Federal Aviation Administration, DOT § 25.149

Federal Aviation Administration, DOT Pt. 25

applicable wear tolerances, and work rec- be included in the principal manual. This ommended at these periods. However, the ap- section must contain a legible statement in plicant may refer to an accessory, instru- a prominent location that reads ‘‘The Air- ment, or equipment manufacturer as the worthiness Limitations section is FAA ap- source of this information if the applicant proved and specifies maintenance required shows that the item has an exceptionally under §§ 43.16 and 91.403 of Title 14 of the Code high degree of complexity requiring special- of Federal Regulations unless an alternative ized maintenance techniques, test equip- program has been FAA approved.’’ ment, or expertise. The recommended overhaul periods and necessary cross reference to the Airworthiness Limitations section of the PART 25—AIRWORTHINESS STAND- manual must also be included. In addition, ARDS: TRANSPORT CATEGORY the applicant must include an inspection AIRPLANES program that includes the frequency and extent of the inspections necessary to provide SPECIAL FEDERAL AVIATION REGULATION NO. for the continued airworthiness of the air- 13 plane. SPECIAL FEDERAL AVIATION REGULATION NO. (2) Troubleshooting information describing 109 probable malfunctions, how to recognize those malfunctions, and the remedial action for those malfunctions. Subpart A—General (3) Information describing the order and Sec. method of removing and replacing products 25.1 Applicability. and parts with any necessary precautions to 25.2 Special retroactive requirements. be taken. 25.3 Special provisions for ETOPS type de- (4) Other general procedural instructions sign approvals. including procedures for system testing dur- 25.5 Incorporations by reference. ing ground running, symmetry checks, weighing and determining the center of grav- Subpart B—Flight ity, lifting and shoring, and storage limitations. GENERAL (c) Diagrams of structural access plates 25.21 Proof of compliance. and information needed to gain access for in- 25.23 Load distribution limits. spections when access plates are not pro- 25.25 Weight limits. vided. 25.27 Center of gravity limits. (d) Details for the application of special in- 25.29 Empty weight and corresponding cen- spection techniques including radiographic ter of gravity. and ultrasonic testing where such processes 25.31 Removable ballast. are specified by the applicant. 25.33 Propeller speed and pitch limits. (e) Information needed to apply protective treatments to the structure after inspection. PERFORMANCE (f) All data relative to structural fasteners such as identification, discard recommenda- 25.101 General. tions, and torque values. 25.103 Stall speed. (g) A list of special tools needed. 25.105 Takeoff. (h) In addition, for level 4 airplanes, the 25.107 Takeoff speeds. following information must be furnished— 25.109 Accelerate-stop distance. (1) Electrical loads applicable to the var- 25.111 Takeoff path. ious systems; 25.113 Takeoff distance and takeoff run. (2) Methods of balancing control surfaces; 25.115 Takeoff flight path. (3) Identification of primary and secondary 25.117 Climb: general. structures; and 25.119 Landing climb: All-engines-operating. (4) Special repair methods applicable to 25.121 Climb: One-engine-inoperative. the airplane. 25.123 En route flight paths. 25.125 Landing. A23.4 Airworthiness limitations section. CONTROLLABILITY AND MANEUVERABILITY The Instructions for Continued Airworthi- ness must contain a section titled Airworthi- 25.143 General. ness Limitations that is segregated and 25.145 Longitudinal control. clearly distinguishable from the rest of the 25.147 Directional and lateral control. document. This section must set forth each 25.149 Minimum control speed. mandatory replacement time, structural in- TRIM spection interval, and related structural inspection procedure required for type certifi- 25.161 Trim. cation. If the Instructions for Continued Air- STABILITY worthiness consist of multiple documents, the section required by this paragraph must 25.171 General.

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25.173 Static longitudinal stability. 25.405 Secondary control system. 25.175 Demonstration of static longitudinal 25.407 Trim tab effects. stability. 25.409 Tabs. 25.177 Static lateral-directional stability. 25.415 Ground gust conditions. 25.181 Dynamic stability. 25.427 Unsymmetrical loads. 25.445 Auxiliary aerodynamic surfaces. STALLS 25.457 Wing flaps. 25.201 Stall demonstration. 25.459 Special devices. 25.203 Stall characteristics. GROUND LOADS 25.207 Stall warning. 25.471 General. GROUND AND WATER HANDLING 25.473 Landing load conditions and assump- CHARACTERISTICS tions. 25.231 Longitudinal stability and control. 25.477 Landing gear arrangement. 25.233 Directional stability and control. 25.479 Level landing conditions. 25.235 Taxiing condition. 25.481 Tail-down landing conditions. 25.237 Wind velocities. 25.483 One-gear landing conditions. 25.239 Spray characteristics, control, and 25.485 Side load conditions. stability on water. 25.487 Rebound landing condition. 25.489 Ground handling conditions. MISCELLANEOUS FLIGHT REQUIREMENTS 25.491 Taxi, takeoff and landing roll. 25.251 Vibration and buffeting. 25.493 Braked roll conditions. 25.253 High-speed characteristics. 25.495 Turning. 25.255 Out-of-trim characteristics. 25.497 Tail-wheel yawing. 25.499 Nose-wheel yaw and steering. Subpart C—Structure 25.503 Pivoting. 25.507 Reversed braking. GENERAL 25.509 Towing loads. 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

FLIGHT LOADS 25.521 General. 25.523 Design weights and center of gravity 25.321 General. positions. FLIGHT MANEUVER AND GUST CONDITIONS 25.525 Application of loads. 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. 25.349 Rolling conditions. 25.351 Yaw maneuver conditions. EMERGENCY LANDING CONDITIONS

SUPPLEMENTARY CONDITIONS 25.561 General. 25.562 Emergency landing dynamic condi- 25.361 Engine and auxiliary power unit tions. torque. 25.563 Structural ditching provisions. 25.362 Engine failure loads. 25.363 Side load on engine and auxiliary FATIGUE EVALUATION power unit mounts. 25.365 Pressurized compartment loads. 25.571 Damage—tolerance and fatigue eval- 25.367 Unsymmetrical loads due to engine uation of structure. failure. LIGHTNING PROTECTION 25.371 Gyroscopic loads. 25.373 Speed control devices. 25.581 Lightning protection.

CONTROL SURFACE AND SYSTEM LOADS Subpart D—Design and Construction 25.391 Control surface loads: General. GENERAL 25.393 Loads parallel to hinge line. 25.395 Control system. 25.601 General. 25.397 Control system loads. 25.603 Materials. 25.399 Dual control system. 25.605 Fabrication methods.

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25.607 Fasteners. 25.793 Floor surfaces. 25.609 Protection of structure. 25.795 Security considerations. 25.611 Accessibility provisions. 25.613 Material strength properties and ma- EMERGENCY PROVISIONS terial design values. 25.801 Ditching. 25.619 Special factors. 25.803 Emergency evacuation. 25.621 Casting factors. 25.807 Emergency exits. 25.623 Bearing factors. 25.809 Emergency exit arrangement. 25.625 Fitting factors. 25.810 Emergency egress assist means and 25.629 Aeroelastic stability requirements. escape routes. 25.631 Bird strike damage. 25.811 Emergency exit marking. CONTROL SURFACES 25.812 Emergency lighting. 25.813 Emergency exit access. 25.651 Proof of strength. 25.815 Width of aisle. 25.655 Installation. 25.817 Maximum number of seats abreast. 25.657 Hinges. 25.819 Lower deck service compartments (including galleys). CONTROL SYSTEMS 25.820 Lavatory doors. 25.671 General. 25.672 Stability augmentation and auto- VENTILATION AND HEATING matic and power-operated systems. 25.831 Ventilation. 25.675 Stops. 25.832 Cabin ozone concentration. 25.677 Trim systems. 25.833 Combustion heating systems. 25.679 Control system gust locks. 25.681 Limit load static tests. PRESSURIZATION 25.683 Operation tests. 25.685 Control system details. 25.841 Pressurized cabins. 25.689 Cable systems. 25.843 Tests for pressurized cabins. 25.693 Joints. FIRE PROTECTION 25.697 Lift and drag devices, controls. 25.699 Lift and drag device indicator. 25.851 Fire extinguishers. 25.701 Flap and slat interconnection. 25.853 Compartment interiors. 25.703 Takeoff warning system. 25.854 Lavatory fire protection. 25.855 Cargo or baggage compartments. LANDING GEAR 25.856 Thermal/Acoustic insulation mate- 25.721 General. rials. 25.723 Shock absorption tests. 25.857 Cargo compartment classification. 25.725–25.727 [Reserved] 25.858 Cargo or baggage compartment 25.729 Retracting mechanism. smoke or fire detection systems. 25.731 Wheels. 25.859 Combustion heater fire protection. 25.733 Tires. 25.863 Flammable fluid fire protection. 25.735 Brakes and braking systems. 25.865 Fire protection of flight controls, en- 25.737 Skis. gine mounts, and other flight structure. 25.867 Fire protection: other components. FLOATS AND HULLS 25.869 Fire protection: systems.

25.751 Main float buoyancy. MISCELLANEOUS 25.753 Main float design. 25.755 Hulls. 25.871 Leveling means. 25.875 Reinforcement near propellers. PERSONNEL AND CARGO ACCOMMODATIONS 25.899 Electrical bonding and protection 25.771 Pilot compartment. against static electricity. 25.772 Pilot compartment doors. 25.773 Pilot compartment view. Subpart E—Powerplant 25.775 Windshields and windows. GENERAL 25.777 Cockpit controls. 25.779 Motion and effect of cockpit controls. 25.901 Installation. 25.781 Cockpit control knob shape. 25.903 Engines. 25.783 Fuselage doors. 25.904 Automatic takeoff thrust control sys- 25.785 Seats, berths, safety belts, and har- tem (ATTCS). nesses. 25.905 Propellers. 25.787 Stowage compartments. 25.907 Propeller vibration and fatigue. 25.789 Retention of items of mass in pas- 25.925 Propeller clearance. senger and crew compartments and gal- 25.929 Propeller deicing. leys. 25.933 Reversing systems. 25.791 Passenger information signs and plac- 25.934 Turbojet engine thrust reverser sys- ards. tem tests.

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25.937 Turbopropeller-drag limiting sys- EXHAUST SYSTEM tems. 25.1121 General. 25.939 Turbine engine operating characteris- 25.1123 Exhaust piping. tics. 25.1125 Exhaust heat exchangers. 25.941 Inlet, engine, and exhaust compat- 25.1127 Exhaust driven turbo-superchargers. ibility. 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. 25.1145 Ignition switches. 25.951 General. 25.1147 Mixture controls. 25.952 Fuel system analysis and test. 25.1149 Propeller speed and pitch controls. 25.953 Fuel system independence. 25.1153 Propeller feathering controls. 25.954 Fuel system lightning protection. 25.1155 Reverse thrust and propeller pitch 25.955 Fuel flow. settings below the flight regime. 25.957 Flow between interconnected tanks. 25.1157 Carburetor air temperature controls. 25.959 Unusable fuel supply. 25.1159 Supercharger controls. 25.961 Fuel system hot weather operation. 25.1161 Fuel jettisoning system controls. 25.963 Fuel tanks: general. 25.1163 Powerplant accessories. 25.965 Fuel tank tests. 25.1165 Engine ignition systems. 25.967 Fuel tank installations. 25.1167 Accessory gearboxes. 25.969 Fuel tank expansion space. 25.971 Fuel tank sump. POWERPLANT FIRE PROTECTION 25.973 Fuel tank filler connection. 25.1181 Designated fire zones; regions in- 25.975 Fuel tank vents and carburetor vapor cluded. vents. 25.1182 Nacelle areas behind firewalls, and 25.977 Fuel tank outlet. engine pod attaching structures con- 25.979 Pressure fueling system. taining flammable fluid lines. 25.981 Fuel tank explosion prevention. 25.1183 Flammable fluid-carrying components. FUEL SYSTEM COMPONENTS 25.1185 Flammable fluids. 25.991 Fuel pumps. 25.1187 Drainage and ventilation of fire 25.993 Fuel system lines and fittings. zones. 25.994 Fuel system components. 25.1189 Shutoff means. 25.995 Fuel valves. 25.1191 Firewalls. 25.997 Fuel strainer or filter. 25.1192 Engine accessory section diaphragm. 25.999 Fuel system drains. 25.1193 Cowling and nacelle skin. 25.1001 Fuel jettisoning system. 25.1195 Fire extinguishing systems. 25.1197 Fire extinguishing agents. OIL SYSTEM 25.1199 Extinguishing agent containers. 25.1201 Fire extinguishing system materials. 25.1011 General. 25.1203 Fire detector system. 25.1013 Oil tanks. 25.1207 Compliance. 25.1015 Oil tank tests. 25.1017 Oil lines and fittings. Subpart F—Equipment 25.1019 Oil strainer or filter. 25.1021 Oil system drains. GENERAL 25.1023 Oil radiators. 25.1301 Function and installation. 25.1025 Oil valves. 25.1302 Installed systems and equipment for 25.1027 Propeller feathering system. use by the flightcrew. 25.1303 Flight and navigation instruments. COOLING 25.1305 Powerplant instruments. 25.1041 General. 25.1307 Miscellaneous equipment. 25.1043 Cooling tests. 25.1309 Equipment, systems, and installa- 25.1045 Cooling test procedures. tions. 25.1310 Power source capacity and distribu- INDUCTION SYSTEM tion. 25.1091 Air induction. 25.1316 Electrical and electronic system 25.1093 Induction system icing protection. lightning protection. 25.1101 Carburetor air preheater design. 25.1317 High-intensity Radiated Fields 25.1103 Induction system ducts and air duct (HIRF) Protection. systems. INSTRUMENTS: INSTALLATION 25.1105 Induction system screens. 25.1107 Inter-coolers and after-coolers. 25.1321 Arrangement and visibility.

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25.1322 Flightcrew alerting. 25.1447 Equipment standards for oxygen dis- 25.1323 Airspeed indicating system. pensing units. 25.1324 Angle of attack system. 25.1449 Means for determining use of oxy- 25.1325 Static pressure systems. gen. 25.1326 Pitot heat indication systems. 25.1450 Chemical oxygen generators. 25.1327 Magnetic direction indicator. 25.1453 Protection of oxygen equipment 25.1329 Flight guidance system. from rupture. 25.1331 Instruments using a power supply. 25.1455 Draining of fluids subject to freez- 25.1333 Instrument systems. ing. 25.1337 Powerplant instruments. 25.1457 Cockpit voice recorders. 25.1459 Flight data recorders. ELECTRICAL SYSTEMS AND EQUIPMENT 25.1461 Equipment containing high energy 25.1351 General. rotors. 25.1353 Electrical equipment and installations. Subpart G—Operating Limitations and 25.1355 Distribution system. Information 25.1357 Circuit protective devices. 25.1360 Precautions against injury. 25.1501 General. 25.1362 Electrical supplies for emergency OPERATING LIMITATIONS conditions. 25.1363 Electrical system tests. 25.1503 Airspeed limitations: general. 25.1365 Electrical appliances, motors, and 25.1505 Maximum operating limit speed. transformers. 25.1507 Maneuvering speed. 25.1511 Flap extended speed. LIGHTS 25.1513 Minimum control speed. 25.1515 Landing gear speeds. 25.1381 Instrument lights. 25.1516 Other speed limitations. 25.1383 Landing lights. 25.1517 Rough air speed, V 25.1385 Position light system installation. RA. 25.1519 Weight, center of gravity, and 25.1387 Position light system dihedral an- weight distribution. gles. 25.1521 Powerplant limitations. 25.1389 Position light distribution and in- 25.1522 Auxiliary power unit limitations. tensities. 25.1523 Minimum flight crew. 25.1391 Minimum intensities in the hori- 25.1525 Kinds of operation. zontal plane of forward and rear position 25.1527 Ambient air temperature and oper- lights. ating altitude. 25.1393 Minimum intensities in any vertical 25.1529 Instructions for Continued Air- plane of forward and rear position lights. worthiness. 25.1395 Maximum intensities in overlapping 25.1531 Maneuvering flight load factors. beams of forward and rear position 25.1533 Additional operating limitations. lights. 25.1535 ETOPS approval. 25.1397 Color specifications. 25.1399 Riding light. MARKINGS AND PLACARDS 25.1401 Anticollision light system. 25.1541 General. 25.1403 Wing icing detection lights. 25.1543 Instrument markings: general. SAFETY EQUIPMENT 25.1545 Airspeed limitation information. 25.1547 Magnetic direction indicator. 25.1411 General. 25.1549 Powerplant and auxiliary power unit 25.1415 Ditching equipment. instruments. 25.1419 Ice protection. 25.1551 Oil quantity indication. 25.1420 Supercooled large drop icing condi- 25.1553 Fuel quantity indicator. tions. 25.1555 Control markings. 25.1421 Megaphones. 25.1557 Miscellaneous markings and plac- 25.1423 Public address system. ards. 25.1561 Safety equipment. MISCELLANEOUS EQUIPMENT 25.1563 Airspeed placard. 25.1431 Electronic equipment. 25.1433 Vacuum systems. AIRPLANE FLIGHT MANUAL 25.1435 Hydraulic systems. 25.1581 General. 25.1438 Pressurization and pneumatic sys- 25.1583 Operating limitations. tems. 25.1585 Operating procedures. 25.1439 Protective breathing equipment. 25.1587 Performance information. 25.1441 Oxygen equipment and supply. 25.1443 Minimum mass flow of supplemental Subpart H—Electrical Wiring oxygen. Interconnection Systems (EWIS) 25.1445 Equipment standards for the oxygen distributing system. 25.1701 Definition.

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25.1703 Function and installation: EWIS. notwithstanding, 1 this regulation shall pro- 25.1705 Systems and functions: EWIS. vide the basis for approval by the Adminis- 25.1707 System separation: EWIS. trator of modifications of individual Douglas 25.1709 System safety: EWIS. DC–3 and Lockheed L–18 airplanes subse- 25.1711 Component identification: EWIS. quent to the effective date of this regulation. 25.1713 Fire protection: EWIS. 2. General modifications. Except as modified 25.1715 Electrical bonding and protection in sections 3 and 4 of this regulation, an ap- against static electricity: EWIS. plicant for approval of modifications to a 25.1717 Circuit protective devices: EWIS. DC–3 or L–18 airplane which result in 25.1719 Accessibility provisions: EWIS. changes in design or in changes to approved 25.1721 Protection of EWIS. limitations shall show that the modifica- 25.1723 Flammable fluid fire protection: tions were accomplished in accordance with EWIS. the rules of either Part 4a or Part 4b in ef- 25.1725 Powerplants: EWIS. fect on September 1, 1953, which are applica- 25.1727 Flammable fluid shutoff means: ble to the modification being made: Provided, EWIS. That an applicant may elect to accomplish a 25.1729 Instructions for Continued Air- modification in accordance with the rules of worthiness: EWIS. Part 4b in effect on the date of application 25.1731 Powerplant and APU fire detector for the modification in lieu of Part 4a or system: EWIS. Part 4b as in effect on September 1, 1953: And 25.1733 Fire detector systems, general: provided further, That each specific modifica- EWIS. tion must be accomplished in accordance with all of the provisions contained in the Subpart I—Special Federal Aviation elected rules relating to the particular modi- Regulations fication. 3. Specific conditions for approval. An appli- 25.1801 SFAR No. 111—Lavatory Oxygen cant for any approval of the following spe- Systems. cific changes shall comply with section 2 of APPENDIX A TO PART 25 this regulation as modified by the applicable APPENDIX B TO PART 25 provisions of this section. APPENDIX C TO PART 25 (a) Increase in take-off power limitation— APPENDIX D TO PART 25 1,200 to 1,350 horsepower. The engine take-off APPENDIX E TO PART 25 power limitation for the airplane may be in- APPENDIX F TO PART 25 creased to more than 1,200 horsepower but APPENDIX G TO PART 25 [RESERVED] not to more than 1,350 horsepower per engine if the increase in power does not adversely APPENDIX H TO PART 25—INSTRUCTIONS FOR affect the flight characteristics of the air- CONTINUED AIRWORTHINESS plane. APPENDIX I TO PART 25—INSTALLATION OF AN AUTOMATIC TAKEOFF THRUST CONTROL (b) Increase in take-off power limitation to SYSTEM (ATTCS) more than 1,350 horsepower. The engine take- APPENDIX J TO PART 25—EMERGENCY EVACU- off power limitation for the airplane may be ATION increased to more than 1,350 horsepower per engine if compliance is shown with the flight APPENDIX K TO PART 25—EXTENDED OPER- characteristics and ground handling require- ATIONS (ETOPS) ments of Part 4b. APPENDIX L TO PART 25—HIRF ENVIRON- MENTS AND EQUIPMENT HIRF TEST LEV- (c) Installation of engines of not more than ELS 1,830 cubic inches displacement and not having a certificated take-off rating of more than 1,350 APPENDIX M TO PART 25—FUEL TANK SYSTEM horsepower. Engines of not more than 1,830 FLAMMABILITY REDUCTION MEANS cubic inches displacement and not having a APPENDIX N TO PART 25—FUEL TANK FLAM- certificated take-off rating of more than MABILITY EXPOSURE AND RELIABILITY 1,350 horsepower which necessitate a major ANALYSIS modification of redesign of the engine instal- APPENDIX O TO PART 25—SUPERCOOLED LARGE lation may be installed, if the engine fire DROP ICING CONDITIONS prevention and fire protection are equivalent AUTHORITY: 49 U.S.C. 106(f), 106(g), 40113, to that on the prior engine installation. 44701, 44702 and 44704. (d) Installation of engines of more than 1,830 cubic inches displacement or having certificated SOURCE: Docket No. 5066, 29 FR 18291, Dec. 24, 1964, unless otherwise noted. take-off rating of more than 1,350 horsepower.

SPECIAL FEDERAL AVIATION REGULATION 1 It is not intended to waive compliance NO. 13 with such airworthiness requirements as are included in the operating parts of the Civil 1. Applicability. Contrary provisions of the Air Regulations for specific types of oper- Civil Air Regulations regarding certification ation.

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Engines of more than 1,830 cubic inches dis- the application of the take-off, en route, and placement or having certificated take-off landing limitations prescribed for transport rating of more than 1,350 horsepower may be category airplanes in the operating parts of installed if compliance is shown with the en- the Civil Air Regulations. gine installation requirements of Part 4b: (d) Performance operating limitations. Each Provided, That where literal compliance with airplane for which new maximum certifi- the engine installation requirements of Part cated weights are established in accordance 4b is extremely difficult to accomplish and with paragraphs (a) or (b) of this section would not contribute materially to the ob- shall be considered a transport category air- jective sought, and the Administrator finds plane for the purpose of complying with the that the experience with the DC–3 or L–18 performance operating limitations applica- airplanes justifies it, he is authorized to ac- ble to the operations in which it is utilized. cept such measures of compliance as he finds 5. Reference. Unless otherwise provided, all will effectively accomplish the basic objec- references in this regulation to Part 4a and tive. Part 4b are those parts of the Civil Air Regu- 4. Establishment of new maximum certificated lations in effect on September 1, 1953. 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, and the appropriate manual material ob- [19 FR 5039, Aug. 11, 1954. Redesignated at 29 tained, as provided in this section. FR 19099, Dec. 30, 1964] NOTE: Transport category performance requirements result in the establishment of SPECIAL FEDERAL AVIATION REGULATION maximum certificated weights for various NO. 109 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. (c) Airplane flight manual-performance oper- Additionally, there must be a placard in- ating information. An approved airplane flight stalled adjacent to each door that can be manual shall be provided for each DC–3 and used as a passenger boarding door that states L–18 airplane which has had new maximum that the maximum passenger capacity is 60. certificated weights established under this The placard must be clearly legible to pas- section. The airplane flight manual shall sengers entering the airplane. contain the applicable performance informa- (b) For airplanes outfitted with interior tion prescribed in that part of the regula- doors under paragraph 10 of this SFAR, the tions under which the new certificated airplane flight manual (AFM) must include weights were established and such additional an appropriate limitation that the airplane information as may be necessary to enable must be staffed with at least the following

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number of flight attendants who meet the re- (1) Existing Criteria. All injury protection quirements of 14 CFR 91.533(b): criteria of § 25.562(c)(1) through (c)(6) apply to (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 49 CFR part 572, subpart F. Torso contact (f) For passenger capacities of 45 to 60 pas- during rebound is acceptable and need not be sengers, analysis must be submitted that measured. demonstrates that the airplane can be evacu- (4) Pelvis. If the pelvis of an ATD at any ated in less than 90 seconds under the condi- seat place impacts seat and/or adjacent tions specified in § 25.803 and appendix J to structure during testing, pelvic lateral accel- part 25. eration injury criteria must be substantiated (g) In order for any airplane certified under by dynamic test or by rational analysis this SFAR to be placed in part 135 or part 121 based on previous test(s) of a similar seat in- operations, the airplane must be brought stallation. Pelvic lateral acceleration may back into full compliance with the applica- not exceed 130g. Pelvic acceleration data ble operational part. must be processed as defined in FMVSS part Equipment and Design 571.214, section S6.13.5 (49 CFR 571.214). (5) Body-to-Wall/Furnishing Contact. If the 3. General. Unless otherwise noted, compli- seat is installed aft of a structure—such as ance is required with the applicable certifi- an interior wall or furnishing that may con- cation basis for the airplane. Some provi- tact the pelvis, upper arm, chest, or head of sions of this SFAR impose alternative re- an occupant seated next to the structure— quirements to certain airworthiness stand- the structure or a conservative representa- ards that do not apply to airplanes certifi- tion of the structure and its stiffness must cated to earlier standards. Those airplanes be included in the tests. It is recommended, with an earlier certification basis are not re- but not required, that the contact surface of quired to comply with those alternative re- the actual structure be covered with at least quirements. two inches of energy absorbing protective 4. Occupant Protection. padding (foam or equivalent) such as (a) Firm Handhold. In lieu of the require- Ensolite. ments of § 25.785(j), there must be means pro- (6) Shoulder Strap Loads. Where upper torso vided to enable persons to steady themselves straps (shoulder straps) are used for sofa oc- in moderately rough air while occupying cupants, the tension loads in individual aisles that are along the cabin sidewall, or straps may not exceed 1,750 pounds. If dual where practicable, bordered by seats (seat straps are used for restraining the upper backs providing a 25-pound minimum break- torso, the total strap tension loads may not away force are an acceptable means of com- exceed 2,000 pounds. pliance). (7) Occupant Retention. All side-facing seats (b) Injury criteria for multiple occupancy require end closures or other means to pre- side-facing seats. The following require- vent the ATD’s pelvis from translating be- ments are only applicable to airplanes that yond the end of the seat at any time during are subject to § 25.562. testing.

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(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 compartment, easily visible to all persons of at least 160 microlamberts. The color may entering the cabin in the immediate vicinity be reversed in the case of a sign that is self- of each passenger entry door. illuminated by other than electrical means. 7. Distance Between Exits. For an airplane (b) Floor Proximity Escape Path Marking. In that is required to comply with § 25.807(f)(4), lieu of the requirements of § 25.812(e)(1), for in effect as of July 24, 1989, which has more cabin seating compartments that do not than one passenger emergency exit on each have the main cabin aisle entering and side of the fuselage, no passenger emergency exiting the compartment, the following are exit may be more than 60 feet from any adja- applicable: cent passenger emergency exit on the same (1) After a passenger leaves any passenger side of the same deck of the fuselage, as seat in the compartment, he/she must be measured parallel to the airplane’s longitu- able to exit the compartment to the main

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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 require- 14. Cooktops. Each cooktop must be de- ments of § 25.813(e), interior doors may be in- signed and installed to minimize any poten- stalled between passenger seats and exits, tial threat to the airplane, passengers, and provided the following requirements are met. crew. Compliance with this requirement (a) Each door between any passenger seat, must be found in accordance with the fol- occupiable for taxi, takeoff, and landing, and lowing criteria: any emergency exit must have a means to (a) Means, such as conspicuous burner-on signal to the flightcrew, at the flightdeck, indicators, physical barriers, or handholds, that the door is in the open position for taxi, must be installed to minimize the potential takeoff and landing. for inadvertent personnel contact with hot (b) Appropriate procedures/limitations surfaces of both the cooktop and cookware. must be established to ensure that any such Conditions of turbulence must be considered. door is in the open configuration for takeoff (b) Sufficient design means must be in- and landing. cluded to restrain cookware while in place (c) Each door between any passenger seat on the cooktop, as well as representative and any exit must have dual means to retain contents, e.g., soup, sauces, etc., from the ef- it in the open position, each of which is capa- fects of flight loads and turbulence. Re- ble of reacting the inertia loads specified in straints must be provided to preclude haz- § 25.561. ardous movement of cookware and contents. (d) Doors installed across a longitudinal These restraints must accommodate any aisle must translate laterally to open and cookware that is identified for use with the close, e.g., pocket doors. cooktop. Restraints must be designed to be (e) Each door between any passenger seat easily utilized and effective in service. The and any exit must be frangible in either di- cookware restraint system should also be de- rection. signed so that it will not be easily disabled, (f) Each door between any passenger seat thus rendering it unusable. Placarding must and any exit must be operable from either be installed which prohibits the use of side, and if a locking mechanism is installed, cookware that cannot be accommodated by it must be capable of being unlocked from ei- the restraint system. ther side without the use of special tools. (c) Placarding must be installed which pro- 11. Width of Aisle. Compliance is required hibits the use of cooktops (i.e., power on any with § 25.815, except that aisle width may be burner) during taxi, takeoff, and landing. reduced to 0 inches between passenger seats (d) Means must be provided to address the during in-flight operations only, provided possibility of a fire occurring on or in the that the applicant demonstrates that all immediate vicinity of the cooktop. Two ac- areas of the cabin are easily accessible by a ceptable means of complying with this re- crew member in the event of an emergency quirement are as follows: (e.g., in-flight fire, decompression). Addition- (1) Placarding must be installed that pro- ally, instructions must be provided at each hibits any burner from being powered when passenger seat for restoring the aisle width the cooktop is unattended. (NOTE: This required by § 25.815. Procedures must be es- would prohibit a single person from cooking tablished and documented in the AFM to en- on the cooktop and intermittently serving

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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 that comply with the flammability require- [Doc. No. FAA–2007–28250, 74 FR 21541, May 8, ments of part III of appendix F to part 25. 2009] 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, consideration must also be given to ensure that the § 25.1 Applicability. 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 system independent of the airplane cabin and (b) Each person who applies under cargo ventilation system. Procedures and Part 21 for such a certificate or change time intervals must be established to inspect must show compliance with the appli- and clean or replace the ventilation system cable requirements in this part. to prevent a fire hazard from the accumulation of flammable oils and be included in the § 25.2 Special retroactive require- instructions for continued airworthiness. ments. The ventilation system ducting must be pro- The following special retroactive re- tected by a flame arrestor. [NOTE: The applicant may find additional useful information quirements are applicable to an air- in Society of Automotive Engineers, Aero- plane for which the regulations ref- space Recommended Practice 85, Rev. E, en- erenced in the type certificate predate titled ‘‘Air Conditioning Systems for Sub- the sections specified below— 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 occupants and will not lead to the loss of type certificate) involving an increase structural strength due to airplane corro- in passenger seating capacity to a total sion. greater than that for which the air- (h) Cooktop installations must provide plane has been type certificated must adequate space for the user to immediately show that the airplane concerned escape a hazardous cooktop condition. meets the requirements of: (i) A means to shut off power to the (1) Sections 25.721(d), 25.783(g), cooktop must be provided at the galley con- 25.785(c), 25.803(c)(2) through (9), 25.803 taining the cooktop and in the cockpit. If additional switches are introduced in the cock- (d) and (e), 25.807 (a), (c), and (d), 25.809 pit, revisions to smoke or fire emergency (f) and (h), 25.811, 25.812, 25.813 (a), (b), procedures of the AFM will be required. and (c), 25.815, 25.817, 25.853 (a) and (b), (j) If the cooktop is required to have a lid 25.855(a), 25.993(f), and 25.1359(c) in ef- to enclose the cooktop there must be a fect on October 24, 1967, and

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(2) Sections 25.803(b) and 25.803(c)(1) § 25.5 Incorporations by reference. in effect on April 23, 1969. (a) The materials listed in this sec- (b) Irrespective of the date of application are incorporated by reference in tion, each applicant for a supplemental the corresponding sections noted. type certificate (or an amendment to a These incorporations by reference were type certificate) for an airplane manu- approved by the Director of the Federal factured after October 16, 1987, must Register in accordance with 5 U.S.C. show that the airplane meets the re- 552(a) and 1 CFR part 51. These mate- quirements of § 25.807(c)(7) in effect on rials are incorporated as they exist on July 24, 1989. the date of the approval, and notice of (c) Compliance with subsequent revi- any change in these materials will be sions to the sections specified in para- published in the FEDERAL REGISTER. graph (a) or (b) of this section may be The materials are available for pur- 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, For information on the availability of 2000] this material at NARA, call 202–741– § 25.3 Special provisions for ETOPS 6030, or go to: http://www.archives.gov/ type design approvals. federal-register/cfr/ibr-locations.html. (b) The following materials are avail- (a) Applicability. This section applies able for purchase from the following to an applicant for ETOPS type design address: The National Technical Infor- approval of an airplane: mation Services (NTIS), Springfield, (1) That has an existing type certifi- Virginia 22166. cate on February 15, 2007; or (1) Fuel Tank Flammability Assess- (2) For which an application for an ment Method User’s Manual, dated original type certificate was submitted May 2008, document number DOT/FAA/ before February 15, 2007. AR–05/8, IBR approved for § 25.981 and (b) Airplanes with two engines. (1) For Appendix N. It can also be obtained at ETOPS type design approval of an air- the following Web site: http:// plane up to and including 180 minutes, www.fire.tc.faa.gov/systems/fueltank/ an applicant must comply with FTFAM.stm. § 25.1535, except that it need not comply (2) [Reserved] with the following provisions of Appen- [73 FR 42494, July 21, 2008, as amended by dix K, K25.1.4, of this part: Doc. No. FAA–2018–0119, Amdt. 21–101, 83 FR (i) K25.1.4(a), fuel system pressure 9169, Mar. 5, 2018] and flow requirements; (ii) K25.1.4(a)(3), low fuel alerting; Subpart B—Flight and (iii) K25.1.4(c), engine oil tank design. GENERAL (2) For ETOPS type design approval § 25.21 Proof of compliance. of an airplane beyond 180 minutes an applicant must comply with § 25.1535. (a) Each requirement of this subpart (c) Airplanes with more than two en- must be met at each appropriate com- gines. An applicant for ETOPS type de- bination of weight and center of grav- sign approval must comply with ity within the range of loading conditions for which certification is re- § 25.1535 for an airplane manufactured quested. This must be shown— on or after February 17, 2015, except (1) By tests upon an airplane of the that, for an airplane configured for a type for which certification is re- three person flight crew, the applicant quested, or by calculations based on, need not comply with Appendix K, and equal in accuracy to, the results of K25.1.4(a)(3), of this part, low fuel alert- testing; and ing. (2) By systematic investigation of [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, each probable combination of weight 2007] and center of gravity, if compliance

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cannot be reasonably inferred from tions defined in Appendix O of this combinations investigated. part, each requirement of this subpart, (b) [Reserved] except §§ 25.105, 25.107, 25.109, 25.111, (c) The controllability, stability, 25.113, 25.115, 25.121, 25.123, 25.143(b)(1), trim, and stalling characteristics of (b)(2), and (c)(1), 25.149, 25.201(c)(2), the airplane must be shown for each al- 25.207(c), (d), and (e)(1), 25.239, and titude up to the maximum expected in 25.251(b) through (e), must be met in operation. the Appendix O icing conditions for (d) Parameters critical for the test which certification is not sought in being conducted, such as weight, load- order to allow a safe exit from those ing (center of gravity and inertia), air- conditions. Compliance must be shown speed, power, and wind, must be main- using the ice accretions defined in part tained within acceptable tolerances of II, paragraphs (b) and (d) of Appendix the critical values during flight test- O, assuming normal operation of the ing. airplane and its ice protection system (e) If compliance with the flight in accordance with the operating limi- characteristics requirements is depend- tations and operating procedures estab- ent upon a stability augmentation sys- lished by the applicant and provided in tem or upon any other automatic or the airplane flight manual. power-operated system, compliance (4) If the applicant seeks certifi- must be shown with §§ 25.671 and 25.672. cation for flight in any portion of the (f) In meeting the requirements of icing conditions of Appendix O of this §§ 25.105(d), 25.125, 25.233, and 25.237, the part, each requirement of this subpart, wind velocity must be measured at a except §§ 25.121(a), 25.123(c), 25.143(b)(1) height of 10 meters above the surface, and (2), 25.149, 25.201(c)(2), 25.239, and or corrected for the difference between 25.251(b) through (e), must be met in the height at which the wind velocity the Appendix O icing conditions for is measured and the 10-meter height. (g) The requirements of this subpart which certification is sought. Section associated with icing conditions apply 25.207(c) and (d) must be met in the only if the applicant is seeking certifi- landing configuration in the Appendix cation for flight in icing conditions. O icing conditions for which certifi- (1) Paragraphs (g)(3) and (4) of this cation is sought, but need not be met section apply only to airplanes with for other configurations. Compliance one or both of the following attributes: must be shown using the ice accretions (i) Maximum takeoff gross weight is defined in part II, paragraphs (c) and less than 60,000 lbs; or (d) of Appendix O, assuming normal op- (ii) The airplane is equipped with re- eration of the airplane and its ice pro- versible flight controls. tection system in accordance with the (2) Each requirement of this subpart, operating limitations and operating except §§ 25.121(a), 25.123(c), 25.143(b)(1) procedures established by the applicant and (2), 25.149, 25.201(c)(2), 25.239, and and provided in the airplane flight 25.251(b) through (e), must be met in manual. the icing conditions specified in Appen- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as dix C of this part. Section 25.207(c) and amended by Amdt. 25–23, 35 FR 5671, Apr. 8, (d) must be met in the landing configu- 1970; Amdt. 25–42, 43 FR 2320, Jan. 16, 1978; ration in the icing conditions specified Amdt. 25–72, 55 FR 29774, July 20, 1990; Amdt. in Appendix C, but need not be met for 25–121, 72 FR 44665, Aug. 8, 2007 Amdt. 25–135, other configurations. Compliance must 76 FR 74654, Dec. 1, 2011; Amdt. 25–140, 79 FR be shown using the ice accretions de- 65524, Nov. 4, 2014] fined in part II of Appendix C of this part, assuming normal operation of the § 25.23 Load distribution limits. airplane and its ice protection system (a) Ranges of weights and centers of in accordance with the operating limi- gravity within which the airplane may tations and operating procedures estab- be safely operated must be established. lished by the applicant and provided in If a weight and center of gravity com- the airplane flight manual. bination is allowable only within cer- (3) If the applicant does not seek certain load distribution limits (such as tification for flight in all icing condi- spanwise) that could be inadvertently

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exceeded, these limits and the cor- § 25.27 Center of gravity limits. responding weight and center of grav- The extreme forward and the extreme ity combinations must be established. aft center of gravity limitations must (b) The load distribution limits may be established for each practicably sep- not exceed— arable operating condition. No such (1) The selected limits; limit may lie beyond— (2) The limits at which the structure (a) The extremes selected by the ap- is proven; or plicant; (3) The limits at which compliance (b) The extremes within which the with each applicable flight require- structure is proven; or ment of this subpart is shown. (c) The extremes within which compliance with each applicable flight re- § 25.25 Weight limits. quirement is shown.

(a) Maximum weights. Maximum § 25.29 Empty weight and cor- weights corresponding to the airplane responding center of gravity. operating conditions (such as ramp, (a) The empty weight and cor- ground or water taxi, takeoff, en route, responding center of gravity must be and landing), environmental conditions determined by weighing the airplane (such as altitude and temperature), and with— loading conditions (such as zero fuel (1) Fixed ballast; weight, center of gravity position and (2) Unusable fuel determined under weight distribution) must be estab- § 25.959; and lished so that they are not more than— (3) Full operating fluids, including— (1) The highest weight selected by (i) Oil; the applicant for the particular condi- (ii) Hydraulic fluid; and tions; or (iii) Other fluids required for normal (2) The highest weight at which com- operation of airplane systems, except pliance with each applicable structural potable water, lavatory precharge 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 limits. (1) The lowest weight selected by the 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— (1) Safe operation under normal oper- with each structural loading condition ating conditions; and of this part is shown); or (2) Compliance with the performance (3) The lowest weight at which com- requirements of §§ 25.101 through 25.125. pliance with each applicable flight re- (b) There must be a propeller speed quirement is shown. limiting means at the governor. It [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as must limit the maximum possible gov- amended by Amdt. 25–23, 35 FR 5671, Apr. 8, erned engine speed to a value not ex- 1970; Amdt. 25–63, 53 FR 16365, May 6, 1988] ceeding the maximum allowable r.p.m.

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(c) The means used to limit the low Altitude Vapor Specific humidity Density ratio pitch position of the propeller blades pressure e w (Lb. moisture r / s = H (ft.) (In. Hg.) per lb. dry air) 0.0023769 must be set so that the engine does not exceed 103 percent of the maximum al- 20,000 .01978 .000896 .53263 lowable engine rpm or 99 percent of an 25,000 .00778 .000436 .44806 approved maximum overspeed, which- (c) The performance must correspond ever is greater, with— to the propulsive thrust available (1) The propeller blades at the low under the particular ambient atmos- pitch limit and governor inoperative; pheric conditions, the particular flight (2) The airplane stationary under condition, and the relative humidity standard atmospheric conditions with specified in paragraph (b) of this sec- no wind; and tion. The available propulsive thrust (3) The engines operating at the take- must correspond to engine power or off manifold pressure limit for recipro- thrust, not exceeding the approved cating engine powered airplanes or the power or thrust less— maximum takeoff torque limit for tur- (1) Installation losses; and bopropeller engine-powered airplanes. (2) The power or equivalent thrust [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as absorbed by the accessories and serv- amended by Amdt. 25–57, 49 FR 6848, Feb. 23, ices appropriate to the particular am- 1984; Amdt. 25–72, 55 FR 29774, July 20, 1990] bient atmospheric conditions and the particular flight condition. PERFORMANCE (d) Unless otherwise prescribed, the applicant must select the takeoff, en § 25.101 General. route, approach, and landing configura- (a) Unless otherwise prescribed, air- tions for the airplane. planes must meet the applicable per- (e) The airplane configurations may formance requirements of this subpart vary with weight, altitude, and tem- for ambient atmospheric conditions perature, to the extent they are com- and still air. patible with the operating procedures (b) The performance, as affected by required by paragraph (f) of this sec- engine power or thrust, must be based tion. on the following relative humidities; (f) Unless otherwise prescribed, in de- (1) For turbine engine powered air- termining the accelerate-stop dis- planes, a relative humidity of— tances, takeoff flight paths, takeoff (i) 80 percent, at and below standard distances, and landing distances, temperatures; and changes in the airplane’s configura- (ii) 34 percent, at and above standard tion, speed, power, and thrust, must be temperatures plus 50 °F. made in accordance with procedures es- Between these two temperatures, the tablished by the applicant for oper- relative humidity must vary linearly. ation in service. (2) For reciprocating engine powered (g) Procedures for the execution of airplanes, a relative humidity of 80 per- balked landings and missed approaches cent in a standard atmosphere. Engine associated with the conditions pre- power corrections for vapor pressure scribed in §§ 25.119 and 25.121(d) must be must be made in accordance with the established. following table: (h) The procedures established under paragraphs (f) and (g) of this section Altitude Vapor Specific humidity Density ratio must— pressure e w (Lb. moisture r / s = H (ft.) (In. Hg.) per lb. dry air) 0.0023769 (1) Be able to be consistently executed in service by crews of average 0 0.403 0.00849 0.99508 1,000 .354 .00773 .96672 skill; 2,000 .311 .00703 .93895 (2) Use methods or devices that are 3,000 .272 .00638 .91178 safe and reliable; and 4,000 .238 .00578 .88514 (3) Include allowance for any time 5,000 .207 .00523 .85910 6,000 .1805 .00472 .83361 delays, in the execution of the proce- 7,000 .1566 .00425 .80870 dures, that may reasonably be expected 8,000 .1356 .00382 .78434 in service. 9,000 .1172 .00343 .76053 10,000 .1010 .00307 .73722 (i) The accelerate-stop and landing 15,000 .0463 .001710 .62868 distances prescribed in §§ 25.109 and

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25.125, respectively, must be deter- (c) Starting from the stabilized trim mined with all the airplane wheel condition, apply the longitudinal con- brake assemblies at the fully worn trol to decelerate the airplane so that limit of their allowable wear range. the speed reduction does not exceed one knot per second. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (d) In addition to the requirements of amended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; Amdt. 25–92, 63 FR 8318, Feb. 18, 1998] paragraph (a) of this section, when a device that abruptly pushes the nose § 25.103 Stall speed. down at a selected angle of attack (e.g., a stick pusher) is installed, the ref- (a) The reference stall speed, V , is a SR erence stall speed, V , may not be less calibrated airspeed defined by the ap- SR than 2 knots or 2 percent, whichever is plicant. V may not be less than a 1-g SR greater, above the speed at which the stall speed. V is expressed as: SR device operates. V [Doc. No. 28404, 67 FR 70825, Nov. 26, 2002, as ≥ CLMAX VSR amended by Amdt. 25–121, 72 FR 44665, Aug. 8, nZW 2007] where: § 25.105 Takeoff. VCLMAX = Calibrated airspeed obtained when the load factor-corrected lift coefficient (a) The takeoff speeds prescribed by § 25.107, the accelerate-stop distance ⎛ nW⎞ prescribed by § 25.109, the takeoff path ⎜ ZW ⎟ prescribed by § 25.111, the takeoff dis- ⎝ qS ⎠ tance and takeoff run prescribed by is first a maximum during the maneuver § 25.113, and the net takeoff flight path prescribed in paragraph (c) of this section. In prescribed by § 25.115, must be deter- addition, when the maneuver is limited by a mined in the selected configuration for device that abruptly pushes the nose down at takeoff at each weight, altitude, and a selected angle of attack (e.g., a stick push- ambient temperature within the oper-

er), VCLMAX may not be less than the speed ex- ational limits selected by the appli- isting at the instant the device operates; cant— 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; S = Aerodynamic reference wing area; and figuration used to show compliance q = Dynamic pressure. with § 25.121(b), and with the most critical of the takeoff ice accretion(s) de- (b) VCLMAX is determined with: fined in appendices C and O of this (1) Engines idling, or, if that result- part, as applicable, in accordance with ant thrust causes an appreciable de- § 25.21(g): crease in stall speed, not more than (i) The stall speed at maximum take- zero thrust at the stall speed; off weight exceeds that in non-icing (2) Propeller pitch controls (if appli- conditions by more than the greater of cable) in the takeoff position; 3 knots CAS or 3 percent of VSR; or (3) The airplane in other respects (ii) The degradation of the gradient (such as flaps, landing gear, and ice ac- of climb determined in accordance with cretions) in the condition existing in § 25.121(b) is greater than one-half of the test or performance standard in the applicable actual-to-net takeoff which VSR is being used; flight path gradient reduction defined (4) The weight used when VSR is being in § 25.115(b). used as a factor to determine compli- (b) No takeoff made to determine the ance with a required performance data required by this section may re- standard; quire exceptional piloting skill or (5) The center of gravity position alertness. that results in the highest value of ref- (c) The takeoff data must be based erence stall speed; and on— (6) The airplane trimmed for straight (1) In the case of land planes and am- flight at a speed selected by the appli- phibians: cant, but not less than 1.13VSR and not (i) Smooth, dry and wet, hard-sur- greater than 1.3VSR. faced runways; and 210

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(ii) At the option of the applicant, (ii) Turbojet powered airplanes with grooved or porous friction course wet, provisions for obtaining a significant hard-surfaced runways. reduction in the one-engine-inoper- (2) Smooth water, in the case of sea- ative power-on stall speed; and planes and amphibians; and (3) 1.10 times VMC established under (3) Smooth, dry snow, in the case of § 25.149. skiplanes. (c) V2, in terms of calibrated air- (d) The takeoff data must include, speed, must be selected by the appli- within the established operational lim- cant to provide at least the gradient of its of the airplane, the following oper- climb required by § 25.121(b) but may ational correction factors: not be less than— (1) Not more than 50 percent of nomi- (1) V2MIN; nal wind components along the takeoff (2) VR plus the speed increment at- path opposite to the direction of take- tained (in accordance with § 25.111(c)(2)) off, and not less than 150 percent of before reaching a height of 35 feet nominal wind components along the above the takeoff surface; and takeoff path in the direction of takeoff. (3) A speed that provides the maneu- (2) Effective runway gradients. vering capability specified in § 25.143(h). [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (d) VMU is the calibrated airspeed at amended by Amdt. 25–92, 63 FR 8318, Feb. 18, and above which the airplane can safe- 1998; Amdt. 25–121, 72 FR 44665, Aug. 8, 2007; ly lift off the ground, and continue Amdt. 25–140, 79 FR 65525, Nov. 4, 2014] the takeoff. VMU speeds must be se- § 25.107 Takeoff speeds. lected by the applicant throughout the range of thrust-to-weight ratios to be (a) V1 must be established in relation certificated. These speeds may be es- to VEF as follows: tablished from free air data if these (1) VEF is the calibrated airspeed at data are verified by ground takeoff which the critical engine is assumed to tests. fail. VEF must be selected by the appli- (e) VR, in terms of calibrated air- cant, but may not be less than VMCG de- speed, must be selected in accordance termined under § 25.149(e). with the conditions of paragraphs (e)(1) (2) V1, in terms of calibrated air- through (4) of this section: speed, is selected by the applicant; (1) V may not be less than— however, V may not be less than V R 1 EF (i) V ; plus the speed gained with critical en- 1 gine inoperative during the time inter- (ii) 105 percent of VMC; val between the instant at which the (iii) The speed (determined in accord- critical engine is failed, and the in- ance with § 25.111(c)(2)) that allows stant at which the pilot recognizes and reaching V2 before reaching a height of reacts to the engine failure, as indi- 35 feet above the takeoff surface; or cated by the pilot’s initiation of the (iv) A speed that, if the airplane is first action (e.g., applying brakes, re- rotated at its maximum practicable ducing thrust, deploying speed brakes) rate, will result in a VLOF of not less to stop the airplane during accelerate- than — stop tests. (A) 110 percent of VMU in the all-en- (b) V2MIN, in terms of calibrated air- gines-operating condition, and 105 per- speed, may not be less than— cent of VMU determined at the thrust- (1) 1.13 VSR for— to-weight ratio corresponding to the (i) Two-engine and three-engine tur- one-engine-inoperative condition; or bopropeller and reciprocating engine (B) If the VMU attitude is limited by powered airplanes; and the geometry of the airplane (i.e., tail (ii) Turbojet powered airplanes with- contact with the runway), 108 percent out provisions for obtaining a signifi- of VMU in the all-engines-operating cant reduction in the one-engine-inop- condition, and 104 percent of VMU deter- erative power-on stall speed; mined at the thrust-to-weight ratio (2) 1.08 VSR for— corresponding to the one-engine-inop- (i) Turbopropeller and reciprocating erative condition. engine powered airplanes with more (2) For any given set of conditions than three engines; and (such as weight, configuration, and

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temperature), a single value of VR, ob- (ii) Allow the airplane to accelerate tained in accordance with this para- from VEF to the highest speed reached graph, must be used to show compli- during the rejected takeoff, assuming ance with both the one-engine-inoper- the critical engine fails at VEF and the ative and the all-engines-operating pilot takes the first action to reject takeoff provisions. the takeoff at the V1 for takeoff from a (3) It must be shown that the one-en- dry runway; and gine-inoperative takeoff distance, (iii) Come to a full stop on a dry run- using a rotation speed of 5 knots less way from the speed reached as pre- than VR established in accordance with scribed in paragraph (a)(1)(ii) of this paragraphs (e)(1) and (2) of this section, section; plus does not exceed the corresponding one- (iv) A distance equivalent to 2 sec- engine-inoperative takeoff distance onds at the V1 for takeoff from a dry using the established VR. The takeoff runway. distances must be determined in ac- (2) The sum of the distances nec- cordance with § 25.113(a)(1). essary to— (4) Reasonably expected variations in service from the established takeoff (i) Accelerate the airplane from a procedures for the operation of the air- standing start with all engines oper- plane (such as over-rotation of the air- ating to the highest speed reached dur- plane and out-of-trim conditions) may ing the rejected takeoff, assuming the not result in unsafe flight characteris- pilot takes the first action to reject tics or in marked increases in the the takeoff at the V1 for takeoff from a scheduled takeoff distances established dry runway; and in accordance with § 25.113(a). (ii) With all engines still operating, come to a full stop on dry runway from (f) VLOF is the calibrated airspeed at which the airplane first becomes air- the speed reached as prescribed in para- borne. graph (a)(2)(i) of this section; plus (g) VFTO, in terms of calibrated air- (iii) A distance equivalent to 2 sec- speed, must be selected by the appli- onds at the V1 for takeoff from a dry cant to provide at least the gradient of runway. climb required by § 25.121(c), but may (b) The accelerate-stop distance on a not be less than— wet runway is the greater of the fol- (1) 1.18 VSR; and lowing distances: (2) A speed that provides the maneu- (1) The accelerate-stop distance on a vering capability specified in § 25.143(h). dry runway determined in accordance (h) In determining the takeoff speeds with paragraph (a) of this section; or V1, VR, and V2 for flight in icing condi- (2) The accelerate-stop distance de- tions, the values of VMCG, VMC, and VMU termined in accordance with paragraph determined for non-icing conditions (a) of this section, except that the run- may be used. way is wet and the corresponding wet [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as runway values of VEF and V1 are used. amended by Amdt. 25–38, 41 FR 55466, Dec. 20, In determining the wet runway accel- 1976; Amdt. 25–42, 43 FR 2320, Jan. 16, 1978; erate-stop distance, the stopping force Amdt. 25–92, 63 FR 8318, Feb. 18, 1998; Amdt. from the wheel brakes may never ex- 25–94, 63 FR 8848, Feb. 23, 1998; Amdt. 25–108, ceed: 67 FR 70826, Nov. 26, 2002; Amdt. 25–121, 72 FR 44665, Aug. 8, 2007; Amdt. 25–135, 76 FR 74654, (i) The wheel brakes stopping force Dec. 1, 2011] determined in meeting the requirements of § 25.101(i) and paragraph (a) of § 25.109 Accelerate-stop distance. this section; and (a) The accelerate-stop distance on a (ii) The force resulting from the wet dry runway is the greater of the fol- runway braking coefficient of friction lowing distances: determined in accordance with para- (1) The sum of the distances nec- graphs (c) or (d) of this section, as ap- essary to— plicable, taking into account the dis- (i) Accelerate the airplane from a tribution of the normal load between standing start with all engines oper- braked and unbraked wheels at the ating to VEF for takeoff from a dry run- most adverse center-of-gravity position way; approved for takeoff.

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(c) The wet runway braking coeffi- (1) The maximum tire-to-ground wet cient of friction for a smooth wet run- runway braking coefficient of friction way is defined as a curve of friction co- is defined as: efficient versus ground speed and must be computed as follows:

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

μt/gMAX = maximum tire-to-ground braking On-Off ...... 0.30 coefficient; 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 temperature, rain, snow, and ice) with- coefficient; 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 ment and must correspond to the most the takeoff path must be positive at critical condition prevailing in the seg- each point; ment; (2) The airplane must reach V2 before (3) The flight path must be based on it is 35 feet above the takeoff surface the airplane’s performance without and must continue at a speed as close ground effect; and as practical to, but not less than V2, (4) The takeoff path data must be until it is 400 feet above the takeoff checked by continuous demonstrated surface; takeoffs up to the point at which the (3) At each point along the takeoff airplane is out of ground effect and its path, starting at the point at which the speed is stabilized, to ensure that the airplane reaches 400 feet above the path is conservative relative to the takeoff surface, the available gradient continous path. of climb may not be less than— (i) 1.2 percent for two-engine air- The airplane is considered to be out of planes; the ground effect when it reaches a (ii) 1.5 percent for three-engine air- height equal to its wing span. planes; and (e) For airplanes equipped with (iii) 1.7 percent for four-engine air- standby power rocket engines, the planes. takeoff path may be determined in ac- (4) The airplane configuration may cordance with section II of appendix E. not be changed, except for gear retrac- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tion and automatic propeller feath- amended by Amdt. 25–6, 30 FR 8468, July 2, ering, and no change in power or thrust 1965; Amdt. 25–42, 43 FR 2321, Jan. 16, 1978; that requires action by the pilot may Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; Amdt. be made until the airplane is 400 feet 25–72, 55 FR 29774, July 20, 1990; Amdt. 25–94, above the takeoff surface; and 63 FR 8848, Feb. 23, 1998; Amdt. 25–108, 67 FR (5) If § 25.105(a)(2) requires the takeoff 70826, Nov. 26, 2002; Amdt. 25–115, 69 FR 40527, path to be determined for flight in July 2, 2004; Amdt. 25–121, 72 FR 44666; Aug. 8, 2007; Amdt. 25–140, 79 FR 65525, Nov. 4, 2014] icing conditions, the airborne part of the takeoff must be based on the air- § 25.113 Takeoff distance and takeoff plane drag: run. (i) With the most critical of the takeoff ice accretion(s) defined in Appen- (a) Takeoff distance on a dry runway dices C and O of this part, as applica- is the greater of— ble, in accordance with § 25.21(g), from (1) The horizontal distance along the a height of 35 feet above the takeoff takeoff path from the start of the take- surface up to the point where the air- off to the point at which the airplane is plane is 400 feet above the takeoff sur- 35 feet above the takeoff surface, deter- face; and mined under § 25.111 for a dry runway; (ii) With the most critical of the final or takeoff ice accretion(s) defined in Ap- (2) 115 percent of the horizontal dis- pendices C and O of this part, as appli- tance along the takeoff path, with all cable, in accordance with § 25.21(g), engines operating, from the start of the from the point where the airplane is 400 takeoff to the point at which the air- feet above the takeoff surface to the plane is 35 feet above the takeoff sur- end of the takeoff path. face, as determined by a procedure con- (d) The takeoff path must be deter- sistent with § 25.111. mined by a continuous demonstrated (b) Takeoff distance on a wet runway takeoff or by synthesis from segments. is the greater of— If the takeoff path is determined by the (1) The takeoff distance on a dry run- segmental method— way determined in accordance with (1) The segments must be clearly de- paragraph (a) of this section; or fined and must be related to the dis- (2) The horizontal distance along the tinct changes in the configuration, takeoff path from the start of the take- power or thrust, and speed; off to the point at which the airplane is (2) The weight of the airplane, the 15 feet above the takeoff surface, configuration, and the power or thrust achieved in a manner consistent with must be constant throughout each seg- the achievement of V2 before reaching 215

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35 feet above the takeoff surface, deter- (determined in accordance with § 25.111 mined under § 25.111 for a wet runway. and with paragraph (a) of this section) (c) If the takeoff distance does not in- reduced at each point by a gradient of clude a clearway, the takeoff run is climb equal to— equal to the takeoff distance. If the (1) 0.8 percent for two-engine air- takeoff distance includes a clearway— planes; (1) The takeoff run on a dry runway (2) 0.9 percent for three-engine air- is the greater of— planes; and (i) The horizontal distance along the (3) 1.0 percent for four-engine air- takeoff path from the start of the take- planes. off to a point equidistant between the (c) The prescribed reduction in climb point at which 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 dis- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tance along the takeoff path, with all amended by Amdt. 25–92, 63 FR 8320, Feb. 18, engines operating, from the start of the 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 Compliance with the requirements of feet above the takeoff surface, deter- §§ 25.119 and 25.121 must be shown at mined by a procedure consistent with each weight, altitude, and ambient § 25.111. temperature within the operational (2) The takeoff run on a wet runway limits established for the airplane and is the greater of— with the most unfavorable center of (i) The horizontal distance along the gravity for each configuration. takeoff path from the start of the takeoff 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 steady gradient of climb may not be 35 feet above the takeoff surface, as de- less than 3.2 percent, with the engines termined under § 25.111 for a wet run- at the power or thrust that is available way; or 8 seconds after initiation of movement (ii) 115 percent of the horizontal dis- of the power or thrust controls from tance along the takeoff path, with all the minimum flight idle to the go- engines operating, from the start of the around power or thrust setting— takeoff to a point equidistant between (a) In non-icing conditions, with a the point at which VLOF is reached and climb speed of VREF determined in ac- the point at which the airplane is 35 cordance with § 25.125(b)(2)(i); and feet above the takeoff surface, deter- (b) In icing conditions with the most mined by a procedure consistent with critical of the landing ice accretion(s) § 25.111. defined in Appendices C and O of this [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as part, as applicable, in accordance with amended by Amdt. 25–23, 35 FR 5671, Apr. 8, § 25.21(g), and with a climb speed of 1970; Amdt. 25–92, 63 FR 8320, Feb. 18, 1998] VREF determined in accordance with § 25.125(b)(2)(ii). § 25.115 Takeoff flight path. [Amdt. 25–121, 72 FR 44666; Aug. 8, 2007, as (a) The takeoff flight path shall be amended by Amdt. 25–,140, 79 FR 65525, Nov. considered to begin 35 feet above the 4, 2014] takeoff surface at the end of the takeoff distance determined in accordance § 25.121 Climb: One-engine-inoper- with § 25.113(a) or (b), as appropriate for ative. the runway surface condition. (a) Takeoff; landing gear extended. In (b) The net takeoff flight path data the critical takeoff configuration exist- must be determined so that they rep- ing along the flight path (between the resent the actual takeoff flight paths points at which the airplane reaches

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VLOF and at which the landing gear is icing conditions by more than the fully retracted) and in the configura- greater of 3 knots CAS or 3 percent of tion used in § 25.111 but without ground VSR; or effect, the steady gradient of climb (B) The degradation of the gradient must be positive for two-engine air- of climb determined in accordance with planes, and not less than 0.3 percent for § 25.121(b) is greater than one-half of three-engine airplanes or 0.5 percent the applicable actual-to-net takeoff for four-engine airplanes, at VLOF and flight path gradient reduction defined with— in § 25.115(b). (1) The critical engine inoperative (c) Final takeoff. In the en route con- and the remaining engines at the power figuration at the end of the takeoff or thrust available when retraction of path determined in accordance with the landing gear is begun in accordance § 25.111: with § 25.111 unless there is a more crit- (1) The steady gradient of climb may ical power operating condition existing not be less than 1.2 percent for two-en- later along the flight path but before gine airplanes, 1.5 percent for three-en- the point at which the landing gear is gine airplanes, and 1.7 percent for four- fully retracted; and engine airplanes, at VFTO with— (2) The weight equal to the weight (i) The critical engine inoperative existing when retraction of the landing and the remaining engines at the avail- gear is begun, determined under able maximum continuous power or § 25.111. thrust; and (b) Takeoff; landing gear retracted. In (ii) The weight equal to the weight the takeoff configuration existing at existing at the end of the takeoff path, the point of the flight path at which determined under § 25.111. the landing gear is fully retracted, and (2) The requirements of paragraph in the configuration used in § 25.111 but (c)(1) of this section must be met: without ground effect: (i) In non-icing conditions; and (1) The steady gradient of climb may (ii) In icing conditions with the most not be less than 2.4 percent for two-en- critical of the final takeoff ice accre- gine airplanes, 2.7 percent for three-en- tion(s) defined in Appendices C and O of gine airplanes, and 3.0 percent for four- this part, as applicable, in accordance engine airplanes, at V2 with: with § 25.21(g), if in the configuration (i) The critical engine inoperative, used to show compliance with § 25.121(b) the remaining engines at the takeoff with the takeoff ice accretion used to power or thrust available at the time show compliance with § 25.111(c)(5)(i): the landing gear is fully retracted, de- (A) The stall speed at maximum termined under § 25.111, unless there is takeoff weight exceeds that in non- a more critical power operating condi- icing conditions by more than the tion existing later along the flight path greater of 3 knots CAS or 3 percent of but before the point where the airplane VSR; or reaches a height of 400 feet above the (B) The degradation of the gradient takeoff surface; and of climb determined in accordance with (ii) The weight equal to the weight § 25.121(b) is greater than one-half of existing when the airplane’s landing the applicable actual-to-net takeoff gear is fully retracted, determined flight path gradient reduction defined under § 25.111. in § 25.115(b). (2) The requirements of paragraph (d) Approach. In a configuration cor- (b)(1) of this section must be met: responding to the normal all-engines- (i) In non-icing conditions; and operating procedure in which VSR for (ii) In icing conditions with the most this configuration does not exceed 110 critical of the takeoff ice accretion(s) percent of the VSR for the related all- defined in Appendices C and O of this engines-operating landing configura- part, as applicable, in accordance with tion: § 25.21(g), if in the configuration used to (1) The steady gradient of climb may show compliance with § 25.121(b) with not be less than 2.1 percent for two-en- this takeoff ice accretion: gine airplanes, 2.4 percent for three-en- (A) The stall speed at maximum gine airplanes, and 2.7 percent for four- takeoff weight exceeds that in non- engine airplanes, with—

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(i) The critical engine inoperative, three-engine airplanes, and 1.6 percent the remaining engines at the go-around for four-engine airplanes— power or thrust setting; (1) In non-icing conditions; and (ii) The maximum landing weight; (2) In icing conditions with the most (iii) A climb speed established in con- critical of the en route ice accretion(s) nection with normal landing proce- defined in Appendices C and O of this part, as applicable, in accordance with dures, but not exceeding 1.4 VSR; and (iv) Landing gear retracted. § 25.21(g), if: (2) The requirements of paragraph (i) A speed of 1.18 ‘‘VSR0 with the en (d)(1) of this section must be met: route ice accretion exceeds the en (i) In non-icing conditions; and route speed selected for non-icing conditions by more than the greater of 3 (ii) In icing conditions with the most knots CAS or 3 percent of V ; or critical of the approach ice accretion(s) SR (ii) The degradation of the gradient defined in Appendices C and O of this of climb is greater than one-half of the part, as applicable, in accordance with applicable actual-to-net flight path re- § 25.21(g). The climb speed selected for duction defined in paragraph (b) of this non-icing conditions may be used if the section. climb speed for icing conditions, com- (c) For three- or four-engine air- puted in accordance with paragraph planes, the two-engine-inoperative net (d)(1)(iii) of this section, does not ex- flight path data must represent the ac- ceed that for non-icing conditions by tual climb performance diminished by more than the greater of 3 knots CAS a gradient of climb of 0.3 percent for or 3 percent. three-engine airplanes and 0.5 percent [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as for four-engine airplanes. amended by Amdt. 25–84, 60 FR 30749, June 9, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as 1995; Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; amended by Amdt. 25–121, 72 FR 44666; Aug. 8, Amdt. 25–121, 72 FR 44666; Aug. 8, 2007; Amdt. 2007; Amdt. 25–140, 79 FR 65525, Nov. 4, 2014] 25–140, 79 FR 65525, Nov. 4, 2014] § 25.125 Landing. § 25.123 En route flight paths. (a) The horizontal distance necessary (a) For the en route configuration, to land and to come to a complete stop the flight paths prescribed in para- (or to a speed of approximately 3 knots graph (b) and (c) of this section must for water landings) from a point 50 feet be determined at each weight, altitude, above the landing surface must be de- and ambient temperature, within the termined (for standard temperatures, operating limits established for the at each weight, altitude, and wind airplane. The variation of weight along within the operational limits estab- the flight path, accounting for the pro- lished by the applicant for the air- gressive consumption of fuel and oil by plane): the operating engines, may be included (1) In non-icing conditions; and in the computation. The flight paths (2) In icing conditions with the most must be determined at a speed not less critical of the landing ice accretion(s) than VFTO, with— defined in Appendices C and O of this (1) The most unfavorable center of part, as applicable, in accordance with gravity; § 25.21(g), if VREF for icing conditions ex- (2) The critical engines inoperative; ceeds VREF for non-icing conditions by (3) The remaining engines at the more than 5 knots CAS at the max- available maximum continuous power imum landing weight. or thrust; and (b) In determining the distance in (4) The means for controlling the en- paragraph (a) of this section: gine-cooling air supply in the position (1) The airplane must be in the land- that provides adequate cooling in the ing configuration. hot-day condition. (2) A stabilized approach, with a cali- (b) The one-engine-inoperative net brated airspeed of not less than VREF, flight path data must represent the ac- must be maintained down to the 50-foot tual climb performance diminished by height. a gradient of climb of 1.1 percent for (i) In non-icing conditions, VREF may two-engine airplanes, 1.4 percent for not be less than:

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(A) 1.23 VSR0; posite to the direction of landing, and (B) VMCL established under § 25.149(f); not less than 150 percent of the nomi- and nal wind components along the landing (C) A speed that provides the maneu- path in the direction of landing. vering capability specified in § 25.143(h). (g) If any device is used that depends (ii) In icing conditions, VREF may not on the operation of any engine, and if be less than: the landing distance would be notice- (A) The speed determined in para- ably increased when a landing is made graph (b)(2)(i) of this section; with that engine inoperative, the land- (B) 1.23 VSR0 with the most critical of ing distance must be determined with the landing ice accretion(s) defined in that engine inoperative unless the use Appendices C and O of this part, as ap- of compensating means will result in a plicable, in accordance with § 25.21(g), if landing distance not more than that that speed exceeds VREF selected for with each engine operating. non-icing conditions by more than 5 [Amdt. 25–121, 72 FR 44666; Aug. 8, 2007; 72 FR knots CAS; and 50467, Aug. 31, 2007; Amdt. 25–140, 79 FR 65525, (C) A speed that provides the maneu- Nov. 4, 2014] vering capability specified in § 25.143(h) with the most critical of the landing CONTROLLABILITY AND ice accretion(s) defined in Appendices C MANEUVERABILITY and O of this part, as applicable, in accordance with § 25.21(g). § 25.143 General. (3) Changes in configuration, power (a) The airplane must be safely con- or thrust, and speed, must be made in trollable and maneuverable during— accordance with the established proce- (1) Takeoff; dures for service operation. (2) Climb; (4) The landing must be made with- (3) Level flight; out excessive vertical acceleration, (4) Descent; and tendency to bounce, nose over, ground (5) Landing. loop, porpoise, or water loop. (b) It must be possible to make a (5) The landings may not require ex- smooth transition from one flight con- ceptional piloting skill or alertness. dition to any other flight condition (c) For landplanes and amphibians, without exceptional piloting skill, the landing distance on land must be alertness, or strength, and without determined on a level, smooth, dry, danger of exceeding the airplane limit- hard-surfaced runway. In addition— load factor under any probable oper- (1) The pressures on the wheel brak- ating conditions, including— ing systems may not exceed those spec- (1) The sudden failure of the critical ified by the brake manufacturer; engine; (2) The brakes may not be used so as (2) For airplanes with three or more to cause excessive wear of brakes or engines, the sudden failure of the sec- tires; and ond critical engine when the airplane is (3) Means other than wheel brakes in the en route, approach, or landing may be used if that means— configuration and is trimmed with the (i) Is safe and reliable; critical engine inoperative; and (ii) Is used so that consistent results (3) Configuration changes, including can be expected in service; and deployment or retraction of decelera- (iii) Is such that exceptional skill is tion devices. not required to control the airplane. (c) The airplane must be shown to be (d) For seaplanes and amphibians, safely controllable and maneuverable the landing distance on water must be with the most critical of the ice accre- determined on smooth water. tion(s) appropriate to the phase of (e) For skiplanes, the landing dis- flight as defined in Appendices C and O tance on snow must be determined on of this part, as applicable, in accord- smooth, dry, snow. ance with § 25.21(g), and with the crit- (f) The landing distance data must ical engine inoperative and its pro- include correction factors for not more peller (if applicable) in the minimum than 50 percent of the nominal wind drag position: components along the landing path op- (1) At the minimum V2 for takeoff; 219

