Pt. 23 14 CFR Ch. I (1–1–11 Edition)

(c) Changes by persons other than the manu- Subpart B—Flight facturer. No design change by any person (other than the manufacturer who provided GENERAL the statement of conformance for the arti- 23.21 Proof of compliance. cle) is eligible for approval under this part 23.23 Load distribution limits. unless the person seeking the approval is a 23.25 Weight limits. manufacturer and applies under § 21.603(a) for 23.29 Empty weight and corresponding cen- a separate TSO authorization. Persons other ter of gravity. than a manufacturer may obtain approval 23.31 Removable ballast. for design changes under part 43 or under the 23.33 Propeller speed and pitch limits. applicable airworthiness regulations of this chapter. PERFORMANCE § 21.620 Changes in quality system. 23.45 General. After the issuance of a TSO authoriza- 23.49 Stalling period. tion— 23.51 Takeoff speeds. (a) Each change to the quality system is 23.53 Takeoff performance. subject to review by the FAA; and 23.55 Accelerate-stop distance. (b) The holder of the TSO authorization 23.57 Takeoff path. must immediately notify the FAA, in writ- 23.59 Takeoff distance and takeoff run. ing, of any change that may affect the in- 23.61 Takeoff flight path. spection, conformity, or airworthiness of its 23.63 Climb: General. article. 23.65 Climb: All engines operating. 23.66 Takeoff climb: One-engine inoperative. § 21.621 Issuance of letters of TSO design 23.67 Climb: One engine inoperative. approval: import articles. 23.69 Enroute climb/descent. (a) The FAA may issue a letter of TSO de- 23.71 Glide: Single-engine airplanes. sign approval for an article— 23.73 Reference landing approach speed. (1) Designed and manufactured in a foreign 23.75 Landing distance. country or jurisdiction subject to the export 23.77 Balked landing. provisions of an agreement with the United States for the acceptance of these articles FLIGHT CHARACTERISTICS for import; and 23.141 General. (2) For import into the United States if— (i) The State of Design certifies that the CONTROLLABILITY AND MANEUVERABILITY article has been examined, tested, and found 23.143 General. to meet the applicable TSO or the applicable 23.145 Longitudinal control. performance standards of the State of Design 23.147 Directional and lateral control. and any other performance standards the 23.149 Minimum control speed. FAA may prescribe to provide a level of safe- 23.151 Acrobatic maneuvers. ty equivalent to that provided by the TSO; 23.153 Control during landings. and 23.155 Elevator control force in maneuvers. (ii) The manufacturer has provided to the 23.157 Rate of roll. FAA one copy of the technical data required in the applicable performance standard TRIM through its State of Design. (b) The FAA issues the letter of TSO de- 23.161 Trim. sign approval that lists any deviation grant- ed under § 21.618. STABILITY 23.171 General. PART 23—AIRWORTHINESS STAND- 23.173 Static longitudinal stability. 23.175 Demonstration of static longitudinal ARDS: NORMAL, UTILITY, ACRO- stability. BATIC, AND COMMUTER CAT- 23.177 Static directional and lateral sta- EGORY AIRPLANES bility. 23.181 Dynamic stability.

SPECIAL FEDERAL AVIATION REGULATION NO. STALLS 23 23.201 Wings level stall. Subpart A—General 23.203 Turning flight and accelerated turn- ing stalls. Sec. 23.207 Stall warning. 23.1 Applicability. SPINNING 23.2 Special retroactive requirements. 23.3 Airplane categories. 23.221 Spinning.

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GROUND AND WATER HANDLING AILERONS AND SPECIAL DEVICES CHARACTERISTICS 23.455 Ailerons. 23.231 Longitudinal stability and control. 23.459 Special devices. 23.233 Directional stability and control. 23.235 Operation on unpaved surfaces. GROUND LOADS 23.237 Operation on water. 23.471 General. 23.239 Spray characteristics. 23.473 Ground load conditions and assump- tions. MISCELLANEOUS FLIGHT REQUIREMENTS 23.477 Landing gear arrangement. 23.251 Vibration and buffeting. 23.479 Level landing conditions. 23.253 High speed characteristics. 23.481 Tail down landing conditions. 23.483 One-wheel landing conditions. Subpart C—Structure 23.485 Side load conditions. 23.493 Braked roll conditions. GENERAL 23.497 Supplementary conditions for tail 23.301 Loads. wheels. 23.302 Canard or tandem wing configura- 23.499 Supplementary conditions for nose tions. wheels. 23.303 Factor of safety. 23.505 Supplementary conditions for ski- 23.305 Strength and deformation. planes. 23.307 Proof of structure. 23.507 Jacking loads. 23.509 Towing loads. FLIGHT LOADS 23.511 Ground load; unsymmetrical loads on multiple-wheel units. 23.321 General. 23.331 Symmetrical flight conditions. WATER LOADS 23.333 Flight envelope. 23.521 Water load conditions. 23.335 Design airspeeds. 23.523 Design weights and center of gravity 23.337 Limit maneuvering load factors. positions. 23.341 Gust loads factors. 23.525 Application of loads. 23.343 Design fuel loads. 23.527 Hull and main float load factors. 23.345 High lift devices. 23.529 Hull and main float landing condi- 23.347 Unsymmetrical flight conditions. tions. 23.349 Rolling conditions. 23.531 Hull and main float takeoff condi- 23.351 Yawing conditions. tion. 23.361 Engine torque. 23.533 Hull and main float bottom pressures. 23.363 Side load on engine mount. 23.535 Auxiliary float loads. 23.365 Pressurized cabin loads. 23.537 Seawing loads. 23.367 Unsymmetrical loads due to engine failure. EMERGENCY LANDING CONDITIONS 23.369 Rear lift truss. 23.371 Gyroscopic and aerodynamic loads. 23.561 General. 23.373 Speed control devices. 23.562 Emergency landing dynamic condi- tions. CONTROL SURFACE AND SYSTEM LOADS FATIGUE EVALUATION 23.391 Control surface loads. 23.393 Loads parallel to hinge line. 23.571 Metallic pressurized cabin structures. 23.395 Control system loads. 23.572 Metallic wing, empennage, and asso- 23.397 Limit control forces and torques. ciated structures. 23.399 Dual control system. 23.573 Damage tolerance and fatigue evalua- 23.405 Secondary control system. tion of structure. 23.407 Trim tab effects. 23.574 Metallic damage tolerance and fa- 23.409 Tabs. tigue evaluation of commuter category 23.415 Ground gust conditions. airplanes. 23.575 Inspections and other procedures. HORIZONTAL STABILIZING AND BALANCING SURFACES Subpart D—Design and Construction 23.421 Balancing loads. 23.601 General. 23.423 Maneuvering loads. 23.603 Materials and workmanship. 23.425 Gust loads. 23.605 Fabrication methods. 23.427 Unsymmetrical loads. 23.607 Fasteners. 23.609 Protection of structure. VERTICAL SURFACES 23.611 Accessibility provisions. 23.441 Maneuvering loads. 23.613 Material strength properties and de- 23.443 Gust loads. sign values. 23.445 Outboard fins or winglets. 23.619 Special factors.

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23.621 Casting factors. 23.787 Baggage and cargo compartments. 23.623 Bearing factors. 23.791 Passenger information signs. 23.625 Fitting factors. 23.803 Emergency evacuation. 23.627 Fatigue strength. 23.805 Flightcrew emergency exits. 23.629 Flutter. 23.807 Emergency exits. 23.811 Emergency exit marking. WINGS 23.812 Emergency lighting. 23.641 Proof of strength. 23.813 Emergency exit access. 23.815 Width of aisle. CONTROL SURFACES 23.831 Ventilation.

23.651 Proof of strength. PRESSURIZATION 23.655 Installation. 23.657 Hinges. 23.841 Pressurized cabins. 23.659 Mass balance. 23.843 Pressurization tests.

CONTROL SYSTEMS FIRE PROTECTION 23.671 General. 23.851 Fire extinguishers. 23.672 Stability augmentation and auto- 23.853 Passenger and crew compartment in- matic and power-operated systems. teriors. 23.673 Primary flight controls. 23.855 Cargo and baggage compartment fire 23.675 Stops. protection. 23.677 Trim systems. 23.859 Combustion heater fire protection. 23.679 Control system locks. 23.863 Flammable fluid fire protection. 23.681 Limit load static tests. 23.865 Fire protection of flight controls, en- 23.683 Operation tests. gine mounts, and other flight structure. 23.685 Control system details. ELECTRICAL BONDING AND LIGHTNING 23.687 Spring devices. PROTECTION 23.689 Cable systems. 23.691 Artificial stall barrier system. 23.867 Electrical bonding and protection 23.693 Joints. against lightning and static electricity. 23.697 Wing flap controls. 23.699 Wing flap position indicator. MISCELLANEOUS 23.701 Flap interconnection. 23.871 Leveling means. 23.703 Takeoff warning system. Subpart E—Powerplant LANDING GEAR 23.721 General. GENERAL 23.723 Shock absorption tests. 23.901 Installation. 23.725 Limit drop tests. 23.903 Engines. 23.726 Ground load dynamic tests. 23.904 Automatic power reserve system. 23.727 Reserve energy absorption drop test. 23.905 Propellers. 23.729 Landing gear extension and retrac- 23.907 Propeller vibration and fatigue. tion system. 23.909 Turbocharger systems. 23.731 Wheels. 23.925 Propeller clearance. 23.733 Tires. 23.929 Engine installation ice protection. 23.735 Brakes. 23.933 Reversing systems. 23.737 Skis. 23.934 Turbojet and turbofan engine thrust 23.745 Nose/tail wheel steering. reverser systems tests. 23.937 Turbopropeller-drag limiting sys- FLOATS AND HULLS tems. 23.751 Main float buoyancy. 23.939 Powerplant operating characteristics. 23.753 Main float design. 23.943 Negative acceleration. 23.755 Hulls. 23.757 Auxiliary floats. FUEL SYSTEM 23.951 General. PERSONNEL AND CARGO ACCOMMODATIONS 23.953 Fuel system independence. 23.771 Pilot compartment. 23.954 Fuel system lightning protection. 23.773 Pilot compartment view. 23.955 Fuel flow. 23.775 Windshields and windows. 23.957 Flow between interconnected tanks. 23.777 Cockpit controls. 23.959 Unusable fuel supply. 23.779 Motion and effect of cockpit controls. 23.961 Fuel system hot weather operation. 23.781 Cockpit control knob shape. 23.963 Fuel tanks: General. 23.783 Doors. 23.965 Fuel tank tests. 23.785 Seats, berths, litters, safety belts, 23.967 Fuel tank installation. and shoulder harnesses. 23.969 Fuel tank expansion space.

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23.971 Fuel tank sump. 23.1153 Propeller feathering controls. 23.973 Fuel tank filler connection. 23.1155 Turbine engine reverse thrust and 23.975 Fuel tank vents and carburetor vapor propeller pitch settings below the flight vents. regime. 23.977 Fuel tank outlet. 23.1157 Carburetor air temperature controls. 23.979 Pressure fueling systems. 23.1163 Powerplant accessories. 23.1165 Engine ignition systems. FUEL SYSTEM COMPONENTS POWERPLANT FIRE PROTECTION 23.991 Fuel pumps. 23.993 Fuel system lines and fittings. 23.1181 Designated fire zones; regions in- 23.994 Fuel system components. cluded. 23.995 Fuel valves and controls. 23.1182 Nacelle areas behind firewalls. 23.997 Fuel strainer or filter. 23.1183 Lines, fittings, and components. 23.999 Fuel system drains. 23.1189 Shutoff means. 23.1001 Fuel jettisoning system. 23.1191 Firewalls. 23.1192 Engine accessory compartment dia- OIL SYSTEM phragm. 23.1011 General. 23.1193 Cowling and nacelle. 23.1013 Oil tanks. 23.1195 Fire extinguishing systems. 23.1015 Oil tank tests. 23.1197 Fire extinguishing agents. 23.1017 Oil lines and fittings. 23.1199 Extinguishing agent containers. 23.1019 Oil strainer or filter. 23.1201 Fire extinguishing systems mate- 23.1021 Oil system drains. rials. 23.1023 Oil radiators. 23.1203 Fire detector system. 23.1027 Propeller feathering system. Subpart F—Equipment COOLING GENERAL 23.1041 General. 23.1301 Function and installation. 23.1043 Cooling tests. 23.1303 Flight and navigation instruments. 23.1045 Cooling test procedures for turbine 23.1305 Powerplant instruments. engine powered airplanes. 23.1307 Miscellaneous equipment. 23.1047 Cooling test procedures for recipro- 23.1308 High-intensity Radiated Fields cating engine powered airplanes. (HIRF) Protection. LIQUID COOLING 23.1309 Equipment, systems, and installa- tions. 23.1061 Installation. 23.1063 Coolant tank tests. INSTRUMENTS: INSTALLATION

INDUCTION SYSTEM 23.1311 Electronic display instrument sys- tems. 23.1091 Air induction system. 23.1321 Arrangement and visibility. 23.1093 Induction system icing protection. 23.1322 Warning, caution, and advisory 23.1095 Carburetor deicing fluid flow rate. lights. 23.1097 Carburetor deicing fluid system ca- 23.1323 Airspeed indicating system. pacity. 23.1325 Static pressure system. 23.1099 Carburetor deicing fluid system de- 23.1326 Pitot heat indication systems. tail design. 23.1327 Magnetic direction indicator. 23.1101 Induction air preheater design. 23.1329 Automatic pilot system. 23.1103 Induction system ducts. 23.1331 Instruments using a power source. 23.1105 Induction system screens. 23.1335 Flight director systems. 23.1107 Induction system filters. 23.1337 Powerplant instruments installa- 23.1109 Turbocharger bleed air system. tion. 23.1111 Turbine engine bleed air system. ELECTRICAL SYSTEMS AND EQUIPMENT EXHAUST SYSTEM 23.1351 General. 23.1121 General. 23.1353 Storage battery design and installa- 23.1123 Exhaust system. tion. 23.1125 Exhaust heat exchangers. 23.1357 Circuit protective devices. 23.1359 Electrical system fire protection. POWERPLANT CONTROLS AND ACCESSORIES 23.1361 Master switch arrangement. 23.1141 Powerplant controls: General. 23.1365 Electric cables and equipment. 23.1142 Auxiliary power unit controls. 23.1367 Switches. 23.1143 Engine controls. LIGHTS 23.1145 Ignition switches. 23.1147 Mixture controls. 23.1381 Instrument lights. 23.1149 Propeller speed and pitch controls. 23.1383 Taxi and landing lights.

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23.1385 Position light system installation. 23.1543 Instrument markings: General. 23.1387 Position light system dihedral an- 23.1545 Airspeed indicator. gles. 23.1547 Magnetic direction indicator. 23.1389 Position light distribution and in- 23.1549 Powerplant and auxiliary power unit tensities. instruments. 23.1391 Minimum intensities in the hori- 23.1551 Oil quantity indicator. zontal plane of position lights. 23.1553 Fuel quantity indicator. 23.1393 Minimum intensities in any vertical 23.1555 Control markings. plane of position lights. 23.1557 Miscellaneous markings and plac- 23.1395 Maximum intensities in overlapping ards. beams of position lights. 23.1559 Operating limitations placard. 23.1397 Color specifications. 23.1561 Safety equipment. 23.1399 Riding light. 23.1563 Airspeed placards. 23.1401 Anticollision light system. 23.1567 Flight maneuver placard. SAFETY EQUIPMENT AIRPLANE FLIGHT MANUAL AND APPROVED 23.1411 General. MANUAL MATERIAL 23.1415 Ditching equipment. 23.1581 General. 23.1416 Pneumatic de-icer boot system. 23.1583 Operating limitations. 23.1419 Ice protection. 23.1585 Operating procedures. MISCELLANEOUS EQUIPMENT 23.1587 Performance information. 23.1589 Loading information. 23.1431 Electronic equipment. APPENDIX A TO PART 23—SIMPLIFIED DESIGN 23.1435 Hydraulic systems. LOAD CRITERIA 23.1437 Accessories for multiengine air- APPENDIX B TO PART 23 [RESERVED] planes. APPENDIX C TO PART 23—BASIC LANDING CON- 23.1438 Pressurization and pneumatic sys- DITIONS tems. APPENDIX D TO PART 23—WHEEL SPIN-UP AND 23.1441 Oxygen equipment and supply. 23.1443 Minimum mass flow of supplemental SPRING-BACK LOADS oxygen. APPENDIX E TO PART 23 [RESERVED] 23.1445 Oxygen distribution system. APPENDIX F TO PART 23—TEST PROCEDURE 23.1447 Equipment standards for oxygen dis- APPENDIX G TO PART 23—INSTRUCTIONS FOR pensing units. CONTINUED AIRWORTHINESS 23.1449 Means for determining use of oxy- APPENDIX H TO PART 23—INSTALLATION OF AN gen. AUTOMATIC POWER RESERVE (APR) SYS- 23.1450 Chemical oxygen generators. TEM 23.1451 Fire protection for oxygen equip- APPENDIX I TO PART 23—SEAPLANE LOADS ment. APPENDIX J TO PART 23—HIRF ENVIRONMENTS 23.1453 Protection of oxygen equipment AND EQUIPMENT HIRF TEST LEVELS from rupture. AUTHORITY: 49 U.S.C. 106(g), 40113, 44701– 23.1457 Cockpit voice recorders. 44702, 44704. 23.1459 Flight data recorders. 23.1461 Equipment containing high energy SOURCE: Docket No. 4080, 29 FR 17955, Dec. rotors. 18. 1964; 30 FR 258, Jan. 9, 1965, unless other- wise noted. Subpart G—Operating Limitations and Information SPECIAL FEDERAL AVIATION REGULATION NO. 23 23.1501 General. 23.1505 Airspeed limitations. 1. Applicability. An applicant is entitled to 23.1507 Operating maneuvering speed. a type certificate in the normal category for 23.1511 Flap extended speed. a reciprocating or turbopropeller multien- 23.1513 Minimum control speed. gine powered small airplane that is to be cer- 23.1519 Weight and center of gravity. tificated to carry more than 10 occupants 23.1521 Powerplant limitations. and that is intended for use in operations 23.1522 Auxiliary power unit limitations. under Part 135 of the Federal Aviation Regu- 23.1523 Minimum flight crew. lations if he shows compliance with the ap- 23.1524 Maximum passenger seating configu- plicable requirements of Part 23 of the Fed- ration. eral Aviation Regulations, as supplemented 23.1525 Kinds of operation. or modified by the additional airworthiness 23.1527 Maximum operating altitude. requirements of this regulation. 23.1529 Instructions for Continued Air- 2. References. Unless otherwise provided, all worthiness. references in this regulation to specific sec- tions of Part 23 of the Federal Aviation Reg- MARKINGS AND PLACARDS ulations are those sections of Part 23 in ef- 23.1541 General. fect on March 30, 1967.

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FLIGHT REQUIREMENTS (2) The selected configuration for takeoff; (3) The center of gravity in the most unfa- 3. General. Compliance must be shown with vorable position; the applicable requirements of Subpart B of Part 23 of the Federal Aviation Regulations (4) The operating engine within approved in effect on March 30, 1967, as supplemented operating limitation; and or modified in sections 4 through 10 of this (5) Takeoff data based on smooth, dry, regulation. hard-surface runway. (b) Takeoff speeds. (1) The decision speed V1 PERFORMANCE is the calibrated airspeed on the ground at which, as a result of engine failure or other 4. General. (a) Unless otherwise prescribed reasons, the pilot is assumed to have made a in this regulation, compliance with each ap- decision to continue or discontinue the take- plicable performance requirement in sections off. The speed V1 must be selected by the ap- 4 through 7 of this regulation must be shown plicant but may not be less than— for ambient atmospheric conditions and still (i) 1.10 V ; air. s1 (ii) 1.10 V ; (b) The performance must correspond to MC (iii) A speed that permits acceleration to the propulsive thrust available under the V and stop in accordance with paragraph (c) particular ambient atmospheric conditions 1 allowing credit for an overrun distance equal and the particular flight condition. The to that required to stop the airplane from a available propulsive thrust must correspond ground speed of 35 knots utilizing maximum to engine power or thrust, not exceeding the braking; or approved power or thrust less— (1) Installation losses; and (iv) A speed at which the airplane can be (2) The power or equivalent thrust ab- rotated for takeoff and shown to be adequate sorbed by the accessories and services appro- to safely continue the takeoff, using normal priate to the particular ambient atmospheric piloting skill, when the critical engine is conditions and the particular flight condi- suddenly made inoperative. tion. (2) Other essential takeoff speeds necessary (c) Unless otherwise prescribed in this reg- for safe operation of the airplane must be de- ulation, the applicant must select the take- termined and shown in the Airplane Flight off, en route, and landing configurations for Manual. the airplane. (c) Accelerate-stop distance. (1) The accel- (d) The airplane configuration may vary erate-stop distance is the sum of the dis- with weight, altitude, and temperature, to tances necessary to— the extent they are compatible with the op- (i) Accelerate the airplane from a standing erating procedures required by paragraph (e) start to V1; and of this section. (ii) Decelerate the airplane from V1 to a (e) Unless otherwise prescribed in this reg- speed not greater than 35 knots, assuming ulation, in determining the critical engine that in the case of engine failure, failure of inoperative takeoff performance, the accel- the critical engine is recognized by the pilot erate-stop distance, takeoff distance, at the speed V1. The landing gear must re- changes in the airplane’s configuration, main in the extended position and maximum speed, power, and thrust, must be made in braking may be utilized during deceleration. accordance with procedures established by (2) Means other than wheel brakes may be the applicant for operation in service. used to determine the accelerate-stop dis- (f) Procedures for the execution of balked tance if that means is available with the landings must be established by the appli- critical engine inoperative and— cant and included in the Airplane Flight (i) Is safe and reliable; Manual. (ii) Is used so that consistent results can (g) The procedures established under para- be expected under normal operating condi- graphs (e) and (f) of this section must— tions; and (1) Be able to be consistently executed in (iii) Is such that exceptional skill is not re- service by a crew of average skill; quired to control the airplane. (2) Use methods or devices that are safe (d) All engines operating takeoff distance. and reliable; and The all engine operating takeoff distance is (3) Include allowance for any time delays, the horizontal distance required to takeoff in the execution of the procedures, that may and climb to a height of 50 feet above the reasonably be expected in service. takeoff surface according to procedures in 5. Takeoff—(a) General. The takeoff speeds FAR 23.51(a). described in paragraph (b), the accelerate- (e) One-engine-inoperative takeoff. The max- stop distance described in paragraph (c), and imum weight must be determined for each the takeoff distance described in paragraph altitude and temperature within the oper- (d), must be determined for— ational limits established for the airplane, at (1) Each weight, altitude, and ambient which the airplane has takeoff capability temperature within the operational limits after failure of the critical engine at or selected by the applicant; above V1 determined in accordance with 187

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paragraph (b) of this section. This capability maintain trim at a speed greater than VMO/ may be established— MMO: (1) By demonstrating a measurably posi- (1) In the approach conditions specified in tive rate of climb with the airplane in the FAR 23.161(c)(3) through (5), except that in- takeoff configuration, landing gear extended; stead of the speeds specified therein, trim or must be maintained with a stick force of not (2) By demonstrating the capability of more than 10 pounds down to a speed used in maintaining flight after engine failure uti- showing compliance with section 7 of this lizing procedures prescribed by the appli- regulation or 1.4 Vs1 whichever is lower. cant. (2) In level flight at any speed from VH or 6. Climb—(a) Landing climb: All-engines-oper- VMO/MMO, whichever is lower, to either Vx or ating. The maximum weight must be deter- 1.4 Vs1, with the landing gear and wing flaps mined with the airplane in the landing con- retracted. figuration, for each altitude, and ambient temperature within the operational limits STABILITY established for the airplane and with the most unfavorable center of gravity and out- 9. Static longitudinal stability. (a) In showing of-ground effect in free air, at which the compliance with the provisions of FAR steady gradient of climb will not be less than 23.175(b) and with paragraph (b) of this sec- 3.3 percent, with: tion, the airspeed must return to within ±71⁄2 (1) The engines at the power that is avail- percent of the trim speed. able 8 seconds after initiation of movement (b) Cruise stability. The stick force curve of the power or thrust controls from the must have a stable slope for a speed range of mimimum flight idle to the takeoff position. ±50 knots from the trim speed except that (2) A climb speed not greater than the ap- the speeds need not exceed VFC/MFC or be proach speed established under section 7 of less than 1.4 Vs1. This speed range will be this regulation and not less than the greater considered to begin at the outer extremes of of 1.05MC or 1.10VS1. the friction band and the stick force may not (b) En route climb, one-engine-inoperative. (1) exceed 50 pounds with— the maximum weight must be determined (i) Landing gear retracted; with the airplane in the en route configura- (ii) Wing flaps retracted; tion, the critical engine inoperative, the re- (iii) The maximum cruising power as se- maining engine at not more than maximum lected by the applicant as an operating limi- continuous power or thrust, and the most tation for turbine engines or 75 percent of unfavorable center of gravity, at which the maximum continuous power for recipro- gradient at climb will be not less than— cating engines except that the power need (i) 1.2 percent (or a gradient equivalent to not exceed that required at VMO/MMO: 0.20 Vso2, if greater) at 5,000 feet and an ambi- (iv) Maximum takeoff weight; and ° ent temperature of 41 F. or (v) The airplane trimmed for level flight (ii) 0.6 percent (or a gradient equivalent to with the power specified in subparagraph 0.01 V 2, if greater) at 5,000 feet and ambient so (iii) of this paragraph. temperature of 81 °F. (2) The minimum climb gradient specified VFC/MFC may not be less than a speed in subdivisions (i) and (ii) of subparagraph (1) midway between VMO/MMO and VDF/MDF, ex- of this paragraph must vary linearly between cept that, for altitudes where Mach number 41 °F. and 81 °F. and must change at the same is the limiting factor, MFC need not exceed rate up to the maximum operational tem- the Mach number at which effective speed perature approved for the airplane. warning occurs. 7. Landing. The landing distance must be (c) Climb stability. For turbopropeller powered determined for standard atmosphere at each airplanes only. In showing compliance with weight and altitude in accordance with FAR FAR 23.175(a), an applicant must in lieu of 23.75(a), except that instead of the gliding ap- the power specified in FAR 23.175(a)(4), use proach specified in FAR 23.75(a)(1), the land- the maximum power or thrust selected by ing may be preceded by a steady approach the applicant as an operating limitation for down to the 50-foot height at a gradient of use during climb at the best rate of climb descent not greater than 5.2 percent (3°) at a speed except that the speed need not be less than 1.4 V . calibrated airspeed not less than 1.3s1. s1

TRIM STALLS 8. Trim—(a) Lateral and directional trim. The 10. Stall warning. If artificial stall warning airplane must maintain lateral and direc- is required to comply with the requirements tional trim in level flight at a speed of Vh of FAR 23.207, the warning device must give or VMO/MMO, whichever is lower, with land- clearly distinguishable indications under ex- ing gear and wing flaps retracted. pected conditions of flight. The use of a vis- (b) Longitudinal trim. The airplane must ual warning device that requires the atten- maintain longitudinal trim during the fol- tion of the crew within the cockpit is not ac- lowing conditions, except that it need not ceptable by itself.

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CONTROL SYSTEMS tion will be determined assuming an engine failure at the mimimum value of V . 11. Electric trim tabs. The airplane must 1 (c) The airspeed error of the installation meet the requirements of FAR 23.677 and in excluding the instrument calibration error, addition it must be shown that the airplane must not exceed 3 percent or 5 knots which- is safely controllable and that a pilot can ever is greater, throughout the speed range perform all the maneuvers and operations from V to 1.3S with flaps retracted and necessary to effect a safe landing following MO 1 from 1.3 VS to V with flaps in the land- any probable electric trim tab runaway O FE ing position. which might be reasonably expected in serv- (d) Information showing the relationship ice allowing for appropriate time delay after between IAS and CAS must be shown in the pilot recognition of the runaway. This dem- Airplane Flight Manual. onstration must be conducted at the critical 14. Static air vent system. The static air vent airplane weights and center of gravity posi- system must meet the requirements of FAR tions. 23.1325. The altimeter system calibration INSTRUMENTS: INSTALLATION must be determined and shown in the Air- plane Flight Manual. 12. Arrangement and visibility. Each instru- ment must meet the requirements of FAR OPERATING LIMITATIONS AND INFORMATION 23.1321 and in addition— 15. Maximum operating limit speed VMO/MMO. (a) Each flight, navigation, and powerplant Instead of establishing operating limitations instrument for use by any pilot must be based on VME and VNO, the applicant must plainly visible to him from his station with establish a maximum operating limit speed the minimum practicable deviation from his VMO/MMO in accordance with the following: normal position and line of vision when he is (a) The maximum operating limit speed looking forward along the flight path. must not exceed the design cruising speed Vc (b) The flight instruments required by FAR and must be sufficiently below VD/MD or 23.1303 and by the applicable operating rules VDF/MDF to make it highly improbable that must be grouped on the instrument panel the latter speeds will be inadvertently ex- and centered as nearly as practicable about ceeded in flight. the vertical plane of each pilot’s forward vi- (b) The speed Vmo must not exceed 0.8 VD/ sion. In addition— MD or 0.8 VDF/MDF unless flight dem- (1) The instrument that most effectively onstrations involving upsets as specified by indicates the attitude must be on the panel the Administrator indicates a lower speed in the top center position; margin will not result in speeds exceeding (2) The instrument that most effectively VD/MD or VDF. Atmospheric variations, hor- indicates airspeed must be adjacent to and izontal gusts, and equipment errors, and air- directly to the left of the instrument in the frame production variations will be taken top center position; into account. (3) The instrument that most effectively 16. Minimum flight crew. In addition to indicates altitude must be adjacent to and meeting the requirements of FAR 23.1523, the directly to the right of the instrument in the applicant must establish the minimum num- top center position; and ber and type of qualified flight crew per- (4) The instrument that most effectively sonnel sufficient for safe operation of the indicates direction of flight must be adjacent airplane considering— to and directly below the instrument in the (a) Each kind of operation for which the top center position. applicant desires approval; 13. Airspeed indicating system. Each airspeed (b) The workload on each crewmember con- indicating system must meet the require- sidering the following: ments of FAR 23.1323 and in addition— (1) Flight path control. (a) Airspeed indicating instruments must (2) Collision avoidance. be of an approved type and must be cali- (3) Navigation. brated to indicate true airspeed at sea level (4) Communications. in the standard atmosphere with a (5) Operation and monitoring of all essen- mimimum practicable instrument calibra- tial aircraft systems. tion error when the corresponding pilot and (6) Command decisions; and static pressures are supplied to the instru- (c) The accessibility and ease of operation ments. of necessary controls by the appropriate (b) The airspeed indicating system must be crewmember during all normal and emer- calibrated to determine the system error, gency operations when at his flight station. i.e., the relation between IAS and CAS, in 17. Airspeed indicator. The airspeed indi- flight and during the accelerate takeoff cator must meet the requirements of FAR ground run. The ground run calibration must 23.1545 except that, the airspeed notations be obtained between 0.8 of the mimimum and markings in terms of VNO and VNE value of V1 and 1.2 times the maximum value must be replaced by the VMO/MMO nota- of V1, considering the approved ranges of al- tions. The airspeed indicator markings must titude and weight. The ground run calibra- be easily read and understood by the pilot. A

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placard adjacent to the airspeed indicator is (c) The performance information (deter- an acceptable means of showing compliance mined by extrapolation and computed for the with the requirements of FAR 23.1545(c). range of weights between the maximum landing and takeoff weights) for— AIRPLANE FLIGHT MANUAL (1) Climb in the landing configuration; and 18. General. The Airplane Flight Manual (2) Landing distance. must be prepared in accordance with the re- (d) Procedure established under section 4 of quirements of FARs 23.1583 and 23.1587, and this regulation related to the limitations in addition the operating limitations and and information required by this section in performance information set forth in sec- the form of guidance material including any tions 19 and 20 must be included. relevant limitations or information. (e) An explanation of significant or un- 19. Operating limitations. The Airplane usual flight or ground handling characteris- Flight Manual must include the following tics of the airplane. limitations— (f) Airspeeds, as indicated airspeeds, cor- (a) Airspeed limitations. (1) The maximum responding to those determined for takeoff operating limit speed V /M and a state- MO MO in accordance with section 5(b). ment that this speed limit may not be delib- 21. Maximum operating altitudes. The max- erately exceeded in any regime of flight imum operating altitude to which operation (climb, cruise, or descent) unless a higher is permitted, as limited by flight, structural, speed is authorized for flight test or pilot powerplant, functional, or equipment char- training; acteristics, must be specified in the Airplane (2) If an airspeed limitation is based upon Flight Manual. compressibility effects, a statement to this 22. Stowage provision for Airplane Flight effect and information as to any symptoms, Manual. Provision must be made for stowing the probable behavior of the airplane, and the Airplane Flight Manual in a suitable the recommended recovery procedures; and fixed container which is readily accessible to (3) The airspeed limits, shown in terms of the pilot. VMO/MMO instead of VNO and VNE. 23. Operating procedures. Procedures for re- (b) Takeoff weight limitations. The max- starting turbine engines in flight (including imum takeoff weight for each airport ele- the effects of altitude) must be set forth in vation, ambient temperature, and available the Airplane Flight Manual. takeoff runway length within the range se- lected by the applicant. This weight may not AIRFRAME REQUIREMENTS exceed the weight at which: (1) The all-engine operating takeoff dis- FLIGHT LOADS tance determined in accordance with section 24. Engine torque. (a) Each turbopropeller 5(d) or the accelerate-stop distance deter- engine mount and its supporting structure mined in accordance with section 5(c), which must be designed for the torque effects of— ever is greater, is equal to the available run- (1) The conditions set forth in FAR way length; 23.361(a). (2) The airplane complies with the one-en- (2) The limit engine torque corresponding gine-inoperative takeoff requirements speci- to takeoff power and propeller speed, multi- fied in section 5(e); and plied by a factor accounting for propeller (3) The airplane complies with the one-en- control system malfunction, including quick gine-inoperative en route climb require- feathering action, simultaneously with 1 g ments specified in section 6(b), assuming level flight loads. In the absence of a ration- that a standard temperature lapse rate ex- al analysis, a factor of 1.6 must be used. ists from the airport elevation to the alti- (b) The limit torque is obtained by multi- tude of 5,000 feet, except that the weight may plying the mean torque by a factor of 1.25. not exceed that corresponding to a tempera- 25. Turbine engine gyroscopic loads. Each ture of 41 °F at 5,000 feet. turbopropeller engine mount and its sup- 20. Performance information. The Airplane porting structure must be designed for the Flight Manual must contain the performance gyroscopic loads that result, with the en- information determined in accordance with gines at maximum continuous r.p.m., under the provisions of the performance require- either— ments of this regulation. The information (a) The conditions prescribed in FARs must include the following: 23.351 and 23.423; or (a) Sufficient information so that the take- (b) All possible combinations of the fol- off weight limits specified in section 19(b) lowing: can be determined for all temperatures and (1) A yaw velocity of 2.5 radius per second. altitudes within the operation limitations (2) A pitch velocity of 1.0 radians per sec- selected by the applicant. ond. (b) The conditions under which the per- (3) A normal load factor of 2.5. formance information was obtained, includ- (4) Maximum continuous thrust. ing the airspeed at the 50-foot height used to 26. Unsymmetrical loads due to engine failure. determine landing distances. (a) Turbopropeller powered airplanes must

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be designed for the unsymmetrical loads re- FATIGUE EVALUATION sulting from the failure of the critical engine 28. Fatigue evaluation of wing and associated including the following conditions in com- structure. Unless it is shown that the struc- bination with a single malfunction of the ture, operating stress levels, materials, and propeller drag limiting system, considering expected use are comparable from a fatigue the probable pilot corrective action on the standpoint to a similar design which has had flight controls. substantial satisfactory service experience, (1) At speeds between VMC and VD, the the strength, detail design, and the fabrica- loads resulting from power failure because of tion of those parts of the wing, wing carry- fuel flow interruption are considered to be through, and attaching structure whose fail- limit loads. ure would be catastrophic must be evaluated (2) At speeds between VMC and VC, the under either— loads resulting from the disconnection of the (a) A fatigue strength investigation in engine compressor from the turbine or from which the structure is shown by analysis, loss of the turbine blades are considered to tests, or both to be able to withstand the re- be ultimate loads. peated loads of variable magnitude expected (3) The time history of the thrust decay in service; or and drag buildup occurring as a result of the (b) A fail-safe strength investigation in prescribed engine failures must be substan- which it is shown by analysis, tests, or both tiated by test or other data applicable to the that catastrophic failure of the structure is particular engine-propeller combination. not probable after fatigue, or obvious partial (4) The timing and magnitude of the prob- failure, of a principal structural element, able pilot corrective action must be conserv- and that the remaining structure is able to atively estimated, considering the character- withstand a static ultimate load factor of 75 istics of the particular engine-propeller-air- percent of the critical limit load factor at Vc. plane combination. These loads must be multiplied by a factor of 1.15 unless the dynamic effects of failure (b) Pilot corrective action may be assumed under static load are otherwise considered. to be initiated at the time maximum yawing velocity is reached, but not earlier than two DESIGN AND CONSTRUCTION seconds after the engine failure. The mag- nitude of the corrective action may be based 29. Flutter. For Multiengine turbopropeller on the control forces specified in FAR 23.397 powered airplanes, a dynamic evaluation except that lower forces may be assumed must be made and must include— where it is shown by analysis or test that (a) The significant elastic, inertia, and aer- these forces can control the yaw and roll re- odynamic forces associated with the rota- sulting from the prescribed engine failure tions and displacements of the plane of the propeller; and conditions. (b) Engine-propeller-nacelle stiffness and GROUND LOADS damping variations appropriate to the par- ticular configuration. 27. Dual wheel landing gear units. Each dual wheel landing gear unit and its supporting LANDING GEAR structure must be shown to comply with the 30. Flap operated landing gear warning de- following: vice. Airplanes having retractable landing (a) Pivoting. The airplane must be assumed gear and wing flaps must be equipped with a to pivot about one side of the main gear with warning device that functions continuously the brakes on that side locked. The limit when the wing flaps are extended to a flap vertical load factor must be 1.0 and the coef- position that activates the warning device to ficient of friction 0.8. This condition need give adequate warning before landing, using apply only to the main gear and its sup- normal landing procedures, if the landing porting structure. gear is not fully extended and locked. There (b) Unequal tire inflation. A 60–40 percent may not be a manual shut off for this warn- distribution of the loads established in ac- ing device. The flap position sensing unit cordance with FAR 23.471 through FAR 23.483 may be installed at any suitable location. must be applied to the dual wheels. The system for this device may use any part (c) Flat tire. (1) Sixty percent of the loads of the system (including the aural warning specified in FAR 23.471 through FAR 23.483 device) provided for other landing gear warn- must be applied to either wheel in a unit. ing devices. (2) Sixty percent of the limit drag and side loads and 100 percent of the limit vertical PERSONNEL AND CARGO ACCOMMODATIONS load established in accordance with FARs 31. Cargo and baggage compartments. Cargo 23.493 and 23.485 must be applied to either and baggage compartments must be designed wheel in a unit except that the vertical load to meet the requirements of FAR 23.787 (a) need not exceed the maximum vertical load and (b), and in addition means must be pro- in paragraph (c)(1) of this section. vided to protect passengers from injury by

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the contents of any cargo or baggage com- Minimum main passenger aisle partment when the ultimate forward inertia width Total seating capacity force is 9g. Less than 25 25 inches and 32. Doors and exits. The airplane must meet inches from floor more from floor the requirements of FAR 23.783 and FAR 23.807 (a)(3), (b), and (c), and in addition: 10 through 23 ...... 9 inches ...... 15 inches. (a) There must be a means to lock and safeguard each external door and exit MISCELLANEOUS against opening in flight either inadvert- 33. Lightning strike protection. Parts that ently by persons, or as a result of mechan- are electrically insulated from the basic air- ical failure. Each external door must be op- frame must be connected to it through light- erable from both the inside and the outside. ning arrestors unless a lightning strike on (b) There must be means for direct visual the insulated part— inspection of the locking mechanism by (a) Is improbable because of shielding by crewmembers to determine whether external other parts; or doors and exits, for which the initial opening (b) Is not hazardous. movement is outward, are fully locked. In 34. Ice protection. If certification with ice protection provisions is desired, compliance addition, there must be a visual means to with the following requirements must be signal to crewmembers when normally used shown: external doors are closed and fully locked. (a) The recommended procedures for the (c) The passenger entrance door must qual- use of the ice protection equipment must be ify as a floor level emergency exit. Each ad- set forth in the Airplane Flight Manual. ditional required emergency exit except floor (b) An analysis must be performed to es- level exits must be located over the wing or tablish, on the basis of the airplane’s oper- must be provided with acceptable means to ational needs, the adequacy of the ice protec- assist the occupants in descending to the tion system for the various components of ground. In addition to the passenger en- the airplane. In addition, tests of the ice pro- trance door: tection system must be conducted to dem- (1) For a total seating capacity of 15 or onstrate that the airplane is capable of oper- less, an emergency exit as defined in FAR ating safely in continuous maximum and 23.807(b) is required on each side of the cabin. intermittent maximum icing conditions as (2) For a total seating capacity of 16 described in FAR 25, appendix C. through 23, three emergency exits as defined (c) Compliance with all or portions of this section may be accomplished by reference, in 23.807(b) are required with one on the same where applicable because of similarity of the side as the door and two on the side opposite designs, to analysis and tests performed by the door. the applicant for a type certificated model. (d) An evacuation demonstration must be 35. Maintenance information. The applicant conducted utilizing the maximum number of must make available to the owner at the occupants for which certification is desired. time of delivery of the airplane the informa- It must be conducted under simulated night tion he considers essential for the proper conditions utilizing only the emergency maintenance of the airplane. That informa- exits on the most critical side of the aircraft. tion must include the following: The participants must be representative of (a) Description of systems, including elec- average airline passengers with no prior trical, hydraulic, and fuel controls. practice or rehearsal for the demonstration. (b) Lubrication instructions setting forth Evacuation must be completed within 90 sec- the frequency and the lubricants and fluids onds. which are to be used in the various systems. (e) Each emergency exit must be marked (c) Pressures and electrical loads applica- with the word ‘‘Exit’’ by a sign which has ble to the various systems. white letters 1 inch high on a red back- (d) Tolerances and adjustments necessary ground 2 inches high, be self-illuminated or for proper functioning. independently internally electrically illumi- (e) Methods of leveling, raising, and tow- ing. nated, and have a minimum luminescence (f) Methods of balancing control surfaces. (brightness) of at least 160 microlamberts. (g) Identification of primary and secondary The colors may be reversed if the passenger structures. compartment illumination is essentially the (h) Frequency and extent of inspections same. necessary to the proper operation of the air- (f) Access to window type emergency exits plane. must not be obstructed by seats or seat (i) Special repair methods applicable to the backs. airplane. (g) The width of the main passenger aisle (j) Special inspection techniques, including at any point between seats must equal or ex- those that require X-ray, ultrasonic, and ceed the values in the following table. magnetic particle inspection.

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(k) List of special tools. ground operation must be designed so that no single failure or malfunction of the sys- PROPULSION tem will result in unwanted reverse thrust under any expected operating condition. GENERAL Failure of structural elements need not be 36. Vibration characteristics. For turbo- considered if the probability of this kind of propeller powered airplanes, the engine in- failure is extremely remote. stallation must not result in vibration char- (b) Turbopropeller reversing systems in- acteristics of the engine exceeding those es- tended for in-flight use must be designed so tablished during the type certification of the that no unsafe condition will result during engine. normal operation of the system, or from any 37. In-flight restarting of engine. If the en- failure (or reasonably likely combination of gine on turbopropeller powered airplanes failures) of the reversing system, under any cannot be restarted at the maximum cruise anticipated condition of operation of the air- altitude, a determination must be made of plane. Failure of structural elements need the altitude below which restarts can be con- not be considered if the probability of this sistently accomplished. Restart information kind of failure is extremely remote. must be provided in the Airplane Flight (c) Compliance with this section may be Manual. shown by failure analysis, testing, or both 38. Engines—(a) For turbopropeller powered for propeller systems that allow propeller airplanes. The engine installation must com- blades to move from the flight low-pitch po- ply with the following requirements: sition to a position that is substantially less (1) Engine isolation. The powerplants must than that at the normal flight low-pitch stop be arranged and isolated from each other to position. The analysis may include or be sup- allow operation, in at least one configura- ported by the analysis made to show compli- tion, so that the failure or malfunction of ance with the type certification of the pro- any engine, or of any system that can affect peller and associated installation compo- the engine, will not— nents. Credit will be given for pertinent (i) Prevent the continued safe operation of analysis and testing completed by the engine the remaining engines; or and propeller manufacturers. (ii) Require immediate action by any crew- member for continued safe operation. 40. Turbopropeller drag-limiting systems. Tur- (2) Control of engine rotation. There must be bopropeller drag-limiting systems must be a means to individually stop and restart the designed so that no single failure or malfunc- rotation of any engine in flight except that tion of any of the systems during normal or engine rotation need not be stopped if con- emergency operation results in propeller tinued rotation could not jeopardize the safe- drag in excess of that for which the airplane ty of the airplane. Each component of the was designed. Failure of structural elements stopping and restarting system on the engine of the drag-limiting systems need not be con- side of the firewall, and that might be ex- sidered if the probability of this kind of fail- posed to fire, must be at least fire resistant. ure is extremely remote. If hydraulic propeller feathering systems are 41. Turbine engine powerplant operating used for this purpose, the feathering lines characteristics. For turbopropeller powered must be at least fire resistant under the op- airplanes, the turbine engine powerplant op- erating conditions that may be expected to erating characteristics must be investigated exist during feathering. in flight to determine that no adverse char- (3) Engine speed and gas temperature control acteristics (such as stall, surge, or flameout) devices. The powerplant systems associated are present to a hazardous degree, during with engine control devices, systems, and in- normal and emergency operation within the strumentation must provide reasonable as- range of operating limitations of the air- surance that those engine operating limita- plane and of the engine. tions that adversely affect turbine rotor 42. Fuel flow. (a) For turbopropeller pow- structural integrity will not be exceeded in ered airplanes— service. (1) The fuel system must provide for con- (b) For reciprocating-engine powered air- tinuous supply of fuel to the engines for nor- planes. To provide engine isolation, the pow- mal operation without interruption due to erplants must be arranged and isolated from depletion of fuel in any tank other than the each other to allow operation, in at least one main tank; and configuration, so that the failure or malfunc- (2) The fuel flow rate for turbopropeller en- tion of any engine, or of any system that can gine fuel pump systems must not be less affect that engine, will not— than 125 percent of the fuel flow required to (1) Prevent the continued safe operation of develop the standard sea level atmospheric the remaining engines; or conditions takeoff power selected and in- (2) Require immediate action by any crew- cluded as an operating limitation in the Air- member for continued safe operation. plane Flight Manual. 39. Turbopropeller reversing systems. (a) Tur- (b) For reciprocating engine powered air- bopropeller reversing systems intended for planes, it is acceptable for the fuel flow rate

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for each pump system (main and reserve sup- (b) Temperatures must be stabilized under ply) to be 125 percent of the takeoff fuel con- the conditions from which entry is made into sumption of the engine. each stage of flight being investigated unless the entry condition is not one during which FUEL SYSTEM COMPONENTS component and engine fluid temperatures 43. Fuel pumps. For turbopropeller powered would stabilize, in which case, operation airplanes, a reliable and independent power through the full entry condition must be source must be provided for each pump used conducted before entry into the stage of with turbine engines which do not have pro- flight being investigated in order to allow visions for mechanically driving the main temperatures to reach their natural levels at pumps. It must be demonstrated that the the time of entry. The takeoff cooling test pump installations provide a reliability and must be preceded by a period during which durability equivalent to that provided by the powerplant component and engine fluid FAR 23.991(a). temperatures are stabilized with the engines at ground idle. 44. Fuel strainer or filter. For turbopropeller (c) Cooling tests for each stage of flight powered airplanes, the following apply: must be continued until— (a) There must be a fuel strainer or filter (1) The component and engine fluid tem- between the tank outlet and the fuel meter- peratures stabilize; ing device of the engine. In addition, the fuel (2) The stage of flight is completed; or strainer or filter must be— (3) An operating limitation is reached. (1) Between the tank outlet and the en- gine-driven positive displacement pump INDUCTION SYSTEM inlet, if there is an engine-driven positive 47. Air induction. For turbopropeller pow- displacement pump; ered airplanes— (2) Accessible for drainage and cleaning (a) There must be means to prevent haz- and, for the strainer screen, easily remov- ardous quantities of fuel leakage or overflow able; and from drains, vents, or other components of (3) Mounted so that its weight is not sup- flammable fluid systems from entering the ported by the connecting lines or by the engine intake system; and inlet or outlet connections of the strainer or (b) The air inlet ducts must be located or filter itself. protected so as to minimize the ingestion of (b) Unless there are means in the fuel sys- foreign matter during takeoff, landing, and tem to prevent the accumulation of ice on taxiing. the filter, there must be means to automati- 48. Induction system icing protection. For cally maintain the fuel flow if ice-clogging of turbopropeller powered airplanes, each tur- the filter occurs; and bine engine must be able to operate through- (c) The fuel strainer or filter must be of out its flight power range without adverse adequate capacity (with respect to operating effect on engine operation or serious loss of limitations established to insure proper serv- power or thrust, under the icing conditions ice) and of appropriate mesh to insure proper specified in appendix C of FAR 25. In addi- engine operation, with the fuel contaminated tion, there must be means to indicate to ap- to a degree (with respect to particle size and propriate flight crewmembers the func- density) that can be reasonably expected in tioning of the powerplant ice protection sys- service. The degree of fuel filtering may not tem. be less than that established for the engine 49. Turbine engine bleed air systems. Turbine type certification. engine bleed air systems of turbopropeller 45. Lightning strike protection. Protection powered airplanes must be investigated to must be provided against the ignition of determine— flammable vapors in the fuel vent system (a) That no hazard to the airplane will re- due to lightning strikes. sult if a duct rupture occurs. This condition must consider that a failure of the duct can COOLING occur anywhere between the engine port and 46. Cooling test procedures for turbopropeller the airplane bleed service; and powered airplanes. (a) Turbopropeller powered (b) That if the bleed air system is used for airplanes must be shown to comply with the direct cabin pressurization, it is not possible requirements of FAR 23.1041 during takeoff, for hazardous contamination of the cabin air climb en route, and landing stages of flight system to occur in event of lubrication sys- that correspond to the applicable perform- tem failure. ance requirements. The cooling test must be EXHAUST SYSTEM conducted with the airplane in the configu- ration and operating under the conditions 50. Exhaust system drains. Turbopropeller that are critical relative to cooling during engine exhaust systems having low spots or each stage of flight. For the cooling tests a pockets must incorporate drains at such lo- temperature is ‘‘stabilized’’ when its rate of cations. These drains must discharge clear of change is less than 2 °F. per minute. the airplane in normal and ground attitudes

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to prevent the accumulation of fuel after the other region where it would create addi- failure of an attempted engine start. tional hazards. 57. Flammable fluid fire protection. If flam- POWERPLANT CONTROLS AND ACCESSORIES mable fluids or vapors might be liberated by the leakage of fluid systems in areas other 51. Engine controls. If throttles or power le- than engine compartments, there must be vers for turbopropeller powered airplanes are means to— such that any position of these controls will (a) Prevent the ignition of those fluids or reduce the fuel flow to the engine(s) below vapors by any other equipment; or that necessary for satisfactory and safe idle (b) Control any fire resulting from that ig- operation of the engine while the airplane is nition. in flight, a means must be provided to pre- vent inadvertent movement of the control EQUIPMENT into this position. The means provided must incorporate a positive lock or stop at this 58. Powerplant instruments. (a) The fol- idle position and must require a separate and lowing are required for turbopropeller air- distinct operation by the crew to displace planes: the control from the normal engine oper- (1) The instruments required by FAR ating range. 23.1305 (a)(1) through (4), (b)(2) and (4). (2) A gas temperature indicator for each 52. Reverse thrust controls. For turbo- engine. propeller powered airplanes, the propeller re- (3) Free air temperature indicator. verse thrust controls must have a means to (4) A fuel flowmeter indicator for each en- prevent their inadvertent operation. The gine. means must have a positive lock or stop at (5) Oil pressure warning means for each en- the idle position and must require a separate gine. and distinct operation by the crew to dis- (6) A torque indicator or adequate means place the control from the flight regime. for indicating power output for each engine. 53. Engine ignition systems. Each turbo- (7) Fire warning indicator for each engine. propeller airplane ignition system must be (8) A means to indicate when the propeller considered an essential electrical load. blade angle is below the low-pitch position 54. Powerplant accessories. The powerplant corresponding to idle operation in flight. accessories must meet the requirements of (9) A means to indicate the functioning of FAR 23.1163, and if the continued rotation of the ice protection system for each engine. any accessory remotely driven by the engine (b) For turbopropeller powered airplanes, is hazardous when malfunctioning occurs, the turbopropeller blade position indicator there must be means to prevent rotation must begin indicating when the blade has without interfering with the continued oper- moved below the flight low-pitch position. ation of the engine. (c) The following instruments are required for reciprocating-engine powered airplanes: POWERPLANT FIRE PROTECTION (1) The instruments required by FAR 55. Fire detector system. For turbopropeller 23.1305. powered airplanes, the following apply: (2) A cylinder head temperature indicator (a) There must be a means that ensures for each engine. (3) A manifold pressure indicator for each prompt detection of fire in the engine com- engine. partment. An overtemperature switch in each engine cooling air exit is an acceptable SYSTEMS AND EQUIPMENTS method of meeting this requirement. (b) Each fire detector must be constructed GENERAL and installed to withstand the vibration, in- 59. Function and installation. The systems ertia, and other loads to which it may be and equipment of the airplane must meet the subjected in operation. requirements of FAR 23.1301, and the fol- (c) No fire detector may be affected by any lowing: oil, water, other fluids, or fumes that might (a) Each item of additional installed equip- be present. ment must— (d) There must be means to allow the flight (1) Be of a kind and design appropriate to crew to check, in flight, the functioning of its intended function; each fire detector electric circuit. (2) Be labeled as to its identification, func- (e) Wiring and other components of each tion, or operating limitations, or any appli- fire detector system in a fire zone must be at cable combination of these factors, unless least fire resistant. misuse or inadvertent actuation cannot cre- 56. Fire protection, cowling and nacelle skin. ate a hazard; For reciprocating engine powered airplanes, (3) Be installed according to limitations the engine cowling must be designed and specified for that equipment; and constructed so that no fire originating in the (4) Function properly when installed. engine compartment can enter, either (b) Systems and installations must be de- through openings or by burn through, any signed to safeguard against hazards to the

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aircraft in the event of their malfunction or and collective disconnection of the electrical failure. power sources from the system; and (c) Where an installation, the functioning (4) There are means to indicate to appro- of which is necessary in showing compliance priate crewmembers the generating system with the applicable requirements, requires a quantities essential for the safe operation of power supply, such installation must be con- the system, including the voltage and cur- sidered an essential load on the power sup- rent supplied by each generator. ply, and the power sources and the distribu- 62. Electrical equipment and installation. tion system must be capable of supplying the Electrical equipment controls, and wiring following power loads in probable operation must be installed so that operation of any combinations and for probable durations: one unit or system of units will not ad- (1) All essential loads after failure of any versely affect the simultaneous operation of prime mover, power converter, or energy to the safe operation. storage device. 63. Distribution system. (a) For the purpose (2) All essential loads after failure of any of complying with this section, the distribu- one engine on two-engine airplanes. tion system includes the distribution busses, (3) In determining the probable operating their associated feeders and each control and combinations and durations of essential protective device. loads for the power failure conditions de- (b) Each system must be designed so that scribed in subparagraphs (1) and (2) of this essential load circuits can be supplied in the paragraph, it is permissible to assume that event of reasonably probable faults or open the power loads are reduced in accordance circuits, including faults in heavy current with a monitoring procedure which is con- carrying cables. sistent with safety in the types of operations (c) If two independent sources of electrical authorized. power for particular equipment or systems 60. Ventilation. The ventilation system of are required by this regulation, their elec- the airplane must meet the requirements of trical energy supply must be insured by FAR 23.831, and in addition, for pressurized means such as duplicate electrical equip- aircraft the ventilating air in flight crew and ment, throwover switching, or multichannel passenger compartments must be free of or loop circuits separately routed. harmful or hazardous concentrations of 64. Circuit protective devices. The circuit gases and vapors in normal operation and in protective devices for the electrical circuits the event of reasonably probable failures or of the airplane must meet the requirements malfunctioning of the ventilating, heating, of FAR 23.1357, and in addition circuits for pressurization, or other systems, and equip- loads which are essential to safe operation ment. If accumulation of hazardous quan- must have individual and exclusive circuit tities of smoke in the cockpit area is reason- protection. ably probable, smoke evacuation must be readily accomplished. [Doc. No. 8070, 34 FR 189, Jan. 7, 1969, as amended by SFAR 23–1, 34 FR 20176, Dec. 24, ELECTRICAL SYSTEMS AND EQUIPMENT 1969; 35 FR 1102, Jan. 28, 1970] 61. General. The electrical systems and equipment of the airplane must meet the re- Subpart A—General quirements of FAR 23.1351, and the following: (a) Electrical system capacity. The required § 23.1 Applicability. generating capacity, and number and kinds (a) This part prescribes airworthiness of power sources must— standards for the issue of type certifi- (1) Be determined by an electrical load analysis, and cates, and changes to those certifi- (2) Meet the requirements of FAR 23.1301. cates, for airplanes in the normal, util- (b) Generating system. The generating sys- ity, acrobatic, and commuter cat- tem includes electrical power sources, main egories. power busses, transmission cables, and asso- (b) Each person who applies under ciated control, regulation, and protective de- Part 21 for such a certificate or change vices. It must be designed so that— must show compliance with the appli- (1) The system voltage and frequency (as applicable) at the terminals of all essential cable requirements of this part. load equipment can be maintained within [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as the limits for which the equipment is de- amended by Amdt. 23–34, 52 FR 1825, Jan. 15, signed, during any probable operating condi- 1987] tions; (2) System transients due to switching, § 23.2 Special retroactive require- fault clearing, or other causes do not make ments. essential loads inoperative, and do not cause a smoke or fire hazard; (a) Notwithstanding §§ 21.17 and 21.101 (3) There are means, accessible in flight to of this chapter and irrespective of the appropriate crewmembers, for the individual type certification basis, each normal,

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utility, and acrobatic category air- (1) Any maneuver incident to normal plane having a passenger seating con- flying; figuration, excluding pilot seats, of (2) Stalls (except whip stalls); and nine or less, manufactured after De- (3) Lazy eights, chandelles, and steep cember 12, 1986, or any such foreign air- turns, in which the angle of bank is not plane for entry into the United States more than 60 degrees. must provide a safety belt and shoulder (b) The utility category is limited to harness for each forward- or aft-facing airplanes that have a seating configu- seat which will protect the occupant ration, excluding pilot seats, of nine or from serious head injury when sub- less, a maximum certificated takeoff jected to the inertia loads resulting weight of 12,500 pounds or less, and in- from the ultimate static load factors tended for limited acrobatic operation. prescribed in § 23.561(b)(2) of this part, Airplanes certificated in the utility or which will provide the occupant pro- category may be used in any of the op- tection specified in § 23.562 of this part erations covered under paragraph (a) of when that section is applicable to the this section and in limited acrobatic airplane. For other seat orientations, operations. Limited acrobatic oper- the seat/restraint system must be de- ation includes: signed to provide a level of occupant (1) Spins (if approved for the par- protection equivalent to that provided ticular type of airplane); and for forward- or aft-facing seats with a (2) Lazy eights, chandelles, and steep safety belt and shoulder harness in- turns, or similar maneuvers, in which stalled. the angle of bank is more than 60 de- (b) Each shoulder harness installed at grees but not more than 90 degrees. a flight crewmember station, as re- (c) The acrobatic category is limited quired by this section, must allow the to airplanes that have a seating con- crewmember, when seated with the figuration, excluding pilot seats, of safety belt and shoulder harness fas- nine or less, a maximum certificated tened, to perform all functions nec- takeoff weight of 12,500 pounds or less, essary for flight operations. and intended for use without restric- (c) For the purpose of this section, tions, other than those shown to be the date of manufacture is: necessary as a result of required flight (1) The date the inspection accept- tests. ance records, or equivalent, reflect (d) The commuter category is limited that the airplane is complete and to propeller-driven, multiengine air- meets the FAA approved type design planes that have a seating configura- data; or tion, excluding pilot seats, of 19 or less, (2) In the case of a foreign manufac- and a maximum certificated takeoff tured airplane, the date the foreign weight of 19,000 pounds or less. The civil airworthiness authority certifies commuter category operation is lim- the airplane is complete and issues an ited to any maneuver incident to nor- original standard airworthiness certifi- mal flying, stalls (except whip stalls), cate, or the equivalent in that country. and steep turns, in which the angle of bank is not more than 60 degrees. [Amdt. 23–36, 53 FR 30812, Aug. 15, 1988] (e) Except for commuter category, airplanes may be type certificated in § 23.3 Airplane categories. more than one category if the require- (a) The normal category is limited to ments of each requested category are airplanes that have a seating configu- met. ration, excluding pilot seats, of nine or [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as less, a maximum certificated takeoff amended by Amdt. 23–4, 32 FR 5934, Apr. 14, weight of 12,500 pounds or less, and in- 1967; Amdt. 23–34, 52 FR 1825, Jan. 15, 1987; 52 tended for nonacrobatic operation. FR 34745, Sept. 14, 1987; Amdt. 23–50, 61 FR Nonacrobatic operation includes: 5183, Feb. 9, 1996]

