 2.5 5800 6000 0.0 0.0 0.0 0.0 0.0 0.2 1.8 2.4 2.8 2.9 6000 6200 0.0 0.0 0.0 0.0 0.0 0.0 1.7 2.6 3.1 3.3 6200 6400 0.0 0.0 0.0 0.0 0.0 0.0 1.4 2.4 2.9 3.1 6400 6600 0.0 0.0 0.0 0.0 0.0 0.0 0.9 1.8 2.2 2.5 6600 6800 0.0 0.0 0.0 0.0 0.0 0.0 0.5 1.2 1.6 1.9 6800 7000 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.8 1.1 1.3 7000 7200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.4 0.7 0.8 7200 7400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.3 0.5 0.7 7400 7600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.5 0.6 7600 7800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 0.5 0.7 7800 8000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 0.6 0.8 8000 8200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 0.8 8200 8400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 1.0 8400 8600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.6 1.3 8600 8800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.4 1.1 8800 9000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.8 9000 9200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5

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TABLE 2.—FLIGHT LENGTH DISTRIBUTION—Continued

Flight length (NM) Airplane maximum range—nautical miles (NM) From To 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

9200 9400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 9400 9600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 9600 9800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 9800 10000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1

(c) Overnight Temperature Drop. For air- TABLE 5.—FLAMMABILITY EXPOSURE LIMIT planes on which FRM is installed, the overnight temperature drop for this appendix is Maximum Maximum defined using: acceptable Monte acceptable Monte Carlo average fuel Carlo average fuel (1) A temperature at the beginning of the Minimum number of flights in Monte tank flammability tank flammability overnight period that equals the landing exposure exposure Carlo analysis (percent) to meet (percent) to meet temperature of the previous flight that is a 3 percent 7 percent part 26 random value based on a Gaussian distribu- requirements requirements tion; and (2) An overnight temperature drop that is 10,000 ...... 2.91 6.79 a random value based on a Gaussian distribu- 100,000 ...... 2.98 6.96 tion. 1,000,000 ...... 3.00 7.00 (3) For any flight that will end with an overnight ground period (one flight per day [Doc. No. FAA–2005–22997, 73 FR 42495, July out of an average number of flights per day, 21, 2008] depending on utilization of the particular airplane model being evaluated), the landing outside air temperature (OAT) is to be cho- PART 26—CONTINUED AIRWORTHI- sen as a random value from the following NESS AND SAFETY IMPROVE- Gaussian curve: MENTS FOR TRANSPORT CAT- EGORY AIRPLANES TABLE 3.—LANDING OUTSIDE AIR TEMPERATURE

Landing outside Subpart A—General Parameter air temperature °F Sec. Mean Temperature ...... 58.68 26.1 Purpose and scope. negative 1 std dev ...... 20.55 26.3 Definitions. positive 1 std dev ...... 13.21 26.5 Applicability table. (4) The outside ambient air temperature (OAT) overnight temperature drop is to be Subpart B—Enhanced Airworthiness chosen as a random value from the following Program for Airplane Systems Gaussian curve: 26.11 Electrical wiring interconnection systems (EWIS) maintenance program. TABLE 4.—OUTSIDE AIR TEMPERATURE (OAT) DROP Subpart C—Aging Airplane Safety— OAT drop Widespread Fatigue Damage Parameter temperature °F 26.21 Limit of validity. Mean Temp ...... 12.0 26.23 Extended limit of validity. 1 std dev ...... 6.0 Subpart D—Fuel Tank Flammability (d) Number of Simulated Flights Required in Analysis. In order for the Monte Carlo 26.31 Definitions. analysis to be valid for showing compliance 26.33 Holders of type certificates: Fuel tank with the fleet average and warm day flam- flammability. mability exposure requirements, the appli- 26.35 Changes to type certificates affecting cant must run the analysis for a minimum fuel tank flammability. number of flights to ensure that the fleet av- 26.37 Pending type certification projects: erage and warm day flammability exposure Fuel tank flammability. for the fuel tank under evaluation meets the applicable flammability limits defined in Table 5 of this appendix.

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