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(2) During an approach and go- (f) When demonstrating compliance around; and with the control force limitations for (3) During an approach and landing. long term application that are pre- (d) The following table prescribes, for scribed in paragraph (d) of this section, conventional wheel type controls, the the airplane must be in trim, or as near maximum control forces permitted to being in trim as practical. during the testing required by para- (g) When maneuvering at a constant graph (a) through (c) of this section: airspeed or Mach number (up to VFC/ Force, in pounds, applied to the MFC), the stick forces and the gradient control wheel or rudder pedals Pitch Roll Yaw of the stick force versus maneuvering For short term application for load factor must lie within satisfactory pitch and roll control—two limits. The stick forces must not be so hands available for control .... 75 50 great as to make excessive demands on For short term application for pitch and roll control—one the pilot’s strength when maneuvering hand available for control ...... 50 25 the airplane, and must not be so low For short term application for that the airplane can easily be over- yaw control ...... 150 stressed inadvertently. Changes of gra- For long term application ...... 10 5 20 dient that occur with changes of load (e) Approved operating procedures or factor must not cause undue difficulty conventional operating practices must in maintaining control of the airplane, be followed when demonstrating com- and local gradients must not be so low pliance with the control force limita- as to result in a danger of overcontrol- tions for short term application that ling. are prescribed in paragraph (d) of this (h) The maneuvering capabilities in a section. The airplane must be in trim, constant speed coordinated turn at for- or as near to being in trim as practical, ward center of gravity, as specified in in the preceding steady flight condi- the following table, must be free of tion. For the takeoff condition, the air- stall warning or other characteristics plane must be trimmed according to that might interfere with normal ma- the approved operating procedures. 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 ly recover from the maneuver without with § 25.143 in icing conditions— exceeding a pull control force of 50 (1) Controllability must be dem- pounds; and onstrated with the most critical of the (3) Any changes in force that the ice accretion(s) for the particular pilot must apply to the pitch control to flight phase as defined in Appendices C maintain speed with increasing sideslip and O of this part, as applicable, in ac- angle must be steadily increasing with cordance with § 25.21(g); no force reversals, unless the change in (2) It must be shown that a push force control force is gradual and easily con- is required throughout a pushover ma- trollable by the pilot without using ex- neuver down to a zero g load factor, or ceptional piloting skill, alertness, or the lowest load factor obtainable if strength. limited by elevator power or other de- (j) For flight in icing conditions be- sign characteristic of the flight control fore the ice protection system has been system. It must be possible to prompt-

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activated and is performing its in- (5) Repeat paragraph (b)(4) except tended function, it must be dem- with flaps extended. onstrated in flight with the most crit- (6) With power off, flaps extended, ical of the ice accretion(s) defined in and the airplane trimmed at 1.3 VSR1, Appendix C, part II, paragraph (e) of obtain and maintain airspeeds between this part and Appendix O, part II, para- VSW and either 1.6 VSR1 or VFE, which- graph (d) of this part, as applicable, in ever is lower. accordance with § 25.21(g), that: (c) It must be possible, without ex- (1) The airplane is controllable in a ceptional piloting skill, to prevent loss pull-up maneuver up to 1.5 g load fac- of altitude when complete retraction of tor; and the high lift devices from any position (2) There is no pitch control force re- is begun during steady, straight, level versal during a pushover maneuver flight at 1.08 VSR1 for propeller powered down to 0.5 g load factor. airplanes, or 1.13 VSR1 for turbojet pow- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ered airplanes, with— amended by Amdt. 25–42, 43 FR 2321, Jan. 16, (1) Simultaneous movement of the 1978; Amdt. 25–84, 60 FR 30749, June 9, 1995; power or thrust controls to the go- Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; around power or thrust setting; Amdt. 25–121, 72 FR 44667, Aug. 8, 2007; Amdt. 25–129, 74 FR 38339, Aug. 3, 2009; Amdt. 25–140, (2) The landing gear extended; and 79 FR 65525, Nov. 4, 2014] (3) The critical combinations of landing weights and altitudes. § 25.145 Longitudinal control. (d) If gated high-lift device control (a) It must be possible, at any point positions are provided, paragraph (c) of between the trim speed prescribed in this section applies to retractions of § 25.103(b)(6) and stall identification (as the high-lift devices from any position defined in § 25.201(d)), to pitch the nose from the maximum landing position to downward so that the acceleration to the first gated position, between gated this selected trim speed is prompt with positions, and from the last gated posi- (1) The airplane trimmed at the trim tion to the fully retracted position. speed prescribed in § 25.103(b)(6); The requirements of paragraph (c) of (2) The landing gear extended; this section also apply to retractions (3) The wing flaps (i) retracted and from each approved landing position to (ii) extended; and the control position(s) associated with (4) Power (i) off and (ii) at maximum the high-lift device configuration(s) continuous power on the engines. used to establish the go-around proce- (b) With the landing gear extended, dure(s) from that landing position. In no change in trim control, or exertion addition, the first gated control posi- of more than 50 pounds control force tion from the maximum landing posi- (representative of the maximum short tion must correspond with a configura- term force that can be applied readily tion of the high-lift devices used to es- by one hand) may be required for the tablish a go-around procedure from a following maneuvers: landing configuration. Each gated con- (1) With power off, flaps retracted, trol position must require a separate and the airplane trimmed at 1.3 VSR1, and distinct motion of the control to extend the flaps as rapidly as possible pass through the gated position and while maintaining the airspeed at ap- must have features to prevent inad- proximately 30 percent above the ref- vertent movement of the control erence stall speed existing at each in- through the gated position. It must stant throughout the maneuver. only be possible to make this separate (2) Repeat paragraph (b)(1) except ini- and distinct motion once the control tially extend the flaps and then retract has reached the gated position. them as rapidly as possible. (3) Repeat paragraph (b)(2), except at [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as 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; (4) With power off, flaps retracted, Amdt. 25–84, 60 FR 30749, June 9, 1995; Amdt. and the airplane trimmed at 1.3 VSR1, 25–98, 64 FR 6164, Feb. 8, 1999; 64 FR 10740, rapidly set go-around power or thrust Mar. 5, 1999; Amdt. 25–108, 67 FR 70827, Nov. while maintaining the same airspeed. 26, 2002]

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§ 25.147 Directional and lateral con- the speeds likely to be used with one trol. engine inoperative, to provide a roll (a) Directional control; general. It must rate necessary for safety without ex- be possible, with the wings level, to cessive control forces or travel. yaw into the operative engine and to (e) Lateral control; airplanes with four safely make a reasonably sudden or more engines. Airplanes with four or ° change in heading of up to 15 degrees in more engines must be able to make 20 the direction of the critical inoperative banked turns, with and against the inoperative engines, from steady flight at engine. This must be shown at 1.3 VSR1 for heading changes up to 15 degrees a speed equal to 1.3 VSR1, with max- (except that the heading change at imum continuous power, and with the which the rudder pedal force is 150 airplane in the configuration pre- pounds need not be exceeded), and scribed by paragraph (b) of this section. with— (f) Lateral control; all engines oper- (1) The critical engine inoperative ating. With the engines operating, roll and its propeller in the minimum drag response must allow normal maneuvers position; (such as recovery from upsets produced (2) The power required for level flight by gusts and the initiation of evasive maneuvers). There must be enough ex- at 1.3 VSR1, but not more than maximum continuous power; cess lateral control in sideslips (up to (3) The most unfavorable center of sideslip angles that might be required gravity; in normal operation), to allow a lim- (4) Landing gear retracted; ited amount of maneuvering and to (5) Flaps in the approach position; correct for gusts. Lateral control must and be enough at any speed up to VFC/MFC (6) Maximum landing weight. to provide a peak roll rate necessary (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 (3) The airplane trimmed for takeoff; the critical engine inoperative, roll re- (4) The maximum sea level takeoff sponse must allow normal maneuvers. weight (or any lesser weight necessary Lateral control must be sufficient, at to show VMC); 222

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(5) The airplane in the most critical (4) The airplane trimmed for takeoff; takeoff configuration existing along and the flight path after the airplane be- (5) The most unfavorable weight in comes airborne, except with the land- the range of takeoff weights.

ing gear retracted; (f) VMCL, the minimum control speed (6) The airplane airborne and the during approach and landing with all ground effect negligible; and engines operating, is the calibrated air- (7) If applicable, the propeller of the speed at which, when the critical en- inoperative engine— gine is suddenly made inoperative, it is (i) Windmilling; possible to maintain control of the air- (ii) In the most probable position for plane with that engine still inoper- the specific design of the propeller con- ative, and maintain straight flight trol; or with an angle of bank of not more than (iii) Feathered, if the airplane has an 5 degrees. VMCL must be established automatic feathering device acceptable with— for showing compliance with the climb (1) The airplane in the most critical requirements of § 25.121. configuration (or, at the option of the (d) The rudder forces required to applicant, each configuration) for ap- maintain control at VMC may not ex- proach and landing with all engines op- ceed 150 pounds nor may it be nec- erating; essary to reduce power or thrust of the (2) The most unfavorable center of operative engines. During recovery, the gravity; airplane may not assume any dan- (3) The airplane trimmed for ap- gerous attitude or require exceptional proach with all engines operating; piloting skill, alertness, or strength to (4) The most favorable weight, or, at prevent a heading change of more than the option of the applicant, as a func- 20 degrees. tion of weight; (e) VMCG, the minimum control speed 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 speed during approach and landing skill. In the determination of VMCG, assuming 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— any point. VMCG must be established (1) The airplane in the most critical 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 ap- (3) The most unfavorable center of proach with one critical engine inoper- gravity; ative;

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(4) The most unfavorable weight, or, ther the primary controls or their cor- at the option of the applicant, as a responding trim controls by the pilot function of weight; or the automatic pilot. (5) For propeller airplanes, the pro- (b) Lateral and directional trim. The peller of the more critical inoperative airplane must maintain lateral and di- engine in the position it achieves with- rectional trim with the most adverse out pilot action, assuming the engine lateral displacement of the center of fails while at the power or thrust nec- gravity within the relevant operating essary to maintain a three degree ap- limitations, during normally expected proach path angle, and the propeller of the other inoperative engine feathered; conditions of operation (including op- (6) The power or thrust on the oper- eration at any speed from 1.3 VSR1 to ating engine(s) necessary to maintain VMO/MMO). an approach path angle of three de- (c) Longitudinal trim. The airplane grees when one critical engine is inop- must maintain longitudinal trim dur- erative; and ing— (7) The power or thrust on the oper- (1) A climb with maximum contin- ating engine(s) rapidly changed, imme- uous power at a speed not more than diately after the second critical engine 1.3 VSR1, with the landing gear re- is made inoperative, from the power or tracted, and the flaps (i) retracted and thrust prescribed in paragraph (g)(6) of (ii) in the takeoff position; this section to— (2) Either a glide with power off at a (i) Minimum power or thrust; and speed not more than 1.3 V , or an ap- (ii) Go-around power or thrust set- SR1 proach within the normal range of ap- ting. proach speeds appropriate to the (h) In demonstrations of V and MCL weight and configuration with power VMCL-2— (1) The rudder force may not exceed settings corresponding to a 3 degree 150 pounds; glidepath, whichever is the most se- (2) The airplane may not exhibit haz- vere, with the landing gear extended, ardous flight characteristics or require the wing flaps (i) retracted and (ii) ex- exceptional piloting skill, alertness, or tended, and with the most unfavorable strength; combination of center of gravity posi- (3) Lateral control must be sufficient tion and weight approved for landing; to roll the airplane, from an initial and condition of steady flight, through an (3) Level flight at any speed from 1.3 angle of 20 degrees in the direction nec- VSR1, to VMO/MMO, with the landing gear essary to initiate a turn away from the and flaps retracted, and from 1.3 VSR1 to inoperative engine(s), in not more than V with the landing gear extended. 5 seconds; and LE (d) Longitudinal, directional, and lat- (4) For propeller airplanes, hazardous eral trim. The airplane must maintain flight characteristics must not be ex- hibited due to any propeller position longitudinal, directional, and lateral achieved when the engine fails or dur- trim (and for the lateral trim, the ing any likely subsequent movements angle of bank may not exceed five de- of the engine or propeller controls. grees) at 1.3 VSR1 during climbing flight with— [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (1) The critical engine inoperative; amended by Amdt. 25–42, 43 FR 2321, Jan. 16, 1978; Amdt. 25–72, 55 FR 29774, July 20, 1990; 55 (2) The remaining engines at max- FR 37607, Sept. 12, 1990; Amdt. 25–84, 60 FR imum continuous power; and 30749, June 9, 1995; Amdt. 25–108, 67 FR 70827, (3) The landing gear and flaps re- Nov. 26, 2002] tracted. TRIM (e) Airplanes with four or more engines. Each airplane with four or more § 25.161 Trim. engines must also maintain trim in (a) General. Each airplane must meet rectilinear flight with the most unfa- the trim requirements of this section vorable center of gravity and at the after being trimmed, and without fur- climb speed, configuration, and power ther pressure upon, or movement of, ei- required by § 25.123(a) for the purpose of

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establishing the en route flight paths speeds above or below the desired trim with two engines inoperative. speeds if exceptional attention on the part of the pilot is not required to re- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5671, Apr. 8, turn to and maintain the desired trim 1970; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; speed and altitude. Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; [Amdt. 25–7, 30 FR 13117, Oct. 15, 1965] Amdt. 25–115, 69 FR 40527, July 2, 2004]

STABILITY § 25.175 Demonstration of static longitudinal stability. § 25.171 General. Static longitudinal stability must be The airplane must be longitudinally, shown as follows: directionally, and laterally stable in (a) Climb. The stick force curve must accordance with the provisions of have a stable slope at speeds between §§ 25.173 through 25.177. In addition, 85 and 115 percent of the speed at which suitable stability and control feel the airplane— (static stability) is required in any con- (1) Is trimmed, with— dition normally encountered in service, (i) Wing flaps retracted; if flight tests show it is necessary for (ii) Landing gear retracted; safe operation. (iii) Maximum takeoff weight; and (iv) 75 percent of maximum contin- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–7, 30 FR 13117, Oct. 15, uous power for reciprocating engines or 1965] the maximum power or thrust selected by the applicant as an operating limi- § 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

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

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stable sense. This factor of proportion- STALLS ality must lie between limits found necessary for safe operation. The range § 25.201 Stall demonstration. of sideslip angles evaluated must in- (a) Stalls must be shown in straight clude those sideslip angles resulting flight and in 30 degree banked turns from the lesser of: with— (1) One-half of the available rudder (1) Power off; and control input; and (2) The power necessary to maintain (2) A rudder control force of 180 level flight at 1.5 VSR1 (where VSR1 cor- pounds. responds to the reference stall speed at (d) For sideslip angles greater than maximum landing weight with flaps in those prescribed by paragraph (c) of the approach position and the landing this section, up to the angle at which gear retracted). full rudder control is used or a rudder (b) In each condition required by control force of 180 pounds is obtained, paragraph (a) of this section, it must the rudder control forces may not re- be possible to meet the applicable reverse, and increased rudder deflection quirements of § 25.203 with— must be needed for increased angles of (1) Flaps, landing gear, and decelera- sideslip. Compliance with this require- tion devices in any likely combination ment must be shown using straight, of positions approved for operation; 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 (1) Starting at a speed sufficiently position with all engines operating. 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 does not exceed one knot per second (a) Any short period oscillation, not until the airplane is stalled. including combined lateral-directional (2) In addition, for turning flight oscillations, occurring between 1.13 VSR stalls, apply the longitudinal control and maximum allowable speed appro- to achieve airspeed deceleration rates priate to the configuration of the air- up to 3 knots per second. plane must be heavily damped with the (3) As soon as the airplane is stalled, primary controls— recover by normal recovery techniques. (1) Free; and (d) The airplane is considered stalled (2) In a fixed position. when the behavior of the airplane gives (b) Any combined lateral-directional the pilot a clear and distinctive indica- oscillations (‘‘Dutch roll’’) occurring tion of an acceptable nature that the between 1.13 VSR and maximum allow- airplane is stalled. Acceptable indica- able speed appropriate to the configu- tions of a stall, occurring either indi- ration of the airplane must be posi- vidually or in combination, are— tively damped with controls free, and (1) A nose-down pitch that cannot be must be controllable with normal use readily arrested; of the primary controls without requir- (2) Buffeting, of a magnitude and se- ing exceptional pilot skill. verity that is a strong and effective de- [Amdt. 25–42, 43 FR 2322, Jan. 16, 1978, as terrent to further speed reduction; or amended by Amdt. 25–72, 55 FR 29775, July 20, (3) The pitch control reaches the aft 1990; 55 FR 37607, Sept. 12, 1990; Amdt. 25–108, stop and no further increase in pitch 67 FR 70827, Nov. 26, 2002] attitude occurs when the control is

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held full aft for a short time before re- acceptable by itself. If a warning de- covery is initiated. vice is used, it must provide a warning [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as in each of the airplane configurations amended by Amdt. 25–84, 60 FR 30750, June 9, prescribed in paragraph (a) of this sec- 1995; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002] tion at the speed prescribed in paragraphs (c) and (d) of this section. Ex- § 25.203 Stall characteristics. cept for the stall warning prescribed in (a) It must be possible to produce and paragraph (h)(3)(ii) of this section, the to correct roll and yaw by unreversed stall warning for flight in icing condi- use of the aileron and rudder controls, tions must be provided by the same up to the time the airplane is stalled. means as the stall warning for flight in No abnormal nose-up pitching may non-icing conditions. occur. The longitudinal control force (c) When the speed is reduced at rates must be positive up to and throughout not exceeding one knot per second, the stall. In addition, it must be pos- stall warning must begin, in each nor- sible to promptly prevent stalling and mal configuration, at a speed, VSW, ex- to recover from a stall by normal use ceeding the speed at which the stall is of the controls. identified in accordance with § 25.201(d) (b) For level wing stalls, the roll oc- by not less than five knots or five per- curring between the stall and the com- cent CAS, whichever is greater. Once pletion of the recovery may not exceed initiated, stall warning must continue approximately 20 degrees. until the angle of attack is reduced to (c) For turning flight stalls, the ac- approximately that at which stall tion of the airplane after the stall may warning began. not be so violent or extreme as to (d) In addition to the requirement of make it difficult, with normal piloting paragraph (c) of this section, when the skill, to effect a prompt recovery and speed is reduced at rates not exceeding to regain control of the airplane. The one knot per second, in straight flight maximum bank angle that occurs dur- with engines idling and at the center- ing the recovery may not exceed— of-gravity position specified in (1) Approximately 60 degrees in the 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 recovery maneuver not less than three [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as seconds after the onset of stall warn- amended by Amdt. 25–84, 60 FR 30750, June 9, 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 most 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 Appendices C and O of this that will give clearly distinguishable part, as applicable, in accordance with indications under expected conditions § 25.21(g), for each configuration used in of flight. However, a visual stall warn- the takeoff phase of flight; ing device that requires the attention (2) The most critical of the en route of the crew within the cockpit is not ice accretion(s) defined in Appendices C

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and O of this part, as applicable, in ac- gin in straight and turning flight must cordance with § 25.21(g), for the en be sufficient to allow the pilot to pre- route configuration; vent stalling without encountering any (3) The most critical of the holding adverse flight characteristics when: ice accretion(s) defined in Appendices C (1) The speed is reduced at rates not and O of this part, as applicable, in ac- exceeding one knot per second; cordance with § 25.21(g), for the holding (2) The pilot performs the recovery configuration(s); maneuver in the same way as for flight (4) The most critical of the approach in non-icing conditions; and ice accretion(s) defined in Appendices C (3) The recovery maneuver is started and O of this part, as applicable, in ac- no earlier than: cordance with § 25.21(g), for the ap- (i) One second after the onset of stall proach configuration(s); and warning if stall warning is provided by (5) The most critical of the landing the same means as for flight in non- ice accretion(s) defined in Appendices C icing conditions; or and O of this part, as applicable, in ac- (ii) Three seconds after the onset of cordance with § 25.21(g), for the landing stall warning if stall warning is pro- and go-around configuration(s). vided by a different means than for (f) The stall warning margin must be flight in non-icing conditions. sufficient in both non-icing and icing (i) In showing compliance with para- conditions to allow the pilot to prevent graph (h) of this section, if stall warn- stalling when the pilot starts a recov- ing is provided by a different means in ery maneuver not less than one second icing conditions than for non-icing con- after the onset of stall warning in slow- ditions, compliance with § 25.203 must down turns with at least 1.5 g load fac- be shown using the accretion defined in tor normal to the flight path and air- appendix C, part II(e) of this part. Com- speed deceleration rates of at least 2 pliance with this requirement must be knots per second. When demonstrating shown using the demonstration pre- compliance with this paragraph for scribed by § 25.201, except that the deicing conditions, the pilot must per- celeration rates of § 25.201(c)(2) need not form the recovery maneuver in the be demonstrated. same way as for the airplane in non- icing conditions. Compliance with this [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as requirement must be demonstrated in amended by Amdt. 25–7, 30 FR 13118, Oct. 15, 1965; Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; flight with— Amdt. 25–108, 67 FR 70827, Nov. 26, 2002; (1) The flaps and landing gear in any Amdt. 25–121, 72 FR 44668, Aug. 8, 2007; Amdt. normal position; 25–129, 74 FR 38339, Aug. 3, 2009; Amdt. 25–140, (2) The airplane trimmed for straight 79 FR 65526, Nov. 4, 2014] flight at a speed of 1.3 VSR; and (3) The power or thrust necessary to GROUND AND WATER HANDLING maintain level flight at 1.3 VSR. CHARACTERISTICS (g) Stall warning must also be provided in each abnormal configuration § 25.231 Longitudinal stability and of the high lift devices that is likely to control. be used in flight following system fail- (a) Landplanes may have no uncon- ures (including all configurations cov- trollable tendency to nose over in any ered by Airplane Flight Manual proce- reasonably expected operating condi- dures). tion or when rebound occurs during (h) The following stall warning mar- landing or takeoff. In addition— gin is required for flight in icing condi- (1) Wheel brakes must operate tions before the ice protection system smoothly and may not cause any undue has been activated and is performing tendency to nose over; and its intended function. Compliance must (2) If a tail-wheel landing gear is be shown using the most critical of the used, it must be possible, during the ice accretion(s) defined in Appendix C, takeoff ground run on concrete, to part II, paragraph (e) of this part and maintain any attitude up to thrust line Appendix O, part II, paragraph (d) of level, at 75 percent of VSR1. this part, as applicable, in accordance (b) For seaplanes and amphibians, with § 25.21(g). The stall warning mar- the most adverse water conditions safe

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for takeoff, taxiing, and landing, must (3) The landing crosswind component be established. must be established for: (i) Non-icing conditions, and [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–108, 67 FR 70828, Nov. (ii) Icing conditions with the most 26, 2002] critical of the landing ice accretion(s) defined in Appendices C and O of this § 25.233 Directional stability and con- part, as applicable, in accordance with trol. § 25.21(g). (a) There may be no uncontrollable (b) For seaplanes and amphibians, ground-looping tendency in 90° cross the following applies: winds, up to a wind velocity of 20 knots (1) A 90-degree cross component of or 0.2 VSR0, whichever is greater, except wind velocity, up to which takeoff and that the wind velocity need not exceed landing is safe under all water condi- 25 knots at any speed at which the air- tions that may reasonably be expected plane may be expected to be operated in normal operation, must be estab- on the ground. This may be shown lished and must be at least 20 knots or while establishing the 90° cross compo- 0.2 VSR0, whichever is greater, except nent of wind velocity required by that it need not exceed 25 knots. § 25.237. (2) A wind velocity, for which taxiing (b) Landplanes must be satisfactorily is safe in any direction under all water controllable, without exceptional pilot- conditions that may reasonably be ex- ing skill or alertness, in power-off land- pected in normal operation, must be es- ings at normal landing speed, without tablished and must be at least 20 knots using brakes or engine power to main- or 0.2 VSR0, whichever is greater, except tain a straight path. This may be that it need not exceed 25 knots. shown during power-off landings made [Amdt. 25–42, 43 FR 2322, Jan. 16, 1978, as in conjunction with other tests. amended by Amdt. 25–108, 67 FR 70827, Nov. (c) The airplane must have adequate 26, 2002; Amdt. 25–121, 72 FR 44668, Aug. 8, directional control during taxiing. This 2007; Amdt. 25–140, 79 FR 65525, Nov. 4, 2014] may be shown during taxiing prior to takeoffs made in conjunction with § 25.239 Spray characteristics, control, other tests. and stability on water. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) For seaplanes and amphibians, amended by Amdt. 25–23, 35 FR 5671, Apr. 8, during takeoff, taxiing, and landing, 1970; Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; and in the conditions set forth in para- Amdt. 25–94, 63 FR 8848, Feb. 23, 1998; Amdt. graph (b) of this section, there may be 25–108, 67 FR 70828, Nov. 26, 2002] no— (1) Spray characteristics that would § 25.235 Taxiing condition. impair the pilot’s view, cause damage, The shock absorbing mechanism may or result in the taking in of an undue not damage the structure of the air- quantity of water; plane when the airplane is taxied on (2) Dangerously uncontrollable the roughest ground that may reason- porpoising, bounding, or swinging tend- ably be expected in normal operation. ency; or (3) Immersion of auxiliary floats or § 25.237 Wind velocities. sponsons, wing tips, propeller blades, (a) For land planes and amphibians, or other parts not designed to with- the following applies: stand the resulting water loads. (1) A 90-degree cross component of (b) Compliance with the require- wind velocity, demonstrated to be safe ments of paragraph (a) of this section for takeoff and landing, must be estab- must be shown— lished for dry runways and must be at (1) In water conditions, from smooth least 20 knots or 0.2 VSR0, whichever is to the most adverse condition estab- greater, except that it need not exceed lished in accordance with § 25.231; 25 knots. (2) In wind and cross-wind velocities, (2) The crosswind component for water currents, and associated waves takeoff established without ice accre- and swells that may reasonably be ex- tions is valid in icing conditions. pected in operation on water;

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(3) At speeds that may reasonably be cient range of speeds and load factors expected in operation on water; for normal operations. Probable inad- (4) With sudden failure of the critical vertent excursions beyond the bound- engine at any time while on water; and aries of the buffet onset envelopes may (5) At each weight and center of grav- not result in unsafe conditions. ity position, relevant to each operating [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as condition, within the range of loading amended by Amdt. 25–23, 35 FR 5671, Apr. 8, conditions for which certification is re- 1970; Amdt. 25–72, 55 FR 29775, July 20, 1990; quested. Amdt. 25–77, 57 FR 28949, June 29, 1992] (c) In the water conditions of paragraph (b) of this section, and in the § 25.253 High-speed characteristics. corresponding wind conditions, the sea- (a) Speed increase and recovery charac- plane or amphibian must be able to teristics. The following speed increase drift for five minutes with engines in- and recovery characteristics must be operative, aided, if necessary, by a sea met: anchor. (1) Operating conditions and characteristics likely to cause inadvertent MISCELLANEOUS FLIGHT REQUIREMENTS speed increases (including upsets in pitch and roll) must be simulated with § 25.251 Vibration and buffeting. the airplane trimmed at any likely (a) The airplane must be dem- cruise speed up to VMO/MMO. These con- onstrated in flight to be free from any ditions and characteristics include gust vibration and buffeting that would pre- upsets, inadvertent control move- vent continued safe flight in any likely ments, low stick force gradient in rela- operating condition. tion to control friction, passenger (b) Each part of the airplane must be movement, leveling off from climb, and demonstrated in flight to be free from descent from Mach to airspeed limit al- excessive vibration under any appro- titudes. priate speed and power conditions up to (2) Allowing for pilot reaction time VDF/MDF. The maximum speeds shown after effective inherent or artificial must be used in establishing the oper- speed warning occurs, it must be shown ating limitations of the airplane in ac- that the airplane can be recovered to a cordance with § 25.1505. normal attitude and its speed reduced (c) Except as provided in paragraph to VMO/MMO, without— (d) of this section, there may be no buf- (i) Exceptional piloting strength or feting condition, in normal flight, in- skill; cluding configuration changes during (ii) Exceeding VD/MD, VDF/MDF, or the cruise, severe enough to interfere with structural limitations; and the control of the airplane, to cause ex- (iii) Buffeting that would impair the cessive fatigue to the crew, or to cause pilot’s ability to read the instruments structural damage. Stall warning buf- or control the airplane for recovery. feting within these limits is allowable. (3) With the airplane trimmed at any (d) There may be no perceptible buf- speed up to VMO/MMO, there must be no feting condition in the cruise configu- reversal of the response to control ration in straight flight at any speed input about any axis at any speed up to up to VMO/MMO, except that stall warn- VDF/MDF. Any tendency to pitch, roll, or ing buffeting is allowable. yaw must be mild and readily control- (e) For an airplane with MD greater lable, using normal piloting tech- than .6 or with a maximum operating niques. When the airplane is trimmed altitude greater than 25,000 feet, the at VMO/MMO, the slope of the elevator positive maneuvering load factors at control force versus speed curve need which the onset of perceptible buf- not be stable at speeds greater than feting occurs must be determined with VFC/MFC, but there must be a push force the airplane in the cruise configuration at all speeds up to VDF/MDF and there for the ranges of airspeed or Mach must be no sudden or excessive reduc- number, weight, and altitude for which tion of elevator control force as VDF/ the airplane is to be certificated. The MDF is reached. envelopes of load factor, speed, alti- (4) Adequate roll capability to assure tude, and weight must provide a suffi- a prompt recovery from a lateral upset

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condition must be available at any controllability with the degree of out- speed up to VDF/MDF. of-trim in both the airplane nose-up (5) With the airplane trimmed at and nose-down directions, which re- VMO/MMO, extension of the speedbrakes sults from the greater of— over the available range of movements (1) A three-second movement of the of the pilot’s control, at all speeds longitudinal trim system at its normal above VMO/MMO, but not so high that rate for the particular flight condition VDF/MDF would be exceeded during the with no aerodynamic load (or an equiv- maneuver, must not result in: alent degree of trim for airplanes that (i) An excessive positive load factor do not have a power-operated trim sys- when the pilot does not take action to tem), except as limited by stops in the counteract the effects of extension; trim system, including those required (ii) Buffeting that would impair the by § 25.655(b) for adjustable stabilizers; pilot’s ability to read the instruments or or control the airplane for recovery; or (2) The maximum mistrim that can (iii) A nose down pitching moment, be sustained by the autopilot while unless it is small. maintaining level flight in the high (b) Maximum speed for stability charac- speed cruising condition. teristics, VFC/MFC. VFC/MFC is the max- (b) In the out-of-trim condition speci- imum speed at which the requirements fied in paragraph (a) of this section, of §§ 25.143(g), 25.147(f), 25.175(b)(1), when the normal acceleration is varied 25.177(a) through (c), and 25.181 must be from + 1 g to the positive and negative met with flaps and landing gear re- values specified in paragraph (c) of this tracted. Except as noted in § 25.253(c), section— VFC/MFC may not be less than a speed (1) The stick force vs. g curve must midway between VMO/MMO and VDF/MDF, have a positive slope at any speed up to except that, for altitudes where Mach and including VFC/MFC; and number is the limiting factor, MFC need (2) At speeds between V /M and not exceed the Mach number at which FC FC VDF/MDF the direction of the primary effective speed warning occurs. longitudinal control force may not re- (c) Maximum speed for stability charac- verse. teristics in icing conditions. The max- (c) Except as provided in paragraphs imum speed for stability characteris- (d) and (e) of this section, compliance tics with the most critical of the ice with the provisions of paragraph (a) of accretions defined in Appendices C and this section must be demonstrated in O of this part, as applicable, in accord- flight over the acceleration range— ance with § 25.21(g), at which the re- (1) ¥1 g to + 2.5 g; or quirements of §§ 25.143(g), 25.147(f), (2) 0 g to 2.0 g, and extrapolating by 25.175(b)(1), 25.177(a) through (c), and an acceptable method to ¥1 g and + 2.5 25.181 must be met, is the lower of: g. (1) 300 knots CAS; (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. 25–84, 60 FR 30750, June 9, 1995; Amdt. 25–121, specified in paragraph (b)(1) of this sec- 72 FR 44668, Aug. 8, 2007; Amdt. 25–135, 76 FR tion. 74654, Dec. 1, 2011; Amdt. 25–140,79 FR 65525, (e) During flight tests required by Nov. 4, 2014] paragraph (a) of this section, the limit maneuvering load factors prescribed in § 25.255 Out-of-trim characteristics. §§ 25.333(b) and 25.337, and the maneu- (a) From an initial condition with vering load factors associated with the airplane trimmed at cruise speeds probable inadvertent excursions be- up to VMO/MMO, the airplane must have yond the boundaries of the buffet onset satisfactory maneuvering stability and envelopes determined under § 25.251(e),

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need not be exceeded. In addition, the termining those loading conditions are entry speeds for flight test demonstra- shown to be reliable. tions at normal acceleration values (c) If deflections under load would less than 1 g must be limited to the ex- significantly change the distribution of tent necessary to accomplish a recov- external or internal loads, this redis- ery without exceeding VDF/MDF. tribution must be taken into account. (f) In the out-of-trim condition specified in paragraph (a) of this section, it [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5672, Apr. 8, must be possible from an overspeed 1970] condition at VDF/MDF to produce at least 1.5 g for recovery by applying not § 25.303 Factor of safety. more than 125 pounds of longitudinal Unless otherwise specified, a factor of control force using either the primary safety of 1.5 must be applied to the pre- longitudinal control alone or the pri- scribed limit load which are considered mary longitudinal control and the longitudinal trim system. If the longitu- external loads on the structure. When a dinal trim is used to assist in pro- loading condition is prescribed in ducing the required load factor, it must terms of ultimate loads, a factor of safety need not be applied unless other- be shown at VDF/MDF that the longitudinal trim can be actuated in the air- wise specified. plane nose-up direction with the pri- [Amdt. 25–23, 35 FR 5672, Apr. 8, 1970] mary surface loaded to correspond to the least of the following airplane § 25.305 Strength and deformation. nose-up control forces: (a) The structure must be able to (1) The maximum control forces ex- support limit loads without detri- pected in service as specified in §§ 25.301 mental permanent deformation. At any and 25.397. load up to limit loads, the deformation (2) The control force required to may not interfere with safe operation. produce 1.5 g. (b) The structure must be able to (3) The control force corresponding to support ultimate loads without failure buffeting or other phenomena of such for at least 3 seconds. However, when 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 Subpart C—Structure deflections and ultimate deformation 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

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that might occur in any likely oper- weight of the airplane. A positive load ating condition up to VD/MD, including factor is one in which the aerodynamic stall and probable inadvertent excur- force acts upward with respect to the sions beyond the boundaries of the buf- airplane. fet onset envelope. This must be shown (b) Considering compressibility ef- by analysis, flight tests, or other tests fects at each speed, compliance with found necessary by the Administrator. the flight load requirements of this (f) Unless shown to be extremely im- subpart must be shown— probable, the airplane must be designed (1) At each critical altitude within to withstand any forced structural vi- the range of altitudes selected by the bration resulting from any failure, applicant; malfunction or adverse condition in (2) At each weight from the design the flight control system. These must minimum weight to the design max- be considered limit loads and must be imum weight appropriate to each par- investigated at airspeeds up to VC/MC. ticular flight load condition; and (3) For each required altitude and [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5672, Apr. 8, weight, for any practicable distribution 1970; Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; of disposable load within the operating Amdt. 25–77, 57 FR 28949, June 29, 1992; Amdt. limitations recorded in the Airplane 25–86, 61 FR 5220, Feb. 9, 1996] Flight Manual. (c) Enough points on and within the § 25.307 Proof of structure. boundaries of the design envelope must (a) Compliance with the strength and be investigated to ensure that the max- deformation requirements of this sub- imum load for each part of the airplane part must be shown for each critical structure is obtained. loading condition. Structural analysis (d) The significant forces acting on may be used only if the structure con- the airplane must be placed in equi- forms to that for which experience has librium in a rational or conservative shown this method to be reliable. In manner. The linear inertia forces must other cases, substantiating tests must be considered in equilibrium with the be made to load levels that are suffi- thrust and all aerodynamic loads, cient to verify structural behavior up while the angular (pitching) inertia to loads specified in § 25.305. forces must be considered in equi- (b)–(c) [Reserved] librium with thrust and all aero- (d) When static or dynamic tests are dynamic moments, including moments used to show compliance with the re- due to loads on components such as quirements of § 25.305(b) for flight tail surfaces and nacelles. Critical structures, appropriate material cor- thrust values in the range from zero to rection factors must be applied to the maximum continuous thrust must be test results, unless the structure, or considered. 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 element results in the redistribution of FLIGHT MANEUVER AND GUST the load through alternate load paths. CONDITIONS [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; 79 FR (a) Procedure. For the analysis of the 59429, Oct. 2, 2014] maneuvering flight conditions specified in paragraphs (b) and (c) of this sec- FLIGHT LOADS tion, the following provisions apply: (1) Where sudden displacement of a § 25.321 General. control is specified, the assumed rate (a) Flight load factors represent the of control surface displacement may ratio of the aerodynamic force compo- not be less than the rate that could be nent (acting normal to the assumed applied by the pilot through the con- longitudinal axis of the airplane) to the trol system.

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(2) In determining elevator angles positive limit load factor prescribed in and chordwise load distribution in the § 25.337 is achieved. As a separate condi- maneuvering conditions of paragraphs tion, nose-down checked pitching ma- (b) and (c) of this section, the effect of neuvers must be analyzed in which a corresponding pitching velocities must limit load factor of 0g is achieved. In be taken into account. The in-trim and defining the airplane loads, the flight out-of-trim flight conditions specified deck pitch control motions described in § 25.255 must be considered. in paragraphs (c)(2)(i) through (iv) of (b) Maneuvering balanced conditions. this section must be used: Assuming the airplane to be in equi- (i) The airplane is assumed to be fly- librium with zero pitching accelera- ing in steady level flight at any speed tion, the maneuvering conditions A between VA and VD and the flight deck through I on the maneuvering envelope pitch control is moved in accordance in § 25.333(b) must be investigated. with the following formula: (c) Maneuvering pitching conditions. d(t) = d sin(wt) for 0 ≤ t ≤ t The following conditions must be in- 1 max vestigated: Where—

(1) Maximum pitch control displacement d1 = the maximum available displacement of at VA. The airplane is assumed to be the flight deck pitch control in the ini- flying in steady level flight (point A1, tial direction, as limited by the control § 25.333(b)) and the cockpit pitch con- system stops, control surface stops, or by trol is suddenly moved to obtain ex- pilot effort in accordance with § 25.397(b); treme nose up pitching acceleration. In d(t) = the displacement of the flight deck pitch control as a function of time. In defining the tail load, the response of the initial direction, d(t) is limited to d1. the airplane must be taken into ac- In the reverse direction, d(t) may be count. Airplane loads that occur subse- truncated at the maximum available dis- quent to the time when normal accel- placement of the flight deck pitch con- eration at the c.g. exceeds the positive trol as limited by the control system limit maneuvering load factor (at point stops, control surface stops, or by pilot A in § 25.333(b)), or the resulting effort in accordance with 25.397(b); 2 π tailplane normal load reaches its max- tmax = 3 /2w; imum, whichever occurs first, need not w = the circular frequency (radians/second) of the control deflection taken equal to the be considered. undamped natural frequency of the short (2) Checked maneuver between VA and period rigid mode of the airplane, with VD. Nose-up checked pitching maneu- active control system effects included vers must be analyzed in which the where appropriate; but not less than:

Where necessary to ensure that the normal V = the speed of the airplane at entry to the acceleration at the center of gravity maneuver. does not go below 0g.

VA = the design maneuvering speed pre- (iii) In addition, for cases where the scribed in § 25.335(c). airplane response to the specified flight deck pitch control motion does not (ii) For nose-up pitching maneuvers, achieve the prescribed limit load fac- the complete flight deck pitch control tors, then the following flight deck displacement history may be scaled pitch control motion must be used: down in amplitude to the extent nec- ≤ ≤ essary to ensure that the positive limit d(t) = d1 sin(wt) for 0 t t1 ≤ ≤ load factor prescribed in § 25.337 is not d(t) = d1 for t1 t t2 (t) = sin( [t + t ¥ t ]) for t ≤ t ≤ exceeded. For nose-down pitching ma- d d1 w 1 2 2 t neuvers, the complete flight deck con- max trol displacement history may be Where—

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t2 = t1 + Dt to be achieved in the initial direction, tmax = t2 + π/w; but it need not exceed five seconds (see Dt = the minimum period of time necessary figure below). to allow the prescribed limit load factor

(iv) In cases where the flight deck (C) tmax.. pitch control motion may be affected [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as by inputs from systems (for example, amended by Amdt. 25–23, 35 FR 5672, Apr. 8, by a stick pusher that can operate at 1970; Amdt. 25–46, 43 FR 50594, Oct. 30, 1978; 43 high load factor as well as at 1g), then FR 52495, Nov. 13, 1978; 43 FR 54082, Nov. 20, the effects of those systems shall be 1978; Amdt. 25–72, 55 FR 29775, July 20, 1990; 55 taken into account. FR 37607, Sept. 12, 1990; Amdt. 25–86, 61 FR 5220, Feb. 9, 1996; Amdt. 25–91, 62 FR 40704, (v) Airplane loads that occur beyond July 29, 1997; Amdt. 25–141, 79 FR 73466, Dec. the following times need not be consid- 11, 2014] ered: (A) For the nose-up pitching maneu- § 25.333 Flight maneuvering envelope. ver, the time at which the normal ac- (a) General. The strength require- celeration at the center of gravity goes ments must be met at each combina- below 0g; tion of airspeed and load factor on and (B) For the nose-down pitching ma- within the boundaries of the represent- neuver, the time at which the normal ative maneuvering envelope (V-n dia- acceleration at the center of gravity gram) of paragraph (b) of this section. goes above the positive limit load fac- This envelope must also be used in de- tor prescribed in § 25.337; termining the airplane structural operating limitations as specified in § 25.1501. (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 vice positioning or load limiting means airspeed); 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 airplane speeds between VB and speeds up to VC; and VC: Positive and negative gusts with (2) Must vary linearly with speed reference gust velocities of 56.0 ft/sec from the value at VC to zero at VD. EAS must be considered at sea level. (d) Maneuvering load factors lower The reference gust velocity may be re- than those specified in this section duced linearly from 56.0 ft/sec EAS at may be used if the airplane has design sea level to 44.0 ft/sec EAS at 15,000 features that make it impossible to ex- feet. The reference gust velocity may ceed these values in flight. be further reduced linearly from 44.0 ft/ [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as sec EAS at 15,000 feet to 20.86 ft/sec amended by Amdt. 25–23, 35 FR 5672, Apr. 8, EAS at 60,000 feet. 1970] (ii) At the airplane design speed VD: The reference gust velocity must be 0.5 § 25.341 Gust and turbulence loads. times the value obtained under (a) Discrete Gust Design Criteria. The § 25.341(a)(5)(i). airplane is assumed to be subjected to (6) The flight profile alleviation fac- symmetrical vertical and lateral gusts tor, Fg, must be increased linearly from in level flight. Limit gust loads must the sea level value to a value of 1.0 at be determined in accordance with the the maximum operating altitude de- provisions: fined in § 25.1527. At sea level, the flight (1) Loads on each part of the struc- profile alleviation factor is determined ture must be determined by dynamic by the following equation: analysis. The analysis must take into account unsteady aerodynamic charac- =+ teristics and all significant structural FFFggzgm05. () degrees of freedom including rigid body motions. Where: (2) The shape of the gust must be: Z =− mo U ⎡ ⎛ πs⎞ ⎤ Fgz 1 ; U = ds ⎢1- Cos⎜ ⎟ ⎥ 250000 2 ⎣ ⎝ H ⎠ ⎦ ⎛πR ⎞ for 0 ≤s ≤2H F= R Tan 1 ; where— gm 2 ⎝ 4 ⎠ s = distance penetrated into the gust (feet); Uds = the design gust velocity in equivalent Maximum Landing Weight airspeed specified in paragraph (a)(4) of R = ; this section; and 1 H = the gust gradient which is the distance Maximum Take- off Weight (feet) parallel to the airplane’s flight path for the gust to reach its peak veloc- = Maximum Zero Fuel Weight ity. R2 ; Maximum Take- off Weight (3) A sufficient number of gust gradient distances in the range 30 feet to Zmo = Maximum operating altitude defined in 350 feet must be investigated to find § 25.1527 (feet). the critical response for each load (7) When a stability augmentation quantity. system is included in the analysis, the (4) The design gust velocity must be: effect of any significant system non- linearities should be accounted for 16 UUF= ()H when deriving limit loads from limit ds ref g 350 gust conditions. where— (b) Continuous turbulence design cri-

Uref = the reference gust velocity in equiva- teria. The dynamic response of the air- lent airspeed defined in paragraph (a)(5) plane to vertical and lateral contin- of this section. uous turbulence must be taken into ac- Fg = the flight profile alleviation factor de- count. The dynamic analysis must take fined in paragraph (a)(6) of this section. into account unsteady aerodynamic (5) The following reference gust ve- characteristics and all significant locities apply: structural degrees of freedom including

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rigid body motions. The limit loads Where—

must be determined for all critical alti- PL = limit load; tudes, weights, and weight distribu- PL¥1g = steady 1g load for the condition; tions as specified in § 25.321(b), and all A = ratio of root-mean-square incremental load for the condition to root-mean- critical speeds within the ranges indi- square turbulence velocity; and cated in § 25.341(b)(3). Uσ = limit turbulence intensity in true air- (1) Except as provided in paragraphs speed, specified in paragraph (b)(3) of this (b)(4) and (5) of this section, the fol- section. lowing equation must be used: (2) Values of A must be determined ¯ PL = PL¥1g ± UσA according to the following formula:

Where— lates the loads in the aircraft structure H(W) = the frequency response function, de- to the atmospheric turbulence; and termined by dynamic analysis, that re- F(W) = normalized power spectral density of atmospheric turbulence given by—

Where— (iv) At all speeds, both positive and W = reduced frequency, radians per foot; and negative incremental loads due to con- L = scale of turbulence = 2,500 ft. tinuous turbulence must be considered. (4) When an automatic system affect- (3) The limit turbulence intensities, ing the dynamic response of the air- Uσ, in feet per second true airspeed re- plane is included in the analysis, the quired for compliance with this para- effects of system non-linearities on graph are— loads at the limit load level must be (i) At airplane speeds between VB and taken into account in a realistic or VC: conservative manner. Uσ = Uσref Fg (5) If necessary for the assessment of Where— loads on airplanes with significant non- linearities, it must be assumed that Uσref is the reference turbulence intensity that varies linearly with altitude from 90 the turbulence field has a root-mean- fps (TAS) at sea level to 79 fps (TAS) at square velocity equal to 40 percent of 24,000 feet and is then constant at 79 fps the Uσ values specified in paragraph (TAS) up to the altitude of 60,000 feet. (b)(3) of this section. The value of limit

Fg is the flight profile alleviation factor de- load is that load with the same prob- fined in paragraph (a)(6) of this section; ability of exceedance in the turbulence field as AUσ of the same load quantity (ii) At speed V : Uσ is equal to 1⁄2 the D in a linear approximated model. values obtained under paragraph (c) Supplementary gust conditions for (b)(3)(i) of this section. wing-mounted engines. For airplanes

(iii) At speeds between VC and VD: Uσ equipped with wing-mounted engines, is equal to a value obtained by linear the engine mounts, pylons, and wing interpolation. supporting structure must be designed

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for the maximum response at the na- ently tuned to the maximum response celle center of gravity derived from the in accordance with § 25.341(a). The pene- following dynamic gust conditions ap- tration of the airplane in the combined plied to the airplane: gust field and the phasing of the (1) A discrete gust determined in ac- vertical and lateral component gusts cordance with § 25.341(a) at each angle must be established to develop the normal to the flight path, and sepa- maximum response to the gust pair. In rately, the absence of a more rational anal- (2) A pair of discrete gusts, one ysis, the following formula must be vertical and one lateral. The length of used for each of the maximum engine each of these gusts must be independ- loads in all six degrees of freedom:

Where— design condition of paragraph (b)(1) of

PL = limit load; this section; and PL-1g = steady 1g load for the condition; (3) The flutter, deformation, and vi- LV = peak incremental response load due to bration requirements must also be met a vertical gust according to § 25.341(a); with zero fuel. and LL = peak incremental response load due to [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as a lateral gust according to § 25.341(a). amended by Amdt. 25–18, 33 FR 12226, Aug. 30, 1968; Amdt. 25–72, 55 FR 37607, Sept. 12, 1990; [Doc. No. 27902, 61 FR 5221, Feb. 9, 1996; 61 FR Amdt. 25–86, 61 FR 5221, Feb. 9, 1996; Amdt. 9533, Mar. 8, 1996; Doc. No. FAA–2013–0142; 79 25–141, 79 FR 73468, Dec. 11, 2014] FR 73467, Dec. 11, 2014; Amdt. 25–141, 80 FR 4762, Jan. 29, 2015; 80 FR 6435, Feb. 5, 2015] § 25.345 High lift devices. § 25.343 Design fuel and oil loads. (a) If wing flaps are to be used during (a) The disposable load combinations takeoff, approach, or landing, at the must include each fuel and oil load in design flap speeds established for these stages of flight under § 25.335(e) and the range from zero fuel and oil to the with the wing flaps in the cor- selected maximum fuel and oil load. A responding positions, the airplane is structural reserve fuel condition, not assumed to be subjected to symmet- exceeding 45 minutes of fuel under the rical maneuvers and gusts. The result- operating conditions in § 25.1001(e) and ing limit loads must correspond to the (f), as applicable, may be selected. conditions determined as follows: (b) If a structural reserve fuel condi- (1) Maneuvering to a positive limit tion is selected, it must be used as the load factor of 2.0; and minimum fuel weight condition for (2) Positive and negative gusts of 25 showing compliance with the flight ft/sec EAS acting normal to the flight load requirements as prescribed in this path in level flight. Gust loads result- subpart. In addition— ing on each part of the structure must (1) The structure must be designed be determined by rational analysis. for a condition of zero fuel and oil in The analysis must take into account the wing at limit loads corresponding the unsteady aerodynamic characteris- to— tics and rigid body motions of the air- (i) A maneuvering load factor of + craft. The shape of the gust must be as 2.25; and described in § 25.341(a)(2) except that— (ii) The gust and turbulence conditions of § 25.341(a) and (b), but assuming Uds = 25 ft/sec EAS; 85% of the gust velocities prescribed in H = 12.5 c; and § 25.341(a)(4) and 85% of the turbulence c = mean geometric chord of the wing (feet). intensities prescribed in § 25.341(b)(3). (b) The airplane must be designed for (2) Fatigue evaluation of the struc- the conditions prescribed in paragraph ture must account for any increase in (a) of this section, except that the air- operating stresses resulting from the plane load factor need not exceed 1.0,

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taking into account, as separate condi- (1) Conditions corresponding to tions, the effects of— steady rolling velocities must be inves- (1) Propeller slipstream cor- tigated. In addition, conditions corresponding to maximum continuous responding to maximum angular accel- power at the design flap speeds VF, and eration must be investigated for air- with takeoff power at not less than 1.4 planes with engines or other weight times the stalling speed for the par- concentrations outboard of the fuse- ticular flap position and associated lage. For the angular acceleration con- maximum weight; and ditions, zero rolling velocity may be (2) A head-on gust of 25 feet per sec- assumed in the absence of a rational ond velocity (EAS). time history investigation of the ma- (c) If flaps or other high lift devices neuver. are to be used in en route conditions, (2) At VA, a sudden deflection of the and with flaps in the appropriate posi- aileron to the stop is assumed. tion at speeds up to the flap design (3) At VC, the aileron deflection must speed chosen for these conditions, the be that required to produce a rate of airplane is assumed to be subjected to roll not less than that obtained in symmetrical maneuvers and gusts paragraph (a)(2) of this section. within the range determined by— (4) At VD, the aileron deflection must (1) Maneuvering to a positive limit be that required to produce a rate of load factor as prescribed in § 25.337(b); roll not less than one-third of that in and paragraph (a)(2) of this section. (2) The vertical gust and turbulence (b) Unsymmetrical gusts. The airplane conditions prescribed in § 25.341(a) and is assumed to be subjected to unsym- (b). metrical vertical gusts in level flight. (d) The airplane must be designed for The resulting limit loads must be de- a maneuvering load factor of 1.5 g at termined from either the wing max- the maximum take-off weight with the imum airload derived directly from wing-flaps and similar high lift devices § 25.341(a), or the wing maximum air- in the landing configurations. load derived indirectly from the vertical load factor calculated from [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.341(a). It must be assumed that 100 amended by Amdt. 25–46, 43 FR 50595, Oct. 30, percent of the wing air load acts on one 1978; Amdt. 25–72, 55 FR 37607, Sept. 17, 1990; side of the airplane and 80 percent of Amdt. 25–86, 61 FR 5221, Feb. 9, 1996; Amdt. 25–91, 62 FR 40704, July 29, 1997; Amdt. 25–141, the wing air load acts on the other 79 FR 73468, Dec. 11, 2014] side. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.349 Rolling conditions. amended by Amdt. 25–23, 35 FR 5672, Apr. 8, The airplane must be designed for 1970; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; loads resulting from the rolling condi- Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] tions specified in paragraphs (a) and (b) of this section. Unbalanced aero- § 25.351 Yaw maneuver conditions. dynamic moments about the center of The airplane must be designed for gravity must be reacted in a rational loads resulting from the yaw maneuver or conservative manner, considering conditions specified in paragraphs (a) the principal masses furnishing the re- through (d) of this section at speeds acting inertia forces. from VMC to VD. Unbalanced aero- (a) Maneuvering. The following condi- dynamic moments about the center of tions, speeds, and aileron deflections gravity must be reacted in a rational (except as the deflections may be lim- or conservative manner considering the ited by pilot effort) must be considered airplane inertia forces. In computing in combination with an airplane load the tail loads the yawing velocity may factor of zero and of two-thirds of the be assumed to be zero. positive maneuvering factor used in de- (a) With the airplane in unacceler- sign. In determining the required aile- ated flight at zero yaw, it is assumed ron deflections, the torsional flexi- that the cockpit rudder control is sud- bility of the wing must be considered denly displaced to achieve the result- in accordance with § 25.301(b): ing rudder deflection, as limited by:

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(1) The control system on control a rational analysis, a factor of 1.6 must surface stops; or be used. (2) A limit pilot force of 300 pounds (2) The limit engine torque to be con- from VMC to VA and 200 pounds from VC/ sidered under paragraph (a)(1) of this MC to VD/MD, with a linear variation section must be obtained by— between VA and VC/MC. (i) For turbopropeller installations, (b) With the cockpit rudder control multiplying mean engine torque for the deflected so as always to maintain the specified power/thrust and speed by a maximum rudder deflection available factor of 1.25; within the limitations specified in (ii) For other turbine engines, the paragraph (a) of this section, it is as- limit engine torque must be equal to sumed that the airplane yaws to the the maximum accelerating torque for overswing sideslip angle. the case considered. (c) With the airplane yawed to the (3) The engine mounts, pylons, and static equilibrium sideslip angle, it is adjacent supporting airframe structure assumed that the cockpit rudder con- must be designed to withstand 1g level trol is held so as to achieve the max- flight loads acting simultaneously with imum rudder deflection available with- the limit engine torque loads imposed in the limitations specified in para- by each of the following conditions to graph (a) of this section. be considered separately: (d) With the airplane yawed to the (i) Sudden maximum engine decelera- static equilibrium sideslip angle of tion due to malfunction or abnormal paragraph (c) of this section, it is as- condition; and sumed that the cockpit rudder control (ii) The maximum acceleration of en- is suddenly returned to neutral. gine. (b) For auxiliary power unit installa- [Amdt. 25–91, 62 FR 40704, July 29, 1997] tions, the power unit mounts and adjacent supporting airframe structure SUPPLEMENTARY CONDITIONS must be designed to withstand 1g level § 25.361 Engine and auxiliary power flight loads acting simultaneously with unit torque. the limit torque loads imposed by each of the following conditions to be con- (a) For engine installations— sidered separately: (1) Each engine mount, pylon, and ad- (1) Sudden maximum auxiliary power jacent supporting airframe structures unit deceleration due to malfunction, must be designed for the effects of— abnormal condition, or structural fail- (i) A limit engine torque cor- ure; and responding to takeoff power/thrust and, (2) The maximum acceleration of the if applicable, corresponding propeller auxiliary power unit. speed, acting simultaneously with 75% of the limit loads from flight condition [Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] A of § 25.333(b); (ii) A limit engine torque cor- § 25.362 Engine failure loads. responding to the maximum contin- (a) For engine mounts, pylons, and uous power/thrust and, if applicable, adjacent supporting airframe struc- corresponding propeller speed, acting ture, an ultimate loading condition simultaneously with the limit loads must be considered that combines 1g from flight condition A of § 25.333(b); flight loads with the most critical and transient dynamic loads and vibra- (iii) For turbopropeller installations tions, as determined by dynamic anal- only, in addition to the conditions ysis, resulting from failure of a blade, specified in paragraphs (a)(1)(i) and (ii) shaft, bearing or bearing support, or of this section, a limit engine torque bird strike event. Any permanent de- corresponding to takeoff power and formation from these ultimate load propeller speed, multiplied by a factor conditions must not prevent continued accounting for propeller control sys- safe flight and landing. tem malfunction, including quick (b) The ultimate loads developed feathering, acting simultaneously with from the conditions specified in para- 1g level flight loads. In the absence of graph (a) of this section are to be—

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(1) Multiplied by a factor of 1.0 when (e) Any structure, component or part, applied to engine mounts and pylons; inside or outside a pressurized com- and partment, the failure of which could (2) Multiplied by a factor of 1.25 when interfere with continued safe flight and applied to adjacent supporting air- landing, must be designed to withstand frame structure. the effects of a sudden release of pressure through an opening in any com- [Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] partment at any operating altitude re- § 25.363 Side load on engine and auxil- sulting from each of the following con- iary power unit mounts. ditions: (1) The penetration of the compart- (a) Each engine and auxiliary power ment by a portion of an engine fol- unit mount and its supporting struc- lowing an engine disintegration; ture must be designed for a limit load (2) Any opening in any pressurized factor in lateral direction, for the side compartment up to the size H in load on the engine and auxiliary power o square feet; however, small compart- unit mount, at least equal to the max- ments may be combined with an adja- imum load factor obtained in the yaw- cent pressurized compartment and both ing conditions but not less than— considered as a single compartment for (1) 1.33; or openings that cannot reasonably be ex- (2) One-third of the limit load factor pected to be confined to the small com- for flight condition A as prescribed in partment. The size Ho must be com- § 25.333(b). puted by the following formula: (b) The side load prescribed in paragraph (a) of this section may be as- Ho = PAs sumed to be independent of other flight where, conditions. Ho = Maximum opening in square feet, need [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as not exceed 20 square feet. amended by Amdt. 25–23, 35 FR 5672, Apr. 8, P = (As/6240) + .024 1970; Amdt. 25–91, 62 FR 40704, July 29, 1997] As = Maximum cross-sectional area of the pressurized shell normal to the longitu- § 25.365 Pressurized compartment dinal axis, in square feet; and loads. (3) The maximum opening caused by For airplanes with one or more pres- airplane or equipment failures not surized compartments the following shown to be extremely improbable. apply: (f) In complying with paragraph (e) of (a) The airplane structure must be this section, the fail-safe features of strong enough to withstand the flight the design may be considered in deter- loads combined with pressure differen- mining the probability of failure or tial loads from zero up to the max- penetration and probable size of open- imum relief valve setting. ings, provided that possible improper (b) The external pressure distribution operation of closure devices and inad- in flight, and stress concentrations and vertent door openings are also consid- fatigue effects must be accounted for. ered. Furthermore, the resulting dif- (c) If landings may be made with the ferential pressure loads must be com- compartment pressurized, landing bined in a rational and conservative loads must be combined with pressure manner with 1–g level flight loads and differential loads from zero up to the any loads arising from emergency de- maximum allowed during landing. pressurization conditions. These loads (d) The airplane structure must be may be considered as ultimate condi- designed to be able to withstand the tions; however, any deformations asso- pressure differential loads cor- ciated with these conditions must not responding to the maximum relief interfere with continued safe flight and valve setting multiplied by a factor of landing. The pressure relief provided by 1.33 for airplanes to be approved for op- intercompartment venting may also be eration to 45,000 feet or by a factor of considered. 1.67 for airplanes to be approved for op- (g) Bulkheads, floors, and partitions eration above 45,000 feet, omitting in pressurized compartments for occu- other loads. pants must be designed to withstand

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the conditions specified in paragraph § 25.371 Gyroscopic loads. (e) of this section. In addition, reason- The structure supporting any engine able design precautions must be taken or auxiliary power unit must be de- to minimize the probability of parts signed for the loads, including gyro- becoming detached and injuring occu- scopic loads, arising from the condi- pants while in their seats. tions specified in §§ 25.331, 25.341, 25.349, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as 25.351, 25.473, 25.479, and 25.481, with the amended by Amdt. 25–54, 45 FR 60172, Sept. engine or auxiliary power unit at the 11, 1980; Amdt. 25–71, 55 FR 13477, Apr. 10, maximum rotating speed appropriate 1990; Amdt. 25–72, 55 FR 29776, July 20, 1990; to the condition. For the purposes of Amdt. 25–87, 61 FR 28695, June 5, 1996] compliance with this paragraph, the pitch maneuver in § 25.331(c)(1) must be § 25.367 Unsymmetrical loads due to carried out until the positive limit ma- engine failure. neuvering load factor (point A2 in (a) The airplane must be designed for § 25.333(b)) is reached. the unsymmetrical loads resulting from the failure of the critical engine. [Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] Turbopropeller airplanes must be de- § 25.373 Speed control devices. signed for the following conditions in combination with a single malfunction If speed control devices (such as of the propeller drag limiting system, spoilers and drag flaps) are installed considering the probable pilot correc- for use in en route conditions— tive action on the flight controls: (a) The airplane must be designed for the symmetrical maneuvers prescribed (1) At speeds between VMC and VD, the loads resulting from power failure be- in §§ 25.333 and 25.337, the yawing ma- cause of fuel flow interruption are con- neuvers in § 25.351, and the vertical and sidered to be limit loads. lateral gust and turbulence conditions prescribed in § 25.341(a) and (b) at each (2) At speeds between V and V the MC C, setting and the maximum speed associ- loads resulting from the disconnection ated with that setting; and of the engine compressor from the tur- (b) If the device has automatic oper- bine or from loss of the turbine blades ating or load limiting features, the air- are considered to be ultimate loads. plane must be designed for the maneu- (3) The time history of the thrust ver and gust conditions prescribed in decay and drag build-up occurring as a paragraph (a) of this section, at the result of the prescribed engine failures speeds and corresponding device posi- must be substantiated by test or other tions that the mechanism allows. data applicable to the particular engine-propeller combination. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (4) The timing and magnitude of the amended by Amdt. 25–72, 55 FR 29776, July 20, probable pilot corrective action must 1990; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] be conservatively estimated, considering the characteristics of the par- CONTROL SURFACE AND SYSTEM LOADS ticular engine-propeller-airplane combination. § 25.391 Control surface loads: Gen- (b) Pilot corrective action may be as- eral. sumed to be initiated at the time max- The control surfaces must be de- imum yawing velocity is reached, but signed for the limit loads resulting not earlier than two seconds after the from the flight conditions in §§ 25.331, engine failure. The magnitude of the 25.341(a) and (b), 25.349, and 25.351, con- corrective action may be based on the sidering the requirements for— control forces specified in § 25.397(b) ex- (a) Loads parallel to hinge line, in cept that lower forces may be assumed § 25.393; where it is shown by anaylsis or test (b) Pilot effort effects, in § 25.397; that these forces can control the yaw (c) Trim tab effects, in § 25.407; and roll resulting from the prescribed (d) Unsymmetrical loads, in § 25.427; engine failure conditions. and

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(e) Auxiliary aerodynamic surfaces, imum values specified for the aileron in § 25.445. and elevator may be used if control [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as surface hinge moments are based on re- amended by Amdt. 25–86, 61 FR 5222, Feb. 9, liable data. In applying this criterion, 1996; Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] the effects of servo mechanisms, tabs, and automatic pilot systems, must be § 25.393 Loads parallel to hinge line. considered. (a) Control surfaces and supporting (c) Limit pilot forces and torques. The hinge brackets must be designed for in- limit pilot forces and torques are as ertia loads acting parallel to the hinge follows: line. (b) In the absence of more rational Maximum Minimum Control forces or forces or data, the inertia loads may be assumed torques torques to be equal to KW, where— (1) K = 24 for vertical surfaces; Aileron: (2) K = 12 for horizontal surfaces; and Stick ...... 100 lbs ...... 40 lbs. (3) W = weight of the movable sur- Wheel 1 ...... 80 D in.-lbs 2 ... 40 D in.-lbs. faces. Elevator: Stick ...... 250 lbs ...... 100 lbs. Wheel (symmetrical) ..... 300 lbs ...... 100 lbs. § 25.395 Control system. Wheel (unsymmetrical) 3 ...... 100 lbs. (a) Longitudinal, lateral, directional, Rudder ...... 300 lbs ...... 130 lbs. and drag control system and their sup- 1 The critical parts of the aileron control system must be de- porting structures must be designed for signed for a single tangential force with a limit value equal to loads corresponding to 125 percent of 1.25 times the couple force determined from these criteria. 2 D = wheel diameter (inches). the computed hinge moments of the 3 The unsymmetrical forces must be applied at one of the movable control surface in the condi- normal handgrip points on the periphery of the control wheel. tions prescribed in § 25.391. (b) The system limit loads of para- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as graph (a) of this section need not ex- amended by Amdt. 25–38, 41 FR 55466, Dec. 20, ceed the loads that can be produced by 1976; Amdt. 25–72, 55 FR 29776, July 20, 1990] the pilot (or pilots) and by automatic § 25.399 Dual control system. or power devices operating the controls. (a) Each dual control system must be (c) The loads must not be less than designed for the pilots operating in op- those resulting from application of the position, using individual pilot forces minimum forces prescribed in not less than— § 25.397(c). (1) 0.75 times those obtained under [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.395; or amended by Amdt. 25–23, 35 FR 5672, Apr. 8, (2) The minimum forces specified in 1970; Amdt. 25–72, 55 FR 29776, July 20, 1990; § 25.397(c). Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] (b) The control system must be de- § 25.397 Control system loads. signed for pilot forces applied in the same direction, using individual pilot (a) General. The maximum and min- forces not less than 0.75 times those ob- imum pilot forces, specified in para- tained under § 25.395. graph (c) of this section, are assumed to act at the appropriate control grips § 25.405 Secondary control system. or pads (in a manner simulating flight conditions) and to be reacted at the at- Secondary controls, such as wheel tachment of the control system to the brake, spoiler, and tab controls, must control surface horn. be designed for the maximum forces (b) Pilot effort effects. In the control that a pilot is likely to apply to those surface flight loading condition, the air controls. The following values may be loads on movable surfaces and the cor- used: responding deflections need not exceed those that would result in flight from the application of any pilot force within the ranges specified in paragraph (c) of this section. Two-thirds of the max-

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PILOT CONTROL FORCE LIMITS (SECONDARY loads generated when the airplane is CONTROLS) subjected to a horizontal 65-knot ground gust from any direction while Control Limit pilot forces taxiing and while parked. For airplanes Miscellaneous: equipped with control system gust *Crank, wheel, or lever .. ((1 + R) / 3) × 50 lbs., but not locks, the taxiing condition must be less than 50 lbs. nor more evaluated with the controls locked and than 150 lbs. (R = radius). (Applicable to any angle with- unlocked, and the parked condition in 20° of plane of control). must be evaluated with the controls Twist ...... 133 in.–lbs. locked. Push-pull ...... To be chosen by applicant. (b) The control system and surface *Limited to flap, tab, stabilizer, spoiler, and landing gear op- loads due to ground gust may be as- eration controls. sumed to be static loads, and the hinge § 25.407 Trim tab effects. moments H must be computed from the formula: The effects of trim tabs on the con- H = K (1/2) r V2 c S trol surface design conditions must be o accounted for only where the surface Where— loads are limited by maximum pilot ef- K = hinge moment factor for ground gusts fort. In these cases, the tabs are con- derived in paragraph (c) of this section; sidered to be deflected in the direction ro = density of air at sea level; that would assist the pilot, and the de- V = 65 knots relative to the aircraft; flections are— S = area of the control surface aft of the (a) For elevator trim tabs, those re- hinge line; c = mean aerodynamic chord of the control quired to trim the airplane at any surface aft of the hinge line. point within the positive portion of the pertinent flight envelope in § 25.333(b), (c) The hinge moment factor K for except as limited by the stops; and ground gusts must be taken from the (b) For aileron and rudder trim tabs, following table: those required to trim the airplane in Position of the critical unsymmetrical power and Surface K controls loading conditions, with appropriate allowance for rigging tolerances. (1) Aileron ...... 0 .75 Control column locked or lashed in mid-position. § 25.409 Tabs. (2) Aileron ...... * ±0 .50 Ailerons at full throw. (a) Trim tabs. Trim tabs must be de- (3) Elevator ...... * ±0 .75 Elevator full down. (4) Elevator ...... * ±0 .75 Elevator full up. signed to withstand loads arising from (5) Rudder ...... 0 .75 Rudder in neutral. all likely combinations of tab setting, (6) Rudder ...... 0 .75 Rudder at full throw. primary control position, and airplane * A positive value of K indicates a moment tending to de- speed (obtainable without exceeding press the surface, while a negative value of K indicates a mo- the flight load conditions prescribed ment tending to raise the surface. for the airplane as a whole), when the (d) The computed hinge moment of effect of the tab is opposed by pilot ef- paragraph (b) of this section must be fort forces up to those specified in used to determine the limit loads due § 25.397(b). to ground gust conditions for the con- (b) Balancing tabs. Balancing tabs trol surface. A 1.25 factor on the com- must be designed for deflections con- puted hinge moments must be used in sistent with the primary control sur- calculating limit control system loads. face loading conditions. (e) Where control system flexibility (c) Servo tabs. Servo tabs must be de- is such that the rate of load applica- signed for deflections consistent with tion in the ground gust conditions the primary control surface loading might produce transient stresses appre- conditions obtainable within the pilot ciably higher than those corresponding maneuvering effort, considering pos- to static loads, in the absence of a ra- sible opposition from the trim tabs. tional analysis substantiating a different dynamic factor, an additional § 25.415 Ground gust conditions. factor of 1.6 must be applied to the con- (a) The flight control systems and trol system loads of paragraph (d) of surfaces must be designed for the limit this section to obtain limit loads. If a

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rational analysis is used, the addi- (1) 100 percent of the maximum load- tional factor must not be less than 1.2. ing from the symmetrical maneuver (f) For the condition of the control conditions of § 25.331 and the vertical locks engaged, the control surfaces, the gust conditions of § 25.341(a) acting sep- control system locks, and the parts of arately on the surface on one side of any control systems between the sur- the plane of symmetry; and faces and the locks must be designed to (2) 80 percent of these loadings acting the resultant limit loads. Where con- on the other side. trol locks are not provided, then the (c) For empennage arrangements control surfaces, the control system where the horizontal tail surfaces have stops nearest the surfaces, and the dihedral angles greater than plus or parts of any control systems between minus 10 degrees, or are supported by the surfaces and the stops must be de- the vertical tail surfaces, the surfaces signed to the resultant limit loads. If and the supporting structure must be the control system design is such as to designed for gust velocities specified in allow any part of the control system to § 25.341(a) acting in any orientation at impact with the stops due to flexi- right angles to the flight path. bility, then the resultant impact loads (d) Unsymmetrical loading on the must be taken into account in deriving empennage arising from buffet condi- the limit loads due to ground gust. tions of § 25.305(e) must be taken into (g) For the condition of taxiing with account. the control locks disengaged, or where [Doc. No. 27902, 61 FR 5222, Feb. 9, 1996] control locks are not provided, the following apply: § 25.445 Auxiliary aerodynamic sur- (1) The control surfaces, the control faces. system stops nearest the surfaces, and (a) When significant, the aero- the parts of any control systems be- dynamic influence between auxiliary tween the surfaces and the stops must aerodynamic surfaces, such as out- be designed to the resultant limit board fins and winglets, and their sup- loads. porting aerodynamic surfaces, must be (2) The parts of the control systems taken into account for all loading con- between the stops nearest the surfaces ditions including pitch, roll, and yaw and the flight deck controls must be maneuvers, and gusts as specified in designed to the resultant limit loads, § 25.341(a) acting at any orientation at except that the parts of the control right angles to the flight path. system where loads are eventually re- (b) To provide for unsymmetrical acted by the pilot need not exceed: loading when outboard fins extend (i) The loads corresponding to the above and below the horizontal surface, maximum pilot loads in § 25.397(c) for the critical vertical surface loading each pilot alone; or (load per unit area) determined under (ii) 0.75 times these maximum loads § 25.391 must also be applied as follows: for each pilot when the pilot forces are (1) 100 percent to the area of the applied in the same direction. vertical surfaces above (or below) the [Amdt. 25–141, 79 FR 73468, Dec. 11, 2014] horizontal surface. (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 empennage arising from effects such as § 25.457 Wing flaps. 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 oc- (b) The horizontal tail must be as- curring in the conditions prescribed in sumed to be subjected to unsymmet- § 25.345, accounting for the loads occur- rical loading conditions determined as ring during transition from one flap po- follows: sition and airspeed to another.