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Subpart B—Flight ing weight) is shown. The maximum weight must be established so that it GENERAL is— (1) Not more than the least of— § 23.21 Proof of compliance. (i) The highest weight selected by the (a) Each requirement of this subpart applicant; or must be met at each appropriate com- (ii) The design maximum weight, bination of weight and center of grav- which is the highest weight at which ity within the range of loading condi- compliance with each applicable struc- tions for which certification is re- tural loading condition of this part quested. This must be shown— (other than those complied with at the (1) By tests upon an airplane of the design landing weight) is shown; or type for which certification is re- (iii) The highest weight at which quested, or by calculations based on, compliance with each applicable flight and equal in accuracy to, the results of requirement is shown, and testing; and (2) Not less than the weight with— (2) By systematic investigation of each probable combination of weight (i) Each seat occupied, assuming a and center of gravity, if compliance weight of 170 pounds for each occupant cannot be reasonably inferred from for normal and commuter category air- combinations investigated. planes, and 190 pounds for utility and (b) The following general tolerances acrobatic category airplanes, except are allowed during flight testing. How- that seats other than pilot seats may ever, greater tolerances may be al- be placarded for a lesser weight; and lowed in particular tests: (A) Oil at full capacity, and (B) At least enough fuel for max- Item Tolerance imum continuous power operation of at Weight ...... +5%, –10%. least 30 minutes for day-VFR approved Critical items affected by weight ...... +5%, –1%. airplanes and at least 45 minutes for C.G ...... ±7% total travel. night-VFR and IFR approved airplanes; or § 23.23 Load distribution limits. (ii) The required minimum crew, and (a) Ranges of weights and centers of fuel and oil to full tank capacity. gravity within which the airplane may (b) Minimum weight. The minimum be safely operated must be established. weight (the lowest weight at which If a weight and center of gravity com- compliance with each applicable re- bination is allowable only within cer- quirement of this part is shown) must tain lateral load distribution limits be established so that it is not more that could be inadvertently exceeded, than the sum of— these limits must be established for the (1) The empty weight determined corresponding weight and center of under § 23.29; gravity combinations. (2) The weight of the required min- (b) The load distribution limits may imum crew (assuming a weight of 170 not exceed any of the following: pounds for each crewmember); and (1) The selected limits; (3) The weight of— (2) The limits at which the structure (i) For turbojet powered airplanes, 5 is proven; or percent of the total fuel capacity of (3) The limits at which compliance that particular fuel tank arrangement with each applicable flight require- under investigation, and ment of this subpart is shown. (ii) For other airplanes, the fuel nec- [Doc. No. 26269, 58 FR 42156, Aug. 6, 1993] essary for one-half hour of operation at maximum continuous power. § 23.25 Weight limits. (a) Maximum weight. The maximum [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13086, Aug. 13, weight is the highest weight at which 1969; Amdt. 23–21, 43 FR 2317, Jan. 16, 1978; compliance with each applicable re- Amdt. 23–34, 52 FR 1825, Jan. 15, 1987; Amdt. quirement of this part (other than 23–45, 58 FR 42156, Aug. 6, 1993; Amdt. 23–50, 61 those complied with at the design land- FR 5183, Feb. 9, 1996]

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§ 23.29 Empty weight and cor- not greater than the maximum allow- responding center of gravity. able takeoff r.p.m.; and (a) The empty weight and cor- (2) During a closed throttle glide, at responding center of gravity must be VNE, the propeller may not cause an determined by weighing the airplane engine speed above 110 percent of max- with— imum continuous speed. (1) Fixed ballast; (c) Controllable pitch propellers without (2) Unusable fuel determined under constant speed controls. Each propeller § 23.959; and that can be controlled in flight, but (3) Full operating fluids, including— that does not have constant speed con- (i) Oil; trols, must have a means to limit the (ii) Hydraulic fluid; and pitch range so that— (1) The lowest possible pitch allows (iii) Other fluids required for normal compliance with paragraph (b)(1) of operation of airplane systems, except this section; and potable water, lavatory precharge (2) The highest possible pitch allows water, and water intended for injection compliance with paragraph (b)(2) of in the engines. this section. (b) The condition of the airplane at (d) Controllable pitch propellers with the time of determining empty weight constant speed controls. Each control- must be one that is well defined and lable pitch propeller with constant can be easily repeated. speed controls must have— [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 (1) With the governor in operation, a FR 258, Jan. 9, 1965, as amended by Amdt. 23– means at the governor to limit the 21, 43 FR 2317, Jan. 16, 1978] maximum engine speed to the max- imum allowable takeoff r.p.m.; and § 23.31 Removable ballast. (2) With the governor inoperative, Removable ballast may be used in the propeller blades at the lowest pos- showing compliance with the flight re- sible pitch, with takeoff power, the air- quirements of this subpart, if— plane stationary, and no wind, either— (a) The place for carrying ballast is (i) A means to limit the maximum properly designed and installed, and is engine speed to 103 percent of the max- marked under § 23.1557; and imum allowable takeoff r.p.m., or (b) Instructions are included in the (ii) For an engine with an approved airplane flight manual, approved man- overspeed, a means to limit the max- ual material, or markings and plac- imum engine and propeller speed to not ards, for the proper placement of the more than the maximum approved removable ballast under each loading overspeed. condition for which removable ballast is necessary. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–45, 58 FR 42156, Aug. 6, [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 1993; Amdt. 23–50, 61 FR 5183, Feb. 9, 1996] FR 258, Jan. 9, 1965, as amended by Amdt. 23– 13, 37 FR 20023, Sept. 23, 1972] PERFORMANCE

§ 23.33 Propeller speed and pitch lim- § 23.45 General. its. (a) Unless otherwise prescribed, the (a) General. The propeller speed and performance requirements of this part pitch must be limited to values that must be met for— will assure safe operation under normal (1) Still air and standard atmosphere; operating conditions. and (b) Propellers not controllable in flight. (2) Ambient atmospheric conditions, For each propeller whose pitch cannot for commuter category airplanes, for be controlled in flight— reciprocating engine-powered airplanes (1) During takeoff and initial climb of more than 6,000 pounds maximum at the all engine(s) operating climb weight, and for turbine engine-powered speed specified in § 23.65, the propeller airplanes. must limit the engine r.p.m., at full (b) Performance data must be deter- throttle or at maximum allowable mined over not less than the following takeoff manifold pressure, to a speed ranges of conditions—

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(1) Airport altitudes from sea level to NOTE: The effect on these distances of op- 10,000 feet; and eration on other types of surfaces (for exam- (2) For reciprocating engine-powered ple, grass, gravel) when dry, may be deter- airplanes of 6,000 pounds, or less, max- mined or derived and these surfaces listed in the Airplane Flight Manual in accordance imum weight, temperature from stand- with § 23.1583(p). ard to 30 °C above standard; or (3) For reciprocating engine-powered (h) For commuter category airplanes, airplanes of more than 6,000 pounds the following also apply: maximum weight and turbine engine- (1) Unless otherwise prescribed, the powered airplanes, temperature from applicant must select the takeoff, standard to 30 °C above standard, or enroute, approach, and landing con- the maximum ambient atmospheric figurations for the airplane. temperature at which compliance with (2) The airplane configuration may the cooling provisions of § 23.1041 to vary with weight, altitude, and tem- § 23.1047 is shown, if lower. perature, to the extent that they are (c) Performance data must be deter- compatible with the operating proce- mined with the cowl flaps or other dures required by paragraph (h)(3) of means for controlling the engine cool- this section. ing air supply in the position used in (3) Unless otherwise prescribed, in de- the cooling tests required by § 23.1041 to termining the critical-engine-inoper- § 23.1047. ative takeoff performance, takeoff (d) The available propulsive thrust flight path, and accelerate-stop dis- must correspond to engine power, not tance, changes in the airplane’s con- exceeding the approved power, less— figuration, speed, and power must be (1) Installation losses; and made in accordance with procedures es- (2) The power absorbed by the acces- tablished by the applicant for oper- sories and services appropriate to the ation in service. particular ambient atmospheric condi- (4) Procedures for the execution of tions and the particular flight condi- discontinued approaches and balked tion. landings associated with the conditions (e) The performance, as affected by prescribed in § 23.67(c)(4) and § 23.77(c) engine power or thrust, must be based must be established. on a relative humidity: (5) The procedures established under (1) Of 80 percent at and below stand- paragraphs (h)(3) and (h)(4) of this sec- ard temperature; and tion must— (2) From 80 percent, at the standard (i) Be able to be consistently exe- temperature, varying linearly down to cuted by a crew of average skill in at- 34 percent at the standard temperature mospheric conditions reasonably ex- plus 50 °F. pected to be encountered in service; (f) Unless otherwise prescribed, in de- (ii) Use methods or devices that are termining the takeoff and landing dis- safe and reliable; and tances, changes in the airplane’s con- (iii) Include allowance for any rea- figuration, speed, and power must be sonably expected time delays in the made in accordance with procedures es- execution of the procedures. tablished by the applicant for oper- ation in service. These procedures must [Doc. No. 27807, 61 FR 5184, Feb. 9, 1996] be able to be executed consistently by pilots of average skill in atmospheric § 23.49 Stalling period.

conditions reasonably expected to be (a) VSO and VS1 are the stalling encountered in service. speeds or the minimum steady flight (g) The following, as applicable, must speeds, in knots (CAS), at which the be determined on a smooth, dry, hard- airplane is controllable with— surfaced runway— (1) For reciprocating engine-powered (1) Takeoff distance of § 23.53(b); airplanes, the engine(s) idling, the (2) Accelerate-stop distance of § 23.55; throttle(s) closed or at not more than (3) Takeoff distance and takeoff run the power necessary for zero thrust at of § 23.59; and a speed not more than 110 percent of (4) Landing distance of § 23.75. the stalling speed;

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(2) For turbine engine-powered air- (b) For normal, utility, and acrobatic planes, the propulsive thrust not great- category airplanes, the speed at 50 feet er than zero at the stalling speed, or, if above the takeoff surface level must the resultant thrust has no appreciable not be less than: effect on the stalling speed, with en- (1) or multiengine airplanes, the gine(s) idling and throttle(s) closed; highest of— (3) The propeller(s) in the takeoff po- (i) A speed that is shown to be safe sition; for continued flight (or emergency (4) The airplane in the condition ex- landing, if applicable) under all reason- isting in the test, in which VSO and ably expected conditions, including VS1 are being used; turbulence and complete failure of the (5) The center of gravity in the posi- critical engine; tion that results in the highest value of (ii) 1.10 VMC; or VSO and VS1; and (iii) 1.20 VS1. (6) The weight used when VSO and (2) For single-engine airplanes, the VS1 are being used as a factor to de- higher of— termine compliance with a required (i) A speed that is shown to be safe performance standard. under all reasonably expected condi- (b) VSO and VS1 must be determined tions, including turbulence and com- by flight tests, using the procedure and plete engine failure; or meeting the flight characteristics spec- (ii) 1.20 VS1. ified in § 23.201. (c) For commuter category airplanes, (c) Except as provided in paragraph the following apply: (l) V must be established in relation (d) of this section, VSO and VS1 at 1 maximum weight must not exceed 61 to VEF as follows: knots for— (i) VEF is the calibrated airspeed at (1) Single-engine airplanes; and which the critical engine is assumed to (2) Multiengine airplanes of 6,000 fail. VEF must be selected by the appli- pounds or less maximum weight that cant but must not be less than 1.05 VMC cannot meet the minimum rate of determined under § 23.149(b) or, at the climb specified in § 23.67(a) (1) with the option of the applicant, not less than critical engine inoperative. VMCG determined under § 23.149(f). (d) All single-engine airplanes, and (ii) The takeoff decision speed, V1, is those multiengine airplanes of 6,000 the calibrated airspeed on the ground pounds or less maximum weight with a at which, as a result of engine failure or other reasons, the pilot is assumed VSO of more than 61 knots that do not meet the requirements of § 23.67(a)(1), to have made a decision to continue or must comply with § 23.562(d). discontinue the takeoff. The takeoff decision speed, V1, must be selected by [Doc. No. 27807, 61 FR 5184, Feb. 9, 1996] the applicant but must not be less than VEF plus the speed gained with the crit- § 23.51 Takeoff speeds. ical engine inoperative during the time (a) For normal, utility, and acrobatic interval between the instant at which category airplanes, rotation speed, VR, the critical engine is failed and the in- is the speed at which the pilot makes a stant at which the pilot recognizes and control input, with the intention of reacts to the engine failure, as indi- lifting the airplane out of contact with cated by the pilot’s application of the the runway or water surface. first retarding means during the accel- (1) For multiengine landplanes, VR, erate-stop determination of § 23.55. must not be less than the greater of (2) The rotation speed, VR, in terms 1.05 VMC; or 1.10 VS1; of calibrated airspeed, must be selected (2) For single-engine landplanes, VR, by the applicant and must not be less must not be less than VS1; and than the greatest of the following: (3) For seaplanes and amphibians (i) V1; taking off from water, VR, may be any (ii) 1.05 VMC determined under speed that is shown to be safe under all § 23.149(b); reasonably expected conditions, includ- (iii) 1.10 VS1; or ing turbulence and complete failure of (iv) The speed that allows attaining the critical engine. the initial climb-out speed, V2, before 201

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reaching a height of 35 feet above the (c) For commuter category airplanes, takeoff surface in accordance with takeoff performance, as required by § 23.57(c)(2). §§ 23.55 through 23.59, must be deter- (3) For any given set of conditions, mined with the operating engine(s) such as weight, altitude, temperature, within approved operating limitations. and configuration, a single value of VR must be used to show compliance with [Doc. No. 27807, 61 FR 5185, Feb. 9, 1996] both the one-engine-inoperative take- § 23.55 Accelerate-stop distance. off and all-engines-operating takeoff requirements. For each commuter category air- (4) The takeoff safety speed, V2, in plane, the accelerate-stop distance terms of calibrated airspeed, must be must be determined as follows: selected by the applicant so as to allow (a) The accelerate-stop distance is the gradient of climb required in § 23.67 the sum of the distances necessary to— (c)(1) and (c)(2) but mut not be less (1) Accelerate the airplane from a than 1.10 VMC or less than 1.20 VS1. standing start to VEF with all engines (5) The one-engine-inoperative take- operating; off distance, using a normal rotation (2) Accelerate the airplane from VEF rate at a speed 5 knots less than VR, es- to V1, assuming the critical engine tablished in accordance with paragraph fails at VEF; and (c)(2) of this section, must be shown (3) Come to a full stop from the point not to exceed the corresponding one- at which V1 is reached. engine-inoperative takeoff distance, (b) Means other than wheel brakes determined in accordance with § 23.57 may be used to determine the accel- and § 23.59(a)(1), using the established erate-stop distances if that means— VR. The takeoff, otherwise performed (1) Is safe and reliable; in accordance with § 23.57, must be con- (2) Is used so that consistent results tinued safely from the point at which can be expected under normal oper- the airplane is 35 feet above the takeoff ating conditions; and surface and at a speed not less than the (3) Is such that exceptional skill is established V2 minus 5 knots. (6) The applicant must show, with all not required to control the airplane. engines operating, that marked in- [Amdt. 23–34, 52 FR 1826, Jan. 15, 1987, as creases in the scheduled takeoff dis- amended by Amdt. 23–50, 61 FR 5185, Feb. 9, tances, determined in accordance with 1996] § 23.59(a)(2), do not result from over-ro- tation of the airplane or out-of-trim § 23.57 Takeoff path. conditions. For each commuter category air- [Doc. No. 27807, 61 FR 5184, Feb. 9, 1996] plane, the takeoff path is as follows: (a) The takeoff path extends from a § 23.53 Takeoff performance. standing start to a point in the takeoff (a) For normal, utility, and acrobatic at which the airplane is 1500 feet above category airplanes, the takeoff dis- the takeoff surface at or below which tance must be determined in accord- height the transition from the takeoff ance with paragraph (b) of this section, to the enroute configuration must be using speeds determined in accordance completed; and with § 23.51 (a) and (b). (1) The takeoff path must be based on (b) For normal, utility, and acrobatic the procedures prescribed in § 23.45; category airplanes, the distance re- (2) The airplane must be accelerated quired to takeoff and climb to a height on the ground to VEF at which point the of 50 feet above the takeoff surface critical engine must be made inoper- must be determined for each weight, ative and remain inoperative for the altitude, and temperature within the rest of the takeoff; and operational limits established for take- (3) After reaching VEF, the airplane off with— must be accelerated to V2. (1) Takeoff power on each engine; (b) During the acceleration to speed (2) Wing flaps in the takeoff posi- V2, the nose gear may be raised off the tion(s); and ground at a speed not less than VR. (3) Landing gear extended. However, landing gear retraction must

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not be initiated until the airplane is option of the applicant, the takeoff airborne. run, must be determined. (c) During the takeoff path deter- (a) Takeoff distance is the greater mination, in accordance with para- of— graphs (a) and (b) of this section— (1) The horizontal distance along the (1) The slope of the airborne part of takeoff path from the start of the take- the takeoff path must not be negative off to the point at which the airplane is at any point; 35 feet above the takeoff surface as de- (2) The airplane must reach V2 before termined under § 23.57; or it is 35 feet above the takeoff surface, (2) With all engines operating, 115 and must continue at a speed as close percent of the horizontal distance from as practical to, but not less than V2, the start of the takeoff to the point at until it is 400 feet above the takeoff which the airplane is 35 feet above the surface; takeoff surface, determined by a proce- (3) At each point along the takeoff dure consistent with § 23.57. path, starting at the point at which the airplane reaches 400 feet above the (b) If the takeoff distance includes a takeoff surface, the available gradient clearway, the takeoff run is the greater of climb must not be less than— of— (i) 1.2 percent for two-engine air- (1) The horizontal distance along the planes; takeoff path from the start of the take- (ii) 1.5 percent for three-engine air- off to a point equidistant between the planes; liftoff point and the point at which the (iii) 1.7 percent for four-engine air- airplane is 35 feet above the takeoff planes; and surface as determined under § 23.57; or (4) Except for gear retraction and (2) With all engines operating, 115 automatic propeller feathering, the percent of the horizontal distance from airplane configuration must not be the start of the takeoff to a point equi- changed, and no change in power that distant between the liftoff point and requires action by the pilot may be the point at which the airplane is 35 made, until the airplane is 400 feet feet above the takeoff surface, deter- above the takeoff surface. mined by a procedure consistent with (d) The takeoff path to 35 feet above § 23.57. the takeoff surface must be determined [Amdt. 23–34, 52 FR 1827, Jan. 15, 1987, as by a continuous demonstrated takeoff. amended by Amdt. 23–50, 61 FR 5185, Feb. 9, (e) The takeoff path to 35 feet above 1996] the takeoff surface must be determined by synthesis from segments; and § 23.61 Takeoff flight path. (1) The segments must be clearly de- fined and must be related to distinct For each commuter category air- changes in configuration, power, and plane, the takeoff flight path must be speed; determined as follows: (2) The weight of the airplane, the (a) The takeoff flight path begins 35 configuration, and the power must be feet above the takeoff surface at the assumed constant throughout each seg- end of the takeoff distance determined ment and must correspond to the most in accordance with § 23.59. critical condition prevailing in the seg- (b) The net takeoff flight path data ment; and must be determined so that they rep- (3) The takeoff flight path must be resent the actual takeoff flight paths, based on the airplane’s performance as determined in accordance with without utilizing ground effect. § 23.57 and with paragraph (a) of this section, reduced at each point by a gra- [Amdt. 23–34, 52 FR 1827, Jan. 15, 1987, as amended by Amdt. 23–50, 61 FR 5185, Feb. 9, dient of climb equal to— 1996] (1) 0.8 percent for two-engine air- planes; § 23.59 Takeoff distance and takeoff (2) 0.9 percent for three-engine air- run. planes; and For each commuter category air- (3) 1.0 percent for four-engine air- plane, the takeoff distance and, at the planes.

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(c) The prescribed reduction in climb § 23.65 Climb: All engines operating. gradient may be applied as an equiva- (a) Each normal, utility, and acro- lent reduction in acceleration along batic category reciprocating engine- that part of the takeoff flight path at powered airplane of 6,000 pounds or less which the airplane is accelerated in level flight. maximum weight must have a steady climb gradient at sea level of at least [Amdt. 23–34, 52 FR 1827, Jan. 15, 1987] 8.3 percent for landplanes or 6.7 percet for seaplanes and amphibians with— § 23.63 Climb: General. (1) Not more than maximum contin- (a) Compliance with the require- uous power on each engine; ments of §§ 23.65, 23.66, 23.67, 23.69, and (2) The landing gear retracted; 23.77 must be shown— (3) The wing flaps in the takeoff posi- (1) Out of ground effect; and tion(s); and (2) At speeds that are not less than (4) A climb speed not less than the those at which compliance with the greater of 1.1 VMC and 1.2 VS1 for multi- powerplant cooling requirements of engine airplanes and not less than 1.2 §§ 23.1041 to 23.1047 has been dem- VS1 for single—engine airplanes. onstrated; and (b) Each normal, utility, and acro- (3) Unless otherwise specified, with batic category reciprocating engine- one engine inoperative, at a bank angle powered airplane of more than 6,000 not exceeding 5 degrees. pounds maximum weight and turbine (b) For normal, utility, and acrobatic engine-powered airplanes in the nor- category reciprocating engine-powered mal, utility, and acrobatic category airplanes of 6,000 pounds or less max- must have a steady gradient of climb imum weight, compliance must be after takeoff of at least 4 percent with shown with § 23.65(a), § 23.67(a), where (1) Take off power on each engine; appropriate, and § 23.77(a) at maximum (2) The landing gear extended, except takeoff or landing weight, as appro- that if the landing gear can be re- priate, in a standard atmosphere. tracted in not more than seven sec- (c) For normal, utility, and acrobatic onds, the test may be conducted with category reciprocating engine-powered the gear retracted; airplanes of more than 6,000 pounds (3) The wing flaps in the takeoff posi- maximum weight, and turbine engine- tion(s); and powered airplanes in the normal, util- (4) A climb speed as specified in ity, and acrobatic category, compli- ance must be shown at weights as a § 23.65(a)(4). function of airport altitude and ambi- [Doc. No. 27807, 61 FR 5186, Feb. 9, 1996] ent temperature, within the oper- ational limits established for takeoff § 23.66 Takeoff climb: One-engine inop- and landing, respectively, with— erative. (1) Sections 23.65(b) and 23.67(b) (1) For normal, utility, and acrobatic and (2), where appropriate, for takeoff, category reciprocating engine-powered and airplanes of more than 6,000 pounds (2) Section 23.67(b)(2), where appro- maximum weight, and turbine engine- priate, and § 23.77(b), for landing. powered airplanes in the normal, util- (d) For commuter category airplanes, ity, and acrobatic category, the steady compliance must be shown at weights gradient of climb or descent must be as a function of airport altitude and determined at each weight, altitude, ambient temperature within the oper- and ambient temperature within the ational limits established for takeoff operational limits established by the and landing, respectively, with— applicant with— (1) Sections 23.67(c)(1), 23.67(c)(2), and (a) The critical engine inoperative 23.67(c)(3) for takeoff; and and its propeller in the position it rap- (2) Sections 23.67(c)(3), 23.67(c)(4), and idly and automatically assumes; 23.77(c) for landing. (b) The remaining engine(s) at take- [Doc. No. 27807, 61 FR 5186, Feb. 9, 1996] off power;

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(c) The landing gear extended, except (ii) Remaining engine(s) at takeoff that if the landing gear can be re- power; tracted in not more than seven sec- (iii) Landing gear retracted; onds, the test may be conducted with (iv) Wing flaps in the takeoff posi- the gear retracted; tion(s); and (d) The wing flaps in the takeoff posi- (v) Climb speed equal to that tion(s): achieved at 50 feet in the demonstra- (e) The wings level; and tion of § 23.53. (f) A climb speed equal to that (2) The steady gradient of climb must achieved at 50 feet in the demonstra- not be less than 0.75 percent at an alti- tion of § 23.53. tude of 1,500 feet above the takeoff sur- [Doc. No. 27807, 61 FR 5186, Feb. 9, 1996] face, or landing surface, as appropriate, with the— § 23.67 Climb: One engine inoperative. (i) Critical engine inoperative and its (a) For normal, utility, and acrobatic propeller in the minimum drag posi- category reciprocating engine-powered tion; airplanes of 6,000 pounds or less max- (ii) Remaining engine(s) at not more imum weight, the following apply: than maximum continuous power; (1) Except for those airplanes that (iii) Landing gear retracted; meet the requirements prescribed in (iv) Wing flaps retracted; and (v) Climb speed not less than 1.2 VS1. § 23.562(d), each airplane with a VSO of more than 61 knots must be able to (c) For commuter category airplanes, maintain a steady climb gradient of at the following apply: least 1.5 percent at a pressure altitude (1) Takeoff; landing gear extended. The of 5,000 feet with the— steady gradient of climb at the altitude (i) Critical engine inoperative and its of the takeoff surface must be measur- propeller in the minimum drag posi- ably positive for two-engine airplanes, tion; not less than 0.3 percent for three-en- (ii) Remaining engine(s) at not more gine airplanes, or 0.5 percent for four- than maximum continuous power; engine airplanes with— (iii) Landing gear retracted; (i) The critical engine inoperative (iv) Wing flaps retracted; and and its propeller in the position it rap- idly and automatically assumes; (v) Climb speed not less than 1.2 VS1. (2) For each airplane that meets the (ii) The remaining engine(s) at take- requirements prescribed in § 23.562(d), off power; or that has a VSO of 61 knots or less, (iii) The landing gear extended, and the steady gradient of climb or descent all landing gear doors open; at a pressure altitude of 5,000 feet must (iv) The wing flaps in the takeoff po- be determined with the— sition(s); (i) Critical engine inoperative and its (v) The wings level; and propeller in the minimum drag posi- (vi) A climb speed equal to V2. tion; (2) Takeoff; landing gear retracted. The (ii) Remaining engine(s) at not more steady gradient of climb at an altitude than maximum continuous power; of 400 feet above the takeoff surface (iii) Landing gear retracted; must be not less than 2.0 percent of (iv) Wing flaps retracted; and two-engine airplanes, 2.3 percent for (v) Climb speed not less than 1.2VS1. three-engine airplanes, and 2.6 percent (b) For normal, utility, and acrobatic for four-engine airplanes with— category reciprocating engine-powered (i) The critical engine inoperative airplanes of more than 6,000 pounds and its propeller in the position it rap- maximum weight, and turbine engine- idly and automatically assumes; powered airplanes in the normal, util- (ii) The remaining engine(s) at take- ity, and acrobatic category— off power; (1) The steady gradient of climb at an (iii) The landing gear retracted; altitude of 400 feet above the takeoff (iv) The wing flaps in the takeoff po- must be measurably positive with the— sition(s); (i) Critical engine inoperative and its (v) A climb speed equal to V2. propeller in the minimum drag posi- (3) Enroute. The steady gradient of tion; climb at an altitude of 1,500 feet above

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the takeoff or landing surface, as ap- (1) The critical engine inoperative propriate, must be not less than 1.2 and its propeller in the minimum drag percent for two-engine airplanes, 1.5 position; percent for three-engine airplanes, and (2) The remaining engine(s) at not 1.7 percent for four-engine airplanes more than maximum continuous with— power; (i) The critical engine inoperative (3) The landing gear retracted; and its propeller in the minimum drag (4) The wing flaps retracted; and position; (5) A climb speed not less than 1.2 (ii) The remaining engine(s) at not VS1. more than maximum continuous power; [Doc. No. 27807, 61 FR 5187, Feb. 9, 1996] (iii) The landing gear retracted; § 23.71 Glide: Single-engine airplanes. (iv) The wing flaps retracted; and (v) A climb speed not less than 1.2 The maximum horizontal distance traveled in still air, in nautical miles, VS1. (4) Discontinued approach. The steady per 1,000 feet of altitude lost in a glide, gradient of climb at an altitude of 400 and the speed necessary to achieve this feet above the landing surface must be must be determined with the engine in- not less than 2.1 percent for two-engine operative, its propeller in the min- airplanes, 2.4 percent for three-engine imum drag position, and landing gear airplanes, and 2.7 percent for four-en- and wing flaps in the most favorable gine airplanes, with— available position. (i) The critical engine inoperative [Doc. No. 27807, 61 FR 5187, Feb. 9, 1996] and its propeller in the minimum drag position; § 23.73 Reference landing approach (ii) The remaining engine(s) at take- speed. off power; (a) For normal, utility, and acrobatic (iii) Landing gear retracted; category reciprocating engine-powered (iv) Wing flaps in the approach posi- airplanes of 6,000 pounds or less max- tion(s) in which VS1 for these posi- imum weight, the reference landing ap- tion(s) does not exceed 110 percent of proach speed, VREF, must not be less the VS1 for the related all-engines-oper- than the greater of VMC, determined in ated landing position(s); and § 23.149(b) with the wing flaps in the (v) A climb speed established in con- most extended takeoff position, and 1.3 nection with normal landing proce- VSO. dures but not exceeding 1.5 VS1. (b) For normal, utility, and acrobatic [Doc. No. 27807, 61 FR 5186, Feb. 9, 1996] category reciprocating engine-powered airplanes of more than 6,000 pounds § 23.69 Enroute climb/descent. maximum weight, and turbine engine- powered airplanes in the normal, util- (a) All engines operating. The steady ity, and acrobatic category, the ref- gradient and rate of climb must be de- erence landing approach speed, V , termined at each weight, altitude, and REF must not be less than the greater of ambient temperature within the oper- V , determined in § 23.149(c), and 1.3 ational limits established by the appli- MC V . cant with— SO (c) For commuter category airplanes, (1) Not more than maximum contin- the reference landing approach speed, uous power on each engine; V , must not be less than the greater (2) The landing gear retracted; REF of 1.05 VMC, determined in § 23.149(c), (3) The wing flaps retracted; and and 1.3 V . (4) A climb speed not less than 1.3 SO VS1. [Doc. No. 27807, 61 FR 5187, Feb. 9, 1996] (b) One engine inoperative. The steady gradient and rate of climb/descent § 23.75 Landing distance. must be determined at each weight, al- The horizontal distance necessary to titude, and ambient temperature with- land and come to a complete stop from in the operational limits established by a point 50 feet above the landing sur- the applicant with— face must be determined, for standard

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temperatures at each weight and alti- § 23.77 Balked landing. tude within the operational limits es- (a) Each normal, utility, and acro- tablished for landing, as follows: batic category reciprocating engine- (a) A steady approach at not less powered airplane at 6,000 pounds or less than VREF, determined in accordance maximum weight must be able to with § 23.73 (a), (b), or (c), as appro- maintain a steady gradient of climb at priate, must be maintained down to the sea level of at least 3.3 percent with— 50 foot height and— (1) Takeoff power on each engine; (1) The steady approach must be at a (2) The landing gear extended; gradient of descent not greater than 5.2 (3) The wing flaps in the landing posi- percent (3 degrees) down to the 50-foot tion, except that if the flaps may safely height. be retracted in two seconds or less (2) In addition, an applicant may without loss of altitude and without demonstrate by tests that a maximum sudden changes of angle of attack, they steady approach gradient steeper than may be retracted; and 5.2 percent, down to the 50-foot height, (4) A climb speed equal to VREF, as de- is safe. The gradient must be estab- fined in § 23.73(a). lished as an operating limitation and (b) Each normal, utility, and acro- the information necessary to display batic category reciprocating engine- the gradient must be available to the powered airplane of more than 6,000 pilot by an appropriate instrument. pounds maximum weight and each nor- (b) A constant configuration must be mal, utility, and acrobatic category maintained throughout the maneuver. turbine engine-powered airplane must (c) The landing must be made with- be able to maintain a steady gradient out excessive vertical acceleration or of climb of at least 2.5 percent with— tendency to bounce, nose over, ground (1) Not more than the power that is loop, porpoise, or water loop. available on each engine eight seconds (d) It must be shown that a safe tran- after initiation of movement of the sition to the balked landing conditions power controls from minimum flight- of § 23.77 can be made from the condi- idle position; (2) The landing gear extended; tions that exist at the 50 foot height, at (3) The wing flaps in the landing posi- maximum landing weight, or at the tion; and maximum landing weight for altitude (4) A climb speed equal to V , as de- and temperature of § 23.63 (c)(2) or REF fined in § 23.73(b). (d)(2), as appropriate. (c) Each commuter category airplane (e) The brakes must be used so as to must be able to maintain a steady gra- not cause excessive wear of brakes or dient of climb of at least 3.2 percent tires. with— (f) Retardation means other than (1) Not more than the power that is wheel brakes may be used if that available on each engine eight seconds means— after initiation of movement of the (1) Is safe and reliable; and power controls from the minimum (2) Is used so that consistent results flight idle position; can be expected in service. (2) Landing gear extended; (g) If any device is used that depends (3) Wing flaps in the landing position; on the operation of any engine, and the and landing distance would be increased (4) A climb speed equal to VREF, as de- when a landing is made with that en- fined in § 23.73(c). gine inoperative, the landing distance [Doc. No. 27807, 61 FR 5187, Feb. 9, 1996] must be determined with that engine inoperative unless the use of other FLIGHT CHARACTERISTICS compensating means will result in a landing distance not more than that § 23.141 General. with each engine operating. The airplane must meet the require- [Amdt. 23–21, 43 FR 2318, Jan. 16, 1978, as ments of §§ 23.143 through 23.253 at all amended by Amdt. 23–34, 52 FR 1828, Jan. 15, practical loading conditions and oper- 1987; Amdt. 23–42, 56 FR 351, Jan. 3, 1991; ating altitudes for which certification Amdt. 23–50, 61 FR 5187, Feb. 9, 1996] has been requested, not exceeding the

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maximum operating altitude estab- lows prompt acceleration to the trim lished under § 23.1527, and without re- speed with— quiring exceptional piloting skill, (1) Maximum continuous power on alertness, or strength. each engine; [Doc. No. 26269, 58 FR 42156, Aug. 6, 1993] (2) Power off; and (3) Wing flap and landing gear— CONTROLLABILITY AND (i) retracted, and MANEUVERABILITY (ii) extended. (b) Unless otherwise required, it must § 23.143 General. be possible to carry out the following (a) The airplane must be safely con- maneuvers without requiring the appli- trollable and maneuverable during all cation of single-handed control forces flight phases including— exceeding those specified in § 23.143(c). (1) Takeoff; The trimming controls must not be ad- (2) Climb; justed during the maneuvers: (3) Level flight; (1) With the landing gear extended, (4) Descent; the flaps retracted, and the airplanes (5) Go-around; and as nearly as possible in trim at 1.4 VS1, (6) Landing (power on and power off) extend the flaps as rapidly as possible with the wing flaps extended and re- and allow the airspeed to transition tracted. from 1.4VS1 to 1.4 VSO: (b) It must be possible to make a (i) With power off; and smooth transition from one flight con- (ii) With the power necessary to dition to another (including turns and maintain level flight in the initial con- slips) without danger of exceeding the dition. limit load factor, under any probable (2) With landing gear and flaps ex- operating condition (including, for tended, power off, and the airplane as multiengine airplanes, those condi- nearly as possible in trim at 1.3 VSO: tions normally encountered in the sud- quickly apply takeoff power and re- den failure of any engine). tract the flaps as rapidly as possible to (c) If marginal conditions exist with the recommended go around setting regard to required pilot strength, the and allow the airspeed to transition control forces necessary must be deter- from 1.3 VSO to 1.3 VS1. Retract the gear mined by quantitative tests. In no case when a positive rate of climb is estab- may the control forces under the condi- lished. tions specified in paragraphs (a) and (b) (3) With landing gear and flaps ex- of this section exceed those prescribed tended, in level flight, power necessary in the following table: to attain level flight at 1.1 VSO, and the airplane as nearly as possible in trim, Values in pounds force applied Pitch Roll Yaw to the relevant control it must be possible to maintain ap- proximately level flight while retract- (a) For temporary application: Stick ...... 60 30 ...... ing the flaps as rapidly as possible with Wheel (Two hands on rim) .... 75 50 ...... simultaneous application of not more Wheel (One hand on rim) ...... 50 25 ...... than maximum continuous power. If Rudder Pedal ...... 150 (b) For prolonged application .... 10 5 20 gated flat positions are provided, the flap retraction may be demonstrated in stages with power and trim reset for [Doc. No, 4080, 29 FR 17955, Dec. 18, 1964, as level flight at 1.1 V , in the initial con- amended by Amdt. 23–14, 38 FR 31819, Nov. 19, S1 1973; Amdt. 23–17, 41 FR 55464, Dec. 20, 1976; figuration for each stage— Amdt. 23–45, 58 FR 42156, Aug. 6, 1993; Amdt. (i) From the fully extended position 23–50, 61 FR 5188, Feb. 9, 1996] to the most extended gated position; (ii) Between intermediate gated posi- § 23.145 Longitudinal control. tions, if applicable; and (a) With the airplane as nearly as (iii) From the least extended gated possible in trim at 1.3 VS1, it must be position to the fully retracted position. possible, at speeds below the trim (4) With power off, flaps and landing speed, to pitch the nose downward so gear retracted and the airplane as that the rate of increase in airspeed al- nearly as possible in trim at 1.4 VS1, 208

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apply takeoff power rapidly while § 23.147 Directional and lateral con- maintaining the same airspeed. trol. (5) With power off, landing gear and (a) For each multiengine airplane, it flaps extended, and the airplane as must be possible, while holding the nearly as possible in trim at VREF, ob- wings level within five degrees, to tain and maintain airspeeds between make sudden changes in heading safely 1.1 VSO, and either 1.7 VSO or VFE, in both directions. This ability must be whichever is lower without requiring shown at 1.4 VS1 with heading changes the application of two-handed control up to 15 degrees, except that the head- forces exceeding those specified in ing change at which the rudder force § 23.143(c). corresponds to the limits specified in (6) With maximum takeoff power, § 23.143 need not be exceeded, with the— landing gear retracted, flaps in the (1) Critical engine inoperative and its takeoff position, and the airplane as propeller in the minimum drag posi- nearly as possible in trim at VFE appro- tion; priate to the takeoff flap position, re- (2) Remaining engines at maximum tract the flaps as rapidly as possible continuous power; while maintaining constant speed. (3) Landing gear— (c) At speeds above VMO/MMO, and up (i) Retracted; and to the maximum speed shown under (ii) Extended; and § 23.251, a maneuvering capability of 1.5 (4) Flaps retracted. g must be demonstrated to provide a (b) For each multiengine airplane, it margin to recover from upset or inad- must be possible to regain full control vertent speed increase. of the airplane without exceeding a (d) It must be possible, with a pilot bank angle of 45 degrees, reaching a control force of not more than 10 dangerous attitude or encountering pounds, to maintain a speed of not dangerous characteristics, in the event more than VREF during a power-off glide of a sudden and complete failure of the with landing gear and wing flaps ex- critical engine, making allowance for a tended, for any weight of the airplane, delay of two seconds in the initiation up to and including the maximum of recovery action appropriate to the weight. situation, with the airplane initially in (e) By using normal flight and power trim, in the following condition: controls, except as otherwise noted in (1) Maximum continuous power on paragraphs (e)(1) and (e)(2) of this sec- each engine; tion, it must be possible to establish a (2) The wing flaps retracted; zero rate of descent at an attitude suit- (3) The landing gear retracted; able for a controlled landing without (4) A speed equal to that at which exceeding the operational and struc- compliance with § 23.69(a) has been tural limitations of the airplane, as shown; and follows: (5) All propeller controls in the posi- (1) For single-engine and multiengine tion at which compliance with § 23.69(a) airplanes, without the use of the pri- has been shown. mary longitudinal control system. (c) For all airplanes, it must be (2) For multiengine airplanes— shown that the airplane is safely con- (i) Without the use of the primary di- trollable without the use of the pri- rectional control; and mary lateral control system in any all- (ii) If a single failure of any one con- engine configuration(s) and at any necting or transmitting link would af- speed or altitude within the approved fect both the longitudinal and direc- operating envelope. It must also be tional primary control system, without shown that the airplane’s flight char- the primary longitudinal and direc- acteristics are not impaired below a tional control system. level needed to permit continued safe [Doc. No. 26269, 58 FR 42157, Aug. 6, 1993; flight and the ability to maintain atti- Amdt. 23–45, 58 FR 51970, Oct. 5, 1993, as tudes suitable for a controlled landing amended by Amdt. 23–50, 61 FR 5188, Feb. 9, without exceeding the operational and 1996] structural limitations of the airplane.

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If a single failure of any one con- operative must be established and des- necting or transmitting link in the lat- ignated as the safe, intentional, one- eral control system would also cause engine-inoperative speed, VSSE. the loss of additional control sys- (e) At VMC, the rudder pedal force re- tem(s), compliance with the above re- quired to maintain control must not quirement must be shown with those exceed 150 pounds and it must not be additional systems also assumed to be necessary to reduce power of the opera- inoperative. tive engine(s). During the maneuver, the airplane must not assume any dan- [Doc. No. 27807, 61 FR 5188, Feb. 9, 1996] gerous attitude and it must be possible § 23.149 Minimum control speed. to prevent a heading change of more than 20 degrees. (a) VMC is the calibrated airspeed at (f) At the option of the applicant, to which, when the critical engine is sud- comply with the requirements of denly made inoperative, it is possible § 23.51(c)(1), V may be determined. to maintain control of the airplane MCG VMCG is the minimum control speed on with that engine still inoperative, and the ground, and is the calibrated air- thereafter maintain straight flight at speed during the takeoff run at which, the same speed with an angle of bank when the critical engine is suddenly of not more than 5 degrees. The method made inoperative, it is possible to used to simulate critical engine failure maintain control of the airplane using must represent the most critical mode the rudder control alone (without the of powerplant failure expected in serv- use of nosewheel steering), as limited ice with respect to controllability. by 150 pounds of force, and using the (b) V for takeoff must not exceed MC lateral control to the extent of keeping 1.2 VS1, where VS1 is determined at the the wings level to enable the takeoff to maximum takeoff weight. VMC must be be safely continued. In the determina- determined with the most unfavorable tion of VMCG, assuming that the path of weight and center of gravity position the airplane accelerating with all en- and with the airplane airborne and the gines operating is along the centerline ground effect negligible, for the takeoff of the runway, its path from the point configuration(s) with— at which the critical engine is made in- (1) Maximum available takeoff power operative to the point at which recov- initially on each engine; ery to a direction parallel to the cen- (2) The airplane trimmed for takeoff; terline is completed may not deviate (3) Flaps in the takeoff position(s); more than 30 feet laterally from the (4) Landing gear retracted; and centerline at any point. V must be (5) All propeller controls in the rec- MCG established with— ommended takeoff position through- (1) The airplane in each takeoff con- out. figuration or, at the option of the ap- (c) For all airplanes except recipro- plicant, in the most critical takeoff cating engine-powered airplanes of configuration; 6,000 pounds or less maximum weight, (2) Maximum available takeoff power the conditions of paragraph (a) of this on the operating engines; section must also be met for the land- (3) The most unfavorable center of ing configuration with— gravity; (1) Maximum available takeoff power (4) The airplane trimmed for takeoff; initially on each engine; and (2) The airplane trimmed for an ap- (5) The most unfavorable weight in proach, with all engines operating, at the range of takeoff weights. VREF, at an approach gradient equal to the steepest used in the landing dis- [Doc. No. 27807, 61 FR 5189, Feb. 9, 1996] tance demonstration of § 23.75; (3) Flaps in the landing position; § 23.151 Acrobatic maneuvers. (4) Landing gear extended; and Each acrobatic and utility category (5) All propeller controls in the posi- airplane must be able to perform safely tion recommended for approach with the acrobatic maneuvers for which cer- all engines operating. tification is requested. Safe entry (d) A minimum speed to inten- speeds for these maneuvers must be de- tionally render the critical engine in- termined.

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§ 23.153 Control during landings. § 23.157 Rate of roll. It must be possible, while in the land- (a) Takeoff. It must be possible, using ing configuration, to safely complete a a favorable combination of controls, to landing without exceeding the one- roll the airplane from a steady 30-de- hand control force limits specified in gree banked turn through an angle of § 23.143(c) following an approach to 60 degrees, so as to reverse the direc- land— tion of the turn within: (a) At a speed of VREF minus 5 knots; (1) For an airplane of 6,000 pounds or (b) With the airplane in trim, or as less maximum weight, 5 seconds from nearly as possible in trim and without initiation of roll; and the trimming control being moved (2) For an airplane of over 6,000 throughout the maneuver; pounds maximum weight, (c) At an approach gradient equal to the steepest used in the landing dis- (W+500)/1,300 tance demonstration of § 23.75; and seconds, but not more than 10 seconds, (d) With only those power changes, if where W is the weight in pounds. any, that would be made when landing (b) The requirement of paragraph (a) normally from an approach at V . REF of this section must be met when roll- [Doc. No. 27807, 61 FR 5189, Feb. 9, 1996] ing the airplane in each direction with— § 23.155 Elevator control force in ma- (1) Flaps in the takeoff position; neuvers. (2) Landing gear retracted; (a) The elevator control force needed (3) For a single-engine airplane, at to achieve the positive limit maneu- maximum takeoff power; and for a vering load factor may not be less multiengine airplane with the critical than: engine inoperative and the propeller in (1) For wheel controls, W/100 (where the minimum drag position, and the W is the maximum weight) or 20 other engines at maximum takeoff pounds, whichever is greater, except power; and that it need not be greater than 50 (4) The airplane trimmed at a speed pounds; or equal to the greater of 1.2 V or 1.1 (2) For stick controls, W/140 (where W S1 VMC, or as nearly as possible in trim for is the maximum weight) or 15 pounds, straight flight. whichever is greater, except that it (c) Approach. It must be possible, need not be greater than 35 pounds. using a favorable combination of con- (b) The requirement of paragraph (a) trols, to roll the airplane from a steady of this section must be met at 75 per- 30-degree banked turn through an angle cent of maximum continuous power for of 60 degrees, so as to reverse the direc- reciprocating engines, or the maximum tion of the turn within: continuous power for turbine engines, (1) For an airplane of 6,000 pounds or and with the wing flaps and landing gear retracted— less maximum weight, 4 seconds from initiation of roll; and (1) In a turn, with the trim setting used for wings level flight at V ; and (2) For an airplane of over 6,000 O pounds maximum weight, (2) In a turn with the trim setting used for the maximum wings level (W+2,800)/2,200 flight speed, except that the speed may not exceed VNE or VMO/MMO, whichever seconds, but not more than 7 seconds, is appropriate. where W is the weight in pounds. (c) There must be no excessive de- (d) The requirement of paragraph (c) crease in the gradient of the curve of of this section must be met when roll- stick force versus maneuvering load ing the airplane in each direction in factor with increasing load factor. the following conditions— (1) Flaps in the landing position(s); [Amdt. 23–14, 38 FR 31819, Nov. 19, 1973; 38 FR 32784, Nov. 28, 1973, as amended by Amdt. 23– (2) Landing gear extended; 45, 58 FR 42158, Aug. 6, 1993; Amdt. 23–50, 61 (3) All engines operating at the power FR 5189 Feb. 9, 1996] for a 3 degree approach; and

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(4) The airplane trimmed at VREF. (4) Approach with landing gear ex- tended and with— [Amdt. 23–14, 38 FR 31819, Nov. 19, 1973, as amended by Amdt. 23–45, 58 FR 42158, Aug. 6, (i) A 3 degree angle of descent, with 1993; Amdt. 23–50, 61 FR 5189, Feb. 9, 1996] flaps retracted and at a speed of 1.4 VS1; (ii) A 3 degree angle of descent, flaps TRIM in the landing position(s) at VREF; and (iii) An approach gradient equal to § 23.161 Trim. the steepest used in the landing dis- (a) General. Each airplane must meet tance demonstrations of § 23.75, flaps in the trim requirements of this section the landing position(s) at VREF. after being trimmed and without fur- (d) In addition, each multiple air- ther pressure upon, or movement of, plane must maintain longitudinal and the primary controls or their cor- directional trim, and the lateral con- responding trim controls by the pilot trol force must not exceed 5 pounds at or the automatic pilot. In addition, it the speed used in complying with must be possible, in other conditions of § 23.67(a), (b)(2), or (c)(3), as appro- loading, configuration, speed and power priate, with— to ensure that the pilot will not be un- (1) The critical engine inoperative, duly fatigued or distracted by the need and if applicable, its propeller in the to apply residual control forces exceed- minimum drag position; ing those for prolonged application of (2) The remaining engines at max- § 23.143(c). This applies in normal oper- imum continuous power; ation of the airplane and, if applicable, (3) The landing gear retracted; to those conditions associated with the (4) Wing flaps retracted; and failure of one engine for which per- (5) An angle of bank of not more than formance characteristics are estab- five degrees. lished. (e) In addition, each commuter cat- (b) Lateral and directional trim. The egory airplane for which, in the deter- airplane must maintain lateral and di- mination of the takeoff path in accord- rectional trim in level flight with the ance with § 23.57, the climb in the take- landing gear and wing flaps retracted off configuration at V2 extends beyond as follows: 400 feet above the takeoff surface, it (1) For normal, utility, and acrobatic must be possible to reduce the longitu- category airplanes, at a speed of 0.9 VH, dinal and lateral control forces to 10 VC, or VMO/MO, whichever is lowest; and pounds and 5 pounds, respectively, and (2) For commuter category airplanes, the directional control force must not at all speeds from 1.4 VS1 to the lesser exceed 50 pounds at V2 with— of VH or VMO/MMO. (1) The critical engine inoperative (c) Longitudinal trim. The airplane and its propeller in the minimum drag must maintain longitudinal trim under position; each of the following conditions: (2) The remaining engine(s) at take- (1) A climb with— off power; (i) Takeoff power, landing gear re- (3) Landing gear retracted; tracted, wing flaps in the takeoff posi- (4) Wing flaps in the takeoff posi- tion(s), at the speeds used in deter- tion(s); and mining the climb performance required (5) An angle of bank not exceeding 5 by § 23.65; and degrees. (ii) Maximum continuous power at the speeds and in the configuration [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as used in determining the climb perform- amended by Amdt. 23–21, 43 FR 2318, Jan. 16, 1978; Amdt. 23–34, 52 FR 1828, Jan. 15, 1987; ance required by § 23.69(a). Amdt. 23–42, 56 FR 351, Jan. 3, 1991; 56 FR (2) Level flight at all speeds from the 5455, Feb. 11, 1991; Amdt. 23–50, 61 FR 5189, lesser of VH and either VNO or VMO/MMO Feb. 9, 1996] (as appropriate), to 1.4 VS1, with the landing gear and flaps retracted. STABILITY (3) A descent at VNO or VMO/MMO, whichever is applicable, with power off § 23.171 General. and with the landing gear and flaps re- The airplane must be longitudinally, tracted. directionally, and laterally stable

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under §§ 23.173 through 23.181. In addi- (3) Maximum continuous power; and tion, the airplane must show suitable (4) The airplane trimmed at the speed stability and control ‘‘feel’’ (static sta- used in determining the climb perform- bility) in any condition normally en- ance required by § 23.69(a). countered in service, if flight tests (b) Cruise. With flaps and landing show it is necessary for safe operation. gear retracted and the airplane in trim with power for level flight at represent- § 23.173 Static longitudinal stability. ative cruising speeds at high and low Under the conditions specified in altitudes, including speeds up to VNO or § 23.175 and with the airplane trimmed VMO/MMO, as appropriate, except that as indicated, the characteristics of the the speed need not exceed VH— elevator control forces and the friction (1) For normal, utility, and acrobatic within the control system must be as category airplanes, the stick force follows: curve must have a stable slope at all (a) A pull must be required to obtain speeds within a range that is the great- and maintain speeds below the speci- er of 15 percent of the trim speed plus fied trim speed and a push required to the resulting free return speed range, obtain and maintain speeds above the or 40 knots plus the resulting free re- specified trim speed. This must be turn speed range, above and below the shown at any speed that can be ob- trim speed, except that the slope need tained, except that speeds requiring a not be stable— control force in excess of 40 pounds or (i) At speeds less than 1.3 VS1; or speeds above the maximum allowable (ii) For airplanes with VNE estab- speed or below the minimum speed for lished under § 23.1505(a), at speeds steady unstalled flight, need not be greater than VNE; or considered. (iii) For airplanes with VMO/MMO es- (b) The airspeed must return to with- tablished under § 23.1505(c), at speeds in the tolerances specified for applica- greater than VFC/MFC. ble categories of airplanes when the (2) For commuter category airplanes, control force is slowly released at any the stick force curve must have a sta- speed within the speed range specified ble slope at all speeds within a range of in paragraph (a) of this section. The ap- 50 knots plus the resulting free return plicable tolerances are— speed range, above and below the trim (1) The airspeed must return to with- speed, except that the slope need not be in plus or minus 10 percent of the origi- stable— nal trim airspeed; and (i) At speeds less than 1.4 VS1; or (2) For commuter category airplanes, (ii) At speeds greater than VFC/MFC; the airspeed must return to within plus or or minus 7.5 percent of the original (iii) At speeds that require a stick trim airspeed for the cruising condition force greater than 50 pounds. specified in § 23.175(b). (c) Landing. The stick force curve (c) The stick force must vary with must have a stable slope at speeds be- speed so that any substantial speed tween 1.1 VS1 and 1.8 VS1 with— change results in a stick force clearly (1) Flaps in the landing position; perceptible to the pilot. (2) Landing gear extended; and (3) The airplane trimmed at— [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (i) VREF, or the minimum trim speed amended by Amdt. 23–14, 38 FR 31820 Nov. 19, if higher, with power off; and 1973; Amdt. 23–34, 52 FR 1828, Jan. 15, 1987] (ii) VREF with enough power to main- § 23.175 Demonstration of static longi- tain a 3 degree angle of descent. tudinal stability. [Doc. No. 27807, 61 FR 5190, Feb. 9, 1996] Static longitudinal stability must be shown as follows: § 23.177 Static directional and lateral (a) Climb. The stick force curve must stability. have a stable slope at speeds between (a) The static directional stability, as 85 and 115 percent of the trim speed, shown by the tendency to recover from with— a wings level sideslip with the rudder (1) Flaps retracted; free, must be positive for any landing (2) Landing gear retracted; gear and flap position appropriate to

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the takeoff, climb, cruise, approach, contained in § 23.143 is reached, the ai- and landing configurations. This must leron and rudder control movements be shown with symmetrical power up and forces must not reverse as the to maximum continuous power, and at angle of sideslip is increased. Rapid speeds from 1.2 VS1 up to the maximum entry into, and recovery from, a max- allowable speed for the condition being imum sideslip considered appropriate investigated. The angel of sideslip for for the airplane must not result in un- these tests must be appropriate to the controllable flight characteristics. type of airplane. At larger angles of sideslip, up to that at which full rudder [Doc. No. 27807, 61 FR 5190, Feb. 9, 1996] is used or a control force limit in § 23.143 is reached, whichever occurs § 23.181 Dynamic stability. first, and at speeds from 1.2 VS1 to VO, (a) Any short period oscillation not the rudder pedal force must not re- including combined lateral-directional verse. oscillations occurring between the (b) The static lateral stability, as stalling speed and the maximum allow- shown by the tendency to raise the low able speed appropriate to the configu- wing in a sideslip, must be positive for ration of the airplane must be heavily all landing gear and flap positions. damped with the primary controls— This must be shown with symmetrical (1) Free; and power up to 75 percent of maximum continuous power at speeds above 1.2 (2) In a fixed position. (b) Any combined lateral-directional VS1 in the take off configuration(s) and oscillations (‘‘Dutch roll’’) occurring be- at speeds above 1.3 VS1 in other con- figurations, up to the maximum allow- tween the stalling speed and the max- able speed for the configuration being imum allowable speed appropriate to investigated, in the takeoff, climb, the configuration of the airplane must cruise, and approach configurations. be damped to 1/10 amplitude in 7 cycles For the landing configuration, the with the primary controls— power must be that necessary to main- (1) Free; and tain a 3 degree angle of descent in co- (2) In a fixed position. ordinated flight. The static lateral sta- (c) If it is determined that the func- bility must not be negative at 1.2 VS1 in tion of a stability augmentation sys- the takeoff configuration, or at 1.3 VS1 tem, reference § 23.672, is needed to in other configurations. The angle of meet the flight characteristic require- sideslip for these tests must be appro- ments of this part, the primary control priate to the type of airplane, but in no requirements of paragraphs (a)(2) and case may the constant heading sideslip (b)(2) of this section are not applicable angle be less than that obtainable with to the tests needed to verify the ac- a 10 degree bank, or if less, the max- ceptability of that system. imum bank angle obtainable with full rudder deflection or 150 pound rudder (d) During the conditions as specified force. in § 23.175, when the longitudinal con- (c) Paragraph (b) of this section does trol force required to maintain speeds not apply to acrobatic category air- differing from the trim speed by at planes certificated for inverted flight. least plus and minus 15 percent is sud- (d) In straight, steady slips at 1.2 VS1 denly released, the response of the air- for any landing gear and flap positions, plane must not exhibit any dangerous and for any symmetrical power condi- characteristics nor be excessive in rela- tions up to 50 percent of maximum con- tion to the magnitude of the control tinuous power, the aileron and rudder force released. Any long-period oscilla- control movements and forces must in- tion of flight path, phugoid oscillation, crease steadily, but not necessarily in that results must not be so unstable as constant proportion, as the angle of to increase the pilot’s workload or oth- sideslip is increased up to the max- erwise endanger the airplane. imum appropriate to the type of air- plane. At larger slip angles, up to the [Amdt. 23–21, 43 FR 2318, Jan. 16, 1978, as angle at which full rudder or aileron amended by Amdt. 23–45, 58 FR 42158, Aug. 6, control is used or a control force limit 1993]

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STALLS may not be less than 50 percent of max- imum continuous power. § 23.201 Wings level stall. (5) Trim. The airplane trimmed at a (a) It must be possible to produce and speed as near 1.5 VS1 as practicable. to correct roll by unreversed use of the (6) Propeller. Full increase r.p.m. posi- rolling control and to produce and to tion for the power off condition. correct yaw by unreversed use of the [Doc. No. 27807, 61 FR 5191, Feb. 9, 1996] directional control, up to the time the airplane stalls. § 23.203 Turning flight and accelerated (b) The wings level stall characteris- turning stalls. tics must be demonstrated in flight as Turning flight and accelerated turn- follows. Starting from a speed at least ing stalls must be demonstrated in 10 knots above the stall speed, the ele- tests as follows: vator control must be pulled back so (a) Establish and maintain a coordi- that the rate of speed reduction will nated turn in a 30 degree bank. Reduce not exceed one knot per second until a speed by steadily and progressively stall is produced, as shown by either: tightening the turn with the elevator (1) An uncontrollable downward until the airplane is stalled, as defined pitching motion of the airplane; in § 23.201(b). The rate of speed reduc- (2) A downward pitching motion of tion must be constant, and— the airplane that results from the acti- (1) For a turning flight stall, may not vation of a stall avoidance device (for exceed one knot per second; and example, stick pusher); or (2) For an accelerated turning stall, (3) The control reaching the stop. be 3 to 5 knots per second with steadily (c) Normal use of elevator control for increasing normal acceleration. recovery is allowed after the downward (b) After the airplane has stalled, as pitching motion of paragraphs (b)(1) or defined in § 23.201(b), it must be possible (b)(2) of this section has unmistakably to regain wings level flight by normal been produced, or after the control has use of the flight controls, but without been held against the stop for not less increasing power and without— than the longer of two seconds or the (1) Excessive loss of altitude; time employed in the minimum steady (2) Undue pitchup; slight speed determination of § 23.49. (3) Uncontrollable tendency to spin; (d) During the entry into and the re- (4) Exceeding a bank angle of 60 de- covery from the maneuver, it must be grees in the original direction of the possible to prevent more than 15 de- turn or 30 degrees in the opposite direc- grees of roll or yaw by the normal use tion in the case of turning flight stalls; of controls. (5) Exceeding a bank angle of 90 de- (e) Compliance with the require- grees in the original direction of the ments of this section must be shown turn or 60 degrees in the opposite direc- under the following conditions: tion in the case of accelerated turning (1) Wing flaps. Retracted, fully ex- stalls; and tended, and each intermediate normal (6) Exceeding the maximum permis- operating position. sible speed or allowable limit load fac- (2) Landing gear. Retracted and ex- tor. tended. (c) Compliance with the require- (3) Cowl flaps. Appropriate to configu- ments of this section must be shown ration. under the following conditions: (4) Power: (1) Wing flaps: Retracted, fully ex- (i) Power off; and tended, and each intermediate normal (ii) 75 percent of maximum contin- operating position; uous power. However, if the power-to- (2) Landing gear: Retracted and ex- weight ratio at 75 percent of maximum tended; continuous power result in extreme (3) Cowl flaps: Appropriate to configu- nose-up attitudes, the test may be car- ration; ried out with the power required for (4) Power: level flight in the landing configura- (i) Power off; and tion at maximum landing weight and a (ii) 75 percent of maximum contin- speed of 1.4 VSO, except that the power uous power. However, if the power-to- 215

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weight ratio at 75 percent of maximum SPINNING continuous power results in extreme nose-up attitudes, the test may be car- § 23.221 Spinning. ried out with the power required for (a) Normal category airplanes. A sin- level flight in the landing configura- gle-engine, normal category airplane tion at maximum landing weight and a must be able to recover from a one- speed of 1.4 VSO, except that the power turn spin or a three-second spin, which- may not be less than 50 percent of max- ever takes longer, in not more than one imum continuous power. additional turn after initiation of the (5) Trim: The airplane trimmed at a first control action for recovery, or speed as near 1.5 VS1 as practicable. demonstrate compliance with the op- (6) Propeller. Full increase rpm posi- tional spin resistant requirements of tion for the power off condition. this section. [Amdt. 23–14, 38 FR 31820, Nov. 19, 1973, as (1) The following apply to one turn or amended by Amdt. 23–45, 58 FR 42159, Aug. 6, three second spins: 1993; Amdt. 23–50, 61 FR 5191, Feb. 9, 1996] (i) For both the flaps-retracted and flaps-extended conditions, the applica- § 23.207 Stall warning. ble airspeed limit and positive limit (a) There must be a clear and distinc- maneuvering load factor must not be tive stall warning, with the flaps and exceeded; landing gear in any normal position, in (ii) No control forces or char- straight and turning flight. acteristic encountered during the spin (b) The stall warning may be fur- or recovery may adversely affect nished either through the inherent aer- prompt recovery; odynamic qualities of the airplane or (iii) It must be impossible to obtain by a device that will give clearly dis- unrecoverable spins with any use of the tinguishable indications under ex- flight or engine power controls either pected conditions of flight. However, a at the entry into or during the spin; visual stall warning device that re- and quires the attention of the crew within (iv) For the flaps-extended condition, the cockpit is not acceptable by itself. the flaps may be retracted during the (c) During the stall tests required by recovery but not before rotation has § 23.201(b) and § 23.203(a)(1), the stall ceased. warning must begin at a speed exceed- (2) At the applicant’s option, the air- ing the stalling speed by a margin of plane may be demonstrated to be spin not less than 5 knots and must con- resistant by the following: tinue until the stall occurs. (i) During the stall maneuver con- (d) When following procedures fur- tained in § 23.201, the pitch control nished in accordance with § 23.1585, the must be pulled back and held against stall warning must not occur during a the stop. Then, using ailerons and rud- takeoff with all engines operating, a ders in the proper direction, it must be takeoff continued with one engine in- possible to maintain wings-level flight operative, or during an approach to within 15 degrees of bank and to roll landing. the airplane from a 30 degree bank in (e) During the stall tests required by one direction to a 30 degree bank in the § 23.203(a)(2), the stall warning must other direction; begin sufficiently in advance of the (ii) Reduce the airplane speed using stall for the stall to be averted by pilot pitch control at a rate of approxi- action taken after the stall warning mately one knot per second until the first occurs. pitch control reaches the stop; then, (f) For acrobatic category airplanes, with the pitch control pulled back and an artificial stall warning may be mu- held against the stop, apply full rudder table, provided that it is armed auto- control in a manner to promote spin matically during takeoff and rearmed entry for a period of seven seconds or automatically in the approach configu- through a 360 degree heading change, ration. whichever occurs first. If the 360 degree [Amdt. 23–7, 34 FR 13087, Aug. 13, 1969, as heading change is reached first, it must amended by Amdt. 23–45, 58 FR 42159, Aug. 6, have taken no fewer than four seconds. 1993; Amdt. 23–50, 61 FR 5191, Feb. 9, 1996] This maneuver must be performed first

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with the ailerons in the neutral posi- (4) There must be no characteristics tion, and then with the ailerons de- during the spin (such as excessive rates flected opposite the direction of turn in of rotation or extreme oscillatory mo- the most adverse manner. Power and tion) that might prevent a successful airplane configuration must be set in recovery due to disorientation or inca- accordance with § 23.201(e) without pacitation of the pilot. change during the maneuver. At the [Doc. No. 27807, 61 FR 5191, Feb. 9, 1996] end of seven seconds or a 360 degree heading change, the airplane must re- GROUND AND WATER HANDLING spond immediately and normally to CHARACTERISTICS primary flight controls applied to re- gain coordinated, unstalled flight with- § 23.231 Longitudinal stability and out reversal of control effect and with- control. out exceeding the temporary control (a) A landplane may have no uncon- forces specified by § 23.143(c); and trollable tendency to nose over in any (iii) Compliance with §§ 23.201 and reasonably expected operating condi- 23.203 must be demonstrated with the tion, including rebound during landing airplane in uncoordinated flight, cor- or takeoff. Wheel brakes must operate responding to one ball width displace- smoothly and may not induce any ment on a slip-skid indicator, unless undue tendency to nose over. one ball width displacement cannot be (b) A seaplane or amphibian may not obtained with full rudder, in which have dangerous or uncontrollable case the demonstration must be with porpoising characteristics at any nor- full rudder applied. mal operating speed on the water. (b) Utility category airplanes. A utility § 23.233 Directional stability and con- category airplane must meet the re- trol. quirements of paragraph (a) of this sec- (a) A 90 degree cross-component of tion. In addition, the requirements of wind velocity, demonstrated to be safe paragraph (c) of this section and for taxiing, takeoff, and landing must § 23.807(b)(7) must be met if approval for be established and must be not less spinning is requested. than 0.2 VSO. (c) Acrobatic category airplanes. An ac- (b) The airplane must be satisfac- robatic category airplane must meet torily controllable in power-off land- the spin requirements of paragraph (a) ings at normal landing speed, without of this section and § 23.807(b)(6). In addi- using brakes or engine power to main- tion, the following requirements must tain a straight path until the speed has be met in each configuration for which decreased to at least 50 percent of the approval for spinning is requested: speed at touchdown. (1) The airplane must recover from (c) The airplane must have adequate any point in a spin up to and including directional control during taxiing. six turns, or any greater number of (d) Seaplanes must demonstrate sat- turns for which certification is re- isfactory directional stability and con- quested, in not more than one and one- trol for water operations up to the half additional turns after initiation of maximum wind velocity specified in the first control action for recovery. paragraph (a) of this section. However, beyond three turns, the spin [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as may be discontinued if spiral charac- amended by Amdt. 23–45, 58 FR 42159, Aug. 6, teristics appear. 1993; Amdt. 23–50, 61 FR 5192, Feb. 9, 1996] (2) The applicable airspeed limits and limit maneuvering load factors must § 23.235 Operation on unpaved sur- not be exceeded. For flaps-extended faces. configurations for which approval is re- The airplane must be demonstrated quested, the flaps must not be re- to have satisfactory characteristics tracted during the recovery. and the shock-absorbing mechanism (3) It must be impossible to obtain must not damage the structure of the unrecoverable spins with any use of the airplane when the airplane is taxied on flight or engine power controls either the roughest ground that may reason- at the entry into or during the spin. ably be expected in normal operation

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and when takeoffs and landings are ent or artificial speed warning speci- performed on unpaved runways having fied in § 23.1303, it must be shown that the roughest surface that may reason- the airplane can be recovered to a nor- ably be expected in normal operation. mal attitude and its speed reduced to [Doc. No. 27807, 61 FR 5192, Feb. 9, 1996] VMO/MMO, without— (1) Exceeding VD/MD, the maximum § 23.237 Operation on water. speed shown under § 23.251, or the struc- tural limitations; or A wave height, demonstrated to be (2) Buffeting that would impair the safe for operation, and any necessary pilot’s ability to read the instruments water handling procedures for sea- or to control the airplane for recovery. planes and amphibians must be estab- (c) There may be no control reversal lished. about any axis at any speed up to the [Doc. No. 27807, 61 FR 5192, Feb. 9, 1996] maximum speed shown under § 23.251. Any reversal of elevator control force § 23.239 Spray characteristics. or tendency of the airplane to pitch, Spray may not dangerously obscure roll, or yaw must be mild and readily the vision of the pilots or damage the controllable, using normal piloting propellers or other parts of a seaplane techniques. or amphibian at any time during tax- [Amdt. 23–7, 34 FR 13087, Aug. 13, 1969; as iing, takeoff, and landing. amended by Amdt. 23–26, 45 FR 60170, Sept. 11, 1980; Amdt. 23–45, 58 FR 42160, Aug. 6, 1993; MISCELLANEOUS FLIGHT REQUIREMENTS Amdt. 23–50, 61 FR 5192, Feb. 9, 1996] § 23.251 Vibration and buffeting. There must be no vibration or buf- Subpart C—Structure feting severe enough to result in struc- GENERAL tural damage, and each part of the air- plane must be free from excessive vi- § 23.301 Loads. bration, under any appropriate speed (a) Strength requirements are speci- and power conditions up to VD/MD. In fied in terms of limit loads (the max- addition, there must be no buffeting in imum loads to be expected in service) any normal flight condition severe and ultimate loads (limit loads multi- enough to interfere with the satisfac- plied by prescribed factors of safety). tory control of the airplane or cause Unless otherwise provided, prescribed excessive fatigue to the flight crew. loads are limit loads. Stall warning buffeting within these (b) Unless otherwise provided, the limits is allowable. air, ground, and water loads must be [Doc. No. 26269, 58 FR 42159, Aug. 6, 1993] placed in equilibrium with inertia forces, considering each item of mass § 23.253 High speed characteristics. in the airplane. These loads must be If a maximum operating speed VMO/ distributed to conservatively approxi- MMO is established under § 23.1505(c), mate or closely represent actual condi- the following speed increase and recov- tions. Methods used to determine load ery characteristics must be met: intensities and distribution on canard (a) Operating conditions and charac- and tandem wing configurations must teristics likely to cause inadvertent be validated by flight test measure- speed increases (including upsets in ment unless the methods used for de- pitch and roll) must be simulated with termining those loading conditions are the airplane trimmed at any likely shown to be reliable or conservative on speed up to VMO/MMO. These conditions the configuration under consideration. and characteristics include gust upsets, (c) If deflections under load would inadvertent control movements, low significantly change the distribution of stick force gradients in relation to con- external or internal loads, this redis- trol friction, passenger movement, lev- tribution must be taken into account. eling off from climb, and descent from (d) Simplified structural design cri- Mach to airspeed limit altitude. teria may be used if they result in de- (b) Allowing for pilot reaction time sign loads not less than those pre- after occurrence of the effective inher- scribed in §§ 23.331 through 23.521. For

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airplane configurations described in those for which experience has shown appendix A, § 23.1, the design criteria of this method to be reliable. In other appendix A of this part are an approved cases, substantiating load tests must equivalent of §§ 23.321 through 23.459. If be made. Dynamic tests, including appendix A of this part is used, the en- structural flight tests, are acceptable if tire appendix must be substituted for the design load conditions have been the corresponding sections of this part. simulated. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 (b) Certain parts of the structure FR 258, Jan. 9, 1965, as amended by Amdt. 23– must be tested as specified in Subpart 28, 47 FR 13315, Mar. 29, 1982; Amdt. 23–42, 56 D of this part. FR 352, Jan. 3, 1991; Amdt. 23–48, 61 FR 5143, Feb. 9, 1996] FLIGHT LOADS

§ 23.302 Canard or tandem wing con- § 23.321 General. figurations. (a) Flight load factors represent the The forward structure of a canard or ratio of the aerodynamic force compo- tandem wing configuration must: nent (acting normal to the assumed (a) Meet all requirements of subpart longitudinal axis of the airplane) to the C and subpart D of this part applicable weight of the airplane. A positive flight to a wing; and load factor is one in which the aero- (b) Meet all requirements applicable dynamic force acts upward, with re- to the function performed by these sur- faces. spect to the airplane. (b) Compliance with the flight load [Amdt. 23–42, 56 FR 352, Jan. 3, 1991] requirements of this subpart must be shown— § 23.303 Factor of safety. (1) At each critical altitude within Unless otherwise provided, a factor of the range in which the airplane may be safety of 1.5 must be used. expected to operate; (2) At each weight from the design § 23.305 Strength and deformation. minimum weight to the design max- (a) The structure must be able to imum weight; and support limit loads without detri- (3) For each required altitude and mental, permanent deformation. At weight, for any practicable distribution any load up to limit loads, the defor- of disposable load within the operating mation may not interfere with safe op- limitations specified in §§ 23.1583 eration. through 23.1589. (b) The structure must be able to (c) When significant, the effects of support ultimate loads without failure compressibility must be taken into ac- for at least three seconds, except local count. failures or structural instabilities be- tween limit and ultimate load are ac- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as ceptable only if the structure can sus- amended by Amdt. 23–45, 58 FR 42160, Aug. 6, tain the required ultimate load for at 1993] least three seconds. However when proof of strength is shown by dynamic § 23.331 Symmetrical flight conditions. tests simulating actual load condi- (a) The appropriate balancing hori- tions, the three second limit does not zontal tail load must be accounted for apply. in a rational or conservative manner [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as when determining the wing loads and amended by Amdt. 23–45, 58 FR 42160, Aug. 6, linear inertia loads corresponding to 1993] any of the symmetrical flight condi- tions specified in §§ 23.333 through § 23.307 Proof of structure. 23.341. (a) Compliance with the strength and (b) The incremental horizontal tail deformation requirements of § 23.305 loads due to maneuvering and gusts must be shown for each critical load must be reacted by the angular inertia condition. Structural analysis may be of the airplane in a rational or conserv- used only if the structure conforms to ative manner.

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(c) Mutual influence of the aero- resulting limit load factors must cor- dynamic surfaces must be taken into respond to the conditions determined account when determining flight loads. as follows: (i) Positive (up) and negative (down) [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23– gusts of 50 f.p.s. at VC must be consid- 42, 56 FR 352, Jan. 3, 1991] ered at altitudes between sea level and 20,000 feet. The gust velocity may be § 23.333 Flight envelope. reduced linearly from 50 f.p.s. at 20,000 (a) General. Compliance with the feet to 25 f.p.s. at 50,000 feet. strength requirements of this subpart (ii) Positive and negative gusts of 25 must be shown at any combination of f.p.s. at VD must be considered at alti- airspeed and load factor on and within tudes between sea level and 20,000 feet. the boundaries of a flight envelope The gust velocity may be reduced lin- (similar to the one in paragraph (d) of early from 25 f.p.s. at 20,000 feet to 12.5 this section) that represents the enve- f.p.s. at 50,000 feet. lope of the flight loading conditions (iii) In addition, for commuter cat- specified by the maneuvering and gust egory airplanes, positive (up) and nega- criteria of paragraphs (b) and (c) of this tive (down) rough air gusts of 66 f.p.s. section respectively. at VB must be considered at altitudes (b) Maneuvering envelope. Except between sea level and 20,000 feet. The where limited by maximum (static) lift gust velocity may be reduced linearly coefficients, the airplane is assumed to from 66 f.p.s. at 20,000 feet to 38 f.p.s. at be subjected to symmetrical maneu- 50,000 feet. vers resulting in the following limit (2) The following assumptions must load factors: be made: (1) The positive maneuvering load (i) The shape of the gust is— factor specified in § 23.337 at speeds up U ⎛ 2πs ⎞ to VD; =−de (2) The negative maneuvering load U ⎜1 COS ⎟ 2 ⎝ 25C⎠ factor specified in § 23.337 at VC; and (3) Factors varying linearly with Where— speed from the specified value at VC to s=Distance penetrated into gust (ft.); C=Mean geometric chord of wing (ft.); and 0.0 at VD for the normal and commuter Ude=Derived gust velocity referred to in sub- category, and ¥1.0 at VD for the acro- batic and utility categories. paragraph (1) of this section. (c) Gust envelope. (1) The airplane is (ii) Gust load factors vary linearly assumed to be subjected to symmet- with speed between VC and VD . rical vertical gusts in level flight. The (d) Flight envelope.

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[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13087, Aug. 13, 1969; Amdt. 23–34, 52 FR 1829, Jan. 15, 1987]

§ 23.335 Design airspeeds. (ii) 1.50 VC min (for utility category Except as provided in paragraph airplanes); and (a)(4) of this section, the selected de- (iii) 1.55 VC min (for acrobatic cat- sign airspeeds are equivalent airspeeds egory airplanes). (EAS). (3) For values of W/S more than 20, (a) Design cruising speed, VC. For VC the multiplying factors in paragraph the following apply: (b)(2) of this section may be decreased (1) Where W/S′=wing loading at the linearly with W/S to a value of 1.35 design maximum takeoff weight, Vc (in where W/S=100. knots) may not be less than— (4) Compliance with paragraphs (b)(1) (i) 33 √(W/S) (for normal, utility, and and (2) of this section need not be commuter category airplanes); shown if VD/MD is selected so that the (ii) 36 √(W/S) (for acrobatic category minimum speed margin between VC/MC airplanes). and VD/MD is the greater of the fol- (2) For values of W/S more than 20, lowing: the multiplying factors may be de- (i) The speed increase resulting when, creased linearly with W/S to a value of from the initial condition of stabilized 28.6 where W/S=100. flight at VC/MC, the airplane is assumed (3) VC need not be more than 0.9 VH at sea level. to be upset, flown for 20 seconds along ° (4) At altitudes where an M is estab- a flight path 7.5 below the initial path, D and then pulled up with a load factor of lished, a cruising speed MC limited by compressibility may be selected. 1.5 (0.5 g. acceleration increment). At least 75 percent maximum continuous (b) Design dive speed VD. For VD, the following apply: power for reciprocating engines, and (1) VD/MD may not be less than 1.25 maximum cruising power for turbines, VC/MC; and or, if less, the power required for VC/MC (2) With VC min, the required min- for both kinds of engines, must be as- imum design cruising speed, VD (in sumed until the pullup is initiated, at knots) may not be less than— which point power reduction and pilot- (i) 1.40 Vc min (for normal and com- controlled drag devices may be used; muter category airplanes); and either—

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(ii) Mach 0.05 for normal, utility, and where W=design maximum takeoff acrobatic category airplanes (at alti- weight, except that n need not be more tudes where MD is established); or than 3.8; (iii) Mach 0.07 for commuter category (2) 4.4 for utility category airplanes; airplanes (at altitudes where MD is es- or tablished) unless a rational analysis, (3) 6.0 for acrobatic category air- including the effects of automatic sys- planes. tems, is used to determine a lower mar- (b) The negative limit maneuvering gin. If a rational analysis is used, the load factor may not be less than— minimum speed margin must be (1) 0.4 times the positive load factor enough to provide for atmospheric for the normal utility and commuter variations (such as horizontal gusts), categories; or and the penetration of jet streams or (2) 0.5 times the positive load factor cold fronts), instrument errors, air- for the acrobatic category. frame production variations, and must (c) Maneuvering load factors lower not be less than Mach 0.05. than those specified in this section (c) Design maneuvering speed VA. For may be used if the airplane has design VA, the following applies: features that make it impossible to ex- (1) VA may not be less than VS√n ceed these values in flight. where— [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (i) VS is a computed stalling speed amended by Amdt. 23–7, 34 FR 13088, Aug. 13, with flaps retracted at the design 1969; Amdt. 23–34, 52 FR 1829, Jan. 15, 1987; weight, normally based on the max- Amdt. 23–48, 61 FR 5144, Feb. 9, 1996] imum airplane normal force coeffi- cients, CNA; and § 23.341 Gust loads factors. (ii) n is the limit maneuvering load (a) Each airplane must be designed to factor used in design withstand loads on each lifting surface (2) The value of VA need not exceed resulting from gusts specified in the value of VC used in design. § 23.333(c). (d) Design speed for maximum gust in- (b) The gust load for a canard or tan- tensity, VB. For VB, the following apply: dem wing configuration must be com- (1) VB may not be less than the speed puted using a rational analysis, or may determined by the intersection of the be computed in accordance with para- line representing the maximum posi- graph (c) of this section, provided that tive lift, CNMAX, and the line rep- the resulting net loads are shown to be resenting the rough air gust velocity conservative with respect to the gust on the gust V-n diagram, or VS1√ ng, criteria of § 23.333(c). whichever is less, where: (c) In the absence of a more rational (i) ng the positive airplane gust load analysis, the gust load factors must be factor due to gust, at speed VC (in ac- computed as follows— cordance with § 23.341), and at the par- ticular weight under consideration; and KUVagde (ii) VS1 is the stalling speed with the n =+1 flaps retracted at the particular weight 498 (WS / ) under consideration. Where— (2) VB need not be greater than VC. Kg=0.88μg/5.3+μg=gust alleviation factor; [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as μg=2(W/S)/r Cag=airplane mass ratio; amended by Amdt. 23–7, 34 FR 13088, Aug. 13, Ude=Derived gust velocities referred to in 1969; Amdt. 23–16, 40 FR 2577, Jan. 14, 1975; § 23.333(c) (f.p.s.); Amdt. 23–34, 52 FR 1829, Jan. 15, 1987; Amdt. r=Density of air (slugs/cu.ft.); 23–24, 52 FR 34745, Sept. 14, 1987; Amdt. 23–48, W/S=Wing loading (p.s.f.) due to the applica- 61 FR 5143, Feb. 9, 1996] ble weight of the airplane in the particular load case. § 23.337 Limit maneuvering load fac- W/S=Wing loading (p.s.f.); tors. C=Mean geometric chord (ft.); g=Acceleration due to gravity (ft./sec.2) (a) The positive limit maneuvering V=Airplane equivalent speed (knots); and load factor n may not be less than— a=Slope of the airplane normal force coeffi- ÷ (1) 2.1+(24,000 (W+10,000)) for normal cient curve CNA per radian if the gust loads and commuter category airplanes, are applied to the wings and horizontal tail

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surfaces simultaneously by a rational (1) Maneuvering, to a positive limit method. The wing lift curve slope CL per load factor of 2.0; and radian may be used when the gust load is (2) Positive and negative gust of 25 applied to the wings only and the hori- feet per second acting normal to the zontal tail gust loads are treated as a sepa- rate condition. flight path in level flight. (b) VF must be assumed to be not less [Amdt. 23–7, 34 FR 13088, Aug. 13, 1969, as than 1.4 VS or 1.8 VSF, whichever is amended by Amdt. 23–42, 56 FR 352, Jan. 3, greater, where— 1991; Amdt. 23–48, 61 FR 5144, Feb. 9, 1996] (1) VS is the computed stalling speed § 23.343 Design fuel loads. with flaps retracted at the design weight; and (a) The disposable load combinations (2) VSF is the computed stalling speed must include each fuel load in the with flaps fully extended at the design range from zero fuel to the selected weight. maximum fuel load. (3) If an automatic flap load limiting (b) If fuel is carried in the wings, the device is used, the airplane may be de- maximum allowable weight of the air- signed for the critical combinations of plane without any fuel in the wing airspeed and flap position allowed by tank(s) must be established as ‘‘max- that device. imum zero wing fuel weight,’’ if it is (c) In determining external loads on less than the maximum weight. the airplane as a whole, thrust, slip- (c) For commuter category airplanes, stream, and pitching acceleration may a structural reserve fuel condition, not be assumed to be zero. exceeding fuel necessary for 45 minutes (d) The flaps, their operating mecha- of operation at maximum continuous nism, and their supporting structures, power, may be selected. If a structural must be designed to withstand the con- reserve fuel condition is selected, it ditions prescribed in paragraph (a) of must be used as the minimum fuel this section. In addition, with the flaps weight condition for showing compli- fully extended at VF, the following con- ance with the flight load requirements ditions, taken separately, must be ac- prescribed in this part and— counted for: (1) The structure must be designed to (1) A head-on gust having a velocity withstand a condition of zero fuel in of 25 feet per second (EAS), combined the wing at limit loads corresponding with propeller slipstream cor- to: responding to 75 percent of maximum (i) Ninety percent of the maneu- continuous power; and vering load factors defined in § 23.337, (2) The effects of propeller slipstream and corresponding to maximum takeoff (ii) Gust velocities equal to 85 per- power. cent of the values prescribed in [Doc. No. 27805, 61 FR 5144, Feb. 9, 1996] § 23.333(c). (2) The fatigue evaluation of the § 23.347 Unsymmetrical flight condi- structure must account for any in- tions. crease in operating stresses resulting (a) The airplane is assumed to be sub- from the design condition of paragraph jected to the unsymmetrical flight con- (c)(1) of this section. ditions of §§ 23.349 and 23.351. Unbal- (3) The flutter, deformation, and vi- anced aerodynamic moments about the bration requirements must also be met center of gravity must be reacted in a with zero fuel in the wings. rational or conservative manner, con- [Doc. No. 27805, 61 FR 5144, Feb. 9, 1996] sidering the principal masses fur- nishing the reacting inertia forces. § 23.345 High lift devices. (b) Acrobatic category airplanes cer- (a) If flaps or similar high lift devices tified for flick maneuvers (snap roll) are to be used for takeoff, approach or must be designed for additional asym- landing, the airplane, with the flaps metric loads acting on the wing and the horizontal tail. fully extended at VF, is assumed to be subjected to symmetrical maneuvers [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as and gusts within the range determined amended by Amdt. 23–48, 61 FR 5144, Feb. 9, by— 1996]

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§ 23.349 Rolling conditions. § 23.361 Engine torque. The wing and wing bracing must be (a) Each engine mount and its sup- designed for the following loading con- porting structure must be designed for ditions: the effects of— (a) Unsymmetrical wing loads appro- (1) A limit engine torque cor- priate to the category. Unless the fol- responding to takeoff power and pro- lowing values result in unrealistic peller speed acting simultaneously with 75 percent of the limit loads from loads, the rolling accelerations may be flight condition A of § 23.333(d); obtained by modifying the symmet- (2) A limit engine torque cor- rical flight conditions in § 23.333(d) as responding to maximum continuous follows: power and propeller speed acting si- (1) For the acrobatic category, in multaneously with the limit loads from conditions A and F, assume that 100 flight condition A of § 23.333(d); and percent of the semispan wing airload (3) For turbopropeller installations, acts on one side of the plane of sym- in addition to the conditions specified metry and 60 percent of this load acts in paragraphs (a)(1) and (a)(2) of this on the other side. section, a limit engine torque cor- (2) For normal, utility, and com- responding to takeoff power and pro- muter categories, in Condition A, as- peller speed, multiplied by a factor ac- sume that 100 percent of the semispan counting for propeller control system wing airload acts on one side of the air- malfunction, including quick feath- plane and 75 percent of this load acts ering, acting simultaneously with lg level flight loads. In the absence of a on the other side. rational analysis, a factor of 1.6 must (b) The loads resulting from the aile- be used. ron deflections and speeds specified in (b) For turbine engine installations, § 23.455, in combination with an air- the engine mounts and supporting plane load factor of at least two thirds structure must be designed to with- of the positive maneuvering load factor stand each of the following: used for design. Unless the following (1) A limit engine torque load im- values result in unrealistic loads, the posed by sudden engine stoppage due to effect of aileron displacement on wing malfunction or structural failure (such torsion may be accounted for by adding as compressor jamming). the following increment to the basic (2) A limit engine torque load im- airfoil moment coefficient over the ai- posed by the maximum acceleration of leron portion of the span in the critical the engine. condition determined in § 23.333(d): (c) The limit engine torque to be con- sidered under paragraph (a) of this sec- Dcm=¥0.01d tion must be obtained by multiplying where— the mean torque by a factor of— (1) 1.25 for turbopropeller installa- Dcm is the moment coefficient increment; and tions; d is the down aileron deflection in degrees in (2) 1.33 for engines with five or more the critical condition. cylinders; and [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (3) Two, three, or four, for engines amended by Amdt. 23–7, 34 FR 13088, Aug. 13, with four, three, or two cylinders, re- 1969; Amdt. 23–34, 52 FR 1829, Jan. 15, 1987; spectively. Amdt. 23–48, 61 FR 5144, Feb. 9, 1996] [Amdt. 23–26, 45 FR 60171, Sept. 11, 1980, as § 23.351 Yawing conditions. amended by Amdt. 23–45, 58 FR 42160, Aug. 6, 1993] The airplane must be designed for yawing loads on the vertical surfaces § 23.363 Side load on engine mount. resulting from the loads specified in (a) Each engine mount and its sup- §§ 23.441 through 23.445. porting structure must be designed for a limit load factor in a lateral direc- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 tion, for the side load on the engine FR 258, Jan. 9, 1965, as amended by Amdt. 23– mount, of not less than— 42, 56 FR 352, Jan. 3, 1991] (1) 1.33, or

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(2) One-third of the limit load factor bine or from loss of the turbine blades for flight condition A. are considered to be ultimate loads. (b) The side load prescribed in para- (3) The time history of the thrust graph (a) of this section may be as- decay and drag buildup occurring as a sumed to be independent of other flight result of the prescribed engine failures conditions. must be substantiated by test or other data applicable to the particular en- § 23.365 Pressurized cabin loads. gine-propeller combination. For each pressurized compartment, (4) The timing and magnitude of the the following apply: probable pilot corrective action must (a) The airplane structure must be be conservatively estimated, consid- strong enough to withstand the flight ering the characteristics of the par- loads combined with pressure differen- ticular engine-propeller-airplane com- tial loads from zero up to the max- bination. imum relief valve setting. (b) Pilot corrective action may be as- (b) The external pressure distribution sumed to be initiated at the time max- in flight, and any stress concentra- imum yawing velocity is reached, but tions, must be accounted for. not earlier than 2 seconds after the en- (c) If landings may be made with the gine failure. The magnitude of the cor- cabin pressurized, landing loads must rective action may be based on the be combined with pressure differential limit pilot forces specified in § 23.397 loads from zero up to the maximum al- except that lower forces may be as- lowed during landing. sumed where it is shown by analysis or (d) The airplane structure must be test that these forces can control the strong enough to withstand the pres- yaw and roll resulting from the pre- sure differential loads corresponding to scribed engine failure conditions. the maximum relief valve setting mul- tiplied by a factor of 1.33, omitting [Amdt. 23–7, 34 FR 13089, Aug. 13, 1969] other loads. § 23.369 Rear lift truss. (e) If a pressurized cabin has two or more compartments separated by bulk- (a) If a rear lift truss is used, it must heads or a floor, the primary structure be designed to withstand conditions of must be designed for the effects of sud- reversed airflow at a design speed of— den release of pressure in any compart- V=8.7 √(W/S) + 8.7 (knots), where W/ ment with external doors or windows. S=wing loading at design maximum This condition must be investigated for takeoff weight. the effects of failure of the largest (b) Either aerodynamic data for the opening in the compartment. The ef- particular wing section used, or a value fects of intercompartmental venting of CL equalling ¥0.8 with a chordwise may be considered. distribution that is triangular between a peak at the trailing edge and zero at § 23.367 Unsymmetrical loads due to the leading edge, must be used. engine failure. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (a) Turbopropeller airplanes must be amended by Amdt. 23–7, 34 FR 13089, Aug. 13, designed for the unsymmetrical loads 1969; 34 FR 17509, Oct. 30, 1969; Amdt. 23–45, 58 resulting from the failure of the crit- FR 42160, Aug. 6, 1993; Amdt. 23–48, 61 FR ical engine including the following con- 5145, Feb. 9, 1996] ditions in combination with a single malfunction of the propeller drag lim- § 23.371 Gyroscopic and aerodynamic iting system, considering the probable loads. pilot corrective action on the flight (a) Each engine mount and its sup- controls: porting structure must be designed for (1) At speeds between VMC and VD, the gyroscopic, inertial, and aero- the loads resulting from power failure dynamic loads that result, with the en- because of fuel flow interruption are gine(s) and propeller(s), if applicable, considered to be limit loads. at maximum continuous r.p.m., under (2) At speeds between VMC and VC, the either: loads resulting from the disconnection (1) The conditions prescribed in of the engine compressor from the tur- § 23.351 and § 23.423; or

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(2) All possible combinations of the § 23.393 Loads parallel to hinge line. following— (a) Control surfaces and supporting (i) A yaw velocity of 2.5 radians per hinge brackets must be designed to second; withstand inertial loads acting parallel (ii) A pitch velocity of 1.0 radian per to the hinge line. second; (b) In the absence of more rational (iii) A normal load factor of 2.5; and data, the inertial loads may be as- (iv) Maximum continuous thrust. sumed to be equal to KW, where— (b) For airplanes approved for aero- (1) K=24 for vertical surfaces; batic maneuvers, each engine mount (2) K=12 for horizontal surfaces; and and its supporting structure must meet (3) W=weight of the movable surfaces. the requirements of paragraph (a) of [Doc. No. 27805, 61 FR 5145, Feb. 9, 1996] this section and be designed to with- stand the load factors expected during § 23.395 Control system loads. combined maximum yaw and pitch ve- (a) Each flight control system and its locities. supporting structure must be designed (c) For airplanes certificated in the for loads corresponding to at least 125 commuter category, each engine percent of the computed hinge mo- mount and its supporting structure ments of the movable control surface must meet the requirements of para- in the conditions prescribed in §§ 23.391 graph (a) of this section and the gust through 23.459. In addition, the fol- conditions specified in § 23.341 of this lowing apply: part. (1) The system limit loads need not exceed the higher of the loads that can [Doc. No. 27805, 61 FR 5145, Feb. 9, 1996] be produced by the pilot and automatic devices operating the controls. How- § 23.373 Speed control devices. ever, autopilot forces need not be added If speed control devices (such as to pilot forces. The system must be de- spoilers and drag flaps) are incor- signed for the maximum effort of the porated for use in enroute conditions— pilot or autopilot, whichever is higher. (a) The airplane must be designed for In addition, if the pilot and the auto- the symmetrical maneuvers and gusts pilot act in opposition, the part of the prescribed in §§ 23.333, 23.337, and 23.341, system between them may be designed and the yawing maneuvers and lateral for the maximum effort of the one that gusts in §§ 23.441 and 23.443, with the de- imposes the lesser load. Pilot forces used for design need not exceed the vice extended at speeds up to the maximum forces prescribed in placard device extended speed; and § 23.397(b). (b) If the device has automatic oper- (2) The design must, in any case, pro- ating or load limiting features, the air- vide a rugged system for service use, plane must be designed for the maneu- considering jamming, ground gusts, ver and gust conditions prescribed in taxiing downwind, control inertia, and paragraph (a) of this section at the friction. Compliance with this subpara- speeds and corresponding device posi- graph may be shown by designing for tions that the mechanism allows. loads resulting from application of the [Amdt. 23–7, 34 FR 13089, Aug. 13, 1969] minimum forces prescribed in § 23.397(b). CONTROL SURFACE AND SYSTEM LOADS (b) A 125 percent factor on computed hinge moments must be used to design § 23.391 Control surface loads. elevator, aileron, and rudder systems. However, a factor as low as 1.0 may be The control surface loads specified in used if hinge moments are based on ac- §§ 23.397 through 23.459 are assumed to curate flight test data, the exact reduc- occur in the conditions described in tion depending upon the accuracy and §§ 23.331 through 23.351. reliability of the data. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (c) Pilot forces used for design are as- amended by Amdt. 23–48, 61 FR 5145, Feb. 9, sumed to act at the appropriate control 1996] grips or pads as they would in flight,

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and to react at the attachments of the § 23.399 Dual control system. control system to the control surface (a) Each dual control system must be horns. designed to withstand the force of the [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as pilots operating in opposition, using in- amended by Amdt. 23–7, 34 FR 13089, Aug. 13, dividual pilot forces not less than the 1969] greater of— (1) 0.75 times those obtained under § 23.397 Limit control forces and § 23.395; or torques. (2) The minimum forces specified in (a) In the control surface flight load- § 23.397(b). ing condition, the airloads on movable (b) Each dual control system must be surfaces and the corresponding deflec- designed to withstand the force of the tions need not exceed those that would pilots applied together, in the same di- result in flight from the application of rection, using individual pilot forces any pilot force within the ranges speci- not less than 0.75 times those obtained fied in paragraph (b) of this section. In under § 23.395. applying this criterion, the effects of [Doc. No. 27805, 61 FR 5145, Feb. 9, 1996] control system boost and servo-mecha- nisms, and the effects of tabs must be § 23.405 Secondary control system. considered. The automatic pilot effort Secondary controls, such as wheel must be used for design if it alone can brakes, spoilers, and tab controls, must produce higher control surface loads be designed for the maximum forces than the human pilot. that a pilot is likely to apply to those (b) The limit pilot forces and torques controls. are as follows: § 23.407 Trim tab effects. Maximum forces The effects of trim tabs on the con- or torques for Minimum Control design weight, forces or trol surface design conditions must be weight equal to 2 or less than torques accounted for only where the surface 5,000 pounds 1 loads are limited by maximum pilot ef- fort. In these cases, the tabs are con- Aileron: Stick ...... 67 lbs ...... 40 lbs. sidered to be deflected in the direction Wheel 3 ...... 50 D in.-lbs 4 ..... 40 D in.- that would assist the pilot. These de- lbs.4 flections must correspond to the max- Elevator: imum degree of ‘‘out of trim’’ expected Stick ...... 167 lbs ...... 100 lbs. at the speed for the condition under Wheel (symmetrical) ...... 200 lbs ...... 100 lbs. Wheel (unsymmetrical) 5 ...... 100 lbs. consideration. Rudder ...... 200 lbs ...... 150 lbs. § 23.409 Tabs. 1 For design weight (W) more than 5,000 pounds, the speci- fied maximum values must be increased linearly with weight Control surface tabs must be de- to 1.18 times the specified values at a design weight of 12,500 pounds and for commuter category airplanes, the signed for the most severe combination specified values must be increased linearly with weight to of airspeed and tab deflection likely to 1.35 times the specified values at a design weight of 19,000 pounds. be obtained within the flight envelope 2 If the design of any individual set of control systems or for any usable loading condition. surfaces makes these specified minimum forces or torques in- applicable, values corresponding to the present hinge mo- ments obtained under § 23.415, but not less than 0.6 of the § 23.415 Ground gust conditions. specified minimum forces or torques, may be used. 3 The critical parts of the aileron control system must also (a) The control system must be inves- be designed for a single tangential force with a limit value of tigated as follows for control surface 1.25 times the couple force determined from the above cri- teria. loads due to ground gusts and taxiing 4 D=wheel diameter (inches). downwind: 5 The unsymmetrical force must be applied at one of the (1) If an investigation of the control normal handgrip points on the control wheel. system for ground gust loads is not re- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as quired by paragraph (a)(2) of this sec- amended by Amdt. 23–7, 34 FR 13089, Aug. 13, tion, but the applicant elects to design 1969; Amdt. 23–17, 41 FR 55464, Dec. 20, 1976; a part of the control system of these Amdt. 23–34, 52 FR 1829, Jan. 15, 1987; Amdt. loads, these loads need only be carried 23–45, 58 FR 42160, Aug. 6, 1993] from control surface horns through the

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nearest stops or gust locks and their HORIZONTAL STABILIZING AND supporting structures. BALANCING SURFACES (2) If pilot forces less than the mini- mums specified in § 23.397(b) are used § 23.421 Balancing loads. for design, the effects of surface loads (a) A horizontal surface balancing due to ground gusts and taxiing down- load is a load necessary to maintain wind must be investigated for the en- equilibrium in any specified flight con- tire control system according to the dition with no pitching acceleration. formula: (b) Horizontal balancing surfaces H=K c S q must be designed for the balancing loads occurring at any point on the where— limit maneuvering envelope and in the H=limit hinge moment (ft.-lbs.); flap conditions specified in § 23.345. c=mean chord of the control surface aft of [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as the hinge line (ft.); amended by Amdt. 23–7, 34 FR 13089, Aug. 13, S=area of control surface aft of the hinge 1969; Amdt. 23–42, 56 FR 352, Jan. 3, 1991] line (sq. ft.); q=dynamic pressure (p.s.f.) based on a design § 23.423 Maneuvering loads. speed not less than 14.6 √(W/S) + 14.6 (f.p.s.) where W/S=wing loading at design max- Each horizontal surface and its sup- imum weight, except that the design speed porting structure, and the main wing need not exceed 88 (f.p.s.); of a canard or tandem wing configura- K=limit hinge moment factor for ground tion, if that surface has pitch control, gusts derived in paragraph (b) of this sec- must be designed for the maneuvering tion. (For ailerons and elevators, a positive loads imposed by the following condi- value of K indicates a moment tending to depress the surface and a negative value of tions: K indicates a moment tending to raise the (a) A sudden movement of the pitch- surface). ing control, at the speed VA, to the maximum aft movement, and the max- (b) The limit hinge moment factor K imum forward movement, as limited by for ground gusts must be derived as fol- the control stops, or pilot effort, lows: whichever is critical. Surface K Position of controls (b) A sudden aft movement of the pitching control at speeds above VA, (a) Aileron ...... 0.75 Control column locked lashed in followed by a forward movement of the mid-position. pitching control resulting in the fol- (b) Aileron ...... ±0.50 Ailerons at full throw; + moment on one aileron, ¥ moment on lowing combinations of normal and an- the other. gular acceleration: (c) Elevator ...... ±0.75 (c) Elevator full up (¥). (d) Elevator ...... (d) Elevator full down (+). Normal Angular acceleration (e) Rudder ...... ±0.75 (e) Rudder in neutral. Condition accelera- tion (n) (radian/sec2) (f) Rudder ...... (f) Rudder at full throw.

Nose-up pitching ...... 1.0 +39nm÷V×(nm¥1.5) (c) At all weights between the empty Nose-down pitching .... nm ¥39nm÷V×(nm¥1.5) weight and the maximum weight de- clared for tie-down stated in the appro- where—

priate manual, any declared tie-down (1) nm=positive limit maneuvering points and surrounding structure, con- load factor used in the design of the trol system, surfaces and associated airplane; and gust locks, must be designed to with- (2) V=initial speed in knots. stand the limit load conditions that The conditions in this paragraph in- exist when the airplane is tied down volve loads corresponding to the loads and that result from wind speeds of up that may occur in a ‘‘checked maneu- to 65 knots horizontally from any di- ver’’ (a maneuver in which the pitching rection. control is suddenly displaced in one di- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as rection and then suddenly moved in the amended by Amdt. 23–7, 34 FR 13089, Aug. 13, opposite direction). The deflections and 1969; Amdt. 23–45, 58 FR 42160, Aug. 6, 1993; timing of the ‘‘checked maneuver’’ must Amdt. 23–48, 61 FR 5145, Feb. 9, 1996] avoid exceeding the limit maneuvering

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load factor. The total horizontal sur- § 23.427 Unsymmetrical loads. face load for both nose-up and nose- (a) Horizontal surfaces other than down pitching conditions is the sum of main wing and their supporting struc- the balancing loads at V and the speci- ture must be designed for unsymmet- fied value of the normal load factor n, rical loads arising from yawing and plus the maneuvering load increment slipstream effects, in combination with due to the specified value of the angu- the loads prescribed for the flight con- lar acceleration. ditions set forth in §§ 23.421 through [Amdt. 23–42, 56 FR 353, Jan. 3, 1991; 56 FR 23.425. 5455, Feb. 11, 1991] (b) In the absence of more rational data for airplanes that are conven- § 23.425 Gust loads. tional in regard to location of engines, (a) Each horizontal surface, other wings, horizontal surfaces other than than a main wing, must be designed for main wing, and fuselage shape: loads resulting from— (1) 100 percent of the maximum load- (1) Gust velocities specified in ing from the symmetrical flight condi- § 23.333(c) with flaps retracted; and tions may be assumed on the surface (2) Positive and negative gusts of 25 on one side of the plane of symmetry; f.p.s. nominal intensity at VF cor- and responding to the flight conditions (2) The following percentage of that specified in § 23.345(a)(2). loading must be applied to the opposite (b) [Reserved] side: (c) When determining the total load Percent=100¥10 (n¥1), where n is the spec- on the horizontal surfaces for the con- ified positive maneuvering load factor, but ditions specified in paragraph (a) of this value may not be more than 80 percent. this section, the initial balancing loads for steady unaccelerated flight at the (c) For airplanes that are not conven- tional (such as airplanes with hori- pertinent design speeds VF, VC, and VD must first be determined. The incre- zontal surfaces other than main wing mental load resulting from the gusts having appreciable dihedral or sup- must be added to the initial balancing ported by the vertical tail surfaces) the load to obtain the total load. surfaces and supporting structures (d) In the absence of a more rational must be designed for combined vertical analysis, the incremental load due to and horizontal surface loads resulting the gust must be computed as follows from each prescribed flight condition only on airplane configurations with taken separately. aft-mounted, horizontal surfaces, un- [Amdt. 23–14, 38 FR 31820, Nov. 19, 1973, as less its use elsewhere is shown to be amended by Amdt. 23–42, 56 FR 353, Jan. 3, conservative: 1991]

K U Va S ⎛ dε ⎞ VERTICAL SURFACES Δ =−gdehtht L ht ⎜1 ⎟ § 23.441 Maneuvering loads. 498 ⎝ dα ⎠ (a) At speeds up to V the vertical where— A, surfaces must be designed to withstand DLht=Incremental horizontal tailload (lbs.); K =Gust alleviation factor defined in § 23.341; the following conditions. In computing g the loads, the yawing velocity may be Ude=Derived gust velocity (f.p.s.); V=Airplane equivalent speed (knots); assumed to be zero: aht=Slope of aft horizontal lift curve (per ra- (1) With the airplane in unacceler- dian) ated flight at zero yaw, it is assumed 2 Sht=Area of aft horizontal lift surface (ft ); that the rudder control is suddenly dis- and placed to the maximum deflection, as limited by the control stops or by limit ⎛ dε ⎞ pilot forces. ⎜1− ⎟ = Downwash factor ⎝ α ⎠ (2) With the rudder deflected as speci- d fied in paragraph (a)(1) of this section, [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as it is assumed that the airplane yaws to amended by Amdt. 23–7, 34 FR 13089 Aug. 13, the overswing sideslip angle. In lieu of 1969; Amdt. 23–42, 56 FR 353, Jan. 3, 1991] a rational analysis, an overswing angle

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equal to 1.5 times the static sideslip (1) The airplane must be yawed to the angle of paragraph (a)(3) of this section largest attainable steady state sideslip may be assumed. angle, with the rudder at maximum de- (3) A yaw angle of 15 degrees with the flection caused by any one of the fol- rudder control maintained in the neu- lowing: tral position (except as limited by pilot (i) Control surface stops; strength). (ii) Maximum available booster ef- (b) For commuter category airplanes, fort; the loads imposed by the following ad- (iii) Maximum pilot rudder force as ditional maneuver must be substan- shown below: tiated at speeds from VA to VD/MD. When computing the tail loads—

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(2) The rudder must be suddenly dis- (c) The yaw angles specified in para- placed from the maximum deflection to graph (a)(3) of this section may be re- the neutral position. duced if the yaw angle chosen for a par- ticular speed cannot be exceeded in—

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(1) Steady slip conditions; V=Equivalent airspeed (knots). (2) Uncoordinated rolls from steep [Amdt. 23–7, 34 FR 13090, Aug. 13, 1969, as banks; or amended by Amdt. 23–34, 52 FR 1830, Jan. 15, (3) Sudden failure of the critical en- 1987; 52 FR 7262, Mar. 9, 1987; Amdt. 23–24, 52 gine with delayed corrective action. FR 34745, Sept. 14, 1987; Amdt. 23–42, 56 FR 353, Jan. 3, 1991; Amdt. 23–48, 61 FR 5147, Feb. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as 9, 1996] amended by Amdt. 23–7, 34 FR 13090, Aug. 13, 1969; Amdt. 23–14, 38 FR 31821, Nov. 19, 1973; § 23.445 Outboard fins or winglets. Amdt. 23–28, 47 FR 13315, Mar. 29, 1982; Amdt. 23–42, 56 FR 353, Jan. 3, 1991; Amdt. 23–48, 61 (a) If outboard fins or winglets are in- FR 5145, Feb. 9, 1996] cluded on the horizontal surfaces or wings, the horizontal surfaces or wings § 23.443 Gust loads. must be designed for their maximum load in combination with loads induced (a) Vertical surfaces must be de- by the fins or winglets and moments or signed to withstand, in unaccelerated forces exerted on the horizontal sur- flight at speed VC, lateral gusts of the faces or wings by the fins or winglets. values prescribed for V in § 23.333(c). C (b) If outboard fins or winglets ex- (b) In addition, for commuter cat- tend above and below the horizontal egory airplanes, the airplane is as- surface, the critical vertical surface sumed to encounter derived gusts nor- loading (the load per unit area as de- mal to the plane of symmetry while in termined under §§ 23.441 and 23.443) unaccelerated flight at VB, VC, VD, and must be applied to— VF. The derived gusts and airplane (1) The part of the vertical surfaces speeds corresponding to these condi- above the horizontal surface with 80 tions, as determined by §§ 23.341 and percent of that loading applied to the 23.345, must be investigated. The shape part below the horizontal surface; and of the gust must be as specified in (2) The part of the vertical surfaces § 23.333(c)(2)(i). below the horizontal surface with 80 (c) In the absence of a more rational percent of that loading applied to the analysis, the gust load must be com- part above the horizontal surface. puted as follows: (c) The end plate effects of outboard KUVaS fins or winglets must be taken into ac- = gt de vt vt count in applying the yawing condi- L vt tions of §§ 23.441 and 23.443 to the 498 vertical surfaces in paragraph (b) of Where— this section. Lvt=Vertical surface loads (lbs.); (d) When rational methods are used 088. μ for computing loads, the maneuvering k = gt = gust alleviation factor; loads of § 23.441 on the vertical surfaces gt 53. + μ and the one-g horizontal surface load, gt including induced loads on the hori- zontal surface and moments or forces 2W K 2 μ ==lateralmassratio; exerted on the horizontal surfaces by gt ρ the vertical surfaces, must be applied cgatvtvtvt S l simultaneously for the structural load-

Ude=Derived gust velocity (f.p.s.); ing condition. r=Air density (slugs/cu.ft.); [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as W=the applicable weight of the airplane in amended by Amdt. 23–14, 38 FR 31821, Nov. 19, the particular load case (lbs.); 1973; Amdt. 23–42, 56 FR 353, Jan. 3, 1991] 2 Svt=Area of vertical surface (ft. ); ˘≤ c t=Mean geometric chord of vertical surface AILERONS AND SPECIAL DEVICES (ft.); avt=Lift curve slope of vertical surface (per § 23.455 Ailerons. radian); K=Radius of gyration in yaw (ft.); (a) The ailerons must be designed for the loads to which they are subjected— lvt=Distance from airplane c.g. to lift center of vertical surface (ft.); (1) In the neutral position during g=Acceleration due to gravity (ft./sec.2); and symmetrical flight conditions; and

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(2) By the following deflections (ex- maximum weight and the design land- cept as limited by pilot effort), during ing weight; or unsymmetrical flight conditions: (2) The design maximum weight less (i) Sudden maximum displacement of the weight of 25 percent of the total the aileron control at VA. Suitable al- fuel capacity. lowance may be made for control sys- (c) The design landing weight of a tem deflections. multiengine airplane may be less than (ii) Sufficient deflection at VC, where that allowed under paragraph (b) of VC is more than VA, to produce a rate this section if— of roll not less than obtained in para- (1) The airplane meets the one-en- graph (a)(2)(i) of this section. gine-inoperative climb requirements of (iii) Sufficient deflection at VD to § 23.67(b)(1) or (c); and produce a rate of roll not less than one- (2) Compliance is shown with the fuel third of that obtained in paragraph jettisoning system requirements of (a)(2)(i) of this section. § 23.1001. (b) [Reserved] (d) The selected limit vertical inertia [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as load factor at the center of gravity of amended by Amdt. 23–7, 34 FR 13090, Aug. 13, the airplane for the ground load condi- 1969; Amdt. 23–42, 56 FR 353, Jan. 3, 1991] tions prescribed in this subpart may not be less than that which would be § 23.459 Special devices. obtained when landing with a descent The loading for special devices using velocity (V), in feet per second, equal aerodynamic surfaces (such as slots to 4.4 (W/S)1⁄4, except that this velocity and spoilers) must be determined from need not be more than 10 feet per sec- test data. ond and may not be less than seven feet per second. GROUND LOADS (e) Wing lift not exceeding two-thirds of the weight of the airplane may be § 23.471 General. assumed to exist throughout the land- The limit ground loads specified in ing impact and to act through the cen- this subpart are considered to be exter- ter of gravity. The ground reaction nal loads and inertia forces that act load factor may be equal to the inertia upon an airplane structure. In each load factor minus the ratio of the specified ground load condition, the ex- above assumed wing lift to the airplane ternal reactions must be placed in weight. equilibrium with the linear and angu- (f) If energy absorption tests are lar inertia forces in a rational or con- made to determine the limit load fac- servative manner. tor corresponding to the required limit descent velocities, these tests must be § 23.473 Ground load conditions and made under § 23.723(a). assumptions. (g) No inertia load factor used for de- (a) The ground load requirements of sign purposes may be less than 2.67, nor this subpart must be complied with at may the limit ground reaction load fac- the design maximum weight except tor be less than 2.0 at design maximum that §§ 23.479, 23.481, and 23.483 may be weight, unless these lower values will complied with at a design landing not be exceeded in taxiing at speeds up weight (the highest weight for landing to takeoff speed over terrain as rough conditions at the maximum descent ve- as that expected in service. locity) allowed under paragraphs (b) [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as and (c) of this section. amended by Amdt. 23–7, 34 FR 13090, Aug. 13, (b) The design landing weight may be 1969; Amdt. 23–28, 47 FR 13315, Mar. 29, 1982; as low as— Amdt. 23–45, 58 FR 42160, Aug. 6, 1993; Amdt. (1) 95 percent of the maximum weight 23–48, 61 FR 5147, Feb. 9, 1996] if the minimum fuel capacity is enough for at least one-half hour of operation § 23.477 Landing gear arrangement. at maximum continuous power plus a Sections 23.479 through 23.483, or the capacity equal to a fuel weight which conditions in appendix C, apply to air- is the difference between the design planes with conventional arrangements

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of main and nose gear, or main and tail sponse, an airplane lift equal to the gear. weight of the airplane may be assumed.

§ 23.479 Level landing conditions. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–17, 41 FR 55464, Dec. 20, (a) For a level landing, the airplane 1976; Amdt. 23–45, 58 FR 42160, Aug. 6, 1993] is assumed to be in the following atti- tudes: § 23.481 Tail down landing conditions. (1) For airplanes with tail wheels, a (a) For a tail down landing, the air- normal level flight attitude. plane is assumed to be in the following (2) For airplanes with nose wheels, attitudes: attitudes in which— (1) For airplanes with tail wheels, an (i) The nose and main wheels contact attitude in which the main and tail the ground simultaneously; and wheels contact the ground simulta- (ii) The main wheels contact the neously. ground and the nose wheel is just clear (2) For airplanes with nose wheels, a of the ground. stalling attitude, or the maximum The attitude used in paragraph (a)(2)(i) angle allowing ground clearance by of this section may be used in the anal- each part of the airplane, whichever is ysis required under paragraph (a)(2)(ii) less. of this section. (b) For airplanes with either tail or (b) When investigating landing condi- nose wheels, ground reactions are as- tions, the drag components simulating sumed to be vertical, with the wheels the forces required to accelerate the up to speed before the maximum tires and wheels up to the landing vertical load is attained. speed (spin-up) must be properly com- bined with the corresponding instanta- § 23.483 One-wheel landing conditions. neous vertical ground reactions, and For the one-wheel landing condition, the forward-acting horizontal loads re- the airplane is assumed to be in the sulting from rapid reduction of the level attitude and to contact the spin-up drag loads (spring-back) must ground on one side of the main landing be combined with vertical ground reac- gear. In this attitude, the ground reac- tions at the instant of the peak for- tions must be the same as those ob- ward load, assuming wing lift and a tained on that side under § 23.479. tire-sliding coefficient of friction of 0.8. However, the drag loads may not be § 23.485 Side load conditions. less than 25 percent of the maximum vertical ground reactions (neglecting (a) For the side load condition, the wing lift). airplane is assumed to be in a level at- titude with only the main wheels con- (c) In the absence of specific tests or tacting the ground and with the shock a more rational analysis for deter- mining the wheel spin-up and spring- absorbers and tires in their static posi- back loads for landing conditions, the tions. method set forth in appendix D of this (b) The limit vertical load factor part must be used. If appendix D of this must be 1.33, with the vertical ground part is used, the drag components used reaction divided equally between the for design must not be less than those main wheels. given by appendix C of this part. (c) The limit side inertia factor must (d) For airplanes with tip tanks or be 0.83, with the side ground reaction large overhung masses (such as turbo- divided between the main wheels so propeller or jet engines) supported by that— the wing, the tip tanks and the struc- (1) 0.5 (W) is acting inboard on one ture supporting the tanks or overhung side; and masses must be designed for the effects (2) 0.33 (W) is acting outboard on the of dynamic responses under the level other side. landing conditions of either paragraph (d) The side loads prescribed in para- (a)(1) or (a)(2)(ii) of this section. In graph (c) of this section are assumed to evaluating the effects of dynamic re- be applied at the ground contact point

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and the drag loads may be assumed to (1) Suitable design loads must be es- be zero. tablished for the tail wheel, bumper, or energy absorption device; and [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–45, 58 FR 42160, Aug. 6, (2) The supporting structure of the 1993] tail wheel, bumper, or energy absorp- tion device must be designed to with- § 23.493 Braked roll conditions. stand the loads established in para- graph (c)(1) of this section. Under braked roll conditions, with the shock absorbers and tires in their [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as static positions, the following apply: amended by Amdt. 23–48, 61 FR 5147, Feb. 9, (a) The limit vertical load factor 1996] must be 1.33. (b) The attitudes and ground con- § 23.499 Supplementary conditions for nose wheels. tacts must be those described in § 23.479 for level landings. In determining the ground loads on (c) A drag reaction equal to the nose wheels and affected supporting vertical reaction at the wheel multi- structures, and assuming that the plied by a coefficient of friction of 0.8 shock absorbers and tires are in their must be applied at the ground contact static positions, the following condi- point of each wheel with brakes, except tions must be met: that the drag reaction need not exceed (a) For aft loads, the limit force com- the maximum value based on limiting ponents at the axle must be— brake torque. (1) A vertical component of 2.25 times the static load on the wheel; and § 23.497 Supplementary conditions for (2) A drag component of 0.8 times the tail wheels. vertical load. In determining the ground loads on (b) For forward loads, the limit force the tail wheel and affected supporting components at the axle must be— structures, the following apply: (1) A vertical component of 2.25 times (a) For the obstruction load, the the static load on the wheel; and limit ground reaction obtained in the (2) A forward component of 0.4 times tail down landing condition is assumed the vertical load. to act up and aft through the axle at 45 (c) For side loads, the limit force degrees. The shock absorber and tire components at ground contact must may be assumed to be in their static be— positions. (1) A vertical component of 2.25 times (b) For the side load, a limit vertical the static load on the wheel; and ground reaction equal to the static (2) A side component of 0.7 times the load on the tail wheel, in combination vertical load. with a side component of equal mag- (d) For airplanes with a steerable nitude, is assumed. In addition— nose wheel that is controlled by hy- (1) If a swivel is used, the tail wheel draulic or other power, at design take- is assumed to be swiveled 90 degrees to off weight with the nose wheel in any the airplane longitudinal axis with the steerable position, the application of resultant ground load passing through 1.33 times the full steering torque com- the axle; bined with a vertical reaction equal to (2) If a lock, steering device, or shim- 1.33 times the maximum static reaction my damper is used, the tail wheel is on the nose gear must be assumed. also assumed to be in the trailing posi- However, if a torque limiting device is tion with the side load acting at the installed, the steering torque can be re- ground contact point; and duced to the maximum value allowed (3) The shock absorber and tire are by that device. assumed to be in their static positions. (e) For airplanes with a steerable (c) If a tail wheel, bumper, or an en- nose wheel that has a direct mechan- ergy absorption device is provided to ical connection to the rudder pedals, show compliance with § 23.925(b), the the mechanism must be designed to following apply: withstand the steering torque for the

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maximum pilot forces specified in and their immediate attaching struc- § 23.397(b). ture. (a) The towing loads specified in [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–48, 61 FR 5147, Feb. 9, paragraph (d) of this section must be 1996] considered separately. These loads must be applied at the towing fittings § 23.505 Supplementary conditions for and must act parallel to the ground. In skiplanes. addition: In determining ground loads for ski- (1) A vertical load factor equal to 1.0 planes, and assuming that the airplane must be considered acting at the center is resting on the ground with one main of gravity; and ski frozen at rest and the other skis (2) The shock struts and tires must free to slide, a limit side force equal to be in there static positions. 0.036 times the design maximum weight (b) For towing points not on the must be applied near the tail assembly, landing gear but near the plane of sym- with a factor of safety of 1. metry of the airplane, the drag and side tow load components specified for [Amdt. 23–7, 34 FR 13090, Aug. 13, 1969] the auxiliary gear apply. For towing points located outboard of the main § 23.507 Jacking loads. gear, the drag and side tow load compo- (a) The airplane must be designed for nents specified for the main gear apply. the loads developed when the aircraft Where the specified angle of swivel is supported on jacks at the design cannot be reached, the maximum ob- maximum weight assuming the fol- tainable angle must be used. lowing load factors for landing gear (c) The towing loads specified in jacking points at a three-point attitude paragraph (d) of this section must be and for primary flight structure jack- reacted as follows: ing points in the level attitude: (1) The side component of the towing (1) Vertical-load factor of 1.35 times load at the main gear must be reacted the static reactions. by a side force at the static ground line (2) Fore, aft, and lateral load factors of the wheel to which the load is ap- of 0.4 times the vertical static reac- plied. tions. (2) The towing loads at the auxiliary (b) The horizontal loads at the jack gear and the drag components of the points must be reacted by inertia towing loads at the main gear must be forces so as to result in no change in reacted as follows: the direction of the resultant loads at (i) A reaction with a maximum value the jack points. equal to the vertical reaction must be (c) The horizontal loads must be con- applied at the axle of the wheel to sidered in all combinations with the which the load is applied. Enough air- vertical load. plane inertia to achieve equilibrium must be applied. [Amdt. 23–14, 38 FR 31821, Nov. 19, 1973] (ii) The loads must be reacted by air- plane inertia. § 23.509 Towing loads. (d) The prescribed towing loads are as The towing loads of this section must follows, where W is the design max- be applied to the design of tow fittings imum weight:

Load Tow point Position Magnitude No. Direction

Main gear ...... 0.225W 1 Forward, parallel to drag axis. 2 Forward, at 30° to drag axis. 3 Aft, parallel to drag axis. 4 Aft, at 30° to drag axis.