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§ 25.459 Special devices. § 25.473 Landing load conditions and assumptions. The loading for special devices using aerodynamic surfaces (such as slots, (a) For the landing conditions speci- slats and spoilers) must be determined fied in § 25.479 to § 25.485 the airplane is from test data. assumed to contact the ground— (1) In the attitudes defined in § 25.479 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as and § 25.481; amended by Amdt. 25–72, 55 FR 29776, July 20, (2) With a limit descent velocity of 10 1990] fps at the design landing weight (the maximum weight for landing condi- GROUND LOADS tions at maximum descent velocity); and § 25.471 General. (3) With a limit descent velocity of 6 (a) Loads and equilibrium. For limit fps at the design take-off weight (the ground loads— maximum weight for landing condi- (1) Limit ground loads obtained tions at a reduced descent velocity). under this subpart are considered to be (4) The prescribed descent velocities external forces applied to the airplane may be modified if it is shown that the structure; and airplane has design features that make (2) In each specified ground load con- it impossible to develop these veloci- dition, the external loads must be ties. placed in equilibrium with the linear (b) Airplane lift, not exceeding airplane weight, may be assumed unless and angular inertia loads in a rational the presence of systems or procedures or conservative manner. significantly affects the lift. (b) Critical centers of gravity. The crit- (c) The method of analysis of air- ical centers of gravity within the range plane and landing gear loads must take for which certification is requested into account at least the following ele- must be selected so that the maximum ments: design loads are obtained in each land- (1) Landing gear dynamic character- ing gear element. Fore and aft, istics. vertical, and lateral airplane centers of (2) Spin-up and springback. gravity must be considered. Lateral (3) Rigid body response. displacements of the c.g. from the air- (4) Structural dynamic response of plane centerline which would result in the airframe, if significant. main gear loads not greater than 103 (d) The landing gear dynamic charac- percent of the critical design load for teristics must be validated by tests as symmetrical loading conditions may be defined in § 25.723(a). selected without considering the ef- (e) The coefficient of friction between fects of these lateral c.g. displacements the tires and the ground may be estab- on the loading of the main gear ele- lished by considering the effects of ments, or on the airplane structure skidding velocity and tire pressure. provided— However, this coefficient of friction (1) The lateral displacement of the need not be more than 0.8. c.g. results from random passenger or [Amdt. 25–91, 62 FR 40705, July 29, 1997; Amdt. cargo disposition within the fuselage or 25–91, 62 FR 45481, Aug. 27, 1997; Amdt. 25–103, from random unsymmetrical fuel load- 66 FR 27394, May 16, 2001] ing or fuel usage; and § 25.477 Landing gear arrangement. (2) Appropriate loading instructions for random disposable loads are in- Sections 25.479 through 25.485 apply cluded under the provisions of to airplanes with conventional ar- § 25.1583(c)(1) to ensure that the lateral rangements of main and nose gears, or displacement of the center of gravity is main and tail gears, when normal oper- maintained within these limits. ating techniques are used. (c) Landing gear dimension data. Fig- § 25.479 Level landing conditions. ure 1 of appendix A contains the basic (a) In the level attitude, the airplane landing gear dimension data. is assumed to contact the ground at [Amdt. 25–23, 35 FR 5673, Apr. 8, 1970] forward velocity components, ranging

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from VL1 to 1.25 VL2 parallel to the (ii) The shock absorber and tire de- ground under the conditions prescribed flections must be assumed to be 75% of in § 25.473 with— the deflection corresponding to the (1) VL1 equal to VS0 (TAS) at the ap- maximum ground reaction of propriate landing weight and in stand- § 25.473(a)(2). This load case need not be ard sea level conditions; and considered in combination with flat (2) VL2 equal to VS0 (TAS) at the ap- tires. propriate landing weight and altitudes (3) The combination of vertical and in a hot day temperature of 41 degrees drag components is considered to be F. above standard. acting at the wheel axle centerline. (3) The effects of increased contact [Amdt. 25–91, 62 FR 40705, July 29, 1997; Amdt. speed must be investigated if approval 25–91, 62 FR 45481, Aug. 27, 1997] of downwind landings exceeding 10 knots is requested. § 25.481 Tail-down landing conditions. (b) For the level landing attitude for (a) In the tail-down attitude, the air- airplanes with tail wheels, the condi- plane is assumed to contact the ground tions specified in this section must be at forward velocity components, rang- investigated with the airplane hori- ing from V to V parallel to the zontal reference line horizontal in ac- L1 L2 ground under the conditions prescribed cordance with Figure 2 of Appendix A in § 25.473 with— of this part. (1) (c) For the level landing attitude for VL1 equal to VS0 (TAS) at the ap- airplanes with nose wheels, shown in propriate landing weight and in stand- Figure 2 of Appendix A of this part, the ard sea level conditions; and conditions specified in this section (2) VL2 equal to VS0 (TAS) at the ap- must be investigated assuming the fol- propriate landing weight and altitudes lowing attitudes: in a hot day temperature of 41 degrees (1) An attitude in which the main F. above standard. wheels are assumed to contact the (3) The combination of vertical and ground with the nose wheel just clear drag components considered to be act- of the ground; and ing at the main wheel axle centerline. (2) If reasonably attainable at the (b) For the tail-down landing condi- specified descent and forward veloci- tion for airplanes with tail wheels, the ties, an attitude in which the nose and main and tail wheels are assumed to main wheels are assumed to contact contact the ground simultaneously, in the ground simultaneously. accordance with figure 3 of appendix A. (d) In addition to the loading condi- Ground reaction conditions on the tail tions prescribed in paragraph (a) of this wheel are assumed to act— section, but with maximum vertical (1) Vertically; and ground reactions calculated from para- (2) Up and aft through the axle at 45 graph (a), the following apply: degrees to the ground line. (1) The landing gear and directly af- (c) For the tail-down landing condi- fected attaching structure must be de- tion for airplanes with nose wheels, the signed 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 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as into account. In absence of a more ra- amended by Amdt. 25–91, 62 FR 40705, July 29, tional analysis of this condition, the 1997; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] following must be investigated: (i) A vertical load equal to 75% of the § 25.483 One-gear landing conditions. maximum ground reaction of § 25.473 For the one-gear landing conditions, must be considered in combination the airplane is assumed to be in the with a drag and side load of 40% and level attitude and to contact the 25% respectively of that vertical load. ground on one main landing gear, in

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accordance with Figure 4 of Appendix maximum weight for ground handling A of this part. In this attitude— conditions). No wing lift may be con- (a) The ground reactions must be the sidered. The shock absorbers and tires same as those obtained on that side may be assumed to be in their static under § 25.479(d)(1), and position. (b) Each unbalanced external load must be reacted by airplane inertia in [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5673, Apr. 8, a rational or conservative manner. 1970] [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–91, 62 FR 40705, July 29, § 25.491 Taxi, takeoff and landing roll. 1997] Within the range of appropriate ground speeds and approved weights, § 25.485 Side load conditions. the airplane structure and landing gear In addition to § 25.479(d)(2) the fol- are assumed to be subjected to loads lowing conditions must be considered: not less than those obtained when the (a) For the side load condition, the aircraft is operating over the roughest airplane is assumed to be in the level ground that may reasonably be ex- attitude with only the main wheels pected in normal operation. contacting the ground, in accordance with figure 5 of appendix A. [Amdt. 25–91, 62 FR 40705, July 29, 1997] (b) Side loads of 0.8 of the vertical reaction (on one side) acting inward and § 25.493 Braked roll conditions. 0.6 of the vertical reaction (on the (a) An airplane with a tail wheel is other side) acting outward must be assumed to be in the level attitude combined with one-half of the max- with the load on the main wheels, in imum vertical ground reactions ob- accordance with figure 6 of appendix A. tained in the level landing conditions. The limit vertical load factor is 1.2 at These loads are assumed to be applied the design landing weight and 1.0 at at the ground contact point and to be the design ramp weight. A drag reac- resisted by the inertia of the airplane. tion equal to the vertical reaction mul- The drag loads may be assumed to be tiplied by a coefficient of friction of zero. 0.8, must be combined with the vertical [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ground reaction and applied at the amended by Amdt. 25–91, 62 FR 40705, July 29, ground contact point. 1997] (b) For an airplane with a nose wheel the limit vertical load factor is 1.2 at § 25.487 Rebound landing condition. the design landing weight, and 1.0 at (a) The landing gear and its sup- the design ramp weight. A drag reac- porting structure must be investigated tion equal to the vertical reaction, for the loads occurring during rebound multiplied by a coefficient of friction of the airplane from the landing sur- of 0.8, must be combined with the face. vertical reaction and applied at the (b) With the landing gear fully ex- ground contact point of each wheel tended and not in contact with the with brakes. The following two atti- ground, a load factor of 20.0 must act tudes, in accordance with figure 6 of on the unsprung weights of the landing appendix A, must be considered: gear. This load factor must act in the (1) The level attitude with the wheels direction of motion of the unsprung contacting the ground and the loads weights as they reach their limiting distributed between the main and nose positions in extending with relation to gear. Zero pitching acceleration is as- the sprung parts of the landing gear. sumed. (2) The level attitude with only the § 25.489 Ground handling conditions. main gear contacting the ground and Unless otherwise prescribed, the with the pitching moment resisted by landing gear and airplane structure angular acceleration. must be investigated for the conditions (c) A drag reaction lower than that in §§ 25.491 through 25.509 with the air- prescribed in this section may be used plane at the design ramp weight (the if it is substantiated that an effective

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drag force of 0.8 times the vertical re- § 25.495 Turning. action cannot be attained under any In the static position, in accordance likely loading condition. with figure 7 of appendix A, the air- (d) An airplane equipped with a nose plane is assumed to execute a steady gear must be designed to withstand the turn by nose gear steering, or by appli- loads arising from the dynamic pitch- cation of sufficient differential power, ing motion of the airplane due to sud- so that the limit load factors applied at den application of maximum braking the center of gravity are 1.0 vertically force. The airplane is considered to be and 0.5 laterally. The side ground reac- at design takeoff weight with the nose tion of each wheel must be 0.5 of the and main gears in contact with the vertical reaction. ground, and with a steady-state § 25.497 Tail-wheel yawing. vertical load factor of 1.0. The steady- state nose gear reaction must be com- (a) A vertical ground reaction equal bined with the maximum incremental to the static load on the tail wheel, in nose gear vertical reaction caused by combination with a side component of the sudden application of maximum equal magnitude, is assumed. braking force as described in para- (b) If there is a swivel, the tail wheel is assumed to be swiveled 90° to the air- graphs (b) and (c) of this section. plane longitudinal axis with the result- (e) In the absence of a more rational ant load passing through the axle. analysis, the nose gear vertical reac- (c) If there is a lock, steering device, tion prescribed in paragraph (d) of this or shimmy damper the tail wheel is section must be calculated according also assumed to be in the trailing posi- to the following formula: tion with the side load acting at the ground contact point. W ⎡ fAEμ ⎤ V = T ⎢B + ⎥ § 25.499 Nose-wheel yaw and steering. N + ++μ AB⎣ AB E⎦ (a) A vertical load factor of 1.0 at the Where: airplane center of gravity, and a side component at the nose wheel ground VN = Nose gear vertical reaction. contact equal to 0.8 of the vertical WT = Design takeoff weight. 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 cen- in static equilibrium with the loads re- ters of the main wheels. sulting from the use of brakes on one E = Vertical height of the c.g. of the airplane side of the main landing gear, the nose above the ground in the 1.0 g static con- gear, its attaching structure, and the dition. fuselage structure forward of the cen- μ = Coefficient of friction of 0.80. ter of gravity must be designed for the f = Dynamic response factor; 2.0 is to be used following loads: unless a lower factor is substantiated. In (1) A vertical load factor at the cen- the absence of other information, the dy- ter of gravity of 1.0. namic response factor f may be defined (2) A forward acting load at the air- by the equation: plane center of gravity of 0.8 times the vertical load on one main gear. ⎛ −πξ ⎞ (3) Side and vertical loads at the =+ ⎜ ⎟ f 1 exp⎜ ⎟ ground contact point on the nose gear ⎝ 1− ξ 2 ⎠ that are required for static equilibrium. Where: (4) A side load factor at the airplane x is the effective critical damping ratio of center of gravity of zero. the rigid body pitching mode about the (c) If the loads prescribed in para- main landing gear effective ground con- graph (b) of this section result in a tact point. nose gear side load higher than 0.8 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as times the vertical nose gear load, the amended by Amdt. 25–23, 35 FR 5673, Apr. 8, design nose gear side load may be lim- 1970; Amdt. 25–97, 63 FR 29072, May 27, 1998] ited to 0.8 times the vertical load, with

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unbalanced yawing moments assumed (c) For airplanes with tail wheels, the to be resisted by airplane inertia resultant of the ground reactions must forces. pass through the center of gravity of (d) For other than the nose gear, its the airplane. attaching structure, and the forward fuselage structure, the loading condi- § 25.509 Towing loads. tions are those prescribed in paragraph (a) The towing loads specified in (b) of this section, except that— paragraph (d) of this section must be (1) A lower drag reaction may be used considered separately. These loads if an effective drag force of 0.8 times must be applied at the towing fittings the vertical reaction cannot be reached and must act parallel to the ground. In under any likely loading condition; and addition— (2) The forward acting load at the (1) A vertical load factor equal to 1.0 center of gravity need not exceed the must be considered acting at the center maximum drag reaction on one main of gravity; gear, determined in accordance with (2) The shock struts and tires must § 25.493(b). be in their static positions; and (e) With the airplane at design ramp (3) With WT as the design ramp weight, and the nose gear in any steer- weight, the towing load, FTOW, is— able position, the combined application (i) 0.3 WT for WT less than 30,000 of full normal steering torque and pounds; vertical force equal to 1.33 times the (ii) (6WT + 450,000)/70 for WT between maximum static reaction on the nose 30,000 and 100,000 pounds; and gear must be considered in designing (iii) 0.15 WT for WT over 100,000 the nose gear, its attaching structure, pounds. and the forward fuselage structure. (b) For towing points not on the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as landing gear but near the plane of sym- amended by Amdt. 25–23, 35 FR 5673, Apr. 8, metry of the airplane, the drag and 1970; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; side tow load components specified for Amdt. 25–91, 62 FR 40705, July 29, 1997] the auxiliary gear apply. For towing points located outboard of the main § 25.503 Pivoting. gear, the drag and side tow load compo- (a) The airplane is assumed to pivot nents specified for the main gear apply. about one side of the main gear with Where the specified angle of swivel the brakes on that side locked. The cannot be reached, the maximum ob- limit vertical load factor must be 1.0 tainable angle must be used. and the coefficient of friction 0.8. (c) The towing loads specified in (b) The airplane is assumed to be in paragraph (d) of this section must be static equilibrium, with the loads being reacted as follows: applied at the ground contact points, (1) The side component of the towing in accordance with figure 8 of appendix load at the main gear must be reacted A. by a side force at the static ground line of the wheel to which the load is ap- § 25.507 Reversed braking. plied. (a) The airplane must be in a three (2) The towing loads at the auxiliary point static ground attitude. Hori- gear and the drag components of the zontal reactions parallel to the ground towing loads at the main gear must be and directed forward must be applied reacted as follows: at the ground contact point of each (i) A reaction with a maximum value wheel with brakes. The limit loads equal to the vertical reaction must be must be equal to 0.55 times the vertical applied at the axle of the wheel to load at each wheel or to the load devel- which the load is applied. Enough air- oped by 1.2 times the nominal max- plane inertia to achieve equilibrium imum static brake torque, whichever is must be applied. less. (ii) The loads must be reacted by air- (b) For airplanes with nose wheels, plane inertia. the pitching moment must be balanced (d) The prescribed towing loads are as by rotational inertia. follows:

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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]

§ 25.511 Ground load: unsymmetrical (3) Any unequal tire inflation pres- loads on multiple-wheel units. sure, assuming the maximum variation ± (a) General. Multiple-wheel landing to be 5 percent of the nominal tire in- gear units are assumed to be subjected flation pressure. to the limit ground loads prescribed in (4) A runway crown of zero and a run- this subpart under paragraphs (b) way crown having a convex upward through (f) of this section. In addi- shape that may be approximated by a tion— slope of 11⁄2 percent with the hori- (1) A tandem strut gear arrangement zontal. Runway crown effects must be is a multiple-wheel unit; and considered with the nose gear unit on (2) In determining the total load on a either slope of the crown. gear unit with respect to the provisions (5) The airplane attitude. of paragraphs (b) through (f) of this (6) Any structural deflections. section, the transverse shift in the load (c) Deflated tires. The effect of de- centroid, due to unsymmetrical load flated tires on the structure must be distribution on the wheels, may be ne- considered with respect to the loading glected. conditions specified in paragraphs (d) (b) Distribution of limit loads to wheels; through (f) of this section, taking into tires inflated. The distribution of the account the physical arrangement of limit loads among the wheels of the the gear components. In addition— landing gear must be established for (1) The deflation of any one tire for each landing, taxiing, and ground han- each multiple wheel landing gear unit, dling condition, taking into account and the deflation of any two critical the effects of the following factors: tires for each landing gear unit using (1) The number of wheels and their four or more wheels per unit, must be physical arrangements. For truck type considered; and landing gear units, the effects of any (2) The ground reactions must be ap- seesaw motion of the truck during the plied to the wheels with inflated tires landing impact must be considered in except that, for multiple-wheel gear determining the maximum design loads units with more than one shock strut, for the fore and aft wheel pairs. a rational distribution of the ground (2) Any differentials in tire diameters reactions between the deflated and in- resulting from a combination of manu- flated tires, accounting for the dif- facturing tolerances, tire growth, and ferences in shock strut extensions re- tire wear. A maximum tire-diameter sulting from a deflated tire, may be differential equal to 2⁄3 of the most unused. favorable combination of diameter (d) Landing conditions. For one and variations that is obtained when tak- for two deflated tires, the applied load ing into account manufacturing toler- to each gear unit is assumed to be 60 ances, tire growth, and tire wear, may percent and 50 percent, respectively, of be assumed. the limit load applied to each gear for

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each of the prescribed landing condi- (2) For jacking by other airplane tions. However, for the drift landing structure at maximum approved jack- condition of § 25.485, 100 percent of the ing weight: vertical load must be applied. (i) The airplane structure must be de- (e) Taxiing and ground handling condi- signed for a vertical load of 1.33 times tions. For one and for two deflated the vertical reaction at each jacking tires— point acting singly and in combination (1) The applied side or drag load fac- with a horizontal load of 0.33 times the tor, or both factors, at the center of vertical static reaction applied in any gravity must be the most critical value direction. up to 50 percent and 40 percent, respec- (ii) The jacking pads and local struc- tively, of the limit side or drag load ture must be designed for a vertical factors, or both factors, corresponding load of 2.0 times the vertical static re- to the most severe condition resulting action at each jacking point, acting from consideration of the prescribed singly and in combination with a hori- taxiing and ground handling condi- zontal load of 0.33 times the vertical tions; static reaction applied in any direc- (2) For the braked roll conditions of tion. § 25.493 (a) and (b)(2), the drag loads on (c) Tie-down. If tie-down points are each inflated tire may not be less than provided, the main tie-down points and those at each tire for the symmetrical local structure must withstand the load distribution with no deflated tires; limit loads resulting from a 65-knot (3) The vertical load factor at the horizontal wind from any direction. center of gravity must be 60 percent [Doc. No. 26129, 59 FR 22102, Apr. 28, 1994] and 50 percent, respectively, of the factor with no deflated tires, except that WATER LOADS it may not be less than 1g; and (4) Pivoting need not be considered. § 25.521 General. (f) Towing conditions. For one and for (a) Seaplanes must be designed for two deflated tires, the towing load, the water loads developed during take- FTOW, must be 60 percent and 50 percent, off and landing, with the seaplane in respectively, of the load prescribed. any attitude likely to occur in normal operation, and at the appropriate for- § 25.519 Jacking and tie-down provi- ward and sinking velocities under the sions. most severe sea conditions likely to be (a) General. The airplane must be de- encountered. signed to withstand the limit load con- (b) Unless a more rational analysis of ditions resulting from the static the water loads is made, or the stand- ground load conditions of paragraph (b) ards in ANC–3 are used, §§ 25.523 of this section and, if applicable, para- through 25.537 apply. graph (c) of this section at the most (c) The requirements of this section critical combinations of airplane and §§ 25.523 through 25.537 apply also to weight and center of gravity. The max- amphibians. imum allowable load at each jack pad must be specified. § 25.523 Design weights and center of (b) Jacking. The airplane must have gravity positions. provisions for jacking and must with- (a) Design weights. The water load re- stand the following limit loads when quirements must be met at each oper- the airplane is supported on jacks— ating weight up to the design landing (1) For jacking by the landing gear at weight except that, for the takeoff con- the maximum ramp weight of the air- dition prescribed in § 25.531, the design plane, the airplane structure must be water takeoff weight (the maximum designed for a vertical load of 1.33 weight for water taxi and takeoff run) times the vertical static reaction at must be used. each jacking point acting singly and in (b) Center of gravity positions. The combination with a horizontal load of critical centers of gravity within the 0.33 times the vertical static reaction limits for which certification is re- applied in any direction. quested must be considered to reach

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maximum design loads for each part of (3) VS0 = seaplane stalling speed in the seaplane structure. knots with flaps extended in the appropriate landing position and with no [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5673, Apr. 8, slipstream effect. 1970] (4) b = angle of dead rise at the longitudinal station at which the load fac- § 25.525 Application of loads. tor is being determined in accordance with figure 1 of appendix B. (a) Unless otherwise prescribed, the (5) W= seaplane design landing seaplane as a whole is assumed to be weight in pounds. subjected to the loads corresponding to (6) K = empirical hull station weigh- the load factors specified in § 25.527. 1 ing factor, in accordance with figure 2 (b) In applying the loads resulting of appendix B. from the load factors prescribed in (7) rx = ratio of distance, measured § 25.527, the loads may be distributed parallel to hull reference axis, from the over the hull or main float bottom (in center of gravity of the seaplane to the order to avoid excessive local shear hull longitudinal station at which the loads and bending moments at the lo- load factor is being computed to the ra- cation of water load application) using dius of gyration in pitch of the sea- pressures not less than those pre- plane, the hull reference axis being a scribed in § 25.533(b). straight line, in the plane of sym- (c) For twin float seaplanes, each metry, tangential to the keel at the float must be treated as an equivalent main step. hull on a fictitious seaplane with a (c) For a twin float seaplane, because weight equal to one-half the weight of of the effect of flexibility of the attach- the twin float seaplane. ment of the floats to the seaplane, the (d) Except in the takeoff condition of factor K1 may be reduced at the bow § 25.531, the aerodynamic lift on the and stern to 0.8 of the value shown in seaplane during the impact is assumed figure 2 of appendix B. This reduction 2 to be ⁄3 of the weight of the seaplane. applies only to the design of the carry- through and seaplane structure. § 25.527 Hull and main float load factors. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) Water reaction load factors n amended by Amdt. 25–23, 35 FR 5673, Apr. 8, W 1970] must be computed in the following manner: § 25.529 Hull and main float landing (1) For the step landing case conditions. (a) Symmetrical step, bow, and stern CV1 2 n = S0 landing. For symmetrical step, bow, w ⎛ 2 ⎞ 1 and stern landings, the limit water re- 3 β 3 ⎝Tan ⎠ W action load factors are those computed under § 25.527. In addition— (2) For the bow and stern landing (1) For symmetrical step landings, cases the resultant water load must be applied at the keel, through the center of CV gravity, and must be directed per- = 1 S02 × K1 nw 1 2 pendicularly to the keel line; ⎛ 2 ⎞ 3 3 (2) For symmetrical bow landings, Tan 3 β W + 2 ⎝ ⎠ ()1 rx the resultant water load must be applied at the keel, one-fifth of the longi- (b) The following values are used: tudinal distance from the bow to the (1) nW = water reaction load factor step, and must be directed perpendicu- (that is, the water reaction divided by larly to the keel line; and seaplane weight). (3) For symmetrical stern landings, (2) C1 = empirical seaplane operations the resultant water load must be ap- factor equal to 0.012 (except that this plied at the keel, at a point 85 percent factor may not be less than that nec- of the longitudinal distance from the essary to obtain the minimum value of step to the stern post, and must be di- step load factor of 2.33). rected perpendicularly to the keel line.

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(b) Unsymmetrical landing for hull and W = design water takeoff weight in pounds. single float seaplanes. Unsymmetrical [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as step, bow, and stern landing conditions amended by Amdt. 25–23, 35 FR 5673, Apr. 8, must be investigated. In addition— 1970] (1) The loading for each condition consists of an upward component and a § 25.533 Hull and main float bottom pressures. side component equal, respectively, to 0.75 and 0.25 tan b times the resultant (a) General. The hull and main float load in the corresponding symmetrical structure, including frames and bulk- landing condition; and heads, stringers, and bottom plating, (2) The point of application and di- must be designed under this section. rection of the upward component of the (b) Local pressures. For the design of the bottom plating and stringers and load is the same as that in the sym- their attachments to the supporting metrical condition, and the point of ap- structure, the following pressure dis- plication of the side component is at tributions must be applied: the same longitudinal station as the (1) For an unflared bottom, the pres- upward component but is directed insure at the chine is 0.75 times the pres- ward perpendicularly to the plane of sure at the keel, and the pressures be- symmetry at a point midway between tween the keel and chine vary linearly, the keel and chine lines. in accordance with figure 3 of appendix (c) Unsymmetrical landing; twin float B. The pressure at the keel (psi) is seaplanes. The unsymmetrical loading computed as follows: consists of an upward load at the step of each float of 0.75 and a side load of KV 2 S12 0.25 tan b at one float times the step PC=× k 2 β landing load reached under § 25.527. The tan k side load is directed inboard, per- where— pendicularly to the plane of symmetry Pk = pressure (p.s.i.) at the keel; midway between the keel and chine C2 = 0.00213; lines of the float, at the same longitu- K2 = hull station weighing factor, in accord- dinal station as the upward load. ance with figure 2 of appendix B; VS1 = seaplane stalling speed (Knots) at the § 25.531 Hull and main float takeoff design water takeoff weight with flaps condition. extended in the appropriate takeoff position; and For the wing and its attachment to bK = angle of dead rise at keel, in accordance the hull or main float— with figure 1 of appendix B. (a) The aerodynamic wing lift is as- (2) For a flared bottom, the pressure sumed to be zero; and at the beginning of the flare is the (b) A downward inertia load, cor- same as that for an unflared bottom, responding to a load factor computed and the pressure between the chine and from the following formula, must be the beginning of the flare varies lin- applied: early, in accordance with figure 3 of appendix B. The pressure distribution is CV2 the same as that prescribed in para- n = TO S1 ⎛ 2 ⎞ 1 graph (b)(1) of this section for an tan 3 β W 3 unflared bottom except that the pres- ⎝ ⎠ sure at the chine is computed as fol- where— lows: n = inertia load factor; KV2 2 CTO = empirical seaplane operations factor PC=× S1 equal to 0.004; ch 3 tan β VS1 = seaplane stalling speed (knots) at the design takeoff weight with the flaps ex- where—

tended in the appropriate takeoff posi- Pch = pressure (p.s.i.) at the chine; tion; C3 = 0.0016; b = angle of dead rise at the main step (de- K2 = hull station weighing factor, in accord- grees); and ance with figure 2 of appendix B;

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VS1 = seaplane stalling speed at the design must be designed for the conditions water takeoff weight with flaps extended prescribed in this section. In the cases in the appropriate takeoff position; and specified in paragraphs (b) through (e) = angle of dead rise at appropriate station. b of this section, the prescribed water The area over which these pressures loads may be distributed over the float are applied must simulate pressures oc- bottom to avoid excessive local loads, curring during high localized impacts using bottom pressures not less than on the hull or float, but need not ex- those prescribed in paragraph (g) of tend over an area that would induce this section. critical stresses in the frames or in the (b) Step loading. The resultant water overall structure. load must be applied in the plane of (c) Distributed pressures. For the de- symmetry of the float at a point three- sign of the frames, keel, and chine fourths of the distance from the bow to structure, the following pressure dis- the step and must be perpendicular to tributions apply: the keel. The resultant limit load is (1) Symmetrical pressures are com- computed as follows, except that the puted as follows: value of L need not exceed three times the weight of the displaced water when KV the float is completely submerged: =×2 S02 PC4 tan β 2 3 CV5 2 W where— S0 P = pressure (p.s.i.); L = 2 2 C4 = 0.078 C1 (with C1 computed under 3 β + 2 3 § 25.527); tan sy()1 r K = hull station weighing factor, deter- 2 where— mined in accordance with figure 2 of appendix B; L = limit load (lbs.); C = 0.0053; VS0 = seaplane stalling speed (Knots) with 5 landing flaps extended in the appropriate VS0 = seaplane stalling speed (knots) with position and with no slipstream effect; landing flaps extended in the appropriate and position and with no slipstream effect; W = seaplane design landing weight in VS0 = seaplane stalling speed with landing flaps extended in the appropriate posi- pounds; 3 tion and with no slipstream effect; and b bS = angle of dead rise at a station ⁄4 of the = angle of dead rise at appropriate sta- distance from the bow to the step, but tion. need not be less than 15 degrees; and ry = ratio of the lateral distance between the (2) The unsymmetrical pressure dis- center of gravity and the plane of sym- tribution consists of the pressures pre- metry of the float to the radius of gyra- scribed in paragraph (c)(1) of this section in roll. tion on one side of the hull or main (c) Bow loading. The resultant limit float centerline and one-half of that load must be applied in the plane of pressure on the other side of the hull or symmetry of the float at a point one- main float centerline, in accordance fourth of the distance from the bow to with figure 3 of appendix B. the step and must be perpendicular to These pressures are uniform and must the tangent to the keel line at that be applied simultaneously over the en- point. The magnitude of the resultant tire hull or main float bottom. The load is that specified in paragraph (b) loads obtained must be carried into the of this section. sidewall structure of the hull proper, (d) Unsymmetrical step loading. The re- but need not be transmitted in a fore sultant water load consists of a compo- and aft direction as shear and bending nent equal to 0.75 times the load speci- loads. fied in paragraph (a) of this section and [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as a side component equal to 3.25 tan b amended by Amdt. 25–23, 35 FR 5673, Apr. 8, times the load specified in paragraph 1970] (b) of this section. The side load must be applied perpendicularly to the plane § 25.535 Auxiliary float loads. of symmetry of the float at a point (a) General. Auxiliary floats and their midway between the keel and the attachments and supporting structures chine.

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(e) Unsymmetrical bow loading. The re- EMERGENCY LANDING CONDITIONS sultant water load consists of a component equal to 0.75 times the load speci- § 25.561 General. fied in paragraph (b) of this section and (a) The airplane, although it may be a side component equal to 0.25 tan b damaged in emergency landing condi- times the load specified in paragraph tions on land or water, must be de- (c) of this section. The side load must signed as prescribed in this section to be applied perpendicularly to the plane protect each occupant under those con- of symmetry at a point midway be- ditions. tween the keel and the chine. (b) The structure must be designed to (f) Immersed float condition. The re- give each occupant every reasonable chance of escaping serious injury in a sultant load must be applied at the minor crash landing when— centroid of the cross section of the (1) Proper use is made of seats, belts, float at a point one-third of the dis- and all other safety design provisions; tance from the bow to the step. The (2) The wheels are retracted (where limit load components are as follows: applicable); and = (3) The occupant experiences the fol- vertical ρgV lowing ultimate inertia forces acting separately relative to the surrounding 2 2 ρ 3 ⎛ ⎞ aft = CV KV structure: x 2 ⎝ S ⎠ (i) Upward, 3.0g 0 (ii) Forward, 9.0g 2 2 (iii) Sideward, 3.0g on the airframe; ρ 3 ⎛ ⎞ side = CV KV and 4.0g on the seats and their attach- y ⎝ S ⎠ 2 0 ments. where— (iv) Downward, 6.0g r = mass density of water (slugs/ft.2); (v) Rearward, 1.5g V = volume of float (ft.2); (c) For equipment, cargo in the passenger compartments and any other Cx = coefficient of drag force, equal to 0.133; large masses, the following apply: Cy = coefficient of side force, equal to 0.106; K = 0.8, except that lower values may be used (1) Except as provided in paragraph if it is shown that the floats are incapa- (c)(2) of this section, these items must

ble of submerging at a speed of 0.8 VS0 in be positioned so that if they break normal operations; loose they will be unlikely to:

VS0 = seaplane stalling speed (knots) with (i) Cause direct injury to occupants; landing flaps extended in the appropriate (ii) Penetrate fuel tanks or lines or position and with no slipstream effect; cause fire or explosion hazard by dam- and age to adjacent systems; or g = acceleration due to gravity (ft./sec.2). (iii) Nullify any of the escape facili- (g) Float bottom pressures. The float ties provided for use after an emer- bottom pressures must be established gency landing. under § 25.533, except that the value of (2) When such positioning is not practical (e.g. fuselage mounted engines or K in the formulae may be taken as 1.0. 2 auxiliary power units) each such item The angle of dead rise to be used in de- of mass shall be restrained under all termining the float bottom pressures is loads up to those specified in paragraph set forth in paragraph (b) of this sec- (b)(3) of this section. The local attach- tion. ments for these items should be de- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as signed to withstand 1.33 times the spec- amended by Amdt. 25–23, 35 FR 5673, Apr. 8, ified loads if these items are subject to 1970] severe wear and tear through frequent removal (e.g. quick change interior § 25.537 Seawing loads. items). Seawing design loads must be based (d) Seats and items of mass (and on applicable test data. their supporting structure) must not deform under any loads up to those specified in paragraph (b)(3) of this section in any manner that would impede

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subsequent rapid evacuation of occu- grees either right or left, whichever pants. would cause the greatest likelihood of [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the upper torso restraint system amended by Amdt. 25–23, 35 FR 5673, Apr. 8, (where installed) moving off the occu- 1970; Amdt. 25–64, 53 FR 17646, May 17, 1988; pant’s shoulder, and with the wings Amdt. 25–91, 62 FR 40706, July 29, 1997] level. Peak floor deceleration must occur in not more than 0.09 seconds § 25.562 Emergency landing dynamic after impact and must reach a min- conditions. imum of 16g. Where floor rails or floor (a) The seat and restraint system in fittings are used to attach the seating the airplane must be designed as pre- devices to the test fixture, the rails or scribed in this section to protect each fittings must be misaligned with re- occupant during an emergency landing spect to the adjacent set of rails or fit- condition when— tings by at least 10 degrees vertically (1) Proper use is made of seats, safety (i.e., out of Parallel) with one rolled 10 belts, and shoulder harnesses provided degrees. for in the design; and (c) The following performance meas- (2) The occupant is exposed to loads ures must not be exceeded during the resulting from the conditions pre- dynamic tests conducted in accordance scribed in this section. with paragraph (b) of this section: (b) Each seat type design approved (1) Where upper torso straps are used for crew or passenger occupancy during for crewmembers, tension loads in indi- takeoff and landing must successfully vidual straps must not exceed 1,750 complete dynamic tests or be dem- pounds. If dual straps are used for re- onstrated by rational analysis based on straining the upper torso, the total dynamic tests of a similar type seat, in strap tension loads must not exceed accordance with each of the following 2,000 pounds. emergency landing conditions. The (2) The maximum compressive load tests must be conducted with an occu- measured between the pelvis and the pant simulated by a 170-pound anthropomorphic test dummy, as de- lumbar column of the anthropomorphic fined by 49 CFR Part 572, Subpart B, or dummy must not exceed 1,500 pounds. its equivalent, sitting in the normal (3) The upper torso restraint straps upright position. (where installed) must remain on the (1) A change in downward vertical ve- occupant’s shoulder during the impact. locity (D v) of not less than 35 feet per (4) The lap safety belt must remain second, with the airplane’s longitu- on the occupant’s pelvis during the im- dinal axis canted downward 30 degrees pact. with respect to the horizontal plane (5) Each occupant must be protected and with the wings level. Peak floor de- from serious head injury under the con- celeration must occur in not more than ditions prescribed in paragraph (b) of 0.08 seconds after impact and must this section. Where head contact with reach a minimum of 14g. seats or other structure can occur, pro- (2) A change in forward longitudinal tection must be provided so that the velocity (D v) of not less than 44 feet head impact does not exceed a Head In- per second, with the airplane’s longitu- jury Criterion (HIC) of 1,000 units. The dinal axis horizontal and yawed 10 de- level of HIC is defined by the equation:

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

Where: a(t) is the total acceleration vs. time curve for the head strike, and where t1 is the initial integration time, t2 is the final integration time, and

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(t) is in seconds, and (a) is in units of gravity catastrophic failure of the airplane; (g). and (6) Where leg injuries may result (iii) An analysis, supported by test from contact with seats or other struc- evidence, of the principal structural ture, protection must be provided to elements and detail design points iden- prevent axially compressive loads ex- tified in paragraph (a)(1)(ii) of this sec- ceeding 2,250 pounds in each femur. tion. (7) The seat must remain attached at (2) The service history of airplanes of all points of attachment, although the similar structural design, taking due structure may have yielded. account of differences in operating con- (8) Seats must not yield under the ditions and procedures, may be used in tests specified in paragraphs (b)(1) and the evaluations required by this sec- (b)(2) of this section to the extent they tion. would impede rapid evacuation of the (3) Based on the evaluations required airplane occupants. by this section, inspections or other procedures must be established, as nec- [Amdt. 25–64, 53 FR 17646, May 17, 1988] essary, to prevent catastrophic failure, and must be included in the Airworthi- § 25.563 Structural ditching provi- ness Limitations section of the In- sions. structions for Continued Airworthiness Structural strength considerations of required by § 25.1529. The limit of valid- ditching provisions must be in accord- ity of the engineering data that sup- ance with § 25.801(e). ports the structural maintenance program (hereafter referred to as LOV), FATIGUE EVALUATION stated as a number of total accumulated flight cycles or flight hours or § 25.571 Damage—tolerance and fa- both, established by this section must tigue evaluation of structure. also be included in the Airworthiness (a) General. An evaluation of the Limitations section of the Instructions strength, detail design, and fabrication for Continued Airworthiness required must show that catastrophic failure by § 25.1529. Inspection thresholds for due to fatigue, corrosion, manufac- the following types of structure must turing defects, or accidental damage, be established based on crack growth will be avoided throughout the oper- analyses and/or tests, assuming the ational life of the airplane. This eval- structure contains an initial flaw of uation must be conducted in accord- the maximum probable size that could ance with the provisions of paragraphs exist as a result of manufacturing or (b) and (e) of this section, except as service-induced damage: specified in paragraph (c) of this sec- (i) Single load path structure, and tion, for each part of the structure that (ii) Multiple load path ‘‘fail-safe’’ could contribute to a catastrophic fail- structure and crack arrest ‘‘fail-safe’’ ure (such as wing, empennage, control structure, where it cannot be dem- surfaces and their systems, the fuse- onstrated that load path failure, par- lage, engine mounting, landing gear, tial failure, or crack arrest will be de- and their related primary attach- tected and repaired during normal ments). For turbojet powered air- maintenance, inspection, or operation planes, those parts that could con- of an airplane prior to failure of the re- tribute to a catastrophic failure must maining structure. also be evaluated under paragraph (d) (b) Damage-tolerance evaluation. The of this section. In addition, the fol- evaluation must include a determina- lowing apply: tion of the probable locations and (1) Each evaluation required by this modes of damage due to fatigue, corro- section must include— sion, or accidental damage. Repeated (i) The typical loading spectra, tem- load and static analyses supported by peratures, and humidities expected in test evidence and (if available) service service; experience must also be incorporated (ii) The identification of principal in the evaluation. Special consider- structural elements and detail design ation for widespread fatigue damage points, the failure of which could cause must be included where the design is

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such that this type of damage could pressures during 1 g level flight) multi- occur. An LOV must be established plied by a factor of 1.15, omitting other that corresponds to the period of time, loads. stated as a number of total accumu- (6) For landing gear and directly-af- lated flight cycles or flight hours or fected airframe structure, the limit both, during which it is demonstrated ground loading conditions specified in that widespread fatigue damage will §§ 25.473, 25.491, and 25.493. not occur in the airplane structure. If significant changes in structural This demonstration must be by full- stiffness or geometry, or both, follow scale fatigue test evidence. The type from a structural failure, or partial certificate may be issued prior to com- failure, the effect on damage tolerance pletion of full-scale fatigue testing, must be further investigated. provided the Administrator has ap- (c) Fatigue (safe-life) evaluation. Com- proved a plan for completing the re- pliance with the damage-tolerance required tests. In that case, the Air- quirements of paragraph (b) of this sec- worthiness Limitations section of the tion is not required if the applicant es- Instructions for Continued Airworthi- tablishes that their application for par- ness required by § 25.1529 must specify ticular structure is impractical. This that no airplane may be operated be- structure must be shown by analysis, 1 yond a number of cycles equal to ⁄2 the supported by test evidence, to be able number of cycles accumulated on the to withstand the repeated loads of vari- fatigue test article, until such testing able magnitude expected during its is completed. The extent of damage for service life without detectable cracks. residual strength evaluation at any Appropriate safe-life scatter factors time within the operational life of the must be applied. airplane must be consistent with the (d) Sonic fatigue strength. It must be initial detectability and subsequent shown by analysis, supported by test growth under repeated loads. The resid- evidence, or by the service history of ual strength evaluation must show airplanes of similar structural design that the remaining structure is able to and sonic excitation environment, withstand loads (considered as static that— ultimate loads) corresponding to the (1) Sonic fatigue cracks are not prob- following conditions: able in any part of the flight structure (1) The limit symmetrical maneu- subject to sonic excitation; or vering conditions specified in § 25.337 at (2) Catastrophic failure caused by all speeds up to V and in § 25.345. c sonic cracks is not probable assuming (2) The limit gust conditions speci- that the loads prescribed in paragraph fied in § 25.341 at the specified speeds up (b) of this section are applied to all to V and in § 25.345. C areas affected by those cracks. (3) The limit rolling conditions speci- (e) fied in § 25.349 and the limit unsymmet- Damage-tolerance (discrete source) The airplane must be capa- rical conditions specified in §§ 25.367 evaluation. ble of successfully completing a flight and 25.427 (a) through (c), at speeds up during which likely structural damage to V . C occurs as a result of— (4) The limit yaw maneuvering conditions specified in § 25.351(a) at the spec- (1) Impact with a 4-pound bird when the velocity of the airplane relative to ified speeds up to VC. (5) For pressurized cabins, the fol- the bird along the airplane’s flight lowing conditions: path is equal to Vc at sea level or 0.85Vc (i) The normal operating differential at 8,000 feet, whichever is more critical; pressure combined with the expected (2) Uncontained fan blade impact; external aerodynamic pressures applied (3) Uncontained engine failure; or 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

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need not be considered. Corrective ac- (a) Be established on the basis of ex- tion to be taken by the pilot following perience or tests; the incident, such as limiting maneu- (b) Conform to approved specifica- vers, avoiding turbulence, and reducing tions (such as industry or military speed, must be considered. If signifi- specifications, or Technical Standard cant changes in structural stiffness or Orders) that ensure their having the geometry, or both, follow from a struc- strength and other properties assumed tural failure or partial failure, the ef- in the design data; and fect on damage tolerance must be fur- (c) Take into account the effects of ther investigated. environmental conditions, such as tem- [Amdt. 25–45, 43 FR 46242, Oct. 5, 1978, as perature and humidity, expected in amended by Amdt. 25–54, 45 FR 60173, Sept. service. 11, 1980; Amdt. 25–72, 55 FR 29776, July 20, 1990; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as Amdt. 25–96, 63 FR 15714, Mar. 31, 1998; 63 FR amended by Amdt. 25–38, 41 FR 55466, Dec. 20, 23338, Apr. 28, 1998; Amdt. 25–132, 75 FR 69781, 1976; Amdt. 25–46, 43 FR 50595, Oct. 30, 1978] Nov. 15, 2010] § 25.605 Fabrication methods. LIGHTNING PROTECTION (a) The methods of fabrication used must produce a consistently sound § 25.581 Lightning protection. structure. If a fabrication process (such (a) The airplane must be protected as gluing, spot welding, or heat treat- against catastrophic effects from lighting) requires close control to reach this ning. objective, the process must be per- (b) For metallic components, compli- formed under an approved process spec- ance with paragraph (a) of this section ification. may be shown by— (b) Each new aircraft fabrication (1) Bonding the components properly method must be substantiated by a to the airframe; or test program. (2) Designing the components so that a strike will not endanger the airplane. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) For nonmetallic components, amended by Amdt. 25–46, 43 FR 50595, Oct. 30, compliance with paragraph (a) of this 1978] section may be shown by— (1) Designing the components to min- § 25.607 Fasteners. imize the effect of a strike; or (a) Each removable bolt, screw, nut, (2) Incorporating acceptable means of pin, or other removable fastener must diverting the resulting electrical cur- incorporate two separate locking de- rent so as not to endanger the airplane. vices if— (1) Its loss could preclude continued [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] flight and landing within the design limitations of the airplane using nor- Subpart D—Design and mal pilot skill and strength; or Construction (2) Its loss could result in reduction in pitch, yaw, or roll control capability GENERAL or response below that required by § 25.601 General. Subpart B of this chapter. (b) The fasteners specified in para- The airplane may not have design graph (a) of this section and their lock- features or details that experience has ing devices may not be adversely af- shown to be hazardous or unreliable. fected by the environmental conditions The suitability of each questionable associated with the particular installa- design detail and part must be estab- tion. lished by tests. (c) No self-locking nut may be used § 25.603 Materials. on any bolt subject to rotation in operation unless a nonfriction locking de- The suitability and durability of ma- vice is used in addition to the self-lock- terials used for parts, the failure of ing device. which could adversely affect safety, must— [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970]

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§ 25.609 Protection of structure. (2) For redundant structure, in which the failure of individual elements Each part of the structure must— would result in applied loads being (a) Be suitably protected against de- safely distributed to other load car- terioration or loss of strength in serv- rying members, 90 percent probability ice due to any cause, including— with 95 percent confidence. (1) Weathering; (c) The effects of environmental con- (2) Corrosion; and ditions, such as temperature and mois- (3) Abrasion; and ture, on material design values used in (b) Have provisions for ventilation an essential component or structure and drainage where necessary for pro- must be considered where these effects tection. are significant within the airplane op- § 25.611 Accessibility provisions. erating envelope. (d) [Reserved] (a)Means must be provided to allow (e) Greater material design values inspection (including inspection of may be used if a ‘‘premium selection’’ principal structural elements and con- of the material is made in which a trol systems), replacement of parts specimen of each individual item is normally requiring replacement, ad- tested before use to determine that the justment, and lubrication as necessary actual strength properties of that par- for continued airworthiness. The in- ticular item will equal or exceed those spection means for each item must be used in design. practicable for the inspection interval (f) Other material design values may for the item. Nondestructive inspection be used if approved by the Adminis- aids may be used to inspect structural trator. elements where it is impracticable to provide means for direct visual inspec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tion if it is shown that the inspection amended by Amdt. 25–46, 43 FR 50595, Oct. 30, is effective and the inspection proce- 1978; Amdt. 25–72, 55 FR 29776, July 20, 1990; dures are specified in the maintenance Amdt. 25–112, 68 FR 46431, Aug. 5, 2003] manual required by § 25.1529. § 25.619 Special factors. (b) EWIS must meet the accessibility requirements of § 25.1719. The factor of safety prescribed in § 25.303 must be multiplied by the high- [Amdt. 25–23, 35 FR 5674, Apr. 8, 1970, as est pertinent special factor of safety amended by Amdt. 25–123, 72 FR 63404, Nov. 8, prescribed in §§ 25.621 through 25.625 for 2007] each part of the structure whose § 25.613 Material strength properties strength is— and material design values. (a) Uncertain; (a) Material strength properties must (b) Likely to deteriorate in service be based on enough tests of material before normal replacement; or meeting approved specifications to es- (c) Subject to appreciable variability tablish design values on a statistical because of uncertainties in manufac- basis. turing processes or inspection methods. (b) Material design values must be [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as chosen to minimize the probability of amended by Amdt. 25–23, 35 FR 5674, Apr. 8, structural failures due to material var- 1970] iability. Except as provided in paragraphs (e) and (f) of this section, com- § 25.621 Casting factors. pliance must be shown by selecting ma- (a) General. For castings used in terial design values which assure mate- structural applications, the factors, rial strength with the following prob- tests, and inspections specified in para- ability: graphs (b) through (d) of this section (1) Where applied loads are eventu- must be applied in addition to those ally distributed through a single mem- necessary to establish foundry quality ber within an assembly, the failure of control. The inspections must meet ap- which would result in loss of structural proved specifications. Paragraphs (c) integrity of the component, 99 percent and (d) of this section apply to any probability with 95 percent confidence. structural castings, except castings

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that are pressure tested as parts of hy- (iii) One casting undergoes a static draulic or other fluid systems and do test and is shown to meet the strength not support structural loads. and deformation requirements of (b) Bearing stresses and surfaces. The § 25.305(a) and (b). casting factors specified in paragraphs (2) A casting factor of 1.25 or greater (c) and (d) of this section— may be used, provided that— (1) Need not exceed 1.25 with respect (i) Each casting receives: to bearing stresses regardless of the (A) Inspection of 100 percent of its method of inspection used; and surface, using visual inspection and liq- (2) Need not be used with respect to uid penetrant or equivalent inspection the bearing surfaces of a part whose methods; and bearing factor is larger than the appli- (B) Inspection of structurally signifi- cable casting factor. cant internal areas and areas where de- (c) Critical castings. Each casting fects are likely to occur, using radio- whose failure could preclude continued graphic or equivalent inspection meth- safe flight and landing of the airplane ods. or could result in serious injury to oc- (ii) Three castings undergo static cupants is a critical casting. Each crit- tests and are shown to meet: ical casting must have a factor associ- (A) The strength requirements of ated with it for showing compliance § 25.305(b) at an ultimate load cor- with strength and deformation require- responding to a casting factor of 1.25; ments of § 25.305, and must comply with and the following criteria associated with (B) The deformation requirements of that factor: § 25.305(a) at a load of 1.15 times the (1) A casting factor of 1.0 or greater limit load. may be used, provided that— (3) A casting factor of 1.50 or greater (i) It is demonstrated, in the form of may be used, provided that— process qualification, proof of product, (i) Each casting receives: and process monitoring that, for each (A) Inspection of 100 percent of its casting design and part number, the surface, using visual inspection and liq- castings produced by each foundry and uid penetrant or equivalent inspection process combination have coefficients methods; and of variation of the material properties (B) Inspection of structurally signifi- that are equivalent to those of wrought cant internal areas and areas where de- alloy products of similar composition. fects are likely to occur, using radio- Process monitoring must include test- graphic or equivalent inspection meth- ing of coupons cut from the prolonga- ods. tions of each casting (or each set of (ii) One casting undergoes a static castings, if produced from a single pour test and is shown to meet: into a single mold in a runner system) (A) The strength requirements of and, on a sampling basis, coupons cut § 25.305(b) at an ultimate load cor- from critical areas of production cast- responding to a casting factor of 1.50; ings. The acceptance criteria for the and process monitoring inspections and (B) The deformation requirements of tests must be established and included § 25.305(a) at a load of 1.15 times the in the process specifications to ensure limit load. the properties of the production cast- (d) Non-critical castings. For each ings are controlled to within levels casting other than critical castings, as used in design. specified in paragraph (c) of this sec- (ii) Each casting receives: tion, the following apply: (A) Inspection of 100 percent of its (1) A casting factor of 1.0 or greater surface, using visual inspection and liq- may be used, provided that the require- uid penetrant or equivalent inspection ments of (c)(1) of this section are met, methods; and or all of the following conditions are (B) Inspection of structurally signifi- met: cant internal areas and areas where de- (i) Castings are manufactured to ap- fects are likely to occur, using radio- proved specifications that specify the graphic or equivalent inspection meth- minimum mechanical properties of the ods. material in the casting and provides

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for demonstration of these properties (b) No bearing factor need be used for by testing of coupons cut from the a part for which any larger special fac- castings on a sampling basis. tor is prescribed. (ii) Each casting receives: (A) Inspection of 100 percent of its § 25.625 Fitting factors. surface, using visual inspection and liq- For each fitting (a part or terminal uid penetrant or equivalent inspection used to join one structural member to methods; and another), the following apply: (B) Inspection of structurally signifi- (a) For each fitting whose strength is cant internal areas and areas where de- not proven by limit and ultimate load fects are likely to occur, using radio- tests in which actual stress conditions graphic or equivalent inspection meth- are simulated in the fitting and sur- ods. rounding structures, a fitting factor of (iii) Three sample castings undergo at least 1.15 must be applied to each static tests and are shown to meet the part of— strength and deformation requirements (1) The fitting; of § 25.305(a) and (b). (2) The means of attachment; and (2) A casting factor of 1.25 or greater (3) The bearing on the joined mem- may be used, provided that each cast- bers. ing receives: (b) No fitting factor need be used— (i) Inspection of 100 percent of its surface, using visual inspection and liquid (1) For joints made under approved penetrant or equivalent inspection practices and based on comprehensive methods; and test data (such as continuous joints in (ii) Inspection of structurally signifi- metal plating, welded joints, and scarf cant internal areas and areas where de- joints in wood); or fects are likely to occur, using radio- (2) With respect to any bearing sur- graphic or equivalent inspection meth- face for which a larger special factor is ods. used. (3) A casting factor of 1.5 or greater (c) For each integral fitting, the part may be used, provided that each cast- must be treated as a fitting up to the ing receives inspection of 100 percent of point at which the section properties its surface using visual inspection and become typical of the member. liquid penetrant or equivalent inspec- (d) For each seat, berth, safety belt, tion methods. and harness, the fitting factor specified (4) A casting factor of 2.0 or greater in § 25.785(f)(3) applies. may be used, provided that each cast- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ing receives inspection of 100 percent of amended by Amdt. 25–23, 35 FR 5674, Apr. 8, its surface using visual inspection 1970; Amdt. 25–72, 55 FR 29776, July 20, 1990] methods. (5) The number of castings per pro- § 25.629 Aeroelastic stability require- duction batch to be inspected by non- ments. visual methods in accordance with (a) General. The aeroelastic stability paragraphs (d)(2) and (3) of this section evaluations required under this section may be reduced when an approved qual- include flutter, divergence, control re- ity control procedure is established. versal and any undue loss of stability [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as and control as a result of structural de- amended by Amdt. 25–139, 79 FR 59429, Oct. 2, formation. The aeroelastic evaluation 2014] 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- clearance (free fit), and that is subject nel tests, ground vibration tests, flight to pounding or vibration, must have a tests, or other means found necessary bearing factor large enough to provide by the Administrator. for the effects of normal relative mo- (b) Aeroelastic stability envelopes. The tion. airplane must be designed to be free

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from aeroelastic instability for all con- (4) Failure of any single element of figurations and design conditions with- the structure supporting any engine, in the aeroelastic stability envelopes independently mounted propeller shaft, as follows: large auxiliary power unit, or large ex- (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 gyroscopic forces resulting from the most must exist at all speeds up to VD/MD and, there must be no large and rapid adverse combination of feathered propellers or other rotating devices capa- reduction in stability as VD/MD is approached. The enlarged envelope may ble of significant dynamic forces. In addition, the effect of a single feath- be limited to Mach 1.0 when MD is less 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 forces rotating at the highest likely (i) The V /M envelope determined by D D overspeed. § 25.335(b); or, (8) Any damage or failure condition, (ii) An altitude-airspeed envelope de- required or selected for investigation fined by a 15 percent increase in equiv- by § 25.571. The single structural fail- alent airspeed above VC at constant al- ures described in paragraphs (d)(4) and titude, from sea level to the altitude of (d)(5) of this section need not be consid- the intersection of 1.15 VC with the ex- ered in showing compliance with this tension of the constant cruise Mach section if; number line, MC, then a linear vari- (i) The structural element could not + .05 ation in equivalent airspeed to MC fail due to discrete source damage re- at the altitude of the lowest V /M C C sulting from the conditions described intersection; then, at higher altitudes, in § 25.571(e), and up to the maximum flight altitude, the (ii) A damage tolerance investigation boundary defined by a .05 Mach in- in accordance with § 25.571(b) shows crease in M at constant altitude. C that the maximum extent of damage (c) Balance weights. If concentrated assumed for the purpose of residual balance weights are used, their effec- strength evaluation does not involve tiveness and strength, including sup- complete failure of the structural ele- porting structure, must be substan- ment. tiated. (9) Any damage, failure, or malfunc- (d) Failures, malfunctions, and adverse tion considered under §§ 25.631, 25.671, conditions. The failures, malfunctions, 25.672, and 25.1309. and adverse conditions which must be (10) Any other combination of fail- considered in showing compliance with ures, malfunctions, or adverse condi- this section are: tions not shown to be extremely im- (1) Any critical fuel loading condi- probable. tions, not shown to be extremely im- (e) Flight flutter testing. Full scale probable, which may result from mis- flight flutter tests at speeds up to VDF/ management of fuel. MDF must be conducted for new type (2) Any single failure in any flutter designs and for modifications to a type damper system. design unless the modifications have (3) For airplanes not approved for op- been shown to have an insignificant ef- eration in icing conditions, the max- fect on the aeroelastic stability. These imum likely ice accumulation expected tests must demonstrate that the air- as a result of an inadvertent encounter. plane has a proper margin of damping

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

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electrical or hydraulic system failures, (1) The airplane is safely controllable or any single failure in combination when the failure or malfunction occurs 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 failure in the stability augmentation [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as system or in any other automatic or amended by Amdt. 25–38, 41 FR 55466, Dec. 20, power-operated system which could re- 1976] sult in an unsafe condition if the pilot § 25.677 Trim systems. were not aware of the failure. Warning systems must not activate the control (a) Trim controls must be designed to systems. prevent inadvertent or abrupt oper- (b) The design of the stability aug- ation and to operate in the plane, and mentation system or of any other auto- with the sense of motion, of the air- matic or power-operated system must plane. permit initial counteraction of failures (b) There must be means adjacent to of the type specified in § 25.671(c) with- the trim control to indicate the direc- out requiring exceptional pilot skill or tion of the control movement relative strength, by either the deactivation of to the airplane motion. In addition, the system, or a failed portion thereof, there must be clearly visible means to or by overriding the failure by move- indicate the position of the trim device ment of the flight controls in the nor- with respect to the range of adjust- mal sense. ment. The indicator must be clearly (c) It must be shown that after any marked with the range within which it single failure of the stability aug- has been demonstrated that takeoff is mentation system or any other auto- safe for all center of gravity positions matic or power-operated system— approved for takeoff.

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(c) Trim control systems must be de- in normal operation, the system is free signed to prevent creeping in flight. from— Trim tab controls must be irreversible (1) Jamming; unless the tab is appropriately bal- (2) Excessive friction; and anced and shown to be free from flut- (3) Excessive deflection. ter. (b) It must be shown by analysis and, (d) If an irreversible tab control sys- where necessary, by tests, that in the tem is used, the part from the tab to presence of deflections of the airplane the attachment of the irreversible unit structure due to the separate applica- to the airplane structure must consist tion of pitch, roll, and yaw limit ma- of a rigid connection. neuver loads, the control system, when loaded to obtain these limit loads and [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5675, Apr. 8, operated within its operational range 1970; Amdt. 25–115, 69 FR 40527, July 2, 2004] of deflections, can be exercised about all control axes and remain free from— § 25.679 Control system gust locks. (1) Jamming; (a) There must be a device to prevent (2) Excessive friction; damage to the control surfaces (includ- (3) Disconnection; and ing tabs), and to the control system, (4) Any form of permanent damage. from gusts striking the airplane while (c) It must be shown that under vi- it is on the ground or water. If the de- bration loads in the normal flight and vice, when engaged, prevents normal ground operating conditions, no hazard operation of the control surfaces by the can result from interference or contact pilot, it must— with adjacent elements. (1) Automatically disengage when the [Amdt. 25–139, 79 FR 59430, Oct. 2, 2014] pilot operates the primary flight controls in a normal manner; or § 25.685 Control system details. (2) Limit the operation of the air- (a) Each detail of each control sys- plane so that the pilot receives unmis- tem must be designed and installed to takable warning at the start of takeoff. prevent jamming, chafing, and inter- (b) The device must have means to ference from cargo, passengers, loose preclude the possibility of it becoming objects, or the freezing of moisture. inadvertently engaged in flight. (b) There must be means in the cockpit to prevent the entry of foreign ob- § 25.681 Limit load static tests. jects into places where they would jam (a) Compliance with the limit load the system. requirements of this Part must be (c) There must be means to prevent shown by tests in which— the slapping of cables or tubes against (1) The direction of the test loads other parts. produces the most severe loading in the (d) Sections 25.689 and 25.693 apply to control system; and cable systems and joints. (2) Each fitting, pulley, and bracket [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as used in attaching the system to the amended by Amdt. 25–38, 41 FR 55466, Dec. 20, main structure is included. 1976] (b) Compliance must be shown (by analyses or individual load tests) with § 25.689 Cable systems. the special factor requirements for (a) Each cable, cable fitting, turn- control system joints subject to angu- buckle, splice, and pulley must be ap- lar motion. proved. In addition— (1) No cable smaller than 1⁄8 inch in § 25.683 Operation tests. diameter may be used in the aileron, (a) It must be shown by operation elevator, or rudder systems; and tests that when portions of the control (2) Each cable system must be de- system subject to pilot effort loads are signed so that there will be no haz- loaded to 80 percent of the limit load ardous change in cable tension specified for the system and the pow- throughout the range of travel under ered portions of the control system are operating conditions and temperature loaded to the maximum load expected variations.

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(b) Each kind and size of pulley must (c) The rate of motion of the surfaces correspond to the cable with which it is in response to the operation of the con- used. Pulleys and sprockets must have trol and the characteristics of the closely fitted guards to prevent the ca- automatic positioning or load limiting bles and chains from being displaced or device must give satisfactory flight fouled. Each pulley must lie in the and performance characteristics under plane passing through the cable so that steady or changing conditions of air- the cable does not rub against the pul- speed, engine power, and airplane atti- ley flange. tude. (c) Fairleads must be installed so (d) The lift device control must be that they do not cause a change in designed to retract the surfaces from cable direction of more than three de- the fully extended position, during grees. steady flight at maximum continuous (d) Clevis pins subject to load or mo- engine power at any speed below VF + tion and retained only by cotter pins 9.0 (knots). may not be used in the control system. [Amdt. 25–23, 35 FR 5675, Apr. 8, 1970, as (e) Turnbuckles must be attached to amended by Amdt. 25–46, 43 FR 50595, Oct. 30, parts having angular motion in a man- 1978; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] ner that will positively prevent binding throughout the range of travel. § 25.699 Lift and drag device indicator. (f) There must be provisions for vis- (a) There must be means to indicate ual inspection of fairleads, pulleys, ter- to the pilots the position of each lift or minals, and turnbuckles. drag device having a separate control in the cockpit to adjust its position. In § 25.693 Joints. addition, an indication of unsymmet- Control system joints (in push-pull rical operation or other malfunction in systems) that are subject to angular the lift or drag device systems must be motion, except those in ball and roller provided when such indication is nec- bearing systems, must have a special essary to enable the pilots to prevent factor of safety of not less than 3.33 or counteract an unsafe flight or with respect to the ultimate bearing ground condition, considering the ef- strength of the softest material used as fects on flight characteristics and per- a bearing. This factor may be reduced formance. to 2.0 for joints in cable control sys- (b) There must be means to indicate tems. For ball or roller bearings, the to the pilots the takeoff, en route, ap- approved ratings may not be exceeded. proach, and landing lift device positions. [Amdt. 25–72, 55 FR 29777, July 20, 1990] (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 clearly marked to identify this range (a) Each lift device control must be of extension. designed so that the pilots can place 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

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engines on one side of the plane of sym- weights, altitudes, and temperatures metry inoperative and the remaining for which certification is requested. engines at takeoff power. [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978] (c) For airplanes with flaps or slats that are not subjected to slipstream LANDING GEAR conditions, the structure must be designed for the loads imposed when the § 25.721 General. wing flaps or slats on one side are car- (a) The landing gear system must be rying the most severe load occurring in designed so that when it fails due to the prescribed symmetrical conditions overloads during takeoff and landing, and those on the other side are car- the failure mode is not likely to cause rying not more than 80 percent of that spillage of enough fuel to constitute a load. fire hazard. The overloads must be as- (d) The interconnection must be de- sumed to act in the upward and aft di- signed for the loads resulting when rections in combination with side loads interconnected flap or slat surfaces on acting inboard and outboard. In the ab- one side of the plane of symmetry are sence of a more rational analysis, the jammed and immovable while the sur- side loads must be assumed to be up to faces on the other side are free to move 20 percent of the vertical load or 20 per- and the full power of the surface actu- cent of the drag load, whichever is ating system is applied. greater. (b) The airplane must be designed to [Amdt. 25–72, 55 FR 29777, July 20, 1990] avoid any rupture leading to the spillage of enough fuel to constitute a fire § 25.703 Takeoff warning system. hazard as a result of a wheels-up land- A takeoff warning system must be in- ing on a paved runway, under the fol- stalled and must meet the following re- lowing minor crash landing conditions: quirements: (1) Impact at 5 feet-per-second (a) The system must provide to the vertical velocity, with the airplane pilots an aural warning that is auto- under control, at Maximum Design matically activated during the initial Landing Weight— portion of the takeoff roll if the air- (i) With the landing gear fully retracted; and plane is in a configuration, including (ii) With any one or more landing any of the following, that would not gear legs not extended. allow a safe takeoff: (2) Sliding on the ground, with— (1) The wing flaps or leading edge de- (i) The landing gear fully retracted vices are not within the approved range and with up to a 20° yaw angle; and of takeoff positions. (ii) Any one or more landing gear (2) Wing spoilers (except lateral con- legs not extended and with 0° yaw trol spoilers meeting the requirements angle. of § 25.671), speed brakes, or longitu- (c) For configurations where the en- dinal trim devices are in a position gine nacelle is likely to come into con- that would not allow a safe takeoff. tact with the ground, the engine pylon (b) The warning required by para- or engine mounting must be designed graph (a) of this section must continue so that when it fails due to overloads until— (assuming the overloads to act pre- (1) The configuration is changed to dominantly in the upward direction allow a safe takeoff; and separately, predominantly in the (2) Action is taken by the pilot to aft direction), the failure mode is not terminate the takeoff roll; likely to cause the spillage of enough (3) The airplane is rotated for take- fuel to constitute a fire hazard. off; or [Amdt. 25–139, 79 FR 59430, Oct. 2, 2014] (4) The warning is manually deactivated by the pilot. § 25.723 Shock absorption tests. (c) The means used to activate the (a) The analytical representation of system must function properly the landing gear dynamic characteris- throughout the ranges of takeoff tics that is used in determining the

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landing loads must be validated by en- flaps in the approach position at design ergy absorption tests. A range of tests landing weight), and must be conducted to ensure that the (iii) Any load factor up to those spec- analytical representation is valid for ified in § 25.345(a) for the wing-flaps ex- the design conditions specified in tended condition. § 25.473. (2) Unless there are other means to (1) The configurations subjected to decelerate the airplane in flight at this energy absorption tests at limit design speed, the landing gear, the retracting conditions must include at least the mechanism, and the airplane structure design landing weight or the design (including wheel well doors) must be takeoff weight, whichever produces the designed to withstand the flight loads greater value of landing impact energy. occurring with the landing gear in the (2) The test attitude of the landing extended position at any speed up to gear unit and the application of appro- 0.67 VC. priate drag loads during the test must (3) Landing gear doors, their oper- simulate the airplane landing condi- ating mechanism, and their supporting tions in a manner consistent with the structures must be designed for the development of rational or conserv- yawing maneuvers prescribed for the ative limit loads. airplane in addition to the conditions (b) The landing gear may not fail in of airspeed and load factor prescribed a test, demonstrating its reserve en- in paragraphs (a)(1) and (2) of this sec- ergy absorption capacity, simulating a tion. descent velocity of 12 f.p.s. at design (b) Landing gear lock. There must be landing weight, assuming airplane lift positive means to keep the landing not greater than airplane weight act- gear extended in flight and on the ing during the landing impact. ground. There must be positive means (c) In lieu of the tests prescribed in to keep the landing gear and doors in this section, changes in previously ap- the correct retracted position in flight, proved design weights and minor unless it can be shown that lowering of changes in design may be substantiated the landing gear or doors, or flight by analyses based on previous tests with the landing gear or doors ex- conducted on the same basic landing tended, at any speed, is not hazardous. 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 the normal retraction system; or §§ 25.725–25.727 [Reserved] (2) The failure of any single source of hydraulic, electric, or equivalent en- § 25.729 Retracting mechanism. ergy supply. (a) General. For airplanes with re- (d) Operation test. The proper func- tractable landing gear, the following tioning of the retracting mechanism apply: must be shown by operation tests. (1) The landing gear retracting mech- (e) Position indicator and warning de- anism, wheel well doors, and sup- vice. If a retractable landing gear is porting structure, must be designed used, there must be a landing gear po- for— sition indicator easily visible to the (i) The loads occurring in the flight pilot or to the appropriate crew mem- conditions when the gear is in the re- bers (as well as necessary devices to ac- tracted position, tuate the indicator) to indicate with- (ii) The combination of friction out ambiguity that the retractable loads, inertia loads, brake torque loads, units and their associated doors are se- air loads, and gyroscopic loads result- cured in the extended (or retracted) po- ing from the wheels rotating at a pe- sition. The means must be designed as ripheral speed equal to 1.23VSR (with follows: the wing-flaps in take-off position at (1) If switches are used, they must be design take-off weight), occurring dur- located and coupled to the landing gear ing retraction and extension at any mechanical systems in a manner that airspeed up to 1.5 VSR1 (with the wing- prevents an erroneous indication of 273

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‘‘down and locked’’ if the landing gear corresponding static ground reaction is not in a fully extended position, or of with— ‘‘up and locked’’ if the landing gear is (1) Design maximum weight; and not in the fully retracted position. The (2) Critical center of gravity. switches may be located where they (c) The maximum limit load rating of are operated by the actual landing gear each wheel must equal or exceed the locking latch or device. maximum radial limit load determined (2) The flightcrew must be given an under the applicable ground load re- aural warning that functions continu- quirements of this part. ously, or is periodically repeated, if a (d) Overpressure burst prevention. landing is attempted when the landing Means must be provided in each wheel gear is not locked down. to prevent wheel failure and tire burst (3) The warning must be given in suf- that may result from excessive pressur- ficient time to allow the landing gear ization of the wheel and tire assembly. to be locked down or a go-around to be (e) Braked wheels. Each braked wheel made. must meet the applicable requirements (4) There must not be a manual shut- of § 25.735. off means readily available to the flightcrew for the warning required by [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as paragraph (e)(2) of this section such amended by Amdt. 25–72, 55 FR 29777, July 20, that it could be operated instinctively, 1990; Amdt. 25–107, 67 FR 20420, Apr. 24, 2002] inadvertently, or by habitual reflexive § 25.733 Tires. action. (5) The system used to generate the (a) When a landing gear axle is fitted aural warning must be designed to with a single wheel and tire assembly, minimize false or inappropriate alerts. the wheel must be fitted with a suit- (6) Failures of systems used to in- able tire of proper fit with a speed rat- hibit the landing gear aural warning, ing approved by the Administrator that would prevent the warning system that is not exceeded under critical con- from operating, must be improbable. ditions and with a load rating approved (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 amended by Amdt. 25–23, 35 FR 5676, Apr. 8, gravity position with a force of 1.0g 1970; Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; acting downward at the center of grav- 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.