Auxiliary gear ...... Swiveled forward ...... 0.3W 5 Forward. 6 Aft. Swiveled aft ...... 0.3W 7 Forward. 8 Aft.

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Load Tow point Position Magnitude No. Direction

Swiveled 45° from forward ...... 0.15W 9 Forward, in plane of wheel. 10 Aft, in plane of wheel. Swiveled 45° from aft ...... 0.15W 11 Forward, in plane of wheel. 12 Aft, in plane of wheel.

[Amdt. 23–14, 38 FR 31821, Nov. 19, 1973] (b) Unless the applicant makes a ra- tional analysis of the water loads, § 23.511 Ground load; unsymmetrical §§ 23.523 through 23.537 apply. loads on multiple-wheel units. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (a) Pivoting loads. The airplane is as- amended by Amdt. 23–45, 58 FR 42160, Aug. 6, sumed to pivot about on side of the 1993; Amdt. 23–48, 61 FR 5147, Feb. 9, 1996] main gear with— (1) The brakes on the pivoting unit § 23.523 Design weights and center of locked; and gravity positions. (2) Loads corresponding to a limit (a) Design weights. The water load re- vertical load factor of 1, and coefficient quirements must be met at each oper- of friction of 0.8 applied to the main ating weight up to the design landing gear and its supporting structure. weight except that, for the takeoff con- (b) Unequal tire loads. The loads es- dition prescribed in § 23.531, the design tablished under §§ 23.471 through 23.483 water takeoff weight (the maximum must be applied in turn, in a 60/40 per- weight for water taxi and takeoff run) cent distribution, to the dual wheels must be used. and tires in each dual wheel landing (b) Center of gravity positions. The gear unit. critical centers of gravity within the (c) Deflated tire loads. For the deflated limits for which certification is re- tire condition— quested must be considered to reach (1) 60 percent of the loads established maximum design loads for each part of under §§ 23.471 through 23.483 must be the seaplane structure. applied in turn to each wheel in a land- ing gear unit; and [Doc. No. 26269, 58 FR 42160, Aug. 6, 1993] (2) 60 percent of the limit drag and § 23.525 Application of loads. side loads, and 100 percent of the limit vertical load established under §§ 23.485 (a) Unless otherwise prescribed, the and 23.493 or lesser vertical load ob- seaplane as a whole is assumed to be tained under paragraph (c)(1) of this subjected to the loads corresponding to section, must be applied in turn to the load factors specified in § 23.527. each wheel in the dual wheel landing (b) In applying the loads resulting gear unit. from the load factors prescribed in § 23.527, the loads may be distributed [Amdt. 23–7, 34 FR 13090, Aug. 13, 1969] over the hull or main float bottom (in order to avoid excessive local shear WATER LOADS loads and bending moments at the lo- § 23.521 Water load conditions. cation of water load application) using pressures not less than those pre- (a) The structure of seaplanes and scribed in § 23.533(c). amphibians must be designed for water (c) For twin float seaplanes, each loads developed during takeoff and float must be treated as an equivalent landing with the seaplane in any atti- hull on a fictitious seaplane with a tude likely to occur in normal oper- weight equal to one-half the weight of ation at appropriate forward and sink- the twin float seaplane. ing velocities under the most severe (d) Except in the takeoff condition of sea conditions likely to be encoun- § 23.531, the aerodynamic lift on the tered.

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seaplane during the impact is assumed factor K1 may be reduced at the bow to be 2⁄3 of the weight of the seaplane. and stern to 0.8 of the value shown in figure 2 of appendix I of this part. This [Doc. No. 26269, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993] reduction applies only to the design of the carrythrough and seaplane struc- § 23.527 Hull and main float load fac- ture. tors. [Doc. No. 26269, 58 FR 42161, Aug. 6, 1993; 58 (a) Water reaction load factors nw FR 51970, Oct. 5, 1993] must be computed in the following manner: § 23.529 Hull and main float landing (1) For the step landing case conditions. (a) Symmetrical step, bow, and stern CV 2 landing. For symmetrical step, bow, n = 1SO and stern landings, the limit water re- w ⎛ 2 ⎞ 1 Tan3 β W 3 action load factors are those computed ⎝ ⎠ under § 23.527. In addition— (2) For the bow and stern landing (1) For symmetrical step landings, cases the resultant water load must be ap- plied at the keel, through the center of CV 2 K gravity, and must be directed per- n = 1SO × 1 pendicularly to the keel line; w ⎛ 2 ⎞ 1 2 (2) For symmetrical bow landings, ⎜ 3 β⎟ 3 ()1r+ 2 3 ⎝Tan⎠ W x the resultant water load must be ap- plied 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. (3) VSO=seaplane stalling speed in (b) Unsymmetrical landing for hull and knots with flaps extended in the appro- single float seaplanes. Unsymmetrical priate landing position and with no step, bow, and stern landing conditions slipstream effect. must be investigated. In addition— (4) b=Angle of dead rise at the longi- (1) The loading for each condition tudinal station at which the load fac- consists of an upward component and a tor is being determined in accordance side component equal, respectively, to with figure 1 of appendix I of this part. 0.75 and 0.25 tan b times the resultant (5) W=seaplane landing weight in load in the corresponding symmetrical pounds. landing condition; and (6) K1=empirical hull station weigh- (2) The point of application and di- ing factor, in accordance with figure 2 rection of the upward component of the of appendix I of this part. load is the same as that in the sym- (7) rx=ratio of distance, measured metrical condition, and the point of ap- parallel to hull reference axis, from the plication of the side component is at center of gravity of the seaplane to the the same longitudinal station as the hull longitudinal station at which the upward component but is directed in- load factor is being computed to the ra- ward perpendicularly to the plane of dius of gyration in pitch of the sea- symmetry at a point midway between plane, the hull reference axis being a the keel and chine lines. straight line, in the plane of sym- (c) Unsymmetrical landing; twin float metry, tangential to the keel at the seaplanes. The unsymmetrical loading main step. consists of an upward load at the step (c) For a twin float seaplane, because of each float of 0.75 and a side load of of the effect of flexibility of the attach- 0.25 tan b at one float times the step ment of the floats to the seaplane, the landing load reached under § 23.527. The

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side load is directed inboard, per- Pk=pressure (p.s.i.) at the keel; pendicularly to the plane of symmetry C2=0.00213; midway between the keel and chine K2=hull station weighing factor, in accord- ance with figure 2 of appendix I of this lines of the float, at the same longitu- part; dinal station as the upward load. VS1=seaplane stalling speed (knots) at the de- [Doc. No. 26269, 58 FR 42161, Aug. 6, 1993] sign water takeoff weight with flaps ex- tended in the appropriate takeoff position; § 23.531 Hull and main float takeoff and condition. bK=angle of dead rise at keel, in accordance with figure 1 of appendix I of this part. For the wing and its attachment to (2) For a flared bottom, the pressure the hull or main float— at the beginning of the flare is the (a) The aerodynamic wing lift is as- same as that for an unflared bottom, sumed to be zero; and and the pressure between the chine and (b) A downward inertia load, cor- the beginning of the flare varies lin- responding to a load factor computed early, in accordance with figure 3 of ap- from the following formula, must be pendix I of this part. The pressure dis- applied: tribution is the same as that prescribed in paragraph (b)(1) of this section for CV2 n = TO S1 an unflared bottom except that the ⎛ 2 ⎞ 1 pressure at the chine is computed as 3 β 3 ⎝Tan⎠ W follows: Where— CKV2 = 32S1 n=inertia load factor; Pch CTO=empirical seaplane operations factor Tan β equal to 0.004; where— VS1=seaplane stalling speed (knots) at the de- sign takeoff weight with the flaps extended Pch=pressure (p.s.i.) at the chine; C =0.0016; in the appropriate takeoff position; 3 K =hull station weighing factor, in accord- b=angle of dead rise at the main step (de- 2 ance with figure 2 of appendix I of this grees); and part; W=design water takeoff weight in pounds. VS1=seaplane stalling speed (knots) at the de- [Doc. No. 26269, 58 FR 42161, Aug. 6, 1993] sign water takeoff weight with flaps ex- tended in the appropriate takeoff position; § 23.533 Hull and main float bottom and pressures. b=angle of dead rise at appropriate station. (a) General. The hull and main float The area over which these pressures structure, including frames and bulk- are applied must simulate pressures oc- heads, stringers, and bottom plating, curring during high localized impacts must be designed under this section. on the hull or float, but need not ex- (b) Local pressures. For the design of tend over an area that would induce the bottom plating and stringers and critical stresses in the frames or in the their attachments to the supporting overall structure. structure, the following pressure dis- (c) Distributed pressures. For the de- tributions must be applied: sign of the frames, keel, and chine (1) For an unflared bottom, the pres- structure, the following pressure dis- sure at the chine is 0.75 times the pres- tributions apply: sure at the keel, and the pressures be- (1) Symmetrical pressures are com- tween the keel and chine vary linearly, puted as follows: in accordance with figure 3 of appendix 2 I of this part. The pressure at the keel CKV42SO (p.s.i.) is computed as follows: P = Tan β CKV 2 where— P = 22S1 P=pressure (p.s.i.); K C4=0.078 C1 (with C1 computed under § 23.527); Tan β K2=hull station weighing factor, determined k in accordance with figure 2 of appendix I of where— this part;

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3 VS0=seaplane stalling speed (knots) with bs=angle of dead rise at a station ⁄4 of the landing flaps extended in the appropriate distance from the bow to the step, but need position and with no slipstream effect; and not be less than 15 degrees; and b=angle of dead rise at appropriate station. ry=ratio of the lateral distance between the center of gravity and the plane of sym- (2) The unsymmetrical pressure dis- metry of the float to the radius of gyration tribution consists of the pressures pre- in roll. scribed in paragraph (c)(1) of this sec- (c) Bow loading. The resultant limit tion on one side of the hull or main load must be applied in the plane of float centerline and one-half of that symmetry of the float at a point one- pressure on the other side of the hull or fourth of the distance from the bow to main float centerline, in accordance the step and must be perpendicular to with figure 3 of appendix I of this part. the tangent to the keel line at that (3) These pressures are uniform and point. The magnitude of the resultant must be applied simultaneously over load is that specified in paragraph (b) the entire hull or main float bottom. of this section. The loads obtained must be carried (d) Unsymmetrical step loading. The re- into the sidewall structure of the hull sultant water load consists of a compo- proper, but need not be transmitted in nent equal to 0.75 times the load speci- a fore and aft direction as shear and fied in paragraph (a) of this section and bending loads. a side component equal to 0.025 tan b [Doc. No. 26269, 58 FR 42161, Aug. 6, 1993; 58 times the load specified in paragraph FR 51970, Oct. 5, 1993] (b) of this section. The side load must be applied perpendicularly to the plane § 23.535 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. must be designed for the conditions (e) Unsymmetrical bow loading. The re- prescribed in this section. In the cases sultant water load consists of a compo- specified in paragraphs (b) through (e) nent equal to 0.75 times the load speci- of this section, the prescribed water fied in paragraph (b) of this section and loads may be distributed over the float a side component equal to 0.25 tan b bottom to avoid excessive local loads, times the load specified in paragraph using bottom pressures not less than (c) of this section. The side load must those prescribed in paragraph (g) of be applied perpendicularly to the plane this section. of symmetry at a point midway be- (b) Step loading. The resultant water tween the keel and the chine. load must be applied in the plane of (f) Immersed float condition. The re- symmetry of the float at a point three- sultant load must be applied at the fourths of the distance from the bow to centroid of the cross section of the the step and must be perpendicular to float at a point one-third of the dis- the keel. The resultant limit load is tance from the bow to the step. The computed as follows, except that the limit load components are as follows: value of L need not exceed three times the weight of the displaced water when vertical= PgV the float is completely submerged: 2 3 2 2 CXSO PV (KV ) CV2 W3 aft = L = 5 SO 2 2 2 3 β 2 3 Tan() 1+ ry 2 S 3 2 CYSO PV (KV ) where— side = L=limit load (lbs.); 2 where— C5=0.0053; 3 VS0=seaplane stalling speed (knots) with P=mass density of water (slugs/ft. ) landing flaps extended in the appropriate V=volume of float (ft.3);

position and with no slipstream effect; CX=coefficient of drag force, equal to 0.133; W=seaplane design landing weight in pounds; Cy=coefficient of side force, equal to 0.106; 240

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K=0.8, except that lower values may be used (c) Each airplane with retractable if it is shown that the floats are incapable landing gear must be designed to pro- of submerging at a speed of 0.8 Vso in nor- tect each occupant in a landing— mal operations; (1) With the wheels retracted; Vso=seaplane stalling speed (knots) with landing flaps extended in the appropriate (2) With moderate descent velocity; position and with no slipstream effect; and and g=acceleration due to gravity (ft/sec2). (3) Assuming, in the absence of a (g) Float bottom pressures. The float more rational analysis— bottom pressures must be established (i) A downward ultimate inertia force under § 23.533, except that the value of of 3 g; and K2 in the formulae may be taken as 1.0. (ii) A coefficient of friction of 0.5 at The angle of dead rise to be used in de- the ground. termining the float bottom pressures is (d) If it is not established that a set forth in paragraph (b) of this sec- turnover is unlikely during an emer- tion. gency landing, the structure must be [Doc. No. 26269, 58 FR 42162, Aug. 6, 1993; 58 designed to protect the occupants in a FR 51970, Oct. 5, 1993] complete turnover as follows: (1) The likelihood of a turnover may § 23.537 Seawing loads. be shown by an analysis assuming the Seawing design loads must be based following conditions— on applicable test data. (i) The most adverse combination of weight and center of gravity position; [Doc. No. 26269, 58 FR 42163, Aug. 6, 1993] (ii) Longitudinal load factor of 9.0g; EMERGENCY LANDING CONDITIONS (iii) Vertical load factor of 1.0g; and (iv) For airplanes with tricycle land- § 23.561 General. ing gear, the nose wheel strut failed (a) The airplane, although it may be with the nose contacting the ground. damaged in emergency landing condi- (2) For determining the loads to be tions, must be designed as prescribed in applied to the inverted airplane after a this section to protect each occupant turnover, an upward ultimate inertia under those conditions. load factor of 3.0g and a coefficient of (b) The structure must be designed to friction with the ground of 0.5 must be give each occupant every reasonable used. chance of escaping serious injury (e) Except as provided in § 23.787(c), when— the supporting structure must be de- (1) Proper use is made of the seats, signed to restrain, under loads up to safety belts, and shoulder harnesses those specified in paragraph (b)(3) of provided for in the design; this section, each item of mass that (2) The occupant experiences the could injure an occupant if it came static inertia loads corresponding to loose in a minor crash landing. the following ultimate load factors— [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (i) Upward, 3.0g for normal, utility, amended by Amdt. 23–7, 34 FR 13090, Aug. 13, and commuter category airplanes, or 1969; Amdt. 23–24, 52 FR 34745, Sept. 14, 1987; 4.5g for acrobatic category airplanes; Amdt. 23–36, 53 FR 30812, Aug. 15, 1988; Amdt. (ii) Forward, 9.0g; 23–46, 59 FR 25772, May 17, 1994; Amdt. 23–48, (iii) Sideward, 1.5g; and 61 FR 5147, Feb. 9, 1996] (iv) Downward, 6.0g when certifi- cation to the emergency exit provi- § 23.562 Emergency landing dynamic sions of § 23.807(d)(4) is requested; and conditions. (3) The items of mass within the (a) Each seat/restraint system for use cabin, that could injure an occupant, in a normal, utility, or acrobatic cat- experience the static inertia loads cor- egory airplane must be designed to pro- responding to the following ultimate tect each occupant during an emer- load factors— gency landing when— (i) Upward, 3.0g; (1) Proper use is made of seats, safety (ii) Forward, 18.0g; and belts, and shoulder harnesses provided (iii) Sideward, 4.5g. for in the design; and

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(2) The occupant is exposed to the the airframe structure must be pre- loads resulting from the conditions loaded to misalign with respect to each prescribed in this section. other by at least 10 degrees vertically (b) Except for those seat/restraint (i.e., pitch out of parallel) and one of systems that are required to meet the rails or attachment devices must paragraph (d) of this section, each seat/ be preloaded to misalign by 10 degrees restraint system for crew or passenger in roll prior to conducting the test de- occupancy in a normal, utility, or acro- fined by paragraph (b)(2) of this sec- batic category airplane, must success- tion. fully complete dynamic tests or be (c) Compliance with the following re- demonstrated by rational analysis sup- quirements must be shown during the ported by dynamic tests, in accordance dynamic tests conducted in accordance with each of the following conditions. with paragraph (b) of this section: These tests must be conducted with an (1) The seat/restraint system must occupant simulated by an restrain the ATD although seat/re- anthropomorphic test dummy (ATD) straint system components may experi- defined by 49 CFR Part 572, Subpart B, ence deformation, elongation, displace- or an FAA-approved equivalent, with a ment, or crushing intended as part of nominal weight of 170 pounds and seat- the design. ed in the normal upright position. (2) The attachment between the seat/ (1) For the first test, the change in restraint system and the test fixture velocity may not be less than 31 feet must remain intact, although the seat per second. The seat/restraint system structure may have deformed. must be oriented in its nominal posi- (3) Each shoulder harness strap must tion with respect to the airplane and remain on the ATD’s shoulder during with the horizontal plane of the air- the impact. plane pitched up 60 degrees, with no (4) The safety belt must remain on yaw, relative to the impact vector. For the ATD’s pelvis during the impact. seat/restraint systems to be installed (5) The results of the dynamic tests in the first row of the airplane, peak must show that the occupant is pro- deceleration must occur in not more tected from serious head injury. than 0.05 seconds after impact and (i) When contact with adjacent seats, must reach a minimum of 19g. For all structure, or other items in the cabin other seat/restraint systems, peak de- can occur, protection must be provided celeration must occur in not more than so that the head impact does not ex- 0.06 seconds after impact and must ceed a head injury criteria (HIC) of reach a minimum of 15g. 1,000. (2) For the second test, the change in (ii) The value of HIC is defined as— velocity may not be less than 42 feet per second. The seat/restraint system ⎧ 2.5 ⎫ ⎡ t ⎤ must be oriented in its nominal posi- ⎪ ⎢ 1 2 ⎥ ⎪ tion with respect to the airplane and =−⎨() ⎬ HIC t21 t ⎢ ∫ a(t)dt⎥ with the vertical plane of the airplane ⎪ ()tt− ⎪ ⎣ 21t1 ⎦ yawed 10 degrees, with no pitch, rel- ⎩ ⎭Max ative to the impact vector in a direc- Where: tion that results in the greatest load t is the initial integration time, expressed on the shoulder harness. For seat/re- 1 in seconds, t2 is the final integration time, straint systems to be installed in the expressed in seconds, (t2¥ t1) is the time first row of the airplane, peak decelera- duration of the major head impact, ex- tion must occur in not more than 0.05 pressed in seconds, and a(t) is the resultant seconds after impact and must reach a deceleration at the center of gravity of the minimum of 26g. For all other seat/re- head form expressed as a multiple of g straint systems, peak deceleration (units of gravity). must occur in not more than 0.06 sec- (iii) Compliance with the HIC limit onds after impact and must reach a must be demonstrated by measuring minimum of 21g. the head impact during dynamic test- (3) To account for floor warpage, the ing as prescribed in paragraphs (b)(1) floor rails or attachment devices used and (b)(2) of this section or by a sepa- to attach the seat/restraint system to rate showing of compliance with the

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head injury criteria using test or anal- quired by this section may be used if ysis procedures. substantiated on a rational basis. (6) Loads in individual shoulder har- ness straps may not exceed 1,750 [Amdt. 23–36, 53 FR 30812, Aug. 15, 1988, as amended by Amdt. 23–44, 58 FR 38639, July 19, pounds. If dual straps are used for re- 1993; Amdt. 23–50, 61 FR 5192, Feb. 9, 1996] taining the upper torso, the total strap loads may not exceed 2,000 pounds. FATIGUE EVALUATION (7) The compression load measured between the pelvis and the lumbar § 23.571 Metallic pressurized cabin spine of the ATD may not exceed 1,500 structures. pounds. For normal, utility, and acrobatic (d) For all single-engine airplanes category airplanes, the strength, detail with a V of more than 61 knots at SO design, and fabrication of the metallic maximum weight, and those multien- structure of the pressure cabin must be gine airplanes of 6,000 pounds or less evaluated under one of the following: maximum weight with a VSO of more than 61 knots at maximum weight that (a) A fatigue strength investigation do not comply with § 23.67(a)(1); in which the structure is shown by (1) The ultimate load factors of tests, or by analysis supported by test § 23.561(b) must be increased by multi- evidence, to be able to withstand the plying the load factors by the square of repeated loads of variable magnitude the ratio of the increased stall speed to expected in service; or 61 knots. The increased ultimate load (b) A fail safe strength investigation, factors need not exceed the values in which it is shown by analysis, tests, or both that catastrophic failure of the reached at a VS0 of 79 knots. The up- ward ultimate load factor for acrobatic structure is not probable after fatigue category airplanes need not exceed failure, or obvious partial failure, of a 5.0g. principal structural element, and that (2) The seat/restraint system test re- the remaining structures are able to quired by paragraph (b)(1) of this sec- withstand a static ultimate load factor tion must be conducted in accordance of 75 percent of the limit load factor at with the following criteria: VC, considering the combined effects of (i) The change in velocity may not be normal operating pressures, expected less than 31 feet per second. external aerodynamic pressures, and (ii)(A) The peak deceleration (gp) of flight loads. These loads must be mul- 19g and 15g must be increased and mul- tiplied by a factor of 1.15 unless the dy- tiplied by the square of the ratio of the namic effects of failure under static increased stall speed to 61 knots: load are otherwise considered. (c) The damage tolerance evaluation g =19.0 (V /61)2 or g =15.0 (V /61)2 p S0 p S0 of § 23.573(b). (B) The peak deceleration need not exceed the value reached at a V of 79 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as S0 amended by Amdt. 23–14, 38 FR 31821, Nov. 19, knots. 1973; Amdt. 23–45, 58 FR 42163, Aug. 6, 1993; (iii) The peak deceleration must Amdt. 23–48, 61 FR 5147, Feb. 9, 1996] occur in not more than time (tr), which must be computed as follows: § 23.572 Metallic wing, empennage, and associated structures. 31 .96 = = (a) For normal, utility, and acrobatic t r 32.2() g g category airplanes, the strength, detail pp design, and fabrication of those parts where— of the airframe structure whose failure

gp=The peak deceleration calculated in ac- would be catastrophic must be evalu- cordance with paragraph (d)(2)(ii) of this ated under one of the following unless section it is shown that the structure, oper- tr=The rise time (in seconds) to the peak de- ating stress level, materials and ex- celeration. pected uses are comparable, from a fa- (e) An alternate approach that tigue standpoint, to a similar design achieves an equivalent, or greater, that has had extensive satisfactory level of occupant protection to that re- service experience:

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(1) A fatigue strength investigation structure, the structure must be evalu- in which the structure is shown by ated in accordance with paragraphs tests, or by analysis supported by test (a)(1) and (a)(6) of this section. Where evidence, to be able to withstand the bonded joints are used, the structure repeated loads of variable magnitude must also be evaluated in accordance expected in service; or with paragraph (a)(5) of this section. (2) A fail-safe strength investigation The effects of material variability and in which it is shown by analysis, tests, environmental conditions on the or both, that catastrophic failure of strength and durability properties of the structure is not probable after fa- the composite materials must be ac- tigue failure, or obvious partial failure, counted for in the evaluations required of a principal structural element, and by this section. that the remaining structure is able to (1) It must be demonstrated by tests, withstand a static ultimate load factor or by analysis supported by tests, that of 75 percent of the critical limit load the structure is capable of carrying ul- factor at V These loads must be mul- c. timate load with damage up to the tiplied by a factor of 1.15 unless the dy- namic effects of failure under static threshold of detectability considering load are otherwise considered. the inspection procedures employed. (3) The damage tolerance evaluation (2) The growth rate or no-growth of of § 23.573(b). damage that may occur from fatigue, (b) Each evaluation required by this corrosion, manufacturing flaws or im- section must— pact damage, under repeated loads ex- (1) Include typical loading spectra pected in service, must be established (e.g. taxi, ground-air-ground cycles, by tests or analysis supported by tests. maneuver, gust); (3) The structure must be shown by (2) Account for any significant effects residual strength tests, or analysis sup- due to the mutual influence of aero- ported by residual strength tests, to be dynamic surfaces; and able to withstand critical limit flight (3) Consider any significant effects loads, considered as ultimate loads, from propeller slipstream loading, and with the extent of detectable damage buffet from vortex impingements. consistent with the results of the dam- [Amdt. 23–7, 34 FR 13090, Aug. 13, 1969, as age tolerance evaluations. For pressur- amended by Amdt. 23–14, 38 FR 31821, Nov. 19, ized cabins, the following loads must be 1973; Amdt. 23–34, 52 FR 1830, Jan. 15, 1987; withstood: Amdt. 23–38, 54 FR 39511, Sept. 26, 1989; Amdt. (i) Critical limit flight loads with the 23–45, 58 FR 42163, Aug. 6, 1993; Amdt. 23–48, 61 combined effects of normal operating FR 5147, Feb. 9, 1996] pressure and expected external aero- § 23.573 Damage tolerance and fatigue dynamic pressures. evaluation of structure. (ii) The expected external aero- dynamic pressures in 1g flight com- (a) Composite airframe structure. Com- bined with a cabin differential pressure posite airframe structure must be eval- equal to 1.1 times the normal operating uated under this paragraph instead of §§ 23.571 and 23.572. The applicant must differential pressure without any other evaluate the composite airframe struc- load. ture, the failure of which would result (4) The damage growth, between ini- in catastrophic loss of the airplane, in tial detectability and the value se- each wing (including canards, tandem lected for residual strength demonstra- wings, and winglets), empennage, their tions, factored to obtain inspection in- carrythrough and attaching structure, tervals, must allow development of an moveable control surfaces and their at- inspection program suitable for appli- taching structure fuselage, and pres- cation by operation and maintenance sure cabin using the damage-tolerance personnel. criteria prescribed in paragraphs (a)(1) (5) For any bonded joint, the failure through (a)(4) of this section unless of which would result in catastrophic shown to be impractical. If the appli- loss of the airplane, the limit load ca- cant establishes that damage-tolerance pacity must be substantiated by one of criteria is impractical for a particular the following methods—

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(i) The maximum disbonds of each (1) The normal operating differential bonded joint consistent with the capa- pressure combined with the expected bility to withstand the loads in para- external aerodynamic pressures applied graph (a)(3) of this section must be de- simultaneously with the flight loading termined by analysis, tests, or both. conditions specified in this part, and Disbonds of each bonded joint greater (2) The expected external aero- than this must be prevented by design dynamic pressures in 1g flight com- features; or bined with a cabin differential pressure (ii) Proof testing must be conducted equal to 1.1 times the normal operating on each production article that will differential pressure without any other apply the critical limit design load to load. each critical bonded joint; or [Doc. No. 26269, 58 FR 42163, Aug. 6, 1993; 58 (iii) Repeatable and reliable non-de- FR 51970, Oct. 5, 1993, as amended by Amdt. structive inspection techniques must 23–48, 61 FR 5147, Feb. 9, 1996; 73 FR 19746, be established that ensure the strength Apr. 11, 2008] of each joint. § 23.574 Metallic damage tolerance and (6) Structural components for which fatigue evaluation of commuter cat- the damage tolerance method is shown egory airplanes. to be impractical must be shown by For commuter category airplanes— component fatigue tests, or analysis (a) Metallic damage tolerance. An eval- supported by tests, to be able to with- uation of the strength, detail design, stand the repeated loads of variable and fabrication must show that cata- magnitude expected in service. Suffi- strophic failure due to fatigue, corro- cient component, subcomponent, ele- sion, defects, or damage will be avoided ment, or coupon tests must be done to throughout the operational life of the establish the fatigue scatter factor and airplane. This evaluation must be con- the environmental effects. Damage up ducted in accordance with the provi- to the threshold of detectability and sions of § 23.573, except as specified in ultimate load residual strength capa- paragraph (b) of this section, for each bility must be considered in the dem- part of the structure that could con- onstration. tribute to a catastrophic failure. (b) Metallic airframe structure. If the (b) Fatigue (safe-life) evaluation. Com- applicant elects to use § 23.571(c) or pliance with the damage tolerance re- § 23.572(a)(3), then the damage tolerance quirements of paragraph (a) of this sec- evaluation must include a determina- tion is not required if the applicant es- tion of the probable locations and tablishes that the application of those modes of damage due to fatigue, corro- requirements is impractical for a par- sion, or accidental damage. Damage at ticular structure. This structure must multiple sites due to fatigue must be be shown, by analysis supported by test included where the design is such that evidence, to be able to withstand the this type of damage can be expected to repeated loads of variable magnitude occur. The evaluation must incor- expected during its service life without porate repeated load and static anal- detectable cracks. Appropriate safe-life yses supported by test evidence. The scatter factors must be applied. extent of damage for residual strength evaluation at any time within the [Doc. No. 27805, 61 FR 5148, Feb. 9, 1996] operational life of the airplane must be § 23.575 Inspections and other proce- consistent with the initial detect- dures. ability and subsequent growth under repeated loads. The residual strength Each inspection or other procedure, evaluation must show that the remain- based on an evaluation required by ing structure is able to withstand crit- §§ 23.571, 23.572, 23.573 or 23.574, must be ical limit flight loads, considered as ul- established to prevent catastrophic timate, with the extent of detectable failure and must be included in the damage consistent with the results of Limitations Section of the Instructions the damage tolerance evaluations. For for Continued Airworthiness required by § 23.1529. pressurized cabins, the following load must be withstood: [Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]

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Subpart D—Design and vice is used in addition to the self-lock- Construction ing device. [Doc. No. 27805, 61 FR 5148, Feb. 9, 1996] § 23.601 General. The suitability of each questionable § 23.609 Protection of structure. design detail and part having an impor- Each part of the structure must— tant bearing on safety in operations, (a) Be suitably protected against de- must be established by tests. terioration or loss of strength in serv- § 23.603 Materials and workmanship. ice due to any cause, including— (1) Weathering; (a) The suitability and durability of (2) Corrosion; and materials used for parts, the failure of (3) Abrasion; and which could adversely affect safety, (b) Have adequate provisions for ven- must— tilation and drainage. (1) Be established by experience or tests; § 23.611 Accessibility provisions. (2) Meet approved specifications that ensure their having the strength and For each part that requires mainte- other properties assumed in the design nance, inspection, or other servicing, data; and appropriate means must be incor- (3) Take into account the effects of porated into the aircraft design to environmental conditions, such as tem- allow such servicing to be accom- perature and humidity, expected in plished. service. [Doc. No. 27805, 61 FR 5148, Feb. 9, 1996] (b) Workmanship must be of a high standard. § 23.613 Material strength properties [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as and design values. amended by Amdt. 23–17, 41 FR 55464, Dec. 20, (a) Material strength properties must 1976; Amdt. 23–23, 43 FR 50592, Oct. 10, 1978] be based on enough tests of material meeting specifications to establish de- § 23.605 Fabrication methods. sign values on a statistical basis. (a) The methods of fabrication used (b) Design values must be chosen to must produce consistently sound struc- minimize the probability of structural tures. If a fabrication process (such as failure due to material variability. Ex- gluing, spot welding, or heat-treating) cept as provided in paragraph (e) of requires close control to reach this ob- this section, compliance with this jective, the process must be performed paragraph must be shown by selecting under an approved process specifica- design values that ensure material tion. strength with the following prob- (b) Each new aircraft fabrication ability: method must be substantiated by a (1) Where applied loads are eventu- test program. ally distributed through a single mem- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 ber within an assembly, the failure of FR 258, Jan. 9, 1965, as amended by Amdt. 23– which would result in loss of structural 23, 43 FR 50592, Oct. 10, 1978] integrity of the component; 99 percent probability with 95 percent confidence. § 23.607 Fasteners. (2) For redundant structure, in which (a) Each removable fastener must in- the failure of individual elements corporate two retaining devices if the would result in applied loads being loss of such fastener would preclude safely distributed to other load car- continued safe flight and landing. rying members; 90 percent probability (b) Fasteners and their locking de- with 95 percent confidence. vices must not be adversely affected by (c) The effects of temperature on al- the environmental conditions associ- lowable stresses used for design in an ated with the particular installation. essential component or structure must (c) No self-locking nut may be used be considered where thermal effects are on any bolt subject to rotation in oper- significant under normal operating ation unless a non-friction locking de- conditions.

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(d) The design of the structure must (c) Critical castings. For each casting minimize the probability of cata- whose failure would preclude continued strophic fatigue failure, particularly at safe flight and landing of the airplane points of stress concentration. or result in serious injury to occu- (e) Design values greater than the pants, the following apply: guaranteed minimums required by this (1) Each critical casting must ei- section may be used where only guar- ther— anteed minimum values are normally (i) Have a casting factor of not less allowed if a ‘‘premium selection’’ of the than 1.25 and receive 100 percent in- material is made in which a specimen spection by visual, radiographic, and of each individual item is tested before either magnetic particle, penetrant or use to determine that the actual other approved equivalent non-destruc- strength properties of that particular tive inspection method; or item will equal or exceed those used in (ii) Have a casting factor of not less design. than 2.0 and receive 100 percent visual inspection and 100 percent approved [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 non-destructive inspection. When an FR 258, Jan. 9, 1965, as amended by Amdt. 23– approved quality control procedure is 23, 43 FR 50592, Oct. 30, 1978; Amdt. 23–45, 58 FR 42163, Aug. 6, 1993] established and an acceptable statis- tical analysis supports reduction, non- § 23.619 Special factors. destructive inspection may be reduced from 100 percent, and applied on a sam- The factor of safety prescribed in pling basis. § 23.303 must be multiplied by the high- (2) For each critical casting with a est pertinent special factors of safety casting factor less than 1.50, three sam- prescribed in §§ 23.621 through 23.625 for ple castings must be static tested and each part of the structure whose shown to meet— strength is— (i) The strength requirements of (a) Uncertain; § 23.305 at an ultimate load cor- (b) Likely to deteriorate in service responding to a casting factor of 1.25; before normal replacement; or and (c) Subject to appreciable variability (ii) The deformation requirements of because of uncertainties in manufac- § 23.305 at a load of 1.15 times the limit turing processes or inspection methods. load. [Amdt. 23–7, 34 FR 13091, Aug. 13, 1969] (3) Examples of these castings are structural attachment fittings, parts of § 23.621 Casting factors. flight control systems, control surface (a) General. The factors, tests, and in- hinges and balance weight attach- spections specified in paragraphs (b) ments, seat, berth, safety belt, and fuel through (d) of this section must be ap- and oil tank supports and attachments, plied in addition to those necessary to and cabin pressure valves. establish foundry quality control. The (d) Non-critical castings. For each inspections must meet approved speci- casting other than those specified in fications. Paragraphs (c) and (d) of this paragraph (c) or (e) of this section, the section apply to any structural cast- following apply: ings except castings that are pressure (1) Except as provided in paragraphs tested as parts of hydraulic or other (d)(2) and (3) of this section, the casting fluid systems and do not support struc- factors and corresponding inspections tural loads. must meet the following table: (b) Bearing stresses and surfaces. The Casting factor Inspection casting factors specified in paragraphs 2.0 or more ...... 100 percent visual. (c) and (d) of this section— Less than 2.0 but more 100 percent visual, and magnetic (1) Need not exceed 1.25 with respect than 1.5. particle or penetrant or equiva- to bearing stresses regardless of the lent nondestructive inspection method of inspection used; and methods. 1.25 through 1.50 ...... 100 percent visual, magnetic par- (2) Need not be used with respect to ticle or penetrant, and radio- the bearing surfaces of a part whose graphic, or approved equivalent bearing factor is larger than the appli- nondestructive inspection meth- cable casting factor. ods.

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(2) The percentage of castings in- test data (such as continuous joints in spected by nonvisual methods may be metal plating, welded joints, and scarf reduced below that specified in sub- joints in wood). paragraph (d)(1) of this section when an (c) For each integral fitting, the part approved quality control procedure is must be treated as a fitting up to the established. point at which the section properties (3) For castings procured to a speci- become typical of the member. fication that guarantees the mechan- (d) For each seat, berth, safety belt, ical properties of the material in the and harness, its attachment to the casting and provides for demonstration structure must be shown, by analysis, of these properties by test of coupons tests, or both, to be able to withstand cut from the castings on a sampling the inertia forces prescribed in § 23.561 basis— multiplied by a fitting factor of 1.33. (i) A casting factor of 1.0 may be used; and [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (ii) The castings must be inspected as amended by Amdt. 23–7, 34 FR 13091, Aug. 13, provided in paragraph (d)(1) of this sec- 1969] tion for casting factors of ‘‘1.25 through § 23.627 Fatigue strength. 1.50’’ and tested under paragraph (c)(2) of this section. The structure must be designed, as (e) Non-structural castings. Castings far as practicable, to avoid points of used for non-structural purposes do not stress concentration where variable require evaluation, testing or close in- stresses above the fatigue limit are spection. likely to occur in normal service.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as § 23.629 Flutter. amended by Amdt. 23–45, 58 FR 42164, Aug. 6, 1993] (a) It must be shown by the methods of paragraph (b) and either paragraph § 23.623 Bearing factors. (c) or (d) of this section, that the air- (a) Each part that has clearance (free plane is free from flutter, control re- fit), and that is subject to pounding or versal, and divergence for any condi- vibration, must have a bearing factor tion of operation within the limit V-n large enough to provide for the effects envelope and at all speeds up to the of normal relative motion. speed specified for the selected method. (b) For control surface hinges and In addition— control system joints, compliance with (1) Adequate tolerances must be es- the factors prescribed in §§ 23.657 and tablished for quantities which affect 23.693, respectively, meets paragraph flutter, including speed, damping, mass (a) of this section. balance, and control system stiffness; and [Amdt. 23–7, 34 FR 13091, Aug. 13, 1969] (2) The natural frequencies of main § 23.625 Fitting factors. structural components must be deter- For each fitting (a part or terminal mined by vibration tests or other ap- used to join one structural member to proved methods. another), the following apply: (b) Flight flutter tests must be made (a) For each fitting whose strength is to show that the airplane is free from not proven by limit and ultimate load flutter, control reversal and divergence tests in which actual stress conditions and to show that— are simulated in the fitting and sur- (1) Proper and adequate attempts to rounding structures, a fitting factor of induce flutter have been made within at least 1.15 must be applied to each the speed range up to VD; part of— (2) The vibratory response of the (1) The fitting; structure during the test indicates (2) The means of attachment; and freedom from flutter; (3) The bearing on the joined mem- (3) A proper margin of damping exists bers. at VD; and (b) No fitting factor need be used for (4) There is no large and rapid reduc- joint designs based on comprehensive tion in damping as VD is approached. 248

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(c) Any rational analysis used to pre- (g) For airplanes showing compliance dict freedom from flutter, control re- with the fail-safe criteria of §§ 23.571 versal and divergence must cover all and 23.572, the airplane must be shown speeds up to 1.2 VD. by analysis to be free from flutter up (d) Compliance with the rigidity and to VD/MD after fatigue failure, or obvi- mass balance criteria (pages 4–12), in ous partial failure, of a principal struc- Airframe and Equipment Engineering tural element. Report No. 45 (as corrected) ‘‘Simplified (h) For airplanes showing compliance Flutter Prevention Criteria’’ (published with the damage tolerance criteria of by the Federal Aviation Administra- § 23.573, the airplane must be shown by tion) may be accomplished to show analysis to be free from flutter up to that the airplane is free from flutter, VD/MD with the extent of damage for control reversal, or divergence if— which residual strength is dem- (1) VD/MD for the airplane is less than onstrated. 260 knots (EAS) and less than Mach 0.5, (i) For modifications to the type de- (2) The wing and aileron flutter pre- sign that could affect the flutter char- vention criteria, as represented by the acteristics, compliance with paragraph wing torsional stiffness and aileron (a) of this section must be shown, ex- balance criteria, are limited in use to cept that analysis based on previously airplanes without large mass con- approved data may be used alone to centrations (such as engines, floats, or show freedom from flutter, control re- fuel tanks in outer wing panels) along versal and divergence, for all speeds up the wing span, and to the speed specified for the selected (3) The airplane— method. (i) Does not have a T-tail or other un- conventional tail configurations; [Amdt. 23–23, 43 FR 50592, Oct. 30, 1978, as (ii) Does not have unusual mass dis- amended by Amdt. 23–31, 49 FR 46867, Nov. 28, tributions or other unconventional de- 1984; Amdt. 23–45, 58 FR 42164, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993; Amdt. 23–48, 61 FR 5148, sign features that affect the applica- Feb. 9, 1996] bility of the criteria, and (iii) Has fixed-fin and fixed-stabilizer WINGS surfaces. (e) For turbopropeller-powered air- § 23.641 Proof of strength. planes, the dynamic evaluation must The strength of stressed-skin wings include— must be proven by load tests or by (1) Whirl mode degree of freedom combined structural analysis and load which takes into account the stability tests. of the plane of rotation of the propeller and significant elastic, inertial, and CONTROL SURFACES aerodynamic forces, and (2) Propeller, engine, engine mount, § 23.651 Proof of strength. and airplane structure stiffness and damping variations appropriate to the (a) Limit load tests of control sur- particular configuration. faces are required. These tests must in- (f) Freedom from flutter, control re- clude the horn or fitting to which the control system is attached. versal, and divergence up to VD/MD must be shown as follows: (b) In structural analyses, rigging (1) For airplanes that meet the cri- loads due to wire bracing must be ac- teria of paragraphs (d)(1) through (d)(3) counted for in a rational or conserv- of this section, after the failure, mal- ative manner. function, or disconnection of any single element in any tab control system. § 23.655 Installation. (2) For airplanes other than those de- (a) Movable surfaces must be in- scribed in paragraph (f)(1) of this sec- stalled so that there is no interference tion, after the failure, malfunction, or between any surfaces, their bracing, or disconnection of any single element in adjacent fixed structure, when one sur- the primary flight control system, any face is held in its most critical clear- tab control system, or any flutter ance positions and the others are oper- damper. ated through their full movement.

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(b) If an adjustable stabilizer is used, provided for any failure in the stability it must have stops that will limit its augmentation system or in any other range of travel to that allowing safe automatic or power-operated system flight and landing. that could result in an unsafe condi- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as tion if the pilot was not aware of the amended by Amdt. 23–45, 58 FR 42164, Aug. 6, failure. Warning systems must not ac- 1993] tivate the control system. (b) The design of the stability aug- § 23.657 Hinges. mentation system or of any other auto- (a) Control surface hinges, except matic or power-operated system must ball and roller bearing hinges, must permit initial counteraction of failures have a factor of safety of not less than without requiring exceptional pilot 6.67 with respect to the ultimate bear- skill or strength, by either the deacti- ing strength of the softest material vation of the system or a failed portion used as a bearing. thereof, or by overriding the failure by (b) For ball or roller bearing hinges, movement of the flight controls in the the approved rating of the bearing may normal sense. not be exceeded. (c) It must be shown that, after any single failure of the stability aug- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as mentation system or any other auto- amended by Amdt. 23–48, 61 FR 5148, Feb. 9, 1996] matic or power-operated system— (1) The airplane is safely controllable § 23.659 Mass balance. when the failure or malfunction occurs The supporting structure and the at- at any speed or altitude within the ap- tachment of concentrated mass bal- proved operating limitations that is ance weights used on control surfaces critical for the type of failure being must be designed for— considered; (a) 24 g normal to the plane of the (2) The controllability and maneuver- control surface; ability requirements of this part are (b) 12 g fore and aft; and met within a practical operational (c) 12 g parallel to the hinge line. flight envelope (for example, speed, al- titude, normal acceleration, and air- CONTROL SYSTEMS plane configuration) that is described in the Airplane Flight Manual (AFM); § 23.671 General. and (a) Each control must operate easily, (3) The trim, stability, and stall char- smoothly, and positively enough to acteristics are not impaired below a allow proper performance of its func- level needed to permit continued safe tions. flight and landing. (b) Controls must be arranged and [Doc. No. 26269, 58 FR 42164, Aug. 6, 1993] identified to provide for convenience in operation and to prevent the possi- § 23.673 Primary flight controls. bility of confusion and subsequent in- Primary flight controls are those advertent operation. used by the pilot for the immediate § 23.672 Stability augmentation and control of pitch, roll, and yaw. automatic and power-operated sys- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as tems. amended by Amdt. 23–48, 61 FR 5148, Feb. 9, If the functioning of stability aug- 1996] mentation or other automatic or power-operated systems is necessary to § 23.675 Stops. show compliance with the flight char- (a) Each control system must have acteristics requirements of this part, stops that positively limit the range of such systems must comply with § 23.671 motion of each movable aerodynamic and the following: surface controlled by the system. (a) A warning, which is clearly dis- (b) Each stop must be located so that tinguishable to the pilot under ex- wear, slackness, or takeup adjustments pected flight conditions without re- will not adversely affect the control quiring the pilot’s attention, must be characteristics of the airplane because

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of a change in the range of surface plane weights and center of gravity po- travel. sitions. (c) Each stop must be able to with- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as stand any loads corresponding to the amended by Amdt. 23–7, 34 FR 13091, Aug. 13, design conditions for the control sys- 1969; Amdt. 23–34, 52 FR 1830, Jan. 15, 1987; tem. Amdt. 23–42, 56 FR 353, Jan. 3, 1991; Amdt. 23– [Amdt. 23–17, 41 FR 55464, Dec. 20, 1976] 49, 61 FR 5165, Feb. 9, 1996]

§ 23.677 Trim systems. § 23.679 Control system locks. (a) Proper precautions must be taken If there is a device to lock the con- to prevent inadvertent, improper, or trol system on the ground or water: abrupt trim tab operation. There must (a) There must be a means to— be means near the trim control to indi- (1) Give unmistakable warning to the cate to the pilot the direction of trim pilot when lock is engaged; or control movement relative to airplane (2) Automatically disengage the de- motion. In addition, there must be vice when the pilot operates the pri- means to indicate to the pilot the posi- mary flight controls in a normal man- tion of the trim device with respect to ner. both the range of adjustment and, in (b) The device must be installed to the case of lateral and directional limit the operation of the airplane so trim, the neutral position. This means that, when the device is engaged, the must be visible to the pilot and must pilot receives unmistakable warning at be located and designed to prevent con- the start of the takeoff. fusion. The pitch trim indicator must (c) The device must have a means to be clearly marked with a position or preclude the possibility of it becoming range within which it has been dem- inadvertently engaged in flight. onstrated that take-off is safe for all [Doc. No. 26269, 58 FR 42164, Aug. 6, 1993] center of gravity positions and each flap position approved for takeoff. § 23.681 Limit load static tests. (b) Trimming devices must be de- signed so that, when any one con- (a) Compliance with the limit load necting or transmitting element in the requirements of this part must be primary flight control system fails, shown by tests in which— adequate control for safe flight and (1) The direction of the test loads landing is available with— produces the most severe loading in the (1) For single-engine airplanes, the control system; and longitudinal trimming devices; or (2) Each fitting, pulley, and bracket (2) For multiengine airplanes, the used in attaching the system to the longitudinal and directional trimming main structure is included. devices. (b) Compliance must be shown (by (c) Tab controls must be irreversible analyses or individual load tests) with unless the tab is properly balanced and the special factor requirements for has no unsafe flutter characteristics. control system joints subject to angu- Irreversible tab systems must have lar motion. adequate rigidity and reliability in the portion of the system from the tab to § 23.683 Operation tests. the attachment of the irreversible unit (a) It must be shown by operation to the airplane structure. tests that, when the controls are oper- (d) It must be demonstrated that the ated from the pilot compartment with airplane is safely controllable and that the system loaded as prescribed in the pilot can perform all maneuvers paragraph (b) of this section, the sys- and operations necessary to effect a tem is free from— safe landing following any probable (1) Jamming; powered trim system runaway that (2) Excessive friction; and reasonably might be expected in serv- (3) Excessive deflection. ice, allowing for appropriate time (b) The prescribed test loads are— delay after pilot recognition of the (1) For the entire system, loads cor- trim system runaway. The demonstra- responding to the limit airloads on the tion must be conducted at critical air- appropriate surface, or the limit pilot

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forces in § 23.397(b), whichever are less; (b) Each kind and size of pulley must and correspond to the cable with which it is (2) For secondary controls, loads not used. Each pulley must have closely less than those corresponding to the fitted guards to prevent the cables maximum pilot effort established from being misplaced or fouled, even under § 23.405. when slack. Each pulley must lie in the plane passing through the cable so that [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13091, Aug. 13, the cable does not rub against the pul- 1969] ley flange. (c) Fairleads must be installed so § 23.685 Control system details. that they do not cause a change in (a) Each detail of each control sys- cable direction of more than three de- tem must be designed and installed to grees. prevent jamming, chafing, and inter- (d) Clevis pins subject to load or mo- ference from cargo, passengers, loose tion and retained only by cotter pins objects, or the freezing of moisture. may not be used in the control system. (b) There must be means in the cock- (e) Turnbuckles must be attached to pit to prevent the entry of foreign ob- parts having angular motion in a man- jects into places where they would jam ner that will positively prevent binding the system. throughout the range of travel. (c) There must be means to prevent (f) Tab control cables are not part of the slapping of cables or tubes against the primary control system and may be other parts. less than 1⁄8 inch diameter in airplanes (d) Each element of the flight control that are safely controllable with the system must have design features, or tabs in the most adverse positions. must be distinctively and permanently [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as marked, to minimize the possibility of amended by Amdt. 23–7, 34 FR 13091, Aug. 13, incorrect assembly that could result in 1969] malfunctioning of the control system. § 23.691 Artificial stall barrier system. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–17, 41 FR 55464, Dec. 20, If the function of an artificial stall 1976] barrier, for example, stick pusher, is used to show compliance with § 23.687 Spring devices. § 23.201(c), the system must comply The reliability of any spring device with the following: used in the control system must be es- (a) With the system adjusted for op- tablished by tests simulating service eration, the plus and minus airspeeds conditions unless failure of the spring at which downward pitching control will not cause flutter or unsafe flight will be provided must be established. characteristics. (b) Considering the plus and minus airspeed tolerances established by § 23.689 Cable systems. paragraph (a) of this section, an air- (a) Each cable, cable fitting, turn- speed must be selected for the activa- buckle, splice, and pulley used must tion of the downward pitching control meet approved specifications. In addi- that provides a safe margin above any tion— airspeed at which any unsatisfactory (1) No cable smaller than 1⁄8 inch di- stall characteristics occur. ameter may be used in primary control (c) In addition to the stall warning systems; required § 23.07, a warning that is clear- (2) Each cable system must be de- ly distinguishable to the pilot under all signed so that there will be no haz- expected flight conditions without re- ardous change in cable tension quiring the pilot’s attention, must be throughout the range of travel under provided for faults that would prevent operating conditions and temperature the system from providing the required variations; and pitching motion. (3) There must be means for visual (d) Each system must be designed so inspection at each fairlead, pulley, ter- that the artificial stall barrier can be minal, and turnbuckle. quickly and positively disengaged by

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the pilots to prevent unwanted down- flap will not move from that position ward pitching of the airplane by a unless the control is adjusted or is quick release (emergency) control that moved by the automatic operation of a meets the requirements of § 23.1329(b). flap load limiting device. (e) A preflight check of the complete (b) The rate of movement of the flaps system must be established and the in response to the operation of the pi- procedure for this check made avail- lot’s control or automatic device must able in the Airplane Flight Manual give satisfactory flight and perform- (AFM). Preflight checks that are crit- ance characteristics under steady or ical to the safety of the airplane must changing conditions of airspeed, engine be included in the limitations section power, and attitude. of the AFM. (c) If compliance with § 23.145(b)(3) (f) For those airplanes whose design necessitates wing flap retraction to po- includes an autopilot system: sitions that are not fully retracted, the (1) A quick release (emergency) con- wing flap control lever settings cor- trol installed in accordance with responding to those positions must be § 23.1329(b) may be used to meet the re- positively located such that a definite quirements of paragraph (d), of this change of direction of movement of the section, and lever is necessary to select settings be- (2) The pitch servo for that system yond those settings. may be used to provide the stall down- ward pitching motion. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–49, 61 FR 5165, Feb. 9, (g) In showing compliance with 1996] § 23.1309, the system must be evaluated to determine the effect that any an- § 23.699 Wing flap position indicator. nounced or unannounced failure may There must be a wing flap position have on the continued safe flight and indicator for— landing of the airplane or the ability of (a) Flap installations with only the the crew to cope with any adverse con- retracted and fully extended position, ditions that may result from such fail- unless— ures. This evaluation must consider (1) A direct operating mechanism the hazards that would result from the provides a sense of ‘‘feel’’ and position airplane’s flight characteristics if the (such as when a mechanical linkage is system was not provided, and the haz- employed); or ard that may result from unwanted (2) The flap position is readily deter- downward pitching motion, which mined without seriously detracting could result from a failure at airspeeds from other piloting duties under any above the selected stall speed. flight condition, day or night; and [Doc. No. 27806, 61 FR 5165, Feb. 9, 1996] (b) Flap installation with inter- mediate flap positions if— § 23.693 Joints. (1) Any flap position other than re- Control system joints (in push-pull tracted or fully extended is used to systems) that are subject to angular show compliance with the performance motion, except those in ball and roller requirements of this part; and bearing systems, must have a special (2) The flap installation does not factor of safety of not less than 3.33 meet the requirements of paragraph with respect to the ultimate bearing (a)(1) of this section. strength of the softest material used as a bearing. This factor may be reduced § 23.701 Flap interconnection. to 2.0 for joints in cable control sys- (a) The main wing flaps and related tems. For ball or roller bearings, the movable surfaces as a system must— approved ratings may not be exceeded. (1) Be synchronized by a mechanical interconnection between the movable § 23.697 Wing flap controls. flap surfaces that is independent of the (a) Each wing flap control must be flap drive system; or by an approved designed so that, when the flap has equivalent means; or been placed in any position upon which (2) Be designed so that the occur- compliance with the performance re- rence of any failure of the flap system quirements of this part is based, the that would result in an unsafe flight

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characteristic of the airplane is ex- LANDING GEAR tremely improbable; or (b) The airplane must be shown to § 23.721 General. have safe flight characteristics with For commuter category airplanes any combination of extreme positions that have a passenger seating configu- of individual movable surfaces (me- ration, excluding pilot seats, of 10 or chanically interconnected surfaces are more, the following general require- to be considered as a single surface). ments for the landing gear apply: (c) If an interconnection is used in (a) The main landing-gear system multiengine airplanes, it must be de- must be designed so that if it fails due signed to account for the to overloads during takeoff and landing unsummetrical loads resulting from (assuming the overloads to act in the flight with the engines on one side of upward and aft directions), the failure the plane of symmetry inoperative and mode is not likely to cause the spillage the remaining engines at takeoff of enough fuel from any part of the fuel power. For single-engine airplanes, and system to consitute a fire hazard. multiengine airplanes with no slip- (b) Each airplane must be designed so stream effects on the flaps, it may be that, with the airplane under control, assumed that 100 percent of the critical it can be landed on a paved runway air load acts on one side and 70 percent with any one or more landing-gear legs on the other. not extended without sustaining a structural component failure that is [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as likely to cause the spillage of enough amended by Amdt. 23–14, 38 FR 31821, Nov. 19, fuel to consitute a fire hazard. 1973; Amdt. 23–42, 56 FR 353, Jan. 3, 1991; 56 (c) Compliance with the provisions of FR 5455, Feb. 11, 1991; Amdt. 23–49, 61 FR 5165, this section may be shown by analysis Feb. 9, 1996] or tests, or both. § 23.703 Takeoff warning system. [Amdt. 23–34, 52 FR 1830, Jan. 15, 1987] For commuter category airplanes, unless it can be shown that a lift or § 23.723 Shock absorption tests. longitudinal trim device that affects (a) It must be shown that the limit the takeoff performance of the aircraft load factors selected for design in ac- would not give an unsafe takeoff con- cordance with § 23.473 for takeoff and figuration when selection out of an ap- landing weights, respectively, will not proved takeoff position, a takeoff be exceeded. This must be shown by en- warning system must be installed and ergy absorption tests except that anal- meet the following requirements: ysis based on tests conducted on a (a) The system must provide to the landing gear system with identical en- pilots an aural warning that is auto- ergy absorption characteristics may be matically activated during the initial used for increases in previously ap- portion of the takeoff role if the air- proved takeoff and landing weights. plane is in a configuration that would (b) The landing gear may not fail, but not allow a safe takeoff. The warning may yield, in a test showing its reserve must continue until— energy absorption capacity, simulating a descent velocity of 1.2 times the limit (1) The configuration is changed to descent velocity, assuming wing lift allow safe takeoff, or equal to the weight of the airplane. (2) Action is taken by the pilot to abandon the takeoff roll. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 (b) The means used to activate the FR 258, Jan. 9, 1965, as amended by Amdt. 23– system must function properly for all 23, 43 FR 50593, Oct. 30, 1978; Amdt. 23–49, 61 FR 5166, Feb. 9, 1996] authorized takeoff power settings and procedures and throughout the ranges § 23.725 Limit drop tests. of takeoff weights, altitudes, and tem- peratures for which certification is re- (a) If compliance with § 23.723(a) is shown by free drop tests, these tests quested. must be made on the complete air- [Doc. No. 27806, 61 FR 5166, Feb. 9, 1996] plane, or on units consisting of wheel,

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tire, and shock absorber, in their prop- nj=the load factor developed in the drop test er relation, from free drop heights not (that is, the acceleration (dv/dt) in gs re- less than those determined by the fol- corded in the drop test) plus 1.0; and lowing formula: We, W, and L are the same as in the drop test computation. h (inches)=3.6 (W/S) 1 2 ⁄ (f) The value of n determined in ac- However, the free drop height may not cordance with paragraph (e) may not be less than 9.2 inches and need not be be more than the limit inertia load fac- more than 18.7 inches. tor used in the landing conditions in § 23.473. (b) If the effect of wing lift is pro- vided for in free drop tests, the landing [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as gear must be dropped with an effective amended by Amdt. 23–7, 34 FR 13091, Aug. 13, weight equal to 1969; Amdt. 23–48, 61 FR 5148, Feb. 9, 1996] § 23.726 Ground load dynamic tests. []hLd+−()1 = (a) If compliance with the ground WWe ()hd+ load requirements of §§ 23.479 through 23.483 is shown dynamically by drop where— test, one drop test must be conducted We=the effective weight to be used in the that meets § 23.725 except that the drop drop test (lbs.); height must be— h=specified free drop height (inches); (1) 2.25 times the drop height pre- d=deflection under impact of the tire (at the approved inflation pressure) plus the scribed in § 23.725(a); or vertical component of the axle travel rel- (2) Sufficient to develop 1.5 times the ative to the drop mass (inches); limit load factor. W=WM for main gear units (lbs), equal to the (b) The critical landing condition for static weight on that unit with the air- each of the design conditions specified plane in the level attitude (with the nose in §§ 23.479 through 23.483 must be used wheel clear in the case of nose wheel type for proof of strength. airplanes); W=WT for tail gear units (lbs.), equal to the [Amdt. 23–7, 34 FR 13091, Aug. 13, 1969] static weight on the tail unit with the air- plane in the tail-down attitude; § 23.727 Reserve energy absorption W=WN for nose wheel units lbs.), equal to the drop test. vertical component of the static reaction (a) If compliance with the reserve en- that would exist at the nose wheel, assum- ing that the mass of the airplane acts at ergy absorption requirement in the center of gravity and exerts a force of § 23.723(b) is shown by free drop tests, 1.0 g downward and 0.33 g forward; and the drop height may not be less than L= the ratio of the assumed wing lift to the 1.44 times that specified in § 23.725. airplane weight, but not more than 0.667. (b) If the effect of wing lift is pro- (c) The limit inertia load factor must vided for, the units must be dropped be determined in a rational or conserv- with an effective mass equal to We=Wh/ ative manner, during the drop test, (h+d), when the symbols and other de- using a landing gear unit attitude, and tails are the same as in § 23.725. applied drag loads, that represent the [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as landing conditions. amended by Amdt. 23–7, 34 FR 13091, Aug. 13, (d) The value of d used in the com- 1969] putation of We in paragraph (b) of this section may not exceed the value actu- § 23.729 Landing gear extension and retraction system. ally obtained in the drop test. (e) The limit inertia load factor must (a) General. For airplanes with re- be determined from the drop test in tractable landing gear, the following paragraph (b) of this section according apply: to the following formula: (1) Each landing gear retracting mechanism and its supporting struc- W ture must be designed for maximum =+e nnj L flight load factors with the gear re- W tracted and must be designed for the where— combination of friction, inertia, brake

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torque, and air loads, occurring during ing device prescribed in this paragraph, retraction at any airspeed up to 1.6 VS1 the warning system must be designed with flaps retracted, and for any load so that when the warning has been sus- factor up to those specified in § 23.345 pended after one or more throttles are for the flaps-extended condition. closed, subsequent retardation of any (2) The landing gear and retracting throttle to, or beyond, the position for mechanism, including the wheel well normal landing approach will activate doors, must withstand flight loads, in- the warning device. cluding loads resulting from all yawing (2) A device that functions continu- conditions specified in § 23.351, with the ously when the wing flaps are extended landing gear extended at any speed up beyond the maximum approach flap po- to at least 1.6 VS1 with the flaps re- sition, using a normal landing proce- tracted. dure, if the landing gear is not fully ex- (b) Landing gear lock. There must be tended and locked. There may not be a positive means (other than the use of manual shutoff for this warning device. hydraulic pressure) to keep the landing The flap position sensing unit may be gear extended. installed at any suitable location. The (c) Emergency operation. For a land- system for this device may use any plane having retractable landing gear part of the system (including the aural that cannot be extended manually, warning device) for the device required there must be means to extend the in paragraph (f)(1) of this section. landing gear in the event of either— (1) Any reasonably probable failure in (g) Equipment located in the landing the normal landing gear operation sys- gear bay. If the landing gear bay is used tem; or as the location for equipment other (2) Any reasonably probable failure in than the landing gear, that equipment a power source that would prevent the must be designed and installed to mini- operation of the normal landing gear mize damage from items such as a tire operation system. burst, or rocks, water, and slush that (d) Operation test. The proper func- may enter the landing gear bay. tioning of the retracting mechanism [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as must be shown by operation tests. amended by Amdt. 23–7, 34 FR 13091, Aug. 13, (e) Position indicator. If a retractable 1969; Amdt. 23–21, 43 FR 2318, Jan. 1978; Amdt. landing gear is used, there must be a 23–26, 45 FR 60171, Sept. 11, 1980; Amdt. 23–45, landing gear position indicator (as well 58 FR 42164, Aug. 6, 1993; Amdt. 23–49, 61 FR as necessary switches to actuate the 5166, Feb. 9, 1996] indicator) or other means to inform the pilot that each gear is secured in the § 23.731 Wheels. extended (or retracted) position. If (a) The maximum static load rating switches are used, they must be located of each wheel may not be less than the and coupled to the landing gear me- corresponding static ground reaction chanical system in a manner that pre- with— vents an erroneous indication of either (1) Design maximum weight; and ‘‘down and locked’’ if each gear is not in (2) Critical center of gravity. the fully extended position, or ‘‘up and (b) The maximum limit load rating of locked’’ if each landing gear is not in each wheel must equal or exceed the the fully retracted position. maximum radial limit load determined (f) Landing gear warning. For land- under the applicable ground load re- planes, the following aural or equally quirements of this part. effective landing gear warning devices must be provided: [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (1) A device that functions continu- amended by Amdt. 23–45, 58 FR 42165, Aug. 6, ously when one or more throttles are 1993] closed beyond the power settings nor- mally used for landing approach if the § 23.733 Tires. landing gear is not fully extended and (a) Each landing gear wheel must locked. A throttle stop may not be have a tire whose approved tire ratings used in place of an aural device. If (static and dynamic) are not exceed- there is a manual shutoff for the warn- ed—

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(1) By a load on each main wheel tire) where— to be compared to the static rating ap- KE=Kinetic energy per wheel (ft.-lb.); proved for such tires) equal to the cor- W=Design landing weight (lb.); responding static ground reaction V=Airplane speed in knots. V must be not

under the design maximum weight and less than VS√, the poweroff stalling speed critical center of gravity; and of the airplane at sea level, at the design (2) By a load on nose wheel tires (to landing weight, and in the landing configu- be compared with the dynamic rating ration; and approved for such tires) equal to the re- N=Number of main wheels with brakes. action obtained at the nose wheel, as- (b) Brakes must be able to prevent suming the mass of the airplane to be the wheels from rolling on a paved run- concentrated at the most critical cen- way with takeoff power on the critical ter of gravity and exerting a force of engine, but need not prevent movement 1.0 W downward and 0.31 W forward of the airplane with wheels locked. (where W is the design maximum (c) During the landing distance deter- weight), with the reactions distributed mination required by § 23.75, the pres- to the nose and main wheels by the principles of statics and with the drag sure on the wheel braking system must reaction at the ground applied only at not exceed the pressure specified by the wheels with brakes. brake manufacturer. (b) If specially constructed tires are (d) If antiskid devices are installed, used, the wheels must be plainly and the devices and associated systems conspicuously marked to that effect. must be designed so that no single The markings must include the make, probable malfunction or failure will re- size, number of plies, and identification sult in a hazardous loss of braking abil- marking of the proper tire. ity or directional control of the air- (c) Each tire installed on a retract- plane. able landing gear system must, at the (e) In addition, for commuter cat- maximum size of the tire type expected egory airplanes, the rejected takeoff in service, have a clearance to sur- brake kinetic energy capacity rating of rounding structure and systems that is each main wheel brake assembly must adequate to prevent contact between not be less than the kinetic energy ab- the tire and any part of the structure sorption requirements determined of systems. under either of the following methods— [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (1) The brake kinetic energy absorp- amended by Amdt. 23–7, 34 FR 13092, Aug. 13, tion requirements must be based on a 1969; Amdt. 23–17, 41 FR 55464, Dec. 20, 1976; conservative rational analysis of the Amdt. 23–45, 58 FR 42165, Aug. 6, 1993] sequence of events expected during a rejected takeoff at the design takeoff § 23.735 Brakes. weight. (a) Brakes must be provided. The (2) Instead of a rational analysis, the landing brake kinetic energy capacity kinetic energy absorption require- rating of each main wheel brake assem- ments for each main wheel brake as- bly must not be less than the kinetic sembly may be derived from the fol- energy absorption requirements deter- lowing formula— mined under either of the following methods: KE=0.0443 WV2N (1) The brake kinetic energy absorp- where, tion requirements must be based on a conservative rational analysis of the KE=Kinetic energy per wheel (ft.-lbs.); sequence of events expected during W=Design takeoff weight (lbs.); V=Ground speed, in knots, associated with landing at the design landing weight. the maximum value of V1 selected in ac- (2) Instead of a rational analysis, the cordance with § 23.51(c)(1); kinetic energy absorption require- N=Number of main wheels with brakes. ments for each main wheel brake as- sembly may be derived from the fol- [Amdt. 23–7, 34 FR 13092, Aug. 13, 1969, as lowing formula: amended by Amdt. 23–24, 44 FR 68742, Nov. 29, 1979; Amdt. 23–42, 56 FR 354, Jan. 3, 1991; KE=0.0443 WV2/N Amdt. 23–49, 61 FR 5166, Feb. 9, 1996]

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§ 23.737 Skis. afloat without capsizing in fresh water when— The maximum limit load rating for each ski must equal or exceed the max- (1) For airplanes of 5,000 pounds or imum limit load determined under the more maximum weight, any two adja- applicable ground load requirements of cent compartments are flooded; and this part. (2) For airplanes of 1,500 pounds up to, but not including, 5,000 pounds max- [Doc. No. 26269, 58 FR 42165, Aug. 6, 1993] imum weight, any single compartment is flooded. § 23.745 Nose/tail wheel steering. (b) Watertight doors in bulkheads (a) If nose/tail wheel steering is in- may be used for communication be- stalled, it must be demonstrated that tween compartments. its use does not require exceptional pilot skill during takeoff and landing, [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as in crosswinds, or in the event of an en- amended by Amdt. 23–45, 58 FR 42165, Aug. 6, 1993; Amdt. 23–48, 61 FR 5148, Feb. 9, 1996] gine failure; or its use must be limited to low speed maneuvering. § 23.757 Auxiliary floats. (b) Movement of the pilot’s steering control must not interfere with the re- Auxiliary floats must be arranged so traction or extension of the landing that, when completely submerged in gear. fresh water, they provide a righting moment of at least 1.5 times the upset- [Doc. No. 27806, 61 FR 5166, Feb. 9, 1996] ting moment caused by the seaplane or amphibian being tilted. FLOATS AND HULLS PERSONNEL AND CARGO § 23.751 Main float buoyancy. ACCOMMODATIONS (a) Each main float must have— (1) A buoyancy of 80 percent in excess § 23.771 Pilot compartment. of the buoyancy required by that float For each pilot compartment— to support its portion of the maximum (a) The compartment and its equip- weight of the seaplane or amphibian in ment must allow each pilot to perform fresh water; and his duties without unreasonable con- (2) Enough watertight compartments centration or fatigue; to provide reasonable assurance that (b) Where the flight crew are sepa- the seaplane or amphibian will stay rated from the passengers by a parti- afloat without capsizing if any two tion, an opening or openable window or compartments of any main float are door must be provided to facilitate flooded. communication between flight crew (b) Each main float must contain at and the passengers; and least four watertight compartments (c) The aerodynamic controls listed approximately equal in volume. in § 23.779, excluding cables and control [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as rods, must be located with respect to amended by Amdt. 23–45, 58 FR 42165, Aug. 6, the propellers so that no part of the 1993] pilot or the controls lies in the region between the plane of rotation of any § 23.753 Main float design. inboard propeller and the surface gen- Each seaplane main float must meet erated by a line passing through the the requirements of § 23.521. center of the propeller hub making an angle of 5 degrees forward or aft of the [Doc. No. 26269, 58 FR 42165, Aug. 6, 1993] plane of rotation of the propeller. § 23.755 Hulls. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (a) The hull of a hull seaplane or am- amended by Amdt. 23–14, 38 FR 31821, Nov. 19, phibian of 1,500 pounds or more max- 1973] imum weight must have watertight compartments designed and arranged § 23.773 Pilot compartment view. so that the hull auxiliary floats, and (a) Each pilot compartment must tires (if used), will keep the airplane be—

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(1) Arranged with sufficiently exten- fail-safe requirements of paragraph (d) sive, clear and undistorted view to en- of this section must be shown. able the pilot to safely taxi, takeoff, (d) If certification for operation approach, land, and perform any ma- above 25,000 feet is requested the wind- neuvers within the operating limita- shields, window panels, and canopies tions of the airplane. must be strong enough to withstand (2) Free from glare and reflections the maximum cabin pressure differen- that could interfere with the pilot’s vi- tial loads combined with critical aero- sion. Compliance must be shown in all dynamic pressure and temperature ef- operations for which certification is re- fects, after failure of any load-carrying quested; and element of the windshield, window (3) Designed so that each pilot is pro- panel, or canopy. tected from the elements so that mod- (e) The windshield and side windows erate rain conditions do not unduly im- forward of the pilot’s back when the pair the pilot’s view of the flight path pilot is seated in the normal flight po- in normal flight and while landing. sition must have a luminous transmit- (b) Each pilot compartment must tance value of not less than 70 percent. have a means to either remove or pre- (f) Unless operation in known or fore- vent the formation of fog or frost on an cast icing conditions is prohibited by area of the internal portion of the operating limitations, a means must be windshield and side windows suffi- provided to prevent or to clear accumu- ciently large to provide the view speci- lations of ice from the windshield so fied in paragraph (a)(1) of this section. that the pilot has adequate view for Compliance must be shown under all taxi, takeoff, approach, landing, and to expected external and internal ambient perform any maneuvers within the op- operating conditions, unless it can be erating limitations of the airplane. shown that the windshield and side (g) In the event of any probable sin- windows can be easily cleared by the gle failure, a transparency heating sys- pilot without interruption of normal tem must be incapable of raising the pilot duties. temperature of any windshield or win- [Doc. No. 26269, 58 FR 42165, Aug. 6, 1993; 71 dow to a point where there would be— FR 537, Jan. 5, 2006] (1) Structural failure that adversely affects the integrity of the cabin; or § 23.775 Windshields and windows. (2) There would be a danger of fire. (a) The internal panels of windshields (h) In addition, for commuter cat- and windows must be constructed of a egory airplanes, the following applies: nonsplintering material, such as non- (1) Windshield panes directly in front splintering safety glass. of the pilots in the normal conduct of (b) The design of windshields, win- their duties, and the supporting struc- dows, and canopies in pressurized air- tures for these panes, must withstand, planes must be based on factors pecu- without penetration, the impact of a liar to high altitude operation, includ- two-pound bird when the velocity of ing— the airplane (relative to the bird along (1) The effects of continuous and cy- the airplane’s flight path) is equal to clic pressurization loadings; the airplane’s maximum approach flap (2) The inherent characteristics of speed. the material used; and (2) The windshield panels in front of (3) The effects of temperatures and the pilots must be arranged so that, as- temperature gradients. suming the loss of vision through any (c) On pressurized airplanes, if cer- one panel, one or more panels remain tification for operation up to and in- available for use by a pilot seated at a cluding 25,000 feet is requested, an en- pilot station to permit continued safe closure canopy including a representa- flight and landing. tive part of the installation must be [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as subjected to special tests to account amended by Amdt. 23–7, 34 FR 13092, Aug. 13, for the combined effects of continuous 1969; Amdt. 23–45, 58 FR 42165, Aug. 6, 1993; 58 and cyclic pressurization loadings and FR 51970, Oct. 5, 1993; Amdt. 23–49, 61 FR 5166, flight loads, or compliance with the Feb. 9, 1996]

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§ 23.777 Cockpit controls. the left control(s) operates the left en- (a) Each cockpit control must be lo- gines(s) and the right control(s) oper- cated and (except where its function is ates the right engine(s). obvious) identified to provide conven- (2) On twin-engine airplanes with ient operation and to prevent confusion front and rear engine locations (tan- and inadvertent operation. dem), the left powerplant controls (b) The controls must be located and must operate the front engine and the arranged so that the pilot, when seat- right powerplant controls must operate ed, has full and unrestricted movement the rear engine. of each control without interference (f) Wing flap and auxiliary lift device from either his clothing or the cockpit controls must be located— structure. (1) Centrally, or to the right of the (c) Powerplant controls must be lo- pedestal or powerplant throttle control cated— centerline; and (1) For multiengine airplanes, on the (2) Far enough away from the landing pedestal or overhead at or near the gear control to avoid confusion. center of the cockpit; (g) The landing gear control must be (2) For single and tandem seated sin- located to the left of the throttle cen- gle-engine airplanes, on the left side terline or pedestal centerline. console or instrument panel; (3) For other single-engine airplanes (h) Each fuel feed selector control at or near the center of the cockpit, on must comply with § 23.995 and be lo- the pedestal, instrument panel, or cated and arranged so that the pilot overhead; and can see and reach it without moving (4) For airplanes, with side-by-side any seat or primary flight control pilot seats and with two sets of power- when his seat is at any position in plant controls, on left and right con- which it can be placed. soles. (1) For a mechanical fuel selector: (d) The control location order from (i) The indication of the selected fuel left to right must be power (thrust) valve position must be by means of a lever, propeller (rpm control), and mix- pointer and must provide positive iden- ture control (condition lever and fuel tification and feel (detent, etc.) of the cutoff for turbine-powered airplanes). selected position. Power (thrust) levers must be at least (ii) The position indicator pointer one inch higher or longer to make must be located at the part of the han- them more prominent than propeller dle that is the maximum dimension of (rpm control) or mixture controls. Car- the handle measured from the center of buretor heat or alternate air control rotation. must be to the left of the throttle or at (2) For electrical or electronic fuel least eight inches from the mixture selector: control when located other than on a (i) Digital controls or electrical pedestal. Carburetor heat or alternate switches must be properly labelled. air control, when located on a pedestal (ii) Means must be provided to indi- must be aft or below the power (thrust) cate to the flight crew the tank or lever. Supercharger controls must be function selected. Selector switch posi- located below or aft of the propeller controls. Airplanes with tandem seat- tion is not acceptable as a means of in- ing or single-place airplanes may uti- dication. The ‘‘off’’ or ‘‘closed’’ position lize control locations on the left side of must be indicated in red. the cabin compartment; however, loca- (3) If the fuel valve selector handle or tion order from left to right must be electrical or digital selection is also a power (thrust) lever, propeller (rpm fuel shut-off selector, the off position control) and mixture control. marking must be colored red. If a sepa- (e) Identical powerplant controls for rate emergency shut-off means is pro- each engine must be located to prevent vided, it also must be colored red. confusion as to the engines they con- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as trol. amended by Amdt. 23–7, 34 FR 13092, Aug. 13, (1) Conventional multiengine power- 1969; Amdt. 23–33, 51 FR 26656, July 24, 1986; plant controls must be located so that Amdt. 23–51, 61 FR 5136, Feb. 9, 1996]

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§ 23.779 Motion and effect of cockpit Motion and effect controls. (1) Powerplant Cockpit controls must be designed so controls: that they operate in accordance with Power Forward to increase for- the following movement and actuation: (thrust) ward thrust and rear- (a) Aerodynamic controls: lever. ward to increase rear- ward thrust. Motion and effect Propellers .. Forward to increase rpm. Mixture ...... Forward or upward for (1) Primary con- rich. trols: Fuel ...... Forward for open. Aileron ...... Right (clockwise) for right wing down. Carburetor, Forward or upward for air heat cold. Elevator ..... Rearward for nose up. or alter- Rudder ...... Right pedal forward for nate air. nose right. Super- Forward or upward for low (2) Secondary charger. blower. controls: Flaps (or Forward or up for flaps up Turbosuper- Forward, upward, or auxiliary or auxiliary device chargers. clockwise to increase lift de- stowed; rearward or pressure. vices). down for flaps down or Rotary con- Clockwise from off to full auxiliary device de- trols. on. ployed. (2) Auxiliary Trim tabs Switch motion or mechan- controls: (or equiv- ical rotation of control Fuel tank Right for right tanks, left alent). to produce similar rota- selector. for left tanks. tion of the airplane Landing Down to extend. about an axis parallel to gear. the axis control. Axis of Speed Aft to extend. roll trim control may be brakes. displaced to accommo- date comfortable actu- [Amdt. 23–33, 51 FR 26656, July 24, 1986, as ation by the pilot. For amended by Amdt. 23–51, 61 FR 5136, Feb. 9, single-engine airplanes, 1996] direction of pilot’s hand movement must be in the same sense as air- § 23.781 Cockpit control knob shape. plane response for rud- (a) Flap and landing gear control der trim if only a por- tion of a rotational ele- knobs must conform to the general ment is accessible. shapes (but not necessarily the exact sizes or specific proportions) in the fol- (b) Powerplant and auxiliary con- lowing figure: trols:

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(b) Powerplant control knobs must disk or any other potential hazard so conform to the general shapes (but not as to endanger persons using the door. necessarily the exact sizes or specific (c) Each external passenger or crew proportions) in the following figure: door must comply with the following [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 requirements: FR 258, Jan. 9, 1965, as amended by Amdt. 23– (1) There must be a means to lock 33, 51 FR 26657, July 24, 1986] and safeguard the door against inad- vertent opening during flight by per- § 23.783 Doors. sons, by cargo, or as a result of me- (a) Each closed cabin with passenger chanical failure. accommodations must have at least (2) The door must be openable from one adequate and easily accessible ex- the inside and the outside when the in- ternal door. ternal locking mechanism is in the (b) Passenger doors must not be lo- locked position. cated with respect to any propeller

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(3) There must be a means of opening (3) There must be a visual warning which is simple and obvious and is ar- means to signal a flight crewmember if ranged and marked inside and outside the external door is not fully closed so that the door can be readily located, and locked. The means must be de- unlocked, and opened, even in dark- signed so that any failure, or combina- ness. tion of failures, that would result in an (4) The door must meet the marking erroneous closed and locked indication requirements of § 23.811 of this part. is improbable for doors for which the (5) The door must be reasonably free initial opening movement is not in- from jamming as a result of fuselage ward. deformation in an emergency landing. (f) In addition, for commuter cat- (6) Auxiliary locking devices that are egory airplanes, the following require- actuated externally to the airplane ments apply: may be used but such devices must be (1) Each passenger entry door must overridden by the normal internal qualify as a floor level emergency exit. opening means. This exit must have a rectangular opening of not less than 24 inches wide (d) In addition, each external pas- by 48 inches high, with corner radii not senger or crew door, for a commuter greater than one-third the width of the category airplane, must comply with exit. the following requirements: (2) If an integral stair is installed at (1) Each door must be openable from a passenger entry door, the stair must both the inside and outside, even be designed so that, when subjected to though persons may be crowded the inertia loads resulting from the ul- against the door on the inside of the timate static load factors in airplane. § 23.561(b)(2) and following the collapse (2) If inward opening doors are used, of one or more legs of the landing gear, there must be a means to prevent occu- it will not reduce the effectiveness of pants from crowding against the door emergency egress through the pas- to the extent that would interfere with senger entry door. opening the door. (g) If lavatory doors are installed, (3) Auxiliary locking devices may be they must be designed to preclude an used. occupant from becoming trapped inside (e) Each external door on a com- the lavatory. If a locking mechanism is muter category airplane, each external installed, it must be capable of being door forward of any engine or propeller unlocked from outside of the lavatory. on a normal, utility, or acrobatic cat- egory airplane, and each door of the [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23– pressure vessel on a pressurized air- 36, 53 FR 30813, Aug. 15, 1988; Amdt. 23–46, 59 plane must comply with the following FR 25772, May 17, 1994; Amdt. 23–49, 61 FR requirements: 5166, Feb. 9, 1996] (1) There must be a means to lock and safeguard each external door, in- § 23.785 Seats, berths, litters, safety cluding cargo and service type doors, belts, and shoulder harnesses. against inadvertent opening in flight, There must be a seat or berth for by persons, by cargo, or as a result of each occupant that meets the fol- mechanical failure or failure of a single lowing: structural element, either during or (a) Each seat/restraint system and after closure. the supporting structure must be de- (2) There must be a provision for di- signed to support occupants weighing rect visual inspection of the locking at least 215 pounds when subjected to mechanism to determine if the exter- the maximum load factors cor- nal door, for which the initial opening responding to the specified flight and movement is not inward, is fully closed ground load conditions, as defined in and locked. The provisions must be dis- the approved operating envelope of the cernible, under operating lighting con- airplane. In addition, these loads must ditions, by a crewmember using a be multiplied by a factor of 1.33 in de- flashlight or an equivalent lighting termining the strength of all fittings source. and the attachment of—

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(1) Each seat to the structure; and (i) The cabin area surrounding each (2) Each safety belt and shoulder har- seat, including the structure, interior ness to the seat or structure. walls, instrument panel, control wheel, (b) Each forward-facing or aft-facing pedals, and seats within striking dis- seat/restraint system in normal, util- tance of the occupant’s head or torso ity, or acrobatic category airplanes (with the restraint system fastened) must consist of a seat, a safety belt, must be free of potentially injurious and a shoulder harness, with a metal- objects, sharp edges, protuberances, to-metal latching device, that are de- and hard surfaces. If energy absorbing signed to provide the occupant protec- designs or devices are used to meet this tion provisions required in § 23.562. requirement, they must protect the oc- Other seat orientations must provide cupant from serious injury when the the same level of occupant protection occupant is subjected to the inertia as a forward-facing or aft-facing seat loads resulting from the ultimate stat- with a safety belt and a shoulder har- ic load factors prescribed in ness, and must provide the protection § 23.561(b)(2) of this part, or they must provisions of § 23.562. comply with the occupant protection (c) For commuter category airplanes, provisions of § 23.562 of this part, as re- each seat and the supporting structure quired in paragraphs (b) and (c) of this must be designed for occupants weigh- section. (j) Each seat track must be fitted ing at least 170 pounds when subjected with stops to prevent the seat from to the inertia loads resulting from the sliding off the track. ultimate static load factors prescribed (k) Each seat/restraint system may in § 23.561(b)(2) of this part. Each occu- use design features, such as crushing or pant must be protected from serious separation of certain components, to head injury when subjected to the iner- reduce occupant loads when showing tia loads resulting from these load fac- compliance with the requirements of tors by a safety belt and shoulder har- § 23.562 of this part; otherwise, the sys- ness, with a metal-to-metal latching tem must remain intact. device, for the front seats and a safety (l) For the purposes of this section, a belt, or a safety belt and shoulder har- front seat is a seat located at a flight ness, with a metal-to-metal latching crewmember station or any seat lo- device, for each seat other than the cated alongside such a seat. front seats. (m) Each berth, or provisions for a (d) Each restraint system must have litter, installed parallel to the longitu- a single-point release for occupant dinal axis of the airplane, must be de- evacuation. signed so that the forward part has a (e) The restraint system for each padded end-board, canvas diaphragm, crewmember must allow the crew- or equivalent means that can with- member, when seated with the safety stand the load reactions from a 215- belt and shoulder harness fastened, to pound occupant when subjected to the perform all functions necessary for inertia loads resulting from the ulti- flight operations. mate static load factors of § 23.561(b)(2) (f) Each pilot seat must be designed of this part. In addition— for the reactions resulting from the ap- (1) Each berth or litter must have an plication of pilot forces to the primary occupant restraint system and may not flight controls as prescribed in § 23.395 have corners or other parts likely to of this part. cause serious injury to a person occu- (g) There must be a means to secure pying it during emergency landing con- each safety belt and shoulder harness, ditions; and when not in use, to prevent inter- (2) Occupant restraint system attach- ference with the operation of the air- ments for the berth or litter must plane and with rapid occupant egress in withstand the inertia loads resulting an emergency. from the ultimate static load factors of (h) Unless otherwise placarded, each § 23.561(b)(2) of this part. seat in a utility or acrobatic category (n) Proof of compliance with the stat- airplane must be designed to accommo- ic strength requirements of this sec- date an occupant wearing a parachute. tion for seats and berths approved as

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part of the type design and for seat and other occupants’ seats or where the berth installations may be shown by— flightcrew members’ compartment is (1) Structural analysis, if the struc- separated from the passenger compart- ture conforms to conventional airplane ment, there must be at least one illu- types for which existing methods of minated sign (using either letters or analysis are known to be reliable; symbols) notifying all passengers when (2) A combination of structural anal- seat belts should be fastened. Signs ysis and static load tests to limit load; that notify when seat belts should be or fastened must: (3) Static load tests to ultimate (a) When illuminated, be legible to loads. each person seated in the passenger [Amdt. 23–36, 53 FR 30813, Aug. 15, 1988; Amdt. compartment under all probable light- 23–36, 54 FR 50737, Dec. 11, 1989; Amdt. 23–49, ing conditions; and 61 FR 5167, Feb. 9, 1996] (b) Be installed so that a flightcrew member can, when seated at the § 23.787 Baggage and cargo compart- ments. flightcrew member’s station, turn the illumination on and off. (a) Each baggage and cargo compart- ment must: [Doc. No. 27806, 61 FR 5167, Feb. 9, 1996] (1) Be designed for its placarded max- imum weight of contents and for the § 23.803 Emergency evacuation. critical load distributions at the appro- (a) For commuter category airplanes, priate maximum load factors cor- an evacuation demonstration must be responding to the flight and ground conducted utilizing the maximum load conditions of this part. number of occupants for which certifi- (2) Have means to prevent the con- cation is desired. The demonstration tents of any compartment from becom- must be conducted under simulated ing a hazard by shifting, and to protect night conditions using only the emer- any controls, wiring, lines, equipment gency exits on the most critical side of or accessories whose damage or failure the airplane. The participants must be would affect safe operations. representative of average airline pas- (3) Have a means to protect occu- sengers with no prior practice or re- pants from injury by the contents of hearsal for the demonstration. Evacu- any compartment, located aft of the ation must be completed within 90 sec- occupants and separated by structure, onds. when the ultimate forward inertial load factor is 9g and assuming the max- (b) In addition, when certification to imum allowed baggage or cargo weight the emergency exit provisions of for the compartment. § 23.807(d)(4) is requested, only the (b) Designs that provide for baggage emergency lighting system required by or cargo to be carried in the same com- § 23.812 may be used to provide cabin in- partment as passengers must have a terior illumination during the evacu- means to protect the occupants from ation demonstration required in para- injury when the baggage or cargo is graph (a) of this section. subjected to the inertial loads result- [Amdt. 23–34, 52 FR 1831, Jan. 15, 1987, as ing from the ultimate static load fac- amended by Amdt. 23–46, 59 FR 25773, May 17, tors of § 23.561(b)(3), assuming the max- 1994] imum allowed baggage or cargo weight for the compartment. § 23.805 Flightcrew emergency exits. (c) For airplanes that are used only For airplanes where the proximity of for the carriage of cargo, the flightcrew the passenger emergency exits to the emergency exits must meet the re- flightcrew area does not offer a conven- quirements of § 23.807 under any cargo ient and readily accessible means of loading conditions. evacuation for the flightcrew, the fol- [Doc. No. 27806, 61 FR 5167, Feb. 9, 1996] lowing apply: (a) There must be either one emer- § 23.791 Passenger information signs. gency exit on each side of the airplane, For those airplanes in which the or a top hatch emergency exit, in the flightcrew members cannot observe the flightcrew area;

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(b) Each emergency exit must be lo- must be designed to be overridden by cated to allow rapid evacuation of the the normal internal opening means. crew and have a size and shape of at The inside handles of emergency exits least a 19- by 20-inch unobstructed rec- that open outward must be adequately tangular opening; and protected against inadvertent oper- (c) For each emergency exit that is ation. In addition, each emergency exit not less than six feet from the ground, must— an assisting means must be provided. (1) Be readily accessible, requiring no The assisting means may be a rope or exceptional agility to be used in emer- any other means demonstrated to be gencies; suitable for the purpose. If the assist- (2) Have a method of opening that is ing means is a rope, or an approved de- simple and obvious; vice equivalent to a rope, it must be— (3) Be arranged and marked for easy (1) Attached to the fuselage structure location and operation, even in dark- at or above the top of the emergency ness; exit opening or, for a device at a pilot’s (4) Have reasonable provisions emergency exit window, at another ap- proved location if the stowed device, or against jamming by fuselage deforma- its attachment, would reduce the pi- tion; and lot’s view; and (5) In the case of acrobatic category (2) Able (with its attachment) to airplanes, allow each occupant to aban- withstand a 400-pound static load. don the airplane at any speed between VSO and VD; and [Doc. No. 26324, 59 FR 25773, May 17, 1994] (6) In the case of utility category air- § 23.807 Emergency exits. planes certificated for spinning, allow each occupant to abandon the airplane (a) Number and location. Emergency at the highest speed likely to be exits must be located to allow escape achieved in the maneuver for which the without crowding in any probable airplane is certificated. crash attitude. The airplane must have (c) Tests. The proper functioning of at least the following emergency exits: each emergency exit must be shown by (1) For all airplanes with a seating tests. capacity of two or more, excluding air- planes with canopies, at least one (d) Doors and exits. In addition, for emergency exit on the opposite side of commuter category airplanes, the fol- the cabin from the main door specified lowing requirements apply: in § 23.783 of this part. (1) In addition to the passenger entry (2) [Reserved] door— (3) If the pilot compartment is sepa- (i) For an airplane with a total pas- rated from the cabin by a door that is senger seating capacity of 15 or fewer, likely to block the pilot’s escape in a an emergency exit, as defined in para- minor crash, there must be an exit in graph (b) of this section, is required on the pilot’s compartment. The number each side of the cabin; and of exits required by paragraph (a)(1) of (ii) For an airplane with a total pas- this section must then be separately senger seating capacity of 16 through determined for the passenger compart- 19, three emergency exits, as defined in ment, using the seating capacity of paragraph (b) of this section, are re- that compartment. quired with one on the same side as the (4) Emergency exits must not be lo- passenger entry door and two on the cated with respect to any propeller side opposite the door. disk or any other potential hazard so (2) A means must be provided to lock as to endanger persons using that exit. each emergency exit and to safeguard (b) Type and operation. Emergency against its opening in flight, either in- exits must be movable windows, panels, advertently by persons or as a result of canopies, or external doors, openable mechanical failure. In addition, a from both inside and outside the air- means for direct visual inspection of plane, that provide a clear and unob- the locking mechanism must be pro- structed opening large enough to admit vided to determine that each emer- a 19-by-26-inch ellipse. Auxiliary lock- gency exit for which the initial opening ing devices used to secure the airplane movement is outward is fully locked.

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(3) Each required emergency exit, ex- greater than one-third the width of the cept floor level exits, must be located exit. over the wing or, if not less than six [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as feet from the ground, must be provided amended by Amdt. 23–7, 34 FR 13092, Aug. 13, with an acceptable means to assist the 1969; Amdt. 23–10, 36 FR 2864, Feb. 11, 1971; occupants to descend to the ground. Amdt. 23–34, 52 FR 1831, Jan. 15, 1987; Amdt. Emergency exits must be distributed as 23–36, 53 FR 30814, Aug. 15, 1988; 53 FR 34194, uniformly as practical, taking into ac- Sept. 2, 1988; Amdt. 23–46, 59 FR 25773, May count passenger seating configuration. 17, 1994; Amdt. 23–49, 61 FR 5167, Feb. 9, 1996] (4) Unless the applicant has complied § 23.811 Emergency exit marking. with paragraph (d)(1) of this section, (a) Each emergency exit and external there must be an emergency exit on door in the passenger compartment the side of the cabin opposite the pas- must be externally marked and readily senger entry door, provided that— identifiable from outside the airplane (i) For an airplane having a pas- by— senger seating configuration of nine or (1) A conspicuous visual identifica- fewer, the emergency exit has a rectan- tion scheme; and gular opening measuring not less than (2) A permanent decal or placard on 19 inches by 26 inches high with corner or adjacent to the emergency exit radii not greater than one-third the which shows the means of opening the width of the exit, located over the emergency exit, including any special wing, with a step up inside the airplane instructions, if applicable. of not more than 29 inches and a step (b) In addition, for commuter cat- down outside the airplane of not more egory airplanes, these exits and doors than 36 inches; must be internally marked with the (ii) For an airplane having a pas- word ‘‘exit’’ by a sign which has white senger seating configuration of 10 to 19 letters 1 inch high on a red background passengers, the emergency exit has a 2 inches high, be self-illuminated or rectangular opening measuring not less independently, internally electrically than 20 inches wide by 36 inches high, illuminated, and have a minimum with corner radii not greater than one- brightness of at least 160 micro- third the width of the exit, and with a lamberts. The color may be reversed if step up inside the airplane of not more the passenger compartment illumina- than 20 inches. If the exit is located tion is essentially the same. over the wing, the step down outside (c) In addition, when certification to the airplane may not exceed 27 inches; the emergency exit provisions of and § 23.807(d)(4) is requested, the following (iii) The airplane complies with the apply: additional requirements of (1) Each emergency exit, its means of access, and its means of opening, must §§ 23.561(b)(2)(iv), 23.803(b), 23.811(c), be conspicuously marked; 23.812, 23.813(b), and 23.815. (2) The identity and location of each (e) For multiengine airplanes, ditch- emergency exit must be recognizable ing emergency exits must be provided from a distance equal to the width of in accordance with the following re- the cabin; quirements, unless the emergency exits (3) Means must be provided to assist required by paragraph (a) or (d) of this occupants in locating the emergency section already comply with them: exits in conditions of dense smoke; (1) One exit above the waterline on (4) The location of the operating han- each side of the airplane having the di- dle and instructions for opening each mensions specified in paragraph (b) or emergency exit from inside the air- (d) of this section, as applicable; and plane must be shown by marking that (2) If side exits cannot be above the is readable from a distance of 30 inches; waterline, there must be a readily ac- (5) Each passenger entry door oper- cessible overhead hatch emergency exit ating handle must— that has a rectangular opening meas- (i) Be self-illuminated with an initial uring not less than 20 inches wide by 36 brightness of at least 160 micro- inches long, with corner radii not lamberts; or

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(ii) Be conspicuously located and well ing system is independent of the power illuminated by the emergency lighting supply to the main lighting system. even in conditions of occupant crowd- (b) There must be a crew warning ing at the door; light that illuminates in the cockpit (6) Each passenger entry door with a when power is on in the airplane and locking mechanism that is released by the emergency lighting control device rotary motion of the handle must be is not armed. marked— (c) The emergency lights must be op- (i) With a red arrow, with a shaft of erable manually from the flightcrew at least three-fourths of an inch wide station and be provided with automatic and a head twice the width of the shaft, activation. The cockpit control device extending along at least 70 degrees of must have ‘‘on,’’ ‘‘off,’’ and ‘‘armed’’ posi- arc at a radius approximately equal to tions so that, when armed in the cock- three-fourths of the handle length; pit, the lights will operate by auto- (ii) So that the center line of the exit matic activation. handle is within ± one inch of the pro- (d) There must be a means to safe- jected point of the arrow when the han- guard against inadvertent operation of dle has reached full travel and has re- the cockpit control device from the leased the locking mechanism; ‘‘armed’’ or ‘‘on’’ positions. (iii) With the word ‘‘open’’ in red let- (e) The cockpit control device must ters, one inch high, placed horizontally have provisions to allow the emergency near the head of the arrow; and lighting system to be armed or acti- (7) In addition to the requirements of vated at any time that it may be need- paragraph (a) of this section, the exter- ed. nal marking of each emergency exit (f) When armed, the emergency light- must— ing system must activate and remain (i) Include a 2-inch colorband out- lighted when— lining the exit; and (1) The normal electrical power of the airplane is lost; or (ii) Have a color contrast that is (2) The airplane is subjected to an readily distinguishable from the sur- impact that results in a deceleration in rounding fuselage surface. The contrast excess of 2g and a velocity change in must be such that if the reflectance of excess of 3.5 feet-per-second, acting the darker color is 15 percent or less, along the longitudinal axis of the air- the reflectance of the lighter color plane; or must be at least 45 percent. ‘‘Reflec- (3) Any other emergency condition tance’’ is the ratio of the luminous flux exists where automatic activation of reflected by a body to the luminous the emergency lighting is necessary to flux it receives. When the reflectance aid with occupant evacuation. of the darker color is greater than 15 (g) The emergency lighting system percent, at least a 30 percent difference must be capable of being turned off and between its reflectance and the reflec- reset by the flightcrew after automatic tance of the lighter color must be pro- activation. vided. (h) The emergency lighting system [Amdt. 23–36, 53 FR 30814, Aug. 15, 1988; 53 FR must provide internal lighting, includ- 34194, Sept. 2, 1988, as amended by Amdt. 23– ing— 46, 59 FR 25773, May 17, 1994] (1) Illuminated emergency exit mark- ing and locating signs, including those § 23.812 Emergency lighting. required in § 23.811(b); When certification to the emergency (2) Sources of general illumination in exit provisions of § 23.807(d)(4) is re- the cabin that provide an average illu- quested, the following apply: mination of not less than 0.05 foot-can- (a) An emergency lighting system, dle and an illumination at any point of independent of the main cabin lighting not less than 0.01 foot-candle when system, must be installed. However, measured along the center line of the the source of general cabin illumina- main passenger aisle(s) and at the seat tion may be common to both the emer- armrest height; and gency and main lighting systems if the (3) Floor proximity emergency escape power supply to the emergency light- path marking that provides emergency

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evacuation guidance for the airplane assistance in evacuation of passengers occupants when all sources of illumina- without reducing the unobstructed tion more than 4 feet above the cabin width of the passageway below 20 aisle floor are totally obscured. inches. (i) The energy supply to each emer- (3) If it is necessary to pass through gency lighting unit must provide the a passageway between passenger com- required level of illumination for at partments to reach a required emer- least 10 minutes at the critical ambient gency exit from any seat in the pas- conditions after activation of the senger cabin, the passageway must be emergency lighting system. unobstructed; however, curtains may (j) If rechargeable batteries are used be used if they allow free entry as the energy supply for the emergency through the passageway. lighting system, they may be re- charged from the main electrical power (4) No door may be installed in any system of the airplane provided the partition between passenger compart- charging circuit is designed to preclude ments unless that door has a means to inadvertent battery discharge into the latch it in the open position. The latch- charging circuit faults. If the emer- ing means must be able to withstand gency lighting system does not include the loads imposed upon it by the door a charging circuit, battery condition when the door is subjected to the iner- monitors are required. tia loads resulting from the ultimate (k) Components of the emergency static load factors prescribed in lighting system, including batteries, § 23.561(b)(2). wiring, relays, lamps, and switches, (5) If it is necessary to pass through must be capable of normal operation a doorway separating the passenger after being subjected to the inertia cabin from other areas to reach a re- forces resulting from the ultimate load quired emergency exit from any pas- factors prescribed in § 23.561(b)(2). senger seat, the door must have a (l) The emergency lighting system means to latch it in the open position. must be designed so that after any sin- The latching means must be able to gle transverse vertical separation of withstand the loads imposed upon it by the fuselage during a crash landing: the door when the door is subjected to (1) At least 75 percent of all elec- the inertia loads resulting from the ul- trically illuminated emergency lights timate static load factors prescribed in required by this section remain opera- § 23.561(b)(2). tive; and (2) Each electrically illuminated exit [Amdt. 23–36, 53 FR 30815, Aug. 15, 1988, as sign required by § 23.811 (b) and (c) re- amended by Amdt. 23–46, 59 FR 25774, May 17, mains operative, except those that are 1994] directly damaged by the fuselage sepa- ration. § 23.815 Width of aisle. (a) Except as provided in paragraph [Doc. No. 26324, 59 FR 25774, May 17, 1994] (b) of this section, for commuter cat- § 23.813 Emergency exit access. egory airplanes, the width of the main (a) For commuter category airplanes, passenger aisle at any point between access to window-type emergency exits seats must equal or exceed the values may not be obstructed by seats or seat in the following table: backs. Minimum main passenger aisle width (b) In addition, when certification to Number of pas- senger seats Less than 25 25 inches and the emergency exit provisions of inches from floor more from floor § 23.807(d)(4) is requested, the following emergency exit access must be pro- 10 through 19 ...... 9 inches ...... 15 inches. vided: (1) The passageway leading from the (b) When certification to the emer- aisle to the passenger entry door must gency exist provisions of § 23.807(d)(4) is be unobstructed and at least 20 inches requested, the main passenger aisle wide. width at any point between the seats (2) There must be enough space next must equal or exceed the following val- to the passenger entry door to allow ues:

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Minimum main passenger large enough so that the failure of any aisle width (inches) one valve would not cause an appre- Number of passenger seats Less than 25 inches ciable rise in the pressure differential. 25 inches and more The pressure differential is positive from floor from floor when the internal pressure is greater 10 or fewer ...... 1 12 15 than the external. 11 through 19 ...... 12 20 (2) Two reverse pressure differential 1 A narrower width not less than 9 inches may be approved relief valves (or their equivalent) to when substantiated by tests found necessary by the Administrator. automatically prevent a negative pres- sure differential that would damage [Amdt. 23–34, 52 FR 1831, Jan. 15, 1987, as the structure. However, one valve is amended by Amdt. 23–46, 59 FR 25774, May 17, enough if it is of a design that reason- 1994] ably precludes its malfunctioning. (3) A means by which the pressure § 23.831 Ventilation. differential can be rapidly equalized. (a) Each passenger and crew compart- (4) An automatic or manual regulator ment must be suitably ventilated. Car- for controlling the intake or exhaust bon monoxide concentration may not airflow, or both, for maintaining the exceed one part in 20,000 parts of air. required internal pressures and airflow (b) For pressurized airplanes, the rates. ventilating air in the flightcrew and (5) Instruments to indicate to the passenger compartments must be free pilot the pressure differential, the of harmful or hazardous concentrations cabin pressure altitude, and the rate of of gases and vapors in normal oper- change of cabin pressure altitude. ations and in the event of reasonably (6) Warning indication at the pilot probable failures or malfunctioning of station to indicate when the safe or the ventilating, heating, pressuriza- preset pressure differential is exceeded tion, or other systems and equipment. and when a cabin pressure altitude of If accumulation of hazardous quan- 10,000 feet is exceeded. tities of smoke in the cockpit area is (7) A warning placard for the pilot if reasonably probable, smoke evacuation the structure is not designed for pres- must be readily accomplished starting sure differentials up to the maximum with full pressurization and without relief valve setting in combination depressurizing beyond safe limits. with landing loads. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 (8) A means to stop rotation of the FR 258, Jan. 9, 1965, as amended by Amdt. 23– compressor or to divert airflow from 34, 52 FR 1831, Jan. 15, 1987; Amdt. 23–42, 56 the cabin if continued rotation of an FR 354, Jan. 3, 1991] engine-driven cabin compressor or con- tinued flow of any compressor bleed air PRESSURIZATION will create a hazard if a malfunction § 23.841 Pressurized cabins. occurs. (a) If certification for operation over [Amdt. 23–14, 38 FR 31822, Nov. 19, 1973, as 25,000 feet is requested, the airplane amended by Amdt. 23–17, 41 FR 55464, Dec. 20, must be able to maintain a cabin pres- 1976; Amdt. 23–49, 61 FR 5167, Feb. 9, 1996] sure altitude of not more than 15,000 feet in event of any probable failure or § 23.843 Pressurization tests. malfunction in the pressurization sys- (a) Strength test. The complete pres- tem. surized cabin, including doors, win- (b) Pressurized cabins must have at dows, canopy, and valves, must be test- least the following valves, controls, ed as a pressure vessel for the pressure and indicators, for controlling cabin differential specified in § 23.365(d). pressure: (b) Functional tests. The following (1) Two pressure relief valves to auto- functional tests must be performed: matically limit the positive pressure (1) Tests of the functioning and ca- differential to a predetermined value pacity of the positive and negative at the maximum rate of flow delivered pressure differential valves, and of the by the pressure source. The combined emergency release valve, to simulate capacity of the relief valves must be the effects of closed regulator valves.

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(2) Tests of the pressurization system (1) There must be an adequate num- to show proper functioning under each ber of self-contained, removable ash- possible condition of pressure, tem- trays; and perature, and moisture, up to the max- (2) Where the crew compartment is imum altitude for which certification separated from the passenger compart- is requested. ment, there must be at least one illu- (3) Flight tests, to show the perform- minated sign (using either letters or ance of the pressure supply, pressure symbols) notifying all passengers when and flow regulators, indicators, and smoking is prohibited. Signs which no- warning signals, in steady and stepped tify when smoking is prohibited must— climbs and descents at rates cor- (i) When illuminated, be legible to responding to the maximum attainable each passenger seated in the passenger within the operating limitations of the cabin under all probable lighting condi- airplane, up to the maximum altitude tions; and for which certification is requested. (ii) Be so constructed that the crew (4) Tests of each door and emergency can turn the illumination on and off; exit, to show that they operate prop- and erly after being subjected to the flight (d) In addition, for commuter cat- tests prescribed in paragraph (b)(3) of egory airplanes the following require- this section. ments apply: (1) Each disposal receptacle for tow- FIRE PROTECTION els, paper, or waste must be fully en- closed and constructed of at least fire § 23.851 Fire extinguishers. resistant materials and must contain fires likely to occur in it under normal (a) There must be at least one hand use. The ability of the disposal recep- fire extinguisher for use in the pilot tacle to contain those fires under all compartment that is located within probable conditions of wear, misalign- easy access of the pilot while seated. ment, and ventilation expected in serv- (b) There must be at least one hand ice must be demonstrated by test. A fire extinguisher located conveniently placard containing the legible words in the passenger compartment— ‘‘No Cigarette Disposal’’ must be lo- (1) Of each airplane accommodating cated on or near each disposal recep- more than 6 passengers; and tacle door. (2) Of each commuter category air- (2) Lavatories must have ‘‘No Smok- plane. ing’’ or ‘‘No Smoking in Lavatory’’ plac- (c) For hand fire extinguishers, the ards located conspicuously on each side following apply: of the entry door and self-contained, (1) The type and quantity of each ex- removable ashtrays located conspicu- tinguishing agent used must be appro- ously on or near the entry side of each priate to the kinds of fire likely to lavatory door, except that one ashtray occur where that agent is to be used. may serve more than one lavatory door (2) Each extinguisher for use in a per- if it can be seen from the cabin side of sonnel compartment must be designed each lavatory door served. The plac- to minimize the hazard of toxic gas ards must have red letters at least 1⁄2 concentrations. inch high on a white background at least 1 inch high (a ‘‘No Smoking’’ sym- [Doc. No. 26269, 58 FR 42165, Aug. 6, 1993] bol may be included on the placard). (3) Materials (including finishes or § 23.853 Passenger and crew compart- ment interiors. decorative surfaces applied to the ma- terials) used in each compartment oc- For each compartment to be used by cupied by the crew or passengers must the crew or passengers: meet the following test criteria as ap- (a) The materials must be at least plicable: flame-resistant; (i) Interior ceiling panels, interior (b) [Reserved] wall panels, partitions, galley struc- (c) If smoking is to be prohibited, ture, large cabinet walls, structural there must be a placard so stating, and flooring, and materials used in the con- if smoking is to be allowed— struction of stowage compartments

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(other than underseat stowage com- ments, may not have an average burn partments and compartments for stow- rate greater than 2.5 inches per minute ing small items such as magazines and when tested horizontally in accordance maps) must be self-extinguishing when with the applicable portions of appen- tested vertically in accordance with dix F of this part or by other approved the applicable portions of appendix F equivalent methods. of this part or by other equivalent (v) Except for electrical wire cable methods. The average burn length may insulation, and for small parts (such as not exceed 6 inches and the average knobs, handles, rollers, fasteners, clips, flame time after removal of the flame grommets, rub strips, pulleys, and source may not exceed 15 seconds. small electrical parts) that the Admin- Drippings from the test specimen may istrator finds would not contribute sig- not continue to flame for more than an nificantly to the propagation of a fire, average of 3 seconds after falling. materials in items not specified in (ii) Floor covering, textiles (includ- paragraphs (d)(3)(i), (ii), (iii), or (iv) of ing draperies and upholstery), seat this section may not have a burn rate cushions, padding, decorative and non- greater than 4.0 inches per minute decorative coated fabrics, leather, when tested horizontally in accordance trays and galley furnishings, electrical with the applicable portions of appen- conduit, thermal and acoustical insula- dix F of this part or by other approved tion and insulation covering, air duct- equivalent methods. ing, joint and edge covering, cargo (e) Lines, tanks, or equipment con- compartment liners, insulation blan- taining fuel, oil, or other flammable kets, cargo covers and transparencies, fluids may not be installed in such molded and thermoformed parts, air compartments unless adequately ducting joints, and trim strips (decora- shielded, isolated, or otherwise pro- tive and chafing), that are constructed tected so that any breakage or failure of materials not covered in paragraph of such an item would not create a haz- (d)(3)(iv) of this section must be self ex- ard. tinguishing when tested vertically in accordance with the applicable por- (f) Airplane materials located on the tions of appendix F of this part or cabin side of the firewall must be self- other approved equivalent methods. extinguishing or be located at such a The average burn length may not ex- distance from the firewall, or otherwise ceed 8 inches and the average flame protected, so that ignition will not time after removal of the flame source occur if the firewall is subjected to a may not exceed 15 seconds. Drippings flame temperature of not less than from the test specimen may not con- 2,000 degrees F for 15 minutes. For self- tinue to flame for more than an aver- extinguishing materials (except elec- age of 5 seconds after falling. trical wire and cable insulation and (iii) Motion picture film must be small parts that the Administrator safety film meeting the Standard Spec- finds would not contribute signifi- ifications for Safety Photographic cantly to the propagation of a fire), a Film PH1.25 (available from the Amer- vertifical self-extinguishing test must ican National Standards Institute, 1430 be conducted in accordance with appen- Broadway, New York, N.Y. 10018) or an dix F of this part or an equivalent FAA approved equivalent. If the film method approved by the Adminis- travels through ducts, the ducts must trator. The average burn length of the meet the requirements of paragraph material may not exceed 6 inches and (d)(3)(ii) of this section. the average flame time after removal (iv) Acrylic windows and signs, parts of the flame source may not exceed 15 constructed in whole or in part of elas- seconds. Drippings from the material tomeric materials, edge-lighted instru- test specimen may not continue to ment assemblies consisting of two or flame for more than an average of 3 more instruments in a common hous- seconds after falling. ing, seatbelts, shoulder harnesses, and [Amdt. 23–14, 23 FR 31822, Nov. 19, 1973, as cargo and baggage tiedown equipment, amended by Amdt. 23–23, 43 FR 50593, Oct. 30, including containers, bins, pallets, etc., 1978; Amdt. 23–25, 45 FR 7755, Feb. 4, 1980; used in passenger or crew compart- Amdt. 23–34, 52 FR 1831, Jan. 15, 1987]

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§ 23.855 Cargo and baggage compart- (i) Be damaged by heater malfunc- ment fire protection. tioning; or (a) Sources of heat within each cargo (ii) Allow flammable fluids or vapors and baggage compartment that are ca- to reach the heater in case of leakage. pable of igniting the compartment con- (2) The region surrounding the heat- tents must be shielded and insulated to er, if the heater fuel system has fit- prevent such ignition. tings that, if they leaked, would allow (b) Each cargo and baggage compart- fuel vapor to enter this region. ment must be constructed of materials (3) The part of the ventilating air that meet the appropriate provisions of passage that surrounds the combustion § 23.853(d)(3). chamber. (c) In addition, for commuter cat- (b) Ventilating air ducts. Each ven- egory airplanes, each cargo and bag- tilating air duct passing through any gage compartment must: fire region must be fireproof. In addi- (1) Be located where the presence of a tion— fire would be easily discovered by the (1) Unless isolation is provided by pilots when seated at their duty sta- fireproof valves or by equally effective tion, or it must be equipped with a means, the ventilating air duct down- smoke or fire detector system to give a stream of each heater must be fireproof warning at the pilots’ station, and pro- for a distance great enough to ensure vide sufficient access to enable a pilot that any fire originating in the heater to effectively reach any part of the can be contained in the duct; and compartment with the contents of a (2) Each part of any ventilating duct hand held fire extinguisher, or passing through any region having a (2) Be equipped with a smoke or fire flammable fluid system must be con- detector system to give a warning at structed or isolated from that system the pilots’ station and have ceiling and so that the malfunctioning of any com- sidewall liners and floor panels con- ponent of that system cannot intro- structed of materials that have been duce flammable fluids or vapors into subjected to and meet the 45 degree the ventilating airstream. angle test of appendix F of this part. (c) Combustion air ducts. Each com- The flame may not penetrate (pass bustion air duct must be fireproof for a through) the material during applica- distance great enough to prevent dam- tion of the flame or subsequent to its age from backfiring or reverse flame removal. The average flame time after propagation. In addition— removal of the flame source may not (1) No combustion air duct may have exceed 15 seconds, and the average glow a common opening with the ventilating time may not exceed 10 seconds. The airstream unless flames from backfires compartment must be constructed to or reverse burning cannot enter the provide fire protection that is not less ventilating airstream under any oper- than that required of its individual ating condition, including reverse flow panels; or or malfunctioning of the heater or its (3) Be constructed and sealed to con- associated components; and tain any fire within the compartment. (2) No combustion air duct may re- [Doc. No. 27806, 61 FR 5167, Feb. 9, 1996] strict the prompt relief of any backfire that, if so restricted, could cause heat- § 23.859 Combustion heater fire pro- er failure. tection. (d) Heater controls: general. Provision (a) Combustion heater fire regions. The must be made to prevent the hazardous following combustion heater fire re- accumulation of water or ice on or in gions must be protected from fire in ac- any heater control component, control cordance with the applicable provisions system tubing, or safety control. of §§ 23.1182 through 23.1191 and 23.1203: (e) Heater safety controls. (1) Each (1) The region surrounding the heat- combustion heater must have the fol- er, if this region contains any flam- lowing safety controls: mable fluid system components (ex- (i) Means independent of the compo- cluding the heater fuel system) that nents for the normal continuous con- could— trol of air temperature, airflow, and

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fuel flow must be provided to auto- by shrouds so that no leakage from matically shut off the ignition and fuel those components can enter the ven- supply to that heater at a point remote tilating airstream. from that heater when any of the fol- (i) Drains. There must be means to lowing occurs: safely drain fuel that might accumu- (A) The heater exchanger tempera- late within the combustion chamber or ture exceeds safe limits. the heater exchanger. In addition— (B) The ventilating air temperature (1) Each part of any drain that oper- exceeds safe limits. ates at high temperatures must be pro- (C) The combustion airflow becomes tected in the same manner as heater inadequate for safe operation. exhausts; and (D) The ventilating airflow becomes (2) Each drain must be protected inadequate for safe operation. from hazardous ice accumulation under (ii) Means to warn the crew when any any operating condition. heater whose heat output is essential [Amdt. 23–27, 45 FR 70387, Oct. 23, 1980] for safe operation has been shut off by the automatic means prescribed in § 23.863 Flammable fluid fire protec- paragraph (e)(1)(i) of this section. tion. (2) The means for complying with (a) In each area where flammable paragraph (e)(1)(i) of this section for fluids or vapors might escape by leak- any individual heater must— age of a fluid system, there must be (i) Be independent of components means to minimize the probability of serving any other heater whose heat ignition of the fluids and vapors, and output is essential for safe operations; the resultant hazard if ignition does and occur. (ii) Keep the heater off until re- (b) Compliance with paragraph (a) of started by the crew. this section must be shown by analysis (f) Air intakes. Each combustion and or tests, and the following factors must ventilating air intake must be located be considered: so that no flammable fluids or vapors (1) Possible sources and paths of fluid can enter the heater system under any leakage, and means of detecting leak- operating condition— age. (1) During normal operation; or (2) Flammability characteristics of (2) As a result of the malfunctioning fluids, including effects of any combus- of any other component. tible or absorbing materials. (g) Heater exhaust. Heater exhaust systems must meet the provisions of (3) Possible ignition sources, includ- §§ 23.1121 and 23.1123. In addition, there ing electrical faults, overheating of must be provisions in the design of the equipment, and malfunctioning of pro- heater exhaust system to safely expel tective devices. the products of combustion to prevent (4) Means available for controlling or the occurrence of— extinguishing a fire, such as stopping (1) Fuel leakage from the exhaust to flow of fluids, shutting down equip- surrounding compartments; ment, fireproof containment, or use of (2) Exhaust gas impingement on sur- extinguishing agents. rounding equipment or structure; (5) Ability of airplane components (3) Ignition of flammable fluids by that are critical to safety of flight to the exhaust, if the exhaust is in a com- withstand fire and heat. partment containing flammable fluid (c) If action by the flight crew is re- lines; and quired to prevent or counteract a fluid (4) Restrictions in the exhaust sys- fire (e.g. equipment shutdown or actu- tem to relieve backfires that, if so re- ation of a fire extinguisher), quick act- stricted, could cause heater failure. ing means must be provided to alert the crew. (h) Heater fuel systems. Each heater fuel system must meet each power- (d) Each area where flammable fluids plant fuel system requirement affect- or vapors might escape by leakage of a ing safe heater operation. Each heater fluid system must be identified and de- fuel system component within the ven- fined. tilating airstream must be protected [Amdt. 23–23, 43 FR 50593, Oct. 30, 1978]

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§ 23.865 Fire protection of flight con- Subpart E—Powerplant trols, engine mounts, and other flight structure. GENERAL Flight controls, engine mounts, and § 23.901 Installation. other flight structure located in des- ignated fire zones, or in adjacent areas (a) For the purpose of this part, the airplane powerplant installation in- that would be subjected to the effects cludes each component that— of fire in the designated fire zones, (1) Is necessary for propulsion; and must be constructed of fireproof mate- (2) Affects the safety of the major rial or be shielded so that they are ca- propulsive units. pable of withstanding the effects of a (b) Each powerplant installation fire. Engine vibration isolators must must be constructed and arranged to— incorporate suitable features to ensure (1) Ensure safe operation to the max- that the engine is retained if the non- imum altitude for which approval is re- fireproof portions of the isolators dete- quested. riorate from the effects of a fire. (2) Be accessible for necessary inspec- tions and maintenance. [Doc. No. 27805, 61 FR 5148, Feb. 9, 1996] (c) Engine cowls and nacelles must be easily removable or openable by the ELECTRICAL BONDING AND LIGHTNING pilot to provide adequate access to and PROTECTION exposure of the engine compartment for preflight checks. § 23.867 Electrical bonding and protec- (d) Each turbine engine installation tion against lightning and static electricity. must be constructed and arranged to— (1) Result in carcass vibration char- (a) The airplane must be protected acteristics that do not exceed those es- against catastrophic effects from light- tablished during the type certification ning. of the engine. (b) For metallic components, compli- (2) Ensure that the capability of the ance with paragraph (a) of this section installed engine to withstand the in- may be shown by— gestion of rain, hail, ice, and birds into (1) Bonding the components properly the engine inlet is not less than the ca- to the airframe; or pability established for the engine (2) Designing the components so that itself under § 23.903(a)(2). (e) The installation must comply a strike will not endanger the airplane. with— (c) For nonmetallic components, (1) The instructions provided under compliance with paragraph (a) of this the engine type certificate and the pro- section may be shown by— peller type certificate. (1) Designing the components to min- (2) The applicable provisions of this imize the effect of a strike; or subpart. (2) Incorporating acceptable means of (f) Each auxiliary power unit instal- diverting the resulting electrical cur- lation must meet the applicable por- rent so as not to endanger the airplane. tions of this part. [Amdt. 23–7, 34 FR 13092, Aug. 13, 1969] [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13092, Aug. 13, MISCELLANEOUS 1969; Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–29, 49 FR 6846, Feb. 23, 1984; Amdt. § 23.871 Leveling means. 23–34, 52 FR 1832, Jan. 15, 1987; Amdt. 23–34, 52 FR 34745, Sept. 14, 1987; Amdt. 23–43, 58 FR There must be means for determining 18970, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, when the airplane is in a level position Feb. 9, 1996; Amdt. 23–53, 63 FR 14797, Mar. 26, on the ground. 1998] [Amdt. 23–7, 34 FR 13092, Aug. 13, 1969] § 23.903 Engines. (a) Engine type certificate. (1) Each en- gine must have a type certificate and must meet the applicable requirements of part 34 of this chapter.