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The reactions in this case must be dis- shown that the tire liner material will tributed to the nose and main wheels not produce a volatile gas when heated by the principles of statics with a drag or that means are provided to prevent reaction equal to 0.31 times the tire temperatures from reaching unsafe vertical load at each wheel with brakes levels. capable of producing this ground reac- [Amdt. 25–48, 44 FR 68752, Nov. 29, 1979; Amdt. tion. This nose tire load may not ex- 25–72, 55 FR 29777, July 20, 1990, as amended ceed 1.5 times the load rating of the by Amdt. 25–78, 58 FR 11781, Feb. 26, 1993] tire. (3) The ground reaction of the tire § 25.735 Brakes and braking systems. corresponding to the most critical (a) Approval. Each assembly con- combination of airplane weight (up to sisting of a wheel(s) and brake(s) must maximum ramp weight) and center of be approved. gravity position combined with forces of 1.0g downward and 0.20g forward act- (b) Brake system capability. The brake ing at the center of gravity. The reac- system, associated systems and compo- tions in this case must be distributed nents must be designed and con- to the nose and main wheels by the structed so that: principles of statics with a drag reac- (1) If any electrical, pneumatic, hy- tion equal to 0.20 times the vertical draulic, or mechanical connecting or load at each wheel with brakes capable transmitting element fails, or if any of producing this ground reaction. This single source of hydraulic or other nose tire load may not exceed 1.5 times brake operating energy supply is lost, the load rating of the tire. it is possible to bring the airplane to (c) When a landing gear axle is fitted rest with a braked roll stopping dis- with more than one wheel and tire as- tance of not more than two times that sembly, such as dual or dual-tandem, obtained in determining the landing each wheel must be fitted with a suit- distance as prescribed in § 25.125. able tire of proper fit with a speed rat- (2) Fluid lost from a brake hydraulic ing approved by the Administrator system following a failure in, or in the that is not exceeded under critical con- vicinity of, the brakes is insufficient to ditions, and with a load rating ap- cause or support a hazardous fire on proved by the Administrator that is the ground or in flight. not exceeded by— (c) Brake controls. The brake controls (1) The loads on each main wheel must be designed and constructed so tire, corresponding to the most critical that: combination of airplane weight (up to (1) Excessive control force is not re- maximum weight) and center of grav- quired for their operation. ity position, when multiplied by a fac- (2) If an automatic braking system is tor of 1.07; and installed, means are provided to: (2) Loads specified in paragraphs (i) Arm and disarm the system, and (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:

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(1) It must operate satisfactorily over (g) Brake condition after high kinetic the range of expected runway condi- energy dynamometer stop(s). Following tions, without external adjustment. the high kinetic energy stop dem- (2) It must, at all times, have pri- onstration(s) required by paragraph (f) ority over the automatic braking sys- of this section, with the parking brake tem, if installed. promptly and fully applied for at least (f) Kinetic energy capacity—(1) Design 3 minutes, it must be demonstrated landing stop. The design landing stop is that for at least 5 minutes from appli- an operational landing stop at max- cation of the parking brake, no condi- imum landing weight. The design land- tion occurs (or has occurred during the ing stop brake kinetic energy absorp- stop), including fire associated with tion requirement of each wheel, brake, the tire or wheel and brake assembly, and tire assembly must be determined. that could prejudice the safe and com- It must be substantiated by dynamometer testing that the wheel, brake and plete evacuation of the airplane. tire assembly is capable of absorbing (h) Stored energy systems. An indica- not less than this level of kinetic en- tion to the flightcrew of the usable ergy throughout the defined wear stored energy must be provided if a range of the brake. The energy absorp- stored energy system is used to show tion rate derived from the airplane compliance with paragraph (b)(1) of manufacturer’s braking requirements this section. The available stored en- must be achieved. The mean decelera- ergy must be sufficient for: tion must not be less than 10 fps 2. (1) At least 6 full applications of the (2) Maximum kinetic energy accelerate- brakes when an antiskid system is not stop. The maximum kinetic energy ac- operating; and celerate-stop is a rejected takeoff for (2) Bringing the airplane to a com- the most critical combination of air- plete stop when an antiskid system is plane takeoff weight and speed. The ac- operating, under all runway surface celerate-stop brake kinetic energy ab- conditions for which the airplane is sorption requirement of each wheel, certificated. 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 absorbing not less than this level of ki- the permissible limit. The means must netic energy throughout the defined be reliable and readily visible. wear range of the brake. The energy (j) Overtemperature burst prevention. absorption rate derived from the air- Means must be provided in each braked plane manufacturer’s braking require- wheel to prevent a wheel failure, a tire ments must be achieved. The mean de- burst, or both, that may result from celeration must not be less than 6 fps2. elevated brake temperatures. Addition- (3) Most severe landing stop. The most ally, all wheels must meet the require- severe landing stop is a stop at the ments of § 25.731(d). most critical combination of airplane (k) Compatibility. Compatibility of landing weight and speed. The most se- the wheel and brake assemblies with vere landing stop brake kinetic energy the airplane and its systems must be absorption requirement of each wheel, substantiated. brake, and tire assembly must be determined. It must be substantiated by [Doc. No. FAA–1999–6063, 67 FR 20420, Apr. 24, dynamometer testing that, at the de- 2002, as amended by Amdt. 25–108, 67 FR clared fully worn limit(s) of the brake 70827, Nov. 26, 2002; 68 FR 1955, Jan. 15, 2003] heat sink, the wheel, brake and tire as- § 25.737 Skis. sembly is capable of absorbing not less than this level of kinetic energy. The Each ski must be approved. The max- most severe landing stop need not be imum limit load rating of each ski considered for extremely improbable must equal or exceed the maximum failure conditions or if the maximum limit load determined under the appli- kinetic energy accelerate-stop energy cable ground load requirements of this is more severe. part.

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FLOATS AND HULLS rain or snow, it will not leak in a manner that will distract the crew or harm § 25.751 Main float buoyancy. the structure. Each main float must have— (e) Vibration and noise characteris- (a) A buoyancy of 80 percent in excess tics of cockpit equipment may not of that required to support the max- interfere with safe operation of the air- imum weight of the seaplane or am- plane. phibian in fresh water; and (b) Not less than five watertight com- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as partments approximately equal in vol- amended by Amdt. 25–4, 30 FR 6113, Apr. 30, 1965] ume.

§ 25.753 Main float design. § 25.772 Pilot compartment doors. Each main float must be approved For an airplane that has a lockable and must meet the requirements of door installed between the pilot com- § 25.521. partment and the passenger compartment: § 25.755 Hulls. (a) For airplanes with a maximum (a) Each hull must have enough wa- passenger seating configuration of tertight compartments so that, with more than 20 seats, the emergency exit any two adjacent compartments flood- configuration must be designed so that ed, the buoyancy of the hull and auxil- neither crewmembers nor passengers iary floats (and wheel tires, if used) require use of the flightdeck door in provides a margin of positive stability order to reach the emergency exits pro- great enough to minimize the prob- vided for them; and ability of capsizing in rough, fresh (b) Means must be provided to enable water. flight crewmembers to directly enter (b) Bulkheads with watertight doors the passenger compartment from the may be used for communication be- pilot compartment if the cockpit door tween compartments. becomes jammed. (c) There must be an emergency PERSONNEL AND CARGO means to enable a flight attendant to ACCOMMODATIONS enter the pilot compartment in the event that the flightcrew becomes in- § 25.771 Pilot compartment. capacitated. (a) Each pilot compartment and its equipment must allow the minimum [Doc. No. 24344, 55 FR 29777, July 20, 1990, as flight crew (established under § 25.1523) amended by Amdt. 25–106, 67 FR 2127, Jan. 15, 2002] to perform their duties without unrea- sonable concentration or fatigue. § 25.773 Pilot compartment view. (b) The primary controls listed in § 25.779(a), excluding cables and control (a) Nonprecipitation conditions. For rods, must be located with respect to nonprecipitation conditions, the fol- the propellers so that no member of the lowing apply: minimum flight crew (established (1) Each pilot compartment must be under § 25.1523), or part of the controls, arranged to give the pilots a suffi- lies in the region between the plane of ciently extensive, clear, and undis- rotation of any inboard propeller and torted view, to enable them to safely the surface generated by a line passing perform any maneuvers within the op- through the center of the propeller hub erating limitations of the airplane, in- making an angle of five degrees for- cluding taxiing takeoff, approach, and ward or aft of the plane of rotation of landing. the propeller. (2) Each pilot compartment must be (c) If provision is made for a second free of glare and reflection that could pilot, the airplane must be controllable interfere with the normal duties of the with equal safety from either pilot minimum flight crew (established seat. under § 25.1523). This must be shown in (d) The pilot compartment must be day and night flight tests under non- constructed so that, when flying in precipitation conditions.

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(b) Precipitation conditions. For pre- (i) Any system failure or combina- cipitation conditions, the following tion of failures which is not extremely apply: improbable, in accordance with (1) The airplane must have a means § 25.1309, under the precipitation condi- to maintain a clear portion of the tions specified in paragraph (b)(1) of windshield, during precipitation condi- this section. tions, sufficient for both pilots to have (ii) An encounter with severe hail, a sufficiently extensive view along the birds, or insects. flight path in normal flight attitudes (c) Internal windshield and window of the airplane. This means must be de- fogging. The airplane must have a signed to function, without continuous means to prevent fogging of the inter- attention on the part of the crew, in— nal portions of the windshield and win- (i) Heavy rain at speeds up to 1.5 VSR1 dow panels over an area which would with lift and drag devices retracted; provide the visibility specified in para- and graph (a) of this section under all in- (ii) The icing conditions specified in ternal and external ambient condi- Appendix C of this part and the fol- tions, including precipitation condi- lowing icing conditions specified in Ap- tions, in which the airplane is intended pendix O of this part, if certification to be operated. for flight in icing conditions is sought: (d) Fixed markers or other guides (A) For airplanes certificated in ac- must be installed at each pilot station cordance with § 25.1420(a)(1), the icing to enable the pilots to position them- conditions that the airplane is certified selves in their seats for an optimum to safely exit following detection. combination of outside visibility and (B) For airplanes certificated in ac- instrument scan. If lighted markers or cordance with § 25.1420(a)(2), the icing conditions that the airplane is certified guides are used they must comply with to safely operate in and the icing con- the requirements specified in § 25.1381. ditions that the airplane is certified to (e) Vision systems with transparent dis- safely exit following detection. plays. A vision system with a trans- (C) For airplanes certificated in ac- parent display surface located in the cordance with § 25.1420(a)(3) and for air- pilot’s outside field of view, such as a planes not subject to § 25.1420, all icing head up-display, head mounted display, conditions. or other equivalent display, must meet (2) No single failure of the systems the following requirements in non- used to provide the view required by precipitation and precipitation condi- paragraph (b)(1) of this section must tions: cause the loss of that view by both pi- (1) While the vision system display is lots in the specified precipitation con- in operation, it must compensate for ditions. interference with the pilot’s outside (3) The first pilot must have a win- field of view such that the combination dow that— of what is visible in the display and (i) Is openable under the conditions what remains visible through and prescribed in paragraph (b)(1) of this around it, enables the pilot to perform section when the cabin is not pressur- the maneuvers and normal duties of ized; paragraph (a) of this section. (ii) Provides the view specified in (2) The pilot’s view of the external paragraph (b)(1) of this section; and scene may not be distorted by the (iii) Provides sufficient protection transparent display surface or by the from the elements against impairment vision system imagery. When the vi- of the pilot’s vision. sion system displays imagery or any (4) The openable window specified in symbology that is referenced to the im- paragraph (b)(3) of this section need agery and outside scene topography, not be provided if it is shown that an including attitude symbology, flight area of the transparent surface will re- path vector, and flight path angle ref- main clear sufficient for at least one erence cue, that imagery and sym- pilot to land the airplane safely in the bology must be aligned with, and event of— scaled to, the external scene.

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(3) The vision system must provide a of the material used, and the effects of means to allow the pilot using the dis- temperatures and temperature dif- play to immediately deactivate and re- ferentials. The windshield and window activate the vision system imagery, on panels must be capable of withstanding demand, without removing the pilot’s the maximum cabin pressure differen- hands from the primary flight controls tial loads combined with critical aero- or thrust controls. dynamic pressure and temperature ef- (4) When the vision system is not in fects after any single failure in the in- operation it may not restrict the pilot stallation or associated systems. It from performing the maneuvers speci- may be assumed that, after a single fied in paragraph (a)(1) of this section failure that is obvious to the flight or the pilot compartment from meet- crew (established under § 25.1523), the ing the provisions of paragraph (a)(2) of cabin pressure differential is reduced this section. from the maximum, in accordance with appropriate operating limitations, to [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as 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 the pilots must be arranged so that, as- FR 1618, Jan. 11, 2012; Amdt. 25–140, 79 FR suming the loss of vision through any 65525, Nov. 4, 2014; Docket FAA–2013–0485, one panel, one or more panels remain Amdt. 25–144, 81 FR 90169, Dec. 13, 2016] available for use by a pilot seated at a § 25.775 Windshields and windows. pilot station to permit continued safe flight and landing. (a) Internal panes must be made of nonsplintering material. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (b) Windshield panes directly in front amended by Amdt. 25–23, 35 FR 5676, Apr. 8, of the pilots in the normal conduct of 1970; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976] their duties, and the supporting struc- § 25.777 Cockpit controls. tures for these panes, must withstand, without penetration, the impact of a (a) Each cockpit control must be lo- four-pound bird when the velocity of cated to provide convenient operation the airplane (relative to the bird along and to prevent confusion and inad- the airplane’s flight path) is equal to vertent operation. the value of VC, at sea level, selected (b) The direction of movement of under § 25.335(a). cockpit controls must meet the re- (c) Unless it can be shown by analysis quirements of § 25.779. Wherever prac- or tests that the probability of occur- ticable, the sense of motion involved in rence of a critical windshield frag- the operation of other controls must mentation condition is of a low order, correspond to the sense of the effect of the airplane must have a means to the operation upon the airplane or minimize the danger to the pilots from upon the part operated. Controls of a flying windshield fragments due to bird variable nature using a rotary motion impact. This must be shown for each must move clockwise from the off posi- transparent pane in the cockpit that— tion, through an increasing range, to (1) Appears in the front view of the the full on position. airplane; (c) The controls must be located and (2) Is inclined 15 degrees or more to arranged, with respect to the pilots’ the longitudinal axis of the airplane; seats, so that there is full and unre- and stricted movement of each control (3) Has any part of the pane located without interference from the cockpit where its fragmentation will constitute structure or the clothing of the min- a hazard to the pilots. imum flight crew (established under (d) The design of windshields and § 25.1523) when any member of this windows in pressurized airplanes must flight crew, from 5′2″ to 6′3″ in height, be based on factors peculiar to high al- is seated with the seat belt and shoul- titude operation, including the effects der harness (if provided) fastened. of continuous and cyclic pressurization (d) Identical powerplant controls for loadings, the inherent characteristics each engine must be located to prevent

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confusion as to the engines they con- (1) Primary. trol. (e) Wing flap controls and other aux- Controls Motion and effect iliary lift device controls must be lo- Aileron ...... Right (clockwise) for right wing cated on top of the pedestal, aft of the down. throttles, centrally or to the right of Elevator ...... Rearward for nose up. the pedestal centerline, and not less Rudder ...... Right pedal forward for nose right. than 10 inches aft of the landing gear (2) Secondary. control. (f) The landing gear control must be Controls Motion and effect located forward of the throttles and must be operable by each pilot when Flaps (or auxiliary lift Forward for flaps up; rearward for devices). flaps down. seated with seat belt and shoulder har- Trim tabs (or equiva- Rotate to produce similar rotation of ness (if provided) fastened. lent). the airplane about an axis parallel (g) Control knobs must be shaped in to the axis of the control. accordance with § 25.781. In addition, the knobs must be of the same color, (b) Powerplant and auxiliary con- and this color must contrast with the trols: color of control knobs for other pur- (1) Powerplant. poses and the surrounding cockpit. Controls Motion and effect (h) If a flight engineer is required as part of the minimum flight crew (es- Power or thrust ...... Forward to increase forward thrust tablished under § 25.1523), the airplane and rearward to increase rearward thrust. must have a flight engineer station lo- Propellers ...... Forward to increase rpm. cated and arranged so that the flight Mixture ...... Forward or upward for rich. crewmembers can perform their func- Carburetor air heat ...... Forward or upward for cold. tions efficiently and without inter- Supercharger ...... Forward or upward for low blower. For turbosuperchargers, forward, fering with each other. upward, or clockwise, to increase pressure. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–46, 43 FR 50596, Oct. 30, 1978] (2) Auxiliary.

§ 25.779 Motion and effect of cockpit Controls Motion and effect controls. Landing gear ...... Down to extend. Cockpit controls must be designed so that they operate in accordance with [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the following movement and actuation: amended by Amdt. 25–72, 55 FR 29778, July 20, (a) Aerodynamic controls: 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 distributed among the required floor- critical load distribution at the appro- level emergency exits to the extent priate maximum load factors cor- feasible. responding to the specified flight and (2) To the extent possible, without ground load conditions, and to those compromising proximity to a required emergency landing conditions of floor level emergency exit, located to § 25.561(b)(3) for which the breaking provide a direct view of the cabin area loose of the contents of such compart- for which the flight attendant is re- ments in the specified direction could— sponsible. (1) Cause direct injury to occupants; (3) Positioned so that the seat will (2) Penetrate fuel tanks or lines or not interfere with the use of a passage- cause fire or explosion hazard by dam- way or exit when the seat is not in use. age to adjacent systems; or (4) Located to minimize the prob- (3) Nullify any of the escape facilities ability that occupants would suffer in- provided for use after an emergency jury by being struck by items dislodged landing. from service areas, stowage compartments, or service equipment. If the airplane has a passenger-seating (5) Either forward or rearward facing configuration, excluding pilot seats, of with an energy absorbing rest that is 10 seats or more, each stowage com- designed to support the arms, shoul- partment in the passenger cabin, ex- ders, head, and spine. cept for under seat and overhead com- (6) Equipped with a restraint system partments for passenger convenience, consisting of a combined safety belt must be completely enclosed. and shoulder harness unit with a single (b) There must be a means to prevent point release. There must be means to the contents in the compartments from secure each restraint system when not becoming a hazard by shifting, under in use to prevent interference with the loads specified in paragraph (a) of rapid egress in an emergency. this section. For stowage compart- (i) Each safety belt must be equipped ments in the passenger and crew cabin, with a metal to metal latching device. if the means used is a latched door, the (j) If the seat backs do not provide a design must take into consideration firm handhold, there must be a hand- the wear and deterioration expected in grip or rail along each aisle to enable service. persons to steady themselves while (c) If cargo compartment lamps are using the aisles in moderately rough installed, each lamp must be installed air. so as to prevent contact between lamp (k) Each projecting object that would bulb and cargo. injure persons seated or moving about [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the airplane in normal flight must be amended by Amdt. 25–32, 37 FR 3969, Feb. 24, padded. 1972; Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; (l) Each forward observer’s seat re- Amdt. 25–51, 45 FR 7755, Feb. 4, 1980; Amdt. quired by the operating rules must be 25–139, 79 FR 59430, Oct. 2, 2014] shown to be suitable for use in con-

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

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(3) Cargo compartment fire suppression. limited by the forward bulkhead and An extinguishing agent must be capa- the aft bulkhead of the passenger cabin ble of suppressing a fire. All cargo- and cargo compartment beyond which compartment fire suppression systems only one-half the sphere is applied. must be designed to withstand the fol- (ii) Where compliance with paragraph lowing effects, including support struc- (c)(2)(i) of this section is impracticable, ture displacements or adjacent mate- other design precautions must be taken rials displacing against the distribu- to maximize the survivability of those tion system: systems. (i) Impact or damage from a 0.5-inch (3) Interior design to facilitate searches. diameter aluminum sphere traveling at Design features must be incorporated 430 feet per second (131.1 meters per that will deter concealment or promote second); discovery of weapons, explosives, or (ii) A 15-pound per square-inch (103.4 other objects from a simple inspection kPa) pressure load if the projected sur- in the following areas of the airplane face area of the component is greater cabin: than 4 square feet. Any single dimen- (i) Areas above the overhead bins sion greater than 4 feet (1.22 meters) must be designed to prevent objects 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 satisfy this requirement. force applied anywhere along the distribution 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) Each chemical oxygen generator this section or are installed remotely or its installation must be designed to from the cargo compartment. be secure from deliberate manipulation (c) An airplane with a maximum cer- by one of the following: tificated passenger seating capacity of (1) By providing effective resistance more than 60 persons or a maximum to tampering, certificated takeoff gross weight of (2) By providing an effective com- over 100,000 pounds (45,359 Kilograms) bination of resistance to tampering and must comply with the following: active tamper-evident features, (1) Least risk bomb location. An air- (3) By installation in a location or plane must be designed with a des- manner whereby any attempt to access ignated location where a bomb or other the generator would be immediately explosive device could be placed to best obvious, or protect flight-critical structures and (4) By a combination of approaches systems from damage in the case of specified in paragraphs (d)(1), (d)(2) and detonation. (d)(3) of this section that the Adminis- (2) Survivability of systems. (i) Except trator finds provides a secure installa- where impracticable, redundant air- tion. plane systems necessary for continued (e) Exceptions. Airplanes used solely safe flight and landing must be phys- to transport cargo only need to meet ically separated, at a minimum, by an the requirements of paragraphs (b)(1), amount equal to a sphere of diameter (b)(3), and (c)(2) of this section. (f) Material Incorporated by Reference. = ()π DH2 0 / You must use National Institute of Justice (NIJ) Standard 0101.04, Ballistic (where H0 is defined under § 25.365(e)(2) Resistance of Personal Body Armor, of this part and D need not exceed 5.05 June 2001, Revision A, to establish bal- feet (1.54 meters)). The sphere is ap- listic resistance as required by para- plied everywhere within the fuselage— graph (a)(3) of this section.

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(1) The Director of the Federal Reg- lowances must be made for probable ister approved the incorporation by ref- structural damage and leakage. If the erence of this document under 5 U.S.C. airplane has fuel tanks (with fuel jetti- 552(a) and 1 CFR part 51. soning provisions) that can reasonably (2) You may review copies of NIJ be expected to withstand a ditching Standard 0101.04 at the: without leakage, the jettisonable vol- (i) National Institute of Justice ume of fuel may be considered as buoy- (NIJ), http://www.ojp.usdoj.gov/nij, tele- ancy volume. phone (202) 307–2942; or (e) Unless the effects of the collapse (ii) National Archives and Records of external doors and windows are ac- Administration (NARA). For informa- counted for in the investigation of the tion on the availability of this mate- probable behavior of the airplane in a rial at NARA, call (202) 741–6030, or go water landing (as prescribed in para- to http://www.archives.gov/federal-reg- graphs (c) and (d) of this section), the ister/cfr/ibr-locations.html. external doors and windows must be (3) You may obtain copies of NIJ designed to withstand the probable Standard 0101.04 from the National maximum local pressures. Criminal Justice Reference Service, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as P.O. Box 6000, Rockville, MD 20849–6000, amended by Amdt. 25–72, 55 FR 29781, July 20, telephone (800) 851–3420. 1990] [Amdt. 25–127; 121–341, 73 FR 63879, Oct. 28, 2008, as amended at 74 FR 22819, May 15, 2009; § 25.803 Emergency evacuation. Amdt. 25–138, 79 FR 13519, Mar. 11, 2014; Doc. (a) Each crew and passenger area No. FAA–2018–0119, Amdt. 25–145, 83 FR 9169, must have emergency means to allow Mar. 5, 2018] rapid evacuation in crash landings, with the landing gear extended as well EMERGENCY PROVISIONS as with the landing gear retracted, con- § 25.801 Ditching. sidering the possibility of the airplane 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 of the airplane, must be considered. [Doc. No. 24344, 55 FR 29781, July 20, 1990] (d) It must be shown that, under reasonably probable water conditions, the § 25.807 Emergency exits. flotation time and trim of the airplane (a) Type. For the purpose of this part, will allow the occupants to leave the the types of exits are defined as fol- airplane and enter the liferafts re- lows: quired by § 25.1415. If compliance with (1) Type I. This type is a floor-level this provision is shown by buoyancy exit with a rectangular opening of not and trim computations, appropriate al- less than 24 inches wide by 48 inches

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high, with corner radii not greater high, with corner radii not greater than eight inches. than 10 inches. (2) Type II. This type is a rectangular (b) Step down distance. Step down dis- opening of not less than 20 inches wide tance, as used in this section, means by 44 inches high, with corner radii not the actual distance between the bot- greater than seven inches. Type II exits tom of the required opening and a usa- must be floor-level exits unless located ble foot hold, extending out from the over the wing, in which case they must fuselage, that is large enough to be ef- not have a step-up inside the airplane fective without searching by sight or of more than 10 inches nor a step-down feel. outside the airplane of more than 17 (c) Over-sized exits. Openings larger inches. than those specified in this section, (3) Type III. This type is a rectan- whether or not of rectangular shape, gular opening of not less than 20 inches may be used if the specified rectan- wide by 36 inches high with corner gular opening can be inscribed within radii not greater than seven inches, the opening and the base of the in- and with a step-up inside the airplane scribed rectangular opening meets the of not more than 20 inches. If the exit specified step-up and step-down is located over the wing, the step-down heights. outside the airplane may not exceed 27 (d) Asymmetry. Exits of an exit pair inches. need not be diametrically opposite (4) Type IV. This type is a rectan- each other nor of the same size; how- gular opening of not less than 19 inches ever, the number of passenger seats wide by 26 inches high, with corner permitted under paragraph (g) of this radii not greater than 6.3 inches, lo- section is based on the smaller of the cated over the wing, with a step-up in- two exits. side the airplane of not more than 29 (e) Uniformity. Exits must be distrib- inches and a step-down outside the air- uted as uniformly as practical, taking plane of not more than 36 inches. into account passenger seat distribu- (5) Ventral. This type is an exit from tion. the passenger compartment through (f) Location. (1) Each required pas- the pressure shell and the bottom fuse- senger emergency exit must be acces- lage skin. The dimensions and physical sible to the passengers and located configuration of this type of exit must where it will afford the most effective allow at least the same rate of egress means of passenger evacuation. as a Type I exit with the airplane in (2) If only one floor-level exit per side the normal ground attitude, with land- is prescribed, and the airplane does not ing gear extended. have a tailcone or ventral emergency (6) Tailcone. This type is an aft exit exit, the floor-level exits must be in from the passenger compartment the rearward part of the passenger through the pressure shell and through compartment unless another location an openable cone of the fuselage aft of affords a more effective means of pas- the pressure shell. The means of open- senger evacuation. ing the tailcone must be simple and ob- (3) If more than one floor-level exit vious and must employ a single oper- per side is prescribed, and the airplane ation. does not have a combination cargo and (7) Type A. This type is a floor-level passenger configuration, at least one exit with a rectangular opening of not floor-level exit must be located in each less than 42 inches wide by 72 inches side near each end of the cabin. high, with corner radii not greater (4) For an airplane that is required to than seven inches. 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.

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(g) Type and number required. The two Type C or larger exits in each side maximum number of passenger seats of the fuselage. permitted depends on the type and (9) If a passenger ventral or tailcone number of exits installed in each side exit is installed and that exit provides of the fuselage. Except as further re- at least the same rate of egress as a stricted in paragraphs (g)(1) through Type III exit with the airplane in the (g)(9) of this section, the maximum most adverse exit opening condition number of passenger seats permitted that would result from the collapse of for each exit of a specific type installed one or more legs of the landing gear, an in each side of the fuselage is as fol- increase in the passenger seating con- lows: figuration is permitted as follows: Type A 110 (i) For a ventral exit, 12 additional Type B 75 passenger seats. Type C 55 (ii) For a tailcone exit incorporating Type I 45 a floor level opening of not less than 20 Type II 40 inches wide by 60 inches high, with cor- Type III 35 ner radii not greater than seven inches, Type IV 9 in the pressure shell and incorporating (1) For a passenger seating configura- an approved assist means in accordance tion of 1 to 9 seats, there must be at with § 25.810(a), 25 additional passenger least one Type IV or larger overwing seats. exit in each side of the fuselage or, if (iii) For a tailcone exit incorporating overwing exits are not provided, at an opening in the pressure shell which least one exit in each side that meets is at least equivalent to a Type III the minimum dimensions of a Type III emergency exit with respect to dimen- exit. sions, step-up and step-down distance, (2) For a passenger seating configura- and with the top of the opening not tion of more than 9 seats, each exit less than 56 inches from the passenger must be a Type III or larger exit. compartment floor, 15 additional pas- (3) For a passenger seating configura- senger seats. tion of 10 to 19 seats, there must be at (h) Other exits. The following exits least one Type III or larger exit in each also must meet the applicable emer- side of the fuselage. gency exit requirements of §§ 25.809 (4) For a passenger seating configura- through 25.812, and must be readily action of 20 to 40 seats, there must be at cessible: least two exits, one of which must be a (1) Each emergency exit in the pas- Type II or larger exit, in each side of senger compartment in excess of the the fuselage. minimum number of required emer- (5) For a passenger seating configura- gency exits. tion of 41 to 110 seats, there must be at (2) Any other floor-level door or exit least two exits, one of which must be a that is accessible from the passenger Type I or larger exit, in each side of compartment and is as large or larger the fuselage. than a Type II exit, but less than 46 (6) For a passenger seating configura- inches wide. tion of more than 110 seats, the emer- (3) Any other ventral or tail cone gency exits in each side of the fuselage passenger exit. must include at least two Type I or (i) Ditching emergency exits for pas- larger exits. sengers. Whether or not ditching cer- (7) The combined maximum number tification is requested, ditching emer- of passenger seats permitted for all gency exits must be provided in accord- Type III exits is 70, and the combined ance with the following requirements, maximum number of passenger seats unless the emergency exits required by permitted for two Type III exits in paragraph (g) of this section already each side of the fuselage that are sepa- meet them: rated by fewer than three passenger (1) For airplanes that have a pas- seat rows is 65. senger seating configuration of nine or (8) If a Type A, Type B, or Type C fewer seats, excluding pilot seats, one exit is installed, there must be at least exit above the waterline in each side of

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the airplane, meeting at least the di- an unobstructed opening to the out- mensions of a Type IV exit. side. In addition, each emergency exit (2) For airplanes that have a pas- must have means to permit viewing of senger seating configuration of 10 of the conditions outside the exit when more seats, excluding pilot seats, one the exit is closed. The viewing means exit above the waterline in a side of the may be on or adjacent to the exit pro- airplane, meeting at least the dimen- vided no obstructions exist between the sions of a Type III exit for each unit (or exit and the viewing means. Means part of a unit) of 35 passenger seats, must also be provided to permit view- but no less than two such exits in the ing of the likely areas of evacuee passenger cabin, with one on each side ground contact. The likely areas of of the airplane. The passenger seat/ evacuee ground contact must be exit ratio may be increased through viewable during all lighting conditions the use of larger exits, or other means, with the landing gear extended as well provided it is shown that the evacu- as in all conditions of landing gear col- ation capability during ditching has lapse. been improved accordingly. (b) Each emergency exit must be (3) If it is impractical to locate side openable from the inside and the out- exits above the waterline, the side side except that sliding window emer- exits must be replaced by an equal gency exits in the flight crew area need number of readily accessible overhead not be openable from the outside if hatches of not less than the dimensions other approved exits are convenient of a Type III exit, except that for air- and readily accessible to the flight planes with a passenger configuration crew area. Each emergency exit must of 35 or fewer seats, excluding pilot be capable of being opened, when there seats, the two required Type III side is no fuselage deformation— exits need be replaced by only one (1) With the airplane in the normal overhead hatch. ground attitude and in each of the atti- (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 system for operating more than one § 25.809 Emergency exit arrangement. 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

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failure of the primary system) is ac- must be capable of carrying simulta- ceptable. neously two parallel lines of evacuees. (e) Each emergency exit must be In addition, the assisting means must shown by tests, or by a combination of be designed to meet the following re- analysis and tests, to meet the require- quirements— ments of paragraphs (b) and (c) of this (i) It must be automatically deployed section. and deployment must begin during the (f) Each door must be located where interval between the time the exit persons using them will not be endan- opening means is actuated from inside gered by the propellers when appro- the airplane and the time the exit is priate operating procedures are used. fully opened. However, each passenger (g) There must be provisions to mini- emergency exit which is also a pas- mize the probability of jamming of the senger entrance door or a service door emergency exits resulting from fuse- must be provided with means to pre- lage deformation in a minor crash vent deployment of the assisting means landing. when it is opened from either the in- (h) When required by the operating side or the outside under non- rules for any large passenger-carrying emergency conditions for normal use. turbojet-powered airplane, each ven- (ii) Except for assisting means in- tral exit and tailcone exit must be— stalled at Type C exits, it must be (1) Designed and constructed so that automatically erected within 6 seconds it cannot be opened during flight; and after deployment is begun. Assisting (2) Marked with a placard readable means installed at Type C exits must from a distance of 30 inches and in- be automatically erected within 10 sec- stalled at a conspicuous location near onds from the time the opening means the means of opening the exit, stating of the exit is actuated. that the exit has been designed and (iii) It must be of such length after constructed so that it cannot be opened full deployment that the lower end is during flight. self-supporting on the ground and pro- (i) Each emergency exit must have a vides safe evacuation of occupants to means to retain the exit in the open the ground after collapse of one or position, once the exit is opened in an more legs of the landing gear. emergency. The means must not re- (iv) It must have the capability, in quire separate action to engage when 25-knot winds directed from the most the exit is opened, and must require critical angle, to deploy and, with the positive action to disengage. 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. evacuate occupants safely to the 20, 1967; Amdt. 25–32, 37 FR 3970, Feb. 24, 1972; ground. Amdt. 25–34, 37 FR 25355, Nov. 30, 1972; Amdt. (v) For each system installation 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. 25–47, 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.

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(2) The assisting means for flightcrew nates is more than 6 feet from the emergency exits may be a rope or any ground with the airplane on the ground other means demonstrated to be suit- and the landing gear extended, for all able for the purpose. If the assisting other exit types. means is a rope, or an approved device (1) If the escape route is over the equivalent to a rope, it must be— flap, the height of the terminal edge (i) Attached to the fuselage structure must be measured with the flap in the at or above the top of the emergency takeoff or landing position, whichever exit opening, or, for a device at a pi- is higher from the ground. lot’s emergency exit window, at an- (2) The assisting means must be usa- other approved location if the stowed ble and self-supporting with one or device, or its attachment, would reduce more landing gear legs collapsed and the pilot’s view in flight; under a 25-knot wind directed from the (ii) Able (with its attachment) to most critical angle. withstand a 400-pound static load. (3) The assisting means provided for (b) Assist means from the cabin to each escape route leading from a Type the wing are required for each type A A or B emergency exit must be capable or Type B exit located above the wing of carrying simultaneously two par- and having a stepdown unless the exit allel lines of evacuees; and, the assist- without an assist-means can be shown ing means leading from any other exit to have a rate of passenger egress at type must be capable of carrying as least equal to that of the same type of many parallel lines of evacuees as non over-wing exit. If an assist means there are required escape routes. is required, it must be automatically (4) The assisting means provided for deployed and automatically erected each escape route leading from a Type concurrent with the opening of the C exit must be automatically erected 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 per- [Amdt. 25–72, 55 FR 29782, July 20, 1990, as cent, and must be defined by markings amended by Amdt. 25–88, 61 FR 57958, Nov. 8, with a surface-to-marking contrast 1996; 62 FR 1817, Jan. 13, 1997; Amdt. 25–114, 69 ratio of at least 5:1. 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, (a) Each passenger emergency exit, if the place on the airplane structure its means of access, and its means of at which the escape route required in opening must be conspicuously paragraph (c) of this section termi- marked.

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(b) The identity and location of each tending along at least 70 degrees of arc passenger emergency exit must be rec- at a radius approximately equal to ognizable from a distance equal to the three-fourths of the handle length. width of the cabin. (ii) So that the centerline of the exit (c) Means must be provided to assist handle is within ±1 inch of the pro- the occupants in locating the exits in jected point of the arrow when the han- conditions of dense smoke. dle has reached full travel and has re- (d) The location of each passenger leased the locking mechanism, and emergency exit must be indicated by a (iii) With the word ‘‘open’’ in red let- sign visible to occupants approaching ters 1 inch high, placed horizontally along the main passenger aisle (or near the head of the arrow. aisles). There must be— (f) Each emergency exit that is re- (1) A passenger emergency exit loca- quired to be openable from the outside, tor sign above the aisle (or aisles) near and its means of opening, must be each passenger emergency exit, or at marked on the outside of the airplane. another overhead location if it is more In addition, the following apply: practical because of low headroom, ex- (1) The outside marking for each pas- cept that one sign may serve more senger emergency exit in the side of than one exit if each exit can be seen the fuselage must include a 2-inch col- readily from the sign; ored band outlining the exit. (2) A passenger emergency exit marking sign next to each passenger emer- (2) Each outside marking including gency exit, except that one sign may the band, must have color contrast to serve two such exits if they both can be be readily distinguishable from the sur- seen readily from the sign; and rounding fuselage surface. The contrast (3) A sign on each bulkhead or divider must be such that if the reflectance of that prevents fore and aft vision along the darker color is 15 percent or less, the passenger cabin to indicate emer- the reflectance of the lighter color gency exits beyond and obscured by the must be at least 45 percent. ‘‘Reflec- bulkhead or divider, except that if this tance’’ is the ratio of the luminous flux is not possible the sign may be placed reflected by a body to the luminous at another appropriate location. flux it receives. When the reflectance (e) The location of the operating han- of the darker color is greater than 15 dle and instructions for opening exits percent, at least a 30-percent difference from the inside of the airplane must be between its reflectance and the reflec- shown in the following manner: tance of the lighter color must be pro- (1) Each passenger emergency exit vided. must have, on or near the exit, a mark- (3) In the case of exists other than ing that is readable from a distance of those in the side of the fuselage, such 30 inches. as ventral or tailcone exists, the exter- (2) Each Type A, Type B, Type C or nal means of opening, including in- Type I passenger emergency exit oper- structions if applicable, must be con- ating handle must— spicuously marked in red, or bright (i) Be self-illuminated with an initial chrome yellow if the background color brightness of at least 160 micro- is such that red is inconspicuous. When lamberts; or the opening means is located on only (ii) Be conspicuously located and well one side of the fuselage, a conspicuous illuminated by the emergency lighting marking to that effect must be pro- even in conditions of occupant crowd- vided on the other side. ing at the exit. (g) Each sign required by paragraph (3) [Reserved] (d) of this section may use the word (4) Each Type A, Type B, Type C, ‘‘exit’’ in its legend in place of the Type I, or Type II passenger emergency term ‘‘emergency exit’’. exit with a locking mechanism released by rotary motion of the handle [Amdt. 25–15, 32 FR 13264, Sept. 20, 1967, as amended by Amdt. 25–32, 37 FR 3970, Feb. 24, must be marked— 1972; Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; 43 (i) With a red arrow, with a shaft at FR 52495, Nov. 13, 1978; Amdt. 25–79, 58 FR least three-fourths of an inch wide and 45229, Aug. 26, 1993; Amdt. 25–88, 61 FR 57958, a head twice the width of the shaft, ex- Nov. 8, 1996]

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§ 25.812 Emergency lighting. inches high. These signs may be inter- (a) An emergency lighting system, nally electrically illuminated, or self- independent of the main lighting sys- illuminated by other than electrical tem, must be installed. However, the means, with an initial brightness of at sources of general cabin illumination least 160 microlamberts. The colors may be common to both the emergency may be reversed in the case of a sign and the main lighting systems if the that is self-illuminated by other than power supply to the emergency light- electrical means. ing system is independent of the power (c) General illumination in the pas- supply to the main lighting system. senger cabin must be provided so that The emergency lighting system must when measured along the centerline of include: main passenger aisle(s), and cross (1) Illuminated emergency exit mark- aisle(s) between main aisles, at seat ing and locating signs, sources of gen- arm-rest height and at 40-inch inter- eral cabin illumination, interior light- vals, the average illumination is not ing in emergency exit areas, and floor less than 0.05 foot-candle and the illu- proximity escape path marking. mination at each 40-inch interval is not (2) Exterior emergency lighting. less than 0.01 foot-candle. A main pas- (b) Emergency exit signs— senger aisle(s) is considered to extend (1) For airplanes that have a pas- along the fuselage from the most for- senger seating configuration, excluding ward passenger emergency exit or pilot seats, of 10 seats or more must cabin occupant seat, whichever is far- meet the following requirements: ther forward, to the most rearward pas- (i) Each passenger emergency exit lo- senger emergency exit or cabin occu- cator sign required by § 25.811(d)(1) and pant seat, whichever is farther aft. 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 pas- (f) Except for subsystems provided in senger seating configuration, excluding accordance with paragraph (h) of this pilot seats, of nine seats or less, that section that serve no more than one as- are required by § 25.811(d)(1), (2), and (3) sist means, are independent of the air- must have red letters at least 1 inch plane’s main emergency lighting sys- high on a white background at least 2 tem, and are automatically activated

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when the assist means is erected, the where an evacuee is likely to make emergency lighting system must be de- first contact with the ground outside signed as follows. the cabin. (1) The lights must be operable (h) The means required in manually from the flight crew station §§ 25.810(a)(1) and (d) to assist the occu- and from a point in the passenger com- pants in descending to the ground must partment that is readily accessible to a be illuminated so that the erected as- normal flight attendant seat. sist means is visible from the airplane. (2) There must be a flight crew warn- (1) If the assist means is illuminated ing light which illuminates when power by exterior emergency lighting, it is on in the airplane and the emergency must provide illumination of not less lighting control device is not armed. than 0.03 foot-candle (measured normal (3) The cockpit control device must to the direction of the incident light) have an ‘‘on,’’ ‘‘off,’’ and ‘‘armed’’ posi- at the ground end of the erected assist tion so that when armed in the cockpit means where an evacuee using the es- or turned on at either the cockpit or tablished escape route would normally flight attendant station the lights will make first contact with the ground, either light or remain lighted upon with the airplane in each of the atti- interruption (except an interruption tudes corresponding to the collapse of caused by a transverse vertical separa- one or more legs of the landing gear. tion of the fuselage during crash land- (2) If the emergency lighting sub- ing) of the airplane’s normal electric system illuminating the assist means power. There must be a means to safe- serves no other assist means, is inde- guard against inadvertent operation of pendent of the airplane’s main emer- 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 normally make first contact with the (measured normal to the direction of ground, with the airplane in each of the incident light) for a minimum the attitudes corresponding to the col- width of 42 inches for a Type A lapse of one or more legs of the landing overwing emergency exit and two feet gear. for all other overwing emergency exits (i) The energy supply to each emer- along the 30 percent of the slip-resist- gency lighting unit must provide the ant portion of the escape route re- required level of illumination for at quired in § 25.810(c) that is farthest least 10 minutes at the critical ambient from the exit; and conditions after emergency landing. (iii) Not less than 0.03 foot-candle on (j) If storage batteries are used as the the ground surface with the landing energy supply for the emergency light- gear extended (measured normal to the ing system, they may be recharged direction of the incident light) where from the airplane’s main electric power an evacuee using the established escape system: Provided, That, the charging route would normally make first con- circuit is designed to preclude inad- tact with the ground. vertent battery discharge into charg- (2) At each non-overwing emergency ing circuit faults. exit not required by § 25.810(a) to have (k) Components of the emergency descent assist means the illumination lighting system, including batteries, must be not less than 0.03 foot-candle wiring relays, lamps, and switches (measured normal to the direction of must be capable of normal operation the incident light) on the ground sur- after having been subjected to the iner- face with the landing gear extended tia forces listed in § 25.561(b).

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(l) The emergency lighting system must be unobstructed and at least 36 must be designed so that after any sin- inches wide. Passageways between indi- gle transverse vertical separation of vidual passenger areas and those lead- the fuselage during crash landing— ing to Type I, Type II, or Type C emer- (1) Not more than 25 percent of all gency exits must be unobstructed and electrically illuminated emergency at least 20 inches wide. Unless there lights required by this section are ren- are two or more main aisles, each Type dered inoperative, in addition to the A or B exit must be located so that lights that are directly damaged by the there is passenger flow along the main separation; aisle to that exit from both the forward (2) Each electrically illuminated exit and aft directions. If two or more main sign required under § 25.811(d)(2) re- aisles are provided, there must be un- mains operative exclusive of those that obstructed cross-aisles at least 20 are directly damaged by the separa- inches wide between main aisles. There tion; and must be— (3) At least one required exterior (1) A cross-aisle which leads directly emergency light for each side of the to each passageway between the near- airplane remains operative exclusive of est main aisle and a Type A or B exit; those that are directly damaged by the and separation. (2) A cross-aisle which leads to the [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967, as immediate vicinity of each passageway amended by Amdt. 25–28, 36 FR 16899, Aug. 26, between the nearest main aisle and a 1971; Amdt. 25–32, 37 FR 3971, Feb. 24, 1972; Type 1, Type II, or Type III exit; except Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. that when two Type III exits are lo- 25–58, 49 FR 43186, Oct. 26, 1984; Amdt. 25–88, cated within three passenger rows of 61 FR 57958, Nov. 8, 1996; Amdt. 25–116, 69 FR each other, a single cross-aisle may be 62788, Oct. 27, 2004; Amdt. 25–128, 74 FR 25645, 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),

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an assist space must be provided, if an provided must not be obstructed and assist means is required by § 25.810(a). there must be no interference in open- (6) There must be a handle, or han- ing the exit by seats, berths, or other dles, at each assist space, located to protrusions (including any seatback in enable the crewmember to steady him- the most adverse position) for a dis- self or herself: tance from that exit not less than the (i) While manually activating the as- width of the narrowest passenger seat sist means (where applicable) and, installed on the airplane. (ii) While assisting passengers during (ii) For airplanes that have a pas- an evacuation. senger seating configuration of 19 or (c) The following must be provided fewer, there may be minor obstructions for each Type III or Type IV exit—(1) in this region, if there are compen- There must be access from the nearest aisle to each exit. In addition, for each sating factors to maintain the effec- Type III exit in an airplane that has a tiveness of the exit. passenger seating configuration of 60 or (3) For each Type III exit, regardless more— of the passenger capacity of the air- (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 from that of the exit. partments to reach any required emer- (ii) In lieu of one 10- or 20-inch pas- gency exit from any seat in the passageway, there may be two passage- senger cabin, the passageway must be ways, between seat rows only, that unobstructed. However, curtains may must be at least 6 inches in width and be used if they allow free entry lead to an unobstructed space adjacent through the passageway. to each exit. (Adjacent exits must not (e) No door may be installed between share a common passageway.) The any passenger seat that is occupiable width of the passageways must be for takeoff and landing and any pas- measured with adjacent seats adjusted senger emergency exit, such that the to their most adverse position. The un- door crosses any egress path (including obstructed space adjacent to the exit aisles, crossaisles and passageways). must extend vertically from the floor (f) If it is necessary to pass through a to the ceiling (or bottom of sidewall doorway separating any crewmember stowage bins), inboard from the exit for seat (except those seats on the a distance not less than the width of flightdeck), occupiable for takeoff and the narrowest passenger seat installed landing, from any emergency exit, the on the airplane, and from the forward door must have a means to latch it in edge of the forward passageway to the the open position. The latching means aft edge of the aft passageway. The exit must be able to withstand the loads opening must be totally within the fore and aft bounds of the unobstructed imposed upon it when the door is sub- space. jected to the ultimate inertia forces, (2) In addition to the access— (i) For airplanes that have a passenger seating configuration of 20 or more, the projected opening of the exit

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relative to the surrounding structure, evacuation routes may not be depend- listed in § 25.561(b). ent on any powered device. The routes must be designed to minimize the pos- [Amdt. 25–1, 30 FR 3204, Mar. 9, 1965, as amended by Amdt. 25–15, 32 FR 13265, Sept. sibility of blockage which might result 20, 1967; Amdt. 25–32, 37 FR 3971, Feb. 24, 1972; from fire, mechanical or structural Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. failure, or persons standing on top of or 25–72, 55 FR 29783, July 20, 1990; Amdt. 25–76, against the escape routes. In the event 57 FR 19244, May 4, 1992; Amdt. 25–76, 57 FR the airplane’s main power system or 29120, June 30, 1992; Amdt. 25–88, 61 FR 57958, compartment main lighting system Nov. 8, 1996; Amdt. 25–116, 69 FR 62788, Oct. should fail, emergency illumination for 27, 2004; Amdt. 25–128, 74 FR 25645, May 29, 2009] each lower deck service compartment must be automatically provided. § 25.815 Width of aisle. (b) There must be a means for two- The passenger aisle width at any way voice communication between the point between seats must equal or ex- flight deck and each lower deck service ceed the values in the following table: compartment, which remains available following loss of normal electrical Minimum passenger power generating system. aisle width (inches) (c) There must be an aural emer- Passenger seating capacity Less than 25 in. and gency alarm system, audible during 25 in. from more from floor floor normal and emergency conditions, to enable crewmembers on the flight deck 10 or less ...... 1 12 15 11 through 19 ...... 12 20 and at each required floor level emer- 20 or more ...... 15 20 gency exit to alert occupants of each 1 A narrower width not less than 9 inches may be approved lower deck service compartment of an when substantiated by tests found necessary by the emergency situation. Administrator. (d) There must be a means, readily [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967, as detectable by occupants of each lower amended by Amdt. 25–38, 41 FR 55466, Dec. 20, deck service compartment, that indi- 1976] cates when seat belts should be fastened. § 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 [Amdt. 25–15, 32 FR 13265, Sept. 20, 1967] must be a forward or aft facing seat 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 oc- For airplanes with a service compart- cupied. ment 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.

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(3) There must be a hatch capable of after reasonably probable failures or being used for evacuating persons from malfunctioning of the ventilating, the lift that is openable from inside heating, pressurization, or other sys- and outside the lift without tools, with tems and equipment. the lift in any position. (d) If accumulation of hazardous [Amdt. 25–53, 45 FR 41593, June 19, 1980; 45 FR quantities of smoke in the cockpit area 43154, June 26, 1980; Amdt. 25–110; 68 FR 36883, is reasonably probable, smoke evacu- June 19, 2003] ation must be readily accomplished, starting with full pressurization and § 25.820 Lavatory doors. without depressurizing beyond safe All lavatory doors must be designed limits. to preclude anyone from becoming (e) Except as provided in paragraph trapped inside the lavatory. If a lock- (f) of this section, means must be pro- ing mechanism is installed, it must be vided to enable the occupants of the capable of being unlocked from the following compartments and areas to outside without the aid of special tools. control the temperature and quantity [Doc. No. 2003–14193, 69 FR 24502, May 3, 2004] of ventilating air supplied to their compartment or area independently of VENTILATION AND HEATING the temperature and quantity of air supplied to other compartments and § 25.831 Ventilation. areas: (a) Under normal operating condi- (1) The flight crew compartment. tions and in the event of any probable (2) Crewmember compartments and failure conditions of any system which areas other than the flight crew com- would adversely affect the ventilating partment unless the crewmember com- air, the ventilation system must be de- partment or area is ventilated by air signed to provide a sufficient amount interchange with other compartments of uncontaminated air to enable the or areas under all operating conditions. crewmembers to perform their duties (f) Means to enable the flight crew to without undue discomfort or fatigue control the temperature and quantity and to provide reasonable passenger of ventilating air supplied to the flight comfort. For normal operating condi- crew compartment independently of tions, the ventilation system must be the temperature and quantity of ven- designed to provide each occupant with tilating air supplied to other compart- an airflow containing at least 0.55 ments are not required if all of the fol- pounds of fresh air per minute. lowing conditions are met: (b) Crew and passenger compartment air must be free from harmful or haz- (1) The total volume of the flight ardous concentrations of gases or va- crew and passenger compartments is pors. In meeting this requirement, the 800 cubic feet or less. following apply: (2) The air inlets and passages for air (1) Carbon monoxide concentrations to flow between flight crew and pas- in excess of 1 part in 20,000 parts of air senger compartments are arranged to are considered hazardous. For test pur- provide compartment temperatures poses, any acceptable carbon monoxide within 5 degrees F. of each other and detection method may be used. adequate ventilation to occupants in (2) Carbon dioxide concentration dur- both compartments. ing flight must be shown not to exceed (3) The temperature and ventilation 0.5 percent by volume (sea level equiva- controls are accessible to the flight lent) in compartments normally occu- crew. pied by passengers or crewmembers. (g) The exposure time at any given (c) There must be provisions made to temperature must not exceed the val- ensure that the conditions prescribed ues shown in the following graph after in paragraph (b) of this section are met 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. The ca- 601 through 700 ...... 8 pacity of each system is adequate if there is sufficient quantity of agent to (2) At least one hand fire extin- extinguish the fire or suppress the fire guisher must be conveniently located anywhere baggage or cargo is placed in the pilot compartment. within the cargo compartment for the (3) At least one readily accessible duration required to land and evacuate hand fire extinguisher must be avail- the airplane. able for use in each Class A or Class B cargo or baggage compartment and in [Amdt. 25–74, 56 FR 15456, Apr. 16, 1991, as amended by Doc. No. Docket FAA–2014–0001, each Class E or Class F cargo or bag- Amdt. 25–142, 81 FR 7703, Feb. 16, 2016] gage compartment that is accessible to crewmembers in flight. § 25.853 Compartment interiors. (4) At least one hand fire extin- For each compartment occupied by guisher must be located in, or readily the crew or passengers, the following accessible for use in, each galley lo- apply: cated above or below the passenger (a) Materials (including finishes or compartment. decorative surfaces applied to the ma- (5) Each hand fire extinguisher must terials) must meet the applicable test be approved. criteria prescribed in part I of appendix (6) At least one of the required fire F of this part, or other approved equiv- extinguishers located in the passenger alent methods, regardless of the pas- compartment of an airplane with a passenger capacity of the airplane. senger capacity of at least 31 and not (b) [Reserved] more than 60, and at least two of the (c) In addition to meeting the re- fire extinguishers located in the pas- quirements of paragraph (a) of this sec- senger compartment of an airplane tion, seat cushions, except those on with a passenger capacity of 61 or more flight crewmember seats, must meet must contain Halon 1211 the test requirements of part II of ap- (bromochlorodifluoromethane CBrC1 pendix F of this part, or other equiva- F2), or equivalent, as the extinguishing lent methods, regardless of the pas- agent. The type of extinguishing agent senger capacity of the airplane. used in any other extinguisher required (d) Except as provided in paragraph by this section must be appropriate for (e) of this section, the following inte- the kinds of fires likely to occur where rior components of airplanes with pas- used. senger capacities of 20 or more must (7) The quantity of extinguishing also meet the test requirements of agent used in each extinguisher re- parts IV and V of appendix F of this quired by this section must be appro- part, or other approved equivalent priate for the kinds of fires likely to method, in addition to the flamma- occur where used. bility requirements prescribed in para- (8) Each extinguisher intended for graph (a) of this section: use in a personnel compartment must (1) Interior ceiling and wall panels, be designed to minimize the hazard of other than lighting lenses and win- toxic gas concentration. dows; (b) Built-in fire extinguishers. If a (2) Partitions, other than transparent built-in fire extinguisher is provided— panels needed to enhance cabin safety; (1) Each built-in fire extinguishing (3) Galley structure, including ex- system must be installed so that— posed surfaces of stowed carts and

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standard containers and the cavity senger cabin that would be readily de- walls that are exposed when a full com- tected by a flight attendant; and plement of such carts or containers is (b) Each lavatory must be equipped not carried; and with a built-in fire extinguisher for (4) Large cabinets and cabin stowage each disposal receptacle for towels, compartments, other than underseat paper, or waste, located within the lav- stowage compartments for stowing atory. The extinguisher must be de- small items such as magazines and signed to discharge automatically into maps. each disposal receptacle upon occur- (e) The interiors of compartments, rence of a fire in that receptacle. such as pilot compartments, galleys, 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 compartments. this section, provided the interiors of 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) Each of the following 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 that is sepa- area occupied by the crew or pas- rate from, but may be attached to, the sengers, an adequate number of self- airplane structure: contained, removable ashtrays must be (1) Any Class B through Class E cargo provided in designated smoking sec- or baggage compartment, and tions for all seated occupants. (2) Any Class F cargo or baggage (g) Regardless of whether smoking is compartment, unless other means of allowed in any other part of the air- containing a fire and protecting crit- plane, lavatories must have self-con- ical systems and structure are pro- tained, removable ashtrays located vided. conspicuously on or near the entry side (c) Ceiling and sidewall liner panels of each lavatory door, except that one of Class C cargo or baggage compart- ashtray may serve more than one lava- ments, and ceiling and sidewall liner tory door if the ashtray can be seen panels in Class F cargo or baggage readily from the cabin side of each lav- compartments, if installed to meet the atory served. requirements of paragraph (b)(2) of this (h) Each receptacle used for the dis- section, must meet the test require- posal of flammable waste material ments of part III of appendix F of this must be fully enclosed, constructed of part or other approved equivalent at least fire resistant materials, and methods. must contain fires likely to occur in it (d) All other materials used in the under normal use. The capability of the construction of the cargo or baggage receptacle to contain those fires under compartment must meet the applicable all probable conditions of wear, mis- test criteria prescribed in part I of ap- alignment, and ventilation expected in pendix F of this part or other approved service must be demonstrated by test. equivalent methods. [Amdt. 25–83, 60 FR 6623, Feb. 2, 1995, as (e) No compartment may contain any amended by Amdt. 25–116, 69 FR 62788, Oct. controls, lines, equipment, or acces- 27, 2004] sories whose damage or failure would affect safe operation, unless those § 25.854 Lavatory fire protection. items are protected so that— For airplanes with a passenger capac- (1) They cannot be damaged by the ity of 20 or more: movement of cargo in the compart- (a) Each lavatory must be equipped ment, and with a smoke detector system or equiv- (2) Their breakage or failure will not alent that provides a warning light in create a fire hazard. the cockpit, or provides a warning (f) There must be means to prevent light or audible warning in the pas- cargo or baggage from interfering with

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the functioning of the fire protective ment does not apply to thermal/acous- features of the compartment. tic insulation installations that the (g) Sources of heat within the com- FAA finds would not contribute to fire partment must be shielded and insu- penetration resistance. lated to prevent igniting the cargo or [Amdt. 25–111, 68 FR 45059, July 31, 2003] baggage. (h) Flight tests must be conducted to § 25.857 Cargo compartment classifica- show compliance with the provisions of tion. § 25.857 concerning— (a) Class A; A Class A cargo or bag- (1) Compartment accessibility, gage compartment is one in which— (2) The entries of hazardous quan- (1) The presence of a fire would be tities of smoke or extinguishing agent easily discovered by a crewmember into compartments occupied by the while at his station; and crew or passengers, and (2) Each part of the compartment is (3) The dissipation of the extin- easily accessible in flight. guishing agent in all Class C compart- (b) Class B. A Class B cargo or bag- ments and, if applicable, in any Class F gage compartment is one in which— compartments. (1) There is sufficient access in flight (i) During the above tests, it must be to enable a crewmember, standing at shown that no inadvertent operation of any one access point and without step- smoke or fire detectors in any com- ping into the compartment, to extin- partment would occur as a result of guish a fire occurring in any part of fire contained in any other compart- the compartment using a hand fire ex- ment, either during or after extin- tinguisher; guishment, unless the extinguishing (2) When the access provisions are system floods each such compartment being used, no hazardous quantity of simultaneously. smoke, flames, or extinguishing agent, (j) Cargo or baggage compartment will enter any compartment occupied electrical wiring interconnection sys- by the crew or passengers; tem components must meet the re- (3) There is a separate approved quirements of § 25.1721. smoke detector or fire detector system [Amdt. 25–72, 55 FR 29784, July 20, 1990, as to give warning at the pilot or flight amended by Amdt. 25–93, 63 FR 8048, Feb. 17, engineer station. 1998; Amdt. 25–116, 69 FR 62788, Oct. 27, 2004; (c) Class C. A Class C cargo or bag- Amdt. 25–123, 72 FR 63405, Nov. 8, 2007; Doc. gage compartment is one not meeting No. Docket FAA–2014–0001, Amdt. 25–142, 81 the requirements for either a Class A FR 7704, Feb. 16, 2016] or B compartment but in which— (1) There is a separate approved § 25.856 Thermal/Acoustic insulation smoke detector or fire detector system materials. to give warning at the pilot or flight (a) Thermal/acoustic insulation ma- engineer station; terial installed in the fuselage must (2) There is an approved built-in fire meet the flame propagation test re- extinguishing or suppression system quirements of part VI of Appendix F to controllable from the cockpit. this part, or other approved equivalent (3) There are means to exclude haz- test requirements. This requirement ardous quantities of smoke, flames, or does not apply to ‘‘small parts,’’ as de- extinguishing agent, from any com- fined in part I of Appendix F of this partment occupied by the crew or pas- part. sengers; (b) For airplanes with a passenger ca- (4) There are means to control ven- pacity of 20 or greater, thermal/acous- tilation and drafts within the compart- tic insulation materials (including the ment so that the extinguishing agent means of fastening the materials to the used can control any fire that may fuselage) installed in the lower half of start within the compartment. the airplane fuselage must meet the (d) [Reserved] flame penetration resistance test re- (e) Class E. A Class E cargo compart- quirements of part VII of Appendix F ment is one on airplanes used only for to this part, or other approved equiva- the carriage of cargo and in which— lent test requirements. This require- (1) [Reserved]

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(2) There is a separate approved operating configurations and condi- smoke or fire detector system to give tions. warning at the pilot or flight engineer [Amdt. 25–54, 45 FR 60173, Sept. 11, 1980, as station; amended by Amdt. 25–93, 63 FR 8048, Feb. 17, (3) There are means to shut off the 1998] ventilating airflow to, or within, the compartment, and the controls for § 25.859 Combustion heater fire pro- these means are accessible to the flight tection. crew in the crew compartment; (a) Combustion heater fire zones. The (4) There are means to exclude haz- following combustion heater fire zones ardous quantities of smoke, flames, or must be protected from fire in accord- noxious gases, from the flight crew ance with the applicable provisions of compartment; and §§ 25.1181 through 25.1191 and §§ 25.1195 (5) The required crew emergency through 25.1203; exits are accessible under any cargo (1) The region surrounding the heat- loading condition. er, if this region contains any flam- (f) Class F. A Class F cargo or bag- mable fluid system components (ex- gage compartment must be located on cluding the heater fuel system), that the main deck and is one in which— could— (1) There is a separate approved (i) Be damaged by heater malfunc- smoke detector or fire detector system tioning; or to give warning at the pilot or flight (ii) Allow flammable fluids or vapors engineer station; to reach the heater in case of leakage. (2) There are means to extinguish or (2) The region surrounding the heat- control a fire without requiring a crew- er, if the heater fuel system has fit- member to enter the compartment; and tings that, if they leaked, would allow (3) There are means to exclude haz- fuel or vapors to enter this region. ardous quantities of smoke, flames, or (3) The part of the ventilating air extinguishing agent from any compart- passage that surrounds the combustion ment occupied by the crew or pas- chamber. However, no fire extinguish- sengers. ment is required in cabin ventilating [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as air passages. amended by Amdt. 25–32, 37 FR 3972, Feb. 24, (b) Ventilating air ducts. Each ven- 1972; Amdt. 25–60, 51 FR 18243, May 16, 1986; tilating air duct passing through any Amdt. 25–93, 63 FR 8048, Feb. 17, 1998; Doc. fire zone must be fireproof. In addi- No. Docket FAA–2014–0001, Amdt. 25–142, 81 FR 7704, Feb. 16, 2016] tion— (1) Unless isolation is provided by § 25.858 Cargo or baggage compart- fireproof valves or by equally effective ment smoke or fire detection sys- means, the ventilating air duct down- tems. stream of each heater must be fireproof If certification with cargo or baggage for a distance great enough to ensure compartment smoke or fire detection that any fire originating in the heater provisions is requested, the following can be contained in the duct; and must be met for each cargo or baggage (2) Each part of any ventilating duct compartment with those provisions: passing through any region having a (a) The detection system must pro- flammable fluid system must be con- vide a visual indication to the flight structed or isolated from that system crew within one minute after the start so that the malfunctioning of any com- of a fire. ponent of that system cannot intro- (b) The system must be capable of de- duce flammable fluids or vapors into tecting a fire at a temperature signifi- the ventilating airstream. cantly below that at which the struc- (c) Combustion air ducts. Each com- tural integrity of the airplane is sub- bustion air duct must be fireproof for a stantially decreased. distance great enough to prevent dam- (c) There must be means to allow the age from backfiring or reverse flame crew to check in flight, the functioning propagation. In addition— of each fire detector circuit. (1) No combustion air duct may have (d) The effectiveness of the detection a common opening with the ventilating system must be shown for all approved airstream unless flames from backfires

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or reverse burning cannot enter the §§ 25.1121 and 25.1123. In addition, there ventilating airstream under any oper- must be provisions in the design of the ating condition, including reverse flow heater exhaust system to safely expel or malfunctioning of the heater or its the products of combustion to prevent associated components; and the occurrence of— (2) No combustion air duct may re- (1) Fuel leakage from the exhaust to strict the prompt relief of any backfire surrounding compartments; that, if so restricted, could cause heat- (2) Exhaust gas impingement on sur- er failure. rounding equipment or structure; (d) Heater controls; general. Provision (3) Ignition of flammable fluids by must be made to prevent the hazardous the exhaust, if the exhaust is in a com- accumulation of water or ice on or in partment containing flammable fluid any heater control component, control lines; and system tubing, or safety control. (4) Restriction by the exhaust of the (e) Heater safety controls. For each prompt relief of backfires that, if so re- combustion heater there must be the stricted, could cause heater failure. following safety control means: (h) Heater fuel systems. Each heater (1) Means independent of the compo- fuel system must meet each power- nents provided for the normal contin- plant fuel system requirement affect- uous control of air temperature, air- ing safe heater operation. Each heater flow, and fuel flow must be provided, fuel system component within the ven- for each heater, to automatically shut tilating airstream must be protected off the ignition and fuel supply to that by shrouds so that no leakage from heater at a point remote from that those components can enter the ven- heater when any of the following oc- tilating airstream. curs: (i) Drains. There must be means to (i) The heat exchanger temperature safely drain fuel that might accumu- exceeds safe limits. late within the combustion chamber or (ii) The ventilating air temperature the heat exchanger. In addition— exceeds safe limits. (1) Each part of any drain that oper- (iii) The combustion airflow becomes ates at high temperatures must be pro- inadequate for safe operation. tected in the same manner as heater (iv) The ventilating airflow becomes exhausts; and inadequate for safe operation. (2) Each drain must be protected (2) The means of complying with from hazardous ice accumulation under paragraph (e)(1) of this section for any any operating condition. individual heater must— [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (i) Be independent of components amended by Amdt. 25–11, 32 FR 6912, May 5, serving any other heater whose heat 1967; Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] output is essential for safe operation; and § 25.863 Flammable fluid fire protec- (ii) Keep the heater off until re- tion. started by the crew. (a) In each area where flammable (3) There must be means to warn the fluids or vapors might escape by leak- crew when any heater whose heat out- age of a fluid system, there must be put is essential for safe operation has means to minimize the probability of been shut off by the automatic means ignition of the fluids and vapors, and prescribed in paragraph (e)(1) of this the resultant hazards if ignition does section. occur. (f) Air intakes. Each combustion and (b) Compliance with paragraph (a) of ventilating air intake must be located this section must be shown by analysis so that no flammable fluids or vapors or tests, and the following factors must can enter the heater system under any be considered: operating condition— (1) Possible sources and paths of fluid (1) During normal operation; or leakage, and means of detecting leak- (2) As a result of the malfunctioning age. of any other component. (2) Flammability characteristics of (g) Heater exhaust. Heater exhaust fluids, including effects of any combus- systems must meet the provisions of tible or absorbing materials.