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(2) Each turbine engine and its in- chanical damage to the engine or air- stallation must comply with one of the plane, as a result of starting the engine following: in any conditions in which starting is (i) Sections 33.76, 33.77 and 33.78 of to be permitted, is reduced to a min- this chapter in effect on December 13, imum. Any techniques and associated 2000, or as subsequently amended; or limitations for engine starting must be (ii) Sections 33.77 and 33.78 of this established and included in the Air- chapter in effect on April 30, 1998, or as plane Flight Manual, approved manual subsequently amended before Decem- material, or applicable operating plac- ber 13, 2000; or ards. Means must be provided for— (iii) Section 33.77 of this chapter in (i) Restarting any engine of a multi- effect on October 31, 1974, or as subse- engine airplane in flight, and quently amended before April 30, 1998, (ii) Stopping any engine in flight, unless that engine’s foreign object in- after engine failure, if continued en- gestion service history has resulted in gine rotation would cause a hazard to an unsafe condition; or the airplane. (iv) Be shown to have a foreign object (2) In addition, for commuter cat- ingestion service history in similar in- egory airplanes, the following apply: stallation locations which has not re- (i) Each component of the stopping sulted in any unsafe condition. system on the engine side of the fire- NOTE: § 33.77 of this chapter in effect on Oc- wall that might be exposed to fire must tober 31, 1974, was published in 14 CFR parts be at least fire resistant. 1 to 59, Revised as of January 1, 1975. See 39 (ii) If hydraulic propeller feathering FR 35467, October 1, 1974. systems are used for this purpose, the (b) Turbine engine installations. For feathering lines must be at least fire turbine engine installations— resistant under the operating condi- (1) Design precautions must be taken tions that may be expected to exist to minimize the hazards to the airplane during feathering. in the event of an engine rotor failure (e) Starting and stopping (turbine en- or of a fire originating inside the en- gine). Turbine engine installations gine which burns through the engine must comply with the following: case. (1) The design of the installation (2) The powerplant systems associ- must be such that risk of fire or me- ated with engine control devices, sys- chanical damage to the engine or the tems, and instrumentation must be de- airplane, as a result of starting the en- signed to give reasonable assurance gine in any conditions in which start- that those operating limitations that ing is to be permitted, is reduced to a adversely affect turbine rotor struc- minimum. Any techniques and associ- tural integrity will not be exceeded in ated limitations must be established service. and included in the Airplane Flight (c) Engine isolation. The powerplants Manual, approved manual material, or must be arranged and isolated from applicable operating placards. each other to allow operation, in at (2) There must be means for stopping least one configuration, so that the combustion within any engine and for failure or malfunction of any engine, or stopping the rotation of any engine if the failure or malfunction (including continued rotation would cause a haz- destruction by fire in the engine com- ard to the airplane. Each component of partment) of any system that can af- the engine stopping system located in fect an engine (other than a fuel tank any fire zone must be fire resistant. If if only one fuel tank is installed), will hydraulic propeller feathering systems not: are used for stopping the engine, the (1) Prevent the continued safe oper- hydraulic feathering lines or hoses ation of the remaining engines; or must be fire resistant. (2) Require immediate action by any (3) It must be possible to restart an crewmember for continued safe oper- engine in flight. Any techniques and ation of the remaining engines. associated limitations must be estab- (d) Starting and stopping (piston en- lished and included in the Airplane gine). (1) The design of the installation Flight Manual, approved manual mate- must be such that risk of fire or me- rial, or applicable operating placards.

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(4) It must be demonstrated in flight shown that any ice shed into the pro- that when restarting engines following peller disc will not create a hazardous a false start, all fuel or vapor is dis- condition. charged in such a way that it does not (f) Each pusher propeller must be constitute a fire hazard. marked so that the disc is conspicuous (f) Restart envelope. An altitude and under normal daylight ground condi- airspeed envelope must be established tions. for the airplane for in-flight engine re- (g) If the engine exhaust gases are starting and each installed engine discharged into the pusher propeller must have a restart capability within disc, it must be shown by tests, or that envelope. analysis supported by tests, that the (g) Restart capability. For turbine en- propeller is capable of continuous safe gine powered airplanes, if the min- operation. imum windmilling speed of the en- (h) All engine cowling, access doors, gines, following the in-flight shutdown and other removable items must be de- of all engines, is insufficient to provide signed to ensure that they will not sep- the necessary electrical power for en- arate from the airplane and contact gine ignition, a power source inde- the pusher propeller. pendent of the engine-driven electrical [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as power generating system must be pro- amended by Amdt. 23–26, 45 FR 60171, Sept. vided to permit in-flight engine igni- 11, 1980; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; tion for restarting. Amdt. 23–43, 58 FR 18970, Apr. 9, 1993; Amdt. No. 23–59, 73 FR 63345, Oct. 24, 2008] [Amdt. 23–14, 38 FR 31822, Nov. 19, 1973]

EDITORIAL NOTE: For FEDERAL REGISTER ci- § 23.907 Propeller vibration and fa- tations affecting § 23.903, see the List of CFR tigue. Sections Affected, which appears in the This section does not apply to fixed- Finding Aids section of the printed volume and at www.fdsys.gov. pitch wood propellers of conventional design. § 23.904 Automatic power reserve sys- (a) The applicant must determine the tem. magnitude of the propeller vibration If installed, an automatic power re- stresses or loads, including any stress serve (APR) system that automatically peaks and resonant conditions, advances the power or thrust on the op- throughout the operational envelope of erating engine(s), when any engine the airplane by either: fails during takeoff, must comply with (1) Measurement of stresses or loads appendix H of this part. through direct testing or analysis based on direct testing of the propeller [Doc. No. 26344, 58 FR 18970, Apr. 9, 1993] on the airplane and engine installation for which approval is sought; or § 23.905 Propellers. (2) Comparison of the propeller to (a) Each propeller must have a type similar propellers installed on similar certificate. airplane installations for which these (b) Engine power and propeller shaft measurements have been made. rotational speed may not exceed the (b) The applicant must demonstrate limits for which the propeller is certifi- by tests, analysis based on tests, or cated. previous experience on similar designs (c) Each featherable propeller must that the propeller does not experience have a means to unfeather it in flight. harmful effects of flutter throughout (d) The propeller blade pitch control the operational envelope of the air- system must meet the requirements of plane. §§ 35.21, 35.23, 35.42 and 35.43 of this (c) The applicant must perform an chapter. evaluation of the propeller to show (e) All areas of the airplane forward that failure due to fatigue will be of the pusher propeller that are likely avoided throughout the operational life to accumulate and shed ice into the of the propeller using the fatigue and propeller disc during any operating structural data obtained in accordance condition must be suitably protected with part 35 of this chapter and the vi- to prevent ice formation, or it must be bration data obtained from compliance

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with paragraph (a) of this section. For no hazard to the airplane under all op- the purpose of this paragraph, the pro- erating conditions. peller includes the hub, blades, blade (e) Engine power, cooling character- retention component and any other istics, operating limits, and procedures propeller component whose failure due affected by the turbocharger system in- to fatigue could be catastrophic to the stallations must be evaluated. Turbo- airplane. This evaluation must include: charger operating procedures and limi- (1) The intended loading spectra in- tations must be included in the Air- cluding all reasonably foreseeable pro- plane Flight Manual in accordance peller vibration and cyclic load pat- with § 23.1581. terns, identified emergency conditions, [Amdt. 23–7, 34 FR 13092, Aug. 13, 1969, as allowable overspeeds and overtorques, amended by Amdt. 23–43, 58 FR 18970, Apr. 9, and the effects of temperatures and hu- 1993] midity expected in service. (2) The effects of airplane and pro- § 23.925 Propeller clearance. peller operating and airworthiness lim- Unless smaller clearances are sub- itations. stantiated, propeller clearances, with the airplane at the most adverse com- [Amdt. No. 23–59, 73 FR 63345, Oct. 24, 2008] bination of weight and center of grav- § 23.909 Turbocharger systems. ity, and with the propeller in the most adverse pitch position, may not be less (a) Each turbocharger must be ap- than the following: proved under the engine type certifi- (a) Ground clearance. There must be a cate or it must be shown that the tur- clearance of at least seven inches (for bocharger system, while in its normal each airplane with nose wheel landing engine installation and operating in gear) or nine inches (for each airplane the engine environment— with tail wheel landing gear) between (1) Can withstand, without defect, an each propeller and the ground with the endurance test of 150 hours that meets landing gear statically deflected and in the applicable requirements of § 33.49 of the level, normal takeoff, or taxing at- this subchapter; and titude, whichever is most critical. In (2) Will have no adverse effect upon addition, for each airplane with con- the engine. ventional landing gear struts using (b) Control system malfunctions, vi- fluid or mechanical means for absorb- brations, and abnormal speeds and ing landing shocks, there must be posi- temperatures expected in service may tive clearance between the propeller not damage the turbocharger com- and the ground in the level takeoff at- pressor or turbine. titude with the critical tire completely (c) Each turbocharger case must be deflated and the corresponding landing able to contain fragments of a com- gear strut bottomed. Positive clear- pressor or turbine that fails at the ance for airplanes using leaf spring highest speed that is obtainable with struts is shown with a deflection cor- normal speed control devices inoper- responding to 1.5g. ative. (b) Aft-mounted propellers. In addition (d) Each intercooler installation, to the clearances specified in para- where provided, must comply with the graph (a) of this section, an airplane following— with an aft mounted propeller must be (1) The mounting provisions of the designed such that the propeller will intercooler must be designed to with- not contact the runway surface when stand the loads imposed on the system; the airplane is in the maximum pitch (2) It must be shown that, under the attitude attainable during normal installed vibration environment, the takeoffs and landings. intercooler will not fail in a manner al- (c) Water clearance. There must be a lowing portions of the intercooler to be clearance of at least 18 inches between ingested by the engine; and each propeller and the water, unless (3) Airflow through the intercooler compliance with § 23.239 can be shown must not discharge directly on any air- with a lesser clearance. plane component (e.g., windshield) un- (d) Structural clearance. There must less such discharge is shown to cause be—

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(1) At least one inch radial clearance (3) Each system must have a means between the blade tips and the airplane to prevent the engine from producing structure, plus any additional radial more than idle thrust when the revers- clearance necessary to prevent harmful ing system malfunctions; except that it vibration; may produce any greater thrust that is (2) At least one-half inch longitudinal shown to allow directional control to clearance between the propeller blades be maintained, with aerodynamic or cuffs and stationary parts of the air- means alone, under the most critical plane; and reversing condition expected in oper- (3) Positive clearance between other ation. rotating parts of the propeller or spin- (b) For propeller reversing systems. (1) ner and stationary parts of the air- Each system must be designed so that plane. no single failure, likely combination of failures or malfunction of the system [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18971, Apr. 9, will result in unwanted reverse thrust 1993; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996; under any operating condition. Failure Amdt. 23–48, 61 FR 5148, Feb. 9, 1996] of structural elements need not be con- sidered if the probability of this type of § 23.929 Engine installation ice protec- failure is extremely remote. tion. (2) Compliance with paragraph (b)(1) Propellers (except wooden propellers) of this section must be shown by fail- and other components of complete en- ure analysis, or testing, or both, for gine installations must be protected propeller systems that allow the pro- against the accumulation of ice as nec- peller blades to move from the flight essary to enable satisfactory func- low-pitch position to a position that is tioning without appreciable loss of substantially less than the normal thrust when operated in the icing con- flight, low-pitch position. The analysis ditions for which certification is re- may include or be supported by the quested. analysis made to show compliance with § 35.21 for the type certification of the [Amdt. 23–14, 33 FR 31822, Nov. 19, 1973, as propeller and associated installation amended by Amdt. 23–51, 61 FR 5136, Feb. 9, 1996] components. Credit will be given for pertinent analysis and testing com- § 23.933 Reversing systems. pleted by the engine and propeller manufacturers. (a) For turbojet and turbofan reversing systems. (1) Each system intended for [Doc. No. 26344, 58 FR 18971, Apr. 9, 1993, as ground operation only must be de- amended by Amdt. 23–51, 61 FR 5136, Feb. 9, signed so that, during any reversal in 1996] flight, the engine will produce no more § 23.934 Turbojet and turbofan engine than flight idle thrust. In addition, it thrust reverser systems tests. must be shown by analysis or test, or both, that— Thrust reverser systems of turbojet (i) Each operable reverser can be re- or turbofan engines must meet the re- stored to the forward thrust position; quirements of § 33.97 of this chapter or or it must be demonstrated by tests that (ii) The airplane is capable of contin- engine operation and vibratory levels ued safe flight and landing under any are not affected. possible position of the thrust reverser. [Doc. No. 26344, 58 FR 18971, Apr. 9, 1993] (2) Each system intended for in-flight use must be designed so that no unsafe § 23.937 Turbopropeller-drag limiting condition will result during normal op- systems. eration of the system, or from any fail- (a) Turbopropeller-powered airplane ure, or likely combination of failures, propeller-drag limiting systems must of the reversing system under any op- be designed so that no single failure or erating condition including ground op- malfunction of any of the systems dur- eration. Failure of structural elements ing normal or emergency operation re- need not be considered if the prob- sults in propeller drag in excess of that ability of this type of failure is ex- for which the airplane was designed tremely remote. under the structural requirements of

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this part. Failure of structural ele- value and duration of the acceleration ments of the drag limiting systems expected in service. need not be considered if the prob- [Amdt. 23–18, 42 FR 15041, Mar. 17, 1977, as ability of this kind of failure is ex- amended by Amdt. 23–43, 58 FR 18971, Apr. 9, tremely remote. 1993] (b) As used in this section, drag lim- iting systems include manual or auto- FUEL SYSTEM matic devices that, when actuated § 23.951 General. after engine power loss, can move the (a) Each fuel system must be con- propeller blades toward the feather po- structed and arranged to ensure fuel sition to reduce windmilling drag to a flow at a rate and pressure established safe level. for proper engine and auxiliary power [Amdt. 23–7, 34 FR 13093, Aug. 13, 1969, as unit functioning under each likely op- amended by Amdt. 23–43, 58 FR 18971, Apr. 9, erating condition, including any ma- 1993] neuver for which certification is re- quested and during which the engine or § 23.939 Powerplant operating charac- auxiliary power unit is permitted to be teristics. in operation. (a) Turbine engine powerplant oper- (b) Each fuel system must be ar- ating characteristics must be inves- ranged so that— tigated in flight to determine that no (1) No fuel pump can draw fuel from more than one tank at a time; or adverse characteristics (such as stall, (2) There are means to prevent intro- surge, or flameout) are present, to a ducing air into the system. hazardous degree, during normal and (c) Each fuel system for a turbine en- emergency operation within the range gine must be capable of sustained oper- of operating limitations of the airplane ation throughout its flow and pressure and of the engine. range with fuel initially saturated with (b) Turbocharged reciprocating en- water at 80 °F and having 0.75cc of free gine operating characteristics must be water per gallon added and cooled to investigated in flight to assure that no the most critical condition for icing adverse characteristics, as a result of likely to be encountered in operation. an inadvertent overboost, surge, flood- (d) Each fuel system for a turbine en- ing, or vapor lock, are present during gine powered airplane must meet the normal or emergency operation of the applicable fuel venting requirements of engine(s) throughout the range of oper- part 34 of this chapter. ating limitations of both airplane and [Amdt. 23–15, 39 FR 35459, Oct. 1, 1974, as engine. amended by Amdt. 23–40, 55 FR 32861, Aug. 10, (c) For turbine engines, the air inlet 1990; Amdt. 23–43, 58 FR 18971, Apr. 9, 1993] system must not, as a result of airflow distortion during normal operation, § 23.953 Fuel system independence. cause vibration harmful to the engine. (a) Each fuel system for a multien- gine airplane must be arranged so that, [Amdt. 23–7, 34 FR 13093 Aug. 13, 1969, as in at least one system configuration, amended by Amdt. 23–14, 38 FR 31823, Nov. 19, the failure of any one component 1973; Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–42, 56 FR 354, Jan. 3, 1991] (other than a fuel tank) will not result in the loss of power of more than one § 23.943 Negative acceleration. engine or require immediate action by the pilot to prevent the loss of power of No hazardous malfunction of an en- more than one engine. gine, an auxiliary power unit approved (b) If a single fuel tank (or series of for use in flight, or any component or fuel tanks interconnected to function system associated with the powerplant as a single fuel tank) is used on a mul- or auxiliary power unit may occur tiengine airplane, the following must when the airplane is operated at the be provided: negative accelerations within the (1) Independent tank outlets for each flight envelopes prescribed in § 23.333. engine, each incorporating a shut-off This must be shown for the greatest valve at the tank. This shutoff valve

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may also serve as the fire wall shutoff failure mode that would restrict fuel valve required if the line between the flow below the level required for this valve and the engine compartment does fuel demonstration. not contain more than one quart of (4) The fuel flow must include that fuel (or any greater amount shown to flow necessary for vapor return flow, be safe) that can escape into the engine jet pump drive flow, and for all other compartment. purposes for which fuel is used. (2) At least two vents arranged to (b) Gravity systems. The fuel flow rate minimize the probability of both vents for gravity systems (main and reserve becoming obstructed simultaneously. supply) must be 150 percent of the (3) Filler caps designed to minimize takeoff fuel consumption of the engine. the probability of incorrect installa- (c) Pump systems. The fuel flow rate tion or inflight loss. for each pump system (main and re- (4) A fuel system in which those parts serve supply) for each reciprocating en- of the system from each tank outlet to gine must be 125 percent of the fuel any engine are independent of each flow required by the engine at the max- part of the system supplying fuel to imum takeoff power approved under any other engine. this part. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (1) This flow rate is required for each amended by Amdt. 23–7, 34 FR 13093 Aug. 13, main pump and each emergency pump, 1969; Amdt. 23–43, 58 FR 18971, Apr. 9, 1993] and must be available when the pump is operating as it would during takeoff; § 23.954 Fuel system lightning protec- (2) For each hand-operated pump, tion. this rate must occur at not more than The fuel system must be designed 60 complete cycles (120 single strokes) and arranged to prevent the ignition of per minute. fuel vapor within the system by— (3) The fuel pressure, with main and (a) Direct lightning strikes to areas emergency pumps operating simulta- having a high probability of stroke at- neously, must not exceed the fuel inlet tachment; pressure limits of the engine unless it (b) Swept lightning strokes on areas can be shown that no adverse effect oc- where swept strokes are highly prob- curs. able; and (d) Auxiliary fuel systems and fuel (c) Corona or streamering at fuel transfer systems. Paragraphs (b), (c), and vent outlets. (f) of this section apply to each auxil- [Amdt. 23–7, 34 FR 13093, Aug. 13, 1969] iary and transfer system, except that— (1) The required fuel flow rate must § 23.955 Fuel flow. be established upon the basis of max- (a) General. The ability of the fuel imum continuous power and engine ro- system to provide fuel at the rates tational speed, instead of takeoff power specified in this section and at a pres- and fuel consumption; and sure sufficient for proper engine oper- (2) If there is a placard providing op- ation must be shown in the attitude erating instructions, a lesser flow rate that is most critical with respect to may be used for transferring fuel from fuel feed and quantity of unusable fuel. any auxiliary tank into a larger main These conditions may be simulated in a tank. This lesser flow rate must be ade- suitable mockup. In addition— quate to maintain engine maximum (1) The quantity of fuel in the tank continuous power but the flow rate may not exceed the amount established must not overfill the main tank at as the unusable fuel supply for that lower engine powers. tank under § 23.959(a) plus that quan- (e) Multiple fuel tanks. For recipro- tity necessary to show compliance with cating engines that are supplied with this section. fuel from more than one tank, if engine (2) If there is a fuel flowmeter, it power loss becomes apparent due to must be blocked during the flow test fuel depletion from the tank selected, and the fuel must flow through the it must be possible after switching to meter or its bypass. any full tank, in level flight, to obtain (3) If there is a flowmeter without a 75 percent maximum continuous power bypass, it must not have any probable on that engine in not more than—

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(1) 10 seconds for naturally aspirated § 23.957 Flow between interconnected single-engine airplanes; tanks. (2) 20 seconds for turbocharged sin- (a) It must be impossible, in a grav- gle-engine airplanes, provided that 75 ity feed system with interconnected percent maximum continuous natu- tank outlets, for enough fuel to flow rally aspirated power is regained with- between the tanks to cause an overflow in 10 seconds; or of fuel from any tank vent under the (3) 20 seconds for multiengine air- conditions in § 23.959, except that full planes. tanks must be used. (f) Turbine engine fuel systems. Each (b) If fuel can be pumped from one turbine engine fuel system must pro- tank to another in flight, the fuel tank vide at least 100 percent of the fuel flow vents and the fuel transfer system required by the engine under each in- must be designed so that no structural tended operation condition and maneu- damage to any airplane component can ver. The conditions may be simulated occur because of overfilling of any in a suitable mockup. This flow must— tank. (1) Be shown with the airplane in the [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as most adverse fuel feed condition (with amended by Amdt. 23–43, 58 FR 18972, Apr. 9, respect to altitudes, attitudes, and 1993] other conditions) that is expected in § 23.959 Unusable fuel supply. operation; and (2) For multiengine airplanes, not- (a) The unusable fuel supply for each withstanding the lower flow rate al- tank must be established as not less lowed by paragraph (d) of this section, than that quantity at which the first be automatically uninterrupted with evidence of malfunctioning occurs respect to any engine until all the fuel under the most adverse fuel feed condi- scheduled for use by that engine has tion occurring under each intended op- eration and flight maneuver involving been consumed. In addition— that tank. Fuel system component fail- (i) For the purposes of this section, ures need not be considered. ‘‘fuel scheduled for use by that engine’’ (b) The effect on the usable fuel means all fuel in any tank intended for quantity as a result of a failure of any use by a specific engine. pump shall be determined. (ii) The fuel system design must clearly indicate the engine for which [Amdt. 23–7, 34 FR 13093, Aug. 13, 1969, as fuel in any tank is scheduled. amended by Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996] (iii) Compliance with this paragraph must require no pilot action after com- § 23.961 Fuel system hot weather oper- pletion of the engine starting phase of ation. operations. Each fuel system must be free from (3) For single-engine airplanes, re- vapor lock when using fuel at its crit- quire no pilot action after completion ical temperature, with respect to vapor of the engine starting phase of oper- formation, when operating the airplane ations unless means are provided that in all critical operating and environ- unmistakenly alert the pilot to take mental conditions for which approval any needed action at least five minutes is requested. For turbine fuel, the ini- prior to the needed action; such pilot tial temperature must be 110 °F, ¥0°, action must not cause any change in +5 °F or the maximum outside air tem- engine operation; and such pilot action perature for which approval is re- must not distract pilot attention from quested, whichever is more critical. essential flight duties during any phase [Doc. No. 26344, 58 FR 18972, Apr. 9, 1993; 58 of operations for which the airplane is FR 27060, May 6, 1993] approved. § 23.963 Fuel tanks: General. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13093, Aug. 13, (a) Each fuel tank must be able to 1969; Amdt. 23–43, 58 FR 18971, Apr. 9, 1993; withstand, without failure, the vibra- Amdt. 23–51, 61 FR 5136, Feb. 9, 1996] tion, inertia, fluid, and structural loads

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that it may be subjected to in oper- (1) Each complete tank assembly and ation. its support must be vibration tested (b) Each flexible fuel tank liner must while mounted to simulate the actual be shown to be suitable for the par- installation. ticular application. (2) Except as specified in paragraph (c) Each integral fuel tank must have (b)(4) of this section, the tank assembly adequate facilities for interior inspec- must be vibrated for 25 hours at a total tion and repair. displacement of not less than 1⁄32 of an (d) The total usable capacity of the inch (unless another displacement is fuel tanks must be enough for at least substantiated) while 2⁄3 filled with one-half hour of operation at maximum water or other suitable test fluid. continuous power. (3) The test frequency of vibration (e) Each fuel quantity indicator must must be as follows: be adjusted, as specified in § 23.1337(b), (i) If no frequency of vibration result- to account for the unusable fuel supply ing from any rpm within the normal determined under § 23.959(a). operating range of engine or propeller [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 speeds is critical, the test frequency of FR 258, Jan. 9, 1965, as amended by Amdt 23– vibration is: 34, 52 FR 1832, Jan. 15, 1987; Amdt. 23–43, 58 (A) The number of cycles per minute FR 18972, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, obtained by multiplying the maximum Feb. 9, 1996] continuous propeller speed in rpm by § 23.965 Fuel tank tests. 0.9 for propeller-driven airplanes, and (B) For non-propeller driven air- (a) Each fuel tank must be able to planes the test frequency of vibration withstand the following pressures with- is 2,000 cycles per minute. out failure or leakage: (1) For each conventional metal tank (ii) If only one frequency of vibration and nonmetallic tank with walls not resulting from any rpm within the nor- supported by the airplane structure, a mal operating range of engine or pro- pressure of 3.5 p.s.i., or that pressure peller speeds is critical, that frequency developed during maximum ultimate of vibration must be the test fre- acceleration with a full tank, which- quency. ever is greater. (iii) If more than one frequency of vi- (2) For each integral tank, the pres- bration resulting from any rpm within sure developed during the maximum the normal operating range of engine limit acceleration of the airplane with or propeller speeds is critical, the most a full tank, with simultaneous applica- critical of these frequencies must be tion of the critical limit structural the test frequency. loads. (4) Under paragraph (b)(3) (ii) and (iii) (3) For each nonmetallic tank with of this section, the time of test must be walls supported by the airplane struc- adjusted to accomplish the same num- ture and constructed in an acceptable ber of vibration cycles that would be manner using acceptable basic tank accomplished in 25 hours at the fre- material, and with actual or simulated quency specified in paragraph (b)(3)(i) support conditions, a pressure of 2 p.s.i. of this section. for the first tank of a specific design. (5) During the test, the tank assem- The supporting structure must be de- bly must be rocked at a rate of 16 to 20 signed for the critical loads occurring complete cycles per minute, through in the flight or landing strength condi- an angle of 15° on either side of the hor- tions combined with the fuel pressure izontal (30° total), about an axis par- loads resulting from the corresponding allel to the axis of the fuselage, for 25 accelerations. hours. (b) Each fuel tank with large, unsup- (c) Each integral tank using methods ported, or unstiffened flat sur- of construction and sealing not pre- faces,whose failure or deformation viously proven to be adequate by test could cause fuel leakage, must be able data or service experience must be able to withstand the following test without to withstand the vibration test speci- leakage, failure, or excessive deforma- fied in paragraphs (b)(1) through (4) of tion of the tank walls: this section.

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(d) Each tank with a nonmetallic lies immediately behind a major air liner must be subjected to the sloshing opening from the engine compartment test outlined in paragraph (b)(5) of this may act as the wall of an integral section, with the fuel at room tempera- tank. ture. In addition, a specimen liner of (d) Each fuel tank must be isolated the same basic construction as that to from personnel compartments by a be used in the airplane must, when in- fume-proof and fuel-proof enclosure stalled in a suitable test tank, with- that is vented and drained to the exte- stand the sloshing test with fuel at a rior of the airplane. The required en- ° temperature of 110 F. closure must sustain any personnel [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as compartment pressurization loads amended by Amdt. 23–43, 58 FR 18972, Apr. 9, without permanent deformation or fail- 1993; Amdt. 23–43, 61 FR 253, Jan. 4, 1996; ure under the conditions of §§ 23.365 and Amdt. 23–51, 61 FR 5136, Feb. 9, 1996] 23.843 of this part. A bladder-type fuel § 23.967 Fuel tank installation. cell, if used, must have a retaining shell at least equivalent to a metal fuel (a) Each fuel tank must be supported so that tank loads are not con- tank in structural integrity. centrated. In addition— (e) Fuel tanks must be designed, lo- (1) There must be pads, if necessary, cated, and installed so as to retain fuel: to prevent chafing between each tank (1) When subjected to the inertia and its supports; loads resulting from the ultimate stat- (2) Padding must be nonabsorbent or ic load factors prescribed in treated to prevent the absorption of § 23.561(b)(2) of this part; and fuel; (2) Under conditions likely to occur (3) If a flexible tank liner is used, it when the airplane lands on a paved must be supported so that it is not re- runway at a normal landing speed quired to withstand fluid loads; under each of the following conditions: (4) Interior surfaces adjacent to the (i) The airplane in a normal landing liner must be smooth and free from attitude and its landing gear retracted. projections that could cause wear, un- (ii) The most critical landing gear leg less— collapsed and the other landing gear (i) Provisions are made for protection of the liner at those points; or legs extended. (ii) The construction of the liner In showing compliance with paragraph itself provides such protection; and (e)(2) of this section, the tearing away (5) A positive pressure must be main- of an engine mount must be considered tained within the vapor space of each unless all the engines are installed bladder cell under any condition of op- above the wing or on the tail or fuse- eration, except for a particular condi- lage of the airplane. tion for which it is shown that a zero or negative pressure will not cause the [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13903, Aug. 13, bladder cell to collapse; and 1969; Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; (6) Syphoning of fuel (other than Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. minor spillage) or collapse of bladder 23–26, 45 FR 60171, Sept. 11, 1980; Amdt. 23–36, fuel cells may not result from improper 53 FR 30815, Aug. 15, 1988; Amdt. 23–43, 58 FR securing or loss of the fuel filler cap. 18972, Apr. 9, 1993] (b) Each tank compartment must be ventilated and drained to prevent the § 23.969 Fuel tank expansion space. accumulation of flammable fluids or Each fuel tank must have an expan- vapors. Each compartment adjacent to sion space of not less than two percent a tank that is an integral part of the airplane structure must also be venti- of the tank capacity, unless the tank lated and drained. vent discharges clear of the airplane (c) No fuel tank may be on the engine (in which case no expansion space is re- side of the firewall. There must be at quired). It must be impossible to fill least one-half inch of clearance be- the expansion space inadvertently with tween the fuel tank and the firewall. the airplane in the normal ground atti- No part of the engine nacelle skin that tude.

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§ 23.971 Fuel tank sump. filler opening must be no smaller than 2.95 inches. (a) Each fuel tank must have a drain- able sump with an effective capacity, [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 in the normal ground and flight atti- FR 258, Jan. 9, 1965, as amended by Amdt. 23– tudes, of 0.25 percent of the tank capac- 18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–43, 58 FR 18972, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, ity, or 1⁄16 gallon, whichever is greater. Feb. 9, 1996] (b) Each fuel tank must allow drain- age of any hazardous quantity of water § 23.975 Fuel tank vents and carbu- from any part of the tank to its sump retor vapor vents. with the airplane in the normal ground (a) Each fuel tank must be vented attitude. from the top part of the expansion (c) Each reciprocating engine fuel space. In addition— system must have a sediment bowl or (1) Each vent outlet must be located chamber that is accessible for drain- and constructed in a manner that mini- age; has a capacity of 1 ounce for every mizes the possibility of its being ob- 20 gallons of fuel tank capacity; and structed by ice or other foreign matter; each fuel tank outlet is located so that, (2) Each vent must be constructed to in the normal flight attitude, water prevent siphoning of fuel during nor- will drain from all parts of the tank ex- mal operation; cept the sump to the sediment bowl or (3) The venting capacity must allow chamber. the rapid relief of excessive differences (d) Each sump, sediment bowl, and of pressure between the interior and sediment chamber drain required by exterior of the tank; paragraphs (a), (b), and (c) of this sec- (4) Airspaces of tanks with inter- tion must comply with the drain provi- connected outlets must be inter- sions of § 23.999(b)(1) and (b)(2). connected; [Doc. No. 26344, 58 FR 18972, Apr. 9, 1993; 58 (5) There may be no point in any vent FR 27060, May 6, 1993] line where moisture can accumulate with the airplane in either the ground § 23.973 Fuel tank filler connection. or level flight attitudes, unless drain- age is provided. Any drain valve in- (a) Each fuel tank filler connection stalled must be accessible for drainage; must be marked as prescribed in (6) No vent may terminate at a point § 23.1557(c). where the discharge of fuel from the (b) Spilled fuel must be prevented vent outlet will constitute a fire haz- from entering the fuel tank compart- ard or from which fumes may enter ment or any part of the airplane other personnel compartments; and than the tank itself. (7) Vents must be arranged to pre- (c) Each filler cap must provide a vent the loss of fuel, except fuel dis- fuel-tight seal for the main filler open- charged because of thermal expansion, ing. However, there may be small open- when the airplane is parked in any di- ings in the fuel tank cap for venting rection on a ramp having a one-percent purposes or for the purpose of allowing slope. passage of a fuel gauge through the cap (b) Each carburetor with vapor elimi- provided such openings comply with nation connections and each fuel injec- the requirements of § 23.975(a). tion engine employing vapor return (d) Each fuel filling point, except provisions must have a separate vent pressure fueling connection points, line to lead vapors back to the top of must have a provision for electrically one of the fuel tanks. If there is more bonding the airplane to ground fueling than one tank and it is necessary to equipment. use these tanks in a definite sequence (e) For airplanes with engines requir- for any reason, the vapor vent line ing gasoline as the only permissible must lead back to the fuel tank to be fuel, the inside diameter of the fuel used first, unless the relative capac- filler opening must be no larger than ities of the tanks are such that return 2.36 inches. to another tank is preferable. (f) For airplanes with turbine en- (c) For acrobatic category airplanes, gines, the inside diameter of the fuel excessive loss of fuel during acrobatic

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maneuvers, including short periods of (c) A means must be provided to pre- inverted flight, must be prevented. It vent damage to the fuel system in the must be impossible for fuel to siphon event of failure of the automatic shut- from the vent when normal flight has off means prescribed in paragraph (b) been resumed after any acrobatic ma- of this section. neuver for which certification is re- (d) All parts of the fuel system up to quested. the tank which are subjected to fueling pressures must have a proof pressure of [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23– 1.33 times, and an ultimate pressure of 18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–29, 49 at least 2.0 times, the surge pressure FR 6847, Feb. 23, 1984; Amdt. 23–43, 58 FR likely to occur during fueling. 18973, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996] [Amdt. 23–14, 38 FR 31823, Nov. 19, 1973, as amended by Amdt. 23–51, 61 FR 5137, Feb. 9, § 23.977 Fuel tank outlet. 1996] (a) There must be a fuel strainer for FUEL SYSTEM COMPONENTS the fuel tank outlet or for the booster pump. This strainer must— § 23.991 Fuel pumps. (1) For reciprocating engine powered (a) Main pumps. For main pumps, the airplanes, have 8 to 16 meshes per inch; following apply: and (1) For reciprocating engine installa- (2) For turbine engine powered air- tions having fuel pumps to supply fuel planes, prevent the passage of any ob- to the engine, at least one pump for ject that could restrict fuel flow or each engine must be directly driven by damage any fuel system component. the engine and must meet § 23.955. This (b) The clear area of each fuel tank pump is a main pump. outlet strainer must be at least five (2) For turbine engine installations, times the area of the outlet line. each fuel pump required for proper en- (c) The diameter of each strainer gine operation, or required to meet the must be at least that of the fuel tank fuel system requirements of this sub- outlet. part (other than those in paragraph (b) (d) Each strainer must be accessible of this section), is a main pump. In ad- for inspection and cleaning. dition— [Amdt. 23–17, 41 FR 55465, Dec. 20, 1976, as (i) There must be at least one main amended by Amdt. 23–43, 58 FR 18973, Apr. 9, pump for each turbine engine; 1993] (ii) The power supply for the main pump for each engine must be inde- § 23.979 Pressure fueling systems. pendent of the power supply for each For pressure fueling systems, the fol- main pump for any other engine; and lowing apply: (iii) For each main pump, provision (a) Each pressure fueling system fuel must be made to allow the bypass of manifold connection must have means each positive displacement fuel pump to prevent the escape of hazardous other than a fuel injection pump ap- quantities of fuel from the system if proved as part of the engine. the fuel entry valve fails. (b) Emergency pumps. There must be (b) An automatic shutoff means must an emergency pump immediately avail- be provided to prevent the quantity of able to supply fuel to the engine if any fuel in each tank from exceeding the main pump (other than a fuel injection maximum quantity approved for that pump approved as part of an engine) tank. This means must— fails. The power supply for each emer- (1) Allow checking for proper shutoff gency pump must be independent of the operation before each fueling of the power supply for each corresponding tank; and main pump. (2) For commuter category airplanes, (c) Warning means. If both the main indicate at each fueling station, a fail- pump and emergency pump operate ure of the shutoff means to stop the continuously, there must be a means to fuel flow at the maximum quantity ap- indicate to the appropriate flight crew- proved for that tank. members a malfunction of either pump.

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(d) Operation of any fuel pump may (2) Allow appropriate flight crew not affect engine operation so as to members to reopen each valve rapidly create a hazard, regardless of the en- after it has been closed. gine power or thrust setting or the (c) Each valve and fuel system con- functional status of any other fuel trol must be supported so that loads re- pump. sulting from its operation or from ac- celerated flight conditions are not [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as transmitted to the lines connected to amended by Amdt. 23–7, 34 FR 13093, Aug. 13, the valve. 1969; Amdt. 23–26, 45 FR 60171, Sept. 11, 1980; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993] (d) Each valve and fuel system con- trol must be installed so that gravity § 23.993 Fuel system lines and fittings. and vibration will not affect the se- lected position. (a) Each fuel line must be installed (e) Each fuel valve handle and its and supported to prevent excessive vi- connections to the valve mechanism bration and to withstand loads due to must have design features that mini- fuel pressure and accelerated flight mize the possibility of incorrect instal- conditions. lation. (b) Each fuel line connected to com- (f) Each check valve must be con- ponents of the airplane between which structed, or otherwise incorporate pro- relative motion could exist must have visions, to preclude incorrect assembly provisions for flexibility. or connection of the valve. (c) Each flexible connection in fuel (g) Fuel tank selector valves must— lines that may be under pressure and (1) Require a separate and distinct subjected to axial loading must use action to place the selector in the flexible hose assemblies. ‘‘OFF’’ position; and (d) Each flexible hose must be shown (2) Have the tank selector positions to be suitable for the particular appli- located in such a manner that it is im- cation. possible for the selector to pass (e) No flexible hose that might be ad- through the ‘‘OFF’’ position when versely affected by exposure to high changing from one tank to another. temperatures may be used where exces- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as sive temperatures will exist during op- amended by Amdt. 23–14, 38 FR 31823, Nov. 19, eration or after engine shutdown. 1973; Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as 23–29, 49 FR 6847, Feb. 23, 1984] amended by Amdt. 23–43, 58 FR 18973, Apr. 9, 1993] § 23.997 Fuel strainer or filter. There must be a fuel strainer or filter § 23.994 Fuel system components. between the fuel tank outlet and the Fuel system components in an engine inlet of either the fuel metering device nacelle or in the fuselage must be pro- or an engine driven positive displace- tected from damage which could result ment pump, whichever is nearer the in spillage of enough fuel to constitute fuel tank outlet. This fuel strainer or a fire hazard as a result of a wheels-up filter must— landing on a paved runway. (a) Be accessible for draining and cleaning and must incorporate a screen [Amdt. 23–29, 49 FR 6847, Feb. 23, 1984] or element which is easily removable; § 23.995 Fuel valves and controls. (b) Have a sediment trap and drain except that it need not have a drain if (a) There must be a means to allow the strainer or filter is easily remov- appropriate flight crew members to able for drain purposes; rapidly shut off, in flight, the fuel to (c) Be mounted so that its weight is each engine individually. not supported by the connecting lines (b) No shutoff valve may be on the or by the inlet or outlet connections of engine side of any firewall. In addition, the strainer or filter itself, unless ade- there must be means to— quate strength margins under all load- (1) Guard against inadvertent oper- ing conditions are provided in the lines ation of each shutoff valve; and and connections; and

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(d) Have the capacity (with respect to weight per minute, except that the operating limitations established for time required to jettison the fuel need the engine) to ensure that engine fuel not be less than 10 minutes. system functioning is not impaired, (b) Fuel jettisoning must be dem- with the fuel contaminated to a degree onstrated at maximum weight with (with respect to particle size and den- flaps and landing gear up and in— sity) that is greater than that estab- (1) A power-off glide at 1.4 VS1; lished for the engine during its type (2) A climb, at the speed at which the certification. one-engine-inoperative enroute climb (e) In addition, for commuter cat- data have been established in accord- egory airplanes, unless means are pro- ance with § 23.69(b), with the critical vided in the fuel system to prevent the engine inoperative and the remaining accumulation of ice on the filter, a engines at maximum continuous means must be provided to automati- power; and cally maintain the fuel flow if ice clog- (3) Level flight at 1.4 VS1, if the re- ging of the filter occurs. sults of the tests in the conditions [Amdt. 23–15, 39 FR 35459, Oct. 1, 1974, as specified in paragraphs (b)(1) and (2) of amended by Amdt. 23–29, 49 FR 6847, Feb. 23, this section show that this condition 1984; Amdt. 23–34, 52 FR 1832, Jan. 15, 1987; could be critical. Amdt. 23–43, 58 FR 18973, Apr. 9, 1993] (c) During the flight tests prescribed in paragraph (b) of this section, it must § 23.999 Fuel system drains. be shown that— (a) There must be at least one drain (1) The fuel jettisoning system and to allow safe drainage of the entire fuel its operation are free from fire hazard; system with the airplane in its normal (2) The fuel discharges clear of any ground attitude. part of the airplane; (b) Each drain required by paragraph (3) Fuel or fumes do not enter any (a) of this section and § 23.971 must— parts of the airplane; and (1) Discharge clear of all parts of the (4) The jettisoning operation does not airplane; adversely affect the controllability of (2) Have a drain valve— the airplane. (i) That has manual or automatic (d) For reciprocating engine powered means for positive locking in the airplanes, the jettisoning system must closed position; be designed so that it is not possible to (ii) That is readily accessible; jettison the fuel in the tanks used for (iii) That can be easily opened and takeoff and landing below the level al- closed; lowing 45 minutes flight at 75 percent (iv) That allows the fuel to be caught maximum continuous power. However, for examination; if there is an auxiliary control inde- (v) That can be observed for proper pendent of the main jettisoning con- closing; and trol, the system may be designed to (vi) That is either located or pro- jettison all the fuel. tected to prevent fuel spillage in the (e) For turbine engine powered air- event of a landing with landing gear re- planes, the jettisoning system must be tracted. designed so that it is not possible to [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as jettison fuel in the tanks used for take- amended by Amdt. 23–17, 41 FR 55465, Dec. 20, off and landing below the level allow- 1976; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993] ing climb from sea level to 10,000 feet and thereafter allowing 45 minutes § 23.1001 Fuel jettisoning system. cruise at a speed for maximum range. (a) If the design landing weight is (f) The fuel jettisoning valve must be less than that permitted under the re- designed to allow flight crewmembers quirements of § 23.473(b), the airplane to close the valve during any part of must have a fuel jettisoning system in- the jettisoning operation. stalled that is able to jettison enough (g) Unless it is shown that using any fuel to bring the maximum weight means (including flaps, slots, and slats) down to the design landing weight. The for changing the airflow across or average rate of fuel jettisoning must be around the wings does not adversely af- at least 1 percent of the maximum fect fuel jettisoning, there must be a

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placard, adjacent to the jettisoning § 23.1013 Oil tanks. control, to warn flight crewmembers (a) Installation. Each oil tank must be against jettisoning fuel while the installed to— means that change the airflow are (1) Meet the requirements of § 23.967 being used. (a) and (b); and (h) The fuel jettisoning system must (2) Withstand any vibration, inertia, be designed so that any reasonably and fluid loads expected in operation. probable single malfunction in the sys- (b) Expansion space. Oil tank expan- tem will not result in a hazardous con- sion space must be provided so that— dition due to unsymmetrical jetti- (1) Each oil tank used with a recipro- soning of, or inability to jettison, fuel. cating engine has an expansion space of [Amdt. 23–7, 34 FR 13094, Aug. 13, 1969, as not less than the greater of 10 percent amended by Amdt. 23–43, 58 FR 18973, Apr. 9, of the tank capacity or 0.5 gallon, and 1993; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996] each oil tank used with a turbine en- gine has an expansion space of not less OIL SYSTEM than 10 percent of the tank capacity; and § 23.1011 General. (2) It is impossible to fill the expan- (a) For oil systems and components sion space inadvertently with the air- that have been approved under the en- plane in the normal ground attitude. gine airworthiness requirements and (c) Filler connection. Each oil tank where those requirements are equal to filler connection must be marked as or more severe than the corresponding specified in § 23.1557(c). Each recessed requirements of subpart E of this part, oil tank filler connection of an oil tank that approval need not be duplicated. used with a turbine engine, that can re- Where the requirements of subpart E of tain any appreciable quantity of oil, this part are more severe, substan- must have provisions for fitting a tiation must be shown to the require- drain. ments of subpart E of this part. (d) Vent. Oil tanks must be vented as (b) Each engine must have an inde- follows: (1) Each oil tank must be vented to pendent oil system that can supply it the engine from the top part of the ex- with an appropriate quantity of oil at a pansion space so that the vent connec- temperature not above that safe for tion is not covered by oil under any continuous operation. normal flight condition. (c) The usable oil tank capacity may (2) Oil tank vents must be arranged not be less than the product of the en- so that condensed water vapor that durance of the airplane under critical might freeze and obstruct the line can- operating conditions and the maximum not accumulate at any point. oil consumption of the engine under (3) For acrobatic category airplanes, the same conditions, plus a suitable there must be means to prevent haz- margin to ensure adequate circulation ardous loss of oil during acrobatic ma- and cooling. neuvers, including short periods of in- (d) For an oil system without an oil verted flight. transfer system, only the usable oil (e) Outlet. No oil tank outlet may be tank capacity may be considered. The enclosed by any screen or guard that amount of oil in the engine oil lines, would reduce the flow of oil below a the oil radiator, and the feathering re- safe value at any operating tempera- serve, may not be considered. ture. No oil tank outlet diameter may (e) If an oil transfer system is used, be less than the diameter of the engine and the transfer pump can pump some oil pump inlet. Each oil tank used with of the oil in the transfer lines into the a turbine engine must have means to main engine oil tanks, the amount of prevent entrance into the tank itself, oil in these lines that can be pumped or into the tank outlet, of any object by the transfer pump may be included that might obstruct the flow of oil in the oil capacity. through the system. There must be a [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as shutoff valve at the outlet of each oil amended by Amdt. 23–43, 58 FR 18973, Apr. 9, tank used with a turbine engine, unless 1993] the external portion of the oil system

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(including oil tank supports) is fire- (5) The breather outlet is protected proof. against blockage by ice or foreign mat- (f) Flexible liners. Each flexible oil ter. tank liner must be of an acceptable [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as kind. amended by Amdt. 23–7, 34 FR 13094, Aug. 13, (g) Each oil tank filler cap of an oil 1969; Amdt. 23–14, 38 FR 31823, Nov. 19, 1973] tank that is used with an engine must provide an oiltight seal. § 23.1019 Oil strainer or filter. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (a) Each turbine engine installation amended by Amdt. 23–15, 39 FR 35459 Oct. 1, must incorporate an oil strainer or fil- 1974; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993; ter through which all of the engine oil Amdt. 23–51, 61 FR 5137, Feb. 9, 1996] flows and which meets the following re- quirements: § 23.1015 Oil tank tests. (1) Each oil strainer or filter that has Each oil tank must be tested under a bypass, must be constructed and in- § 23.965, except that— stalled so that oil will flow at the nor- (a) The applied pressure must be five mal rate through the rest of the sys- p.s.i. for the tank construction instead tem with the strainer or filter com- of the pressures specified in § 23.965(a); pletely blocked. (b) For a tank with a nonmetallic (2) The oil strainer or filter must liner the test fluid must be oil rather have the capacity (with respect to op- than fuel as specified in § 23.965(d), and erating limitations established for the the slosh test on a specimen liner must engine) to ensure that engine oil sys- be conducted with the oil at 250 °F.; tem functioning is not impaired when and the oil is contaminated to a degree (c) For pressurized tanks used with a (with respect to particle size and den- turbine engine, the test pressure may sity) that is greater than that estab- lished for the engine for its type cer- not be less than 5 p.s.i. plus the max- tification. imum operating pressure of the tank. (3) The oil strainer or filter, unless it [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as is installed at an oil tank outlet, must amended by Amdt. 23–15, 39 FR 35460, Oct. 1, incorporate a means to indicate con- 1974] tamination before it reaches the capac- ity established in accordance with § 23.1017 Oil lines and fittings. paragraph (a)(2) of this section. (a) Oil lines. Oil lines must meet (4) The bypass of a strainer or filter § 23.993 and must accommodate a flow must be constructed and installed so of oil at a rate and pressure adequate that the release of collected contami- for proper engine functioning under nants is minimized by appropriate lo- any normal operating condition. cation of the bypass to ensure that col- (b) Breather lines. Breather lines must lected contaminants are not in the by- be arranged so that— pass flow path. (1) Condensed water vapor or oil that (5) An oil strainer or filter that has might freeze and obstruct the line can- no bypass, except one that is installed not accumulate at any point; at an oil tank outlet, must have a (2) The breather discharge will not means to connect it to the warning constitute a fire hazard if foaming oc- system required in § 23.1305(c)(9). curs, or cause emitted oil to strike the (b) Each oil strainer or filter in a pilot’s windshield; powerplant installation using recipro- (3) The breather does not discharge cating engines must be constructed and into the engine air induction system; installed so that oil will flow at the and normal rate through the rest of the (4) For acrobatic category airplanes, system with the strainer or filter ele- there is no excessive loss of oil from ment completely blocked. the breather during acrobatic maneu- [Amdt. 23–15, 39 FR 35460, Oct. 1, 1974, as vers, including short periods of in- amended by Amdt. 23–29, 49 FR 6847, Feb. 23, verted flight. 1984; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]

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§ 23.1021 Oil system drains. is requested, and after normal engine A drain (or drains) must be provided and auxiliary power unit shutdown. to allow safe drainage of the oil sys- [Doc. No. 26344, 58 FR 18973, Apr. 9, 1993, as tem. Each drain must— amended by Amdt. 23–51, 61 FR 5137, Feb. 9, (a) Be accessible; 1996] (b) Have drain valves, or other clo- sures, employing manual or automatic § 23.1043 Cooling tests. shut-off means for positive locking in (a) General. Compliance with § 23.1041 the closed position; and must be shown on the basis of tests, for (c) Be located or protected to prevent which the following apply: inadvertent operation. (1) If the tests are conducted under [Amdt. 23–29, 49 FR 6847, Feb. 23, 1984, as ambient atmospheric temperature con- amended by Amdt. 23–43, 58 FR 18973, Apr. 9, ditions deviating from the maximum 1993] for which approval is requested, the re- § 23.1023 Oil radiators. corded powerplant temperatures must be corrected under paragraphs (c) and Each oil radiator and its supporting (d) of this section, unless a more ra- structures must be able to withstand tional correction method is applicable. the vibration, inertia, and oil pressure (2) No corrected temperature deter- loads to which it would be subjected in operation. mined under paragraph (a)(1) of this section may exceed established limits. § 23.1027 Propeller feathering system. (3) The fuel used during the cooling (a) If the propeller feathering system tests must be of the minimum grade uses engine oil and that oil supply can approved for the engine. become depleted due to failure of any (4) For turbocharged engines, each part of the oil system, a means must be turbocharger must be operated through incorporated to reserve enough oil to that part of the climb profile for which operate the feathering system. operation with the turbocharger is re- (b) The amount of reserved oil must quested. be enough to accomplish feathering (5) For a reciprocating engine, the and must be available only to the mixture settings must be the leanest feathering pump. recommended for climb. (c) The ability of the system to ac- (b) Maximum ambient atmospheric tem- complish feathering with the reserved perature. A maximum ambient atmos- oil must be shown. pheric temperature corresponding to (d) Provision must be made to pre- sea level conditions of at least 100 de- vent sludge or other foreign matter grees F must be established. The as- from affecting the safe operation of the sumed temperature lapse rate is 3.6 de- propeller feathering system. grees F per thousand feet of altitude [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as above sea level until a temperature of amended by Amdt. 23–14, 38 FR 31823, Nov. 19, ¥69.7 degrees F is reached, above which 1973; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993] altitude the temperature is considered ¥ COOLING constant at 69.7 degrees F. However, for winterization installations, the ap- § 23.1041 General. plicant may select a maximum ambi- The powerplant and auxiliary power ent atmospheric temperature cor- unit cooling provisions must maintain responding to sea level conditions of the temperatures of powerplant compo- less than 100 degrees F. nents and engine fluids, and auxiliary (c) Correction factor (except cylinder power unit components and fluids with- barrels). Temperatures of engine fluids in the limits established for those com- and powerplant components (except ponents and fluids under the most ad- cylinder barrels) for which temperature verse ground, water, and flight oper- limits are established, must be cor- ations to the maximum altitude and rected by adding to them the difference maximum ambient atmospheric tem- perature conditions for which approval

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between the maximum ambient atmos- (3) An operating limitation is pheric temperature for the relevant al- reached. titude for which approval has been re- [Amdt. 23–7, 34 FR 13094, Aug. 13, 1969, as quested and the temperature of the am- amended by Amdt. 23–51, 61 FR 5137, Feb. 9, bient air at the time of the first occur- 1996] rence of the maximum fluid or compo- nent temperature recorded during the § 23.1047 Cooling test procedures for cooling test. reciprocating engine powered air- (d) Correction factor for cylinder barrel planes. temperatures. Cylinder barrel tempera- Compliance with § 23.1041 must be tures must be corrected by adding to shown for the climb (or, for multien- them 0.7 times the difference between gine airplanes with negative one-en- the maximum ambient atmospheric gine-inoperative rates of climb, the de- temperature for the relevant altitude scent) stage of flight. The airplane for which approval has been requested must be flown in the configurations, at and the temperature of the ambient air the speeds and following the procedures at the time of the first occurrence of recommended in the Airplane Flight the maximum cylinder barrel tempera- Manual, that correspond to the appli- ture recorded during the cooling test. cable performance requirements that are critical to cooling. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13094, Aug. 13, [Amdt. 23–51, 61 FR 5137, Feb. 9, 1996] 1969; Amdt. 23–21, 43 FR 2319, Jan. 16, 1978; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996] LIQUID COOLING

§ 23.1045 Cooling test procedures for § 23.1061 Installation. turbine engine powered airplanes. (a) General. Each liquid-cooled engine (a) Compliance with § 23.1041 must be must have an independent cooling sys- shown for all phases of operation. The tem (including coolant tank) installed airplane must be flown in the configu- so that— rations, at the speeds, and following (1) Each coolant tank is supported so the procedures recommended in the that tank loads are distributed over a Airplane Flight Manual for the rel- large part of the tank surface; evant stage of flight, that correspond (2) There are pads or other isolation to the applicable performance require- means between the tank and its sup- ments that are critical to cooling. ports to prevent chafing. (b) Temperatures must be stabilized (3) Pads or any other isolation means under the conditions from which entry that is used must be nonabsorbent or is made into each stage of flight being must be treated to prevent absorption investigated, unless the entry condi- of flammable fluids; and tion normally is not one during which (4) No air or vapor can be trapped in component and engine fluid tempera- any part of the system, except the tures would stabilize (in which case, coolant tank expansion space, during operation through the full entry condi- filling or during operation. tion must be conducted before entry (b) Coolant tank. The tank capacity into the stage of flight being inves- must be at least one gallon, plus 10 per- tigated in order to allow temperatures cent of the cooling system capacity. In to reach their natural levels at the addition— time of entry). The takeoff cooling test (1) Each coolant tank must be able to must be preceded by a period during withstand the vibration, inertia, and which the powerplant component and fluid loads to which it may be sub- engine fluid temperatures are sta- jected in operation; bilized with the engines at ground idle. (2) Each coolant tank must have an (c) Cooling tests for each stage of expansion space of at least 10 percent flight must be continued until— of the total cooling system capacity; (1) The component and engine fluid and temperatures stabilize; (3) It must be impossible to fill the (2) The stage of flight is completed; expansion space inadvertently with the or airplane in the normal ground attitude.

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(c) Filler connection. Each coolant (b) For a tank with a nonmetallic tank filler connection must be marked liner the test fluid must be coolant as specified in § 23.1557(c). In addition— rather than fuel as specified in (1) Spilled coolant must be prevented § 23.965(d), and the slosh test on a speci- from entering the coolant tank com- men liner must be conducted with the partment or any part of the airplane coolant at operating temperature. other than the tank itself; and (2) Each recessed coolant filler con- INDUCTION SYSTEM nection must have a drain that dis- § 23.1091 Air induction system. charges clear of the entire airplane. (d) Lines and fittings. Each coolant (a) The air induction system for each system line and fitting must meet the engine and auxiliary power unit and requirements of § 23.993, except that the their accessories must supply the air inside diameter of the engine coolant required by that engine and auxiliary inlet and outlet lines may not be less power unit and their accessories under than the diameter of the corresponding the operating conditions for which cer- engine inlet and outlet connections. tification is requested. (e) Radiators. Each coolant radiator (b) Each reciprocating engine instal- must be able to withstand any vibra- lation must have at least two separate tion, inertia, and coolant pressure load air intake sources and must meet the to which it may normally be subjected. following: In addition— (1) Primary air intakes may open (1) Each radiator must be supported within the cowling if that part of the to allow expansion due to operating cowling is isolated from the engine ac- temperatures and prevent the trans- cessory section by a fire-resistant dia- mittal of harmful vibration to the radi- phragm or if there are means to pre- ator; and vent the emergence of backfire flames. (2) If flammable coolant is used, the (2) Each alternate air intake must be air intake duct to the coolant radiator located in a sheltered position and may must be located so that (in case of fire) not open within the cowling if the flames from the nacelle cannot strike emergence of backfire flames will re- the radiator. sult in a hazard. (f) Drains. There must be an acces- (3) The supplying of air to the engine sible drain that— through the alternate air intake sys- (1) Drains the entire cooling system tem may not result in a loss of exces- (including the coolant tank, radiator, sive power in addition to the power loss and the engine) when the airplane is in due to the rise in air temperature. the normal ground altitude; (4) Each automatic alternate air door must have an override means acces- (2) Discharges clear of the entire air- sible to the flight crew. plane; and (5) Each automatic alternate air door (3) Has means to positively lock it must have a means to indicate to the closed. flight crew when it is not closed. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (c) For turbine engine powered air- amended by Amdt. 23–43, 58 FR 18973, Apr. 9, planes— 1993] (1) There must be means to prevent hazardous quantities of fuel leakage or § 23.1063 Coolant tank tests. overflow from drains, vents, or other Each coolant tank must be tested components of flammable fluid systems under § 23.965, except that— from entering the engine intake sys- (a) The test required by § 23.965(a)(1) tem; and must be replaced with a similar test (2) The airplane must be designed to using the sum of the pressure devel- prevent water or slush on the runway, oped during the maximum ultimate ac- taxiway, or other airport operating celeration with a full tank or a pres- surfaces from being directed into the sure of 3.5 pounds per square inch, engine or auxiliary power unit air in- whichever is greater, plus the max- take ducts in hazardous quantities. imum working pressure of the system; The air intake ducts must be located or and protected so as to minimize the hazard

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of ingestion of foreign matter during systems not having fuel metering com- takeoff, landing, and taxiing. ponents projecting into the airstream [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as on which ice may form, and intro- amended by Amdt. 23–7, 34 FR 13095, Aug. 13, ducing fuel into the air induction sys- 1969; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993; 58 tem downstream of any components or FR 27060, May 6, 1993; Amdt. 23–51, 61 FR 5137, other obstruction on which ice pro- Feb. 9, 1996] duced by fuel evaporation may form, has a sheltered alternate source of air § 23.1093 Induction system icing pro- with a preheat of not less than 60 °F tection. with the engines at 75 percent of its (a) Reciprocating engines. Each recip- maximum continuous power. rocating engine air induction system (b) Turbine engines. (1) Each turbine must have means to prevent and elimi- engine and its air inlet system must nate icing. Unless this is done by other operate throughout the flight power means, it must be shown that, in air range of the engine (including idling), free of visible moisture at a tempera- ture of 30 °F— without the accumulation of ice on en- (1) Each airplane with sea level en- gine or inlet system components that gines using conventional venturi car- would adversely affect engine oper- buretors has a preheater that can pro- ation or cause a serious loss of power vide a heat rise of 90 °F. with the en- or thrust— gines at 75 percent of maximum contin- (i) Under the icing conditions speci- uous power; fied in appendix C of part 25 of this (2) Each airplane with altitude en- chapter; and gines using conventional venturi car- (ii) In snow, both falling and blowing, buretors has a preheater that can pro- within the limitations established for vide a heat rise of 120 °F. with the en- the airplane for such operation. gines at 75 percent of maximum contin- (2) Each turbine engine must idle for uous power; 30 minutes on the ground, with the air (3) Each airplane with altitude en- bleed available for engine icing protec- gines using fuel metering device tend- tion at its critical condition, without ing to prevent icing has a preheater adverse effect, in an atmosphere that is that, with the engines at 60 percent of at a temperature between 15° and 30 °F maximum continuous power, can pro- (between ¥9° and ¥1 °C) and has a liq- vide a heat rise of— uid water content not less than 0.3 ° (i) 100 F.; or grams per cubic meter in the form of (ii) 40 °F., if a fluid deicing system drops having a mean effective diameter meeting the requirements of §§ 23.1095 not less than 20 microns, followed by through 23.1099 is installed; momentary operation at takeoff power (4) Each airplane with sea level en- or thrust. During the 30 minutes of idle gine(s) using fuel metering device tend- operation, the engine may be run up ing to prevent icing has a sheltered al- ternate source of air with a preheat of periodically to a moderate power or not less than 60 °F with the engines at thrust setting in a manner acceptable 75 percent of maximum continuous to the Administrator. power; (c) Reciprocating engines with Super- (5) Each airplane with sea level or al- chargers. For airplanes with recipro- titude engine(s) using fuel injection cating engines having superchargers to systems having metering components pressurize the air before it enters the on which impact ice may accumulate fuel metering device, the heat rise in has a preheater capable of providing a the air caused by that supercharging at heat rise of 75 °F when the engine is op- any altitude may be utilized in deter- erating at 75 percent of its maximum mining compliance with paragraph (a) continuous power; and of this section if the heat rise utilized (6) Each airplane with sea level or al- is that which will be available, auto- titude engine(s) using fuel injection matically, for the applicable altitudes

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and operating condition because of su- heater is not being used during engine percharging. operation; (b) Allow inspection of the exhaust [Amdt. 23-7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23–15, 39 FR 35460, Oct. 1, manifold parts that it surrounds; and 1974; Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; (c) Allow inspection of critical parts Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. of the preheater itself. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–43, 58 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as FR 18973, Apr. 9, 1993; Amdt. 23–51, 61 FR 5137, amended by Amdt. 23–43, 58 FR 18974, Apr. 9, Feb. 9, 1996] 1993] § 23.1095 Carburetor deicing fluid flow rate. § 23.1103 Induction system ducts. (a) If a carburetor deicing fluid sys- (a) Each induction system duct must tem is used, it must be able to simulta- have a drain to prevent the accumula- neously supply each engine with a rate tion of fuel or moisture in the normal of fluid flow, expressed in pounds per ground and flight attitudes. No drain hour, of not less than 2.5 times the may discharge where it will cause a square root of the maximum contin- fire hazard. uous power of the engine. (b) Each duct connected to compo- (b) The fluid must be introduced into nents between which relative motion the air induction system— could exist must have means for flexi- (1) Close to, and upstream of, the car- bility. buretor; and (c) Each flexible induction system (2) So that it is equally distributed duct must be capable of withstanding over the entire cross section of the in- the effects of temperature extremes, duction system air passages. fuel, oil, water, and solvents to which it is expected to be exposed in service § 23.1097 Carburetor deicing fluid sys- and maintenance without hazardous tem capacity. deterioration or delamination. (a) The capacity of each carburetor (d) For reciprocating engine installa- deicing fluid system— tions, each induction system duct must (1) May not be less than the greater be— of— (1) Strong enough to prevent induc- (i) That required to provide fluid at tion system failures resulting from the rate specified in § 23.1095 for a time normal backfire conditions; and equal to three percent of the maximum (2) Fire resistant in any compart- endurance of the airplane; or ment for which a fire extinguishing (ii) 20 minutes at that flow rate; and system is required. (2) Need not exceed that required for (e) Each inlet system duct for an aux- two hours of operation. iliary power unit must be— (b) If the available preheat exceeds 50 (1) Fireproof within the auxiliary °F. but is less than 100 °F., the capacity power unit compartment; of the system may be decreased in pro- (2) Fireproof for a sufficient distance portion to the heat rise available in ex- upstream of the auxiliary power unit cess of 50 °F. compartment to prevent hot gas re- verse flow from burning through the § 23.1099 Carburetor deicing fluid sys- duct and entering any other compart- tem detail design. ment of the airplane in which a hazard Each carburetor deicing fluid system would be created by the entry of the must meet the applicable requirements hot gases; for the design of a fuel system, except (3) Constructed of materials suitable as specified in §§ 23.1095 and 23.1097. to the environmental conditions ex- pected in service, except in those areas § 23.1101 Induction air preheater de- requiring fireproof or fire resistant ma- sign. terials; and Each exhaust-heated, induction air (4) Constructed of materials that will preheater must be designed and con- not absorb or trap hazardous quantities structed to— of flammable fluids that could be ig- (a) Ensure ventilation of the pre- nited by a surge or reverse-flow condi- heater when the induction air pre- tion.

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(f) Induction system ducts that sup- following any probable failure of the ply air to a cabin pressurization sys- turbocharger or its lubrication system. tem must be suitably constructed of (b) The turbocharger supply air must material that will not produce haz- be taken from a source where it cannot ardous quantities of toxic gases or iso- be contaminated by harmful or haz- lated to prevent hazardous quantities ardous gases or vapors following any of toxic gases from entering the cabin probable failure or malfunction of the during a powerplant fire. engine exhaust, hydraulic, fuel, or oil system. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13095, Aug. 13, [Amdt. 23–42, 56 FR 354, Jan. 3, 1991] 1969; Amdt. 23–43, 58 FR 18974, Apr. 9, 1993] § 23.1111 Turbine engine bleed air sys- § 23.1105 Induction system screens. tem. If induction system screens are For turbine engine bleed air systems, used— the following apply: (a) Each screen must be upstream of (a) No hazard may result if duct rup- the carburetor or fuel injection system. ture or failure occurs anywhere be- tween the engine port and the airplane (b) No screen may be in any part of unit served by the bleed air. the induction system that is the only (b) The effect on airplane and engine passage through which air can reach performance of using maximum bleed the engine, unless— air must be established. (1) The available preheat is at least (c) Hazardous contamination of cabin ° 100 F.; and air systems may not result from fail- (2) The screen can be deiced by heat- ures of the engine lubricating system. ed air; (c) No screen may be deiced by alco- [Amdt. 23–7, 34 FR 13095, Aug. 13, 1969, as hol alone; and amended by Amdt. 23–17, 41 FR 55465, Dec. 20, 1976] (d) It must be impossible for fuel to strike any screen. EXHAUST SYSTEM

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 § 23.1121 General. FR 258, Jan. 9, 1996, as amended by Amdt. 23– 51, 61 FR 5137, Feb. 9, 1996] For powerplant and auxiliary power unit installations, the following § 23.1107 Induction system filters. apply— If an air filter is used to protect the (a) Each exhaust system must ensure engine against foreign material par- safe disposal of exhaust gases without ticles in the induction air supply— fire hazard or carbon monoxide con- (a) Each air filter must be capable of tamination in any personnel compart- withstanding the effects of tempera- ment. ture extremes, rain, fuel, oil, and sol- (b) Each exhaust system part with a vents to which it is expected to be ex- surface hot enough to ignite flammable posed in service and maintenance; and fluids or vapors must be located or shielded so that leakage from any sys- (b) Each air filter shall have a design tem carrying flammable fluids or va- feature to prevent material separated pors will not result in a fire caused by from the filter media from interfering impingement of the fluids or vapors on with proper fuel metering operation. any part of the exhaust system includ- [Doc. No. 26344, 58 FR 18974, Apr. 9, 1993, as ing shields for the exhaust system. amended by Amdt. 23–51, 61 FR 5137, Feb. 9, (c) Each exhaust system must be sep- 1996] arated by fireproof shields from adja- cent flammable parts of the airplane § 23.1109 Turbocharger bleed air sys- that are outside of the engine and aux- tem. iliary power unit compartments. The following applies to (d) No exhaust gases may discharge turbocharged bleed air systems used dangerously near any fuel or oil system for cabin pressurization: drain. (a) The cabin air system may not be (e) No exhaust gases may be dis- subject to hazardous contamination charged where they will cause a glare

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seriously affecting pilot vision at (3) Each exchanger must have cooling night. provisions wherever it is subject to (f) Each exhaust system component contact with exhaust gases. must be ventilated to prevent points of (b) Each heat exchanger used for excessively high temperature. heating ventilating air must be con- (g) If significant traps exist, each structed so that exhaust gases may not turbine engine and auxiliary power enter the ventilating air. unit exhaust system must have drains discharging clear of the airplane, in [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as any normal ground and flight attitude, amended by Amdt. 23–17, 41 FR 55465, Dec. 20, to prevent fuel accumulation after the 1976] failure of an attempted engine or auxil- iary power unit start. POWERPLANT CONTROLS AND (h) Each exhaust heat exchanger ACCESSORIES must incorporate means to prevent blockage of the exhaust port after any § 23.1141 Powerplant controls: Gen- internal heat exchanger failure. eral. (i) For the purpose of compliance (a) Powerplant controls must be lo- with § 23.603, the failure of any part of cated and arranged under § 23.777 and the exhaust system will be considered marked under § 23.1555(a). to adversely affect safety. (b) Each flexible control must be [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as shown to be suitable for the particular amended by Amdt. 23–7, 34 FR 13095, Aug. 13, application. 1969; Amdt. 23–18, 42 FR 15042, Mar. 17, 1977; (c) Each control must be able to Amdt. 23–43, 58 FR 18974, Apr. 9, 1993; Amdt. maintain any necessary position with- 23–51, 61 FR 5137, Feb. 9, 1996] out— § 23.1123 Exhaust system. (1) Constant attention by flight crew (a) Each exhaust system must be fire- members; or proof and corrosion-resistant, and must (2) Tendency to creep due to control have means to prevent failure due to loads or vibration. expansion by operating temperatures. (d) Each control must be able to (b) Each exhaust system must be sup- withstand operating loads without fail- ported to withstand the vibration and ure or excessive deflection. inertia loads to which it may be sub- (e) For turbine engine powered air- jected in operation. planes, no single failure or malfunc- (c) Parts of the system connected to tion, or probable combination thereof, components between which relative in any powerplant control system may motion could exist must have means cause the failure of any powerplant for flexibility. function necessary for safety. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (f) The portion of each powerplant amended by Amdt. 23–43, 58 FR 18974, Apr. 9, control located in the engine compart- 1993] ment that is required to be operated in the event of fire must be at least fire § 23.1125 Exhaust heat exchangers. resistant. For reciprocating engine powered air- (g) Powerplant valve controls located planes the following apply: in the cockpit must have— (a) Each exhaust heat exchanger (1) For manual valves, positive stops must be constructed and installed to or in the case of fuel valves suitable withstand the vibration, inertia, and other loads that it may be subjected to index provisions, in the open and closed in normal operation. In addition— position; and (1) Each exchanger must be suitable (2) For power-assisted valves, a for continued operation at high tem- means to indicate to the flight crew peratures and resistant to corrosion when the valve— from exhaust gases; (i) Is in the fully open or fully closed (2) There must be means for inspec- position; or tion of critical parts of each exchanger; and

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(ii) Is moving between the fully open device, the airplane is capable of con- and fully closed position. tinued safe flight and landing. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as [Amdt. 23–7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23–7, 34 FR 13095, Aug. 13, amended by Amdt. 23–17, 41 FR 55465, Dec. 20, 1969; Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; 1976; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–18, 42 FR 15042, Mar. 17, 1977; Amdt. Amdt. 23–43, 58 FR 18974, Apr. 9, 1993; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996] 23–51, 61 FR 5137, Feb. 9, 1996]

§ 23.1142 Auxiliary power unit con- § 23.1145 Ignition switches. trols. (a) Ignition switches must control Means must be provided on the flight and shut off each ignition circuit on deck for the starting, stopping, moni- toring, and emergency shutdown of each engine. each installed auxiliary power unit. (b) There must be means to quickly shut off all ignition on multiengine air- [Doc. No. 26344, 58 FR 18974, Apr. 9, 1993] planes by the grouping of switches or by a master ignition control. § 23.1143 Engine controls. (c) Each group of ignition switches, (a) There must be a separate power or except ignition switches for turbine en- thrust control for each engine and a gines for which continuous ignition is separate control for each supercharger not required, and each master ignition that requires a control. control must have a means to prevent (b) Power, thrust, and supercharger its inadvertent operation. controls must be arranged to allow— (1) Separate control of each engine [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 and each supercharger; and FR 258, Jan. 9, 1965, as amended by Amdt. 23– (2) Simultaneous control of all en- 18, 42 FR 15042, Mar. 17, 1977; Amdt. 23–43, 58 gines and all superchargers. FR 18974, Apr. 9, 1993] (c) Each power, thrust, or super- charger control must give a positive § 23.1147 Mixture controls. and immediate responsive means of (a) If there are mixture controls, controlling its engine or supercharger. each engine must have a separate con- (d) The power, thrust, or super- trol, and each mixture control must charger controls for each engine or su- have guards or must be shaped or ar- percharger must be independent of ranged to prevent confusion by feel those for every other engine or super- with other controls. charger. (1) The controls must be grouped and (e) For each fluid injection (other arranged to allow— than fuel) system and its controls not (i) Separate control of each engine; provided and approved as part of the and engine, the applicant must show that (ii) Simultaneous control of all en- the flow of the injection fluid is ade- gines. quately controlled. (f) If a power, thrust, or a fuel con- (2) The controls must require a sepa- trol (other than a mixture control) in- rate and distinct operation to move the corporates a fuel shutoff feature, the control toward lean or shut-off posi- control must have a means to prevent tion. the inadvertent movement of the con- (b) For reciprocating single-engine trol into the off position. The means airplanes, each manual engine mixture must— control must be designed so that, if the (1) Have a positive lock or stop at the control separates at the engine fuel idle position; and metering device, the airplane is capa- (2) Require a separate and distinct ble of continued safe flight and land- operation to place the control in the ing. shutoff position. (g) For reciprocating single-engine [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13096, Aug. 13, airplanes, each power or thrust control 1969; Amdt. 23–33, 51 FR 26657, July 24, 1986; must be designed so that if the control Amdt. 23–43, 58 FR 18974, Apr. 9, 1993] separates at the engine fuel metering

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§ 23.1149 Propeller speed and pitch (b) Electrical equipment subject to controls. arcing or sparking must be installed to (a) If there are propeller speed or minimize the probability of contact pitch controls, they must be grouped with any flammable fluids or vapors and arranged to allow— that might be present in a free state. (1) Separate control of each pro- (c) Each generator rated at or more peller; and than 6 kilowatts must be designed and (2) Simultaneous control of all pro- installed to minimize the probability pellers. of a fire hazard in the event it malfunc- (b) The controls must allow ready tions. synchronization of all propellers on (d) If the continued rotation of any multiengine airplanes. accessory remotely driven by the en- § 23.1153 Propeller feathering controls. gine is hazardous when malfunctioning occurs, a means to prevent rotation If there are propeller feathering con- without interfering with the continued trols installed, it must be possible to operation of the engine must be pro- feather each propeller separately. Each vided. control must have a means to prevent inadvertent operation. (e) Each accessory driven by a gear- box that is not approved as part of the [Doc. No. 27804, 61 FR 5138, Feb. 9, 1996] powerplant driving the gearbox must— (1) Have torque limiting means to § 23.1155 Turbine engine reverse thrust and propeller pitch settings prevent the torque limits established below the flight regime. for the affected drive from being ex- For turbine engine installations, ceeded; each control for reverse thrust and for (2) Use the provisions on the gearbox propeller pitch settings below the for mounting; and flight regime must have means to pre- (3) Be sealed to prevent contamina- vent its inadvertent operation. The tion of the gearbox oil system and the means must have a positive lock or accessory system. stop at the flight idle position and [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as must require a separate and distinct amended by Amdt. 23–14, 38 FR 31823, Nov. 19, operation by the crew to displace the 1973; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; control from the flight regime (forward Amdt. 23–34, 52 FR 1832, Jan. 15, 1987; Amdt. thrust regime for turbojet powered air- 23–42, 56 FR 354, Jan. 3, 1991] planes). [Amdt. 23–7, 34 FR 13096, Aug. 13, 1969] § 23.1165 Engine ignition systems. (a) Each battery ignition system § 23.1157 Carburetor air temperature must be supplemented by a generator controls. that is automatically available as an There must be a separate carburetor alternate source of electrical energy to air temperature control for each en- allow continued engine operation if gine. any battery becomes depleted. § 23.1163 Powerplant accessories. (b) The capacity of batteries and gen- erators must be large enough to meet (a) Each engine mounted accessory the simultaneous demands of the en- must— gine ignition system and the greatest (1) Be approved for mounting on the demands of any electrical system com- engine involved and use the provisions ponents that draw from the same on the engines for mounting; or (2) Have torque limiting means on all source. accessory drives in order to prevent the (c) The design of the engine ignition torque limits established for those system must account for— drives from being exceeded; and (1) The condition of an inoperative (3) In addition to paragraphs (a)(1) or generator; (a)(2) of this section, be sealed to pre- (2) The condition of a completely de- vent contamination of the engine oil pleted battery with the generator run- system and the accessory system. ning at its normal operating speed; and

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(3) The condition of a completely de- § 23.1351(e), located behind the engine- pleted battery with the generator oper- compartment firewall must be con- ating at idling speed, if there is only structed of such materials and located one battery. at such distances from the firewall (d) There must be means to warn ap- that they will not suffer damage suffi- propriate crewmembers if malfunc- cient to endanger the airplane if a por- tioning of any part of the electrical tion of the engine side of the firewall is system is causing the continuous dis- subjected to a flame temperature of charge of any battery used for engine not less than 2000 °F for 15 minutes. ignition. (e) Each turbine engine ignition sys- [Amdt. 23–14, 38 FR 31816, Nov. 19, 1973] tem must be independent of any elec- trical circuit that is not used for as- § 23.1183 Lines, fittings, and compo- nents. sisting, controlling, or analyzing the operation of that system. (a) Except as provided in paragraph (f) In addition, for commuter cat- (b) of this section, each component, egory airplanes, each turbopropeller ig- line, and fitting carrying flammable nition system must be an essential fluids, gas, or air in any area subject to electrical load. engine fire conditions must be at least fire resistant, except that flammable [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–17, 41 FR 55465 Dec. 20, fluid tanks and supports which are part 1976; Amdt. 23–34, 52 FR 1833, Jan. 15, 1987] of and attached to the engine must be fireproof or be enclosed by a fireproof POWERPLANT FIRE PROTECTION shield unless damage by fire to any non-fireproof part will not cause leak- § 23.1181 Designated fire zones; re- age or spillage of flammable fluid. gions included. Components must be shielded or lo- Designated fire zones are— cated so as to safeguard against the ig- (a) For reciprocating engines— nition of leaking flammable fluid. (1) The power section; Flexible hose assemblies (hose and end (2) The accessory section; fittings) must be shown to be suitable (3) Any complete powerplant com- for the particular application. An inte- partment in which there is no isolation gral oil sump of less than 25–quart ca- between the power section and the ac- pacity on a reciprocating engine need cessory section. not be fireproof nor be enclosed by a (b) For turbine engines— fireproof shield. (1) The compressor and accessory sec- (b) Paragraph (a) of this section does tions; not apply to— (2) The combustor, turbine and tail- (1) Lines, fittings, and components pipe sections that contain lines or com- which are already approved as part of a ponents carrying flammable fluids or type certificated engine; and gases. (3) Any complete powerplant com- (2) Vent and drain lines, and their fit- partment in which there is no isolation tings, whose failure will not result in, between compressor, accessory, com- or add to, a fire hazard. bustor, turbine, and tailpipe sections. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (c) Any auxiliary power unit com- amended by Amdt. 23–5, 32 FR 6912, May 5, partment; and 1967; Amdt. 23–15, 39 FR 35460, Oct. 1, 1974; (d) Any fuel-burning heater, and Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. other combustion equipment installa- 23–51, 61 FR 5138, Feb. 9, 1996] tion described in § 23.859. § 23.1189 Shutoff means. [Doc. No. 26344, 58 FR 18975, Apr. 9, 1993, as amended by Amdt. 23–51, 61 FR 5138, Feb. 9, (a) For each multiengine airplane the 1996] following apply: (1) Each engine installation must § 23.1182 Nacelle areas behind fire- have means to shut off or otherwise walls. prevent hazardous quantities of fuel, Components, lines, and fittings, ex- oil, deicing fluid, and other flammable cept those subject to the provisions of liquids from flowing into, within, or

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through any engine compartment, ex- firewalls, shrouds, or equivalent cept in lines, fittings, and components means. forming an integral part of an engine. (b) Each firewall or shroud must be (2) The closing of the fuel shutoff constructed so that no hazardous quan- valve for any engine may not make any tity of liquid, gas, or flame can pass fuel unavailable to the remaining en- from the compartment created by the gines that would be available to those firewall or shroud to other parts of the engines with that valve open. airplane. (3) Operation of any shutoff means (c) Each opening in the firewall or may not interfere with the later emer- shroud must be sealed with close fit- gency operation of other equipment ting, fireproof grommets, bushings, or such as propeller feathering devices. firewall fittings. (4) Each shutoff must be outside of (d) [Reserved] the engine compartment unless an (e) Each firewall and shroud must be equal degree of safety is provided with fireproof and protected against corro- the shutoff inside the compartment. sion. (5) Not more than one quart of flam- (f) Compliance with the criteria for mable fluid may escape into the engine compartment after engine shutoff. For fireproof materials or components those installations where the flam- must be shown as follows: mable fluid that escapes after shut- (1) The flame to which the materials down cannot be limited to one quart, it or components are subjected must be must be demonstrated that this greater 2,000 ±150 °F. amount can be safely contained or (2) Sheet materials approximately 10 drained overboard. inches square must be subjected to the (6) There must be means to guard flame from a suitable burner. against inadvertent operation of each (3) The flame must be large enough shutoff means, and to make it possible to maintain the required test tempera- for the crew to reopen the shutoff ture over an area approximately five means in flight after it has been closed. inches square. (b) Turbine engine installations need (g) Firewall materials and fittings not have an engine oil system shutoff must resist flame penetration for at if— least 15 minutes. (1) The oil tank is integral with, or (h) The following materials may be mounted on, the engine; and used in firewalls or shrouds without (2) All oil system components exter- being tested as required by this sec- nal to the engine are fireproof or lo- tion: cated in areas not subject to engine (1) Stainless steel sheet, 0.015 inch fire conditions. thick. (c) Power operated valves must have (2) Mild steel sheet (coated with alu- means to indicate to the flight crew minum or otherwise protected against when the valve has reached the se- corrosion) 0.018 inch thick. lected position and must be designed so (3) Terne plate, 0.018 inch thick. that the valve will not move from the (4) Monel metal, 0.018 inch thick. selected position under vibration con- ditions likely to exist at the valve lo- (5) Steel or copper base alloy firewall cation. fittings. (6) Titanium sheet, 0.016 inch thick. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13096, Aug. 13, [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as 1969; Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; amended by Amdt. 23–43, 58 FR 18975, Apr. 9, Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 1993; 58 FR 27060, May 6, 1993; Amdt. 23–51, 61 23–43, 58 FR 18975, Apr. 9, 1993] FR 5138, Feb. 9, 1996]

§ 23.1191 Firewalls. § 23.1192 Engine accessory compart- (a) Each engine, auxiliary power ment diaphragm. unit, fuel burning heater, and other For aircooled radial engines, the en- combustion equipment, must be iso- gine power section and all portions of lated from the rest of the airplane by the exhaust sytem must be isolated

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from the engine accessory compart- ponents carrying flammable fluids or ment by a diaphragm that meets the gases for which a fire originating in firewall requirements of § 23.1191. these sections is shown to be control- [Amdt. 23–14, 38 FR 31823, Nov. 19, 1973] lable, a fire extinguisher system must serve each engine compartment; § 23.1193 Cowling and nacelle. (2) The fire extinguishing system, the (a) Each cowling must be constructed quantity of the extinguishing agent, and supported so that it can resist any the rate of discharge, and the discharge vibration, inertia, and air loads to distribution must be adequate to extin- which it may be subjected in operation. guish fires. An individual ‘‘one shot’’ (b) There must be means for rapid system may be used. and complete drainage of each part of (3) The fire extinguishing system for the cowling in the normal ground and a nacelle must be able to simulta- flight attitudes. Drain operation may neously protect each compartment of be shown by test, analysis, or both, to the nacelle for which protection is pro- ensure that under normal aerodynamic vided. pressure distribution expected in serv- (b) If an auxiliary power unit is in- ice each drain will operate as designed. stalled in any airplane certificated to No drain may discharge where it will this part, that auxiliary power unit cause a fire hazard. compartment must be served by a fire (c) Cowling must be at least fire re- extinguishing system meeting the re- sistant. quirements of paragraph (a)(2) of this (d) Each part behind an opening in section. the engine compartment cowling must be at least fire resistant for a distance [Amdt. 23–34, 52 FR 1833, Jan. 15, 1987, as of at least 24 inches aft of the opening. amended by Amdt. 23–43, 58 FR 18975, Apr. 9, (e) Each part of the cowling subjected 1993] to high temperatures due to its near- ness to exhaust sytem ports or exhaust § 23.1197 Fire extinguishing agents. gas impingement, must be fire proof. For commuter category airplanes, (f) Each nacelle of a multiengine air- the following applies: plane with supercharged engines must (a) Fire extinguishing agents must— be designed and constructed so that (1) Be capable of extinguishing with the landing gear retracted, a fire flames emanating from any burning of in the engine compartment will not fluids or other combustible materials burn through a cowling or nacelle and in the area protected by the fire extin- enter a nacelle area other than the en- guishing system; and gine compartment. (g) In addition, for commuter cat- (2) Have thermal stability over the egory airplanes, the airplane must be temperature range likely to be experi- designed so that no fire originating in enced in the compartment in which any engine compartment can enter, ei- they are stored. ther through openings or by burn- (b) If any toxic extinguishing agent is through, any other region where it used, provisions must be made to pre- would create additional hazards. vent harmful concentrations of fluid or fluid vapors (from leakage during nor- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 mal operation of the airplane or as a FR 258, Jan. 9, 1965, as amended by Amdt. 23– 18, 42 FR 15042, Mar. 17, 1977; Amdt. 23–34, 52 result of discharging the fire extin- FR 1833, Jan. 15, 1987; 58 FR 18975, Apr. 9, guisher on the ground or in flight) from 1993] entering any personnel compartment, even though a defect may exist in the § 23.1195 Fire extinguishing systems. extinguishing system. This must be (a) For commuter category airplanes, shown by test except for built-in car- fire extinguishing systems must be in- bon dioxide fuselage compartment fire stalled and compliance shown with the extinguishing systems for which— following: (1) Five pounds or less of carbon diox- (1) Except for combustor, turbine, ide will be discharged, under estab- and tailpipe sections of turbine-engine lished fire control procedures, into any installations that contain lines or com- fuselage compartment; or

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(2) Protective breathing equipment is § 23.1203 Fire detector system. available for each flight crewmember (a) There must be means that ensure on flight deck duty. the prompt detection of a fire in— [Amdt. 23–34, 52 FR 1833, Jan. 15, 1987] (1) An engine compartment of— (i) Multiengine turbine powered air- § 23.1199 Extinguishing agent con- planes; tainers. (ii) Multiengine reciprocating engine For commuter category airplanes, powered airplanes incorporating the following applies: turbochargers; (a) Each extinguishing agent con- (iii) Airplanes with engine(s) located tainer must have a pressure relief to where they are not readily visible from prevent bursting of the container by the cockpit; and excessive internal pressures. (iv) All commuter category air- (b) The discharge end of each dis- charge line from a pressure relief con- planes. nection must be located so that dis- (2) The auxiliary power unit compart- charge of the fire extinguishing agent ment of any airplane incorporating an would not damage the airplane. The auxiliary power unit. line must also be located or protected (b) Each fire detector must be con- to prevent clogging caused by ice or structed and installed to withstand the other foreign matter. vibration, inertia, and other loads to (c) A means must be provided for which it may be subjected in operation. each fire extinguishing agent container (c) No fire detector may be affected to indicate that the container has dis- by any oil, water, other fluids, or charged or that the charging pressure fumes that might be present. is below the established minimum nec- (d) There must be means to allow the essary for proper functioning. crew to check, in flight, the func- (d) The temperature of each con- tioning of each fire detector electric tainer must be maintained, under in- circuit. tended operating conditions, to prevent (e) Wiring and other components of the pressure in the container from— each fire detector system in a des- (1) Falling below that necessary to ignated fire zone must be at least fire provide an adequate rate of discharge; resistant. or (2) Rising high enough to cause pre- [Amdt. 23–18, 42 FR 15042, Mar. 17, 1977, as mature discharge. amended by Amdt. 23–34, 52 FR 1833, Jan. 15, (e) If a pyrotechnic capsule is used to 1987; Amdt. 23–43, 58 FR 18975, Apr. 9, 1993; discharge the extinguishing agent, Amdt. 23–51, 61 FR 5138, Feb. 9, 1996] each container must be installed so that temperature conditions will not Subpart F—Equipment cause hazardous deterioration of the pyrotechnic capsule. GENERAL [Amdt. 23–34, 52 FR 1833, Jan. 15, 1987; 52 FR § 23.1301 Function and installation. 34745, Sept. 14, 1987] Each item of installed equipment § 23.1201 Fire extinguishing systems must— materials. (a) Be of a kind and design appro- For commuter category airplanes, priate to its intended function. the following apply: (b) Be labeled as to its identification, (a) No material in any fire extin- function, or operating limitations, or guishing system may react chemically any applicable combination of these with any extinguishing agent so as to factors; create a hazard. (c) Be installed according to limita- (b) Each system component in an en- tions specified for that equipment; and gine compartment must be fireproof. (d) Function properly when installed. [Amdt. 23–34, 52 FR 1833, Jan. 15, 1987; 52 FR [Amdt. 23–20, 42 FR 36968, July 18, 1977] 7262, Mar. 9, 1987]

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§ 23.1303 Flight and navigation instru- (ii) Continues reliable operation for a ments. minimum of 30 minutes after total fail- The following are the minimum re- ure of the electrical generating system; quired flight and navigation instru- (iii) Operates independently of any ments: other attitude indicating system; (a) An airspeed indicator. (iv) Is operative without selection (b) An altimeter. after total failure of the electrical gen- (c) A direction indicator (non- erating system; stabilized magnetic compass). (v) Is located on the instrument (d) For reciprocating engine-powered panel in a position acceptable to the airplanes of more than 6,000 pounds Administrator that will make it plain- maximum weight and turbine engine ly visible to and usable by any pilot at powered airplanes, a free air tempera- the pilot’s station; and ture indicator or an air-temperature (vi) Is appropriately lighted during indicator which provides indications all phases of operation. that are convertible to free-air. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (e) A speed warning device for— amended by Amdt. 23–17, 41 FR 55465, Dec. 20, (1) Turbine engine powered airplanes; 1976; Amdt. 23–43, 58 FR 18975, Apr. 9, 1993; and Amdt. 23–49, 61 FR 5168, Feb. 9, 1996] (2) Other airplanes for which VMO/ MMO and VD/MD are established under § 23.1305 Powerplant instruments. §§ 23.335(b)(4) and 23.1505(c) if VMO/MMO The following are required power- is greater than 0.8 VD/MD. plant instruments: The speed warning device must give (a) For all airplanes. (1) A fuel quan- effective aural warning (differing dis- tity indicator for each fuel tank, in- tinctively from aural warnings used for stalled in accordance with § 23.1337(b). other purposes) to the pilots whenever (2) An oil pressure indicator for each the speed exceeds VMO plus 6 knots or engine. MMO+0.01. The upper limit of the pro- (3) An oil temperature indicator for duction tolerance for the warning de- each engine. vice may not exceed the prescribed (4) An oil quantity measuring device warning speed. The lower limit of the for each oil tank which meets the re- warning device must be set to mini- quirements of § 23.1337(d). mize nuisance warning; (5) A fire warning means for those (f) When an attitude display is in- airplanes required to comply with stalled, the instrument design must § 23.1203. not provide any means, accessible to the flightcrew, of adjusting the relative (b) For reciprocating engine-powered positions of the attitude reference sym- airplanes. In addition to the powerplant bol and the horizon line beyond that instruments required by paragraph (a) necessary for parallax correction. of this section, the following power- plant instruments are required: (g) In addition, for commuter cat- egory airplanes: (1) An induction system air tempera- (1) If airspeed limitations vary with ture indicator for each engine equipped altitude, the airspeed indicator must with a preheater and having induction have a maximum allowable airspeed in- air temperature limitations that can be exceeded with preheat. dicator showing the variation of VMO with altitude. (2) A tachometer indicator for each (2) The altimeter must be a sensitive engine. type. (3) A cylinder head temperature indi- (3) Having a passenger seating con- cator for— figuration of 10 or more, excluding the (i) Each air-cooled engine with cowl pilot’s seats and that are approved for flaps; IFR operations, a third attitude instru- (ii) [Reserved] ment must be provided that: (iii) Each commuter category air- (i) Is powered from a source inde- plane. pendent of the electrical generating (4) For each pump-fed engine, a system; means:

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(i) That continuously indicates, to pilot of the occurrence of contamina- the pilot, the fuel pressure or fuel flow; tion of the strainer or filter screen be- or fore it reaches the capacity established (ii) That continuously monitors the in accordance with § 23.1019(a)(5). fuel system and warns the pilot of any (10) An indicating means to indicate fuel flow trend that could lead to en- the functioning of any heater used to gine failure. prevent ice clogging of fuel system (5) A manifold pressure indicator for components. each altitude engine and for each en- (d) For turbojet/turbofan engine-pow- gine with a controllable propeller. ered airplanes. In addition to the power- (6) For each turbocharger installa- plant instruments required by para- tion: (i) If limitations are established for graphs (a) and (c) of this section, the either carburetor (or manifold) air following powerplant instruments are inlet temperature or exhaust gas or required: turbocharger turbine inlet tempera- (1) For each engine, an indicator to ture, indicators must be furnished for indicate thrust or to indicate a param- each temperature for which the limita- eter that can be related to thrust, in- tion is established unless it is shown cluding a free air temperature indi- that the limitation will not be exceed- cator if needed for this purpose. ed in all intended operations. (2) For each engine, a position indi- (ii) If its oil system is separate from cating means to indicate to the flight the engine oil system, oil pressure and crew when the thrust reverser, if in- oil temperature indicators must be pro- stalled, is in the reverse thrust posi- vided. tion. (7) A coolant temperature indicator (e) For turbopropeller-powered air- for each liquid-cooled engine. planes. In addition to the powerplant (c) For turbine engine-powered air- instruments required by paragraphs (a) planes. In addition to the powerplant and (c) of this section, the following instruments required by paragraph (a) powerplant instruments are required: of this section, the following power- (1) A torque indicator for each en- plant instruments are required: (1) A gas temperature indicator for gine. each engine. (2) A position indicating means to in- (2) A fuel flowmeter indicator for dicate to the flight crew when the pro- each engine. peller blade angle is below the flight (3) A fuel low pressure warning low pitch position, for each propeller, means for each engine. unless it can be shown that such occur- (4) A fuel low level warning means for rence is highly improbable. any fuel tank that should not be de- [Doc. No. 26344, 58 FR 18975, Apr. 9, 1993; 58 pleted of fuel in normal operations. FR 27060, May 6, 1993; Amdt. 23–51, 61 FR 5138, (5) A tachometer indicator (to indi- Feb. 9, 1996; Amdt. 23–52, 61 FR 13644, Mar. 27, cate the speed of the rotors with estab- 1996] lished limiting speeds) for each engine. (6) An oil low pressure warning § 23.1307 Miscellaneous equipment. means for each engine. The equipment necessary for an air- (7) An indicating means to indicate the functioning of the powerplant ice plane to operate at the maximum oper- protection system for each engine. ating altitude and in the kinds of oper- (8) For each engine, an indicating ation and meteorological conditions means for the fuel strainer or filter re- for which certification is requested and quired by § 23.997 to indicate the occur- is approved in accordance with § 23.1559 rence of contamination of the strainer must be included in the type design. or filter before it reaches the capacity [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 established in accordance with FR 258, Jan. 9, 1965, as amended by Amdt. 23– § 23.997(d). 23, 43 FR 50593, Oct. 30, 1978; Amdt. 23–43, 58 (9) For each engine, a warning means FR 18976, Apr. 9, 1993; Amdt. 23–49, 61 FR 5168, for the oil strainer or filter required by Feb. 9, 1996] § 23.1019, if it has no bypass, to warn the

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§ 23.1308 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 44024, Aug. 6, fected during and after the time the 2007] airplane is exposed to HIRF environ- ment I, as described in appendix J to § 23.1309 Equipment, systems, and in- this part; stallations. (2) The system automatically recov- (a) Each item of equipment, each sys- ers normal operation of that function, in a timely manner, after the airplane tem, and each installation: is exposed to HIRF environment I, as (1) When performing its intended described in appendix J to this part, function, may not adversely affect the unless the system’s recovery conflicts response, operation, or accuracy of with other operational or functional any— requirements of the system; and (i) Equipment essential to safe oper- (3) The system is not adversely af- ation; or fected during and after the time the (ii) Other equipment unless there is a airplane is exposed to HIRF environ- means to inform the pilot of the effect. ment II, as described in appendix J to (2) In a single-engine airplane, must this part. be designed to minimize hazards to the (b) Each electrical and electronic airplane in the event of a probable mal- system that performs a function whose function or failure. failure would significantly reduce the (3) In a multiengine airplane, must be capability of the airplane or the ability designed to prevent hazards to the air- of the flightcrew to respond to an ad- plane in the event of a probable mal- verse operating condition must be de- function or failure. signed and installed so the system is (4) In a commuter category airplane, not adversely affected when the equip- must be designed to safeguard against ment providing the function is exposed hazards to the airplane in the event of to equipment HIRF test level 1 or 2, as their malfunction or failure. described in appendix J to this part. (b) The design of each item of equip- (c) Each electrical and electronic sys- ment, each system, and each installa- tem that performs a function whose failure would reduce the capability of tion must be examined separately and the airplane or the ability of the in relationship to other airplane sys- flightcrew to respond to an adverse op- tems and installations to determine if erating condition must be designed and the airplane is dependent upon its func- installed so the system is not adversely tion for continued safe flight and land- affected when the equipment providing ing and, for airplanes not limited to the function is exposed to equipment VFR conditions, if failure of a system HIRF test level 3, as described in ap- would significantly reduce the capa- pendix J to this part. bility of the airplane or the ability of (d) Before December 1, 2012, an elec- the crew to cope with adverse oper- trical or electronic system that per- ating conditions. Each item of equip- forms a function whose failure would ment, each system, and each installa- prevent the continued safe flight and tion identified by this examination as landing of an airplane may be designed one upon which the airplane is depend- and installed without meeting the pro- ent for proper functioning to ensure visions of paragraph (a) provided— continued safe flight and landing, or (1) The system has previously been whose failure would significantly re- shown to comply with special condi- duce the capability of the airplane or

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the ability of the crew to cope with ad- (1) Loads connected to the power dis- verse operating conditions, must be de- tribution system with the system func- signed to comply with the following ad- tioning normally. ditional requirements: (2) Essential loads after failure of— (1) It must perform its intended func- (i) Any one engine on two-engine air- tion under any foreseeable operating planes; or condition. (ii) Any two engines on an airplane (2) When systems and associated with three or more engines; or components are considered separately (iii) Any power converter or energy and in relation to other systems— storage device. (i) The occurrence of any failure con- (3) Essential loads for which an alter- dition that would prevent the contin- nate source of power is required, as ap- ued safe flight and landing of the air- plicable, by the operating rules of this plane must be extremely improbable; chapter, after any failure or malfunc- and tion in any one power supply system, (ii) The occurrence of any other fail- distribution system, or other utiliza- ure condition that would significantly tion system. reduce the capability of the airplane or the ability of the crew to cope with ad- (d) In determining compliance with verse operating conditions must be im- paragraph (c)(2) of this section, the probable. power loads may be assumed to be re- (3) Warning information must be pro- duced under a monitoring procedure vided to alert the crew to unsafe sys- consistent with safety in the kinds of tem operating conditions and to enable operations authorized. Loads not re- them to take appropriate corrective quired in controlled flight need not be action. Systems, controls, and associ- considered for the two-engine-inoper- ated monitoring and warning means ative condition on airplanes with three must be designed to minimize crew er- or more engines. rors that could create additional haz- (e) In showing compliance with this ards. section with regard to the electrical (4) Compliance with the requirements power system and to equipment design of paragraph (b)(2) of this section may and installation, critical environ- be shown by analysis and, where nec- mental and atmospheric conditions, in- essary, by appropriate ground, flight, cluding radio frequency energy and the or simulator tests. The analysis must effects (both direct and indirect) of consider— lightning strikes, must be considered. (i) Possible modes of failure, includ- For electrical generation, distribution, ing malfunctions and damage from ex- and utilization equipment required by ternal sources; or used in complying with this chapter, (ii) The probability of multiple fail- the ability to provide continuous, safe ures, and the probability of undetected service under forseeable environmental faults.; conditions may be shown by environ- (iii) The resulting effects on the air- mental tests, design analysis, or ref- plane and occupants, considering the erence to previous comparable service stage of flight and operating condi- experience on other airplanes. tions; and (f) As used in this section, ‘‘system’’ (iv) The crew warning cues, correc- refers to all pneumatic systems, fluid tive action required, and the crew’s ca- systems, electrical systems, mechan- pability of determining faults. ical systems, and powerplant systems (c) Each item of equipment, each sys- included in the airplane design, except tem, and each installation whose func- for the following: tioning is required by this chapter and that requires a power supply is an ‘‘es- (1) Powerplant systems provided as sential load’’ on the power supply. The part of the certificated engine. power sources and the system must be (2) The flight structure (such a wing, able to supply the following power empennage, control surfaces and their loads in probable operating combina- systems, the fuselage, engine mount- tions and for probable durations: ing, and landing gear and their related

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primary attachments) whose require- (7) Incorporate visual displays of in- ments are specific in subparts C and D strument markings, required by of this part. §§ 23.1541 through 23.1553, or visual dis- plays that alert the pilot to abnormal [Amdt. 23–41, 55 FR 43309, Oct. 26, 1990; 55 FR operational values or approaches to es- 47028, Nov. 8, 1990, as amended by Amdt. 23– 49, 61 FR 5168, Feb. 9, 1996] tablished limitation values, for each parameter required to be displayed by INSTRUMENTS: INSTALLATION this part. (b) The electronic display indicators, § 23.1311 Electronic display instru- including their systems and installa- ment systems. tions, and considering other airplane (a) Electronic display indicators, in- systems, must be designed so that one cluding those with features that make display of information essential for isolation and independence between continued safe flight and landing will powerplant instrument systems im- remain available to the crew, without practical, must: need for immediate action by any pilot for continued safe operation, after any (1) Meet the arrangement and visi- single failure or probable combination bility requirements of § 23.1321. of failures. (2) Be easily legible under all lighting (c) As used in this section, ‘‘instru- conditions encountered in the cockpit, ment’’ includes devices that are phys- including direct sunlight, considering ically contained in one unit, and de- the expected electronic display bright- vices that are composed of two or more ness level at the end of an electronic physically separate units or compo- display indictor’s useful life. Specific nents connected together (such as a re- limitations on display system useful mote indicating gyroscopic direction life must be contained in the Instruc- indicator that includes a magnetic tions for Continued Airworthiness re- sensing element, a gyroscopic unit, an quired by § 23.1529. amplifier, and an indicator connected (3) Not inhibit the primary display of together). As used in this section, ‘‘pri- attitude, airspeed, altitude, or power- mary’’ display refers to the display of a plant parameters needed by any pilot parameter that is located in the instru- to set power within established limita- ment panel such that the pilot looks at tions, in any normal mode of oper- it first when wanting to view that pa- ation. rameter. (4) Not inhibit the primary display of engine parameters needed by any pilot [Doc. No. 27806, 61 FR 5168, Feb. 9, 1996] to properly set or monitor powerplant limitations during the engine starting § 23.1321 Arrangement and visibility. mode of operation. (a) Each flight, navigation, and pow- (5) Have an independent magnetic di- erplant instrument for use by any re- rection indicator and either an inde- quired pilot during takeoff, initial pendent secondary mechanical altim- climb, final approach, and landing eter, airspeed indicator, and attitude must be located so that any pilot seat- instrument or individual electronic ed at the controls can monitor the air- display indicators for the altitude, air- plane’s flight path and these instru- speed, and attitude that are inde- ments with minimum head and eye pendent from the airplane’s primary movement. The powerplant instru- electrical power system. These sec- ments for these flight conditions are ondary instruments may be installed in those needed to set power within pow- panel positions that are displaced from erplant limitations. the primary positions specified by (b) For each multiengine airplane, § 23.1321(d), but must be located where identical powerplant instruments must they meet the pilot’s visibility require- be located so as to prevent confusion as ments of § 23.1321(a). to which engine each instrument re- (6) Incorporate sensory cues for the lates. pilot that are equivalent to those in (c) Instrument panel vibration may the instrument being replaced by the not damage, or impair the accuracy of, electronic display indicators. any instrument.

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(d) For each airplane, the flight in- the color differs sufficiently from the struments required by § 23.1303, and, as colors prescribed in paragraphs (a) applicable, by the operating rules of through (c) of this section to avoid pos- this chapter, must be grouped on the sible confusion. instrument panel and centered as near- (e) Effective under all probable cock- ly as practicable about the vertical pit lighting conditions. plane of each required pilot’s forward vision. In addition: [Amdt. 23–17, 41 FR 55465, Dec. 20, 1976, as (1) The instrument that most effec- amended by Amdt. 23–43, 58 FR 18976, Apr. 9, tively indicates the attitude must be 1993] on the panel in the top center position; § 23.1323 Airspeed indicating system. (2) The instrument that most effec- tively indicates airspeed must be adja- (a) Each airspeed indicating instru- cent to and directly to the left of the ment must be calibrated to indicate instrument in the top center position; true airspeed (at sea level with a stand- (3) The instrument that most effec- ard atmosphere) with a minimum prac- tively indicates altitude must be adja- ticable instrument calibration error cent to and directly to the right of the when the corresponding pitot and stat- instrument in the top center position; ic pressures are applied. (4) The instrument that most effec- (b) Each airspeed system must be tively indicates direction of flight, calibrated in flight to determine the other than the magnetic direction indi- system error. The system error, includ- cator required by § 23.1303(c), must be ing position error, but excluding the adjacent to and directly below the in- airspeed indicator instrument calibra- strument in the top center position; tion error, may not exceed three per- and cent of the calibrated airspeed or five (5) Electronic display indicators may knots, whichever is greater, through- be used for compliance with paragraphs out the following speed ranges: (d)(1) through (d)(4) of this section (1) 1.3 VS1 to VMO/MMO or VNE, which- when such displays comply with re- ever is appropriate with flaps re- quirements in § 23.1311. tracted. (e) If a visual indicator is provided to (2) 1.3 VS1 to VFE with flaps extended. indicate malfunction of an instrument, (c) The design and installation of it must be effective under all probable each airspeed indicating system must cockpit lighting conditions. provide positive drainage of moisture [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as from the pitot static plumbing. amended by Amdt. 23–14, 38 FR 31824, Nov. 19, (d) If certification for instrument 1973; Amdt. 23–20, 42 FR 36968, July 18, 1977; flight rules or flight in icing conditions Amdt. 23–41, 55 FR 43310, Oct. 26, 1990; 55 FR is requested, each airspeed system 46888, Nov. 7, 1990; Amdt. 23–49, 61 FR 5168, Feb. 9, 1996] must have a heated pitot tube or an equivalent means of preventing mal- § 23.1322 Warning, caution, and advi- function due to icing. sory lights. (e) In addition, for commuter cat- If warning, caution, or advisory egory airplanes, the airspeed indi- lights are installed in the cockpit, they cating system must be calibrated to de- must, unless otherwise approved by the termine the system error during the Administrator, be— accelerate-takeoff ground run. The (a) Red, for warning lights (lights in- ground run calibration must be ob- dicating a hazard which may require tained between 0.8 of the minimum immediate corrective action); value of V1, and 1.2 times the maximum (b) Amber, for caution lights (lights value of V1 considering the approved indicating the possible need for future ranges of altitude and weight. The corrective action); ground run calibration must be deter- (c) Green, for safe operation lights; mined assuming an engine failure at and the minimum value of V1. (d) Any other color, including white, (f) For commuter category airplanes, for lights not described in paragraphs where duplicate airspeed indicators are (a) through (c) of this section, provided required, their respective pitot tubes

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must be far enough apart to avoid dam- the maximum cabin differential pres- age to both tubes in a collision with a sure or 100 feet, whichever is greater. bird. (3) If a static pressure system is pro- [Amdt. 23–20, 42 FR 36968, July 18, 1977, as vided for any instrument, device, or amended by Amdt. 23–34, 52 FR 1834, Jan. 15, system required by the operating rules 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23–42, of this chapter, each static pressure 56 FR 354, Jan. 3, 1991; Amdt. 23–49, 61 FR port must be designed or located in 5168, Feb. 9, 1996] such a manner that the correlation be- tween air pressure in the static pres- § 23.1325 Static pressure system. sure system and true ambient atmos- (a) Each instrument provided with pheric static pressure is not altered static pressure case connections must when the airplane encounters icing be so vented that the influence of air- conditions. An antiicing means or an plane speed, the opening and closing of alternate source of static pressure may windows, airflow variations, moisture, be used in showing compliance with or other foreign matter will least af- this requirement. If the reading of the fect the accuracy of the instruments altimeter, when on the alternate static except as noted in paragraph (b)(3) of pressure system differs from the read- this section. (b) If a static pressure system is nec- ing of the altimeter when on the pri- essary for the functioning of instru- mary static system by more than 50 ments, systems, or devices, it must feet, a correction card must be pro- comply with the provisions of para- vided for the alternate static system. graphs (b) (1) through (3) of this sec- (c) Except as provided in paragraph tion. (d) of this section, if the static pressure (1) The design and installation of a system incorporates both a primary static pressure system must be such and an alternate static pressure source, that— the means for selecting one or the (i) Positive drainage of moisture is other source must be designed so provided; that— (ii) Chafing of the tubing, and exces- (1) When either source is selected, the sive distortion or restriction at bends other is blocked off; and in the tubing, is avoided; and (2) Both sources cannot be blocked (iii) The materials used are durable, off simultaneously. suitable for the purpose intended, and (d) For unpressurized airplanes, para- protected against corrosion. graph (c)(1) of this section does not (2) A proof test must be conducted to apply if it can be demonstrated that demonstrate the integrity of the static the static pressure system calibration, pressure system in the following man- when either static pressure source is ner: selected, is not changed by the other (i) Unpressurized airplanes. Evacuate static pressure source being open or the static pressure system to a pres- blocked. sure differential of approximately 1 (e) Each static pressure system must inch of mercury or to a reading on the altimeter, 1,000 feet above the aircraft be calibrated in flight to determine the elevation at the time of the test. With- system error. The system error, in in- out additional pumping for a period of dicated pressure altitude, at sea-level, 1 minute, the loss of indicated altitude with a standard atmosphere, excluding must not exceed 100 feet on the altim- instrument calibration error, may not ± eter. exceed 30 feet per 100 knot speed for (ii) Pressurized airplanes. Evacuate the appropriate configuration in the the static pressure system until a pres- speed range between 1.3 VS0 with flaps sure differential equivalent to the max- extended, and 1.8 VS1 with flaps re- imum cabin pressure differential for tracted. However, the error need not be which the airplane is type certificated less than 30 feet. is achieved. Without additional pump- (f) [Reserved] ing for a period of 1 minute, the loss of (g) For airplanes prohibited from indicated altitude must not exceed 2 flight in instrument meteorological or percent of the equivalent altitude of icing conditions, in accordance with

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§ 23.1559(b) of this part, paragraph (b)(3) grees must be placarded in accordance of this section does not apply. with § 23.1547(e). [Amdt. 23–1, 30 FR 8261, June 29, 1965, as [Amdt. 23–20, 42 FR 36969, July 18, 1977] amended by Amdt. 23–6, 32 FR 7586, May 24, 1967; 32 FR 13505, Sept. 27, 1967; 32 FR 13714, § 23.1329 Automatic pilot system. Sept. 30, 1967; Amdt. 23–20, 42 FR 36968, July If an automatic pilot system is in- 18, 1977; Amdt. 23–34, 52 FR 1834, Jan. 15, 1987; stalled, it must meet the following: Amdt. 23–42, 56 FR 354, Jan. 3, 1991; Amdt. 23– (a) Each system must be designed so 49, 61 FR 5169, Feb. 9, 1996; Amdt. 23–50, 61 FR that the automatic pilot can— 5192, Feb. 9, 1996] (1) Be quickly and positively dis- § 23.1326 Pitot heat indication systems. engaged by the pilots to prevent it from interfering with their control of If a flight instrument pitot heating the airplane; or system is installed to meet the require- (2) Be sufficiently overpowered by ments specified in § 23.1323(d), an indi- one pilot to let him control the air- cation system must be provided to in- plane. dicate to the flight crew when that (b) If the provisions of paragraph pitot heating system is not operating. (a)(1) of this section are applied, the The indication system must comply quick release (emergency) control with the following requirements: must be located on the control wheel (a) The indication provided must in- (both control wheels if the airplane can corporate an amber light that is in be operated from either pilot seat) on clear view of a flightcrew member. the side opposite the throttles, or on (b) The indication provided must be the stick control, (both stick controls, designed to alert the flight crew if ei- if the airplane can be operated from ei- ther of the following conditions exist: ther pilot seat) such that it can be op- (1) The pitot heating system is erated without moving the hand from its normal position on the control. switched ‘‘off.’’ (c) Unless there is automatic syn- (2) The pitot heating system is chronization, each system must have a switched on and any pitot tube heat- ‘‘ ’’ means to readily indicate to the pilot ing element is inoperative. the alignment of the actuating device [Doc. No. 27806, 61 FR 5169, Feb. 9, 1996] in relation to the control system it op- erates. § 23.1327 Magnetic direction indicator. (d) Each manually operated control (a) Except as provided in paragraph for the system operation must be read- ily accessible to the pilot. Each control (b) of this section— must operate in the same plane and (1) Each magnetic direction indicator sense of motion as specified in § 23.779 must be installed so that its accuracy for cockpit controls. The direction of is not excessively affected by the air- motion must be plainly indicated on or plane’s vibration or magnetic fields; near each control. and (e) Each system must be designed and (2) The compensated installation may adjusted so that, within the range of not have a deviation in level flight, adjustment available to the pilot, it greater than ten degrees on any head- cannot produce hazardous loads on the ing. airplane or create hazardous deviations (b) A magnetic nonstabilized direc- in the flight path, under any flight con- tion indicator may deviate more than dition appropriate to its use, either ten degrees due to the operation of during normal operation or in the electrically powered systems such as event of a malfunction, assuming that electrically heated windshields if ei- corrective action begins within a rea- ther a magnetic stabilized direction in- sonable period of time. dicator, which does not have a devi- (f) Each system must be designed so ation in level flight greater than ten that a single malfunction will not degrees on any heading, or a gyroscopic produce a hardover signal in more than direction indicator, is installed. Devi- one control axis. If the automatic pilot ations of a magnetic nonstabilized di- integrates signals from auxiliary con- rection indicator of more than 10 de- trols or furnishes signals for operation

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of other equipment, positive interlocks dicate to the flight crew its current and sequencing of engagement to pre- mode of operation. Selector switch po- vent improper operation are required. sition is not acceptable as a means of (g) There must be protection against indication. adverse interaction of integrated com- ponents, resulting from a malfunction. [Amdt. 23–20, 42 FR 36969, July 18, 1977] (h) If the automatic pilot system can § 23.1337 Powerplant instruments in- be coupled to airborne navigation stallation. equipment, means must be provided to indicate to the flight crew the current (a) Instruments and instrument lines. mode of operation. Selector switch po- (1) Each powerplant and auxiliary sition is not acceptable as a means of power unit instrument line must meet indication. the requirements of § 23.993. (2) Each line carrying flammable [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 fluids under pressure must— FR 258, Jan. 9, 1965, as amended by Amdt. 23– 23, 43 FR 50593, Oct. 30, 1978; Amdt. 23–43, 58 (i) Have restricting orifices or other FR 18976, Apr. 9, 1993; Amdt. 23–49, 61 FR 5169, safety devices at the source of pressure Feb. 9, 1996] to prevent the escape of excessive fluid if the line fails; and § 23.1331 Instruments using a power (ii) Be installed and located so that source. the escape of fluids would not create a For each instrument that uses a hazard. power source, the following apply: (3) Each powerplant and auxiliary (a) Each instrument must have an in- power unit instrument that utilizes tegral visual power annunciator or sep- flammable fluids must be installed and arate power indicator to indicate when located so that the escape of fluid power is not adequate to sustain proper would not create a hazard. instrument performance. If a separate (b) Fuel quantity indication. There indicator is used, it must be located so must be a means to indicate to the that the pilot using the instruments flightcrew members the quantity of us- can monitor the indicator with min- able fuel in each tank during flight. An imum head and eye movement. The indicator calibrated in appropriate power must be sensed at or near the units and clearly marked to indicate point where it enters the instrument. those units must be used. In addition: For electric and vacuum/pressure in- (1) Each fuel quantity indicator must struments, the power is considered to be calibrated to read ‘‘zero’’ during level be adequate when the voltage or the flight when the quantity of fuel re- vacuum/pressure, respectively, is with- maining in the tank is equal to the un- in approved limits. usable fuel supply determined under (b) The installation and power supply § 23.959(a); systems must be designed so that— (2) Each exposed sight gauge used as (1) The failure of one instrument will a fuel quantity indicator must be pro- not interfere with the proper supply of tected against damage; energy to the remaining instrument; (3) Each sight gauge that forms a and trap in which water can collect and (2) The failure of the energy supply freeze must have means to allow drain- from one source will not interfere with age on the ground; the proper supply of energy from any (4) There must be a means to indicate other source. the amount of usable fuel in each tank (c) There must be at least two inde- when the airplane is on the ground pendent sources of power (not driven (such as by a stick gauge); by the same engine on multiengine air- (5) Tanks with interconnected outlets planes), and a manual or an automatic and airspaces may be considered as one means to select each power source. tank and need not have separate indi- [Doc. No. 26344, 58 FR 18976, Apr. 9, 1993] cators; and (6) No fuel quantity indicator is re- § 23.1335 Flight director systems. quired for an auxiliary tank that is If a flight director system is in- used only to transfer fuel to other stalled, means must be provided to in- tanks if the relative size of the tank,

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the rate of fuel transfer, and operating (i) Free from hazards in itself, in its instructions are adequate to— method of operation, and in its effects (i) Guard against overflow; and on other parts of the airplane; (ii) Give the flight crewmembers (ii) Protected from fuel, oil, water, prompt warning if transfer is not pro- other detrimental substances, and me- ceeding as planned. chanical damage; and (c) Fuel flowmeter system. If a fuel (iii) So designed that the risk of elec- flowmeter system is installed, each trical shock to crew, passengers, and metering component must have a ground personnel is reduced to a min- means to by-pass the fuel supply if imum. malfunctioning of that component se- (2) Electric power sources must func- verely restricts fuel flow. tion properly when connected in com- (d) Oil quantity indicator. There must bination or independently. be a means to indicate the quantity of (3) No failure or malfunction of any oil in each tank— (1) On the ground (such as by a stick electric power source may impair the gauge); and ability of any remaining source to sup- (2) In flight, to the flight crew mem- ply load circuits essential for safe oper- bers, if there is an oil transfer system ation. or a reserve oil supply system. (4) In addition, for commuter cat- egory airplanes, the following apply: [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as (i) Each system must be designed so amended by Amdt. 23–7, 34 FR 13096, Aug. 13, 1969; Amdt. 23–18, 42 FR 15042, Mar. 17, 1977; that essential load circuits can be sup- Amdt. 23–43, 58 FR 18976, Apr. 9, 1993; Amdt. plied in the event of reasonably prob- 23–51, 61 FR 5138, Feb. 9, 1996; Amdt. 23–49, 61 able faults or open circuits including FR 5169, Feb. 9, 1996] faults in heavy current carrying cables; (ii) A means must be accessible in ELECTRICAL SYSTEMS AND EQUIPMENT flight to the flight crewmembers for the individual and collective dis- § 23.1351 General. connection of the electrical power (a) Electrical system capacity. Each sources from the system; electrical system must be adequate for (iii) The system must be designed so the intended use. In addition— that voltage and frequency, if applica- (1) Electric power sources, their ble, at the terminals of all essential transmission cables, and their associ- load equipment can be maintained ated control and protective devices, within the limits for which the equip- must be able to furnish the required ment is designed during any probable power at the proper voltage to each operating conditions; load circuit essential for safe oper- (iv) If two independent sources of ation; and electrical power for particular equip- (2) Compliance with paragraph (a)(1) ment or systems are required, their of this section must be shown as fol- lows— electrical energy supply must be en- (i) For normal, utility, and acrobatic sured by means such as duplicate elec- category airplanes, by an electrical trical equipment, throwover switching, load analysis or by electrical measure- or multichannel or loop circuits sepa- ments that account for the electrical rately routed; and loads applied to the electrical system (v) For the purpose of complying in probable combinations and for prob- with paragraph (b)(5) of this section, able durations; and the distribution system includes the (ii) For commuter category air- distribution busses, their associated planes, by an electrical load analysis feeders, and each control and protec- that accounts for the electrical loads tive device. applied to the electrical system in (c) Generating system. There must be probable combinations and for probable at least one generator/alternator if the durations. electrical system supplies power to (b) Function. For each electrical sys- load circuits essential for safe oper- tem, the following apply: ation. In addition— (1) Each system, when installed, (1) Each generator/alternator must be must be— able to deliver its continuous rated

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power, or such power as is limited by having a reverse polarity, or a reverse its regulation system. phase sequence, can supply power to (2) Generator/alternator voltage con- the airplane’s electrical system. The trol equipment must be able to depend- external power connection must be lo- ably regulate the generator/alternator cated so that its use will not result in output within rated limits. a hazard to the airplane or ground per- (3) Automatic means must be pro- sonnel. vided to prevent damage to any gener- (g) It must be shown by analysis, ator/alternator and adverse effects on tests, or both, that the airplane can be the airplane electrical system due to operated safely in VFR conditions, for reverse current. A means must also be a period of not less than five minutes, provided to disconnect each generator/ with the normal electrical power (elec- alternator from the battery and other trical power sources excluding the bat- generators/alternators. tery and any other standby electrical (4) There must be a means to give im- sources) inoperative, with critical type mediate warning to the flight crew of a fuel (from the standpoint of flameout failure of any generator/alternator. and restart capability), and with the (5) Each generator/alternator must airplane initially at the maximum cer- have an overvoltage control designed tificated altitude. Parts of the elec- and installed to prevent damage to the trical system may remain on if— electrical system, or to equipment sup- (1) A single malfunction, including a plied by the electrical system that wire bundle or junction box fire, can- could result if that generator/alter- not result in loss of the part turned off nator were to develop an overvoltage and the part turned on; and condition. (2) The parts turned on are elec- (d) Instruments. A means must exist trically and mechanically isolated to indicate to appropriate flight crew- from the parts turned off. members the electric power system [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as quantities essential for safe operation. amended by Amdt. 23–7, 34 FR 13096, Aug. 13, (1) For normal, utility, and acrobatic 1969; Amdt. 23–14, 38 FR 31824, Nov. 19, 1973; category airplanes with direct current Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. systems, an ammeter that can be 23–20, 42 FR 36969, July 18, 1977; Amdt. 23–34, 52 FR 1834, Jan. 15, 1987; 52 FR 34745, Sept. 14, switched into each generator feeder 1987; Amdt. 23–43, 58 FR 18976, Apr. 9, 1993; may be used and, if only one generator Amdt. 23–49, 61 FR 5169, Feb. 9, 1996] exists, the ammeter may be in the bat- tery feeder. § 23.1353 Storage battery design and (2) For commuter category airplanes, installation. the essential electric power system (a) Each storage battery must be de- quantities include the voltage and cur- signed and installed as prescribed in rent supplied by each generator. this section. (e) Fire resistance. Electrical equip- (b) Safe cell temperatures and pres- ment must be so designed and installed sures must be maintained during any that in the event of a fire in the engine probable charging and discharging con- compartment, during which the surface dition. No uncontrolled increase in cell of the firewall adjacent to the fire is temperature may result when the bat- heated to 2,000 °F for 5 minutes or to a tery is recharged (after previous com- lesser temperature substantiated by plete discharge)— the applicant, the equipment essential (1) At maximum regulated voltage or to continued safe operation and located power; behind the firewall will function satis- (2) During a flight of maximum dura- factorily and will not create an addi- tion; and tional fire hazard. (3) Under the most adverse cooling (f) External power. If provisions are condition likely to occur in service. made for connecting external power to (c) Compliance with paragraph (b) of the airplane, and that external power this section must be shown by tests un- can be electrically connected to equip- less experience with similar batteries ment other than that used for engine and installations has shown that main- starting, means must be provided to taining safe cell temperatures and ensure that no external power supply pressures presents no problem.