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(3) Possible ignition sources, includ- smoke protection requirements of ing electrical faults, overheating of §§ 25.831(c) and 25.863. equipment, and malfunctioning of pro- (2) Equipment that is located in des- tective devices. ignated fire zones and is used during (4) Means available for controlling or emergency procedures must be at least extinguishing a fire, such as stopping fire resistant. flow of fluids, shutting down equip- (3) EWIS components must meet the ment, fireproof containment, or use of requirements of § 25.1713. extinguishing agents. (b) Each vacuum air system line and (5) Ability of airplane components fitting on the discharge side of the that are critical to safety of flight to pump that might contain flammable withstand fire and heat. vapors or fluids must meet the require- (c) If action by the flight crew is re- ments of § 25.1183 if the line or fitting is quired to prevent or counteract a fluid in a designated fire zone. Other vacuum fire (e.g., equipment shutdown or actu- air systems components in designated ation of a fire extinguisher) quick acting means must be provided to alert fire zones must be at least fire resist- the crew. ant. (d) Each area where flammable fluids (c) Oxygen equipment and lines or vapors might escape by leakage of a must— fluid system must be identified and de- (1) Not be located in any designated fined. fire zone, (2) Be protected from heat that may [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970, as be generated in, or escape from, any amended by Amdt. 25–46, 43 FR 50597, Oct. 30, 1978] designated fire zone, and (3) Be installed so that escaping oxy- § 25.865 Fire protection of flight con- gen cannot cause ignition of grease, trols, engine mounts, and other fluid, or vapor accumulations that are flight structure. present in normal operation or as a re- Essential flight controls, engine sult of failure or malfunction of any mounts, and other flight structures lo- system. cated in designated fire zones or in adjacent areas which would be subjected [Amdt. 25–72, 55 FR 29784, July 20, 1990, as amended by Amdt. 25–113, 69 FR 12530, Mar. to the effects of fire in the fire zone 16, 2004; Amdt. 25–123, 72 FR 63405, Nov. 8, must be constructed of fireproof mate- 2007] rial or shielded so that they are capable of withstanding the effects of fire. MISCELLANEOUS [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] § 25.871 Leveling means. § 25.867 Fire protection: other compo- There must be means for determining nents. when the airplane is in a level position (a) Surfaces to the rear of the na- on the ground. celles, within one nacelle diameter of the nacelle centerline, must be at least [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] fire-resistant. § 25.875 Reinforcement near propel- (b) Paragraph (a) of this section does lers. not apply to tail surfaces to the rear of the nacelles that could not be readily (a) Each part of the airplane near the affected by heat, flames, or sparks propeller tips must be strong and stiff coming from a designated fire zone or enough to withstand the effects of the engine compartment of any nacelle. induced vibration and of ice thrown from the propeller. [Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] (b) No window may be near the pro- § 25.869 Fire protection: systems. peller tips unless it can withstand the most severe ice impact likely to occur. (a) Electrical system components: (1) Components of the electrical system must meet the applicable fire and

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§ 25.899 Electrical bonding and protec- function or probable combination of tion against static electricity. failures will jeopardize the safe oper- (a) Electrical bonding and protection ation of the airplane except that the against static electricity must be de- failure of structural elements need not signed to minimize accumulation of be considered if the probability of such electrostatic charge that would cause— failure is extremely remote. (1) Human injury from electrical (d) Each auxiliary power unit instal- shock, lation must meet the applicable provi- (2) Ignition of flammable vapors, or sions of this subpart. (3) Interference with installed elec- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as trical/electronic equipment. amended by Amdt. 25–23, 35 FR 5676, Apr. 8, (b) Compliance with paragraph (a) of 1970; Amdt. 25–40, 42 FR 15042, Mar. 17, 1977; this section may be shown by— Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. (1) Bonding the components properly 25–126, 73 FR 63345, Oct. 24, 2008] to the airframe; or (2) Incorporating other acceptable § 25.903 Engines. means to dissipate the static charge so (a) Engine type certificate. (1) Each en- as not to endanger the airplane, per- gine must have a type certificate and sonnel, or operation of the installed must meet the applicable requirements electrical/electronic systems. of part 34 of this chapter. [Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] (2) Each turbine engine must comply with one of the following: Subpart E—Powerplant (i) Sections 33.76, 33.77 and 33.78 of this chapter in effect on December 13, GENERAL 2000, or as subsequently amended; or (ii) Sections 33.77 and 33.78 of this § 25.901 Installation. chapter in effect on April 30, 1998, or as (a) For the purpose of this part, the subsequently amended before Decem- airplane powerplant installation in- ber 13, 2000; or cludes each component that— (iii) Comply with § 33.77 of this chap- (1) Is necessary for propulsion; ter in effect on October 31, 1974, or as (2) Affects the control of the major subsequently amended prior to April propulsive units; or 30, 1998, unless that engine’s foreign ob- (3) Affects the safety of the major ject ingestion service history has re- propulsive units between normal in- sulted in an unsafe condition; or spections or overhauls. (iv) Be shown to have a foreign object (b) For each powerplant— ingestion service history in similar in- (1) The installation must comply stallation locations which has not re- with— sulted in any unsafe condition. (i) The installation instructions provided under §§ 33.5 and 35.3 of this chap- NOTE: § 33.77 of this chapter in effect on Oc- tober 31, 1974, was published in 14 CFR parts ter; and 1 to 59, Revised as of January 1, 1975. See 39 (ii) The applicable provisions of this FR 35467, October 1, 1974. subpart; (2) The components of the installa- (3) Each turbine engine must comply tion must be constructed, arranged, with one of the following paragraphs: and installed so as to ensure their con- (i) Section 33.68 of this chapter in ef- tinued safe operation between normal fect on January 5, 2015, or as subse- inspections or overhauls; quently amended; or (3) The installation must be acces- (ii) Section 33.68 of this chapter in ef- sible for necessary inspections and fect on February 23, 1984, or as subse- maintenance; and quently amended before January 5, (4) The major components of the in- 2015, unless that engine’s ice accumula- stallation must be electrically bonded tion service history has resulted in an to the other parts of the airplane. unsafe condition; or (c) For each powerplant and auxiliary (iii) Section 33.68 of this chapter in power unit installation, it must be es- effect on October 1, 1974, or as subse- tablished that no single failure or mal- quently amended prior to February 23,

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1984, unless that engine’s ice accumula- a restart capability within that enve- tion service history has resulted in an lope. unsafe condition; or (3) For turbine engine powered air- (iv) Be shown to have an ice accumu- planes, if the minimum windmilling lation service history in similar instal- speed of the engines, following the lation locations which has not resulted inflight shutdown of all engines, is in- in any unsafe conditions. sufficient to provide the necessary (b) Engine isolation. The powerplants electrical power for engine ignition, a must be arranged and isolated from power source independent of the en- each other to allow operation, in at gine-driven electrical power generating least one configuration, so that the system must be provided to permit in- failure or malfunction of any engine, or flight engine ignition for restarting. of any system that can affect the en- (f) Auxiliary Power Unit. Each auxil- gine, will not— iary power unit must be approved or (1) Prevent the continued safe oper- meet the requirements of the category ation of the remaining engines; or for its intended use. (2) Require immediate action by any crewmember for continued safe oper- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ation. amended by Amdt. 25–23, 35 FR 5676, Apr. 8, (c) Control of engine rotation. There 1970; Amdt. 25–40, 42 FR 15042, Mar. 17, 1977; must be means for stopping the rota- Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; Amdt. 25–72, 55 FR 29784, July 20, 1990; Amdt. 25–73, tion of any engine individually in 55 FR 32861, Aug. 10, 1990; Amdt. 25–94, 63 FR flight, except that, for turbine engine 8848, Feb. 23, 1998; Amdt. 25–95, 63 FR 14798, installations, the means for stopping Mar. 26, 1998; Amdt. 25–100, 65 FR 55854, Sept. the rotation of any engine need be pro- 14, 2000; Amdt. 25–140, 79 FR 65525, Nov. 4, vided only where continued rotation 2014] could jeopardize the safety of the airplane. Each component of the stopping § 25.904 Automatic takeoff thrust con- system on the engine side of the fire- trol system (ATTCS). wall that might be exposed to fire must Each applicant seeking approval for be at least fire-resistant. If hydraulic installation of an engine power control propeller feathering systems are used system that automatically resets the for this purpose, the feathering lines power or thrust on the operating en- must be at least fire resistant under gine(s) when any engine fails during the operating conditions that may be expected to exist during feathering. the takeoff must comply with the requirements of appendix I of this part. (d) Turbine engine installations. For turbine engine installations— [Amdt. 25–62, 52 FR 43156, Nov. 9, 1987] (1) Design precautions must be taken to minimize the hazards to the airplane § 25.905 Propellers. in the event of an engine rotor failure (a) Each propeller must have a type or of a fire originating within the en- certificate. gine which burns through the engine (b) Engine power and propeller shaft case. rotational speed may not exceed the (2) The powerplant systems associ- limits for which the propeller is certifi- ated with engine control devices, sys- cated. tems, and instrumentation, must be de- (c) The propeller blade pitch control signed to give reasonable assurance that those engine operating limitations system must meet the requirements of that adversely affect turbine rotor §§ 35.21, 35.23, 35.42 and 35.43 of this structural integrity will not be exceed- chapter. ed in service. (d) Design precautions must be taken (e) Restart capability. (1) Means to re- to minimize the hazards to the airplane start any engine in flight must be pro- in the event a propeller blade fails or is vided. released by a hub failure. The hazards (2) An altitude and airspeed envelope which must be considered include dam- must be established for in-flight engine age to structure and vital systems due restarting, and each engine must have to impact of a failed or released blade

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and the unbalance created by such fail- (2) The effects of airplane and pro- ure or release. peller operating and airworthiness limitations. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–54, 45 FR 60173, Sept. [Amdt. 25–126, 73 FR 63345, Oct. 24, 2008] 11, 1980; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; Amdt. 25–72, 55 FR 29784, July 20, 1990; Amdt. § 25.925 Propeller clearance. 25–126, 73 FR 63345, Oct. 24, 2008] Unless smaller clearances are substantiated, propeller clearances with § 25.907 Propeller vibration and fa- the airplane at maximum weight, with tigue. the most adverse center of gravity, and This section does not apply to fixed- with the propeller in the most adverse pitch wood propellers of conventional pitch position, may not be less than design. the following: (a) The applicant must determine the (a) Ground clearance. There must be a magnitude of the propeller vibration clearance of at least seven inches (for stresses or loads, including any stress each airplane with nose wheel landing peaks and resonant conditions, gear) or nine inches (for each airplane throughout the operational envelope of with tail wheel landing gear) between the airplane by either: each propeller and the ground with the (1) Measurement of stresses or loads landing gear statically deflected and in through direct testing or analysis the level takeoff, or taxiing attitude, based on direct testing of the propeller whichever is most critical. In addition, on the airplane and engine installation there must be positive clearance be- for which approval is sought; or tween the propeller and the ground when in the level takeoff attitude with (2) Comparison of the propeller to the critical tire(s) completely deflated similar propellers installed on similar and the corresponding landing gear airplane installations for which these strut bottomed. measurements have been made. (b) Water clearance. There must be a (b) The applicant must demonstrate clearance of at least 18 inches between by tests, analysis based on tests, or each propeller and the water, unless previous experience on similar designs compliance with § 25.239(a) can be that the propeller does not experience shown with a lesser clearance. harmful effects of flutter throughout (c) Structural clearance. There must the operational envelope of the air- be— plane. (1) At least one inch radial clearance (c) The applicant must perform an between the blade tips and the airplane evaluation of the propeller to show structure, plus any additional radial that failure due to fatigue will be clearance necessary to prevent harmful avoided throughout the operational life vibration; of the propeller using the fatigue and (2) At least one-half inch longitudinal structural data obtained in accordance clearance between the propeller blades with part 35 of this chapter and the vi- or cuffs and stationary parts of the air- bration data obtained from compliance plane; and with paragraph (a) of this section. For (3) Positive clearance between other the purpose of this paragraph, the pro- rotating parts of the propeller or spin- peller includes the hub, blades, blade ner and stationary parts of the air- retention component and any other plane. propeller component whose failure due [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as to fatigue could be catastrophic to the amended by Amdt. 25–72, 55 FR 29784, July 20, airplane. This evaluation must include: 1990] (1) The intended loading spectra including all reasonably foreseeable pro- § 25.929 Propeller deicing. peller vibration and cyclic load pat- (a) If certification for flight in icing terns, identified emergency conditions, is sought there must be a means to pre- allowable overspeeds and overtorques, vent or remove hazardous ice accumu- and the effects of temperatures and hu- lations that could form in the icing midity expected in service. conditions defined in Appendix C of

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this part and in the portions of Appen- ered if this kind of failure is extremely dix O of this part for which the air- remote. plane is approved for flight on propel- (2) Compliance with this section may lers or on accessories where ice accu- be shown by failure analysis or testing, mulation would jeopardize engine per- or both, for propeller systems that formance. allow propeller blades to move from (b) If combustible fluid is used for the flight low-pitch position to a posi- propeller deicing, §§ 25.1181 through tion that is substantially less than 25.1185 and 25.1189 apply. that at the normal flight low-pitch po- [ Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as sition. The analysis may include or be amended by Amdt. 25–140, 79 FR 65525, Nov. 4, supported by the analysis made to 2014] show compliance with the requirements of § 35.21 of this chapter for the § 25.933 Reversing systems. propeller and associated installation (a) For turbojet reversing systems— components. (1) Each system intended for ground [Amdt. 25–72, 55 FR 29784, July 20, 1990] operation only must be designed so that during any reversal in flight the § 25.934 Turbojet engine thrust re- engine will produce no more than flight verser system tests. idle thrust. In addition, it must be Thrust reversers installed on tur- shown by analysis or test, or both, bojet engines must meet the require- that— ments of § 33.97 of this chapter. (i) Each operable reverser can be re- stored to the forward thrust position; [Amdt. 25–23, 35 FR 5677, Apr. 8, 1970] and (ii) The airplane is capable of contin- § 25.937 Turbopropeller-drag limiting systems. ued safe flight and landing under any possible position of the thrust reverser. Turbopropeller power airplane pro- (2) Each system intended for inflight peller-drag limiting systems must be use must be designed so that no unsafe designed so that no single failure or condition will result during normal op- malfunction of any of the systems dur- eration of the system, or from any fail- ing normal or emergency operation re- ure (or reasonably likely combination sults in propeller drag in excess of that of failures) of the reversing system, for which the airplane was designed under any anticipated condition of op- under § 25.367. Failure of structural ele- eration of the airplane including ments of the drag limiting systems ground operation. Failure of structural need not be considered if the prob- elements need not be considered if the ability of this kind of failure is ex- probability of this kind of failure is extremely remote. tremely remote. (3) Each system must have means to § 25.939 Turbine engine operating characteristics. prevent the engine from producing more than idle thrust when the revers- (a) Turbine engine operating charac- ing system malfunctions, except that it teristics must be investigated in flight may produce any greater forward to determine that no adverse charac- thrust that is shown to allow direc- teristics (such as stall, surge, or flame- tional control to be maintained, with out) are present, to a hazardous degree, aerodynamic means alone, under the during normal and emergency oper- most critical reversing condition ex- ation within the range of operating pected in operation. limitations of the airplane and of the (b) For propeller reversing systems— engine. (1) Each system intended for ground (b) [Reserved] operation only must be designed so (c) The turbine engine air inlet sys- that no single failure (or reasonably tem may not, as a result of air flow dis- likely combination of failures) or mal- tortion during normal operation, cause function of the system will result in vibration harmful to the engine. unwanted reverse thrust under any ex- [Amdt. 25–11, 32 FR 6912, May 5, 1967, as pected operating condition. Failure of amended by Amdt. 25–40, 42 FR 15043, Mar. 17, structural elements need not be consid- 1977]

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§ 25.941 Inlet, engine, and exhaust (b) Fluid tanks. Each augmentation compatibility. system fluid tank must meet the following requirements: For airplanes using variable inlet or (1) Each tank must be able to with- exhaust system geometry, or both— stand without failure the vibration, in- (a) The system comprised of the ertia, fluid, and structural loads that it inlet, engine (including thrust aug- may be subject to in operation. mentation systems, if incorporated), (2) The tanks as mounted in the air- and exhaust must be shown to function plane must be able to withstand with- properly under all operating conditions out failure or leakage an internal pres- for which approval is sought, including sure 1.5 times the maximum operating all engine rotating speeds and power pressure. settings, and engine inlet and exhaust (3) If a vent is provided, the venting configurations; must be effective under all normal (b) The dynamic effects of the oper- flight conditions. ation of these (including consideration (4) [Reserved] of probable malfunctions) upon the aer- (5) Each tank must have an expan- odynamic control of the airplane may sion space of not less than 2 percent of not result in any condition that would the tank capacity. It must be impos- require exceptional skill, alertness, or sible to fill the expansion space inad- strength on the part of the pilot to vertently with the airplane in the nor- avoid exceeding an operational or mal ground attitude. structural limitation of the airplane; (c) Augmentation system drains and must be designed and located in ac- (c) In showing compliance with para- cordance with § 25.1455 if— graph (b) of this section, the pilot (1) The augmentation system fluid is strength required may not exceed the subject to freezing; and limits set forth in § 25.143(d), subject to (2) The fluid may be drained in flight the conditions set forth in paragraphs or during ground operation. (e) and (f) of § 25.143. (d) The augmentation liquid tank capacity available for the use of each en- [Amdt. 25–38, 41 FR 55467, Dec. 20, 1976, as gine must be large enough to allow op- amended by Amdt. 25–121, 72 FR 44669, Aug. 8, eration of the airplane under the ap- 2007] proved procedures for the use of liquid- augmented power. The computation of § 25.943 Negative acceleration. liquid consumption must be based on No hazardous malfunction of an en- the maximum approved rate appro- gine, an auxiliary power unit approved priate for the desired engine output for use in flight, or any component or and must include the effect of tempera- system associated with the powerplant ture on engine performance as well as or auxiliary power unit may occur any other factors that might vary the when the airplane is operated at the amount of liquid required. negative accelerations within the (e) This section does not apply to fuel flight envelopes prescribed in § 25.333. injection systems. This must be shown for the greatest [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977, as duration expected for the acceleration. amended by Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–115, 69 FR 40527, July 2, 2004] [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] FUEL SYSTEM § 25.945 Thrust or power augmentation system. § 25.951 General. (a) General. Each fluid injection sys- (a) Each fuel system must be con- tem must provide a flow of fluid at the structed and arranged to ensure a flow rate and pressure established for proper of fuel at a rate and pressure estab- engine functioning under each intended lished for proper engine and auxiliary operating condition. If the fluid can power unit functioning under each freeze, fluid freezing may not damage likely operating condition, including the airplane or adversely affect air- any maneuver for which certification is plane performance. requested and during which the engine

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or auxiliary power unit is permitted to feature or structure, could create an be in operation. ignition source. (b) Each fuel system must be ar- (2) A fuel system includes any compo- ranged so that any air which is intro- nent within either the fuel tank struc- duced into the system will not result ture or the fuel tank systems, and any in— airplane structure or system compo- (1) Power interruption for more than nents that penetrate, connect to, or are 20 seconds for reciprocating engines; or located within a fuel tank. (2) Flameout for turbine engines. (b) The design and installation of a (c) Each fuel system for a turbine en- fuel system must prevent catastrophic gine must be capable of sustained oper- fuel vapor ignition due to lightning and ation throughout its flow and pressure its effects, including: range with fuel initially saturated with (1) Direct lightning strikes to areas water at 80 °F and having 0.75cc of free having a high probability of stroke at- water per gallon added and cooled to tachment; the most critical condition for icing (2) Swept lightning strokes to areas likely to be encountered in operation. where swept strokes are highly prob- (d) Each fuel system for a turbine enable; and gine powered airplane must meet the (3) Lightning-induced or conducted applicable fuel venting requirements of electrical transients. part 34 of this chapter. (c) To comply with paragraph (b) of this section, catastrophic fuel vapor ig- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5677, Apr. 8, nition must be extremely improbable, 1970; Amdt. 25–36, 39 FR 35460, Oct. 1, 1974; taking into account flammability, crit- Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. ical lightning strikes, and failures 25–73, 55 FR 32861, Aug. 10, 1990] within the fuel system. (d) To protect design features that § 25.952 Fuel system analysis and test. prevent catastrophic fuel vapor igni- (a) Proper fuel system functioning tion caused by lightning, the type de- under all probable operating conditions sign must include critical design con- must be shown by analysis and those figuration control limitations tests found necessary by the Adminis- (CDCCLs) identifying those features trator. Tests, if required, must be made and providing information to protect using the airplane fuel system or a test them. To ensure the continued effec- article that reproduces the operating tiveness of those design features, the characteristics of the portion of the type design must also include inspec- fuel system to be tested. tion and test procedures, intervals be- (b) The likely failure of any heat ex- tween repetitive inspections and tests, changer using fuel as one of its fluids and mandatory replacement times for may not result in a hazardous condi- those design features used in dem- tion. onstrating compliance to paragraph (b) [Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] of this section. The applicant must include the information required by this § 25.953 Fuel system independence. paragraph in the Airworthiness Limi- tations section of the Instructions for Each fuel system must meet the re- Continued Airworthiness required by quirements of § 25.903(b) by— § 25.1529. (a) Allowing the supply of fuel to each engine through a system inde- [Doc. No. FAA–2014–1027, Amdt. 25–146, 83 FR pendent of each part of the system sup- 47556, Sept. 20, 2018] plying fuel to any other engine; or (b) Any other acceptable method. § 25.955 Fuel flow. (a) Each fuel system must provide at § 25.954 Fuel system lightning protec- least 100 percent of the fuel flow re- tion. quired under each intended operating (a) For purposes of this section— condition and maneuver. Compliance (1) A critical lightning strike is a must be shown as follows: lightning strike that attaches to the (1) Fuel must be delivered to each en- airplane in a location that, when com- gine at a pressure within the limits bined with the failure of any design specified in the engine type certificate.

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(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 be substituted for each main pump so (a) The fuel system must perform used. satisfactorily in hot weather operation. (4) If there is a fuel flowmeter, it This must be shown by showing that must be blocked and the fuel must flow the fuel system from the tank outlets through the meter or its bypass. to each engine is pressurized, under all (b) If an engine can be supplied with intended operations, so as to prevent fuel from more than one tank, the fuel vapor formation, or must be shown by system must— climbing from the altitude of the air- (1) For each reciprocating engine, port elected by the applicant to the supply the full fuel pressure to that en- maximum altitude established as an gine in not more than 20 seconds after operating limitation under § 25.1527. If a switching to any other fuel tank con- climb test is elected, there may be no taining usable fuel when engine mal- evidence of vapor lock or other malfunctioning becomes apparent due to functioning during the climb test con- the depletion of the fuel supply in any ducted under the following conditions: tank from which the engine can be fed; (1) For reciprocating engine powered and airplanes, the engines must operate at (2) For each turbine engine, in addi- maximum continuous power, except tion to having appropriate manual that takeoff power must be used for the switching capability, be designed to altitudes from 1,000 feet below the crit- prevent interruption of fuel flow to ical altitude through the critical alti- that engine, without attention by the tude. The time interval during which flight crew, when any tank supplying takeoff power is used may not be less fuel to that engine is depleted of usable than the takeoff time limitation. fuel during normal operation, and any (2) For turbine engine powered air- other tank, that normally supplies fuel planes, the engines must operate at to that engine alone, contains usable takeoff power for the time interval se- fuel. lected for showing the takeoff flight 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, 1967] (3) The weight of the airplane must be the weight with full fuel tanks, min- § 25.957 Flow between interconnected imum crew, and the ballast necessary tanks. to maintain the center of gravity with- If fuel can be pumped from one tank in allowable limits. to another in flight, the fuel tank (4) The climb airspeed may not ex- vents and the fuel transfer system ceed— must be designed so that no structural (i) For reciprocating engine powered damage to the tanks can occur because airplanes, the maximum airspeed es- of overfilling. tablished for climbing from takeoff to the maximum operating altitude with § 25.959 Unusable fuel supply. the airplane in the following configura- The unusable fuel quantity for each tion: fuel tank and its fuel system compo- (A) Landing gear retracted. nents must be established at not less (B) Wing flaps in the most favorable than the quantity at which the first position. evidence of engine malfunction occurs (C) Cowl flaps (or other means of con- under the most adverse fuel feed condi- trolling the engine cooling supply) in

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the position that provides adequate L = a reference distance between the point of cooling in the hot-day condition. pressure and the tank farthest boundary (D) Engine operating within the max- in the direction of loading K = 4.5 for the forward loading condition for imum continuous power limitations. those parts of fuel tanks outside the fu- (E) Maximum takeoff weight; and selage pressure boundary (ii) For turbine engine powered air- K = 9 for the forward loading condition for planes, the maximum airspeed estab- those parts of fuel tanks within the fuse- lished for climbing from takeoff to the lage pressure boundary, or that form part maximum operating altitude. of the fuselage pressure boundary K = 1.5 for the aft loading condition (5) The fuel temperature must be at K = 3.0 for the inboard and outboard loading least 110 °F. conditions for those parts of fuel tanks (b) The test prescribed in paragraph within the fuselage pressure boundary, or (a) of this section may be performed in that form part of the fuselage pressure flight or on the ground under closely boundary simulated flight conditions. If a flight K = 1.5 for the inboard and outboard loading test is performed in weather cold conditions for those parts of fuel tanks outside the fuselage pressure boundary enough to interfere with the proper K = 6 for the downward loading condition conduct of the test, the fuel tank sur- K = 3 for the upward loading condition faces, fuel lines, and other fuel system parts subject to cold air must be insu- (2) For those parts of wing fuel tanks lated to simulate, insofar as prac- near the fuselage or near the engines, ticable, flight in hot weather. the greater of the fuel pressures resulting from paragraphs (d)(2)(i) or (d)(2)(ii) [Amdt. 25–11, 32 FR 6912, May 5, 1967, as of this section must be used: amended by Amdt. 25–57, 49 FR 6848, Feb. 23, (i) The fuel pressures resulting from 1984] paragraph (d)(1) of this section, and (ii) The lesser of the two following § 25.963 Fuel tanks: general. conditions: (a) Each fuel tank must be able to (A) Fuel pressures resulting from the withstand, without failure, the vibra- accelerations specified in § 25.561(b)(3) tion, inertia, fluid, and structural loads considering the fuel tank full of fuel at that it may be subjected to in oper- maximum fuel density. Fuel pressures ation. based on the 9.0g forward acceleration (b) Flexible fuel tank liners must be may be calculated using the fuel static approved or must be shown to be suit- head equal to the streamwise local able for the particular application. chord of the tank. For inboard and out- (c) Integral fuel tanks must have fa- board conditions, an acceleration of cilities for interior inspection and re- 1.5g may be used in lieu of 3.0g as speci- pair. fied in § 25.561(b)(3). (d) Fuel tanks must, so far as it is (B) Fuel pressures resulting from the practicable, be designed, located, and accelerations as specified in installed so that no fuel is released in § 25.561(b)(3) considering a fuel volume or near the fuselage, or near the en- beyond 85 percent of the maximum per- gines, in quantities that would con- missible volume in each tank using the stitute a fire hazard in otherwise sur- static head associated with the 85 per- vivable emergency landing conditions, cent fuel level. A typical density of the and— appropriate fuel may be used. For in- (1) Fuel tanks must be able to resist board and outboard conditions, an ac- rupture and retain fuel under ultimate celeration of 1.5g may be used in lieu of hydrostatic design conditions in which 3.0g as specified in § 25.561(b)(3). the pressure P within the tank varies (3) Fuel tank internal barriers and in accordance with the formula: baffles may be considered as solid boundaries if shown to be effective in P = KrgL limiting fuel flow. Where— (4) For each fuel tank and sur- P = fuel pressure at each point within the rounding airframe structure, the ef- tank fects of crushing and scraping actions r = typical fuel density with the ground must not cause the g = acceleration due to gravity spillage of enough fuel, or generate

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temperatures that would constitute a cause fuel leakage, must be able to fire hazard under the conditions speci- withstand the following test, or its fied in § 25.721(b). equivalent, without leakage or exces- (5) Fuel tank installations must be sive deformation of the tank walls: such that the tanks will not rupture as (1) Each complete tank assembly and a result of the landing gear or an en- its supports must be vibration tested gine pylon or engine mount tearing while mounted to simulate the actual away as specified in § 25.721(a) and (c). 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: ments, low energy engine debris, or (i) If no frequency of vibration result- other likely debris. ing from any r.p.m. within the normal (2) All covers must be fire resistant operating range of engine speeds is as defined in part 1 of this chapter. critical, the test frequency of vibration (f) For pressurized fuel tanks, a must be 2,000 cycles per minute. means with fail-safe features must be (ii) If only one frequency of vibration provided to prevent the buildup of an resulting from any r.p.m. within the excessive pressure difference between normal operating range of engine the inside and the outside of the tank. speeds is critical, that frequency of vi- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as bration must be the test frequency. amended by Amdt. 25–40, 42 FR 15043, Mar. 17, (iii) If more than one frequency of vi- 1977; Amdt. 25–69, 54 FR 40354, Sept. 29, 1989; bration resulting from any r.p.m. with- Amdt. 25–139, 79 FR 59430, Oct. 2, 2014] in the normal operating range of engine speeds is critical, the most crit- § 25.965 Fuel tank tests. ical of these frequencies must be the (a) It must be shown by tests that the test frequency. fuel tanks, as mounted in the airplane, (4) Under paragraphs (b)(3)(ii) and can withstand, without failure or leak- (iii) of this section, the time of test age, the more critical of the pressures must be adjusted to accomplish the resulting from the conditions specified same number of vibration cycles that in paragraphs (a)(1) and (2) of this sec- would be accomplished in 25 hours at tion. In addition, it must be shown by the frequency specified in paragraph either analysis or tests, that tank sur- (b)(3)(i) of this section. faces subjected to more critical pres- (5) During the test, the tank assem- sures resulting from the condition of bly must be rocked at the rate of 16 to paragraphs (a)(3) and (4) of this section, 20 complete cycles per minute, through are able to withstand the following an angle of 15° on both sides of the hor- pressures: izontal (30° total), about the most crit- (1) An internal pressure of 3.5 psi. ical axis, for 25 hours. If motion about (2) 125 percent of the maximum air more than one axis is likely to be crit- pressure developed in the tank from ical, the tank must be rocked about ram effect. each critical axis for 121⁄2 hours. (3) Fluid pressures developed during (c) Except where satisfactory oper- maximum limit accelerations, and de- ating experience with a similar tank in flections, of the airplane with a full a similar installation is shown, non- tank. metallic tanks must withstand the test (4) Fluid pressures developed during specified in paragraph (b)(5) of this sec- the most adverse combination of air- tion, with fuel at a temperature of 110 plane roll and fuel load. °F. During this test, a representative (b) Each metallic tank with large un- specimen of the tank must be installed supported or unstiffened flat surfaces, in a supporting structure simulating whose failure or deformation could the installation in the airplane.

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(d) For pressurized fuel tanks, it tion may be shown with the means pro- must be shown by analysis or tests vided to comply with § 25.979(b). that the fuel tanks can withstand the [Amdt. 25–11, 32 FR 6913, May 5, 1967] maximum pressure likely to occur on the ground or in flight. § 25.971 Fuel tank sump. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (a) Each fuel tank must have a sump amended by Amdt. 25–11, 32 FR 6913, May 5, with an effective capacity, in the nor- 1967; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977] mal ground attitude, of not less than the greater of 0.10 percent of the tank § 25.967 Fuel tank installations. capacity or one-sixteenth of a gallon (a) Each fuel tank must be supported unless operating limitations are estab- so that tank loads (resulting from the lished to ensure that the accumulation weight of the fuel in the tanks) are not of water in service will not exceed the concentrated on unsupported tank sur- sump capacity. faces. In addition— (b) Each fuel tank must allow drain- (1) There must be pads, if necessary, age of any hazardous quantity of water to prevent chafing between the tank from any part of the tank to its sump and its supports; with the airplane in the ground atti- (2) Padding must be nonabsorbent or tude. treated to prevent the absorption of (c) Each fuel tank sump must have fluids; an accessible drain that— (3) If a flexible tank liner is used, it (1) Allows complete drainage of the must be supported so that it is not re- sump on the ground; quired to withstand fluid loads; and (2) Discharges clear of each part of (4) Each interior surface of the tank the airplane; and compartment must be smooth and free (3) Has manual or automatic means of projections that could cause wear of for positive locking in the closed posi- the liner unless— tion. (i) Provisions are made for protection of the liner at these points; or § 25.973 Fuel tank filler connection. (ii) The construction of the liner Each fuel tank filler connection must itself provides that protection. prevent the entrance of fuel into any (b) Spaces adjacent to tank surfaces part of the airplane other than the must be ventilated to avoid fume accu- tank itself. In addition— mulation due to minor leakage. If the (a) [Reserved] tank is in a sealed compartment, ven- (b) Each recessed filler connection tilation may be limited to drain holes that can retain any appreciable quan- large enough to prevent excessive pres- tity of fuel must have a drain that dis- sure resulting from altitude changes. charges clear of each part of the air- (c) The location of each tank must plane; meet the requirements of § 25.1185(a). (c) Each filler cap must provide a (d) No engine nacelle skin imme- fuel-tight seal; and diately behind a major air outlet from (d) Each fuel filling point must have the engine compartment may act as a provision for electrically bonding the the wall of an integral tank. airplane to ground fueling equipment. (e) Each fuel tank must be isolated [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as from personnel compartments by a amended by Amdt. 25–40, 42 FR 15043, Mar. 17, fumeproof and fuelproof enclosure. 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–115, 69 FR 40527, July 2, 2004] § 25.969 Fuel tank expansion space. Each fuel tank must have an expan- § 25.975 Fuel tank vents and carbu- sion space of not less than 2 percent of retor vapor vents. the tank capacity. It must be impos- (a) Fuel tank vents. Each fuel tank sible to fill the expansion space inad- must be vented from the top part of the vertently with the airplane in the nor- expansion space so that venting is ef- mal ground attitude. For pressure fuel- fective under any normal flight condi- ing systems, compliance with this section. In addition—

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(1) Each vent must be arranged to § 25.977 Fuel tank outlet. avoid stoppage by dirt or ice forma- (a) There must be a fuel strainer for tion; the fuel tank outlet or for the booster (2) The vent arrangement must pre- pump. This strainer must— vent siphoning of fuel during normal (1) For reciprocating engine powered operation; airplanes, have 8 to 16 meshes per inch; (3) The venting capacity and vent and pressure levels must maintain accept- (2) For turbine engine powered air- able differences of pressure between planes, prevent the passage of any ob- the interior and exterior of the tank, ject that could restrict fuel flow or during— damage any fuel system component. (i) Normal flight operation; (b) [Reserved] (ii) Maximum rate of ascent and de- (c) The clear area of each fuel tank scent; and outlet strainer must be at least five (iii) Refueling and defueling (where times the area of the outlet line. applicable); (d) The diameter of each strainer (4) Airspaces of tanks with inter- must be at least that of the fuel tank connected outlets must be inter- outlet. connected; (e) Each finger strainer must be ac- (5) There may be no point in any vent cessible for inspection and cleaning. line where moisture can accumulate [Amdt. 25–11, 32 FR 6913, May 5, 1967, as with the airplane in the ground atti- amended by Amdt. 25–36, 39 FR 35460, Oct. 1, tude or the level flight attitude, unless 1974] drainage is provided; (6) No vent or drainage provision may § 25.979 Pressure fueling system. end at any point— For pressure fueling systems, the fol- (i) Where the discharge of fuel from lowing apply: the vent outlet would constitute a fire (a) Each pressure fueling system fuel hazard; or manifold connection must have means (ii) From which fumes could enter to prevent the escape of hazardous personnel compartments; and quantities of fuel from the system if (7) Each fuel tank vent system must the fuel entry valve fails. prevent explosions, for a minimum of 2 (b) An automatic shutoff means must minutes and 30 seconds, caused by be provided to prevent the quantity of propagation of flames from outside the fuel in each tank from exceeding the tank through the fuel tank vents into maximum quantity approved for that fuel tank vapor spaces when any fuel tank. This means must— tank vent is continuously exposed to (1) Allow checking for proper shutoff flame. operation before each fueling of the tank; and (b) Carburetor vapor vents. Each car- (2) Provide indication at each fueling buretor with vapor elimination connec- station of failure of the shutoff means tions must have a vent line to lead va- to stop the fuel flow at the maximum pors back to one of the fuel tanks. In quantity approved for that tank. addition— (c) A means must be provided to pre- (1) Each vent system must have vent damage to the fuel system in the means to avoid stoppage by ice; and event of failure of the automatic shut- (2) If there is more than one fuel off means prescribed in paragraph (b) tank, and it is necessary to use the of this section. tanks in a definite sequence, each (d) The airplane pressure fueling sys- vapor vent return line must lead back tem (not including fuel tanks and fuel to the fuel tank used for takeoff and tank vents) must withstand an ulti- landing. mate load that is 2.0 times the load arising from the maximum pressures, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Docket No. FAA–2014–0500, including surge, that is likely to occur Amdt. No. 25–143, 81 FR 41207, June 24, 2016] during fueling. The maximum surge pressure must be established with any

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combination of tank valves being ei- the analysis must be based on an as- ther intentionally or inadvertently sumed Equivalent Conventional closed. Unheated Aluminum Wing Tank. (e) The airplane defueling system (1) Fleet Average Flammability Ex- (not including fuel tanks and fuel tank posure is determined in accordance vents) must withstand an ultimate with Appendix N of this part. The as- load that is 2.0 times the load arising sessment must be done in accordance from the maximum permissible with the methods and procedures set defueling pressure (positive or nega- forth in the Fuel Tank Flammability tive) at the airplane fueling connec- Assessment Method User’s Manual, tion. dated May 2008, document number [Amdt. 25–11, 32 FR 6913, May 5, 1967, as DOT/FAA/AR–05/8 (incorporated by ref- amended by Amdt. 25–38, 41 FR 55467, Dec. 20, erence, see § 25.5). 1976; Amdt. 25–72, 55 FR 29785, July 20, 1990] (2) Any fuel tank other than a main fuel tank on an airplane must meet the § 25.981 Fuel tank explosion preven- flammability exposure criteria of Ap- tion. pendix M to this part if any portion of (a) No ignition source may be present the tank is located within the fuselage at each point in the fuel tank or fuel contour. tank system where catastrophic failure (3) As used in this paragraph, could occur due to ignition of fuel or (i) Equivalent Conventional Unheated vapors. This must be shown by: Aluminum Wing Tank is an integral (1) Determining the highest tempera- tank in an unheated semi-monocoque ture allowing a safe margin below the aluminum wing of a subsonic airplane lowest expected autoignition tempera- that is equivalent in aerodynamic per- ture of the fuel in the fuel tanks. formance, structural capability, fuel (2) Demonstrating that no tempera- tank capacity and tank configuration ture at each place inside each fuel tank to the designed wing. where fuel ignition is possible will ex- (ii) Fleet Average Flammability Expo- ceed the temperature determined under sure is defined in Appendix N to this paragraph (a)(1) of this section. This part and means the percentage of time must be verified under all probable op- each fuel tank ullage is flammable for erating, failure, and malfunction con- a fleet of an airplane type operating ditions of each component whose oper- over the range of flight lengths. ation, failure, or malfunction could in- (iii) Main Fuel Tank means a fuel crease the temperature inside the tank. tank that feeds fuel directly into one (3) Except for ignition sources due to or more engines and holds required fuel lightning addressed by § 25.954, dem- reserves continually throughout each onstrating that an ignition source flight. could not result from each single fail- (c) Paragraph (b) of this section does ure, from each single failure in com- not apply to a fuel tank if means are bination with each latent failure condi- provided to mitigate the effects of an tion not shown to be extremely remote, ignition of fuel vapors within that fuel and from all combinations of failures tank such that no damage caused by an not shown to be extremely improbable, ignition will prevent continued safe taking into account the effects of man- flight and landing. ufacturing variability, aging, wear, (d) To protect design features that corrosion, and likely damage. prevent catastrophic ignition sources (b) Except as provided in paragraphs within the fuel tank or fuel tank sys- (b)(2) and (c) of this section, no fuel tem according to paragraph (a) of this tank Fleet Average Flammability Ex- section, and to prevent increasing the posure on an airplane may exceed three flammability exposure of the tanks percent of the Flammability Exposure above that permitted in paragraph (b) Evaluation Time (FEET) as defined in of this section, the type design must Appendix N of this part, or that of a include critical design configuration fuel tank within the wing of the air- control limitations (CDCCLs) identi- plane model being evaluated, which- fying those features and providing in- ever is greater. If the wing is not a con- structions on how to protect them. To ventional unheated aluminum wing, ensure the continued effectiveness of

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those features, and prevent degrada- relative motion could exist must have tion of the performance and reliability provisions for flexibility. of any means provided according to (c) Each flexible connection in fuel paragraphs (a), (b), or (c) of this sec- lines that may be under pressure and tion, the type design must also include subjected to axial loading must use necessary inspection and test proce- flexible hose assemblies. dures, intervals between repetitive in- (d) Flexible hose must be approved or spections and tests, and mandatory re- must be shown to be suitable for the placement times for those features. particular application. The applicant must include informa- (e) No flexible hose that might be ad- tion required by this paragraph in the versely affected by exposure to high Airworthiness Limitations section of temperatures may be used where exces- the Instructions for Continued Air- sive temperatures will exist during op- worthiness required by § 25.1529. The eration or after engine shut-down. type design must also include visible (f) Each fuel line within the fuselage means of identifying critical features must be designed and installed to allow of the design in areas of the airplane a reasonable degree of deformation and where foreseeable maintenance ac- stretching without leakage. tions, repairs, or alterations may com- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as promise the CDCCLs. amended by Amdt. 25–15, 32 FR 13266, Sept. 20, 1967] [Doc. No. 1999–6411, 66 FR 23129, May 7, 2001, as amended by Doc. No. FAA–2005–22997, 73 § 25.994 Fuel system components. FR 42494, July 21, 2008; Doc. No. FAA– 2014– 1027, Amdt. No. 25–146, 83 FR 47556, Sept. 20, Fuel system components in an engine 2018] nacelle or in the fuselage must be protected from damage that could result FUEL SYSTEM COMPONENTS in spillage of enough fuel to constitute a fire hazard as a result of a wheels-up § 25.991 Fuel pumps. landing on a paved runway under each (a) Main pumps. Each fuel pump re- of the conditions prescribed in quired for proper engine operation, or § 25.721(b). required to meet the fuel system re- [Amdt. 25–139, 79 FR 59430, Oct. 2, 2014] 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 § 25.1189 for shutoff means, each fuel bypass of each positive displacement valve must— fuel pump other than a fuel injection (a) [Reserved] pump (a pump that supplies the proper (b) Be supported so that no loads re- flow and pressure for fuel injection sulting from their operation or from when the injection is not accomplished accelerated flight conditions are trans- in a carburetor) approved as part of the mitted to the lines attached to the engine. valve. (b) Emergency pumps. There must be emergency pumps or another main [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–40, 42 FR 15043, Mar. 17, pump to feed each engine immediately 1977] after failure of any main pump (other than a fuel injection pump approved as § 25.997 Fuel strainer or filter. part of the engine). There must be a fuel strainer or filter between the fuel tank outlet and the § 25.993 Fuel system lines and fittings. inlet of either the fuel metering device (a) Each fuel line must be installed or an engine driven positive displace- and supported to prevent excessive vi- ment pump, whichever is nearer the bration and to withstand loads due to fuel tank outlet. This fuel strainer or fuel pressure and accelerated flight filter must— conditions. (a) Be accessible for draining and (b) Each fuel line connected to com- cleaning and must incorporate a screen ponents of the airplane between which or element which is easily removable;

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(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. event of a landing with landing gear re- (e) For reciprocating engine powered tracted. airplanes, means must be provided to [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as prevent jettisoning the fuel in the amended by Amdt. 25–38, 41 FR 55467, Dec. 20, tanks used for takeoff and landing 1976] below the level allowing 45 minutes flight at 75 percent maximum contin- § 25.1001 Fuel jettisoning system. uous power. However, if there is an (a) A fuel jettisoning system must be auxiliary control independent of the installed on each airplane unless it is main jettisoning control, the system shown that the airplane meets the may be designed to jettison the re- climb requirements of §§ 25.119 and maining fuel by means of the auxiliary 25.121(d) at maximum takeoff weight, jettisoning control. less the actual or computed weight of (f) For turbine engine powered air- fuel necessary for a 15-minute flight planes, means must be provided to pre- comprised of a takeoff, go-around, and vent jettisoning the fuel in the tanks landing at the airport of departure used for takeoff and landing below the with the airplane configuration, speed, level allowing climb from sea level to power, and thrust the same as that 10,000 feet and thereafter allowing 45

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minutes cruise at a speed for maximum (c) Fuel/oil ratios higher than those range. However, if there is an auxiliary prescribed in paragraphs (b)(1) and (2) control independent of the main jetti- of this section may be used if substan- soning control, the system may be de- tiated by data on actual engine oil con- signed to jettison the remaining fuel sumption. by means of the auxiliary jettisoning control. § 25.1013 Oil tanks. (g) The fuel jettisoning valve must be (a) Installation. Each oil tank instal- designed to allow flight personnel to lation must meet the requirements of close the valve during any part of the § 25.967. jettisoning operation. (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. (c) Filler connection. Each recessed oil [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as 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. (b) The usable oil capacity may not (2) Oil tank vents must be arranged be less than the product of the endur- so that condensed water vapor that ance of the airplane under critical op- might freeze and obstruct the line can- erating conditions and the approved not accumulate at any point. maximum allowable oil consumption of (e) Outlet. There must be means to the engine under the same conditions, prevent entrance into the tank itself, plus a suitable margin to ensure sys- or into the tank outlet, of any object tem circulation. Instead of a rational that might obstruct the flow of oil analysis of airplane range for the pur- through the system. No oil tank outlet pose of computing oil requirements for may be enclosed by any screen or guard reciprocating engine powered air- that would reduce the flow of oil below planes, the following fuel/oil ratios a safe value at any operating tempera- may be used: ture. There must be a shutoff valve at (1) For airplanes without a reserve the outlet of each oil tank used with a oil or oil transfer system, a fuel/oil turbine engine, unless the external por- ratio of 30:1 by volume. tion of the oil system (including the oil (2) For airplanes with either a re- tank supports) is fireproof. serve oil or oil transfer system, a fuel/ (f) Flexible oil tank liners. Each flexi- oil ratio of 40:1 by volume. ble oil tank liner must be approved or

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must be shown to be suitable for the (1) Each oil strainer or filter that has particular application. a bypass must be constructed and in- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as stalled so that oil will flow at the nor- amended by Amdt. 25–19, 33 FR 15410, Oct. 17, mal rate through the rest of the sys- 1968; Amdt. 25–23, 35 FR 5677, Apr. 8, 1970; tem with the strainer or filter com- Amdt. 25–36, 39 FR 35460, Oct. 1, 1974; Amdt. pletely blocked. 25–57, 49 FR 6848, Feb. 23, 1984; Amdt. 25–72, 55 (2) The oil strainer or filter must FR 29785, July 20, 1990] have the capacity (with respect to op- § 25.1015 Oil tank tests. erating limitations established for the 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 (b) Breather lines must be arranged system with the strainer or filter ele- so that— ment completely blocked. (1) Condensed water vapor that might freeze and obstruct the line cannot ac- [Amdt. 25–36, 39 FR 35461, Oct. 1, 1974, as cumulate at any point; amended by Amdt. 25–57, 49 FR 6848, Feb. 23, (2) The breather discharge does not 1984] constitute a fire hazard if foaming oc- § 25.1021 Oil system drains. curs or causes emitted oil to strike the pilot’s windshield; and A drain (or drains) must be provided (3) The breather does not discharge to allow safe drainage of the oil sys- into the engine air induction system. tem. Each drain must— (a) Be accessible; and § 25.1019 Oil strainer or filter. (b) Have manual or automatic means (a) Each turbine engine installation for positive locking in the closed posi- must incorporate an oil strainer or fil- tion. ter through which all of the engine oil flows and which meets the following re- [Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] quirements:

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§ 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 max- (c) Each oil valve must have positive imum ambient atmospheric tempera- stops or suitable index provisions in ture, the recorded powerplant tempera- the ‘‘on’’ and ‘‘off’’ positions and must tures must be corrected under para- be supported so that no loads resulting graphs (c) and (d) of this section. from its operation or from accelerated (2) No corrected temperatures deter- flight conditions are transmitted to mined under paragraph (a)(1) of this the lines attached to the valve. section may exceed established limits. (3) For reciprocating engines, the fuel § 25.1027 Propeller feathering system. used during the cooling tests must be 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 source of oil for lubricating the engine constant at ¥69.7 degrees F. However, during operation. for winterization installations, the ap- (d) Provision must be made to pre- plicant may select a maximum ambi- vent sludge or other foreign matter ent atmospheric temperature cor- from affecting the safe operation of the responding to sea level conditions of propeller feathering system. less than 100 degrees F. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) Correction factor (except cylinder amended by Amdt. 25–38, 41 FR 55467, Dec. 20, barrels). Unless a more rational correc- 1976] tion applies, temperatures of engine fluids and powerplant components (ex- COOLING cept cylinder barrels) for which temperature limits are established, must § 25.1041 General. be corrected by adding to them the dif- The powerplant and auxiliary power ference between the maximum ambient unit cooling provisions must be able to atmospheric temperature and the tem- maintain the temperatures of power- perature of the ambient air at the time plant components, engine fluids, and of the first occurrence of the maximum

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component or fluid temperature re- airplane reaches an altitude of 1,500 corded during the cooling test. feet above the takeoff surface or (d) Correction factor for cylinder barrel reaches a point in the takeoff where temperatures. Unless a more rational the transition from the takeoff to the correction applies, cylinder barrel tem- en route configuration is completed peratures must be corrected by adding and a speed is reached at which compli- to them 0.7 times the difference be- ance with § 25.121(c) is shown, which- tween the maximum ambient atmos- ever point is at a higher altitude. The pheric temperature and the tempera- airplane must be in the following con- ture of the ambient air at the time of figuration: the first occurrence of the maximum (1) Landing gear retracted. cylinder barrel temperature recorded (2) Wing flaps in the most favorable during the cooling test. position. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (3) Cowl flaps (or other means of con- amended by Amdt. 25–42, 43 FR 2323, Jan. 16, trolling the engine cooling supply) in 1978] the position that provides adequate cooling in the hot-day condition. § 25.1045 Cooling test procedures. (4) Critical engine inoperative and its (a) Compliance with § 25.1041 must be propeller stopped. shown for the takeoff, climb, en route, (5) Remaining engines at the max- and landing stages of flight that cor- imum continuous power available for respond to the applicable performance the altitude. requirements. The cooling tests must (e) For hull seaplanes and amphib- be conducted with the airplane in the ians, cooling must be shown during configuration, and operating under the taxiing downwind for 10 minutes, at conditions, that are critical relative to five knots above step speed. cooling during each stage of flight. For the cooling tests, a temperature is [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ‘‘stabilized’’ when its rate of change is amended by Amdt. 25–57, 49 FR 6848, Feb. 23, 1984] less than two degrees F. per minute. (b) Temperatures must be stabilized INDUCTION SYSTEM under the conditions from which entry is made into each stage of flight being § 25.1091 Air induction. investigated, unless the entry condi- (a) The air induction system for each tion normally is not one during which engine and auxiliary power unit must component and the engine fluid tem- supply— peratures would stabilize (in which case, operation through the full entry (1) The air required by that engine condition must be conducted before and auxiliary power unit under each entry into the stage of flight being in- operating condition for which certifi- vestigated in order to allow tempera- cation is requested; and tures to reach their natural levels at (2) The air for proper fuel metering the time of entry). The takeoff cooling and mixture distribution with the in- test must be preceded by a period dur- duction system valves in any position. ing which the powerplant component (b) Each reciprocating engine must and engine fluid temperatures are sta- have an alternate air source that pre- bilized with the engines at ground idle. vents the entry of rain, ice, or any (c) Cooling tests for each stage of other foreign matter. flight must be continued until— (c) Air intakes may not open within (1) The component and engine fluid the cowling, unless— temperatures stabilize; (1) That part of the cowling is iso- (2) The stage of flight is completed; lated from the engine accessory section or by means of a fireproof diaphragm; or (3) An operating limitation is (2) For reciprocating engines, there reached. are means to prevent the emergence of (d) For reciprocating engine powered backfire flames. airplanes, it may be assumed, for cool- (d) For turbine engine powered air- ing test purposes, that the takeoff planes and airplanes incorporating aux- stage of flight is complete when the iliary power units—

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(1) There must be means to prevent descent idling speeds, in the icing con- hazardous quantities of fuel leakage or ditions defined in Appendices C and O overflow from drains, vents, or other of this part, and Appendix D of part 33 components of flammable fluid systems of this chapter, and in falling and blow- from entering the engine or auxiliary ing snow within the limitations estab- power unit intake system; and lished for the airplane for such oper- (2) The airplane must be designed to ation, without the accumulation of ice prevent water or slush on the runway, on the engine, inlet system compo- taxiway, or other airport operating nents, or airframe components that surfaces from being directed into the would do any of the following: engine or auxiliary power unit air inlet (i) Adversely affect installed engine ducts in hazardous quantities, and the operation or cause a sustained loss of air inlet ducts must be located or pro- power or thrust; or an unacceptable in- tected so as to minimize the ingestion crease in gas path operating tempera- of foreign matter during takeoff, land- ture; or an airframe/engine incompati- ing, and taxiing. bility; or (e) If the engine induction system (ii) Result in unacceptable temporary contains parts or components that power loss or engine damage; or could be damaged by foreign objects (iii) Cause a stall, surge, or flameout entering the air inlet, it must be shown by tests or, if appropriate, by analysis or loss of engine controllability (for ex- that the induction system design can ample, rollback). withstand the foreign object ingestion (2) Operate at ground idle speed for a test conditions of §§ 33.76, 33.77 and minimum of 30 minutes on the ground 33.78(a)(1) of this chapter without fail- in the following icing conditions shown ure of parts or components that could in Table 1 of this section, unless re- create a hazard. placed by similar test conditions that are more critical. These conditions [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as must be demonstrated with the avail- amended by Amdt. 25–38, 41 FR 55467, Dec. 20, able air bleed for icing protection at its 1976; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. critical condition, without adverse ef- 25–100, 65 FR 55854, Sept. 14, 2000] fect, followed by an acceleration to takeoff power or thrust in accordance § 25.1093 Induction system icing pro- with the procedures defined in the air- tection. plane flight manual. During the idle (a) Reciprocating engines. Each recip- operation, the engine may be run up rocating engine air induction system periodically to a moderate power or must have means to prevent and elimi- thrust setting in a manner acceptable nate icing. Unless this is done by other to the Administrator. Analysis may be means, it must be shown that, in air used to show ambient temperatures free of visible moisture at a tempera- below the tested temperature are less ture of 30 F., each airplane with alti- critical. The applicant must document tude engines using— the engine run-up procedure (including (1) Conventional venturi carburetors the maximum time interval between have a preheater that can provide a run-ups from idle, run-up power set- heat rise of 120 F. with the engine at 60 ting, and duration at power), the asso- percent of maximum continuous power; ciated minimum ambient temperature, or and the maximum time interval. These (2) Carburetors tending to reduce the conditions must be used in the analysis probability of ice formation has a pre- that establishes the airplane operating heater that can provide a heat rise of limitations in accordance with § 25.1521. 100 °F. with the engine at 60 percent of (3) For the purposes of this section, maximum continuous power. the icing conditions defined in appen- (b) Turbine engines. Except as pro- dix O of this part, including the condi- vided in paragraph (b)(3) of this sections specified in Condition 3 of Table 1 tion, each engine, with all icing protec- of this section, are not applicable to tion systems operating, must: airplanes with a maximum takeoff (1) Operate throughout its flight weight equal to or greater than 60,000 power range, including the minimum pounds.

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TABLE 1—ICING CONDITIONS FOR GROUND TESTS

Water concentration Mean effective par- Condition Total air temperature (minimum) ticle diameter Demonstration

1. Rime ice condition 0 to 15 °F (18 to ¥9 °C) Liquid—0.3 g/m3 ...... 15–25 microns ...... By test, analysis or combination of the two. 2. Glaze ice condition 20 to 30 °F (¥7 to ¥1 Liquid—0.3 g/m3 ...... 15–25 microns ...... By test, analysis or com- °C). bination of the two. 3. Large drop condi- 15 to 30 °F (¥9 to ¥1 Liquid—0.3 g/m3 ...... 100 microns (min- By test, analysis or com- tion. °C). imum). bination of the two.

(c) Supercharged reciprocating engines. (2) Fire-resistant if it is in any fire For each engine having a supercharger zone for which a fire-extinguishing sys- to pressurize the air before it enters tem is required, except that ducts for the carburetor, the heat rise in the air auxiliary power units must be fireproof caused by that supercharging at any within the auxiliary power unit fire altitude may be utilized in determining zone. compliance with paragraph (a) of this (c) Each duct connected to compo- section if the heat rise utilized is that nents between which relative motion which will be available, automatically, could exist must have means for flexi- for the applicable altitude and oper- bility. ating condition because of super- (d) For turbine engine and auxiliary charging. power unit bleed air duct systems, no hazard may result if a duct failure oc- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55467, Dec. 20, curs at any point between the air duct 1976; Amdt. 25–40, 42 FR 15043, Mar. 17, 1977; source and the airplane unit served by Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. the air. 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–140, (e) Each auxiliary power unit induc- 79 FR 65526, Nov. 4, 2014] tion system duct must be fireproof for a sufficient distance upstream of the § 25.1101 Carburetor air preheater de- auxiliary power unit compartment to sign. prevent hot gas reverse flow from burn- Each carburetor air preheater must ing through auxiliary power unit ducts be designed and constructed to— and entering any other compartment (a) Ensure ventilation of the pre- or area of the airplane in which a haz- heater when the engine is operated in ard would be created resulting from the cold air; entry of hot gases. The materials used (b) Allow inspection of the exhaust to form the remainder of the induction manifold parts that it surrounds; and system duct and plenum chamber of (c) Allow inspection of critical parts the auxiliary power unit must be capa- of the preheater itself. ble of resisting the maximum heat conditions likely to occur. § 25.1103 Induction system ducts and (f) Each auxiliary power unit induc- air duct systems. tion system duct must be constructed of materials that will not absorb or (a) Each induction system duct up- trap hazardous quantities of flammable stream of the first stage of the engine fluids that could be ignited in the supercharger and of the auxiliary event of a surge or reverse flow condi- power unit compressor must have a tion. drain to prevent the hazardous accumulation of fuel and moisture in the [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ground attitude. No drain may dis- amended by Amdt. 25–46, 43 FR 50597, Oct. 30, charge where it might cause a fire haz- 1978] ard. (b) Each induction system duct must § 25.1105 Induction system screens. be— If induction system screens are (1) Strong enough to prevent induc- used— tion system failures resulting from (a) Each screen must be upstream of normal backfire conditions; and the carburetor;

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(b) No screen may be in any part of (g) Each exhaust shroud must be ven- the induction system that is the only tilated or insulated to avoid, during passage through which air can reach normal operation, a temperature high the engine, unless it can be deiced by enough to ignite any flammable fluids heated air; or vapors external to the shroud. (c) No screen may be deiced by alco- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as hol alone; and amended by Amdt. 25–40, 42 FR 15043, Mar. 17, (d) It must be impossible for fuel to 1977] strike any screen. § 25.1123 Exhaust piping. § 25.1107 Inter-coolers and after-coolers. For powerplant and auxiliary power unit installations, the following apply: Each inter-cooler and after-cooler (a) Exhaust piping must be heat and must be able to withstand any vibra- corrosion resistant, and must have pro- tion, inertia, and air pressure load to visions to prevent failure due to expan- which it would be subjected in oper- sion by operating temperatures. ation. (b) Piping must be supported to withstand any vibration and inertia loads EXHAUST SYSTEM to which it would be subjected in oper- § 25.1121 General. ation; and (c) Piping connected to components For powerplant and auxiliary power between which relative motion could unit installations the following apply: exist must have means for flexibility. (a) Each exhaust system must ensure safe disposal of exhaust gases without [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as fire hazard or carbon monoxide con- amended by Amdt. 25–40, 42 FR 15044, Mar. 17, tamination in any personnel compart- 1977] ment. For test purposes, any acceptable carbon monoxide detection meth- § 25.1125 Exhaust heat exchangers. od may be used to show the absence of For reciprocating engine powered carbon monoxide. airplanes, the following apply: (b) Each exhaust system part with a (a) Each exhaust heat exchanger surface hot enough to ignite flammable must be constructed and installed to fluids or vapors must be located or withstand each vibration, inertia, and 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 re- fluids or vapors that might be present spect to any flammable fluid vent or in case of the failure or malfunction of drain. components carrying flammable fluids. (e) No exhaust gases may discharge (b) If an exhaust heat exchanger is where they will cause a glare seriously used for heating ventilating air— affecting pilot vision at night. (1) There must be a secondary heat (f) Each exhaust system component exchanger between the primary ex- must be ventilated to prevent points of haust gas heat exchanger and the ven- excessively high temperature. tilating air system; or

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(2) Other means must be used to pre- (e) The portion of each powerplant clude the harmful contamination of the control located in a designated fire ventilating air. zone that is required to be operated in the event of fire must be at least fire [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as resistant. amended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976] (f) For powerplant valve controls located in the flight deck there must be § 25.1127 Exhaust driven turbo-super- a means: chargers. (1) For the flightcrew to select each (a) Each exhaust driven turbo-super- intended position or function of the charger must be approved or shown to valve; and be suitable for the particular applica- (2) To indicate to the flightcrew: tion. It must be installed and sup- (i) The selected position or function ported to ensure safe operation be- of the valve; and tween normal inspections and over- (ii) When the valve has not responded hauls. In addition, there must be provi- as intended to the selected position or sions for expansion and flexibility be- function. tween exhaust conduits and the tur- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as bine. amended by Amdt. 25–40, 42 FR 15044, Mar. 17, (b) There must be provisions for lu- 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990; bricating the turbine and for cooling Amdt. 25–115, 69 FR 40527, July 2, 2004] turbine parts where temperatures are critical. § 25.1142 Auxiliary power unit con- (c) If the normal turbo-supercharger trols. control system malfunctions, the tur- Means must be provided on the flight bine speed may not exceed its max- deck for starting, stopping, and emer- imum allowable value. Except for the gency shutdown of each installed auxil- waste gate operating components, the iary power unit. components provided for meeting this requirement must be independent of [Amdt. 25–46, 43 FR 50598, Oct. 30, 1978] the normal turbo-supercharger con- § 25.1143 Engine controls. trols. (a) There must be a separate power or POWERPLANT CONTROLS AND thrust control for each engine. ACCESSORIES (b) Power and thrust controls must be arranged to allow— § 25.1141 Powerplant controls: general. (1) Separate control of each engine; Each powerplant control must be lo- and cated, arranged, and designed under (2) Simultaneous control of all en- §§ 25.777 through 25.781 and marked gines. under § 25.1555. In addition, it must (c) Each power and thrust control meet the following requirements: must provide a positive and imme- (a) Each control must be located so diately responsive means of controlling that it cannot be inadvertently oper- its engine. ated by persons entering, leaving, or (d) For each fluid injection (other moving normally in, the cockpit. than fuel) system and its controls not (b) Each flexible control must be ap- provided and approved as part of the proved or must be shown to be suitable engine, the applicant must show that for the particular application. the flow of the injection fluid is ade- (c) Each control must have sufficient quately controlled. strength and rigidity to withstand op- (e) If a power or thrust control incor- erating loads without failure and with- porates a fuel shutoff feature, the con- out excessive deflection. trol must have a means to prevent the (d) Each control must be able to inadvertent movement of the control maintain any set position without con- into the shutoff position. The means stant attention by flight crewmembers must— and without creep due to control loads (1) Have a positive lock or stop at the or vibration. idle position; and

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(2) Require a separate and distinct (d) The propeller speed and pitch con- operation to place the control in the trols must be to the right of, and at shutoff position. least one inch below, the pilot’s throt- [Amdt. 25–23, 35 FR 5677, Apr. 8, 1970, as tle controls. amended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984] § 25.1153 Propeller feathering controls. (a) There must be a separate pro- § 25.1145 Ignition switches. peller feathering control for each pro- (a) Ignition switches must control peller. The control must have means to each engine ignition circuit on each prevent its inadvertent operation. engine. (b) If feathering is accomplished by (b) There must be means to quickly movement of the propeller pitch or shut off all ignition by the grouping of speed control lever, there must be switches or by a master ignition con- means to prevent the inadvertent trol. movement of this lever to the feath- (c) Each group of ignition switches, ering position during normal oper- except ignition switches for turbine en- ation. gines for which continuous ignition is not required, and each master ignition [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as control must have a means to prevent amended by Amdt. 25–11, 32 FR 6913, May 5, its inadvertent operation. 1967] [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.1155 Reverse thrust and propeller amended by Amdt. 25–40, 42 FR 15044 Mar. 17, pitch settings below the flight re- 1977] gime. § 25.1147 Mixture controls. Each control for reverse thrust and for propeller pitch settings below the (a) If there are mixture controls, flight regime must have means to pre- each engine must have a separate con- vent its inadvertent operation. The trol. The controls must be grouped and arranged to allow— means must have a positive lock or (1) Separate control of each engine; stop at the flight idle position and and must require a separate and distinct (2) Simultaneous control of all en- operation by the crew to displace the gines. control from the flight regime (forward (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 (a) There must be a separate pro- accessible to the pilots or, if there is a peller speed and pitch control for each separate flight engineer station with a propeller. control panel, to the flight engineer. (b) The controls must be grouped and arranged to allow— § 25.1161 Fuel jettisoning system con- (1) Separate control of each pro- trols. peller; and Each fuel jettisoning system control (2) Simultaneous control of all pro- must have guards to prevent inad- pellers. vertent operation. No control may be (c) The controls must allow synchro- near any fire extinguisher control or nization of all propellers. other control used to combat fire.

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§ 25.1163 Powerplant accessories. (e) No ground wire for any engine (a) Each engine mounted accessory may be routed through a fire zone of must— another engine unless each part of that (1) Be approved for mounting on the wire within that zone is fireproof. engine involved; (f) Each ignition system must be (2) Use the provisions on the engine independent of any electrical circuit, for mounting; and not used for assisting, controlling, or (3) Be sealed to prevent contamina- analyzing the operation of that system. tion of the engine oil system and the (g) There must be means to warn ap- accessory system. propriate flight crewmembers if the (b) Electrical equipment subject to malfunctioning of any part of the elec- arcing or sparking must be installed to trical system is causing the continuous minimize the probability of contact discharge of any battery necessary for with any flammable fluids or vapors engine ignition. that might be present in a free state. (h) Each engine ignition system of a (c) If continued rotation of an engine- turbine powered airplane must be con- driven cabin supercharger or of any re- sidered an essential electrical load. mote accessory driven by the engine is [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as hazardous if malfunctioning occurs, amended by Amdt. 25–23, 35 FR 5677, Apr. 8, there must be means to prevent rota- 1970; Amdt. 25–72, 55 FR 29785, July 20, 1990] tion without interfering with the continued operation of the engine. § 25.1167 Accessory gearboxes. For airplanes equipped with an acces- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–57, 49 FR 6849, Feb. 23, sory gearbox that is not certificated as 1984] part of an engine— (a) The engine with gearbox and con- § 25.1165 Engine ignition systems. necting transmissions and shafts at- (a) Each battery ignition system tached must be subjected to the tests must be supplemented by a generator specified in § 33.49 or § 33.87 of this chap- that is automatically available as an ter, as applicable; alternate source of electrical energy to (b) The accessory gearbox must meet allow continued engine operation if the requirements of §§ 33.25 and 33.53 or any battery becomes depleted. 33.91 of this chapter, as applicable; and (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 [Amdt. 25–38, 41 FR 55467, Dec. 20, 1976] from the same source. (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 de- (a) Designated fire zones are— pleted battery with the generator run- (1) The engine power section; ning at its normal operating speed; and (2) The engine accessory section; (3) The condition of a completely de- (3) Except for reciprocating engines, pleted battery with the generator oper- any complete powerplant compartment ating at idling speed, if there is only in which no isolation is provided be- one battery. tween the engine power section and the (d) Magneto ground wiring (for sepa- engine accessory section; rate ignition circuits) that lies on the (4) Any auxiliary power unit com- engine side of the fire wall, must be in- partment; stalled, located, or protected, to mini- (5) Any fuel-burning heater and other mize the probability of simultaneous combustion equipment installation de- failure of two or more wires as a result scribed in § 25.859; of mechanical damage, electrical (6) The compressor and accessory sec- faults, or other cause. tions of turbine engines; and

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(7) Combustor, turbine, and tailpipe (b) Paragraph (a) of this section does sections of turbine engine installations not apply to— that contain lines or components car- (1) Lines, fittings, and components rying flammable fluids or gases. which are already approved as part of a (b) Each designated fire zone must type certificated engine; and meet the requirements of §§ 25.863, (2) Vent and drain lines, and their fit- 25.865, 25.867, 25.869, and 25.1185 through tings, whose failure will not result in, 25.1203. or add to, a fire hazard. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (c) All components, including ducts, amended by Amdt. 25–11, 32 FR 6913, May 5, within a designated fire zone must be 1967; Amdt. 25–23, 35 FR 5677, Apr. 8, 1970; fireproof if, when exposed to or dam- Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. aged by fire, they could— 25–115, 69 FR 40527, July 2, 2004] (1) Result in fire spreading to other § 25.1182 Nacelle areas behind fire- regions of the airplane; or walls, and engine pod attaching (2) Cause unintentional operation of, structures containing flammable or inability to operate, essential serv- fluid lines. ices or equipment. (a) Each nacelle area immediately [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as behind the firewall, and each portion of amended by Amdt. 25–11, 32 FR 6913, May 5, any engine pod attaching structure 1967; Amdt. 25–36, 39 FR 35461, Oct. 1, 1974; containing flammable fluid lines, must Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. meet each requirement of §§ 25.1103(b), 25–101, 65 FR 79710, Dec. 19, 2000] 25.1165 (d) and (e), 25.1183, 25.1185(c), 25.1187, 25.1189, and 25.1195 through § 25.1185 Flammable fluids. 25.1203, including those concerning des- (a) Except for the integral oil sumps ignated fire zones. However, engine pod specified in § 25.1183(a), no tank or res- attaching structures need not contain ervoir that is a part of a system con- fire detection or extinguishing means. taining flammable fluids or gases may (b) For each area covered by para- be in a designated fire zone unless the graph (a) of this section that contains fluid contained, the design of the sys- a retractable landing gear, compliance tem, the materials used in the tank, with that paragraph need only be the shut-off means, and all connec- shown with the landing gear retracted. tions, lines, and control provide a de- [Amdt. 25–11, 32 FR 6913, May 5, 1967] gree of safety equal to that which would exist if the tank or reservoir § 25.1183 Flammable fluid-carrying were outside such a zone. components. (b) There must be at least one-half (a) Except as provided in paragraph inch of clear airspace between each (b) of this section, each line, fitting, tank or reservoir and each firewall or and other component carrying flam- shroud isolating a designated fire zone. mable fluid in any area subject to en- (c) Absorbent materials close to gine fire conditions, and each compo- flammable fluid system components nent which conveys or contains flam- that might leak must be covered or mable fluid in a designated fire zone treated to prevent the absorption of must be fire resistant, except that hazardous quantities of fluids. flammable fluid tanks and supports in a designated fire zone must be fireproof [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as or be enclosed by a fireproof shield un- amended by Amdt. 25–19, 33 FR 15410, Oct. 17, less damage by fire to any non-fire- 1968; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998] proof part will not cause leakage or spillage of flammable fluid. Compo- § 25.1187 Drainage and ventilation of nents must be shielded or located to fire zones. safeguard against the ignition of leak- (a) There must be complete drainage ing flammable fluid. An integral oil of each part of each designated fire sump of less than 25-quart capacity on zone to minimize the hazards resulting a reciprocating engine need not be fire- from failure or malfunctioning of any proof nor be enclosed by a fireproof component containing flammable shield. fluids. The drainage means must be—

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(1) Effective under conditions ex- (e) No hazardous quantity of flam- pected to prevail when drainage is mable fluid may drain into any des- needed; and ignated fire zone after shutoff. (2) Arranged so that no discharged (f) There must be means to guard fluid will cause an additional fire haz- against inadvertent operation of the ard. shutoff means and to make it possible (b) Each designated fire zone must be for the crew to reopen the shutoff ventilated to prevent the accumulation means in flight after it has been closed. of flammable vapors. (g) Each tank-to-engine shutoff valve (c) No ventilation opening may be must be located so that the operation where it would allow the entry of flam- of the valve will not be affected by mable fluids, vapors, or flame from powerplant or engine mount structural other zones. failure. (d) Each ventilation means must be (h) Each shutoff valve must have a arranged so that no discharged vapors means to relieve excessive pressure ac- will cause an additional fire hazard. cumulation unless a means for pressure (e) Unless the extinguishing agent ca- relief is otherwise provided in the sys- pacity and rate of discharge are based tem. on maximum air flow through a zone, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as there must be means to allow the crew amended by Amdt. 25–23, 35 FR 5677, Apr. 8, to shut off sources of forced ventilation 1970; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984] to any fire zone except the engine power section of the nacelle and the § 25.1191 Firewalls. combustion heater ventilating air (a) Each engine, auxiliary power ducts. unit, fuel-burning heater, other combustion equipment intended for oper- § 25.1189 Shutoff means. ation in flight, and the combustion, (a) Each engine installation and each turbine, and tailpipe sections of tur- fire zone specified in § 25.1181(a)(4) and bine engines, must be isolated from the (5) must have a means to shut off or rest of the airplane by firewalls, otherwise prevent hazardous quantities shrouds, or equivalent means. of fuel, oil, deicer, and other flammable (b) Each firewall and shroud must fluids, from flowing into, within, or be— through any designated fire zone, ex- (1) Fireproof; cept that shutoff means are not re- (2) Constructed so that no hazardous quired for— quantity of air, fluid, or flame can pass (1) Lines, fittings, and components from the compartment to other parts forming an integral part of an engine; of the airplane; and (3) Constructed so that each opening (2) Oil systems for turbine engine in- is sealed with close fitting fireproof stallations in which all components of grommets, bushings, or firewall fit- the system in a designated fire zone, tings; and including oil tanks, are fireproof or lo- (4) Protected against corrosion. cated in areas not subject to engine fire conditions. § 25.1192 Engine accessory section dia- (b) The closing of any fuel shutoff phragm. valve for any engine may not make For reciprocating engines, the engine fuel unavailable to the remaining en- power section and all portions of the gines. exhaust system must be isolated from (c) Operation of any shutoff may not the engine accessory compartment by a interfere with the later emergency op- diaphragm that complies with the fire- eration of other equipment, such as the wall requirements of § 25.1191. means for feathering the propeller. (d) Each flammable fluid shutoff [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970] means and control must be fireproof or must be located and protected so that § 25.1193 Cowling and nacelle skin. any fire in a fire zone will not affect its (a) Each cowling must be constructed operation. and supported so that it can resist any

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vibration, inertia, and air load to heaters, and other combustion equip- which it may be subjected in operation. ment. For each other designated fire (b) Cowling must meet the drainage zone, two discharges must be provided and ventilation requirements of each of which produces adequate agent § 25.1187. concentration. (c) On airplanes with a diaphragm (c) The fire extinguishing system for isolating the engine power section from a nacelle must be able to simulta- the engine accessory section, each part neously protect each zone of the na- of the accessory section cowling sub- celle for which protection is provided. ject to flame in case of fire in the en- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as gine power section of the powerplant amended by Amdt. 25–46, 43 FR 50598, Oct. 30, must— 1978] (1) Be fireproof; and (2) Meet the requirements of § 25.1191. § 25.1197 Fire extinguishing agents. (d) Each part of the cowling subject (a) Fire extinguishing agents must— to high temperatures due to its near- (1) Be capable of extinguishing ness to exhaust system parts or ex- flames emanating from any burning of haust gas impingement must be fire- fluids or other combustible materials proof. in the area protected by the fire extin- (e) Each airplane must— guishing system; and (1) Be designed and constructed so (2) Have thermal stability over the that no fire originating in any fire zone temperature range likely to be experi- can enter, either through openings or enced in the compartment in which by burning through external skin, any they are stored. other zone or region where it would (b) If any toxic extinguishing agent is create additional hazards; used, provisions must be made to pre- (2) Meet paragraph (e)(1) of this sec- vent harmful concentrations of fluid or tion with the landing gear retracted (if fluid vapors (from leakage during nor- applicable); and mal operation of the airplane or as a (3) Have fireproof skin in areas sub- result of discharging the fire extin- ject to flame if a fire starts in the en- guisher on the ground or in flight) from gine power or accessory sections. entering any personnel compartment, even though a defect may exist in the § 25.1195 Fire extinguishing systems. extinguishing system. This must be (a) Except for combustor, turbine, shown by test except for built-in car- and tail pipe sections of turbine engine bon dioxide fuselage compartment fire installations that contain lines or com- extinguishing systems for which— ponents carrying flammable fluids or (1) Five pounds or less of carbon diox- gases for which it is shown that a fire ide will be discharged, under estab- originating in these sections can be lished fire control procedures, into any controlled, there must be a fire extin- fuselage compartment; or guisher system serving each designated (2) There is protective breathing fire zone. equipment for each flight crewmember (b) The fire extinguishing system, the on flight deck duty. quantity of the extinguishing agent, [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the rate of discharge, and the discharge amended by Amdt. 25–38, 41 FR 55467, Dec. 20, distribution must be adequate to extin- 1976; Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] guish fires. It must be shown by either actual or simulated flights tests that § 25.1199 Extinguishing agent con- under critical airflow conditions in tainers. flight the discharge of the extin- (a) Each extinguishing agent con- guishing agent in each designated fire tainer must have a pressure relief to zone specified in paragraph (a) of this prevent bursting of the container by section will provide an agent con- excessive internal pressures. centration capable of extinguishing (b) The discharge end of each dis- fires in that zone and of minimizing charge line from a pressure relief con- the probability of reignition. An indi- nection must be located so that dis- vidual ‘‘one-shot’’ system may be used charge of the fire extinguishing agent for auxiliary power units, fuel burning would not damage the airplane. The

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line must also be located or protected sensor or associated wiring within a to prevent clogging caused by ice or designated fire zone, unless the system other foreign matter. continues to function as a satisfactory (c) There must be a means for each detection system after the short cir- fire extinguishing agent container to cuit. indicate that the container has dis- (c) No fire or overheat detector may charged or that the charging pressure be affected by any oil, water, other is below the established minimum nec- fluids or fumes that might be present. essary for proper functioning. (d) There must be means to allow the (d) The temperature of each con- crew to check, in flight, the func- tainer must be maintained, under in- tioning of each fire or overheat detec- tended operating conditions, to prevent tor electric circuit. the pressure in the container from— (1) Falling below that necessary to (e) Components of each fire or over- provide an adequate rate of discharge; heat detector system in a fire zone or must be fire-resistant. (2) Rising high enough to cause pre- (f) No fire or overheat detector sys- mature discharge. tem component for any fire zone may (e) If a pyrotechnic capsule is used to pass through another fire zone, un- discharge the extinguishing agent, less— each container must be installed so (1) It is protected against the possi- that temperature conditions will not bility of false warnings resulting from cause hazardous deterioration of the fires in zones through which it passes; pyrotechnic capsule. or [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (2) Each zone involved is simulta- amended by Amdt. 25–23, 35 FR 5678, Apr. 8, neously protected by the same detector 1970; Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] and extinguishing system. (g) Each fire detector system must be § 25.1201 Fire extinguishing system constructed so that when it is in the materials. configuration for installation it will (a) No material in any fire extin- not exceed the alarm activation time guishing system may react chemically approved for the detectors using the re- with any extinguishing agent so as to sponse time criteria specified in the ap- create a hazard. propriate Technical Standard Order for (b) Each system component in an en- the detector. gine compartment must be fireproof. (h) EWIS for each fire or overheat detector system in a fire zone must meet § 25.1203 Fire detector system. the requirements of § 25.1731. (a) There must be approved, quick acting fire or overheat detectors in [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as each designated fire zone, and in the amended by Amdt. 25–23, 35 FR 5678, Apr. 8, combustion, turbine, and tailpipe sec- 1970; Amdt. 25–26, 36 FR 5493, Mar. 24, 1971; tions of turbine engine installations, in Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] numbers and locations ensuring § 25.1207 Compliance. prompt detection of fire in those zones. (b) Each fire detector system must be Unless otherwise specified, compli- constructed and installed so that— ance with the requirements of §§ 25.1181 (1) It will withstand the vibration, in- through 25.1203 must be shown by a full ertia, and other loads to which it may scale fire test or by one or more of the be subjected in operation; following methods: (2) There is a means to warn the crew (a) Tests of similar powerplant con- in the event that the sensor or associ- figurations; ated wiring within a designated fire (b) Tests of components; zone is severed at one point, unless the (c) Service experience of aircraft system continues to function as a sat- with similar powerplant configura- isfactory detection system after the tions; severing; and (d) Analysis. (3) There is a means to warn the crew in the event of a short circuit in the [Amdt. 25–46, 43 FR 50598, Oct. 30, 1978]

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Subpart F—Equipment ation, of the effects on the airplane or systems resulting from flightcrew ac- GENERAL tions. (c) Operationally-relevant behavior § 25.1301 Function and installation. of the installed equipment must be: (a) Each item of installed equipment (1) Predictable and unambiguous; and must— (2) Designed to enable the flightcrew (1) Be of a kind and design appro- to intervene in a manner appropriate priate to its intended function; to the task. (2) Be labeled as to its identification, (d) To the extent practicable, in- function, or operating limitations, or stalled equipment must incorporate any applicable combination of these means to enable the flightcrew to man- factors; age errors resulting from the kinds of (3) Be installed according to limita- flightcrew interactions with the equip- tions specified for that equipment; and ment that can be reasonably expected (4) Function properly when installed. in service. This paragraph does not (b) EWIS must meet the require- apply to any of the following: ments of subpart H of this part. (1) Skill-related errors associated [Doc. No. 5066, 29 FR 18333, Dec. 24, 1964, as with manual control of the airplane; amended by Amdt. 25–123, 72 FR 63405, Nov. 8, (2) Errors that result from decisions, 2007] actions, or omissions committed with malicious intent; § 25.1302 Installed systems and equip- (3) Errors arising from a crew- ment for use by the flightcrew. member’s reckless decisions, actions, This section applies to installed sys- or omissions reflecting a substantial tems and equipment intended for disregard for safety; and flightcrew members’ use in operating (4) Errors resulting from acts or the airplane from their normally seat- threats of violence, including actions ed positions on the flight deck. The ap- taken under duress. plicant must show that these systems and installed equipment, individually [Doc. No. FAA–2010–1175, 78 FR 25846, May 3, and in combination with other such 2013] systems and equipment, are designed § 25.1303 Flight and navigation instru- so that qualified flightcrew members ments. trained in their use can safely perform all of the tasks associated with the sys- (a) The following flight and naviga- tems’ and equipment’s intended function instruments must be installed so tions. Such installed equipment and that the instrument is visible from systems must meet the following re- each pilot station: quirements: (1) A free air temperature indicator (a) Flight deck controls must be in- or an air-temperature indicator which stalled to allow accomplishment of all provides indications that are convert- the tasks required to safely perform ible to free-air temperature. the equipment’s intended function, and (2) A clock displaying hours, min- information must be provided to the utes, and seconds with a sweep-second flightcrew that is necessary to accom- pointer or digital presentation. plish the defined tasks. (3) A direction indicator (non- (b) Flight deck controls and informa- stabilized magnetic compass). tion intended for the flightcrew’s use (b) The following flight and naviga- must: tion instruments must be installed at (1) Be provided in a clear and unam- each pilot station: biguous manner at a resolution and (1) An airspeed indicator. If airspeed precision appropriate to the task; limitations vary with altitude, the in- (2) Be accessible and usable by the dicator must have a maximum allow- flightcrew in a manner consistent with able airspeed indicator showing the the urgency, frequency, and duration of variation of VMO with altitude. their tasks; and (2) An altimeter (sensitive). (3) Enable flightcrew awareness, if (3) A rate-of-climb indicator (vertical awareness is required for safe oper- speed).

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(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 (c) The following flight and naviga- instruments required by paragraph (a) tion instruments are required as pre- of this section, the following power- scribed in this paragraph: plant instruments are required: (1) A speed warning device is required (1) A carburetor air temperature indi- for turbine engine powered airplanes cator for each engine. (2) A cylinder head temperature indi- and for airplanes with VMO/MMO greater cator for each air-cooled engine. than 0.8 VDF/MDF or 0.8 V D/MD. The speed warning device must give effec- (3) A manifold pressure indicator for tive aural warning (differing distinc- each engine. tively from aural warnings used for (4) A fuel pressure indicator (to indi- other purposes) to the pilots, whenever cate the pressure at which the fuel is supplied) for each engine. the speed exceeds VMO plus 6 knots or (5) A fuel flowmeter, or fuel mixture MMO + 0.01. The upper limit of the production tolerance for the warning de- indicator, for each engine without an vice may not exceed the prescribed automatic altitude mixture control. warning speed. (6) A tachometer for each engine. (2) A machmeter is required at each (7) A device that indicates, to the pilot station for airplanes with com- flight crew (during flight), any change pressibility limitations not otherwise in the power output, for each engine indicated to the pilot by the airspeed with— indicating system required under para- (i) An automatic propeller feathering graph (b)(1) of this section. system, whose operation is initiated by a power output measuring system; or [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as (ii) A total engine piston displace- amended by Amdt. 25–24, 35 FR 7108, May 6, ment of 2,000 cubic inches or more. 1970; Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; (8) A means to indicate to the pilot Amdt. 25–90, 62 FR 13253, Mar. 19, 1997] when the propeller is in reverse pitch, § 25.1305 Powerplant instruments. for each reversing propeller. (c) For turbine engine-powered air- The following are required power- planes. In addition to the powerplant plant instruments: instruments required by paragraph (a) (a) For all airplanes. (1) A fuel pres- of this section, the following power- sure warning means for each engine, or plant instruments are required: a master warning means for all engines (1) A gas temperature indicator for with provision for isolating the indi- each engine. vidual warning means from the master (2) A fuel flowmeter indicator for warning means. each engine. (2) A fuel quantity indicator for each (3) A tachometer (to indicate the fuel tank. speed of the rotors with established (3) An oil quantity indicator for each limiting speeds) for each engine. oil tank. (4) A means to indicate, to the flight (4) An oil pressure indicator for each crew, the operation of each engine independent pressure oil system of starter that can be operated continu- each engine. ously but that is neither designed for

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continuous operation nor designed to means must be provided to indicate the prevent hazard if it failed. proper functioning of that system to (5) An indicator to indicate the func- the flight crew. tioning of the powerplant ice protec- [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as tion system for each engine. amended by Amdt. 25–35, 39 FR 1831, Jan. 15, (6) An indicator for the fuel strainer 1974; Amdt. 25–36, 39 FR 35461, Oct. 1, 1974; or filter required by § 25.997 to indicate Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. the occurrence of contamination of the 25–54, 45 FR 60173, Sept. 11, 1980; Amdt. 25–72, strainer or filter before it reaches the 55 FR 29785, July 20, 1990; Amdt. 25–115, 69 FR capacity established in accordance 40527, July 2, 2004] with § 25.997(d). (7) A warning means for the oil § 25.1307 Miscellaneous equipment. strainer or filter required by § 25.1019, if The following is required miscella- it has no bypass, to warn the pilot of neous equipment: the occurrence of contamination of the (a) [Reserved] strainer or filter screen before it (b) Two or more independent sources reaches the capacity established in ac- of electrical energy. cordance with § 25.1019(a)(2). (c) Electrical protective devices, as (8) An indicator to indicate the prop- prescribed in this part. er functioning of any heater used to (d) Two systems for two-way radio prevent ice clogging of fuel system communications, with controls for components. each accessible from each pilot station, (d) For turbojet engine powered air- designed and installed so that failure of planes. In addition to the powerplant one system will not preclude operation instruments required by paragraphs (a) of the other system. The use of a com- and (c) of this section, the following mon antenna system is acceptable if powerplant instruments are required: adequate reliability is shown. (1) An indicator to indicate thrust, or (e) Two systems for radio navigation, a parameter that is directly related to with controls for each accessible from thrust, to the pilot. The indication each pilot station, designed and in- must be based on the direct measure- stalled so that failure of one system ment of thrust or of parameters that will not preclude operation of the other are directly related to thrust. The indi- system. The use of a common antenna cator must indicate a change in thrust system is acceptable if adequate reli- resulting from any engine malfunction, ability is shown. damage, or deterioration. [Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as (2) A position indicating means to in- amended by Amdt. 25–46, 43 FR 50598, Oct. 30, dicate to the flightcrew when the 1978; Amdt. 25–54, 45 FR 60173, Sept. 11, 1980; thrust reversing device— Amdt. 25–72, 55 FR 29785, July 20, 1990] (i) Is not in the selected position, and (ii) Is in the reverse thrust position, § 25.1309 Equipment, systems, and infor each engine using a thrust revers- stallations. ing device. (a) The equipment, systems, and in- (3) An indicator to indicate rotor sys- stallations whose functioning is re- tem unbalance. quired by this subchapter, must be de- (e) For turbopropeller-powered air- signed to ensure that they perform planes. In addition to the powerplant their intended functions under any instruments required by paragraphs (a) foreseeable operating condition. and (c) of this section, the following (b) The airplane systems and associ- powerplant instruments are required: ated components, considered sepa- (1) A torque indicator for each en- rately and in relation to other systems, gine. must be designed so that— (2) Position indicating means to indi- (1) The occurrence of any failure con- cate to the flight crew when the pro- dition which would prevent the contin- peller blade angle is below the flight ued safe flight and landing of the air- low pitch position, for each propeller. plane is extremely improbable, and (f) For airplanes equipped with fluid (2) The occurrence of any other fail- systems (other than fuel) for thrust or ure conditions which would reduce the power augmentation, an approved capability of the airplane or the ability

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

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manner after the airplane is exposed to landing of an airplane may be designed lightning. and installed without meeting the provisions of paragraph (a) provided— [Doc. No. FAA–2010–0224, Amdt. 25–134, 76 FR 33135, June 8, 2011] (1) The system has previously been shown to comply with special condi- § 25.1317 High-intensity Radiated tions for HIRF, prescribed under § 21.16, Fields (HIRF) Protection. issued before December 1, 2007; (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: ment providing these functions is ex- (3) The instrument that most effec- posed to equipment HIRF test level 1 tively indicates altitude must be adja- or 2, as described in appendix L to this cent to and directly to the right of the part. instrument in the top center position; (c) Each electrical and electronic sys- and tem that performs a function whose (4) The instrument that most effec- failure would reduce the capability of tively indicates direction of flight the airplane or the ability of the must be adjacent to and directly below flightcrew to respond to an adverse op- the instrument in the top center posi- erating condition must be designed and tion. installed so the system is not adversely (c) Required powerplant instruments affected when the equipment providing must be closely grouped on the instru- the function is exposed to equipment ment panel. In addition— HIRF test level 3, as described in ap- (1) The location of identical power- pendix L to this part. plant instruments for the engines must (d) Before December 1, 2012, an elec- prevent confusion as to which engine trical or electronic system that per- each instrument relates; and forms a function whose failure would (2) Powerplant instruments vital to prevent the continued safe flight and the safe operation of the airplane must

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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, 1977] or by habitual reflexive action. When an alert is suppressed, there must be a § 25.1322 Flightcrew alerting. clear and unmistakable annunciation 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 actions, if any. (ii) Amber or yellow for caution alert indications. (2) Be readily and easily detectable and intelligible by the flightcrew under (iii) Any color except red or green for all foreseeable operating conditions, advisory alert indications. including conditions where multiple (2) Use visual coding techniques, to- alerts are provided. gether with other alerting function ele- (3) Be removed when the alerting ments on the flight deck, to distin- condition no longer exists. guish between warning, caution, and advisory alert indications, if they are (b) Alerts must conform to the following prioritization hierarchy based presented on monochromatic displays on the urgency of flightcrew awareness that are not capable of conforming to and response. the color convention in paragraph (e)(1) of this section. (1) Warning: For conditions that require immediate flightcrew awareness (f) Use of the colors red, amber, and and immediate flightcrew response. yellow on the flight deck for functions other than flightcrew alerting must be (2) Caution: For conditions that re- limited and must not adversely affect quire immediate flightcrew awareness flightcrew alerting. and subsequent flightcrew response. (3) Advisory: For conditions that re- [Amdt. 25–131, 75 FR 67209, Nov. 2, 2010] quire flightcrew awareness and may require subsequent flightcrew response. § 25.1323 Airspeed indicating system. (c) Warning and caution alerts must: For each airspeed indicating system, (1) Be prioritized within each cat- the following apply: egory, when necessary. (a) Each airspeed indicating instru- (2) Provide timely attention-getting ment must be approved and must be cues through at least two different calibrated to indicate true airspeed (at senses by a combination of aural, vis- sea level with a standard atmosphere) ual, or tactile indications. with a minimum practicable instru- (3) Permit each occurrence of the at- ment calibration error when the cor- tention-getting cues required by para- responding pitot and static pressures graph (c)(2) of this section to be ac- are applied. knowledged and suppressed, unless (b) Each system must be calibrated they are required to be continuous. to determine the system error (that is, (d) The alert function must be de- the relation between IAS and CAS) in signed to minimize the effects of false flight and during the accelerated take- and nuisance alerts. In particular, it off ground run. The ground run calibra- must be designed to: tion must be determined—

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(1) From 0.8 of the minimum value of the achievement of a steady climbing V1 to the maximum value of V2, consid- condition. ering the approved ranges of altitude (g) The effects of airspeed indicating and weight; and system lag may not introduce signifi- (2) With the flaps and power settings cant takeoff indicated airspeed bias, or corresponding to the values determined significant errors in takeoff or accel- in the establishment of the takeoff erate-stop distances. path under § 25.111 assuming that the (h) Each system must be arranged, so critical engine fails at the minimum far as practicable, to prevent malfunc- value of V1. tion or serious error due to the entry of (c) The airspeed error of the installa- moisture, dirt, or other substances. tion, excluding the airspeed indicator (i) Each system must have a heated instrument calibration error, may not pitot tube or an equivalent means of exceed three percent or five knots, preventing malfunction in the heavy whichever is greater, throughout the rain conditions defined in Table 1 of speed range, from— this section; mixed phase and ice crys- (1) VMO to 1.23 VSR1, with flaps re- tal conditions as defined in part 33, Ap- tracted; and pendix D, of this chapter; the icing con- (2) 1.23 VSR0 to VFE with flaps in the ditions defined in Appendix C of this landing position. part; and the following icing conditions (d) From 1.23 VSR to the speed at specified in Appendix O of this part: which stall warning begins, the IAS (1) For airplanes certificated in ac- must change perceptibly with CAS and cordance with § 25.1420(a)(1), the icing in the same sense, and at speeds below conditions that the airplane is certified stall warning speed the IAS must not to safely exit following detection. change in an incorrect sense. (2) For airplanes certificated in ac- 2 (e) From VMO to VMO + ⁄3 (VDF ¥ cordance with § 25.1420(a)(2), the icing VMO), the IAS must change perceptibly conditions that the airplane is certified with CAS and in the same sense, and at to safely operate in and the icing con- higher speeds up to VDF the IAS must ditions that the airplane is certified to not change in an incorrect sense. safely exit following detection. (f) There must be no indication of (3) For airplanes certificated in ac- airspeed that would cause undue dif- cordance with § 25.1420(a)(3) and for air- ficulty to the pilot during the takeoff planes not subject to § 25.1420, all icing between the initiation of rotation and conditions.

TABLE 1—HEAVY RAIN CONDITIONS FOR AIRSPEED INDICATING SYSTEM TESTS

Altitude range Liquid water Horizontal extent Droplet MVD content μ (ft) (m) (g/m3) (km) (nmiles) ( m)

0 to 10 000 ...... 0 to 3000 ...... 1 100 50 1000 6 5 3 2000 15 1 0 .5 2000

(j) Where duplicate airspeed indica- § 25.1324 Angle of attack system. tors are required, their respective pitot Each angle of attack system sensor tubes must be far enough apart to must be heated or have an equivalent avoid damage to both tubes in a colli- means of preventing malfunction in the sion with a bird. heavy rain conditions defined in Table [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as 1 of § 25.1323, the mixed phase and ice amended by Amdt. 25–57, 49 FR 6849, Feb. 23, crystal conditions as defined in part 33, 1984; Amdt. 25–108, 67 FR 70828, Nov. 26, 2002; Appendix D, of this chapter, the icing Amdt. 25–109, 67 FR 76656, Dec. 12, 2002; Amdt. conditions defined in Appendix C of 25–140, 79 FR 65526, Nov. 4, 2014] this part, and the following icing conditions specified in Appendix O of this part:

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(a) For airplanes certificated in ac- (2) It is airtight except for the port cordance with § 25.1420(a)(1), the icing into the atmosphere. A proof test must conditions that the airplane is certified be conducted to demonstrate the integ- to safely exit following detection. rity of the static pressure system in (b) For airplanes certificated in ac- the following manner: cordance with § 25.1420(a)(2), the icing (i) Unpressurized airplanes. Evacuate conditions that the airplane is certified the static pressure system to a pres- to safely operate in and the icing con- sure differential of approximately 1 ditions that the airplane is certified to inch of mercury or to a reading on the safely exit following detection. altimeter, 1,000 feet above the airplane (c) For airplanes certificated in ac- elevation at the time of the test. With- cordance with § 25.1420(a)(3) and for air- out additional pumping for a period of planes not subject to § 25.1420, all icing 1 minute, the loss of indicated altitude conditions. must not exceed 100 feet on the altim- [Amdt. 25–140, 79 FR 65527, Nov. 4, 2014] eter. (ii) Pressurized airplanes. Evacuate § 25.1325 Static pressure systems. the static pressure system until a pres- (a) Each instrument with static air sure differential equivalent to the max- case connections must be vented to the imum cabin pressure differential for outside atmosphere through an appro- which the airplane is type certificated priate piping system. is achieved. Without additional pump- (b) Each static port must be designed ing for a period of 1 minute, the loss of and located so that: indicated altitude must not exceed 2 (1) The static pressure system per- percent of the equivalent altitude of formance is least affected by airflow the maximum cabin differential pres- variation, or by moisture or other for- sure or 100 feet, whichever is greater. eign matter; and (d) Each pressure altimeter must be (2) The correlation between air pres- approved and must be calibrated to insure in the static pressure system and dicate pressure altitude in a standard true ambient atmospheric static pres- atmosphere, with a minimum prac- sure is not changed when the airplane ticable calibration error when the cor- is exposed to the icing conditions de- responding static pressures are applied. fined in Appendix C of this part, and (e) Each system must be designed and the following icing conditions specified installed so that the error in indicated in Appendix O of this part: pressure altitude, at sea level, with a (i) For airplanes certificated in ac- standard atmosphere, excluding instru- cordance with § 25.1420(a)(1), the icing ment calibration error, does not result conditions that the airplane is certified in an error of more than ±30 feet per 100 to safely exit following detection. knots speed for the appropriate con- (ii) For airplanes certificated in ac- figuration in the speed range between cordance with § 25.1420(a)(2), the icing 1.23 VSR0 with flaps extended and 1.7 conditions that the airplane is certified VSR1 with flaps retracted. However, the to safely operate in and the icing con- error need not be less than ±30 feet. ditions that the airplane is certified to (f) If an altimeter system is fitted safely exit following detection. with a device that provides corrections (iii) For airplanes certificated in ac- to the altimeter indication, the device cordance with § 25.1420(a)(3) and for air- must be designed and installed in such planes not subject to § 25.1420, all icing manner that it can be bypassed when it conditions. malfunctions, unless an alternate al- (c) The design and installation of the timeter system is provided. Each cor- static pressure system must be such rection device must be fitted with a that— means for indicating the occurrence of (1) Positive drainage of moisture is reasonably probable malfunctions, in- provided; chafing of the tubing and excluding power failure, to the flight cessive distortion or restriction at crew. The indicating means must be ef- bends in the tubing is avoided; and the fective for any cockpit lighting condi- materials used are durable, suitable for tion likely to occur. the purpose intended, and protected (g) Except as provided in paragraph against corrosion; and (h) of this section, if the static pressure

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system incorporates both a primary autopilot quick disengagement con- and an alternate static pressure source, trols must be located on both control the means for selecting one or the wheels (or equivalent). The autothrust other source must be designed so quick disengagement controls must be that— located on the thrust control levers. (1) When either source is selected, the Quick disengagement controls must be other is blocked off; and readily accessible to each pilot while (2) Both sources cannot be blocked operating the control wheel (or equiva- off simultaneously. lent) and thrust control levers. (h) For unpressurized airplanes, para- (b) The effects of a failure of the sys- graph (g)(1) of this section does not tem to disengage the autopilot or apply if it can be demonstrated that autothrust functions when manually the static pressure system calibration, commanded by the pilot must be as- when either static pressure source is sessed in accordance with the require- selected, is not changed by the other ments of § 25.1309. static pressure source being open or (c) Engagement or switching of the blocked. flight guidance system, a mode, or a [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as sensor may not cause a transient re- amended by Amdt. 25–5, 30 FR 8261, June 29, sponse of the airplane’s flight path any 1965; Amdt. 25–12, 32 FR 7587, May 24, 1967; greater than a minor transient, as de- Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. fined in paragraph (n)(1) of this section. 25–108, 67 FR 70828, Nov. 26, 2002; Amdt. 25– (d) Under normal conditions, the dis- 140, 79 FR 65527, Nov. 4, 2014] engagement of any automatic control § 25.1326 Pitot heat indication systems. function of a flight guidance system may not cause a transient response of If a flight instrument pitot heating the airplane’s flight path any greater system is installed, an indication sys- than a minor transient. tem must be provided to indicate to (e) Under rare normal and non-nor- the flight crew when that pitot heating mal conditions, disengagement of any system is not operating. The indication automatic control function of a flight system must comply with the following guidance system may not result in a requirements: transient any greater than a signifi- (a) The indication provided must in- cant transient, as defined in paragraph corporate an amber light that is in (n)(2) of this section. clear view of a flight crewmember. (f) The function and direction of mo- (b) The indication provided must be tion of each command reference con- designed to alert the flight crew if ei- trol, such as heading select or vertical ther of the following conditions exist: speed, must be plainly indicated on, or (1) The pitot heating system is adjacent to, each control if necessary switched ‘‘off’’. to prevent inappropriate use or confu- (2) The pitot heating system is sion. switched ‘‘on’’ and any pitot tube heat- (g) Under any condition of flight ap- ing element is inoperative. propriate to its use, the flight guidance [Amdt. 25–43, 43 FR 10339, Mar. 13, 1978] system may not produce hazardous loads on the airplane, nor create haz- § 25.1327 Magnetic direction indicator. ardous deviations in the flight path. (a) Each magnetic direction indicator This applies to both fault-free oper- must be installed so that its accuracy ation and in the event of a malfunc- is not excessively affected by the air- tion, and assumes that the pilot begins plane’s vibration or magnetic fields. corrective action within a reasonable (b) The compensated installation period of time. may not have a deviation, in level (h) When the flight guidance system flight, greater than 10 degrees on any is in use, a means must be provided to heading. avoid excursions beyond an acceptable margin from the speed range of the § 25.1329 Flight guidance system. normal flight envelope. If the airplane (a) Quick disengagement controls for experiences an excursion outside this the autopilot and autothrust functions range, a means must be provided to must be provided for each pilot. The prevent the flight guidance system

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from providing guidance or control to or recover to the normal flight enve- an unsafe speed. lope, any of the following: (i) The flight guidance system func- (i) Exceptional piloting skill, alert- tions, controls, indications, and alerts ness, or strength. must be designed to minimize (ii) Forces applied by the pilot which flightcrew errors and confusion con- are greater than those specified in cerning the behavior and operation of § 25.143(c). the flight guidance system. Means (iii) Accelerations or attitudes in the must be provided to indicate the cur- airplane that might result in further rent mode of operation, including any hazard to secured or non-secured occu- armed modes, transitions, and rever- pants. sions. Selector switch position is not an acceptable means of indication. The [Doc. No. FAA–2004–18775, 71 FR 18191, Apr. 11, 2006] controls and indications must be grouped and presented in a logical and § 25.1331 Instruments using a power consistent manner. The indications supply. must be visible to each pilot under all expected lighting conditions. (a) For each instrument required by (j) Following disengagement of the § 25.1303(b) that uses a power supply, autopilot, a warning (visual and audi- the following apply: tory) must be provided to each pilot (1) Each instrument must have a vis- and be timely and distinct from all ual means integral with, the instru- other cockpit warnings. ment, to indicate when power adequate (k) Following disengagement of the to sustain proper instrument perform- autothrust function, a caution must be ance is not being supplied. The power provided to each pilot. must be measured at or near the point (l) The autopilot may not create a where it enters the instruments. For potential hazard when the flightcrew electric instruments, the power is con- applies an override force to the flight sidered to be adequate when the volt- controls. age is within approved limits. (m) During autothrust operation, it (2) Each instrument must, in the must be possible for the flightcrew to event of the failure of one power move the thrust levers without requir- source, be supplied by another power ing excessive force. The autothrust source. This may be accomplished may not create a potential hazard automatically or by manual means. when the flightcrew applies an override (3) If an instrument presenting navi- force to the thrust levers. gation data receives information from (n) For purposes of this section, a sources external to that instrument transient is a disturbance in the con- and loss of that information would trol or flight path of the airplane that render the presented data unreliable, is not consistent with response to the instrument must incorporate a vis- flightcrew inputs or environmental ual means to warn the crew, when such conditions. loss of information occurs, that the (1) A minor transient would not sig- presented data should not be relied nificantly reduce safety margins and upon. would involve flightcrew actions that (b) As used in this section, ‘‘instru- are well within their capabilities. A ment’’ includes devices that are phys- minor transient may involve a slight ically contained in one unit, and de- increase in flightcrew workload or vices that are composed of two or more some physical discomfort to passengers physically separate units or compo- or cabin crew. nents connected together (such as a re- (2) A significant transient may lead mote indicating gyroscopic direction to a significant reduction in safety indicator that includes a magnetic margins, an increase in flightcrew sensing element, a gyroscopic unit, an workload, discomfort to the flightcrew, amplifier and an indicator connected or physical distress to the passengers together). or cabin crew, possibly including non- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as fatal injuries. Significant transients do amended by Amdt. 25–41, 42 FR 36970, July 18, not require, in order to remain within 1977]

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§ 25.1333 Instrument systems. or equivalent units, of usable fuel in each tank during flight. In addition— For systems that operate the instruments required by § 25.1303(b) which are (1) Each fuel quantity indicator must located at each pilot’s station— be calibrated to read ‘‘zero’’ during (a) Means must be provided to con- level flight when the quantity of fuel nect the required instruments at the remaining in the tank is equal to the first pilot’s station to operating sys- unusable fuel supply determined under tems which are independent of the op- § 25.959; erating systems at other flight crew (2) Tanks with interconnected outlets stations, or other equipment; and airspaces may be treated as one (b) The equipment, systems, and in- tank and need not have separate indi- stallations must be designed so that cators; and one display of the information essen- (3) Each exposed sight gauge, used as tial to the safety of flight which is pro- a fuel quantity indicator, must be provided by the instruments, including at- tected against damage. titude, direction, airspeed, and altitude (c) Fuel flowmeter system. If a fuel will remain available to the pilots, flowmeter system is installed, each without additional crewmember ac- metering component must have a tion, after any single failure or com- means for bypassing the fuel supply if bination of failures that is not shown malfunction of that component se- to be extremely improbable; and verely restricts fuel flow. (c) Additional instruments, systems, (d) Oil quantity indicator. There must or equipment may not be connected to be a stick gauge or equivalent means the operating systems for the required to indicate the quantity of oil in each instruments, unless provisions are tank. If an oil transfer or reserve oil made to ensure the continued normal supply system is installed, there must functioning of the required instru- be a means to indicate to the flight ments in the event of any malfunction crew, in flight, the quantity of oil in of the additional instruments, systems, each tank. or equipment which is not shown to be extremely improbable. (e) Turbopropeller blade position indicator. Required turbopropeller blade [Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as position indicators must begin indi- amended by Amdt. 25–41, 42 FR 36970, July 18, cating before the blade moves more 1977] than eight degrees below the flight low § 25.1337 Powerplant instruments. pitch stop. The source of indication must directly sense the blade position. (a) Instruments and instrument lines. (f) Fuel pressure indicator. There must (1) Each powerplant and auxiliary be means to measure fuel pressure, in power unit instrument line must meet each system supplying reciprocating the requirements of §§ 25.993 and 25.1183. engines, at a point downstream of any (2) Each line carrying flammable fuel pump except fuel injection pumps. fluids under pressure must— In addition— (i) Have restricting orifices or other (1) If necessary for the maintenance safety devices at the source of pressure of proper fuel delivery pressure, there to prevent the escape of excessive fluid if the line fails; and must be a connection to transmit the carburetor air intake static pressure to (ii) Be installed and located so that the proper pump relief valve connec- the escape of fluids would not create a hazard. tion; and (3) Each powerplant and auxiliary (2) If a connection is required under power unit instrument that utilizes paragraph (f)(1) of this section, the flammable fluids must be installed and gauge balance lines must be independ- located so that the escape of fluid ently connected to the carburetor inlet would not create a hazard. pressure to avoid erroneous readings. (b) Fuel quantity indicator. There [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as must be means to indicate to the flight amended by Amdt. 25–40, 42 FR 15044, Mar. 17, crewmembers, the quantity, in gallons 1977]

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ELECTRICAL SYSTEMS AND EQUIPMENT with the normal electrical power (electrical power sources excluding the bat- § 25.1351 General. tery) inoperative, with critical type (a) Electrical system capacity. The re- fuel (from the standpoint of flameout quired generating capacity, and num- and restart capability), and with the ber and kinds of power sources must— airplane initially at the maximum cer- (1) Be determined by an electrical tificated altitude. Parts of the elec- load analysis; and trical system may remain on if— (2) Meet the requirements of § 25.1309. (1) A single malfunction, including a (b) Generating system. The generating wire bundle or junction box fire, can- system includes electrical power not result in loss of both the part sources, main power busses, trans- turned off and the part turned on; and mission cables, and associated control, (2) The parts turned on are elec- regulation, and protective devices. It trically and mechanically isolated must be designed so that— from the parts turned off. (1) Power sources function properly [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as when independent and when connected amended by Amdt. 25–41, 42 FR 36970, July 18, in combination; 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990] (2) No failure or malfunction of any power source can create a hazard or § 25.1353 Electrical equipment and in- impair the ability of remaining sources stallations. to supply essential loads; (a) Electrical equipment and controls (3) The system voltage and frequency must be installed so that operation of (as applicable) at the terminals of all any one unit or system of units will essential load equipment can be main- not adversely affect the simultaneous tained within the limits for which the operation of any other electrical unit equipment is designed, during any or system essential to safe operation. probable operating condition; and Any electrical interference likely to be (4) System transients due to switch- present in the airplane must not result ing, fault clearing, or other causes do in hazardous effects on the airplane or not make essential loads inoperative, its systems. and do not cause a smoke or fire haz- (b) Storage batteries must be de- ard. signed and installed as follows: (5) There are means accessible, in (1) Safe cell temperatures and pres- flight, to appropriate crewmembers for sures must be maintained during any the individual and collective dis- probable charging or discharging con- connection of the electrical power dition. No uncontrolled increase in cell sources from the system. temperature may result when the bat- (6) There are means to indicate to ap- tery is recharged (after previous com- propriate crewmembers the generating plete discharge)— system quantities essential for the safe (i) At maximum regulated voltage or operation of the system, such as the power; voltage and current supplied by each (ii) During a flight of maximum dura- generator. tion; and (c) External power. If provisions are (iii) Under the most adverse cooling made for connecting external power to condition likely to occur in service. the airplane, and that external power (2) Compliance with paragraph (b)(1) can be electrically connected to equip- of this section must be shown by test ment other than that used for engine unless experience with similar bat- starting, means must be provided to teries and installations has shown that ensure that no external power supply maintaining safe cell temperatures and having a reverse polarity, or a reverse pressures presents no problem. phase sequence, can supply power to (3) No explosive or toxic gases emit- the airplane’s electrical system. ted by any battery in normal oper- (d) Operation without normal electrical ation, or as the result of any probable power. It must be shown by analysis, malfunction in the charging system or tests, or both, that the airplane can be battery installation, may accumulate operated safely in VFR conditions, for in hazardous quantities within the air- a period of not less than five minutes, plane.

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(4) No corrosive fluids or gases that the electrical system and hazard to the may escape from the battery may dam- airplane in the event of wiring faults or age surrounding airplane structures or serious malfunction of the system or adjacent essential equipment. connected equipment. (5) Each nickel cadmium battery in- (b) The protective and control de- stallation must have provisions to pre- vices in the generating system must be vent any hazardous effect on structure designed to de-energize and disconnect or essential systems that may be faulty power sources and power trans- caused by the maximum amount of mission equipment from their associ- heat the battery can generate during a ated busses with sufficient rapidity to short circuit of the battery or of indi- provide protection from hazardous vidual cells. over-voltage and other malfunctioning. (6) Nickel cadmium battery installa- (c) Each resettable circuit protective tions must have— device must be designed so that, when (i) A system to control the charging an overload or circuit fault exists, it rate of the battery automatically so as will open the circuit irrespective of the to prevent battery overheating; position of the operating control. (ii) A battery temperature sensing (d) If the ability to reset a circuit and over-temperature warning system with a means for disconnecting the breaker or replace a fuse is essential to battery from its charging source in the safety in flight, that circuit breaker or event of an over-temperature condi- fuse must be located and identified so tion; or that it can be readily reset or replaced (iii) A battery failure sensing and in flight. Where fuses are used, there warning system with a means for dis- must be spare fuses for use in flight connecting the battery from its charg- equal to at least 50% of the number of ing source in the event of battery fail- fuses of each rating required for com- ure. plete circuit protection. (c) Electrical bonding must provide (e) Each circuit for essential loads an adequate electrical return path must have individual circuit protec- under both normal and fault condition. However, individual protection tions, on airplanes having grounded for each circuit in an essential load electrical systems. system (such as each position light circuit in a system) is not required. [Amdt. 25–123, 72 FR 63405, Nov. 8, 2007] (f) For airplane systems for which § 25.1355 Distribution system. the ability to remove or reset power during normal operations is necessary, (a) The distribution system includes the system must be designed so that the distribution busses, their associ- circuit breakers are not the primary ated feeders, and each control and pro- means to remove or reset system power tective device. unless specifically designed for use as a (b) [Reserved] switch. (c) If two independent sources of electrical power for particular equipment (g) Automatic reset circuit breakers or systems are required by this chap- may be used as integral protectors for ter, in the event of the failure of one electrical equipment (such as thermal power source for such equipment or cut-outs) if there is circuit protection system, another power source (includ- to protect the cable to the equipment. ing its separate feeder) must be auto- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as matically provided or be manually se- amended by Amdt. 25–123, 72 FR 63405, Nov. 8, lectable to maintain equipment or sys- 2007] tem operation. § 25.1360 Precautions against injury. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5679, Apr. 8, (a) Shock. The electrical system 1970; Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] must be designed to minimize risk of electric shock to crew, passengers, and § 25.1357 Circuit protective devices. servicing personnel and to mainte- (a) Automatic protective devices nance personnel using normal pre- must be used to minimize distress to cautions.

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(b) Burns. The temperature of any (c) Domestic appliances, particularly part that may be handled by a crew- those in galley areas, must be installed member during normal operations or protected so as to prevent damage or must not cause dangerous inadvertent contamination of other equipment or movement by the crewmember or in- systems from fluids or vapors which jury to the crewmember. may be present during normal operation or as a result of spillage, if such [Amdt. 25–123, 72 FR 63406, Nov. 8, 2007] damage or contamination could create § 25.1362 Electrical supplies for emer- a hazardous condition. gency conditions. (d) Unless compliance with § 25.1309(b) A suitable electrical supply must be is provided by the circuit protective provided to those services required for device required by § 25.1357(a), electric emergency procedures after an emer- motors and transformers, including gency landing or ditching. The circuits those installed in domestic systems, for these services must be designed, must have a suitable thermal protec- protected, and installed so that the tion device to prevent overheating risk of the services being rendered inef- under normal operation and failure fective under these emergency condi- conditions, if overheating could create tions is minimized. a smoke or fire hazard. [Amdt. 25–123, 72 FR 63406, Nov. 8, 2007] [Amdt. 25–123, 72 FR 63406, Nov. 8, 2007]

§ 25.1363 Electrical system tests. LIGHTS (a) When laboratory tests of the elec- § 25.1381 Instrument lights. trical system are conducted— (a) The instrument lights must— (1) The tests must be performed on a (1) Provide sufficient illumination to mock-up using the same generating make each instrument, switch and equipment used in the airplane; other device necessary for safe oper- (2) The equipment must simulate the ation easily readable unless sufficient electrical characteristics of the dis- illumination is available from another tribution wiring and connected loads to source; and the extent necessary for valid test re- (2) Be installed so that— sults; and (i) Their direct rays are shielded from (3) Laboratory generator drives must the pilot’s eyes; and simulate the actual prime movers on (ii) No objectionable reflections are the airplane with respect to their reac- visible to the pilot. tion to generator loading, including (b) Unless undimmed instrument loading due to faults. lights are satisfactory under each ex- (b) For each flight condition that pected flight condition, there must be a cannot be simulated adequately in the means to control the intensity of illu- laboratory or by ground tests on the mination. airplane, flight tests must be made. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as § 25.1365 Electrical appliances, motors, amended by Amdt. 25–72, 55 FR 29785, July 20, and transformers. 1990] (a) Domestic appliances must be designed and installed so that in the § 25.1383 Landing lights. event of failures of the electrical sup- (a) Each landing light must be apply or control system, the require- proved, and must be installed so that— ments of § 25.1309(b), (c), and (d) will be (1) No objectionable glare is visible satisfied. Domestic appliances are to the pilot; items such as cooktops, ovens, coffee (2) The pilot is not adversely affected makers, water heaters, refrigerators, by halation; and and toilet flush systems that are (3) It provides enough light for night placed on the airplane to provide serv- landing. ice amenities to passengers. (b) Except when one switch is used (b) Galleys and cooking appliances for the lights of a multiple light instal- must be installed in a way that mini- lation at one location, there must be a mizes risk of overheat or fire. separate switch for each light.

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(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 far apart as practicable and installed rear position light and whose elements forward on the airplane so that, with make an angle of 30° with a vertical the airplane in the normal flying posi- line passing through the rear position tion, the red light is on the left side light. and the green light is on the right side. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as Each light must be approved. amended by Amdt. 25–30, 36 FR 21278, Nov. 5, (c) Rear position light. The rear posi- 1971] tion light must be a white light mounted as far aft as practicable on the tail § 25.1389 Position light distribution or on each wing tip, and must be ap- and intensities. proved. (a) General. The intensities prescribed (d) Light covers and color filters. Each in this section must be provided by new light cover or color filter must be at equipment with light covers and color least flame resistant and may not filters in place. Intensities must be de- change color or shape or lose any ap- termined with the light source oper- preciable light transmission during ating at a steady value equal to the av- normal use. erage luminous output of the source at [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as the normal operating voltage of the amended by Amdt. 25–38, 41 FR 55468, Dec. 20, airplane. The light distribution and in- 1976] tensity of each position light must meet the requirements of paragraph (b) § 25.1387 Position light system dihe- of this section. dral angles. (b) Forward and rear position lights. (a) Except as provided in paragraph The light distribution and intensities (e) of this section, each forward and of forward and rear position lights rear position light must, as installed, must be expressed in terms of min- show unbroken light within the dihe- imum intensities in the horizontal dral angles described in this section. plane, minimum intensities in any (b) Dihedral angle L (left) is formed vertical plane, and maximum inten- by two intersecting vertical planes, the sities in overlapping beams, within di- first parallel to the longitudinal axis of hedral angles L, R, and A, and must the airplane, and the other at 110 de- meet the following requirements: grees to the left of the first, as viewed (1) Intensities in the horizontal plane. when looking forward along the longi- Each intensity in the horizontal plane tudinal axis. (the plane containing the longitudinal (c) Dihedral angle R (right) is formed axis of the airplane and perpendicular by two intersecting vertical planes, the to the plane of symmetry of the air- first parallel to the longitudinal axis of plane) must equal or exceed the values the airplane, and the other at 110 de- in § 25.1391. grees to the right of the first, as viewed (2) Intensities in any vertical plane. when looking forward along the longi- Each intensity in any vertical plane tudinal axis. (the plane perpendicular to the hori- (d) Dihedral angle A (aft) is formed zontal plane) must equal or exceed the by two intersecting vertical planes appropriate value in § 25.1393, where I is

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the minimum intensity prescribed in § 25.1395 Maximum intensities in over- § 25.1391 for the corresponding angles in lapping beams of forward and rear the horizontal plane. position lights. (3) Intensities in overlaps between adja- No position light intensity may ex- cent signals. No intensity in any over- ceed the applicable values in the fol- lap between adjacent signals may ex- lowing table, except as provided in ceed the values given in § 25.1395, except § 25.1389(b)(3). that higher intensities in overlaps may be used with main beam intensities Maximum intensity substantially greater than the minima Overlaps Area A Area B specified in §§ 25.1391 and 25.1393 if the (candles) (candles) overlap intensities in relation to the Green in dihedral angle L ...... 10 1 main beam intensities do not adversely Red in dihedral angle R ...... 10 1 affect signal clarity. When the peak in- Green in dihedral angle A ...... 5 1 tensity of the forward position lights is Red in dihedral angle A ...... 5 1 more than 100 candles, the maximum Rear white in dihedral angle L ...... 5 1 Rear white in dihedral angle R ..... 5 1 overlap intensities between them may exceed the values given in § 25.1395 if Where— the overlap intensity in Area A is not (a) Area A includes all directions in more than 10 percent of peak position the adjacent dihedral angle that pass light intensity and the overlap inten- through the light source and intersect sity in Area B is not greater than 2.5 the common boundary plane at more percent of peak position light inten- than 10 degrees but less than 20 de- sity. grees; and § 25.1391 Minimum intensities in the (b) Area B includes all directions in horizontal plane of forward and the adjacent dihedral angle that pass rear position lights. through the light source and intersect Each position light intensity must the common boundary plane at more equal or exceed the applicable values in than 20 degrees. the following table: § 25.1397 Color specifications. Angle from right Each position light color must have or left of longitu- Dihedral angle (light in- Intensity the applicable International Commis- cluded) dinal axis, meas- (candles) ured from dead sion on Illumination chromaticity co- ahead ordinates as follows: L and R (forward red and 0° to 10° ...... 40 (a) Aviation red— green). 10° to 20° ...... 30 20° to 110° ...... 5 y is not greater than 0.335; and A (rear white) ...... 110° to 180° ...... 20 z is not greater than 0.002. (b) Aviation green— § 25.1393 Minimum intensities in any vertical plane of forward and rear x is not greater than 0.440¥0.320y ; position lights. x is not greater than y¥0.170; and Each position light intensity must y is not less than 0.390¥0.170x. equal or exceed the applicable values in (c) Aviation white— the following table: x is not less than 0.300 and not greater than Angle above or below the horizontal plane Intensity, l 0.540; y is not less than x¥0.040; or y0¥0.010, which- 0° ...... 1.00 ever is the smaller; and ° ° 0 to 5 ...... 0.90 y is not greater than x + 0.020 nor 5° to 10° ...... 0.80 0.636¥0.400x; 10° to 15° ...... 0.70 15° to 20° ...... 0.50 Where y0 is the y coordinate of the Planckian 20° to 30° ...... 0.30 radiator for the value of x considered. ° ° 30 to 40 ...... 0.10 [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ° ° 40 to 90 ...... 0.05 amended by Amdt. 25–27, 36 FR 12972, July 10, 1971]

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§ 25.1399 Riding light. (e) Light intensity. The minimum light intensities in all vertical planes, (a) Each riding (anchor) light required 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 t moored or drifting on the water. ∫ 2 I(t)dt (b) Externally hung lights may be 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 (1) Consists of one or more approved of time. anticollision lights located so that t2—t1 = flash time interval (seconds). their light will not impair the crew’s Normally, the maximum value of effec- vision or detract from the conspicuity tive intensity is obtained when t and t of the position lights; and 2 1 are chosen so that the effective inten- (2) Meets the requirements of paragraphs (b) through (f) of this section. sity is equal to the instantaneous in- (b) Field of coverage. The system must tensity at t2 and t1. consist of enough lights to illuminate (f) Minimum effective intensities for the vital areas around the airplane anticollision lights. Each anticollision considering the physical configuration light effective intensity must equal or and flight characteristics of the air- exceed the applicable values in the fol- plane. The field of coverage must ex- lowing table. tend in each direction within at least Effective 75 degrees above and 75 degrees below Angle above or below the horizontal plane intensity the horizontal plane of the airplane, (candles) except that a solid angle or angles of 0° to 5° ...... 400 obstructed visibility totaling not more 5° to 10° ...... 240 than 0.03 steradians is allowable within 10° to 20° ...... 80 a solid angle equal to 0.15 steradians 20° to 30° ...... 40 centered about the longitudinal axis in 30° to 75° ...... 20 the rearward direction. (c) Flashing characteristics. The ar- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as rangement of the system, that is, the amended by Amdt. 25–27, 36 FR 12972, July 10, number of light sources, beam width, 1971; Amdt. 25–41, 42 FR 36970, July 18, 1977] speed of rotation, and other characteristics, must give an effective flash fre- § 25.1403 Wing icing detection lights. quency of not less than 40, nor more Unless operations at night in known than 100 cycles per minute. The effec- or forecast icing conditions are prohib- tive flash frequency is the frequency at ited by an operating limitation, a which the airplane’s complete anti- means must be provided for illu- collision light system is observed from minating or otherwise determining the a distance, and applies to each sector of light including any overlaps that formation of ice on the parts of the exist when the system consists of more wings that are critical from the stand- than one light source. In overlaps, point of ice accumulation. Any illu- flash frequencies may exceed 100, but mination that is used must be of a type not 180 cycles per minute. that will not cause glare or reflection (d) Color. Each anticollision light that would handicap crewmembers in must be either aviation red or aviation the performance of their duties. white and must meet the applicable re- [Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] quirements of § 25.1397.

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SAFETY EQUIPMENT (2) Be arranged to allow the life lines to be used to enable the occupants to § 25.1411 General. stay on the wing after ditching. (a) Accessibility. Required safety [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as equipment to be used by the crew in an amended by Amdt. 25–32, 37 FR 3972, Feb. 24, emergency must be readily accessible. 1972; Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; (b) Stowage provisions. Stowage provi- Amdt. 25–53, 45 FR 41593, June 19, 1980; Amdt. sions for required emergency equip- 25–70, 54 FR 43925, Oct. 27, 1989; Amdt. 25–79, 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 equipment is directly accessible and its loca- § 25.1415 Ditching equipment. tion is obvious; and (a) Ditching equipment used in air- (2) Protect the safety equipment planes to be certificated for ditching from inadvertent damage. under § 25.801, and required by the oper- (c) Emergency exit descent device. The ating rules of this chapter, must meet stowage provisions for the emergency the requirements of this section. exit descent devices required by (b) Each liferaft and each life pre- § 25.810(a) must be at each exit for server must be approved. In addition— which they are intended. (1) Unless excess rafts of enough ca- (d) Liferafts. (1) The stowage provi- pacity are provided, the buoyancy and sions for the liferafts described in seating capacity beyond the rated ca- § 25.1415 must accommodate enough pacity of the rafts must accommodate rafts for the maximum number of occu- all occupants of the airplane in the pants for which certification for ditch- event of a loss of one raft of the largest ing is requested. rated capacity; and (2) Liferafts must be stowed near (2) Each raft must have a trailing exits through which the rafts can be line, and must have a static line de- launched during an unplanned ditch- signed to hold the raft near the air- ing. plane but to release it if the airplane (3) Rafts automatically or remotely becomes totally submerged. released outside the airplane must be (c) Approved survival equipment attached to the airplane by means of must be attached to each liferaft. (d) There must be an approved sur- the static line prescribed in § 25.1415. vival type emergency locator trans- (4) The stowage provisions for each mitter for use in one life raft. portable liferaft must allow rapid de- (e) For airplanes not certificated for tachment and removal of the raft for ditching under § 25.801 and not having use at other than the intended exits. approved life preservers, there must be (e) Long-range signaling device. The an approved flotation means for each stowage provisions for the long-range occupant. This means must be within signaling device required by § 25.1415 easy reach of each seated occupant and must be near an exit available during must be readily removable from the an unplanned ditching. airplane. (f) Life preserver stowage provisions. The stowage provisions for life pre- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–29, 36 FR 18722, Sept. servers described in § 25.1415 must ac- 21, 1971; Amdt. 25–50, 45 FR 38348, June 9, 1980; commodate one life preserver for each Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. occupant for which certification for 25–82, 59 FR 32057, June 21, 1994] ditching is requested. Each life preserver must be within easy reach of § 25.1419 Ice protection. each seated occupant. If the applicant seeks certification (g) Life line stowage provisions. If cer- for flight in icing conditions, the air- tification for ditching under § 25.801 is plane must be able to safely operate in requested, there must be provisions to the continuous maximum and inter- store life lines. These provisions mittent maximum icing conditions of must— appendix C. To establish this— (1) Allow one life line to be attached (a) An analysis must be performed to to each side of the fuselage; and establish that the ice protection for

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the various components of the airplane phases of flight, the requirements of is adequate, taking into account the paragraph (e) of this section are appli- various airplane operational configura- cable to all phases of flight. tions; and (g) After the initial activation of the (b) To verify the ice protection anal- airframe ice protection system— ysis, to check for icing anomalies, and (1) The ice protection system must be to demonstrate that the ice protection designed to operate continuously; system and its components are effec- (2) The airplane must be equipped tive, the airplane or its components with a system that automatically cy- must be flight tested in the various cles the ice protection system; or operational configurations, in meas- (3) An ice detection system must be ured natural atmospheric icing condi- provided to alert the flightcrew each tions and, as found necessary, by one time the ice protection system must be or more of the following means: cycled. (1) Laboratory dry air or simulated (h) Procedures for operation of the icing tests, or a combination of both, of ice protection system, including acti- the components or models of the com- vation and deactivation, must be estab- ponents. lished and documented in the Airplane (2) Flight dry air tests of the ice pro- Flight Manual. tection system as a whole, or of its in- [Amdt. 25–72, 55 FR 29785, July 20, 1990, as dividual components. amended by Amdt. 25–121, 72 FR 44669, Aug. 8, (3) Flight tests of the airplane or its 2007; Amdt. 25–129, 74 FR 38339, Aug. 3, 2009] components in measured simulated icing conditions. § 25.1420 Supercooled large drop icing (c) Caution information, such as an conditions. amber caution light or equivalent, (a) If certification for flight in icing must be provided to alert the conditions is sought, in addition to the flightcrew when the anti-ice or de-ice requirements of § 25.1419, an airplane system is not functioning normally. with a maximum takeoff weight less (d) For turbine engine powered air- than 60,000 pounds or with reversible planes, the ice protection provisions of flight controls must be capable of oper- this section are considered to be appli- ating in accordance with paragraphs cable primarily to the airframe. For (a)(1), (2), or (3), of this section. the powerplant installation, certain ad- (1) Operating safely after encoun- ditional provisions of subpart E of this tering the icing conditions defined in part may be found applicable. Appendix O of this part: (e) One of the following methods of (i) The airplane must have a means icing detection and activation of the to detect that it is operating in Appen- airframe ice protection system must be dix O icing conditions; and provided: (ii) Following detection of Appendix (1) A primary ice detection system O icing conditions, the airplane must that automatically activates or alerts be capable of operating safely while the flightcrew to activate the airframe exiting all icing conditions. ice protection system; (2) Operating safely in a portion of (2) A definition of visual cues for rec- the icing conditions defined in Appen- ognition of the first sign of ice accre- dix O of this part as selected by the ap- tion on a specified surface combined plicant: with an advisory ice detection system (i) The airplane must have a means that alerts the flightcrew to activate to detect that it is operating in condi- the airframe ice protection system; or tions that exceed the selected portion (3) Identification of conditions con- of Appendix O icing conditions; and ducive to airframe icing as defined by (ii) Following detection, the airplane an appropriate static or total air tem- must be capable of operating safely perature and visible moisture for use while exiting all icing conditions. by the flightcrew to activate the air- (3) Operating safely in the icing con- frame ice protection system. ditions defined in Appendix O of this (f) Unless the applicant shows that part. the airframe ice protection system (b) To establish that the airplane can need not be operated during specific operate safely as required in paragraph

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(a) of this section, an applicant must § 25.1421 Megaphones. show through analysis that the ice pro- If a megaphone is installed, a re- tection for the various components of straining means must be provided that the airplane is adequate, taking into is capable of restraining the mega- account the various airplane oper- phone when it is subjected to the ulti- ational configurations. To verify the mate inertia forces specified in analysis, one, or more as found nec- § 25.561(b)(3). essary, of the following methods must be used: [Amdt. 25–41, 42 FR 36970, July 18, 1977] (1) Laboratory dry air or simulated icing tests, or a combination of both, of § 25.1423 Public address system. the components or models of the com- A public address system required by ponents. this chapter must— (2) Laboratory dry air or simulated (a) Be powerable when the aircraft is icing tests, or a combination of both, of in flight or stopped on the ground, models of the airplane. after the shutdown or failure of all en- (3) Flight tests of the airplane or its gines and auxiliary power units, or the components in simulated icing condi- disconnection or failure of all power tions, measured as necessary to sup- sources dependent on their continued port the analysis. operation, for— (4) Flight tests of the airplane with (1) A time duration of at least 10 min- simulated ice shapes. utes, including an aggregate time dura- (5) Flight tests of the airplane in nat- tion of at least 5 minutes of announce- ural icing conditions, measured as nec- ments made by flight and cabin crew- essary to support the analysis. members, considering all other loads (c) For an airplane certified in ac- which may remain powered by the cordance with paragraph (a)(2) or (3) of same source when all other power this section, the requirements of sources are inoperative; and § 25.1419(e), (f), (g), and (h) must be met (2) An additional time duration in its for the icing conditions defined in Ap- standby state appropriate or required pendix O of this part in which the air- for any other loads that are powered by plane is certified to operate. the same source and that are essential (d) For the purposes of this section, to safety of flight or required during emergency conditions. the following definitions apply: (b) Be capable of operation within 3 (1) Reversible Flight Controls. Flight seconds from the time a microphone is controls in the normal operating con- removed from its stowage. figuration that have force or motion originating at the airplane’s control (c) Be intelligible at all passenger surface (for example, through aero- seats, lavatories, and flight attendant dynamic loads, static imbalance, or seats and work stations. trim or servo tab inputs) that is trans- (d) Be designed so that no unused, mitted back to flight deck controls. unstowed microphone will render the This term refers to flight deck controls system inoperative. connected to the pitch, roll, or yaw (e) Be capable of functioning inde- control surfaces by direct mechanical pendently of any required crewmember linkages, cables, or push-pull rods in interphone system. such a way that pilot effort produces (f) Be accessible for immediate use motion or force about the hinge line. from each of two flight crewmember (2) Simulated Icing Test. Testing con- stations in the pilot compartment. ducted in simulated icing conditions, (g) For each required floor-level pas- such as in an icing tunnel or behind an senger emergency exit which has an ad- icing tanker. jacent flight attendant seat, have a microphone which is readily accessible (3) Simulated Ice Shape. Ice shape fab- to the seated flight attendant, except ricated from wood, epoxy, or other ma- that one microphone may serve more terials by any construction technique. than one exit, provided the proximity [Amdt. 25–140, 79 FR 65528, Nov. 4, 2014] of the exits allows unassisted verbal

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communication between seated flight Element Proof Ultimate attendants. (xDOP) (xDOP) [Doc. No. 26003, 58 FR 45229, Aug. 26, 1993, as 1. Tubes and fittings...... 1.5 3.0 amended by Amdt. 25–115, 69 FR 40527, July 2, 2. Pressure vessels containing gas: 2004] High pressure (e.g., accumula- tors) ...... 3.0 4.0 Low pressure (e.g., reservoirs) .. 1.5 3.0 MISCELLANEOUS EQUIPMENT 3. Hoses ...... 2.0 4.0 4. All other elements ...... 1.5 2.0 § 25.1431 Electronic equipment. (a) In showing compliance with (2) Withstand, without deformation § 25.1309 (a) and (b) with respect to that would prevent it from performing radio and electronic equipment and its intended function, the design oper- their installations, critical environ- ating pressure in combination with mental conditions must be considered. limit structural loads that may be im- (b) Radio and electronic equipment posed; must be supplied with power under the (3) Withstand, without rupture, the requirements of § 25.1355(c). design operating pressure multiplied by (c) Radio and electronic equipment, a factor of 1.5 in combination with ulti- controls, and wiring must be installed mate structural load that can reason- so that operation of any one unit or ably occur simultaneously; system of units will not adversely af- (4) Withstand the fatigue effects of fect the simultaneous operation of any all cyclic pressures, including tran- other radio or electronic unit, or sys- sients, and associated externally in- tem of units, required by this chapter. duced loads, taking into account the (d) Electronic equipment must be de- consequences of element failure; and signed and installed such that it does (5) Perform as intended under all en- not cause essential loads to become in- vironmental conditions for which the operative as a result of electrical airplane is certificated. power supply transients or transients (b) System design. Each hydraulic sys- from other causes. tem must: (1) Have means located at a [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as flightcrew station to indicate appro- amended by Amdt. 25–113, 69 FR 12530, Mar. priate system parameters, if 16, 2004] (i) It performs a function necessary § 25.1433 Vacuum systems. for continued safe flight and landing; or There must be means, in addition to (ii) In the event of hydraulic system the normal pressure relief, to auto- malfunction, corrective action by the matically relieve the pressure in the crew to ensure continued safe flight discharge lines from the vacuum air and landing is necessary; pump when the delivery temperature of (2) Have means to ensure that system the air becomes unsafe. pressures, including transient pres- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as sures and pressures from fluid volu- amended by Amdt. 25–72, 55 FR 29785, July 20, metric changes in elements that are 1990] likely to remain closed long enough for such changes to occur, are within the § 25.1435 Hydraulic systems. design capabilities of each element, (a) Element design. Each element of such that they meet the requirements the hydraulic system must be designed defined in § 25.1435(a)(1) through (a)(5); to: (3) Have means to minimize the re- (1) Withstand the proof pressure lease of harmful or hazardous con- without permanent deformation that centrations of hydraulic fluid or vapors would prevent it from performing its into the crew and passenger compart- intended functions, and the ultimate ments during flight; pressure without rupture. The proof (4) Meet the applicable requirements and ultimate pressures are defined in of §§ 25.863, 25.1183, 25.1185, and 25.1189 if terms of the design operating pressure a flammable hydraulic fluid is used; (DOP) as follows: and

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(5) Be designed to use any suitable (c) An analysis, or a combination of hydraulic fluid specified by the air- analysis and test, may be substituted plane manufacturer, which must be for any test required by paragraph (a) identified by appropriate markings as or (b) of this section if the Adminis- required by § 25.1541. trator finds it equivalent to the re- (c) Tests. Tests must be conducted on quired test. the hydraulic system(s), and/or sub- [Amdt. 25–41, 42 FR 36971, July 18, 1977] system(s) and elements, except that analysis may be used in place of or to § 25.1439 Protective breathing equip- supplement testing, where the analysis ment. is shown to be reliable and appropriate. (a) Fixed (stationary, or built in) pro- All internal and external influences tective breathing equipment must be must be taken into account to an ex- installed for the use of the flightcrew, tent necessary to evaluate their ef- and at least one portable protective fects, and to assure reliable system and breathing equipment shall be located element functioning and integration. at or near the flight deck for use by a Failure or unacceptable deficiency of flight crewmember. In addition, port- an element or system must be cor- able protective breathing equipment rected and be sufficiently retested, must be installed for the use of appro- where necessary. priate crewmembers for fighting fires (1) The system(s), subsystem(s), or in compartments accessible in flight element(s) must be subjected to per- other than the flight deck. This in- formance, fatigue, and endurance tests cludes isolated compartments and representative of airplane ground and upper and lower lobe galleys, in which flight operations. crewmember occupancy is permitted (2) The complete system must be during flight. Equipment must be in- tested to determine proper functional stalled for the maximum number of performance and relation to the other crewmembers expected to be in the systems, including simulation of rel- area during any operation. evant failure conditions, and to sup- (b) For protective breathing equip- port or validate element design. ment required by paragraph (a) of this (3) The complete hydraulic system(s) section or by the applicable Operating must be functionally tested on the air- Regulations: plane in normal operation over the (1) The equipment must be designed range of motion of all associated user to protect the appropriate crewmember systems. The test must be conducted at from smoke, carbon dioxide, and other the system relief pressure or 1.25 times harmful gases while on flight deck the DOP if a system pressure relief de- duty or while combating fires. vice is not part of the system design. (2) The equipment must include— Clearances between hydraulic system (i) Masks covering the eyes, nose and elements and other systems or struc- mouth, or tural elements must remain adequate (ii) Masks covering the nose and and there must be no detrimental ef- mouth, plus accessory equipment to fects. cover the eyes. (3) Equipment, including portable [Doc. No. 28617, 66 FR 27402, May 16, 2001] equipment, must allow communication with other crewmembers while in use. § 25.1438 Pressurization and pneu- Equipment available at flightcrew as- matic systems. signed duty stations must also enable (a) Pressurization system elements the flightcrew to use radio equipment. must be burst pressure tested to 2.0 (4) The part of the equipment pro- times, and proof pressure tested to 1.5 tecting the eyes shall not cause any ap- times, the maximum normal operating preciable adverse effect on vision and pressure. must allow corrective glasses to be (b) Pneumatic system elements must worn. be burst pressure tested to 3.0 times, (5) The equipment must supply pro- and proof pressure tested to 1.5 times, tective oxygen of 15 minutes duration the maximum normal operating pres- per crewmember at a pressure altitude sure. of 8,000 feet with a respiratory minute

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volume of 30 liters per minute BTPD. spiration, a mean tracheal oxygen par- The equipment and system must be de- tial pressure of 149 mm. Hg. when signed to prevent any inward leakage breathing 15 liters per minute, BTPS, to the inside of the device and prevent and with a maximum tidal volume of any outward leakage causing signifi- 700 cc. with a constant time interval cant increase in the oxygen content of between respirations. the local ambient atmosphere. If a de- (b) If demand equipment is installed mand oxygen system is used, a supply for use by flight crewmembers, the of 300 liters of free oxygen at 70 °F. and minimum mass flow of supplemental 760 mm. Hg. pressure is considered to oxygen required for each crewmember be of 15-minute duration at the pre- may not be less than the flow required scribed altitude and minute volume. If to maintain, during inspiration, a a continuous flow open circuit protec- mean tracheal oxygen partial pressure tive breathing system is used, a flow of 122 mm. Hg., up to and including a rate of 60 liters per minute at 8,000 feet cabin pressure altitude of 35,000 feet, (45 liters per minute at sea level) and a and 95 percent oxygen between cabin supply of 600 liters of free oxygen at 70 pressure altitudes of 35,000 and 40,000 ° F. and 760 mm. Hg. pressure is consid- feet, when breathing 20 liters per ered to be of 15-minute duration at the minute BTPS. In addition, there must prescribed altitude and minute volume. be means to allow the crew to use undi- Continuous flow systems must not in- luted oxygen at their discretion. crease the ambient oxygen content of (c) For passengers and cabin attend- the local atmosphere above that of de- ants, the minimum mass flow of sup- mand systems. BTPD refers to body plemental oxygen required for each temperature conditions (that is, 37 °C., at ambient pressure, dry). person at various cabin pressure alti- (6) The equipment must meet the re- tudes may not be less than the flow requirements of § 25.1441. quired to maintain, during inspiration and while using the oxygen equipment [Doc. No. FAA–2002–13859, 69 FR 40528, July 2, (including masks) provided, the fol- 2004] lowing mean tracheal oxygen partial pressures: § 25.1441 Oxygen equipment and supply. (1) At cabin pressure altitudes above 10,000 feet up to and including 18,500 (a) If certification with supplemental feet, a mean tracheal oxygen partial oxygen equipment is requested, the pressure of 100 mm. Hg. when breathing equipment must meet the requirements 15 liters per minute, BTPS, and with a of this section and §§ 25.1443 through tidal volume of 700 cc. with a constant 25.1453. (b) The oxygen system must be free time interval between respirations. from hazards in itself, in its method of (2) At cabin pressure altitudes above operation, and in its effect upon other 18,500 feet up to and including 40,000 components. feet, a mean tracheal oxygen partial (c) There must be a means to allow pressure of 83.8 mm. Hg. when breath- the crew to readily determine, during ing 30 liters per minute, BTPS, and flight, the quantity of oxygen available with a tidal volume of 1,100 cc. with a in each source of supply. constant time interval between res- (d) The oxygen flow rate and the oxy- pirations. gen equipment for airplanes for which (d) If first-aid oxygen equipment is certification for operation above 40,000 installed, the minimum mass flow of feet is requested must be approved. oxygen to each user may not be less than four liters per minute, STPD. § 25.1443 Minimum mass flow of sup- However, there may be a means to de- plemental oxygen. crease this flow to not less than two li- (a) If continuous flow equipment is ters per minute, STPD, at any cabin al- installed for use by flight crew- titude. The quantity of oxygen re- members, the minimum mass flow of quired is based upon an average flow supplemental oxygen required for each rate of three liters per minute per per- crewmember may not be less than the son for whom first-aid oxygen is re- flow required to maintain, during in- quired.

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(e) If portable oxygen equipment is least two oxygen dispensing units con- installed for use by crewmembers, the nected to oxygen terminals in each lav- minimum mass flow of supplemental atory. The total number of dispensing oxygen is the same as specified in para- units and outlets in the cabin must ex- graph (a) or (b) of this section, which- ceed the number of seats by at least 10 ever is applicable. percent. The extra units must be as uniformly distributed throughout the § 25.1445 Equipment standards for the cabin as practicable. If certification for oxygen distributing system. operation above 30,000 feet is requested, (a) When oxygen is supplied to both the dispensing units providing the re- crew and passengers, the distribution quired oxygen flow must be automati- system must be designed for either— cally presented to the occupants before (1) A source of supply for the flight the cabin pressure altitude exceeds crew on duty and a separate source for 15,000 feet. The crew must be provided the passengers and other crewmembers; with a manual means of making the or dispensing units immediately available (2) A common source of supply with in the event of failure of the automatic means to separately reserve the min- system. imum supply required by the flight (2) Each flight crewmember on flight crew on duty. deck duty must be provided with a (b) Portable walk-around oxygen quick-donning type oxygen dispensing units of the continuous flow, diluter- unit connected to an oxygen supply demand, and straight demand kinds terminal. This dispensing unit must be may be used to meet the crew or pas- immediately available to the flight senger breathing requirements. crewmember when seated at his station, and installed so that it: § 25.1447 Equipment standards for ox- (i) Can be placed on the face from its ygen dispensing units. ready position, properly secured, If oxygen dispensing units are in- sealed, and supplying oxygen upon de- stalled, the following apply: mand, with one hand, within five sec- (a) There must be an individual dis- onds and without disturbing eyeglasses pensing unit for each occupant for or causing delay in proceeding with whom supplemental oxygen is to be emergency duties; and supplied. Units must be designed to (ii) Allows, while in place, the per- cover the nose and mouth and must be formance of normal communication equipped with a suitable means to re- functions. tain the unit in position on the face. (3) The oxygen dispensing equipment Flight crew masks for supplemental for the flight crewmembers must be: oxygen must have provisions for the (i) The diluter demand or pressure de- use of communication equipment. mand (pressure demand mask with a (b) If certification for operation up to diluter demand pressure breathing reg- and including 25,000 feet is requested, ulator) type, or other approved oxygen an oxygen supply terminal and unit of equipment shown to provide the same oxygen dispensing equipment for the degree of protection, for airplanes to be immediate use of oxygen by each crew- operated above 25,000 feet. member must be within easy reach of (ii) The pressure demand (pressure that crewmember. For any other occu- demand mask with a diluter demand pants, the supply terminals and dis- pressure breathing regulator) type with pensing equipment must be located to mask-mounted regulator, or other ap- allow the use of oxygen as required by proved oxygen equipment shown to the operating rules in this chapter. provide the same degree of protection, (c) If certification for operation for airplanes operated at altitudes above 25,000 feet is requested, there where decompressions that are not ex- must be oxygen dispensing equipment tremely improbable may expose the meeting the following requirements: flightcrew to cabin pressure altitudes (1) There must be an oxygen dis- in excess of 34,000 feet. pensing unit connected to oxygen sup- (4) Portable oxygen equipment must ply terminals immediately available to be immediately available for each each occupant, wherever seated, and at cabin attendant. The portable oxygen

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equipment must have the oxygen dis- § 25.1453 Protection of oxygen equip- pensing unit connected to the portable ment from rupture. oxygen supply. Oxygen pressure tanks, and lines be- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tween tanks and the shutoff means, amended by Amdt. 25–41, 42 FR 36971, July 18, must be— 1977; Amdt. 25–87, 61 FR 28696, June 5, 1996; (a) Protected from unsafe tempera- Amdt. 25–116, 69 FR 62789, Oct. 27, 2004] tures; and (b) Located where the probability and § 25.1449 Means for determining use of hazards of rupture in a crash landing oxygen. are minimized. There must be a means to allow the crew to determine whether oxygen is § 25.1455 Draining of fluids subject to being delivered to the dispensing equip- freezing. ment. If fluids subject to freezing may be drained overboard in flight or during § 25.1450 Chemical oxygen generators. ground operation, the drains must be (a) For the purpose of this section, a designed and located to prevent the chemical oxygen generator is defined formation of hazardous quantities of as a device which produces oxygen by ice on the airplane as a result of the drainage. chemical reaction. (b) Each chemical oxygen generator [Amdt. 25–23, 35 FR 5680, Apr. 8, 1970] must be designed and installed in accordance with the following require- § 25.1457 Cockpit voice recorders. ments: (a) Each cockpit voice recorder re- (1) Surface temperature developed by quired by the operating rules of this the generator during operation may chapter must be approved and must be not create a hazard to the airplane or installed so that it will record the fol- to its occupants. lowing: (2) Means must be provided to relieve (1) Voice communications trans- any internal pressure that may be haz- mitted from or received in the airplane ardous. by radio. (3) Except as provided in SFAR 109, (2) Voice communications of flight each chemical oxygen generator instal- crewmembers on the flight deck. lation must meet the requirements of (3) Voice communications of flight § 25.795(d). crewmembers on the flight deck, using (c) In addition to meeting the re- the airplane’s interphone system. quirements in paragraph (b) of this sec- (4) Voice or audio signals identifying tion, each portable chemical oxygen navigation or approach aids introduced into a headset or speaker. generator that is capable of sustained operation by successive replacement of (5) Voice communications of flight crewmembers using the passenger loud- a generator element must be placarded speaker system, if there is such a sys- to show— tem and if the fourth channel is avail- (1) The rate of oxygen flow, in liters able in accordance with the require- per minute; ments of paragraph (c)(4)(ii) of this sec- (2) The duration of oxygen flow, in tion. minutes, for the replaceable generator (6) If datalink communication equip- element; and ment is installed, all datalink commu- (3) A warning that the replaceable nications, using an approved data mes- generator element may be hot, unless sage set. Datalink messages must be the element construction is such that recorded as the output signal from the the surface temperature cannot exceed communications unit that translates 100 degrees F. the signal into usable data. (b) The recording requirements of [Amdt. 25–41, 42 FR 36971, July 18, 1977, as paragraph (a)(2) of this section must be amended at 79 FR 13519, Mar. 11, 2014] met by installing a cockpit-mounted area microphone, located in the best

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position for recording voice commu- (1)(i) It receives its electrical power nications originating at the first and from the bus that provides the max- second pilot stations and voice commu- imum reliability for operation of the nications of other crewmembers on the cockpit voice recorder without jeopard- flight deck when directed to those sta- izing service to essential or emergency tions. The microphone must be so lo- loads. cated and, if necessary, the pre- (ii) It remains powered for as long as amplifiers and filters of the recorder possible without jeopardizing emer- must be so adjusted or supplemented, gency operation of the airplane. that the intelligibility of the recorded (2) There is an automatic means to communications is as high as prac- simultaneously stop the recorder and ticable when recorded under flight prevent each erasure feature from func- cockpit noise conditions and played tioning, within 10 minutes after crash back. Repeated aural or visual play- impact; back of the record may be used in eval- (3) There is an aural or visual means uating intelligibility. for preflight checking of the recorder (c) Each cockpit voice recorder must for proper operation; be installed so that the part of the (4) Any single electrical failure exter- communication or audio signals speci- nal to the recorder does not disable fied in paragraph (a) of this section ob- both the cockpit voice recorder and the tained from each of the following flight data recorder; sources is recorded on a separate chan- (5) It has an independent power nel: source— (1) For the first channel, from each ± boom, mask, or hand-held microphone, (i) That provides 10 1 minutes of headset, or speaker used at the first electrical power to operate both the pilot station. cockpit voice recorder and cockpit- (2) For the second channel from each mounted area microphone; boom, mask, or hand-held microphone, (ii) That is located as close as prac- headset, or speaker used at the second ticable to the cockpit voice recorder; pilot station. and (3) For the third channel—from the (iii) To which the cockpit voice re- cockpit-mounted area microphone. corder and cockpit-mounted area (4) For the fourth channel, from— microphone are switched automati- (i) Each boom, mask, or hand-held cally in the event that all other power microphone, headset, or speaker used to the cockpit voice recorder is inter- at the station for the third and fourth rupted either by normal shutdown or crew members; or by any other loss of power to the elec- (ii) If the stations specified in para- trical power bus; and graph (c)(4)(i) of this section are not re- (6) It is in a separate container from quired or if the signal at such a station the flight data recorder when both are is picked up by another channel, each required. If used to comply with only microphone on the flight deck that is the cockpit voice recorder require- used with the passenger loudspeaker ments, a combination unit may be in- system, if its signals are not picked up stalled. by another channel. (e) The recorder container must be (5) As far as is practicable all sounds located and mounted to minimize the received by the microphone listed in probability of rupture of the container paragraphs (c)(1), (2), and (4) of this as a result of crash impact and con- section must be recorded without sequent heat damage to the recorder interruption irrespective of the posi- from fire. tion of the interphone-transmitter key (1) Except as provided in paragraph switch. The design shall ensure that (e)(2) of this section, the recorder con- sidetone for the flight crew is produced tainer must be located as far aft as only when the interphone, public ad- practicable, but need not be outside of dress system, or radio transmitters are the pressurized compartment, and may in use. not be located where aft-mounted en- (d) Each cockpit voice recorder must gines may crush the container during be installed so that— impact.

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(2) If two separate combination dig- (ii) It remains powered for as long as ital flight data recorder and cockpit possible without jeopardizing emer- voice recorder units are installed in- gency operation of the airplane. stead of one cockpit voice recorder and (4) There is an aural or visual means one digital flight data recorder, the for preflight checking of the recorder combination unit that is installed to for proper recording of data in the stor- comply with the cockpit voice recorder age medium; requirements may be located near the (5) Except for recorders powered sole- cockpit. ly by the engine-driven electrical gen- (f) If the cockpit voice recorder has a erator system, there is an automatic bulk erasure device, the installation means to simultaneously stop a re- must be designed to minimize the prob- corder that has a data erasure feature ability of inadvertent operation and ac- and prevent each erasure feature from tuation of the device during crash im- functioning, within 10 minutes after pact. crash impact; (g) Each recorder container must— (6) There is a means to record data (1) Be either bright orange or bright from which the time of each radio yellow; transmission either to or from ATC can (2) Have reflective tape affixed to its be determined; external surface to facilitate its loca- (7) Any single electrical failure exter- tion under water; and nal to the recorder does not disable (3) Have an underwater locating de- both the cockpit voice recorder and the vice, when required by the operating flight data recorder; and rules of this chapter, on or adjacent to (8) It is in a separate container from the container which is secured in such the cockpit voice recorder when both manner that they are not likely to be are required. If used to comply with separated during crash impact. only the flight data recorder requirements, a combination unit may be in- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–2, 30 FR 3932, Mar. 26, stalled. If a combination unit is in- 1965; Amdt. 25–16, 32 FR 13914, Oct. 6, 1967; stalled as a cockpit voice recorder to Amdt. 25–41, 42 FR 36971, July 18, 1977; Amdt. comply with § 25.1457(e)(2), a combina- 25–65, 53 FR 26143, July 11, 1988; Amdt. 25–124, tion unit must be used to comply with 73 FR 12563, Mar. 7, 2008; 74 FR 32800, July 9, this flight data recorder requirement. 2009] (b) Each nonejectable record container must be located and mounted so § 25.1459 Flight data recorders. as to minimize the probability of con- (a) Each flight recorder required by tainer rupture resulting from crash im- the operating rules of this chapter pact and subsequent damage to the must be installed so that— record from fire. In meeting this re- (1) It is supplied with airspeed, alti- quirement the record container must tude, and directional data obtained be located as far aft as practicable, but from sources that meet the accuracy need not be aft of the pressurized com- requirements of §§ 25.1323, 25.1325, and partment, and may not be where aft- 25.1327, as appropriate; mounted engines may crush the con- (2) The vertical acceleration sensor is tainer upon impact. rigidly attached, and located longitu- (c) A correlation must be established dinally either within the approved cen- between the flight recorder readings of ter of gravity limits of the airplane, or airspeed, altitude, and heading and the at a distance forward or aft of these corresponding readings (taking into ac- limits that does not exceed 25 percent count correction factors) of the first pi- of the airplane’s mean aerodynamic lot’s instruments. The correlation chord; must cover the airspeed range over (3)(i) It receives its electrical power which the airplane is to be operated, from the bus that provides the max- the range of altitude to which the air- imum reliability for operation of the plane is limited, and 360 degrees of flight data recorder without jeopard- heading. Correlation may be estab- izing service to essential or emergency lished on the ground as appropriate. loads. (d) Each recorder container must—

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(1) Be either bright orange or bright Subpart G—Operating Limitations yellow; and Information (2) Have reflective tape affixed to its external surface to facilitate its loca- § 25.1501 General. tion under water; and (a) Each operating limitation speci- (3) Have an underwater locating de- fied in §§ 25.1503 through 25.1533 and vice, when required by the operating other limitations and information nec- rules of this chapter, on or adjacent to essary for safe operation must be es- the container which is secured in such tablished. a manner that they are not likely to be (b) The operating limitations and separated during crash impact. other information necessary for safe (e) Any novel or unique design or operation must be made available to operational characteristics of the air- the crewmembers as prescribed in craft shall be evaluated to determine if §§ 25.1541 through 25.1587. any dedicated parameters must be recorded on flight recorders in addition [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978] to or in place of existing requirements. OPERATING LIMITATIONS [Amdt. 25–8, 31 FR 127, Jan. 6, 1966, as amended by Amdt. 25–25, 35 FR 13192, Aug. 19, 1970; § 25.1503 Airspeed limitations: general. Amdt. 25–37, 40 FR 2577, Jan. 14, 1975; Amdt. 25–41, 42 FR 36971, July 18, 1977; Amdt. 25–65, When airspeed limitations are a func- 53 FR 26144, July 11, 1988; Amdt. 25–124, 73 FR tion of weight, weight distribution, al- 12563, Mar. 7, 2008; 74 FR 32800, July 9, 2009] titude, or Mach number, limitations corresponding to each critical com- § 25.1461 Equipment containing high bination of these factors must be estab- energy rotors. lished. (a) Equipment containing high energy rotors must meet paragraph (b), § 25.1505 Maximum operating limit speed. (c), or (d) of this section. (b) High energy rotors contained in The maximum operating limit speed equipment must be able to withstand (VMO/MMO airspeed or Mach Number, damage caused by malfunctions, vibra- whichever is critical at a particular al- tion, abnormal speeds, and abnormal titude) is a speed that may not be de- temperatures. In addition— liberately exceeded in any regime of (1) Auxiliary rotor cases must be able flight (climb, cruise, or descent), unless to contain damage caused by the fail- a higher speed is authorized for flight ure of high energy rotor blades; and test or pilot training operations. VMO/ (2) Equipment control devices, sys- MMO must be established so that it is tems, and instrumentation must rea- not greater than the design cruising sonably ensure that no operating limi- speed VC and so that it is sufficiently tations affecting the integrity of high below VD/MD or VDF/MDF, to make it energy rotors will be exceeded in serv- highly improbable that the latter ice. speeds will be inadvertently exceeded (c) It must be shown by test that in operations. The speed margin be- equipment containing high energy ro- tween VMO/MMO and VD/MD or VDFM/DF tors can contain any failure of a high may not be less than that determined energy rotor that occurs at the highest under § 25.335(b) or found necessary dur- speed obtainable with the normal speed ing the flight tests conducted under control devices inoperative. § 25.253. (d) Equipment containing high en- [Amdt. 25–23, 35 FR 5680, Apr. 8, 1970] ergy rotors must be located where rotor failure will neither endanger the § 25.1507 Maneuvering speed. occupants nor adversely affect contin- The maneuvering speed must be es- ued safe flight. tablished so that it does not exceed the [Amdt. 25–41, 42 FR 36971, July 18, 1977] design maneuvering speed VA determined under § 25.335(c).

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§ 25.1511 Flap extended speed. to provide an optimum margin between The established flap extended speed low and high speed buffet boundaries. VFE must be established so that it does [Amdt. 25–141, 79 FR 73469, Dec. 11, 2014] not exceed the design flap speed VF chosen under §§ 25.335(e) and 25.345, for § 25.1519 Weight, center of gravity, and the corresponding flap positions and weight distribution. engine powers. The airplane weight, center of gravity, and weight distribution limita- § 25.1513 Minimum control speed. tions determined under §§ 25.23 through

The minimum control speed VMC de- 25.27 must be established as operating termined under § 25.149 must be estab- limitations. lished as an operating limitation. § 25.1521 Powerplant limitations. § 25.1515 Landing gear speeds. (a) General. The powerplant limita- (a) The established landing gear oper- tions prescribed in this section must be ating speed or speeds, VLO, may not ex- established so that they do not exceed ceed the speed at which it is safe both the corresponding limits for which the to extend and to retract the landing engines or propellers are type certifi- gear, as determined under § 25.729 or by cated and do not exceed the values on flight characteristics. If the extension which compliance with any other re- speed is not the same as the retraction quirement of this part is based. speed, the two speeds must be des- (b) Reciprocating engine installations. ignated as VLO(EXT) and VLO(RET), respec- Operating limitations relating to the tively. following must be established for recip- (b) The established landing gear ex- rocating engine installations: tended speed VLE may not exceed the (1) Horsepower or torque, r.p.m., speed at which it is safe to fly with the manifold pressure, and time at critical landing gear secured in the fully ex- pressure altitude and sea level pressure tended position, and that determined altitude for— under § 25.729. (i) Maximum continuous power (relating to unsupercharged operation or [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as to operation in each supercharger mode amended by Amdt. 25–38, 41 FR 55468, Dec. 20, 1976] as applicable); and (ii) Takeoff power (relating to unsu- § 25.1516 Other speed limitations. percharged operation or to operation in Any other limitation associated with each supercharger mode as applicable). speed must be established. (2) Fuel grade or specification. (3) Cylinder head and oil tempera- [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] tures. (4) Any other parameter for which a § 25.1517 Rough air speed, VRA. limitation has been established as part (a) A rough air speed, VRA, for use as of the engine type certificate except the recommended turbulence penetra- that a limitation need not be estab- tion airspeed, and a rough air Mach lished for a parameter that cannot be number, MRA, for use as the rec- exceeded during normal operation due ommended turbulence penetration to the design of the installation or to Mach number, must be established. another established limitation. VRA/MRA must be sufficiently less than (c) Turbine engine installations. Oper- VMO/MMO to ensure that likely speed ating limitations relating to the fol- variation during rough air encounters lowing must be established for turbine will not cause the overspeed warning to engine installations: operate too frequently. (1) Horsepower, torque or thrust, (b) At altitudes where VMO is not lim- r.p.m., gas temperature, and time for— ited by Mach number, in the absence of (i) Maximum continuous power or a rational investigation substantiating thrust (relating to augmented or un- the use of other values, VRA must be augmented operation as applicable). less than VMO—35 KTAS. (ii) Takeoff power or thrust (relating (c) At altitudes where VMO is limited to augmented or unaugmented oper- by Mach number, MRA may be chosen ation as applicable). 365

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(2) Fuel designation or specification. certification and by the installed (3) Maximum time interval between equipment. engine run-ups from idle, run-up power setting and duration at power for § 25.1527 Ambient air temperature and ground operation in icing conditions, operating altitude. as defined in § 25.1093(b)(2). The extremes of the ambient air tem- (4) Any other parameter for which a perature and operating altitude for limitation has been established as part which operation is allowed, as limited of the engine type certificate except by flight, structural, powerplant, func- that a limitation need not be estab- tional, or equipment characteristics, lished for a parameter that cannot be must be established. exceeded during normal operation due to the design of the installation or to [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] another established limitation. § 25.1529 Instructions for Continued (d) Ambient temperature. An ambient Airworthiness. temperature limitation (including limitations for winterization installations, The applicant must prepare Instruc- if applicable) must be established as tions for Continued Airworthiness in the maximum ambient atmospheric accordance with appendix H to this temperature established in accordance part that are acceptable to the Admin- with § 25.1043(b). istrator. The instructions may be in- complete at type certification if a pro- [Amdt. 25–72, 55 FR 29786, July 20, 1990, as gram exists to ensure their completion amended by Amdt. 25–140, 79 FR 65528, Nov. 4, prior to delivery of the first airplane or 2014] issuance of a standard certificate of § 25.1522 Auxiliary power unit limita- airworthiness, whichever occurs later. tions. [Amdt. 25–54, 45 FR 60173, Sept. 11, 1980] If an auxiliary power unit is installed in the airplane, limitations established § 25.1531 Maneuvering flight load fac- for the auxiliary power unit, including tors. categories of operation, must be speci- Load factor limitations, not exceed- fied as operating limitations for the ing the positive limit load factors de- airplane. termined from the maneuvering dia- gram in § 25.333(b), must be established. [Amdt. 25–72, 55 FR 29786, July 20, 1990]

§ 25.1523 Minimum flight crew. § 25.1533 Additional operating limitations. The minimum flight crew must be es- (a) Additional operating limitations tablished so that it is sufficient for safe must be established as follows: operation, considering— (1) The maximum takeoff weights (a) The workload on individual crew- must be established as the weights at members; which compliance is shown with the (b) The accessibility and ease of oper- applicable provisions of this part (in- ation of necessary controls by the ap- cluding the takeoff climb provisions of propriate crewmember; and § 25.121(a) through (c), for altitudes and (c) The kind of operation authorized ambient temperatures). under § 25.1525. (2) The maximum landing weights The criteria used in making the deter- must be established as the weights at minations required by this section are which compliance is shown with the set forth in appendix D. applicable provisions of this part (in- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as cluding the landing and approach climb amended by Amdt. 25–3, 30 FR 6067, Apr. 29, provisions of §§ 25.119 and 25.121(d) for 1965] altitudes and ambient temperatures). (3) The minimum takeoff distances § 25.1525 Kinds of operation. must be established as the distances at The kinds of operation to which the which compliance is shown with the airplane is limited are established by applicable provisions of this part (in- the category in which it is eligible for cluding the provisions of §§ 25.109 and

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25.113, for weights, altitudes, tempera- (b) Each marking and placard pre- tures, wind components, runway sur- scribed in paragraph (a) of this sec- face conditions (dry and wet), and run- tion— way gradients) for smooth, hard-sur- (1) Must be displayed in a con- faced runways. Additionally, at the op- spicuous place; and tion of the applicant, wet runway take- (2) May not be easily erased, dis- off distances may be established for figured, or obscured. runway surfaces that have been grooved or treated with a porous fric- § 25.1543 Instrument markings: gen- tion course, and may be approved for eral. use on runways where such surfaces For each instrument— have been designed constructed, and (a) When markings are on the cover maintained in a manner acceptable to glass of the instrument, there must be the Administrator. means to maintain the correct align- (b) The extremes for variable factors ment of the glass cover with the face of (such as altitude, temperature, wind, the dial; and and runway gradients) are those at (b) Each instrument marking must which compliance with the applicable be clearly visible to the appropriate provisions of this part is shown. crewmember. (c) For airplanes certified in accord- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as ance with § 25.1420(a)(1) or (2), an oper- amended by Amdt. 25–72, 55 FR 29786, July 20, ating limitation must be established 1990] to: (1) Prohibit intentional flight, in- § 25.1545 Airspeed limitation informa- cluding takeoff and landing, into icing tion. conditions defined in Appendix O of The airspeed limitations required by this part for which the airplane has not § 25.1583 (a) must be easily read and un- been certified to safely operate; and derstood by the flight crew. (2) Require exiting all icing conditions if icing conditions defined in Ap- § 25.1547 Magnetic direction indicator. pendix O of this part are encountered (a) A placard meeting the require- for which the airplane has not been ments of this section must be installed certified to safely operate. on, or near, the magnetic direction indicator. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (b) The placard must show the cali- amended by Amdt. 25–38, 41 FR 55468, Dec. 20, bration of the instrument in level 1976; Amdt. 25–72, 55 FR 29786, July 20, 1990; Amdt. 25–92, 63 FR 8321, Feb. 18, 1998; Amdt. flight with the engines operating. 25–140, 79 FR 65528, Nov. 4, 2014] (c) The placard must state whether the calibration was made with radio re- § 25.1535 ETOPS approval. ceivers on or off. (d) Each calibration reading must be Except as provided in § 25.3, each ap- in terms of magnetic heading in not plicant seeking ETOPS type design ap- more than 45 degree increments. proval must comply with the provisions of Appendix K of this part. § 25.1549 Powerplant and auxiliary [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, power unit instruments. 2007] For each required powerplant and auxiliary power unit instrument, as ap- MARKINGS AND PLACARDS propriate to the type of instrument— (a) Each maximum and, if applicable, § 25.1541 General. minimum safe operating limit must be (a) The airplane must contain— marked with a red radial or a red line; (1) The specified markings and plac- (b) Each normal operating range ards; and must be marked with a green arc or (2) Any additional information, in- green line, not extending beyond the strument markings, and placards re- maximum and minimum safe limits; quired for the safe operation if there (c) Each takeoff and precautionary are unusual design, operating, or han- range must be marked with a yellow dling characteristics. arc or a yellow line; and

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(d) Each engine, auxiliary power § 25.1557 Miscellaneous markings and unit, or propeller speed range that is placards. restricted because of excessive vibra- (a) Baggage and cargo compartments tion stresses must be marked with red and ballast location. Each baggage and arcs or red lines. cargo compartment, and each ballast [Amdt. 25–40, 42 FR 15044, Mar. 17, 1977] location must have a placard stating any limitations on contents, including § 25.1551 Oil quantity indication. weight, that are necessary under the 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 [Amdt. 25–72, 55 FR 29786, July 20, 1990] have a loading limitation placard. (b) Powerplant fluid filler openings. § 25.1553 Fuel quantity indicator. The following apply: If the unusable fuel supply for any (1) Fuel filler openings must be tank exceeds one gallon, or five per- marked at or near the filler cover cent of the tank capacity, whichever is with— greater, a red arc must be marked on (i) The word ‘‘fuel’’; its indicator extending from the cali- (ii) For reciprocating engine powered brated zero reading to the lowest read- airplanes, the minimum fuel grade; ing obtainable in level flight. (iii) For turbine engine powered airplanes, the permissible fuel designa- § 25.1555 Control markings. tions; and (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. (2) If safe operation requires the use (d) Doors. Each door that must be of any tanks in a specific sequence, used in order to reach any required that sequence must be marked on, or emergency exit must have a suitable adjacent to, the selector for those placard stating that the door is to be tanks; and latched in the open position during (3) Each valve control for each engine takeoff and landing. must be marked to indicate the posi- [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as tion corresponding to each engine con- amended by Amdt. 25–32, 37 FR 3972, Feb. 24, trolled. 1972; Amdt. 25–38, 41 FR 55468, Dec. 20, 1976; (d) For accessory, auxiliary, and Amdt. 25–72, 55 FR 29786, July 20, 1990] emergency controls— (1) Each emergency control (includ- § 25.1561 Safety equipment. ing each fuel jettisoning and fluid shut- (a) Each safety equipment control to off must be colored red; and be operated by the crew in emergency, (2) Each visual indicator required by such as controls for automatic liferaft § 25.729(e) must be marked so that the releases, must be plainly marked as to pilot can determine at any time when its method of operation. the wheels are locked in either extreme (b) Each location, such as a locker or position, if retractable landing gear is compartment, that carries any fire ex- used. tinguishing, signaling, or other life

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saving equipment must be marked ac- § 25.1583 Operating limitations. cordingly. (a) Airspeed limitations. The following (c) Stowage provisions for required airspeed limitations and any other air- emergency equipment must be con- speed limitations necessary for safe op- spicuously marked to identify the con- eration must be furnished: tents and facilitate the easy removal of (1) The maximum operating limit the equipment. speed VMO/MMO and a statement that (d) Each liferaft must have obviously this speed limit may not be delib- marked operating instructions. erately exceeded in any regime of (e) Approved survival equipment flight (climb, cruise, or descent) unless must be marked for identification and a higher speed is authorized for flight method of operation. test or pilot training. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as (2) If an airspeed limitation is based amended by Amdt. 25–46, 43 FR 50598, Oct. 30, upon compressibility effects, a state- 1978] ment to this effect and information as to any symptoms, the probable behav- § 25.1563 Airspeed placard. ior of the airplane, and the rec- A placard showing the maximum air- ommended recovery procedures. speeds for flap extension for the take- (3) The maneuvering speed estab- off, approach, and landing positions lished under § 25.1507 and statements, must be installed in clear view of each as applicable to the particular design, pilot. explaining that: (i) Full application of pitch, roll, or AIRPLANE FLIGHT MANUAL yaw controls should be confined to speeds below the maneuvering speed; § 25.1581 General. and (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). ble noise standards of part 36 of this (6) The landing gear extended speed chapter. VLE, if greater than VLO, and a state- (b) Approved information. Each part of ment that this is the maximum speed the manual listed in §§ 25.1583 through at which the airplane can be safely 25.1587, that is appropriate to the air- flown with the landing gear extended. plane, must be furnished, verified, and (b) Powerplant limitations. The fol- approved, and must be segregated, lowing information must be furnished: identified, and clearly distinguished (1) Limitations required by § 25.1521 from each unapproved part of that and § 25.1522. manual. (2) Explanation of the limitations, (c) [Reserved] when appropriate. (d) Each Airplane Flight Manual (3) Information necessary for mark- must include a table of contents if the ing the instruments required by complexity of the manual indicates a §§ 25.1549 through 25.1553. need for it. (c) Weight and loading distribution. [Amdt. 25–42, 43 FR 2323, Jan. 16, 1978, as The weight and center of gravity limi- amended by Amdt. 25–72, 55 FR 29786, July 20, tations established under § 25.1519 must 1990] be furnished in the Airplane Flight

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Manual. All of the following informa- involving the use of special systems or tion, including the weight distribution the alternative use of regular systems; limitations established under § 25.1519, and must be presented either in the Air- (3) Emergency procedures for foresee- plane Flight Manual or in a separate able but unusual situations in which weight and balance control and loading immediate and precise action by the document that is incorporated by ref- crew may be expected to substantially erence in the Airplane Flight Manual: reduce the risk of catastrophe. (1) The condition of the airplane and (b) Information or procedures not di- the items included in the empty weight rectly related to airworthiness or not as defined in accordance with § 25.29. under the control of the crew, must not (2) Loading instructions necessary to be included, nor must any procedure ensure loading of the airplane within that is accepted as basic airmanship. the weight and center of gravity limits, (c) Information identifying each op- and to maintain the loading within erating condition in which the fuel sys- these limits in flight. tem independence prescribed in § 25.953 (3) If certification for more than one is necessary for safety must be fur- center of gravity range is requested, nished, together with instructions for the appropriate limitations, with replacing the fuel system in a configura- gard to weight and loading procedures, tion used to show compliance with that for each separate center of gravity section. range. (d) The buffet onset envelopes, deter- (d) Flight crew. The number and func- mined under § 25.251 must be furnished. tions of the minimum flight crew de- The buffet onset envelopes presented termined under § 25.1523 must be fur- may reflect the center of gravity at nished. which the airplane is normally loaded (e) Kinds of operation. The kinds of during cruise if corrections for the ef- operation approved under § 25.1525 must fect of different center of gravity loca- be furnished. tions are furnished. (e) Information must be furnished (f) Ambient air temperatures and oper- that indicates that when the fuel quan- ating altitudes. The extremes of the am- tity indicator reads ‘‘zero’’ in level bient air temperatures and operating flight, any fuel remaining in the fuel altitudes established under § 25.1527 tank cannot be used safely in flight. must be furnished. (f) Information on the total quantity (g) [Reserved] of usable fuel for each fuel tank must (h) Additional operating limitations. be furnished. The operating limitations established under § 25.1533 must be furnished. [Doc. No. 2000–8511, 66 FR 34024, June 26, 2001] (i) Maneuvering flight load factors. The positive maneuvering limit load fac- § 25.1587 Performance information. tors for which the structure is proven, (a) Each Airplane Flight Manual described in terms of accelerations, must contain information to permit must be furnished. conversion of the indicated temperature to free air temperature if other [Doc. No. 5066, 29 FR 1891, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55468, Dec. 20, than a free air temperature indicator is 1976; Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; used to comply with the requirements Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; Amdt. of § 25.1303(a)(1). 25–72, 55 FR 29787, July 20, 1990; Amdt. 25–105, (b) Each Airplane Flight Manual 66 FR 34024, June 26, 2001; 75 FR 49818, Aug. must contain the performance informa- 16, 2010] tion computed under the applicable provisions of this part (including § 25.1585 Operating procedures. §§ 25.115, 25.123, and 25.125 for the (a) Operating procedures must be fur- weights, altitudes, temperatures, wind nished for— components, and runway gradients, as (1) Normal procedures peculiar to the applicable) within the operational lim- particular type or model encountered its of the airplane, and must contain in connection with routine operations; the following: (2) Non-normal procedures for mal- (1) In each case, the conditions of function cases and failure conditions power, configuration, and speeds, and

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the procedures for handling the air- contactors, terminal blocks and circuit plane and any system having a signifi- breakers, and other circuit protection cant effect on the performance infor- devices. mation. (4) Connectors, including feed- (2) VSR determined in accordance through connectors. with § 25.103. (5) Connector accessories. (3) The following performance infor- (6) Electrical grounding and bonding mation (determined by extrapolation devices and their associated connec- and computed for the range of weights tions. between the maximum landing weight (7) Electrical splices. and the maximum takeoff weight): (8) Materials used to provide addi- (i) Climb in the landing configura- tional protection for wires, including tion. wire insulation, wire sleeving, and con- (ii) Climb in the approach configura- duits that have electrical termination tion. for the purpose of bonding. (iii) Landing distance. (9) Shields or braids. (4) Procedures established under (10) Clamps and other devices used to § 25.101(f) and (g) that are related to the route and support the wire bundle. limitations and information required (11) Cable tie devices. by § 25.1533 and by this paragraph (b) in (12) Labels or other means of identi- the form of guidance material, includ- fication. ing any relevant limitations or infor- (13) Pressure seals. mation. (14) EWIS components inside shelves, (5) An explanation of significant or panels, racks, junction boxes, distribu- unusual flight or ground handling char- tion panels, and back-planes of equip- acteristics of the airplane. ment 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: (1) Electrical equipment or avionics Subpart H—Electrical Wiring that are qualified to environmental Interconnection Systems (EWIS) conditions and testing procedures when those conditions and procedures are— SOURCE: Docket No. FAA–2004–18379, 72 FR (i) Appropriate for the intended func- 63406, Nov. 8, 2007, unless otherwise noted. tion and operating environment, and (ii) Acceptable to the FAA. § 25.1701 Definition. (2) Portable electrical devices that (a) As used in this chapter, electrical are not part of the type design of the wiring interconnection system (EWIS) airplane. This includes personal enter- means any wire, wiring device, or com- tainment devices and laptop combination of these, including termi- puters. nation devices, installed in any area of (3) Fiber optics. the airplane for the purpose of transmitting electrical energy, including § 25.1703 Function and installation: data and signals, between two or more EWIS. intended termination points. This in- (a) Each EWIS component installed cludes: in any area of the aircraft must: (1) Wires and cables. (1) Be of a kind and design appro- (2) Bus bars. priate to its intended function. (3) The termination point on elec- (2) Be installed according to limita- trical devices, including those on re- tions specified for the EWIS compo- lays, interrupters, switches, nents.

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(3) Perform the function for which it § 25.1707 System separation: EWIS. was intended without degrading the (a) Each EWIS must be designed and airworthiness of the airplane. installed with adequate physical sepa- (4) Be designed and installed in a way ration from other EWIS and airplane that will minimize mechanical strain. systems so that an EWIS component (b) Selection of wires must take into failure will not create a hazardous con- account known characteristics of the dition. Unless otherwise stated, for the wire in relation to each installation purposes of this section, adequate and application to minimize the risk of physical separation must be achieved wire damage, including any arc track- by separation distance or by a barrier ing phenomena. that provides protection equivalent to (c) The design and installation of the that separation distance. main power cables (including generator (b) Each EWIS must be designed and cables) in the fuselage must allow for a installed so that any electrical inter- reasonable degree of deformation and ference likely to be present in the air- stretching without failure. plane will not result in hazardous ef- (d) EWIS components located in fects upon the airplane or its systems. areas of known moisture accumulation (c) Wires and cables carrying heavy must be protected to minimize any current, and their associated EWIS hazardous effects due to moisture. components, must be designed and installed to ensure adequate physical § 25.1705 Systems and functions: EWIS. separation and electrical isolation so (a) EWIS associated with any system that damage to circuits associated required for type certification or by op- with essential functions will be mini- erating rules must be considered an in- mized under fault conditions. tegral part of that system and must be (d) Each EWIS associated with inde- considered in showing compliance with pendent airplane power sources or the applicable requirements for that power sources connected in combina- system. tion must be designed and installed to (b) For systems to which the fol- ensure adequate physical separation lowing rules apply, the components of and electrical isolation so that a fault EWIS associated with those systems in any one airplane power source EWIS must be considered an integral part of will not adversely affect any other that system or systems and must be independent power sources. In addition: considered in showing compliance with (1) Airplane independent electrical the applicable requirements for that power sources must not share a com- system. mon ground terminating location. (1) § 25.773(b)(2) Pilot compartment (2) Airplane system static grounds view. must not share a common ground ter- (2) § 25.981 Fuel tank ignition pre- minating location with any of the air- vention. plane’s independent electrical power (3) § 25.1165 Engine ignition systems. sources. (4) § 25.1310 Power source capacity (e) Except to the extent necessary to and distribution. provide electrical connection to the (5) § 25.1316 System lightning protec- fuel systems components, the EWIS tion. must be designed and installed with (6) § 25.1331(a)(2) Instruments using a adequate physical separation from fuel power supply. lines and other fuel system compo- (7) § 25.1351 General. nents, so that: (8) § 25.1355 Distribution system. (1) An EWIS component failure will (9) § 25.1360 Precautions against in- not create a hazardous condition. jury. (2) Any fuel leakage onto EWIS com- (10) § 25.1362 Electrical supplies for ponents will not create a hazardous emergency conditions. condition. (11) § 25.1365 Electrical appliances, (f) Except to the extent necessary to motors, and transformers. provide electrical connection to the (12) § 25.1431(c) and (d) Electronic hydraulic systems components, EWIS equipment. must be designed and installed with

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adequate physical separation from hy- those systems must be designed and in- draulic lines and other hydraulic sys- stalled with adequate physical separa- tem components, so that: tion. (1) An EWIS component failure will (l) Each EWIS must be designed and not create a hazardous condition. installed so there is adequate physical (2) Any hydraulic fluid leakage onto separation between it and other air- EWIS components will not create a craft components and aircraft struc- hazardous condition. ture, and so that the EWIS is protected (g) Except to the extent necessary to from sharp edges and corners, to mini- provide electrical connection to the ox- mize potential for abrasion/chafing, vi- ygen systems components, EWIS must bration damage, and other types of me- be designed and installed with ade- chanical damage. quate physical separation from oxygen lines and other oxygen system compo- § 25.1709 System safety: EWIS. nents, so that an EWIS component fail- Each EWIS must be designed and in- ure will not create a hazardous condi- stalled so that: tion. (a) Each catastrophic failure condi- (h) Except to the extent necessary to tion— provide electrical connection to the (1) Is extremely improbable; and water/waste systems components, (2) Does not result from a single fail- EWIS must be designed and installed ure. with adequate physical separation from 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 § 25.1711 Component identification: not create a hazardous condition. EWIS. (2) Any water/waste leakage onto 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 (1) Chafing, jamming, or other inter- operating rules, or as a result of the as- ference are prevented. sessment required by § 25.1709, EWIS (2) An EWIS component failure will components associated with those sys- not create a hazardous condition. tems must be specifically identified (3) Failure of any flight or other me- with component part number, function, chanical control systems cables or sys- and separation requirement for bun- tems components will not damage the dles. EWIS and create a hazardous condi- (1) The identification must be placed tion. along the wire, cable, or wire bundle at (j) EWIS must be designed and inappropriate intervals and in areas of stalled with adequate physical separa- the airplane where it is readily visible tion between the EWIS components to maintenance, repair, or alteration and heated equipment, hot air ducts, personnel. and lines, so that: (2) If an EWIS component cannot be (1) An EWIS component failure will marked physically, then other means not create a hazardous condition. of identification must be provided. (2) Any hot air leakage or heat gen- (c) The identifying markings re- erated onto EWIS components will not quired by paragraphs (a) and (b) of this create a hazardous condition. section must remain legible through- (k) For systems for which redun- out the expected service life of the dancy is required, by certification EWIS component. rules, by operating rules, or as a result (d) The means used for identifying of the assessment required by § 25.1709, each EWIS component as required by EWIS components associated with this section must not have an adverse

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effect on the performance of that com- § 25.1721 Protection of EWIS. ponent throughout its expected service (a) No cargo or baggage compartment life. may contain any EWIS whose damage (e) Identification for EWIS modifica- or failure may affect safe operation, tions to the type design must be con- unless the EWIS is protected so that: sistent with the identification scheme (1) It cannot be damaged by move- of the original type design. ment of cargo or baggage in the compartment. § 25.1713 Fire protection: EWIS. (2) Its breakage or failure will not (a) All EWIS components must meet create a fire hazard. the applicable fire and smoke protec- (b) EWIS must be designed and in- tion requirements of § 25.831(c) of this stalled to minimize damage and risk of part. damage to EWIS by movement of peo- (b) EWIS components that are lo- ple in the airplane during all phases of cated in designated fire zones and are flight, maintenance, and servicing. used during emergency procedures (c) EWIS must be designed and in- must be fire resistant. stalled to minimize damage and risk of (c) Insulation on electrical wire and damage to EWIS by items carried onto electrical cable, and materials used to the aircraft by passengers or cabin provide additional protection for the crew. wire and cable, installed in any area of the airplane, must be self-extin- § 25.1723 Flammable fluid fire protec- guishing when tested in accordance tion: EWIS. with the applicable portions of Appen- EWIS components located in each dix F, part I, of 14 CFR part 25. area where flammable fluid or vapors might escape by leakage of a fluid sys- § 25.1715 Electrical bonding and pro- tem must be considered a potential ig- tection against static electricity: nition source and must meet the re- EWIS. quirements of § 25.863. (a) EWIS components used for electrical bonding and protection against § 25.1725 Powerplants: EWIS. static electricity must meet the re- (a) EWIS associated with any power- quirements of § 25.899. plant must be designed and installed so (b) On airplanes having grounded that the failure of an EWIS component electrical systems, electrical bonding will not prevent the continued safe op- provided by EWIS components must eration of the remaining powerplants provide an electrical return path capa- or require immediate action by any ble of carrying both normal and fault crewmember for continued safe oper- currents without creating a shock haz- ation, in accordance with the require- ard or damage to the EWIS compo- ments of § 25.903(b). nents, other airplane system compo- (b) Design precautions must be taken nents, or airplane structure. to minimize hazards to the airplane due to EWIS damage in the event of a § 25.1717 Circuit protective devices: powerplant rotor failure or a fire origi- EWIS. nating within the powerplant that Electrical wires and cables must be burns through the powerplant case, in designed and installed so they are com- accordance with the requirements of patible with the circuit protection de- § 25.903(d)(1). vices required by § 25.1357, so that a fire or smoke hazard cannot be created § 25.1727 Flammable fluid shutoff under temporary or continuous fault means: EWIS. conditions. EWIS associated with each flammable fluid shutoff means and control § 25.1719 Accessibility provisions: must be fireproof or must be located EWIS. and protected so that any fire in a fire Access must be provided to allow in- zone will not affect operation of the spection and replacement of any EWIS flammable fluid shutoff means, in ac- component as necessary for continued cordance with the requirements of airworthiness. § 25.1189.

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§ 25.1729 Instructions for Continued must meet the requirements of Airworthiness: EWIS. § 25.1203. The applicant must prepare Instruc- tions for Continued Airworthiness ap- § 25.1733 Fire detector systems, general: EWIS. plicable to EWIS in accordance with Appendix H sections H25.4 and H25.5 to EWIS associated with any installed this part that are approved by the fire protection system, including those FAA. required by §§ 25.854 and 25.858, must be considered an integral part of the sys- § 25.1731 Powerplant and APU fire de- tem in showing compliance with the tector system: EWIS. applicable requirements for that sys- (a) EWIS that are part of each fire or tem. overheat detector system in a fire zone must be fire-resistant. Subpart I—Special Federal (b) No EWIS component of any fire or Aviation Regulations overheat detector system for any fire zone may pass through another fire zone, unless: SOURCE: Docket No. FAA–2011–0186, Amdt. (1) It is protected against the possi- 25–133, 76 FR 12555, Mar. 8, 2011, unless other- bility of false warnings resulting from wise noted. fires in zones through which it passes; § 25.1801 SFAR No. 111—Lavatory Oxy- or gen Systems. (2) Each zone involved is simultaneously protected by the same detector The requirements of § 121.1500 of this and extinguishing system. chapter also apply to this part. (c) EWIS that are part of each fire or overheat detector system in a fire zone

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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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(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 (b) In order to reduce the number of ice ac- Compliance With Subpart B. cretions to be considered when demonstrating compliance with the require- (a) Ice accretions—General. The most crit- ments of § 25.21(g), any of the ice accretions ical ice accretion in terms of airplane per- defined in paragraph (a) of this section may formance and handling qualities for each be used for any other flight phase if it is flight phase must be used to show compli- shown to be more critical than the specific ance with the applicable airplane perform- ice accretion defined for that flight phase. ance and handling requirements in icing con- Configuration differences and their effects ditions of subpart B of this part. Applicants on ice accretions must be taken into ac- must demonstrate that the full range of at- count. mospheric icing conditions specified in part I (c) The ice accretion that has the most ad- of this appendix have been considered, in- verse effect on handling qualities may be cluding the mean effective drop diameter, used for airplane performance tests provided liquid water content, and temperature appro- any difference in performance is conserv- priate to the flight conditions (for example, atively taken into account. configuration, speed, angle-of-attack, and al- (d) For both unprotected and protected titude). The ice accretions for each flight parts, the ice accretion for the takeoff phase phase are defined as follows: may be determined by calculation, assuming (1) Takeoff ice is the most critical ice accre- the takeoff maximum icing conditions de- tion on unprotected surfaces and any ice ac- fined in appendix C, and assuming that: cretion on the protected surfaces appropriate (1) Airfoils, control surfaces and, if appli- to normal ice protection system operation, cable, propellers are free from frost, snow, or occurring between the end of the takeoff dis- ice at the start of the takeoff; tance and 400 feet above the takeoff surface, (2) The ice accretion starts at the end of assuming accretion starts at the end of the the takeoff distance. takeoff distance in the takeoff maximum (3) The critical ratio of thrust/power-to- icing conditions defined in part I of this Ap- weight; pendix. (4) Failure of the critical engine occurs at (2) Final takeoff ice is the most critical ice VEF; and accretion on unprotected surfaces, and any (5) Crew activation of the ice protection ice accretion on the protected surfaces ap- system is in accordance with a normal oper- propriate to normal ice protection system ating procedure provided in the Airplane operation, between 400 feet and either 1,500 Flight Manual, except that after beginning feet above the takeoff surface, or the height the takeoff roll, it must be assumed that the at which the transition from the takeoff to crew takes no action to activate the ice pro- the en route configuration is completed and tection system until the airplane is at least VFTO is reached, whichever is higher. Ice ac- 400 feet above the takeoff surface. cretion is assumed to start at the end of the (e) The ice accretion before the ice protec- takeoff distance in the takeoff maximum tion system has been activated and is per- icing conditions of part I, paragraph (c) of forming its intended function is the critical this Appendix. ice accretion formed on the unprotected and (3) En route ice is the critical ice accretion normally protected surfaces before activa- on the unprotected surfaces, and any ice action and effective operation of the ice pro- cretion on the protected surfaces appropriate tection system in continuous maximum at- to normal ice protection system operation, mospheric icing conditions. This ice accre- during the en route phase. tion only applies in showing compliance to (4) Holding ice is the critical ice accretion §§ 25.143(j) and 25.207(h), and 25.207(i). on the unprotected surfaces, and any ice ac- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as cretion on the protected surfaces appropriate amended by Amdt. 25–121, 72 FR 44669, Aug. 8, to normal ice protection system operation, 2007; 72 FR 50467, Aug. 31, 2007; Amdt. 25–129, during the holding flight phase. 74 FR 38340, Aug. 3, 2009; Amdt. 25–140, 79 FR (5) Approach ice is the critical ice accretion 65528, Nov. 4, 2014] on the unprotected surfaces, and any ice accretion on the protected surfaces appropriate APPENDIX D TO PART 25 to normal ice protection system operation following exit from the holding flight phase Criteria for determining minimum flight crew. and transition to the most critical approach The following are considered by the Agency configuration. in determining the minimum flight crew (6) Landing ice is the critical ice accretion under § 25.1523: on the unprotected surfaces, and any ice ac- (a) Basic workload functions. The following cretion on the protected surfaces appropriate basic workload functions are considered: to normal ice protection system operation (1) Flight path control. following exit from the approach flight phase (2) Collision avoidance. and transition to the final landing configura- (3) Navigation. tion. (4) Communications.

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(5) Operation and monitoring of aircraft that each airplane certificated under this engines and systems. Part will operate under IFR conditions. (6) Command decisions. [Amdt. 25–3, 30 FR 6067, Apr. 29, 1965] (b) Workload factors. The following workload factors are considered significant when APPENDIX E TO PART 25 analyzing and demonstrating workload for minimum flight crew determination: I—Limited Weight Credit For Airplanes (1) The accessibility, ease, and simplicity Equipped With Standby Power of operation of all necessary flight, power, (a) Each applicant for an increase in the and equipment controls, including emer- maximum certificated takeoff and landing gency fuel shutoff valves, electrical controls, weights of an airplane equipped with a type- electronic controls, pressurization system certificated standby power rocket engine controls, and engine controls. may obtain an increase as specified in para- (2) The accessibility and conspicuity of all graph (b) if— necessary instruments and failure warning (1) The installation of the rocket engine devices such as fire warning, electrical sys- has been approved and it has been estab- tem malfunction, and other failure or cau- lished by flight test that the rocket engine tion indicators. The extent to which such in- and its controls can be operated safely and struments or devices direct the proper cor- reliably at the increase in maximum weight; rective action is also considered. and (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- pended or discharged. (5) The extent of required monitoring of (b) The currently approved maximum take- the fuel, hydraulic, pressurization, elec- off and landing weights at which an airplane trical, electronic, deicing, and other systems is certificated without a standby power rock- while en route. et engine installation may be increased by (6) The actions requiring a crewmember to an amount that does not exceed any of the be unavailable at his assigned duty station, following: including: observation of systems, emer- (1) An amount equal in pounds to 0.014 IN, gency operation of any control, and emer- where I is the maximum usable impulse in gencies in any compartment. pounds-seconds available from each standby (7) The degree of automation provided in power rocket engine and N is the number of the aircraft systems to afford (after failures rocket engines installed. or malfunctions) automatic crossover or iso- (2) An amount equal to 5 percent of the lation of difficulties to minimize the need for maximum certificated weight approved in flight crew action to guard against loss of accordance with the applicable airworthiness hydraulic or electric power to flight controls regulations without standby power rocket or to other essential systems. engines installed. (8) The communications and navigation (3) An amount equal to the weight of the workload. rocket engine installation. (9) The possibility of increased workload (4) An amount that, together with the cur- associated with any emergency that may rently approved maximum weight, would lead to other emergencies. equal the maximum structural weight estab- (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 pilots. II—Performance Credit for Transport Category 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

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weights, the takeoff distance, and the take- from the start of the takeoff to the point off paths, and may not exceed that found by where the airplane attains a height of 50 feet the Administrator to result in an overall above the takeoff surface for reciprocating- level of safety in the takeoff, approach, and engine-powered airplanes and a height of 35 landing regimes of flight equivalent to that feet above the takeoff surface for turbine- prescribed in the regulations under which powered airplanes. the airplane was originally certificated with- (4) Maximum certificated takeoff weights. The out standby power. For the purposes of this maximum certificated takeoff weights must appendix, ‘‘standby power’’ is power or be determined at all altitudes, and at ambi- thrust, or both, obtained from rocket en- ent temperatures, if applicable, at which per- gines for a relatively short period and actu- formance credit is to be applied and may not ated only in cases of emergency. The fol- exceed the weights established in compliance lowing provisions apply: with paragraphs (a) and (b) of this section. (1) Takeoff; general. The takeoff data pre- (a) The conditions of paragraphs (2)(b) 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.

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

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

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measured by a calibrated thermocouple py- Part II—Flammability of Seat Cushions rometer in the center of the flame must be (a) Criteria for Acceptance. Each seat cush- 1550 °F. Suitable precautions must be taken ion must meet the following criteria: to avoid drafts. The flame must be applied for 30 seconds with one-third contacting the (1) At least three sets of seat bottom and material at the center of the specimen and seat back cushion specimens must be tested. then removed. Flame time, glow time, and (2) If the cushion is constructed with a fire whether the flame penetrates (passes blocking material, the fire blocking material through) the specimen must be recorded. must completely enclose the cushion foam (7) Sixty degree test. A minimum of three core material. specimens of each wire specification (make (3) Each specimen tested must be fab- and size) must be tested. The specimen of ricated using the principal components (i.e., wire or cable (including insulation) must be foam core, flotation material, fire blocking placed at an angle of 60° with the horizontal material, if used, and dress covering) and as- in the cabinet specified in subparagraph (3) sembly processes (representative seams and of this paragraph with the cabinet door open closures) intended for use in the production during the test, or must be placed within a articles. If a different material combination chamber approximately 2 feet high by 1 foot is used for the back cushion than for the bot- by 1 foot, open at the top and at one vertical tom cushion, both material combinations side (front), and which allows sufficient flow must be tested as complete specimen sets, of air for complete combustion, but which is each set consisting of a back cushion speci- free from drafts. The specimen must be par- men and a bottom cushion specimen. If a allel to and approximately 6 inches from the cushion, including outer dress covering, is front of the chamber. The lower end of the demonstrated to meet the requirements of specimen must be held rigidly clamped. The this appendix using the oil burner test, the upper end of the specimen must pass over a dress covering of that cushion may be re- pulley or rod and must have an appropriate placed with a similar dress covering provided weight attached to it so that the specimen is the burn length of the replacement covering, held tautly throughout the flammability as determined by the test specified in test. The test specimen span between lower § 25.853(c), does not exceed the corresponding clamp and upper pulley or rod must be 24 burn length of the dress covering used on the inches and must be marked 8 inches from the cushion subjected to the oil burner test. lower end to indicate the central point for (4) For at least two-thirds of the total flame application. A flame from a Bunsen or number of specimen sets tested, the burn Tirrill burner must be applied for 30 seconds length from the burner must not reach the at the test mark. The burner must be mount- side of the cushion opposite the burner. The ed underneath the test mark on the speci- burn length must not exceed 17 inches. Burn men, perpendicular to the specimen and at length is the perpendicular distance from the an angle of 30° to the vertical plane of the inside edge of the seat frame closest to the specimen. The burner must have a nominal burner to the farthest evidence of damage to bore of 3⁄8-inch and be adjusted to provide a the test specimen due to flame impingement, 3-inch high flame with an inner cone ap- including areas of partial or complete con- proximately one-third of the flame height. sumption, charring, or embrittlement, but The minimum temperature of the hottest not including areas sooted, stained, warped, portion of the flame, as measured with a or discolored, or areas where material has calibrated thermocouple pyrometer, may not shrunk or melted away from the heat source. be less than 1750 °F. The burner must be posi- (5) The average percentage weight loss tioned so that the hottest portion of the must not exceed 10 percent. Also, at least flame is applied to the test mark on the two-thirds of the total number of specimen wire. Flame time, burn length, and flaming sets tested must not exceed 10 percent time of drippings, if any, must be recorded. weight loss. All droppings falling from the The burn length determined in accordance cushions and mounting stand are to be dis- with paragraph (8) of this paragraph must be carded before the after-test weight is deter- measured to the nearest tenth of an inch. mined. The percentage weight loss for a spec- Breaking of the wire specimens is not consid- imen set is the weight of the specimen set ered a failure. before testing less the weight of the speci- (8) Burn length. Burn length is the distance men set after testing expressed as the per- from the original edge to the farthest evi- centage of the weight before testing. dence of damage to the test specimen due to (b) Test Conditions. Vertical air velocity flame impingement, including areas of par- should average 25 fpm±10 fpm at the top of tial or complete consumption, charring, or the back seat cushion. Horizontal air veloc- embrittlement, but not including areas soot- ity should be below 10 fpm just above the ed, stained, warped, or discolored, nor areas bottom seat cushion. Air velocities should be where material has shrunk or melted away measured with the ventilation hood oper- from the heat source. ating and the burner motor off.

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(c) Test Specimens. (1) For each test, one set (ii) Because crumbling of the insulating of cushion specimens representing a seat bot- board with service can result in misalign- tom and seat back cushion must be used. ment of the calorimeter, the calorimeter (2) The seat bottom cushion specimen must must be monitored and the mounting be 18 ±1⁄8 inches (457 ±3 mm) wide by 20 ±1⁄8 shimmed, as necessary, to ensure that the inches (508 ±3 mm) deep by 4 ±1⁄8 inches (102 calorimeter face is flush with the exposed ±3 mm) thick, exclusive of fabric closures plane of the insulating board in a plane par- and seam overlap. allel to the exit of the test burner cone. (3) The seat back cushion specimen must (4) Thermocouples. The seven thermocouples ±1 ± ±1 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 the mounting stand legs is 12 ±1⁄8 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 side up. rimeter and the thermocouples. (2) Test Burner. The burner to be used in (7) Weight Scale. Weighing Device—A device testing must— must be used that with proper procedures (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.

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(5) If the calorimeter and thermocouples do run for 2 minutes to provide adequate not read within range, repeat steps in para- warmup of the burner cone and flame sta- graphs 1 through 4 and adjust the burner air bilization. intake damper until the proper readings are (5) To begin the test, swing the burner into obtained. The thermocouple rake and the the test position and simultaneously start calorimeter should be used frequently to the timing device. maintain and record calibrated test param- (6) Expose the seat bottom cushion speci- eters. Until the specific apparatus has dem- men to the burner flame for 2 minutes and onstrated consistency, each test should be then turn off the burner. Immediately swing calibrated. After consistency has been con- the burner away from the test position. Ter- firmed, several tests may be conducted with minate test 7 minutes after initiating cush- the pre-test calibration before and a calibra- ion exposure to the flame by use of a gaseous tion check after the series. extinguishing agent (i.e., Halon or CO2). (g) Test Procedure. The flammability of (7) Determine the weight of the remains of each set of specimens must be tested as fol- the seat cushion specimen set left on the lows: mounting stand to the nearest 0.02 pound (9 (1) Record the weight of each set of seat grams) excluding all droppings. bottom and seat back cushion specimens to (h) Test Report. With respect to all speci- be tested to the nearest 0.02 pound (9 grams). men sets tested for a particular seat cushion (2) Mount the seat bottom and seat back for which testing of compliance is performed, cushion test specimens on the test stand as the following information must be recorded: shown in Figure 2, securing the seat back (1) An identification and description of the cushion specimen to the test stand at the specimens being tested. top. (2) The number of specimen sets tested. (3) Swing the burner into position and en- (3) The initial weight and residual weight sure that the distance from the exit of the of each set, the calculated percentage weight burner cone to the side of the seat bottom loss of each set, and the calculated average cushion specimen is 4 ±1⁄8 inches (102 ±3 mm). percentage weight loss for the total number (4) Swing the burner away from the test of sets tested. position. Turn on the burner and allow it to (4) The burn length for each set tested.

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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, 5.25 × Testing and Materials (ASTM), ASTM E–906. 2.75 inches (133 × 70 mm) in cross section, and (1) This apparatus is shown in Figures 1A 10 inches (254 mm) long, fabricated from 28 and 1B of this part IV. All exterior surfaces guage stainless steel must be mounted on of the apparatus, except the holding cham- the outlet of the pyramidal section. A. 1.0 × ber, must be insulated with 1 inch (25 mm) 3.0 inch (25 × 76 mm) baffle plate of .018 ±.002 thick, low density, high temperature, fiber- inch (.50 ±.05 mm) stainless steel must be glass board insulation. A gasketed door, centered inside the stack, perpendicular to through which the sample injection rod the air flow, 3 inches (76 mm) above the base slides, must be used to form an airtight clo- of the stack. sure on the specimen hold chamber. (6) Specimen Holders. (i) The specimen must (2) Thermopile. The temperature difference be tested in a vertical orientation. The speci- between the air entering the environmental men holder (Figure 3 of this part IV) must chamber and that leaving must be monitored incorporate a frame that touches the speci- by a thermopile having five hot, and five men (which is wrapped with aluminum foil cold, 24-guage Chromel-Alumel junctions. as required by paragraph (d)(3) of this Part) The hot junctions must be spaced across the along only the .25 inch (6 mm) perimeter. A top of the exhaust stack, .38 inches (10 mm) ‘‘V’’ shaped spring is used to hold the assem- below the top of the chimney. The bly together. A detachable .50 × 50 × 5.91 inch thermocouples must have a .050 ±.010 inch (12 × 12 × 150 mm) drip pan and two .020 inch (1.3 ±.3mm) diameter, ball-type, welded tip. (.5 mm) stainless steel wires (as shown in One thermocouple must be located in the Figure 3 of this part IV) must be used for geometric center, with the other four located testing materials prone to melting and drip- 1.18 inch (30 mm) from the center along the ping. The positioning of the spring and frame diagonal toward each of the corners (Figure may be changed to accommodate different 5 of this part IV). The cold junctions must be specimen thicknesses by inserting the re- located in the pan below the lower air dis- taining rod in different holes on the speci- tribution plate (see paragraph (b)(4) of this men holder. part IV). Thermopile hot junctions must be (ii) Since the radiation shield described in cleared of soot deposits as needed to main- ASTM E–906 is not used, a guide pin must be tain the calibrated sensitivity. added to the injection mechanism. This fits (3) Radiation Source. A radiant heat source into a slotted metal plate on the injection incorporating four Type LL silicon carbide mechanism outside of the holding chamber. elements, 20 inches (508 mm) long by .63 inch It can be used to provide accurate posi- (16 mm) O.D., must be used, as shown in Fig- tioning of the specimen face after injection. ures 2A and 2B of this part IV. The heat The front surface of the specimen must be 3.9 source must have a nominal resistance of 1.4 inches (100 mm) from the closed radiation ohms and be capable of generating a flux up doors after injection. to 100 kW/m2. The silicone carbide elements (iii) The specimen holder clips onto the must be mounted in the stainless steel panel mounted bracket (Figure 3 of this part IV). box by inserting them through .63 inch (16 The mounting bracket must be attached to mm) holes in .03 inch (1 mm) thick ceramic the injection rod by three screws that pass fiber or calcium-silicate millboard. Loca- through a wide-area washer welded onto a 1⁄2- tions of the holes in the pads and stainless inch (13 mm) nut. The end of the injection steel cover plates are shown in Figure 2B of rod must be threaded to screw into the nut,

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and a .020 inch (5.1 mm) thick wide area the tubing, .50 inch (13 mm) apart, for gas washer must be held between two 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

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the specimen must be the same as that of the (4) The specimen must be placed in the aircraft component it represents up to a hold chamber with the radiation doors maximum thickness of 1.75 inches (45 mm). closed. The airtight outer door must be se- Test specimens representing thicker compo- cured, and the recording devices must be nents must be 1.75 inches (45 mm). started. The specimen must be retained in (2) Conditioning. Specimens must be condi- the hold chamber for 60 seconds, plus or tioned as described in Part 1 of this appen- minus 10 seconds, before injection. The ther- dix. mopile ‘‘zero’’ value must be determined dur- (3) Mounting. Each test specimen must be ing the last 20 seconds of the hold period. wrapped tightly on all sides of the specimen, The sample must not be injected before com- except for the one surface that is exposed pletion of the ‘‘zero’’ value determination. with a single layer of .001 inch (.025 mm) alu- (5) When the specimen is to be injected, the minum foil. radiation doors must be opened. After the (e) Procedure. (1) The power supply to the specimen is injected into the environmental radiant panel must be set to produce a radi- chamber, the radiation doors must be closed ± 2 ant flux of 3.5 .05 W/cm , as measured at the behind the specimen. point the center of the specimen surface will (6) [Reserved] occupy when positioned for the test. The ra- (7) Injection of the specimen and closure of diant flux must be measured after the air the inner door marks time zero. A record of flow through the equipment is adjusted to the thermopile output with at least one data the desired rate. (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) F1 = higher preset flow of methane (1pm) V = thermopile voltage at baseline (mv) (3) The integral of the heat release rate is 0 the total heat release as a function of time V1 = thermopile voltage at higher flow (mv) and is calculated by multiplying the rate by Ta = Ambient temperature (K) P = Ambient pressure (mm Hg) the data sampling frequency in minutes and Pv = Water vapor pressure (mm Hg) summing the time from zero to two minutes. (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 voltage at any instant of time as: each of the three or more samples tested must be averaged, and the peak heat release ()VVK− rate for each of the samples must be aver- HRR = mbn aged. The average total heat release must 2 not exceed 65 kilowatt-minutes per square .02323m meter, and the average peak heat release HRR = heat release rate (kw/m2) rate must not exceed 65 kilowatts per square Vb = baseline voltage (mv) meter. Vm = measured thermopile voltage (mv) 410

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

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

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(ii) Attach a 1⁄2 inch (13 mm) piece of (iii) Place the test specimen horizontally Kaowool M TM 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 Bernzomatic TM 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- (3) Basic control and operation information sulation specimens at a point 12 inches (30.5 describing how the airplane components and cm) from the face of the test rig. systems are controlled and how they operate, including any special procedures and [Amdt. 25–32, 37 FR 3972, Feb. 24, 1972] limitations that apply. EDITORIAL NOTE: For FEDERAL REGISTER ci- (4) Servicing information that covers de- tations affecting appendix F to Part 25, see tails regarding servicing points, capacities of the List of CFR Sections Affected, which ap- tanks, reservoirs, types of fluids to be used, pears in the Finding Aids section of the pressures applicable to the various systems, printed volume and at www.govinfo.gov. location of access panels for inspection and servicing, locations of lubrication points, lu- bricants to be used, equipment required for servicing, tow instructions and limitations, mooring, jacking, and leveling information. APPENDIX H TO PART 25—INSTRUCTIONS (b) Maintenance instructions. (1) Scheduling FOR CONTINUED AIRWORTHINESS information for each part of the airplane and its engines, auxiliary power units, propellers, H25.1 General. accessories, instruments, and equipment (a) This appendix specifies requirements that provides the recommended periods at for preparation of Instructions for Continued which they should be cleaned, inspected, ad- Airworthiness as required by §§ 25.1529, justed, tested, and lubricated, and the degree 25.1729, and applicable provisions of parts 21 of inspection, the applicable wear tolerances, and 26 of this chapter. and work recommended at these periods. (b) The Instructions for Continued Air- However, the applicant may refer to an accessory, instrument, or equipment manufac- worthiness for each airplane must include turer as the source of this information if the the Instructions for Continued Airworthiness applicant shows that the item has an excep- for each engine and propeller (hereinafter tionally high degree of complexity requiring designated ‘‘products’’), for each appliance specialized maintenance techniques, test required by this chapter, and any required equipment, or expertise. The recommended information relating to the interface of overhaul periods and necessary cross ref- those appliances and products with the air- erences to the Airworthiness Limitations plane. If Instructions for Continued Air- section of the manual must also be included. worthiness are not supplied by the manufac- In addition, the applicant must include an turer of an appliance or product installed in inspection program that includes the fre- the airplane, the Instructions for Continued quency and extent of the inspections nec- Airworthiness for the airplane must include essary to provide for the continued air- the information essential to the continued worthiness of the airplane. airworthiness of the airplane. (2) Troubleshooting information describing (c) The applicant must submit to the FAA probable malfunctions, how to recognize a program to show how changes to the In- those malfunctions, and the remedial action structions for Continued Airworthiness made for those malfunctions. by the applicant or by the manufacturers or (3) Information describing the order and products and appliances installed in the air- method of removing and replacing products plane will be distributed. and parts with any necessary precautions to H25.2 Format. be taken. (4) Other general procedural instructions (a) The Instructions for Continued Air- including procedures for system testing dur- worthiness must be in the form of a manual ing ground running, symmetry checks, or manuals as appropriate for the quantity weighing and determining the center of grav- of data to be provided. ity, lifting and shoring, and storage limita- (b) The format of the manual or manuals tions. must provide for a practical arrangement. (c) Diagrams of structural access plates H25.3 Content. and information needed to gain access for in- The contents of the manual or manuals spections when access plates are not pro- must be prepared in the English language. vided. The Instructions for Continued Airworthi- (d) Details for the application of special in- ness must contain the following manuals or spection techniques including radiographic sections, as appropriate, and information: and ultrasonic testing where such processes (a) Airplane maintenance manual or section. are specified. (1) Introduction information that includes an (e) Information needed to apply protective explanation of the airplane’s features and treatments to the structure after inspection. data to the extent necessary for mainte- (f) All data relative to structural fasteners nance or preventive maintenance. such as identification, discard recommenda- (2) A description of the airplane and its tions, and torque values. systems and installations including its en- (g) A list of special tools needed. gines, propellers, and appliances. H25.4 Airworthiness Limitations section.

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(a) The Instructions for Continued Air- and back-up hydraulic, mechanical, or elec- worthiness must contain a section titled Air- trical flight controls and lines. worthiness Limitations that is segregated (v) Identification of— and clearly distinguishable from the rest of (A) Tasks, and the intervals for performing the document. This section must set forth— those tasks, that will reduce the likelihood (1) Each mandatory modification time, re- of ignition sources and accumulation of com- placement time, structural inspection inter- bustible material, and val, and related structural inspection proce- (B) Procedures, and the intervals for per- dure approved under § 25.571. forming those procedures, that will effec- (2) Each mandatory replacement time, in- tively clean the EWIS components of com- spection interval, related inspection proce- bustible material if there is not an effective dure, and all critical design configuration task to reduce the likelihood of combustible control limitations approved under § 25.981 material accumulation. for the fuel tank system. (vi) Instructions for protections and cau- (3) Any mandatory replacement time of tion information that will minimize con- EWIS components as defined in section tamination and accidental damage to EWIS, 25.1701. as applicable, during performance of mainte- (4) A limit of validity of the engineering nance, alteration, or repairs. data that supports the structural mainte- (2) Acceptable EWIS maintenance practices in a standard format. nance program (LOV), stated as a total num- (3) Wire separation requirements as deter- ber of accumulated flight cycles or flight mined under § 25.1707. hours or both, approved under § 25.571. Until (4) Information explaining the EWIS iden- the full-scale fatigue testing is completed tification method and requirements for iden- and the FAA has approved the LOV, the tifying any changes to EWIS under § 25.1711. number of cycles accumulated by the air- (5) Electrical load data and instructions for 1 plane cannot be greater than ⁄2 the number updating that data. of cycles accumulated on the fatigue test ar- (b) The EWIS ICA developed in accordance ticle. with the requirements of H25.5(a)(1) must be (5) Each mandatory replacement time, in- in the form of a document appropriate for spection interval, and related inspection and the information to be provided, and they test procedure, and each critical design con- must be easily recognizable as EWIS ICA. figuration control limitation for each light- This document must either contain the re- ning protection feature approved under quired EWIS ICA or specifically reference § 25.954. other portions of the ICA that contain this (b) If the Instructions for Continued Air- information. worthiness consist of multiple documents, the section required by this paragraph must [Amdt. 25–54, 45 FR 60177, Sept. 11, 1980, as be included in the principal manual. This amended by Amdt. 25–68, 54 FR 34329, Aug. 18, section must contain a legible statement in 1989; Amdt. 25–102, 66 FR 23130, May 7, 2001; a prominent location that reads: ‘‘The Air- Amdt. 25–123, 72 FR 63408, Nov. 8, 2007; Amdt. worthiness Limitations section is FAA-ap- 25–132, 75 FR 69782, Nov. 15, 2010; Doc. No. proved and specifies maintenance required FAA–2014–1027, Amdt. No. 25–146, 83 FR 47557, under §§ 43.16 and 91.403 of the Federal Avia- Sept. 20, 2018] tion Regulations, unless an alternative program has been FAA approved.’’ APPENDIX I TO PART 25—INSTALLATION H25.5 Electrical Wiring Interconnection Sys- OF AN AUTOMATIC TAKEOFF THRUST tem (EWIS) Instructions for Continued Air- CONTROL SYSTEM (ATTCS) worthiness. I25.1 General. (a) The applicant must prepare Instruc- tions for Continued Airworthiness (ICA) ap- (a) This appendix specifies additional re- plicable to EWIS as defined by § 25.1701 that quirements for installation of an engine are approved by the FAA and include the fol- power control system that automatically lowing: resets thrust or power on operating engine(s) (1) Maintenance and inspection require- in the event of any one engine failure during ments for the EWIS developed with the use takeoff. of an enhanced zonal analysis procedure that (b) With the ATTCS and associated sys- includes: tems functioning normally as designed, all (i) Identification of each zone of the air- applicable requirements of Part 25, except as plane. provided in this appendix, must be met with- (ii) Identification of each zone that con- out requiring any action by the crew to in- tains EWIS. crease thrust or power. (iii) Identification of each zone containing I25.2 Definitions. EWIS that also contains combustible mate- (a) Automatic Takeoff Thrust Control System rials. (ATTCS). An ATTCS is defined as the entire (iv) Identification of each zone in which automatic system used on takeoff, including EWIS is in close proximity to both primary all devices, both mechanical and electrical,

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that sense engine failure, transmit signals, the minimum performance, all-engine flight actuate fuel controls or power levers or in- path where, assuming a simultaneous occur- crease engine power by other means on oper- rence of an engine and ATTCS failure, the ating engines to achieve scheduled thrust or resulting minimum flight path thereafter power increases, and furnish cockpit infor- intersects the Part 25 required actual flight mation on system operation. path at no less than 400 feet above the take- (b) Critical Time Interval. When conducting off surface. This time interval is shown in an ATTCS takeoff, the critical time interval the following illustration: is between V1 minus 1 second and a point on

I25.3 Performance and System Reliability Re- (2) Shall not result in a significant loss or quirements. reduction in thrust or power, or must be The applicant must comply with the per- shown to be an extremely improbable event. formance and ATTCS reliability require- (b) The concurrent existence of an ATTCS ments as follows: failure and an engine failure during the crit- (a) An ATTCS failure or a combination of ical time interval must be shown to be ex- failures in the ATTCS during the critical tremely improbable. time interval: (c) All applicable performance require- (1) Shall not prevent the insertion of the ments of Part 25 must be met with an engine maximum approved takeoff thrust or power, or failure occurring at the most critical point must be shown to be an improbable event.

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during takeoff with the ATTCS system func- tem that is independent of the ATTCS must tioning. be provided to give the pilot a clear warning I25.4 Thrust Setting. of any engine failure during takeoff. The initial takeoff thrust or power setting [Amdt. 25–62, 52 FR 43156, Nov. 9, 1987] on each engine at the beginning of the takeoff roll may not be less than any of the fol- APPENDIX J TO PART 25—EMERGENCY lowing: EVACUATION (a) Ninety (90) percent of the thrust or power set by the ATTCS (the maximum The following test criteria and procedures takeoff thrust or power approved for the air- must be used for showing compliance with plane under existing ambient conditions); § 25.803: (b) That required to permit normal oper- (a) The emergency evacuation must be con- ation of all safety-related systems and equip- ducted with exterior ambient light levels of ment dependent upon engine thrust or power no greater than 0.3 foot-candles prior to the lever position; or activation of the airplane emergency light- (c) That shown to be free of hazardous en- ing system. The source(s) of the initial exte- gine response characteristics when thrust or rior ambient light level may remain active power is advanced from the initial takeoff or illuminated during the actual demonstra- thrust or power to the maximum approved tion. There must, however, be no increase in takeoff thrust or power. the exterior ambient light level except for I25.5 Powerplant Controls. that due to activation of the airplane emer- (a) In addition to the requirements of gency lighting system. § 25.1141, no single failure or malfunction, or (b) The airplane must be in a normal atti- probable combination thereof, of the ATTCS, tude with landing gear extended. including associated systems, may cause the (c) Unless the airplane is equipped with an failure of any powerplant function necessary off-wing descent means, stands or ramps may for safety. be used for descent from the wing to the (b) The ATTCS must be designed to: ground. Safety equipment such as mats or (1) Apply thrust or power on the operating inverted life rafts may be placed on the floor engine(s), following any one engine failure or ground to protect participants. No other during takeoff, to achieve the maximum ap- equipment that is not part of the emergency proved takeoff thrust or power without ex- evacuation equipment of the airplane may be ceeding engine operating limits; used to aid the participants in reaching the (2) Permit manual decrease or increase in ground. thrust or power up to the maximum takeoff (d) Except as provided in paragraph (a) of thrust or power approved for the airplane this appendix, only the airplane’s emergency under existing conditions through the use of lighting system may provide illumination. the power lever. For airplanes equipped with (e) All emergency equipment required for limiters that automatically prevent engine the planned operation of the airplane must operating limits from being exceeded under be installed. existing ambient conditions, other means (f) Each internal door or curtain must be may be used to increase the thrust or power in the takeoff configuration. in the event of an ATTCS failure provided (g) Each crewmember must be seated in the means is located on or forward of the the normally assigned seat for takeoff and power levers; is easily identified and oper- must remain in the seat until receiving the ated under all operating conditions by a sin- signal for commencement of the demonstra- gle action of either pilot with the hand that tion. Each crewmember must be a person is normally used to actuate the power levers; having knowledge of the operation of exits and meets the requirements of § 25.777 (a), and emergency equipment and, if compliance (b), and (c); with § 121.291 is also being demonstrated, (3) Provide a means to verify to the each flight attendant must be a member of a flightcrew before takeoff that the ATTCS is regularly scheduled line crew. in a condition to operate; and (h) A representative passenger load of per- (4) Provide a means for the flightcrew to sons in normal health must be used as fol- deactivate the automatic function. This lows: means must be designed to prevent inad- (1) At least 40 percent of the passenger load vertent deactivation. must be female. I25.6 Powerplant Instruments. (2) At least 35 percent of the passenger load In addition to the requirements of § 25.1305: must be over 50 years of age. (a) A means must be provided to indicate (3) At least 15 percent of the passenger load when the ATTCS is in the armed or ready must be female and over 50 years of age. condition; and (4) Three life-size dolls, not included as (b) If the inherent flight characteristics of part of the total passenger load, must be car- the airplane do not provide adequate warn- ried by passengers to simulate live infants 2 ing that an engine has failed, a warning sys- years old or younger.

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(5) Crewmembers, mechanics, and training (r) The applicant’s approved procedures personnel, who maintain or operate the air- must be fully utilized, except the flightcrew plane in the normal course of their duties, must take no active role in assisting others may not be used as passengers. inside the cabin during the demonstration. (i) No passenger may be assigned a specific (s) The evacuation time period is com- seat except as the Administrator may re- pleted when the last occupant has evacuated quire. Except as required by subparagraph the airplane and is on the ground. Provided (g) of this paragraph, no employee of the ap- that the acceptance rate of the stand or plicant may be seated next to an emergency ramp is no greater than the acceptance rate exit. of the means available on the airplane for de- (j) Seat belts and shoulder harnesses (as re- scent from the wing during an actual crash quired) must be fastened. situation, evacuees using stands or ramps al- (k) Before the start of the demonstration, lowed by paragraph (c) of this appendix are approximately one-half of the total average considered to be on the ground when they are amount of carry-on baggage, blankets, pil- on the stand or ramp. lows, and other similar articles must be dis- [Amdt. 25–72, 55 FR 29788, July 20, 1990, as tributed at several locations in aisles and amended by Amdt. 25–79, Aug. 26, 1993; Amdt. emergency exit access ways to create minor 25–117, 69 FR 67499, Nov. 17, 2004] obstructions. (l) No prior indication may be given to any APPENDIX K TO PART 25—EXTENDED crewmember or passenger of the particular exits to be used in the demonstration. OPERATIONS (ETOPS) (m) The applicant may not practice, re- This appendix specifies airworthiness re- hearse, or describe the demonstration for the quirements for the approval of an airplane- participants nor may any participant have engine combination for extended operations taken part in this type of demonstration (ETOPS). For two-engine airplanes, the ap- within the preceding 6 months. plicant must comply with sections K25.1 and (n) Prior to entering the demonstration K25.2 of this appendix. For airplanes with aircraft, the passengers may also be advised more than two engines, the applicant must to follow directions of crewmembers but may comply with sections K25.1 and K25.3 of this not be instructed on the procedures to be fol- appendix. lowed in the demonstration, except with re- K25.1 Design requirements. spect to safety procedures in place for the K25.1.1 Part 25 compliance. demonstration or which have to do with the The airplane-engine combination must demonstration site. Prior to the start of the comply with the requirements of part 25 con- demonstration, the pre-takeoff passenger sidering the maximum flight time and the briefing required by § 121.571 may be given. longest diversion time for which the appli- Flight attendants may assign demonstration cant seeks approval. subjects to assist persons from the bottom of K25.1.2 Human factors. a slide, consistent with their approved train- An applicant must consider crew workload, ing program. operational implications, and the crew’s and (o) The airplane must be configured to pre- passengers’ physiological needs during con- vent disclosure of the active emergency exits tinued operation with failure effects for the to demonstration participants in the air- longest diversion time for which it seeks ap- plane until the start of the demonstration. proval. (p) Exits used in the demonstration must K25.1.3 Airplane systems. consist of one exit from each exit pair. The (a) Operation in icing conditions. demonstration may be conducted with the (1) The airplane must be certificated for escape slides, if provided, inflated and the operation in icing conditions in accordance exits open at the beginning of the dem- with § 25.1419. onstration. In this case, all exits must be (2) The airplane must be able to safely con- configured such that the active exits are not duct an ETOPS diversion with the most crit- disclosed to the occupants. If this method is ical ice accretion resulting from: used, the exit preparation time for each exit (i) Icing conditions encountered at an alti- utilized must be accounted for, and exits tude that the airplane would have to fly fol- that are not to be used in the demonstration lowing an engine failure or cabin decompres- must not be indicated before the demonstra- sion. tion has started. The exits to be used must (ii) A 15-minute hold in the continuous be representative of all of the emergency maximum icing conditions specified in Ap- exits on the airplane and must be designated pendix C of this part with a liquid water con- by the applicant, subject to approval by the tent factor of 1.0. Administrator. At least one floor level exit (iii) Ice accumulated during approach and must be used. landing in the icing conditions specified in (q) Except as provided in paragraph (c) of Appendix C of this part. this section, all evacuees must leave the air- (b) Electrical power supply. The airplane plane by a means provided as part of the air- must be equipped with at least three inde- plane’s equipment. pendent sources of electrical power.

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(c) Time limited systems. The applicant must (c) Engine oil tank design. The engine oil define the system time capability of each tank filler cap must comply with § 33.71(c)(4) ETOPS significant system that is time-lim- of this chapter. ited. K25.1.5 Engine-condition monitoring. K25.1.4 Propulsion systems. Procedures for engine-condition moni- (a) Fuel system design. Fuel necessary to toring must be specified and validated in ac- complete an ETOPS flight (including a diver- cordance with Part 33, Appendix A, para- sion for the longest time for which the appli- graph A33.3(c) of this chapter. cant seeks approval) must be available to the K25.1.6 Configuration, maintenance, and operating engines at the pressure and fuel- procedures. flow required by § 25.955 under any airplane The applicant must list any configuration, failure condition not shown to be extremely operating and maintenance requirements, improbable. Types of failures that must be hardware life limits, MMEL constraints, and considered include, but are not limited to: ETOPS approval in a CMP document. crossfeed valve failures, automatic fuel man- K25.1.7 Airplane flight manual. agement system failures, and normal elec- The airplane flight manual must contain trical power generation failures. the following information applicable to the (1) If the engine has been certified for lim- ETOPS type design approval: ited operation with negative engine-fuel- (a) Special limitations, including any limi- pump-inlet pressures, the following require- tation associated with operation of the air- ments apply: plane up to the maximum diversion time (i) Airplane demonstration-testing must being approved. cover worst case cruise and diversion condi- (b) Required markings or placards. tions involving: (c) The airborne equipment required for ex- (A) Fuel grade and temperature. tended operations and flightcrew operating (B) Thrust or power variations. procedures for this equipment. (C) Turbulence and negative G. (d) The system time capability for the fol- (D) Fuel system components degraded lowing: within their approved maintenance limits. (1) The most limiting fire suppression sys- (ii) Unusable-fuel quantity in the suction tem for Class C cargo or baggage compart- feed configuration must be determined in ac- ments. cordance with § 25.959. (2) The most limiting ETOPS significant (2) For two-engine airplanes to be certifi- system other than fire suppression systems cated for ETOPS beyond 180 minutes, one for Class C cargo or baggage compartments. fuel boost pump in each main tank and at (e) This statement: ‘‘The type-design reli- least one crossfeed valve, or other means for ability and performance of this airplane-en- transferring fuel, must be powered by an gine combination has been evaluated under independent electrical power source other 14 CFR 25.1535 and found suitable for (iden- than the three power sources required to tify maximum approved diversion time) ex- comply with section K25.1.3(b) of this appen- tended operations (ETOPS) when the con- dix. This requirement does not apply if the figuration, maintenance, and procedures normal fuel boost pressure, crossfeed valve standard contained in (identify the CMP doc- actuation, or fuel transfer capability is not ument) are met. The actual maximum ap- provided by electrical power. proved diversion time for this airplane may (3) An alert must be displayed to the be less based on its most limiting system flightcrew when the quantity of fuel avail- time capability. This finding does not con- able to the engines falls below the level re- stitute operational approval to conduct quired to fly to the destination. The alert ETOPS.’’ must be given when there is enough fuel re- K25.2. Two-engine airplanes. maining to safely complete a diversion. This An applicant for ETOPS type design ap- alert must account for abnormal fuel man- proval of a two-engine airplane must use one agement or transfer between tanks, and pos- of the methods described in section K25.2.1, sible loss of fuel. This paragraph does not K25.2.2, or K25.2.3 of this appendix. apply to airplanes with a required flight en- K25.2.1 Service experience method. gineer. An applicant for ETOPS type design ap- (b) APU design. If an APU is needed to com- proval using the service experience method ply with this appendix, the applicant must must comply with sections K25.2.1(a) and demonstrate that: K25.2.1(b) of this appendix before conducting (1) The reliability of the APU is adequate the assessments specified in sections to meet those requirements; and K25.2.1(c) and K25.2.1(d) of this appendix, and (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 (a) Service experience. The world fleet for of the airplane, or 45,000 feet, whichever is the airplane-engine combination must accu- lower, and run for the remainder of any mulate a minimum of 250,000 engine-hours. flight . The FAA may reduce this number of hours if

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the applicant identifies compensating fac- (F) Weather and other environmental con- tors that are acceptable to the FAA. The ditions; and compensating factors may include experi- (G) Cause of engine shutdown or occur- ence on another airplane, but experience on rence. the candidate airplane must make up a sig- (ii) A history of unscheduled engine re- nificant portion of the total service experi- moval rates since introduction into service ence. (using 6- and 12-month rolling averages), (b) In-flight shutdown (IFSD) rates. The with a summary of the major causes for the demonstrated 12-month rolling average IFSD removals. rate for the world fleet of the airplane-en- (iii) A list of all propulsion system events gine combination must be commensurate (whether or not caused by maintenance or with the level of ETOPS approval being flightcrew error), including dispatch delays, sought. cancellations, aborted takeoffs, turnbacks, (1) For type design approval up to and in- diversions, and flights that continue to des- cluding 120 minutes: An IFSD rate of 0.05 or tination after the event. less per 1,000 world-fleet engine-hours, unless (iv) The total number of engine hours and otherwise approved by the FAA. Unless the cycles, the number of hours for the engine IFSD rate is 0.02 or less per 1,000 world-fleet with the highest number of hours, the num- engine-hours, the applicant must provide a ber of cycles for the engine with the highest 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. (vi) A history of the IFSD rates since in- (2) For type design approval up to and introduction into service using a 12-month cluding 180 minutes: An IFSD rate of 0.02 or rolling average. less per 1,000 world-fleet engine-hours, unless (2) The cause or potential cause of each otherwise approved by the FAA. If the air- item listed in K25.2.1(c)(1)(i) must have a cor- plane-engine combination does not meet this rective action or actions that are shown to rate by compliance with an existing 120- be effective in preventing future occur- minute CMP document, then new or addi- rences. Each corrective action must be iden- tional CMP requirements that the applicant tified in the CMP document specified in sec- has demonstrated would achieve this IFSD tion K25.1.6. A corrective action is not re- rate must be added to the CMP document. quired: (3) For type design approval beyond 180 (i) For an item where the manufacturer is minutes: An IFSD rate of 0.01 or less per 1,000 unable to determine a cause or potential fleet engine-hours unless otherwise approved cause. by the FAA. If the airplane-engine combina- (ii) For an event where it is technically tion does not meet this rate by compliance unfeasible to develop a corrective action. with an existing 120-minute or 180-minute (iii) If the world-fleet IFSD rate— CMP document, then new or additional CMP (A) Is at or below 0.02 per 1,000 world-fleet requirements that the applicant has dem- 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

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in, an IFSD or diversion. The applicant must during the validation in accordance with the include in this assessment relevant problems problem tracking and resolution system with similar or identical equipment installed specified in section K25.2.2(h) of this appen- on other types of airplanes to the extent dix. such information is reasonably available. (d) Propulsion system validation test. (1) The (e) Airplane flight test. The applicant must installed engine configuration for which ap- conduct a flight test to validate the proval is being sought must comply with flightcrew’s ability to safely conduct an § 33.201(c) of this chapter. The test engine ETOPS diversion with an inoperative engine must be configured with a complete airplane and worst-case ETOPS Significant System nacelle package, including engine-mounted failures and malfunctions that could occur in equipment, except for any configuration dif- service. The flight test must validate the air- ferences necessary to accommodate test plane’s flying qualities and performance stand interfaces with the engine nacelle with the demonstrated failures and malfunc- package. At the conclusion of the test, the tions. propulsion system must be— K25.2.2 Early ETOPS method. (i) Visually inspected according to the ap- An applicant for ETOPS type design applicant’s on-wing inspection recommenda- proval using the Early ETOPS method must tions and limits; and comply with the following requirements: (ii) Completely disassembled and the pro- (a) Assessment of relevant experience with air- pulsion system hardware inspected to deter- planes previously certificated under part 25. mine whether it meets the service limits The applicant must identify specific correc- specified in the Instructions for Continued tive actions taken on the candidate airplane Airworthiness submitted in compliance with to prevent relevant design, manufacturing, § 25.1529. operational, and maintenance problems ex- (2) The applicant must identify, track, and perienced on airplanes previously certificated 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.

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(ii) Maximum duration flights with max- for Continued Airworthiness to establish its imum duration diversions. condition for continued safe operation. Each (iii) Maximum duration engine-inoperative engine must also undergo a gas path inspec- diversions distributed among the engines in- tion. These inspections must be conducted in stalled on the airplanes used for the airplane a manner to identify abnormal conditions demonstration flight test program. At least that could result in an IFSD or diversion. two one-engine-inoperative diversions must The applicant must identify, track and re- be conducted at maximum continuous thrust solve any abnormal conditions in accordance or power using the same engine. with the problem tracking and resolution (iv) Flights under non-normal conditions system specified in section K25.2.2(h) of this to demonstrate the flightcrew’s ability to appendix. safely conduct an ETOPS diversion with (h) Problem tracking and resolution system. worst-case ETOPS significant system fail- (1) The applicant must establish and main- ures or malfunctions that could occur in tain a problem tracking and resolution sys- service. tem. The system must: (v) Diversions to airports that represent (i) Contain a process for prompt reporting airports of the types used for ETOPS diver- to the FAA office responsible for the design sions. (vi) Repeated exposure to humid and in- approval of each occurrence reportable under clement weather on the ground followed by a § 21.4(a)(6) encountered during the phases of long-duration flight at normal cruise alti- airplane and engine development used to as- tude. sess Early ETOPS eligibility. (2) The airplane demonstration flight test (ii) Contain a process for notifying the program must validate the adequacy of the FAA office responsible for the design ap- airplane’s flying qualities and performance, proval of each proposed corrective action and the flightcrew’s ability to safely conduct that the applicant determines necessary for an ETOPS diversion under the conditions each problem identified from the occurrences specified in section K25.2.2(g)(1) of this ap- reported under section K25.2.2. (h)(1)(i) of pendix. this appendix. The timing of the notification (3) During the airplane demonstration must permit appropriate FAA review before flight test program, each test airplane must taking the proposed corrective action. be operated and maintained using the appli- (2) If the applicant is seeking ETOPS type cant’s recommended operating and mainte- design approval of a change to an airplane- nance procedures. engine combination previously approved for (4) At the completion of the airplane dem- ETOPS, the problem tracking and resolution onstration flight test program, each ETOPS system need only address those problems significant system must undergo an on-wing specified in the following table, provided the inspection or test in accordance with the applicant obtains prior authorization from tasks defined in the proposed Instructions 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- fleet of the candidate airplane-engine com- quency of failures and malfunctions on bination. ETOPS significant systems that occur dur- (b) The Early ETOPS requirements of ing the airplane flight test program and the K25.2.2, except for the airplane demonstra- airplane demonstration flight test program tion specified in section K25.2.2(g) of this ap- specified in section K25.2.2(g) of this appen- pendix; and dix must be consistent with the type and fre- (c) The flight test requirement of section quency of failures and malfunctions that K25.2.1(e) of this appendix. would be expected to occur on currently cer- K25.3. Airplanes with more than two engines. tificated airplanes approved for ETOPS. An applicant for ETOPS type design ap- K25.2.3. Combined service experience and proval of an airplane with more than two en- Early ETOPS method. gines must use one of the methods described An applicant for ETOPS type design ap- in section K25.3.1, K25.3.2, or K25.3.3 of this proval using the combined service experience appendix. and Early ETOPS method must comply with K25.3.1 Service experience method. the following requirements. An applicant for ETOPS type design ap- (a) A service experience requirement of not proval using the service experience method less than 15,000 engine-hours for the world must comply with section K25.3.1(a) of this

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appendix before conducting the airplane sys- substantiate its suitability for the airplane tems assessment specified in K25.3.1(b), and design. the flight test specified in section K25.3.1(c) (c) APU validation test. If an APU is needed of this appendix. to comply with this appendix, one APU of (a) Service experience. The world fleet for the type to be certified with the airplane the airplane-engine combination must accu- must be tested for 3,000 equivalent airplane mulate a minimum of 250,000 engine-hours. operational cycles. Following completion of The FAA may reduce this number of hours if the test, the APU must be disassembled and the applicant identifies compensating fac- inspected. The applicant must identify, tors that are acceptable to the FAA. The track, and resolve each cause or potential compensating factors may include experi- cause of an inability to start or operate the ence on another airplane, but experience on APU in flight as intended in accordance with the candidate airplane must make up a sig- the problem tracking and resolution system nificant portion of the total required service specified in section K25.3.2(e) of this appen- experience. dix. (b) Airplane systems assessment. The appli- (d) Airplane demonstration. For each air- cant must conduct an airplane systems as- plane-engine combination to be approved for sessment. The applicant must show that the ETOPS, the applicant must flight test at airplane systems comply with the § 25.1309(b) least one airplane to demonstrate that the using available in-service reliability data for airplane, and its components and equipment ETOPS significant systems on the candidate are capable of functioning properly during airplane-engine combination. Each cause or ETOPS flights and diversions of the longest potential cause of a relevant design, manu- duration for which the applicant seeks ap- facturing, operational or maintenance prob- proval. This flight testing may be performed 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

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in accordance with the tasks defined in the airplane and engine development used to as- proposed Instructions for Continued Air- sess Early ETOPS eligibility. worthiness to establish its condition for con- (ii) Contain a process for notifying the tinued safe operation. Each engine must also FAA office responsible for the design ap- undergo a gas path inspection. These inspec- proval of each proposed corrective action tions must be conducted in a manner to iden- that the applicant determines necessary for tify abnormal conditions that could result in each problem identified from the occurrences an IFSD or diversion. The applicant must reported under section K25.3.2(h)(1)(i) of this identify, track and resolve any abnormal appendix. The timing of the notification conditions in accordance with the problem must permit appropriate FAA review before tracking and resolution system specified in section K25.3.2(e) of this appendix. taking the proposed corrective action. (e) Problem tracking and resolution system. (2) If the applicant is seeking ETOPS type (1) The applicant must establish and main- design approval of a change to an airplane- tain a problem tracking and resolution sys- engine combination previously approved for tem. The system must: ETOPS, the problem tracking and resolution (i) Contain a process for prompt reporting system need only address those problems to the FAA office responsible for the design specified in the following table, provided the approval of each occurrence reportable under applicant obtains prior authorization from § 21.4(a)(6) encountered during the phases of 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- HIRF environments and equipment HIRF quency of failures and malfunctions on test levels are expressed in root-mean-square ETOPS significant systems that occur dur- units measured during the peak of the modu- ing the airplane flight test program and the lation cycle. airplane demonstration flight test program (a) HIRF environment I is specified in the specified in section K25.3.2(d) of this appen- following table: dix must be consistent with the type and frequency of failures and malfunctions that TABLE I.—HIRF ENVIRONMENT I would be expected to occur on currently certificated airplanes approved for ETOPS. Field strength K25.3.3 Combined service experience and Frequency (volts/meter) Early ETOPS method. Peak Average An applicant for ETOPS type design approval using the Early ETOPS method must 10 kHz–2 MHz ...... 50 50 comply with the following requirements: 2 MHz–30 MHz ...... 100 100 30 MHz–100 MHz ...... 50 50 (a) A service experience requirement of 100 MHz–400 MHz ...... 100 100 less than 15,000 engine-hours for the world 400 MHz–700 MHz ...... 700 50 fleet of the candidate airplane-engine com- 700 MHz–1 GHz ...... 700 100 bination; 1 GHz–2 GHz ...... 2,000 200 (b) The Early ETOPS requirements of sec- 2 GHz–6 GHz ...... 3,000 200 tion K25.3.2 of this appendix, except for the 6 GHz–8 GHz ...... 1,000 200 airplane demonstration specified in section 8 GHz–12 GHz ...... 3,000 300 12 GHz–18 GHz ...... 2,000 200 K25.3.2(d) of this appendix; and 18 GHz–40 GHz ...... 600 200 (c) The flight test requirement of section K25.3.1(c) of this appendix. In this table, the higher field strength applies at the frequency band edges. [Doc. No. FAA–2002–6717, 72 FR 1873, Jan. 16, (b) HIRF environment II is specified in the 2007, as amended by Doc. No. FAA–2018–0119, following table: Amdt. 25–145, 83 FR 9169, Mar. 5, 2018] ABLE NVIRONMENT APPENDIX L TO PART 25—HIRF ENVI- T II.–HIRF E II RONMENTS AND EQUIPMENT HIRF Field strength TEST LEVELS Frequency (volts/meter) This appendix specifies the HIRF environ- Peak Average ments and equipment HIRF test levels for 10 kHz–500 kHz ...... 20 20 electrical and electronic systems under 500 kHz–2 MHz ...... 30 30 § 25.1317. The field strength values for the 2 MHz–30 MHz ...... 100 100

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TABLE II.–HIRF ENVIRONMENT II—Continued (4) From 100 MHz to 8 GHz, use radiated susceptibility tests at a minimum of 5 V/m. Field strength (volts/meter) [Doc. No. FAA–2006–23657, 72 FR 44026, Aug. 6, Frequency 2007] Peak Average

30 MHz–100 MHz ...... 10 10 APPENDIX M TO PART 25—FUEL TANK 100 MHz–200 MHz ...... 30 10 SYSTEM FLAMMABILITY REDUCTION 200 MHz–400 MHz ...... 10 10 MEANS 400 MHz–1 GHz ...... 700 40 1 GHz–2 GHz ...... 1,300 160 M25.1 Fuel tank flammability exposure re- 2 GHz–4 GHz ...... 3,000 120 quirements. 4 GHz–6 GHz ...... 3,000 160 (a) The Fleet Average Flammability Expo- 6 GHz–8 GHz ...... 400 170 sure of each fuel tank, as determined in ac- 8 GHz–12 GHz ...... 1,230 230 12 GHz–18 GHz ...... 730 190 cordance with Appendix N of this part, may 18 GHz–40 GHz ...... 600 150 not exceed 3 percent of the Flammability Ex- posure Evaluation Time (FEET), as defined In this table, the higher field strength applies at the fre- in Appendix N of this part. As a portion of quency band edges. this 3 percent, if flammability reduction (c) Equipment HIRF Test Level 1. (1) From 10 means (FRM) are used, each of the following kilohertz (kHz) to 400 megahertz (MHz), use time periods may not exceed 1.8 percent of conducted susceptibility tests with contin- the FEET: uous wave (CW) and 1 kHz square wave mod- (1) When any FRM is operational but the ulation with 90 percent depth or greater. The fuel tank is not inert and the tank is flam- conducted susceptibility current must start mable; and at a minimum of 0.6 milliamperes (mA) at 10 (2) When any FRM is inoperative and the kHz, increasing 20 decibels (dB) per fre- tank is flammable. quency decade to a minimum of 30 mA at 500 (b) The Fleet Average Flammability Expo- kHz. sure, as defined in Appendix N of this part, of (2) From 500 kHz to 40 MHz, the conducted each fuel tank may not exceed 3 percent of susceptibility current must be at least 30 the portion of the FEET occurring during ei- mA. ther ground or takeoff/climb phases of flight (3) From 40 MHz to 400 MHz, use conducted during warm days. The analysis must con- susceptibility tests, starting at a minimum sider the following conditions. of 30 mA at 40 MHz, decreasing 20 dB per fre- (1) The analysis must use the subset of quency decade to a minimum of 3 mA at 400 those flights that begin with a sea level MHz. ground ambient temperature of 80 °F (stand- (4) From 100 MHz to 400 MHz, use radiated ard day plus 21 °F atmosphere) or above, susceptibility tests at a minimum of 20 volts from the flammability exposure analysis per meter (V/m) peak with CW and 1 kHz done for overall performance. square wave modulation with 90 percent (2) For the ground and takeoff/climb phases depth or greater. of flight, the average flammability exposure (5) From 400 MHz to 8 gigahertz (GHz), use must be calculated by dividing the time dur- radiated susceptibility tests at a minimum ing the specific flight phase the fuel tank is of 150 V/m peak with pulse modulation of 4 flammable by the total time of the specific percent duty cycle with a 1 kHz pulse repeti- flight phase. tion frequency. This signal must be switched (3) Compliance with this paragraph may be on and off at a rate of 1 Hz with a duty cycle shown using only those flights for which the of 50 percent. airplane is dispatched with the flammability (d) Equipment HIRF Test Level 2. Equipment reduction means operational. HIRF test level 2 is HIRF environment II in M25.2 Showing compliance. table II of this appendix reduced by accept- (a) The applicant must provide data from able aircraft transfer function and attenu- analysis, ground testing, and flight testing, ation curves. Testing must cover the fre- or any combination of these, that: quency band of 10 kHz to 8 GHz. (1) Validate the parameters used in the (e) Equipment HIRF Test Level 3. (1) From 10 analysis required by paragraph M25.1 of this kHz to 400 MHz, use conducted susceptibility appendix; tests, starting at a minimum of 0.15 mA at 10 (2) Substantiate that the FRM is effective kHz, increasing 20 dB per frequency decade at limiting flammability exposure in all to a minimum of 7.5 mA at 500 kHz. compartments of each tank for which the (2) From 500 kHz to 40 MHz, use conducted FRM is used to show compliance with para- susceptibility tests at a minimum of 7.5 mA. graph M25.1 of this appendix; and (3) From 40 MHz to 400 MHz, use conducted (3) Describe the circumstances under which susceptibility tests, starting at a minimum the FRM would not be operated during each of 7.5 mA at 40 MHz, decreasing 20 dB per fre- phase of flight. quency decade to a minimum of 0.75 mA at (b) The applicant must validate that the 400 MHz. FRM meets the requirements of paragraph

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M25.1 of this appendix with any airplane or in part 26 of this subchapter, to correct any engine configuration affecting the perform- failures of the FRM that occur in service ance of the FRM for which approval is that could increase any fuel tank’s Fleet Av- sought. erage Flammability Exposure to more than M25.3 Reliability indications and mainte- that required by paragraph M25.1 of this ap- nance access. pendix. (a) Reliability indications must be provided to identify failures of the FRM that [Doc. No. FAA–2005–22997, 73 FR 42494, July would otherwise be latent and whose identi- 21, 2008, as amended by Doc. No. FAA–2018– fication is necessary to ensure the fuel tank 0119, Amdt. 25–145, 83 FR 9169, Mar. 5, 2018] with an FRM meets the fleet average flammability exposure requirements listed in APPENDIX N TO PART 25—FUEL TANK paragraph M25.1 of this appendix, including FLAMMABILITY EXPOSURE AND RELI- when the FRM is inoperative. ABILITY ANALYSIS (b) Sufficient accessibility to FRM reliability indications must be provided for N25.1 General. maintenance personnel or the flightcrew. (a) This appendix specifies the require- (c) The access doors and panels to the fuel ments for conducting fuel tank fleet average tanks with FRMs (including any tanks that flammability exposure analyses required to communicate with a tank via a vent sys- meet § 25.981(b) and Appendix M of this part. tem), and to any other confined spaces or en- For fuel tanks installed in aluminum wings, closed areas that could contain hazardous at- a qualitative assessment is sufficient if it mosphere under normal conditions or failure substantiates that the tank is a conven- conditions, must be permanently stenciled, tional unheated wing tank. marked, or placarded to warn maintenance (b) This appendix defines parameters af- personnel of the possible presence of a poten- fecting fuel tank flammability that must be tially hazardous atmosphere. used in performing the analysis. These in- M25.4 Airworthiness limitations and proce- clude parameters that affect all airplanes dures. within the fleet, such as a statistical dis- (a) If FRM is used to comply with para- tribution of ambient temperature, fuel flash graph M25.1 of this appendix, Airworthiness point, flight lengths, and airplane descent Limitations must be identified for all main- rate. Demonstration of compliance also re- tenance or inspection tasks required to iden- quires application of factors specific to the tify failures of components within the FRM airplane model being evaluated. Factors that that are needed to meet paragraph M25.1 of need to be included are maximum range, this appendix. cruise mach number, typical altitude where (b) Maintenance procedures must be devel- the airplane begins initial cruise phase of oped to identify any hazards to be considered flight, fuel temperature during both ground during maintenance of the FRM. These pro- and flight times, and the performance of a cedures must be included in the instructions flammability reduction means (FRM) if infor continued airworthiness (ICA). stalled. M25.5 Reliability reporting. (c) The following definitions, input vari- The effects of airplane component failures ables, and data tables must be used in the on FRM reliability must be assessed on an program to determine fleet average flamma- on-going basis. The applicant/holder must do bility exposure for a specific airplane model. the following: N25.2 Definitions. (a) Demonstrate effective means to ensure (a) Bulk Average Fuel Temperature means collection of FRM reliability data. The the average fuel temperature within the fuel means must provide data affecting FRM reli- tank or different sections of the tank if the ability, such as component failures. tank is subdivided by baffles or compart- (b) Unless alternative reporting procedures ments. are approved by the responsible Aircraft Cer- (b) Flammability Exposure Evaluation Time tification Service office, as defined in part 26 (FEET). The time from the start of preparing of this subchapter, provide a report to the the airplane for flight, through the flight FAA every six months for the first five years and landing, until all payload is unloaded, after service introduction. After that period, and all passengers and crew have dis- continued reporting every six months may embarked. In the Monte Carlo program, the be replaced with other reliability tracking flight time is randomly selected from the methods found acceptable to the FAA or Flight Length Distribution (Table 2), the eliminated if it is established that the reli- pre-flight times are provided as a function of ability of the FRM meets, and will continue the flight time, and the post-flight time is a to meet, the exposure requirements of para- constant 30 minutes. graph M25.1 of this appendix. (c) Flammable. With respect to a fluid or (c) Develop service instructions or revise gas, flammable means susceptible to igniting the applicable airplane manual, according to readily or to exploding (14 CFR Part 1, Defi- a schedule approved by the responsible Air- nitions). A non-flammable ullage is one craft Certification Service office, as defined where the fuel-air vapor is too lean or too

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rich to burn or is inert as defined below. For tribution, equal to the square root of the the purposes of this appendix, a fuel tank arithmetic mean of the squares of the devi- that is not inert is considered flammable ations from the arithmetic means. when the bulk average fuel temperature (m) Transport Effects. For purposes of this within the tank is within the flammable appendix, transport effects are the change in range for the fuel type being used. For any fuel vapor concentration in a fuel tank fuel tank that is subdivided into sections by caused by low fuel conditions and fuel con- baffles or compartments, the tank is consid- densation and vaporization. ered flammable when the bulk average fuel (n) Ullage. The volume within the fuel tank temperature within any section of the tank, not occupied by liquid fuel. that is not inert, is within the flammable N25.3 Fuel tank flammability exposure anal- range for the fuel type being used. ysis. (d) Flash Point. The flash point of a flammable fluid means the lowest temperature at (a) A flammability exposure analysis must which the application of a flame to a heated be conducted for the fuel tank under evalua- sample causes the vapor to ignite momen- tion to determine fleet average flammability tarily, or ‘‘flash.’’ Table 1 of this appendix exposure for the airplane and fuel types provides the flash point for the standard fuel under evaluation. For fuel tanks that are to be used in the analysis. subdivided by baffles or compartments, an (e) Fleet average flammability exposure is the analysis must be performed either for each percentage of the flammability exposure section of the tank, or for the section of the evaluation time (FEET) each fuel tank tank having the highest flammability expo- ullage is flammable for a fleet of an airplane sure. Consideration of transport effects is type operating over the range of flight not allowed in the analysis. The analysis lengths in a world-wide range of environ- must be done in accordance with the meth- mental conditions and fuel properties as de- ods and procedures set forth in the Fuel fined in this appendix. Tank Flammability Assessment Method (f) Gaussian Distribution is another name User’s Manual, dated May 2008, document for the normal distribution, a symmetrical number DOT/FAA/AR–05/8 (incorporated by frequency distribution having a precise reference, see § 25.5). The parameters speci- mathematical formula relating the mean fied in sections N25.3(b) and (c) of this appen- and standard deviation of the samples. dix must be used in the fuel tank flamma- Gaussian distributions yield bell-shaped fre- bility exposure ‘‘Monte Carlo’’ analysis. quency curves having a preponderance of val- (b) The following parameters are defined in ues around the mean with progressively the Monte Carlo analysis and provided in fewer observations as the curve extends out- paragraph N25.4 of this appendix: ward. (1) Cruise Ambient Temperature, as de- (g) Hazardous atmosphere. An atmosphere fined in this appendix. that may expose maintenance personnel, (2) Ground Ambient Temperature, as de- passengers or flight crew to the risk of fined in this appendix. 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. (i) Inerting. A process where a noncombus- must be provided as inputs to the Monte tible gas is introduced into the ullage of a Carlo analysis are: fuel tank so that the ullage becomes non- (1) Airplane cruise altitude. flammable. (2) Fuel tank quantities. If fuel quantity (j) Monte Carlo Analysis. The analytical affects fuel tank flammability, inputs to the method that is specified in this appendix as Monte Carlo analysis must be provided that the compliance means for assessing the fleet represent the actual fuel quantity within the average flammability exposure time for a fuel tank or compartment of the fuel tank fuel tank. throughout each of the flights being evalu- (k) Oxygen evolution occurs when oxygen ated. Input values for this data must be ob- dissolved in the fuel is released into the tained from ground and flight test data or ullage as the pressure and temperature in the approved FAA fuel management proce- the fuel tank are reduced. dures. (l) Standard deviation is a statistical meas- (3) Airplane cruise mach number. ure of the dispersion or variation in a dis- (4) Airplane maximum range.

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(5) Fuel tank thermal characteristics. If must include any times when oxygen evo- fuel temperature affects fuel tank flamma- lution from the fuel in the tank or compart- bility, inputs to the Monte Carlo analysis ment under evaluation would result in a must be provided that represent the actual flammable fuel tank. The oxygen evolution bulk average fuel temperature within the rate that must be used is defined in the Fuel fuel tank at each point in time throughout Tank Flammability Assessment Method each of the flights being evaluated. For fuel User’s Manual, dated May 2008, document tanks that are subdivided by baffles or com- number DOT/FAA/AR–05/8 (incorporated by partments, bulk average fuel temperature reference in § 25.5). inputs must be provided for each section of (6) If an inerting system FRM is used, the the tank. Input values for these data must be effects of any air that may enter the fuel obtained from ground and flight test data or tank following the last flight of the day due a thermal model of the tank that has been to changes in ambient temperature, as de- validated by ground and flight test data. fined in Table 4, during a 12-hour overnight (6) Maximum airplane operating tempera- period. ture limit, as defined by any limitations in (e) The applicant must submit to the re- the airplane flight manual. sponsible Aircraft Certification Service (7) Airplane Utilization. The applicant officefor approval the fuel tank flammability must provide data supporting the number of analysis, including the airplane-specific pa- flights per day and the number of hours per rameters identified under paragraph N25.3(c) flight for the specific airplane model under of this appendix and any deviations from the evaluation. If there is no existing airplane parameters identified in paragraph N25.3(b) fleet data to support the airplane being eval- of this appendix that affect flammability ex- uated, the applicant must provide substan- posure, substantiating data, and any air- tiation that the number of flights per day worthiness limitations and other conditions and the number of hours per flight for that assumed in the analysis. airplane model is consistent with the exist- N25.4 Variables and data tables. ing fleet data they propose to use. The following data must be used when con- (d) Fuel Tank FRM Model. If FRM is used, ducting a flammability exposure analysis to an FAA approved Monte Carlo program must determine the fleet average flammability ex- be used to show compliance with the flam- posure. Variables used to calculate fleet mability requirements of § 25.981 and Appen- flammability exposure must include atmos- dix M of this part. The program must deter- pheric ambient temperatures, flight length, mine the time periods during each flight flammability exposure evaluation time, fuel phase when the fuel tank or compartment flash point, thermal characteristics of the with the FRM would be flammable. The fol- fuel tank, overnight temperature drop, and lowing factors must be considered in estab- oxygen evolution from the fuel into the lishing these time periods: ullage. (1) Any time periods throughout the flam- (a) Atmospheric Ambient Temperatures mability exposure evaluation time and under and Fuel Properties. the full range of expected operating condi- (1) In order to predict flammability expo- tions, when the FRM is operating properly sure during a given flight, the variation of but fails to maintain a non-flammable fuel ground ambient temperatures, cruise ambi- tank because of the effects of the fuel tank ent temperatures, and a method to compute vent system or other causes, the transition from ground to cruise and (2) If dispatch with the system inoperative back again must be used. The variation of under the Master Minimum Equipment List the ground and cruise ambient temperatures (MMEL) is requested, the time period as- and the flash point of the fuel is defined by sumed in the reliability analysis (60 flight a Gaussian curve, given by the 50 percent hours must be used for a 10-day MMEL dis- value and a ±1-standard deviation value. patch limit unless an alternative period has (2) Ambient Temperature: Under the pro- been approved by the Administrator), gram, the ground and cruise ambient tem- (3) Frequency and duration of time periods peratures are linked by a set of assumptions of FRM inoperability, substantiated by test on the atmosphere. The temperature varies or analysis acceptable to the FAA, caused by with altitude following the International latent or known failures, including airplane Standard Atmosphere (ISA) rate of change system shut-downs and failures that could from the ground ambient temperature until cause the FRM to shut down or become inop- the cruise temperature for the flight is erative. reached. Above this altitude, the ambient (4) Effects of failures of the FRM that temperature is fixed at the cruise ambient could increase the flammability exposure of temperature. This results in a variation in the fuel tank. the upper atmospheric temperature. For cold (5) If an FRM is used that is affected by ox- days, an inversion is applied up to 10,000 feet, ygen concentrations in the fuel tank, the and then the ISA rate of change is used. time periods when oxygen evolution from the (3) Fuel properties: fuel results in the fuel tank or compartment (i) For Jet A fuel, the variation of flash exceeding the inert level. The applicant point of the fuel is defined by a Gaussian

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curve, given by the 50 percent value and a ±1- decreases from sea level value with increas- standard deviation, as shown in Table 1 of ing altitude at a rate of 1 °F per 808 feet. this appendix. (B) UFL at sea level = flash point tempera- (ii) The flammability envelope of the fuel ture of the fuel at sea level plus 63.5 °F. UFL that must be used for the flammability expo- decreases from the sea level value with insure analysis is a function of the flash point creasing altitude at a rate of 1 °F per 512 of the fuel selected by the Monte Carlo for a feet. given flight. The flammability envelope for (4) For each flight analyzed, a separate the fuel is defined by the upper flammability random number must be generated for each limit (UFL) and lower flammability limit of the three parameters (ground ambient (LFL) as follows: temperature, cruise ambient temperature, (A) LFL at sea level = flash point tempera- and fuel flash point) using the Gaussian dis- ture of the fuel at sea level minus 10 °F. LFL tribution 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

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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

7000 7200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.4 0.7 0.8 7200 7400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.3 0.5 0.7 7400 7600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.5 0.6 7600 7800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 0.5 0.7 7800 8000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 0.6 0.8 8000 8200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 0.8 8200 8400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 1.0 8400 8600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.6 1.3 8600 8800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.4 1.1 8800 9000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.8 9000 9200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 9200 9400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 9400 9600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 9600 9800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 9800 10000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1

(c) Overnight Temperature Drop. For air- with the fleet average and warm day flam- planes on which FRM is installed, the over- mability exposure requirements, the appli- night temperature drop for this appendix is cant must run the analysis for a minimum defined using: number of flights to ensure that the fleet av- (1) A temperature at the beginning of the erage and warm day flammability exposure overnight period that equals the landing for the fuel tank under evaluation meets the temperature of the previous flight that is a applicable flammability limits defined in random value based on a Gaussian distribu- Table 5 of this appendix. tion; and (2) An overnight temperature drop that is TABLE 5.—FLAMMABILITY EXPOSURE LIMIT a random value based on a Gaussian distribution. Maximum Maximum (3) For any flight that will end with an acceptable Monte acceptable Monte overnight ground period (one flight per day Carlo average fuel Carlo average fuel Minimum number of tank flammability tank flammability out of an average number of flights per day, flights in Monte exposure exposure depending on utilization of the particular Carlo analysis (percent) to meet (percent) to meet airplane model being evaluated), the landing 3 percent 7 percent part 26 requirements requirements outside air temperature (OAT) is to be chosen as a random value from the following 10,000 ...... 2.91 6.79 Gaussian curve: 100,000 ...... 2.98 6.96 1,000,000 ...... 3.00 7.00 TABLE 3.—LANDING OUTSIDE AIR TEMPERATURE [Doc. No. FAA–2005–22997, 73 FR 42495, July Parameter Landing outside air temperature °F 21, 2008, as amended by Doc. No. FAA–2018– 0119, Amdt. 25–145, 83 FR 9169, Mar. 5, 2018] Mean Temperature ...... 58.68 negative 1 std dev ...... 20.55 positive 1 std dev ...... 13.21 APPENDIX O TO PART 25—SUPERCOOLED LARGE DROP ICING CONDITIONS (4) The outside ambient air temperature (OAT) overnight temperature drop is to be This Appendix consists of two parts. Part I chosen as a random value from the following defines this Appendix as a description of Gaussian curve: supercooled large drop icing conditions in which the drop median volume diameter TABLE 4.—OUTSIDE AIR TEMPERATURE (OAT) (MVD) is less than or greater than 40 μm, the DROP maximum mean effective drop diameter (MED) of Appendix C of this part continuous OAT drop maximum (stratiform clouds) icing condi- Parameter ° temperature F tions. For this Appendix, supercooled large Mean Temp ...... 12.0 drop icing conditions consist of freezing driz- 1 std dev ...... 6.0 zle and freezing rain occurring in and/or below stratiform clouds. Part II defines ice (d) Number of Simulated Flights Required accretions used to show compliance with the in Analysis. In order for the Monte Carlo airplane performance and handling qualities analysis to be valid for showing compliance requirements of subpart B of this part.

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PART I—METEOROLOGY (1) Pressure altitude range: 0 to 12,000 ft MSL. In this Appendix icing conditions are de- (2) Maximum vertical extent: 7,000 ft. fined by the parameters of altitude, vertical (3) Horizontal extent: Standard distance of and horizontal extent, temperature, liquid 17.4 nautical miles. water content, and water mass distribution (4) Total liquid water content. as a function of drop diameter distribution. 3 (a) Freezing Drizzle (Conditions with spec- NOTE: LWC in grams per cubic meter (g/m ) tra maximum drop diameters from 100μm to based on horizontal extent standard distance 500 μm): of 17.4 nautical miles. (1) Pressure altitude range: 0 to 22,000 feet (5) Drop Diameter Distribution: Figure 5. MSL. (6) Altitude and temperature envelope: Figure 6. (2) Maximum vertical extent: 12,000 feet. (c) Horizontal extent. (3) Horizontal extent: Standard distance of The liquid water content for freezing driz- 17.4 nautical miles. zle and freezing rain conditions for hori- (4) Total liquid water content. zontal extents other than the standard 17.4 NOTE: Liquid water content (LWC) in nautical miles can be determined by the grams per cubic meter (g/m3) based on hori- value of the liquid water content determined zontal extent standard distance of 17.4 nau- from Figure 1 or Figure 4, multiplied by the tical miles. factor provided in Figure 7, which is defined (5) Drop diameter distribution: Figure 2. by the following equation: (6) Altitude and temperature envelope: S = 1.266 ¥ 0.213 log10(H) Figure 3. Where: (b) Freezing Rain (Conditions with spectra S = Liquid Water Content Scale Factor maximum drop diameters greater than 500 (dimensionless) and μm): H = horizontal extent in nautical miles

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PART II—AIRFRAME ICE ACCRETIONS in part II, paragraph (c) of this Appendix, FOR SHOWING COMPLIANCE WITH SUB- only the portion of part I of this Appendix in PART B OF THIS PART which the airplane is capable of operating (a) General. The most critical ice accretion safely must be considered. in terms of airplane performance and han- (3) For an airplane certified in accordance dling qualities for each flight phase must be with § 25.1420(a)(3), the ice accretions for each used to show compliance with the applicable flight phase are defined in part II, paragraph airplane performance and handling qualities (c) of this Appendix. requirements for icing conditions contained (b) Ice accretions for airplanes certified in in subpart B of this part. Applicants must accordance with § 25.1420(a)(1) or (2). demonstrate that the full range of atmos- (1) En route ice is the en route ice as de- pheric icing conditions specified in part I of fined by part II, paragraph (c)(3), of this Ap- this Appendix have been considered, includ- pendix, for an airplane certified in according drop diameter distributions, liquid water ance with § 25.1420(a)(2), or defined by part II, content, and temperature appropriate to the paragraph (a)(3), of Appendix C of this part, flight conditions (for example, configuration, for an airplane certified in accordance with speed, angle of attack, and altitude). § 25.1420(a)(1), plus: (1) For an airplane certified in accordance with § 25.1420(a)(1), the ice accretions for each (i) Pre-detection ice as defined by part II, flight phase are defined in part II, paragraph paragraph (b)(5), of this Appendix; and (b) of this Appendix. (ii) The ice accumulated during the transit (2) For an airplane certified in accordance of one cloud with a horizontal extent of 17.4 with § 25.1420(a)(2), the most critical ice ac- nautical miles in the most critical of the cretion for each flight phase defined in part icing conditions defined in part I of this Ap- II, paragraphs (b) and (c) of this Appendix, pendix and one cloud with a horizontal ex- must be used. For the ice accretions defined tent of 17.4 nautical miles in the continuous

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maximum icing conditions defined in Appen- (b)(2), of this Appendix, or the ice calculated dix C of this part. in the applicable paragraphs (b)(4)(i) or (ii) of (2) Holding ice is the holding ice defined by part II of this Appendix: part II, paragraph (c)(4), of this Appendix, for (i) For an airplane certified in accordance an airplane certified in accordance with with § 25.1420(a)(2), the ice accretion defined § 25.1420(a)(2), or defined by part II, paragraph by part II, paragraph (c)(5)(i), of this Appen- (a)(4), of Appendix C of this part, for an air- dix, plus a descent from 2,000 feet above the plane certified in accordance with landing surface to a height of 200 feet above § 25.1420(a)(1), plus: the landing surface with a transition to the (i) Pre-detection ice as defined by part II, landing configuration in the icing conditions paragraph (b)(5), of this Appendix; and defined in part I of this Appendix, plus: (ii) The ice accumulated during the transit (A) Pre-detection ice, as defined in part II, of one cloud with a 17.4 nautical miles hori- paragraph (b)(5), of this Appendix; and zontal extent in the most critical of the icing conditions defined in part I of this Ap- (B) The ice accumulated during an exit pendix and one cloud with a horizontal ex- maneuver, beginning with the minimum tent of 17.4 nautical miles in the continuous climb gradient required by § 25.119, from a maximum icing conditions defined in Appen- height of 200 feet above the landing surface dix C of this part. through one cloud with a horizontal extent (iii) Except the total exposure to holding of 17.4 nautical miles in the most critical of ice conditions does not need to exceed 45 the icing conditions defined in part I of this minutes. Appendix and one cloud with a horizontal ex- (3) Approach ice is the more critical of the tent of 17.4 nautical miles in the continuous holding ice defined by part II, paragraph maximum icing conditions defined in Appen- (b)(2), of this Appendix, or the ice calculated dix C of this part. in the applicable paragraphs (b)(3)(i) or (ii) of (ii) For an airplane certified in accordance part II, of this Appendix: with § 25.1420(a)(1), the ice accumulated in (i) For an airplane certified in accordance the maximum continuous icing conditions with § 25.1420(a)(2), the ice accumulated dur- defined in Appendix C of this part, during a ing descent from the maximum vertical ex- descent from the maximum vertical extent tent of the icing conditions defined in part I of the icing conditions defined in Appendix C of this Appendix to 2,000 feet above the land- of this part, to 2,000 feet above the landing ing surface in the cruise configuration, plus surface in the cruise configuration, plus transition to the approach configuration, transition to the approach configuration and plus: flying for 15 minutes at 2,000 feet above the (A) Pre-detection ice, as defined by part II, landing surface, plus a descent from 2,000 paragraph (b)(5), of this Appendix; and feet above the landing surface to a height of (B) The ice accumulated during the transit 200 feet above the landing surface with a at 2,000 feet above the landing surface of one transition to the landing configuration, plus: cloud with a horizontal extent of 17.4 nau- (A) Pre-detection ice, as described by part tical miles in the most critical of the icing II, paragraph (b)(5), of this Appendix; and conditions defined in part I of this Appendix (B) The ice accumulated during an exit and one cloud with a horizontal extent of 17.4 nautical miles in the continuous maximum maneuver, beginning with the minimum icing conditions defined in Appendix C of climb gradient required by § 25.119, from a this part. height of 200 feet above the landing surface (ii) For an airplane certified in accordance through one cloud with a horizontal extent with § 25.1420(a)(1), the ice accumulated dur- of 17.4 nautical miles in the most critical of ing descent from the maximum vertical ex- the icing conditions defined in part I of this tent of the maximum continuous icing condi- Appendix and one cloud with a horizontal ex- tions defined in part I of Appendix C to 2,000 tent of 17.4 nautical miles in the continuous feet above the landing surface in the cruise maximum icing conditions defined in Appen- configuration, plus transition to the ap- dix C of this part. proach configuration, plus: (5) Pre-detection ice is the ice accretion be- (A) Pre-detection ice, as defined by part II, fore detection of flight conditions in this Ap- paragraph (b)(5), of this Appendix; and pendix that require exiting per § 25.1420(a)(1) (B) The ice accumulated during the transit and (2). It is the pre-existing ice accretion at 2,000 feet above the landing surface of one that may exist from operating in icing condi- cloud with a horizontal extent of 17.4 nau- tions in which the airplane is approved to optical miles in the most critical of the icing erate prior to encountering the icing condi- conditions defined in part I of this Appendix tions requiring an exit, plus the ice accumu- and one cloud with a horizontal extent of 17.4 lated during the time needed to detect the nautical miles in the continuous maximum icing conditions, followed by two minutes of icing conditions defined in Appendix C of further ice accumulation to take into ac- this part. count the time for the flightcrew to take ac- (4) Landing ice is the more critical of the tion to exit the icing conditions, including holding ice as defined by part II, paragraph coordination with air traffic control.

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(i) For an airplane certified in accordance flying for 15 minutes at 2,000 feet above the with § 25.1420(a)(1), the pre-existing ice accre- landing surface; or tion must be based on the icing conditions (ii) Holding ice as defined by part II, para- defined in Appendix C of this part. graph (c)(4), of this Appendix. (ii) For an airplane certified in accordance (6) Landing ice is the ice accretion on the with § 25.1420(a)(2), the pre-existing ice accre- unprotected surfaces, and any ice accretion tion must be based on the more critical of on the protected surfaces appropriate to nor- the icing conditions defined in Appendix C of mal ice protection system operation, result- this part, or the icing conditions defined in ing from the more critical of the: part I of this Appendix in which the airplane (i) Ice accretion defined by part II, para- is capable of safely operating. graph (c)(5)(i), of this Appendix, plus ice ac- (c) Ice accretions for airplanes certified in accumulated in the icing conditions defined in cordance with §§ 25.1420(a)(2) or (3). For an air- part I of this Appendix during a descent from plane certified in accordance with 2,000 feet above the landing surface to a § 25.1420(a)(2), only the portion of the icing height of 200 feet above the landing surface conditions of part I of this Appendix in with a transition to the landing configura- which the airplane is capable of operating tion, followed by a go-around at the min- safely must be considered. imum climb gradient required by § 25.119, (1) Takeoff ice is the most critical ice accre- from a height of 200 feet above the landing tion on unprotected surfaces, and any ice ac- surface to 2,000 feet above the landing sur- cretion on the protected surfaces, occurring between the end of the takeoff distance and face, flying for 15 minutes at 2,000 feet above 400 feet above the takeoff surface, assuming the landing surface in the approach configu- accretion starts at the end of the takeoff dis- ration, and a descent to the landing surface tance in the icing conditions defined in part (touchdown) in the landing configuration; or I of this Appendix. (ii) Holding ice as defined by part II, para- (2) Final takeoff ice is the most critical ice graph (c)(4), of this Appendix. accretion on unprotected surfaces, and any (7) For both unprotected and protected ice accretion on the protected surfaces ap- parts, the ice accretion for the takeoff phase propriate to normal ice protection system must be determined for the icing conditions operation, between 400 feet and either 1,500 defined in part I of this Appendix, using the feet above the takeoff surface, or the height following assumptions: at which the transition from the takeoff to (i) The airfoils, control surfaces, and, if ap- the en route configuration is completed and plicable, propellers are free from frost, snow, VFTO is reached, whichever is higher. Ice ac- or ice at the start of takeoff; cretion is assumed to start at the end of the (ii) The ice accretion starts at the end of takeoff distance in the icing conditions de- the takeoff distance; fined in part I of this Appendix. (iii) The critical ratio of thrust/power-to- (3) En route ice is the most critical ice ac- weight; cretion on the unprotected surfaces, and any (iv) Failure of the critical engine occurs at ice accretion on the protected surfaces ap- V ; and propriate to normal ice protection system EF (v) Crew activation of the ice protection operation, during the en route flight phase in system is in accordance with a normal oper- the icing conditions defined in part I of this ating procedure provided in the airplane Appendix. flight manual, except that after beginning (4) Holding ice is the most critical ice ac- the takeoff roll, it must be assumed that the cretion on the unprotected surfaces, and any ice accretion on the protected surfaces ap- crew takes no action to activate the ice pro- propriate to normal ice protection system tection system until the airplane is at least operation, resulting from 45 minutes of flight 400 feet above the takeoff surface. within a cloud with a 17.4 nautical miles hor- (d) The ice accretion before the ice protec- izontal extent in the icing conditions defined tion system has been activated and is per- in part I of this Appendix, during the holding forming its intended function is the critical phase of flight. ice accretion formed on the unprotected and (5) Approach ice is the ice accretion on the normally protected surfaces before activa- unprotected surfaces, and any ice accretion tion and effective operation of the ice pro- on the protected surfaces appropriate to nor- tection system in the icing conditions de- mal ice protection system operation, result- fined in part I of this Appendix. This ice ac- ing from the more critical of the: cretion only applies in showing compliance (i) Ice accumulated in the icing conditions to §§ 25.143(j) and 25.207(h). defined in part I of this Appendix during a (e) In order to reduce the number of ice ac- descent from the maximum vertical extent cretions to be considered when dem- of the icing conditions defined in part I of onstrating compliance with the require- this Appendix, to 2,000 feet above the landing ments of § 25.21(g), any of the ice accretions surface in the cruise configuration, plus defined in this Appendix may be used for any transition to the approach configuration and other flight phase if it is shown to be at least

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as critical as the specific ice accretion de- Subpart A—General fined for that flight phase. Configuration differences and their effects on ice accretions § 26.1 Purpose and scope. must be taken into account. (f) The ice accretion that has the most ad- (a) This part establishes require- verse effect on handling qualities may be ments for support of the continued air- used for airplane performance tests provided worthiness of and safety improvements any difference in performance is conserv- for transport category airplanes. These atively taken into account. requirements may include performing [Amdt. 25–140, 79 FR 65528, Nov. 4, 2014] assessments, developing design changes, developing revisions to In- PART 26—CONTINUED AIRWORTHI- structions for Continued Airworthiness NESS AND SAFETY IMPROVE- (ICA), and making necessary docu- MENTS FOR TRANSPORT CAT- mentation available to affected per- EGORY AIRPLANES sons. Requirements of this part that establish standards for design changes Subpart A—General and revisions to the ICA are considered airworthiness requirements. Sec. (b) Except as provided in paragraph 26.1 Purpose and scope. (c) of this section, this part applies to 26.3 [Reserved] the following persons, as specified in 26.5 Applicability table. each subpart of this part: Subpart B—Enhanced Airworthiness (1) Holders of type certificates and Program for Airplane Systems supplemental type certificates. (2) Applicants for type certificates 26.11 Electrical wiring interconnection sys- and supplemental type certificates and tems (EWIS) maintenance program. changes to those certificates (including service bulletins describing design Subpart C—Aging Airplane Safety— changes). Widespread Fatigue Damage (3) Persons seeking design approval 26.21 Limit of validity. for airplane repairs, alterations, or 26.23 Extended limit of validity. modifications that may affect airworthiness. Subpart D—Fuel Tank Flammability (4) Holders of type certificates and 26.31 Definitions. their licensees producing new air- 26.33 Holders of type certificates: Fuel tank planes. flammability. (c) An applicant for approval of a de- 26.35 Changes to type certificates affecting sign change is not required to comply fuel tank flammability. with any applicable airworthiness re- 26.37 Pending type certification projects: quirement of this part if the applicant Fuel tank flammability. 26.39 Newly produced airplanes: Fuel tank elects or is required to comply with a flammability. corresponding amendment to part 25 of this chapter that is adopted concur- Subpart E—Aging Airplane Safety—Dam- rently or after that airworthiness re- age Tolerance Data for Repairs and quirement. Alterations (d) For the purposes of this part, the word ‘‘type certificate’’ does not in- 26.41 Definitions. 26.43 Holders of and applicants for type cer- clude supplemental type certificates. tificates—Repairs. 26.45 Holders of type certificates—Alter- § 26.3 [Reserved] ations and repairs to alterations. 26.47 Holders of and applicants for a supple- § 26.5 Applicability table. mental type certificate—Alterations and Table 1 of this section provides an repairs to alterations. overview of the applicability of this 26.49 Compliance plan. part. It provides guidance in identi- AUTHORITY: 49 U.S.C. 106(g), 40113, 44701, fying what sections apply to various 44702 and 44704. types of entities. The specific applica- SOURCE: Docket No. FAA–2004–18379, 72 FR bility of each subpart and section is 63409, Nov. 8, 2007, unless otherwise noted. specified in the regulatory text.

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