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(d) No explosive or toxic gases emit- (1) Main circuits of starter motors ted by any battery in normal oper- used during starting only; and ation, or as the result of any probable (2) Circuits in which no hazard is pre- malfunction in the charging system or sented by their omission. battery installation, may accumulate (b) A protective device for a circuit in hazardous quantities within the air- essential to flight safety may not be plane. used to protect any other circuit. (e) No corrosive fluids or gases that (c) Each resettable circuit protective may escape from the battery may dam- device (‘‘trip free’’ device in which the age surrounding structures or adjacent tripping mechanism cannot be over- essential equipment. ridden by the operating control) must (f) Each nickel cadmium battery in- be designed so that— stallation capable of being used to start an engine or auxiliary power unit (1) A manual operation is required to must have provisions to prevent any restore service after tripping; and hazardous effect on structure or essen- (2) If an overload or circuit fault ex- tial systems that may be caused by the ists, the device will open the circuit re- maximum amount of heat the battery gardless of the position of the oper- can generate during a short circuit of ating control. the battery or of its individual cells. (d) If the ability to reset a circuit (g) Nickel cadmium battery installa- breaker or replace a fuse is essential to tions capable of being used to start an safety in flight, that circuit breaker or engine or auxiliary power unit must fuse must be so located and identified have— that it can be readily reset or replaced (1) A system to control the charging in flight. rate of the battery automatically so as (e) For fuses identified as replaceable to prevent battery overheating; in flight— (2) A battery temperature sensing (1) There must be one spare of each and over-temperature warning system rating or 50 percent spare fuses of each with a means for disconnecting the rating, whichever is greater; and battery from its charging source in the (2) The spare fuse(s) must be readily event of an over-temperature condi- accessible to any required pilot. tion; or (3) A battery failure sensing and [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 warning system with a means for dis- FR 258, Jan. 9, 1965, as amended by Amdt. 23– connecting the battery from its charg- 20, 42 FR 36969, July 18, 1977]; Amdt. 23–43, 58 ing source in the event of battery fail- FR 18976, Apr. 9, 1993 ure. § 23.1359 Electrical system fire protec- (h) In the event of a complete loss of tion. the primary electrical power gener- ating system, the battery must be ca- (a) Each component of the electrical pable of providing at least 30 minutes system must meet the applicable fire of electrical power to those loads that protection requirements of §§ 23.863 and are essential to continued safe flight 23.1182. and landing. The 30 minute time period (b) Electrical cables, terminals, and includes the time needed for the pilots equipment in designated fire zones that to recognize the loss of generated are used during emergency procedures power and take appropriate load shed- must be fire-resistant. ding action. (c) Insulation on electrical wire and electrical cable must be self-extin- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23– guishing when tested at an angle of 60 20, 42 FR 36969, July 18, 1977; Amdt. 23–21, 43 degrees in accordance with the applica- FR 2319, Jan. 16, 1978; Amdt. 23–49, 61 FR 5169, ble portions of appendix F of this part, Feb. 9, 1996] or other approved equivalent methods. The average burn length must not ex- § 23.1357 Circuit protective devices. ceed 3 inches (76 mm) and the average (a) Protective devices, such as fuses flame time after removal of the flame or circuit breakers, must be installed source must not exceed 30 seconds. in all electrical circuits other than— Drippings from the test specimen must

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not continue to flame for more than an (c) Main power cables (including gen- average of 3 seconds after falling. erator cables) in the fuselage must be designed to allow a reasonable degree [Doc. No. 27806, 61 FR 5169, Feb. 9, 1996] of deformation and stretching without § 23.1361 Master switch arrangement. failure and must— (1) Be separated from flammable fluid (a) There must be a master switch ar- lines; or rangement to allow ready disconnec- (2) Be shrouded by means of elec- tion of each electric power source from trically insulated flexible conduit, or power distribution systems, except as equivalent, which is in addition to the provided in paragraph (b) of this sec- normal cable insulation. tion. The point of disconnection must (d) Means of identification must be be adjacent to the sources controlled provided for electrical cables, termi- by the switch arrangement. If separate nals, and connectors. switches are incorporated into the (e) Electrical cables must be in- master switch arrangement, a means stalled such that the risk of mechan- must be provided for the switch ar- ical damage and/or damage cased by rangement to be operated by one hand fluids vapors, or sources of heat, is with a single movement. minimized. (b) Load circuits may be connected so (f) Where a cable cannot be protected that they remain energized when the by a circuit protection device or other master switch is open, if the circuits overload protection, it must not cause are isolated, or physically shielded, to a fire hazard under fault conditions. prevent their igniting flammable fluids or vapors that might be liberated by [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as the leakage or rupture of any flam- amended by Amdt. 23–14, 38 FR 31824, Nov. 19, mable fluid system; and 1973; Amdt. 23–43, 58 FR 18977, Apr. 9, 1993; (1) The circuits are required for con- Amdt. 23–49, 61 FR 5169, Feb. 9, 1996] tinued operation of the engine; or § 23.1367 Switches. (2) The circuits are protected by cir- cuit protective devices with a rating of Each switch must be— five amperes or less adjacent to the (a) Able to carry its rated current; electric power source. (b) Constructed with enough distance (3) In addition, two or more circuits or insulating material between current installed in accordance with the re- carrying parts and the housing so that quirements of paragraph (b)(2) of this vibration in flight will not cause short- section must not be used to supply a ing; load of more than five amperes. (c) Accessible to appropriate flight (c) The master switch or its controls crewmembers; and must be so installed that the switch is (d) Labeled as to operation and the easily discernible and accessible to a circuit controlled. crewmember. LIGHTS [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23– § 23.1381 Instrument lights. 20, 42 FR 36969, July 18, 1977; Amdt. 23–43, 58 FR 18977, Apr. 9, 1993; Amdt. 23–49, 61 FR 5169, The instrument lights must— Feb. 9, 1996] (a) Make each instrument and con- trol easily readable and discernible; § 23.1365 Electric cables and equip- (b) Be installed so that their direct ment. rays, and rays reflected from the wind- (a) Each electric connecting cable shield or other surface, are shielded must be of adequate capacity. from the pilot’s eyes; and (b) Any equipment that is associated (c) Have enough distance or insu- with any electrical cable installation lating material between current car- and that would overheat in the event of rying parts and the housing so that vi- circuit overload or fault must be flame bration in flight will not cause short- resistant. That equipment and the elec- ing. trical cables must not emit dangerous A cabin dome light is not an instru- quantities of toxic fumes. ment light.

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§ 23.1383 Taxi and landing lights. (c) Dihedral angle R (right) is formed Each taxi and landing light must be by two intersecting vertical planes, the designed and installed so that: first parallel to the longitudinal axis of (a) No dangerous glare is visible to the airplane, and the other at 110 de- the pilots. grees to the right of the first, as viewed (b) The pilot is not seriously affected when looking forward along the longi- by halation. tudinal axis. (c) It provides enough light for night (d) Dihedral angle A (aft) is formed operations. by two intersecting vertical planes (d) It does not cause a fire hazard in making angles of 70 degrees to the any configuration. right and to the left, respectively, to a vertical plane passing through the lon- [Doc. No. 27806, 61 FR 5169, Feb. 9, 1996] gitudinal axis, as viewed when looking aft along the longitudinal axis. § 23.1385 Position light system installa- (e) If the rear position light, when tion. mounted as far aft as practicable in ac- (a) General. Each part of each posi- cordance with § 23.1385(c), cannot show tion light system must meet the appli- unbroken light within dihedral angle A cable requirements of this section and (as defined in paragraph (d) of this sec- each system as a whole must meet the tion), a solid angle or angles of ob- requirements of §§ 23.1387 through structed visibility totaling not more 23.1397. than 0.04 steradians is allowable within (b) Left and right position lights. Left that dihedral angle, if such solid angle and right position lights must consist is within a cone whose apex is at the of a red and a green light spaced lat- rear position light and whose elements erally as far apart as practicable and make an angle of 30° with a vertical installed on the airplane such that, line passing through the rear position with the airplane in the normal flying light. position, the red light is on the left side and the green light is on the right [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23– side. 12, 36 FR 21278, Nov. 5, 1971; Amdt. 23–43, 58 (c) Rear position light. The rear posi- FR 18977, Apr. 9, 1993] tion light must be a white light mount- ed as far aft as practicable on the tail § 23.1389 Position light distribution or on each wing tip. and intensities. (d) Light covers and color filters. Each (a) General. The intensities prescribed light cover or color filter must be at in this section must be provided by new least flame resistant and may not equipment with each light cover and change color or shape or lose any ap- color filter in place. Intensities must preciable light transmission during be determined with the light source op- normal use. erating at a steady value equal to the [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as average luminous output of the source amended by Amdt. 23–17, 41 FR 55465, Dec. 20, at the normal operating voltage of the 1976; Amdt. 23–43, 58 FR 18977, Apr. 9, 1993] airplane. The light distribution and in- tensity of each position light must § 23.1387 Position light system dihe- meet the requirements of paragraph (b) dral angles. of this section. (a) Except as provided in paragraph (b) Position lights. The light distribu- (e) of this section, each position light tion and intensities of position lights must, as installed, show unbroken light must be expressed in terms of min- within the dihedral angles described in imum intensities in the horizontal 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

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axis of the airplane and perpendicular Angle from right or left of longitu- to the plane of symmetry of the air- Dihedral angle (light in- dinal axis, meas- Intensity plane) must equal or exceed the values cluded) ured from dead (candles) in § 23.1391. ahead (2) Intensities in any vertical plane. L and R (red and green) .... 0° to 10° ...... 40 Each intensity in any vertical plane 10° to 20° ...... 30 20° to 110° ...... 5 (the plane perpendicular to the hori- A (rear white) ...... 110° to 180° ...... 20 zontal plane) must equal or exceed the appropriate value in § 23.1393, where I is [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as the minimum intensity prescribed in amended by Amdt. 23–43, 58 FR 18977, Apr. 9, § 23.1391 for the corresponding angles in 1993] the horizontal plane. (3) Intensities in overlaps between adja- § 23.1393 Minimum intensities in any cent signals. No intensity in any over- vertical plane of position lights. lap between adjacent signals may ex- Each position light intensity must ceed the values in § 23.1395, except that equal or exceed the applicable values in higher intensities in overlaps may be the following table: used with main beam intensities sub- stantially greater than the minima Angle above or below the horizontal plane Intensity, l specified in §§ 23.1391 and 23.1393, if the 0° ...... 1.00 overlap intensities in relation to the 0° to 5° ...... 0.90 main beam intensities do not adversely 5° to 10° ...... 0.80 ° ° affect signal clarity. When the peak in- 10 to 15 ...... 0.70 15° to 20° ...... 0.50 tensity of the left and right position 20° to 30° ...... 0.30 lights is more than 100 candles, the 30° to 40° ...... 0.10 maximum overlap intensities between 40° to 90° ...... 0.05 them may exceed the values in § 23.1395 if the overlap intensity in Area A is [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as not more than 10 percent of peak posi- amended by Amdt. 23–43, 58 FR 18977, Apr. 9, tion light intensity and the overlap in- 1993] tensity in Area B is not more than 2.5 § 23.1395 Maximum intensities in over- percent of peak position light inten- lapping beams of position lights. sity. (c) Rear position light installation. A No position light intensity may ex- ceed the applicable values in the fol- single rear position light may be in- lowing equal or exceed the applicable stalled in a position displaced laterally values in § 23.1389(b)(3): from the plane of symmetry of an air- plane if— Maximum intensity (1) The axis of the maximum cone of Overlaps Area A Area B illumination is parallel to the flight (candles) (candles) path in level flight; and Green in dihedral angle L ...... 10 1 (2) There is no obstruction aft of the Red in dihedral angle R ...... 10 1 light and between planes 70 degrees to Green in dihedral angle A ...... 5 1 the right and left of the axis of max- Red in dihedral angle A ...... 5 1 imum illumination. Rear white in dihedral angle L ...... 5 1 Rear white in dihedral angle R ..... 5 1 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18977, Apr. 9, Where— 1993] (a) Area A includes all directions in the adjacent dihedral angle that pass § 23.1391 Minimum intensities in the through the light source and intersect horizontal plane of position lights. the common boundary plane at more Each position light intensity must than 10 degrees but less than 20 de- equal or exceed the applicable values in grees; and the following table: (b) Area B includes all directions in the adjacent dihedral angle that pass through the light source and intersect

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the common boundary plane at more considering the physical configuration than 20 degrees. and flight characteristics of the air- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as plane. The field of coverage must ex- amended by Amdt. 23–43, 58 FR 18977, Apr. 9, tend in each direction within at least 1993] 75 degrees above and 75 degrees below the horizontal plane of the airplane, § 23.1397 Color specifications. except that there may be solid angles Each position light color must have of obstructed visibility totaling not the applicable International Commis- more than 0.5 steradians. sion on Illumination chromaticity co- (c) Flashing characteristics. The ar- ordinates as follows: rangement of the system, that is, the (a) Aviation red— number of light sources, beam width, y is not greater than 0.335; and speed of rotation, and other character- z is not greater than 0.002. istics, must give an effective flash fre- quency of not less than 40, nor more (b) Aviation green— than 100, cycles per minute. The effec- x is not greater than 0.440¥0.320y; tive flash frequency is the frequency at x is not greater than y¥0.170; and which the airplane’s complete anti- ¥ y is not less than 0.390 0.170x. collision light system is observed from (c) Aviation white— a distance, and applies to each sector x is not less than 0.300 and not greater than of light including any overlaps that 0.540; exist when the system consists of more y is not less than x¥0.040 or y0¥0.010, which- than one light source. In overlaps, ever is the smaller; and flash frequencies may exceed 100, but y is not greater than x+0.020 nor 0.636¥0.400x; not 180, cycles per minute. Where y0 is the y coordinate of the Planckian (d) Color. Each anticollision light radiator for the value of x considered. must be either aviation red or aviation [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, white and must meet the applicable re- amended by Amdt. 23–11, 36 FR 12971, July 10, quirements of § 23.1397. 1971] (e) Light intensity. The minimum § 23.1399 Riding light. light intensities in any vertical plane, measured with the red filter (if used) (a) Each riding (anchor) light re- and expressed in terms of ‘‘effective’’ in- quired for a seaplane or amphibian, tensities, must meet the requirements must be installed so that it can— of paragraph (f) of this section. The fol- (1) Show a white light for at least lowing relation must be assumed: two miles at night under clear atmos- pheric conditions; and t2 (2) Show the maximum unbroken ∫ Itdt() light practicable when the airplane is = t1 Ie moored or drifting on the water. 02. +−()tt (b) Externally hung lights may be 21 used. where:

Ie=effective intensity (candles). § 23.1401 Anticollision light system. I(t)=instantaneous intensity as a function of (a) General. The airplane must have time. an anticollision light system that: t2¥t1=flash time interval (seconds). (1) Consists of one or more approved anticollision lights located so that Normally, the maximum value of effec- their light will not impair the flight tive intensity is obtained when t2 and t1 crewmembers’ vision or detract from are chosen so that the effective inten- the conspicuity of the position lights; sity is equal to the instantaneous in- and tensity at t2 and t1. (2) Meets the requirements of para- (f) Minimum effective intensities for graphs (b) through (f) of this section. anticollision lights. Each anticollision (b) Field of coverage. The system must light effective intensity must equal or consist of enough lights to illuminate exceed the applicable values in the fol- the vital areas around the airplane, lowing table.

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Effective in- § 23.1416 Pneumatic de-icer boot sys- Angle above or below the horizontal plane tensity (can- dles) tem. If certification with ice protection ° ° 0 to 5 ...... 400 provisions is desired and a pneumatic 5° to 10° ...... 240 10° to 20° ...... 80 de-icer boot system is installed— 20° to 30° ...... 40 (a) The system must meet the re- 30° to 75° ...... 20 quirements specified in § 23.1419. (b) The system and its components [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as must be designed to perform their in- amended by Amdt. 23–11, 36 FR 12972, July 10, tended function under any normal sys- 1971; Amdt. 23–20, 42 FR 36969, July 18, 1977; tem operating temperature or pressure, Amdt. 23–49, 61 FR 5169, Feb. 9, 1996] and (c) Means to indicate to the flight SAFETY EQUIPMENT crew that the pneumatic de-icer boot system is receiving adequate pressure § 23.1411 General. and is functioning normally must be (a) Required safety equipment to be provided. used by the flight crew in an emer- [Amdt. 23–23, 43 FR 50593, Oct. 30, 1978] gency, such as automatic liferaft re- leases, must be readily accessible. § 23.1419 Ice protection. (b) Stowage provisions for required If certification with ice protection safety equipment must be furnished provisions is desired, compliance with and must— the requirements of this section and (1) Be arranged so that the equip- other applicable sections of this part ment is directly accessible and its loca- must be shown: tion is obvious; and (a) An analysis must be performed to (2) Protect the safety equipment establish, on the basis of the airplane’s from damage caused by being subjected operational needs, the adequacy of the to the inertia loads resulting from the ice protection system for the various ultimate static load factors specified in components of the airplane. In addi- tion, tests of the ice protection system § 23.561(b)(3) of this part. must be conducted to demonstrate that [Amdt. 23–17, 41 FR 55465, Dec. 20, 1976, as the airplane is capable of operating amended by Amdt. 23–36, 53 FR 30815, Aug. 15, safely in continuous maximum and 1988] intermittent maximum icing condi- tions, as described in appendix C of § 23.1415 Ditching equipment. part 25 of this chapter. As used in this (a) Emergency flotation and sig- section, ‘‘Capable of operating safely,’’ naling equipment required by any oper- means that airplane performance, con- ating rule in this chapter must be in- trollability, maneuverability, and sta- stalled so that it is readily available to bility must not be less than that re- the crew and passengers. quired in part 23, subpart B. (b) Each raft and each life preserver (b) Except as provided by paragraph must be approved. (c) of this section, in addition to the analysis and physical evaluation pre- (c) Each raft released automatically scribed in paragraph (a) of this section, or by the pilot must be attached to the the effectiveness of the ice protection airplane by a line to keep it alongside system and its components must be the airplane. This line must be weak shown by flight tests of the airplane or enough to break before submerging the its components in measured natural at- empty raft to which it is attached. mospheric icing conditions and by one (d) Each signaling device required by or more of the following tests, as found any operating rule in this chapter, necessary to determine the adequacy of must be accessible, function satisfac- the ice protection system— torily, and must be free of any hazard (1) Laboratory dry air or simulated in its operation. icing tests, or a combination of both, of the components or models of the com- ponents.

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(2) Flight dry air tests of the ice pro- (c) For those airplanes required to tection system as a whole, or its indi- have more than one flightcrew mem- vidual components. ber, or whose operation will require (3) Flight test of the airplane or its more than one flightcrew member, the components in measured simulated cockpit must be evaluated to deter- icing conditions. mine if the flightcrew members, when (c) If certification with ice protection seated at their duty station, can con- has been accomplished on prior type verse without difficulty under the ac- certificated airplanes whose designs in- tual cockpit noise conditions when the clude components that are airplane is being operated. If the air- thermodynamically and aero- plane design includes provision for the dynamically equivalent to those used use of communication headsets, the on a new airplane design, certification evaluation must also consider condi- of these equivalent components may be tions where headsets are being used. If accomplished by reference to pre- the evaluation shows conditions under viously accomplished tests, required in which it will be difficult to converse, § 23.1419 (a) and (b), provided that the an intercommunication system must applicant accounts for any differences be provided. in installation of these components. (d) If installed communication equip- (d) A means must be identified or ment includes transmitter ‘‘off-on’’ provided for determining the formation switching, that switching means must of ice on the critical parts of the air- be designed to return from the ‘‘trans- plane. Adequate lighting must be pro- mit’’ to the ‘‘off’’ position when it is re- vided for the use of this means during leased and ensure that the transmitter night operation. Also, when monitoring will return to the off (non transmit- of the external surfaces of the airplane ting) state. by the flight crew is required for oper- (e) If provisions for the use of com- ation of the ice protection equipment, munication headsets are provided, it external lighting must be provided that must be demonstrated that the is adequate to enable the monitoring to flightcrew members will receive all be done at night. Any illumination aural warnings under the actual cock- that is used must be of a type that will pit noise conditions when the airplane not cause glare or reflection that is being operated when any headset is would handicap crewmembers in the being used. performance of their duties. The Air- [Doc. No. 26344, 58 FR 18977, Apr. 9, 1993, as plane Flight Manual or other approved amended by Amdt. 23–49, 61 FR 5169, Feb. 9, manual material must describe the 1996] means of determining ice formation and must contain information for the § 23.1435 Hydraulic systems. safe operation of the airplane in icing (a) Design. Each hydraulic system conditions. must be designed as follows: [Doc. No. 26344, 58 FR 18977, Apr. 9, 1993] (1) Each hydraulic system and its ele- ments must withstand, without yield- MISCELLANEOUS EQUIPMENT ing, the structural loads expected in addition to hydraulic loads. § 23.1431 Electronic equipment. (2) A means to indicate the pressure (a) In showing compliance with in each hydraulic system which sup- § 23.1309(b)(1) and (2) with respect to plies two or more primary functions radio and electronic equipment and must be provided to the flight crew. their installations, critical environ- (3) There must be means to ensure mental conditions must be considered. that the pressure, including transient (b) Radio and electronic equipment, (surge) pressure, in any part of the sys- controls, and wiring must be installed tem will not exceed the safe limit so that operation of any unit or system above design operating pressure and to of units will not adversely affect the si- prevent excessive pressure resulting multaneous operation of any other from fluid volumetric changes in all radio or electronic unit, or system of lines which are likely to remain closed units, required by this chapter. long enough for such changes to occur.

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(4) The minimum design burst pres- § 23.1441 Oxygen equipment and sup- sure must be 2.5 times the operating ply. pressure. (a) If certification with supplemental (b) Tests. Each system must be sub- oxygen equipment is requested, or the stantiated by proof pressure tests. airplane is approved for operations at When proof tested, no part of any sys- or above altitudes where oxygen is re- tem may fail, malfunction, or experi- quired to be used by the operating ence a permanent set. The proof load of rules, oxygen equipment must be pro- each system must be at least 1.5 times vided that meets the requirements of the maximum operating pressure of this section and §§ 23.1443 through that system. 23.1449. Portable oxygen equipment (c) Accumulators. A hydraulic accu- may be used to meet the requirements mulator or reservoir may be installed of this part if the portable equipment on the engine side of any firewall if— is shown to comply with the applicable (1) It is an integral part of an engine requirements, is identified in the air- or propeller system, or plane type design, and its stowage pro- visions are found to be in compliance (2) The reservoir is nonpressurized with the requirements of § 23.561. and the total capacity of all such non- (b) The oxygen system must be free pressurized reservoirs is one quart or from hazards in itself, in its method of less. operation, and its effect upon other [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as components. amended by Amdt. 23–7, 34 FR 13096, Aug. 13, (c) There must be a means to allow 1969; Amdt. 23–14, 38 FR 31824, Nov. 19, 1973; the crew to readily determine, during Amdt. 23–43, 58 FR 18977, Apr. 9, 1993; Amdt. the flight, the quantity of oxygen 23–49, 61 FR 5170, Feb. 9, 1996] available in each source of supply. (d) Each required flight crewmember § 23.1437 Accessories for multiengine must be provided with— airplanes. (1) Demand oxygen equipment if the For multiengine airplanes, engine- airplane is to be certificated for oper- driven accessories essential to safe op- ation above 25,000 feet. eration must be distributed among two (2) Pressure demand oxygen equip- or more engines so that the failure of ment if the airplane is to be certifi- any one engine will not impair safe op- cated for operation above 40,000 feet. eration through the malfunctioning of (e) There must be a means, readily these accessories. available to the crew in flight, to turn on and to shut off the oxygen supply at § 23.1438 Pressurization and pneu- the high pressure source. This shutoff matic systems. requirement does not apply to chem- (a) Pressurization system elements ical oxygen generators. must be burst pressure tested to 2.0 [Amdt. 23–9, 35 FR 6386, Apr. 21, 1970, as times, and proof pressure tested to 1.5 amended by Amdt. 23–43, 58 FR 18978, Apr. 9, times, the maximum normal operating 1993] pressure. § 23.1443 Minimum mass flow of sup- (b) Pneumatic system elements must plemental oxygen. be burst pressure tested to 3.0 times, and proof pressure tested to 1.5 times, (a) If continuous flow oxygen equip- the maximum normal operating pres- ment is installed, an applicant must show compliance with the require- sure. ments of either paragraphs (a)(1) and (c) An analysis, or a combination of (a)(2) or paragraph (a)(3) of this sec- analysis and test, may be substituted tion: for any test required by paragraph (a) (1) For each passenger, the minimum or (b) of this section if the Adminis- mass flow of supplemental oxygen re- trator finds it equivalent to the re- quired at various cabin pressure alti- quired test. tudes may not be less than the flow re- [Amdt. 23–20, 42 FR 36969, July 18, 1977] quired to maintain, during inspiration and while using the oxygen equipment

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(including masks) provided, the fol- stant time interval between respira- lowing mean tracheal oxygen partial tions. pressures: (2) For each flight crewmember, the (i) At cabin pressure altitudes above minimum mass flow may not be less 10,000 feet up to and including 18,500 than the flow required to maintain, feet, a mean tracheal oxygen partial during inspiration, a mean tracheal ox- pressure of 100 mm. Hg when breathing ygen partial pressure of 149 mm. Hg 15 liters per minute, Body Tempera- when breathing 15 liters per minute, ture, Pressure, Saturated (BTPS) and BTPS, and with a maximum tidal vol- with a tidal volume of 700 cc. with a ume of 700 cc. with a constant time in- constant time interval between res- terval between respirations. pirations. (3) The minimum mass flow of sup- (ii) At cabin pressure altitudes above plemental oxygen supplied for each 18,500 feet up to and including 40,000 user must be at a rate not less than feet, a mean tracheal oxygen partial that shown in the following figure for pressure of 83.8 mm. Hg when breathing each altitude up to and including the 30 liters per minute, BTPS, and with a maximum operating altitude of the air- tidal volume of 1,100 cc. with a con- plane.

(b) If demand equipment is installed (c) If first-aid oxygen equipment is for use by flight crewmembers, the installed, the minimum mass flow of minimum mass flow of supplemental oxygen to each user may not be less oxygen required for each flight crew- than 4 liters per minute, STPD. How- member may not be less than the flow ever, there may be a means to decrease required to maintain, during inspira- this flow to not less than 2 liters per tion, a mean tracheal oxygen partial minute, STPD, at any cabin altitude. pressure of 122 mm. Hg up to and in- The quantity of oxygen required is cluding a cabin pressure altitude of based upon an average flow rate of 3 li- 35,000 feet, and 95 percent oxygen be- ters per minute per person for whom tween cabin pressure altitudes of 35,000 first-aid oxygen is required. and 40,000 feet, when breathing 20 liters (d) As used in this section: per minute BTPS. In addition, there (1) BTPS means Body Temperature, must be means to allow the crew to use and Pressure, Saturated (which is, 37 undiluted oxygen at their discretion. °C, and the ambient pressure to which the body is exposed, minus 47 mm. Hg,

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which is the tracheal pressure dis- nasal cannula or its connecting tubing placed by water vapor pressure when must have permanently affixed— the breathed air becomes saturated (i) A visible warning against smoking with water vapor at 37 °C). while in use; (2) STPD means Standard, Tempera- (ii) An illustration of the correct ture, and Pressure, Dry (which is, 0 °C method of donning; and at 760 mm. Hg with no water vapor). (iii) A visible warning against use [Doc. No. 26344, 58 FR 18978, Apr. 9, 1993] with nasal obstructions or head colds with resultant nasal congestion. § 23.1445 Oxygen distribution system. (c) If certification for operation (a) Except for flexible lines from oxy- above 18,000 feet (MSL) is requested, gen outlets to the dispensing units, or each oxygen dispensing unit must where shown to be otherwise suitable cover the nose and mouth of the user. to the installation, nonmetallic tubing (d) For a pressurized airplane de- must not be used for any oxygen line signed to operate at flight altitudes that is normally pressurized during above 25,000 feet (MSL), the dispensing flight. units must meet the following: (b) Nonmetallic oxygen distribution (1) The dispensing units for pas- lines must not be routed where they sengers must be connected to an oxy- may be subjected to elevated tempera- gen supply terminal and be imme- tures, electrical arcing, and released diately available to each occupant flammable fluids that might result wherever seated. from any probable failure. (2) The dispensing units for crew- members must be automatically pre- [Doc. No. 26344, 58 FR 18978, Apr. 9, 1993] sented to each crewmember before the cabin pressure altitude exceeds 15,000 § 23.1447 Equipment standards for ox- ygen dispensing units. feet, or the units must be of the quick- donning type, connected to an oxygen If oxygen dispensing units are in- supply terminal that is immediately stalled, the following apply: available to crewmembers at their sta- (a) There must be an individual dis- tion. pensing unit for each occupant for (e) If certification for operation whom supplemental oxygen is to be above 30,000 feet is requested, the dis- supplied. Each dispensing unit must: pensing units for passengers must be (1) Provide for effective utilization of automatically presented to each occu- the oxygen being delivered to the unit. pant before the cabin pressure altitude (2) Be capable of being readily placed exceeds 15,000 feet. into position on the face of the user. (f) If an automatic dispensing unit (3) Be equipped with a suitable means (hose and mask, or other unit) system to retain the unit in position on the is installed, the crew must be provided face. with a manual means to make the dis- (4) If radio equipment is installed, pensing units immediately available in the flightcrew oxygen dispensing units the event of failure of the automatic must be designed to allow the use of system. that equipment and to allow commu- nication with any other required crew [Amdt. 23–9, 35 FR 6387, Apr. 21, 1970, as member while at their assigned duty amended by Amdt. 23–20, 42 FR 36969, July 18, station. 1977; Amdt. 23–30, 49 FR 7340, Feb. 28, 1984; Amdt. 23–43, 58 FR 18978, Apr. 9, 1993; Amdt. (b) If certification for operation up to 23–49, 61 FR 5170, Feb. 9, 1996] and including 18,000 feet (MSL) is re- quested, each oxygen dispensing unit § 23.1449 Means for determining use of must: oxygen. (1) Cover the nose and mouth of the There must be a means to allow the user; or crew to determine whether oxygen is (2) Be a nasal cannula, in which case being delivered to the dispensing equip- one oxygen dispensing unit covering ment. both the nose and mouth of the user must be available. In addition, each [Amdt. 23–9, 35 FR 6387, Apr. 21, 1970]

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§ 23.1450 Chemical oxygen generators. externally applied loads arising from (a) For the purpose of this section, a consideration of limit structural loads, chemical oxygen generator is defined that may be acting on that part of the as a device which produces oxygen by system. chemical reaction. (b) Oxygen pressure sources and the (b) Each chemical oxygen generator lines between the source and the shut- must be designed and installed in ac- off means must be: cordance with the following require- (1) Protected from unsafe tempera- ments: tures; and (1) Surface temperature developed by (2) Located where the probability and the generator during operation may hazard of rupture in a crash landing not create a hazard to the airplane or are minimized. to its occupants. [Doc. No. 27806, 61 FR 5170, Feb. 9, 1996] (2) Means must be provided to relieve any internal pressure that may be haz- § 23.1457 Cockpit voice recorders. ardous. (a) Each cockpit voice recorder re- (c) In addition to meeting the re- quired by the operating rules of this quirements in paragraph (b) of this sec- chapter must be approved and must be tion, each portable chemical oxygen installed so that it will record the fol- generator that is capable of sustained lowing: operation by successive replacement of (1) Voice communications trans- a generator element must be placarded mitted from or received in the airplane to show— by radio. (1) The rate of oxygen flow, in liters (2) Voice communications of flight per minute; crewmembers on the flight deck. (2) The duration of oxygen flow, in minutes, for the replaceable generator (3) Voice communications of flight element; and crewmembers on the flight deck, using (3) A warning that the replaceable the airplane’s interphone system. generator element may be hot, unless (4) Voice or audio signals identifying the element construction is such that navigation or approach aids introduced the surface temperature cannot exceed into a headset or speaker. 100 °F. (5) Voice communications of flight crewmembers using the passenger loud- [Amdt. 23–20, 42 FR 36969, July 18, 1977] speaker system, if there is such a sys- tem and if the fourth channel is avail- § 23.1451 Fire protection for oxygen able in accordance with the require- equipment. ments of paragraph (c)(4)(ii) of this sec- Oxygen equipment and lines must: tion. (a) Not be installed in any designed (6) If datalink communication equip- fire zones. ment is installed, all datalink commu- (b) Be protected from heat that may nications, using an approved data mes- be generated in, or escape from, any sage set. Datalink messages must be designated fire zone. recorded as the output signal from the (c) Be installed so that escaping oxy- communications unit that translates gen cannot come in contact with and the signal into usable data. cause ignition of grease, fluid, or vapor (b) The recording requirements of accumulations that are present in nor- paragraph (a)(2) of this section must be mal operation or that may result from met by installing a cockpit-mounted the failure or malfunction of any other area microphone, located in the best system. position for recording voice commu- [Doc. No. 27806, 61 FR 5170, Feb. 9, 1996] nications originating at the first and second pilot stations and voice commu- § 23.1453 Protection of oxygen equip- nications of other crewmembers on the ment from rupture. flight deck when directed to those sta- (a) Each element of the oxygen sys- tions. The microphone must be so lo- tem must have sufficient strength to cated and, if necessary, the pre- withstand the maximum pressure and amplifiers and filters of the recorder temperature, in combination with any must be so adjusted or supplemented,

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so that the intelligibility of the re- (2) There is an automatic means to corded communications is as high as simultaneously stop the recorder and practicable 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; and 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 (i) That provides 10 ± 1 minutes of boom, mask, or handheld microphone, electrical power to operate both the headset, or speaker used at the first cockpit voice recorder and cockpit- pilot station. mounted area microphone; (2) For the second channel from each boom, mask, or handheld 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 handheld 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 crewmembers. 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) And that as far as is practicable located and mounted to minimize the all sounds received by the microphone probability of rupture of the container listed in paragraphs (c)(1), (2), and (4) of as a result of crash impact and con- this 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. (1)(i) It receives its electrical power (2) If two separate combination dig- from the bus that provides the max- ital flight data recorder and cockpit imum reliability for operation of the voice recorder units are installed in- cockpit voice recorder without jeopard- stead of one cockpit voice recorder and izing service to essential or emergency one digital flight data recorder, the loads. combination unit that is installed to (ii) It remains powered for as long as comply with the cockpit voice recorder possible without jeopardizing emer- requirements may be located near the gency operation of the airplane. cockpit.

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(f) If the cockpit voice recorder has a functioning, within 10 minutes after bulk erasure device, the installation crash impact; must be designed to minimize the prob- (6) Any single electrical failure exter- ability of inadvertent operation and ac- nal to the recorder does not disable tuation of the device during crash im- both the cockpit voice recorder and the pact. flight data recorder; and (g) Each recorder container must: (7) It is in a separate container from (1) Be either bright orange or bright the cockpit voice recorder when both yellow; are required. If used to comply with (2) Have reflective tape affixed to its only the flight data recorder require- external surface to facilitate its loca- ments, a combination unit may be in- tion under water; and stalled. If a combination unit is in- (3) Have an underwater locating de- stalled as a cockpit voice recorder to vice, when required by the operating comply with § 23.1457(e)(2), a combina- rules of this chapter, on or adjacent to tion unit must be used to comply with the container which is secured in such this flight data recorder requirement. manner that they are not likely to be separated during crash impact. (b) Each nonejectable record con- tainer must be located and mounted so [Amdt. 23–35, 53 FR 26142, July 11, 1988, as as to minimize the probability of con- amended by Amdt. No. 23–58, 73 FR 12562, tainer rupture resulting from crash im- Mar. 7, 2008; 74 FR 32799, July 9, 2009] pact and subsequent damage to the § 23.1459 Flight data recorders. record from fire. In meeting this re- quirement the record container must (a) Each flight recorder required by be located as far aft as practicable, but the operating rules of this chapter need not be aft of the pressurized com- must be installed so that: partment, and may not be where aft- (1) It is supplied with airspeed, alti- mounted engines may crush the con- tude, and directional data obtained tainer upon impact. from sources that meet the accuracy requirements of §§ 23.1323, 23.1325, and (c) A correlation must be established 23.1327, as appropriate; between the flight recorder readings of (2) The vertical acceleration sensor is airspeed, altitude, and heading and the rigidly attached, and located longitu- corresponding readings (taking into ac- dinally either within the approved cen- count correction factors) of the first pi- ter of gravity limits of the airplane, or lot’s instruments. The correlation at a distance forward or aft of these must cover the airspeed range over limits that does not exceed 25 percent which the airplane is to be operated, of the airplane’s mean aerodynamic the range of altitude to which the air- chord; plane is limited, and 360 degrees of (3)(i) It receives its electrical power heading. Correlation may be estab- from the bus that provides the max- lished on the ground as appropriate. imum reliability for operation of the (d) Each recorder container must: flight data recorder without jeopard- (1) Be either bright orange or bright izing service to essential or emergency yellow; loads. (2) Have reflective tape affixed to its (ii) It remains powered for as long as external surface to facilitate its loca- possible without jeopardizing emer- tion under water; and gency operation of the airplane. (3) Have an underwater locating de- (4) There is an aural or visual means vice, when required by the operating for preflight checking of the recorder rules of this chapter, on or adjacent to for proper recording of data in the stor- age medium; the container which is secured in such (5) Except for recorders powered sole- a manner that they are not likely to be ly by the engine-driven electrical gen- separated during crash impact. erator system, there is an automatic (e) Any novel or unique design or means to simultaneously stop a re- operational characteristics of the air- corder that has a data erasure feature craft shall be evaluated to determine if and prevent each erasure feature from

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any dedicated parameters must be re- § 23.1505 Airspeed limitations. corded on flight recorders in addition to or in place of existing requirements. (a) The never-exceed speed VNE must be established so that it is— [Amdt. 23–35, 53 FR 26143, July 11, 1988, as (1) Not less than 0.9 times the min- amended by Amdt. No. 23–58, 73 FR 12562, imum value of V allowed under Mar. 7, 2008; 74 FR 32800, July 9, 2009] D § 23.335; and § 23.1461 Equipment containing high (2) Not more than the lesser of— energy rotors. (i) 0.9 VD established under § 23.335; or (a) Equipment, such as Auxiliary (ii) 0.9 times the maximum speed Power Units (APU) and constant speed shown under § 23.251. drive units, containing high energy ro- (b) The maximum structural cruising tors must meet paragraphs (b), (c), or speed VNO must be established so that (d) of this section. it is— (b) High energy rotors contained in (1) Not less than the minimum value equipment must be able to withstand of VC allowed under § 23.335; and damage caused by malfunctions, vibra- (2) Not more than the lesser of— tion, abnormal speeds, and abnormal (i) V established under § 23.335; or temperatures. In addition— C (ii) 0.89 established under para- (1) Auxiliary rotor cases must be able VNE to contain damage caused by the fail- graph (a) of this section. ure of high energy rotor blades; and (c) Paragraphs (a) and (b) of this sec- (2) Equipment control devices, sys- tion do not apply to turbine airplanes tems, and instrumentation must rea- or to airplanes for which a design div- sonably ensure that no operating limi- ing speed VD/MD is established under tations affecting the integrity of high § 23.335(b)(4). For those airplanes, a energy rotors will be exceeded in serv- maximum operating limit speed (VMO/ ice. MMO-airspeed or Mach number, which- (c) It must be shown by test that ever is critical at a particular altitude) equipment containing high energy ro- must be established as a speed that tors can contain any failure of a high may not be deliberately exceeded in energy rotor that occurs at the highest any regime of flight (climb, cruise, or speed obtainable with the normal speed descent) unless a higher speed is au- control devices inoperative. thorized for flight test or pilot training (d) Equipment containing high en- operations. VMO/MMO must be estab- ergy rotors must be located where lished so that it is not greater than the rotor failure will neither endanger the design cruising speed V /M and so that occupants nor adversely affect contin- C C it is sufficiently below V /M and the ued safe flight. D D maximum speed shown under § 23.251 to [Amdt. 23–20, 42 FR 36969, July 18, 1977, as make it highly improbable that the amended by Amdt. 23–49, 61 FR 5170, Feb. 9, latter speeds will be inadvertently ex- 1996] ceeded in operations. The speed margin between VMO/MMO and VD/MD or the Subpart G—Operating Limitations maximum speed shown under § 23.251 and Information may not be less than the speed margin established between V /M and V /M § 23.1501 General. C C D D under § 23.335(b), or the speed margin (a) Each operating limitation speci- found necessary in the flight test con- fied in §§ 23.1505 through 23.1527 and ducted under § 23.253. other limitations and information nec- essary for safe operation must be es- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as tablished. amended by Amdt. 23–7, 34 FR 13096, Aug. 13, (b) The operating limitations and 1969] other information necessary for safe operation must be made available to § 23.1507 Operating maneuvering the crewmembers as prescribed in speed. §§ 23.1541 through 23.1589. The maximum operating maneu- [Amdt. 23–21, 43 FR 2319, Jan. 16, 1978] vering speed, VO, must be established

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as an operating limitation. VO is a se- (1) The maximum rotational speed; lected speed that is not greater than (2) The maximum allowable manifold VS√n established in § 23.335(c). pressure (for reciprocating engines); [Doc. No. 26269, 58 FR 42165, Aug. 6, 1993] (3) The maximum allowable gas tem- perature (for turbine engines); and § 23.1511 Flap extended speed. (4) The maximum allowable cylinder head, oil, and liquid coolant tempera- (a) The flap extended speed VFE must be established so that it is— tures. (1) Not less than the minimum value (d) Fuel grade or designation. The min- of V allowed in § 23.345(b); and imum fuel grade (for reciprocating en- F gines), or fuel designation (for turbine (2) Not more than VF established under § 23.345(a), (c), and (d). engines), must be established so that it (b) Additional combinations of flap is not less than that required for the setting, airspeed, and engine power operation of the engines within the may be established if the structure has limitations in paragraphs (b) and (c) of been proven for the corresponding de- this section. sign conditions. (e) Ambient temperature. For all air- planes except reciprocating engine- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 powered airplanes of 6,000 pounds or FR 258, Jan. 9, 1965, as amended by Amdt. 23– less maximum weight, ambient tem- 50, 61 FR 5192, Feb. 9, 1996] perature limitations (including limita- § 23.1513 Minimum control speed. tions for winterization installations if applicable) must be established as the The minimum control speed VMC, de- maximum ambient atmospheric tem- termined under § 23.149, must be estab- perature at which compliance with the lished as an operating limitation. cooling provisions of §§ 23.1041 through 23.1047 is shown. § 23.1519 Weight and center of gravity. The weight and center of gravity lim- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 itations determined under § 23.23 must FR 258, Jan. 9, 1965, as amended by Amdt. 23– 21, 43 FR 2319, Jan. 16, 1978; Amdt. 23–45, 58 be established as operating limitations. FR 42165, Aug. 6, 1993; Amdt. 23–50, 61 FR 5192, Feb. 9, 1996] § 23.1521 Powerplant limitations. (a) General. The powerplant limita- § 23.1522 Auxiliary power unit limita- tions prescribed in this section must be tions. established so that they do not exceed If an auxiliary power unit is in- the corresponding limits for which the stalled, the limitations established for engines or propellers are type certifi- the auxiliary power must be specified cated. In addition, other powerplant in the operating limitations for the air- limitations used in determining com- plane. pliance with this part must be estab- lished. [Doc. No. 26269, 58 FR 42166, Aug. 6, 1993] (b) Takeoff operation. The powerplant § 23.1523 Minimum flight crew. takeoff operation must be limited by— (1) The maximum rotational speed The minimum flight crew must be es- (rpm); tablished so that it is sufficient for safe (2) The maximum allowable manifold operation considering— pressure (for reciprocating engines); (a) The workload on individual crew- (3) The maximum allowable gas tem- members and, in addition for com- perature (for turbine engines); muter category airplanes, each crew- (4) The time limit for the use of the member workload determination must power or thrust corresponding to the consider the following: limitations established in paragraphs (1) Flight path control, (b)(1) through (3) of this section; and (2) Collision avoidance, (5) The maximum allowable cylinder (3) Navigation, head (as applicable), liquid coolant and (4) Communications, oil temperatures. (5) Operation and monitoring of all (c) Continuous operation. The contin- essential airplane systems, uous operation must be limited by— (6) Command decisions, and

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(7) The accessibility and ease of oper- MARKINGS AND PLACARDS ation of necessary controls by the ap- propriate crewmember during all nor- § 23.1541 General. mal and emergency operations when at (a) The airplane must contain— the crewmember flight station; (1) The markings and placards speci- (b) The accessibility and ease of oper- fied in §§ 23.1545 through 23.1567; and ation of necessary controls by the ap- (2) Any additional information, in- propriate crewmember; and strument markings, and placards re- (c) The kinds of operation authorized quired for the safe operation if it has under § 23.1525. unusual design, operating, or handling characteristics. [Amdt. 23–21, 43 FR 2319, Jan. 16, 1978, as amended by Amdt. 23–34, 52 FR 1834, Jan. 15, (b) Each marking and placard pre- 1987] scribed in paragraph (a) of this sec- tion— § 23.1524 Maximum passenger seating (1) Must be displayed in a con- configuration. spicuous place; and The maximum passenger seating con- (2) May not be easily erased, dis- figuration must be established. figured, or obscured. (c) For airplanes which are to be cer- [Amdt. 23–10, 36 FR 2864, Feb. 11, 1971] tificated in more than one category— (1) The applicant must select one cat- § 23.1525 Kinds of operation. egory upon which the placards and The kinds of operation authorized markings are to be based; and (e.g. VFR, IFR, day or night) and the (2) The placards and marking infor- meteorological conditions (e.g. icing) mation for all categories in which the to which the operation of the airplane airplane is to be certificated must be is limited or from which it is prohib- furnished in the Airplane Flight Man- ited, must be established appropriate ual. to the installed equipment. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 [Doc. No. 26269, 58 FR 42166, Aug. 6, 1993] FR 258, Jan. 9, 1965, as amended by Amdt. 23– 21, 43 FR 2319, Jan. 16, 1978] § 23.1527 Maximum operating altitude. § 23.1543 Instrument markings: Gen- (a) The maximum altitude up to eral. which operation is allowed, as limited by flight, structural, powerplant, func- For each instrument— tional or equipment characteristics, (a) When markings are on the cover must be established. glass of the instrument, there must be (b) A maximum operating altitude means to maintain the correct align- limitation of not more than 25,000 feet ment of the glass cover with the face of must be established for pressurized air- the dial; and planes unless compliance with (b) Each arc and line must be wide § 23.775(e) is shown. enough and located to be clearly visi- ble to the pilot. [Doc. No. 26269, 58 FR 42166, Aug. 6, 1993] (c) All related instruments must be calibrated in compatible units. § 23.1529 Instructions for Continued Airworthiness. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23– The applicant must prepare Instruc- 50, 61 FR 5192, Feb. 9, 1996] tions for Continued Airworthiness in accordance with appendix G to this § 23.1545 Airspeed indicator. part that are acceptable to the Admin- (a) Each airspeed indicator must be istrator. The instructions may be in- marked as specified in paragraph (b) of complete at type certification if a pro- this section, with the marks located at gram exists to ensure their completion the corresponding indicated airspeeds. prior to delivery of the first airplane or (b) The following markings must be issuance of a standard certificate of made: airworthiness, whichever occurs later. (1) For the never-exceed speed VNE, a [Amdt. 23–26, 45 FR 60171, Sept. 11, 1980] radial red line.

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(2) For the caution range, a yellow § 23.1547 Magnetic direction indicator. arc extending from the red line speci- (a) A placard meeting the require- fied in paragraph (b)(1) of this section ments of this section must be installed to the upper limit of the green arc on or near the magnetic direction indi- specified in paragraph (b)(3) of this sec- cator. tion. (b) The placard must show the cali- (3) For the normal operating range, a bration of the instrument in level green arc with the lower limit at V S1 flight with the engines operating. with maximum weight and with land- (c) The placard must state whether ing gear and wing flaps retracted, and the calibration was made with radio re- the upper limit at the maximum struc- ceivers on or off. tural cruising speed V established NO (d) Each calibration reading must be under § 23.1505(b). in terms of magnetic headings in not (4) For the flap operating range, a more than 30 degree increments. white arc with the lower limit at V at S0 (e) If a magnetic nonstabilized direc- the maximum weight, and the upper tion indicator can have a deviation of limit at the flaps-extended speed V FE more than 10 degrees caused by the op- established under § 23.1511. eration of electrical equipment, the (5) For reciprocating multiengine- placard must state which electrical powered airplanes of 6,000 pounds or loads, or combination of loads, would less maximum weight, for the speed at cause a deviation of more than 10 de- which compliance has been shown with grees when turned on. § 23.69(b) relating to rate of climb at maximum weight and at sea level, a [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 blue radial line. FR 258, Jan. 9, 1965, as amended by Amdt. 23– (6) For reciprocating multiengine- 20, 42 FR 36969, July 18, 1977] powered airplanes of 6,000 pounds or § 23.1549 Powerplant and auxiliary less maximum weight, for the max- power unit instruments. imum value of minimum control speed, For each required powerplant and VMC, (one-engine-inoperative) deter- mined under § 23.149(b), a red radial auxiliary power unit instrument, as ap- line. propriate to the type of instruments— (a) Each maximum and, if applicable, (c) If VNE or VNO vary with altitude, there must be means to indicate to the minimum safe operating limit must be pilot the appropriate limitations marked with a red radial or a red line; throughout the operating altitude (b) Each normal operating range range. must be marked with a green arc or (d) Paragraphs (b)(1) through (b)(3) green line, not extending beyond the and paragraph (c) of this section do not maximum and minimum safe limits; apply to aircraft for which a maximum (c) Each takeoff and precautionary operating speed VMO/MMO is estab- range must be marked with a yellow lished under § 23.1505(c). For those air- arc or a yellow line; and craft there must either be a maximum (d) Each engine, auxiliary power allowable airspeed indication showing unit, or propeller range that is re- the variation of VMO/MMO with altitude stricted because of excessive vibration or compressibility limitations (as ap- stresses must be marked with red arcs propriate), or a radial red line marking or red lines. for VMO/MMO must be made at lowest [Amdt. 23–12, 41 FR 55466, Dec. 20, 1976, as value of VMO/MMO established for any amended by Amdt. 23–28, 47 FR 13315, Mar. 29, altitude up to the maximum operating 1982; Amdt. 23–45, 58 FR 42166, Aug. 6, 1993] altitude for the airplane. § 23.1551 Oil quantity indicator. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–3, 30 FR 14240, Nov. 13, Each oil quantity indicator must be 1965; Amdt. 23–7, 34 FR 13097, Aug. 13, 1969; marked in sufficient increments to in- Amdt. 23–23, 43 FR 50593, Oct. 30, 1978; Amdt. dicate readily and accurately the quan- 23–50, 61 FR 5193, Feb. 9, 1996] tity of oil.

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§ 23.1553 Fuel quantity indicator. tion to its other functions, shall be this color. A red radial line must be marked on each indicator at the calibrated zero [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 reading, as specified in § 23.1337(b)(1). FR 258, Jan. 9, 1965, as amended by Amdt. 23– 21, 43 FR 2319, Jan. 16, 1978; Amdt. 23–50, 61 [Doc. No. 27807, 61 FR 5193, Feb. 9, 1996] FR 5193, Feb. 9, 1996]

§ 23.1555 Control markings. § 23.1557 Miscellaneous markings and placards. (a) Each cockpit control, other than primary flight controls and simple (a) Baggage and cargo compartments, push button type starter switches, and ballast location. Each baggage and must be plainly marked as to its func- cargo compartment, and each ballast tion and method of operation. location, must have a placard stating any limitations on contents, including (b) Each secondary control must be weight, that are necessary under the suitably marked. loading requirements. (c) For powerplant fuel controls— (b) Seats. If the maximum allowable (1) Each fuel tank selector control weight to be carried in a seat is less must be marked to indicate the posi- than 170 pounds, a placard stating the tion corresponding to each tank and to lesser weight must be permanently at- each existing cross feed position; tached to the seat structure. (2) If safe operation requires the use (c) Fuel, oil, and coolant filler open- of any tanks in a specific sequence, ings. The following apply: that sequence must be marked on or (1) Fuel filler openings must be near the selector for those tanks; marked at or near the filler cover (3) The conditions under which the with— full amount of usable fuel in any re- (i) For reciprocating engine-powered stricted usage fuel tank can safely be airplanes— used must be stated on a placard adja- (A) The word ‘‘Avgas’’; and cent to the selector valve for that (B) The minimum fuel grade. tank; and (ii) For turbine engine-powered air- (4) Each valve control for any engine planes— of a multiengine airplane must be (A) The words ‘‘Jet Fuel’’; and marked to indicate the position cor- (B) The permissible fuel designations, responding to each engine controlled. or references to the Airplane Flight (d) Usable fuel capacity must be Manual (AFM) for permissible fuel des- marked as follows: ignations. (iii) For pressure fueling systems, the (1) For fuel systems having no selec- maximum permissible fueling supply tor controls, the usable fuel capacity of pressure and the maximum permissible the system must be indicated at the defueling pressure. fuel quantity indicator. (2) Oil filler openings must be (2) For fuel systems having selector marked at or near the filler cover with controls, the usable fuel capacity the word ‘‘Oil’’ and the permissible oil available at each selector control posi- designations, or references to the Air- tion must be indicated near the selec- plane Flight Manual (AFM) for permis- tor control. sible oil designations. (e) For accessory, auxiliary, and (3) Coolant filler openings must be emergency controls— marked at or near the filler cover with (1) If retractable landing gear is used, the word ‘‘Coolant’’. the indicator required by § 23.729 must (d) Emergency exit placards. Each be marked so that the pilot can, at any placard and operating control for each time, ascertain that the wheels are se- emergency exit must be red. A placard cured in the extreme positions; and must be near each emergency exit con- (2) Each emergency control must be trol and must clearly indicate the loca- red and must be marked as to method tion of that exit and its method of op- of operation. No control other than an eration. emergency control, or a control that (e) The system voltage of each direct serves an emergency function in addi- current installation must be clearly

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marked adjacent to its exernal power § 23.1567 Flight maneuver placard. connection. (a) For normal category airplanes, [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; as there must be a placard in front of and amended by Amdt. 23–21, 42 FR 15042, Mar. 17, in clear view of the pilot stating: ‘‘No 1977; Amdt. 23–23, 43 FR 50594, Oct. 30, 1978; acrobatic maneuvers, including spins, Amdt. 23–45, 58 FR 42166, Aug. 6, 1993; 73 FR approved.’’ 35063, June 20, 2008] (b) For utility category airplanes, there must be— § 23.1559 Operating limitations (1) A placard in clear view of the placard. pilot stating: ‘‘Acrobatic maneuvers are (a) There must be a placard in clear limited to the following view of the pilot stating— lllllllllll;’’ (list approved (1) That the airplane must be oper- maneuvers and the recommended entry ated in accordance with the Airplane speed for each); and Flight Manual; and (2) For those airplanes that do not (2) The certification category of the meet the spin requirements for acro- airplane to which the placards apply. batic category airplanes, an additional (b) For airplanes certificated in more placard in clear view of the pilot stat- than one category, there must be a ing: ‘‘Spins Prohibited.’’ placard in clear view of the pilot stat- (c) For acrobatic category airplanes, ing that other limitations are con- there must be a placard in clear view of tained in the Airplane Flight Manual. the pilot listing the approved acrobatic maneuvers and the recommended entry (c) There must be a placard in clear airspeed for each. If inverted flight ma- view of the pilot that specifies the kind neuvers are not approved, the placard of operations to which the operation of must bear a notation to this effect. the airplane is limited or from which it (d) For acrobatic category airplanes is prohibited under § 23.1525. and utility category airplanes approved [Doc. No. 27807, 61 FR 5193, Feb. 9, 1996] for spinning, there must be a placard in clear view of the pilot— § 23.1561 Safety equipment. (1) Listing the control actions for re- covery from spinning maneuvers; and (a) Safety equipment must be plainly (2) Stating that recovery must be ini- marked as to method of operation. tiated when spiral characteristics ap- (b) Stowage provisions for required pear, or after not more than six turns safety equipment must be marked for or not more than any greater number the benefit of occupants. of turns for which the airplane has been certificated. § 23.1563 Airspeed placards. There must be an airspeed placard in [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23– clear view of the pilot and as close as 13, 37 FR 20023, Sept. 23, 1972; Amdt. 23–21, 43 practicable to the airspeed indicator. FR 2319, Jan. 16, 1978; Amdt. 23–50, 61 FR 5193, This placard must list— Feb. 9, 1996] (a) The operating maneuvering speed, AIRPLANE FLIGHT MANUAL AND VO; and (b) The maximum landing gear oper- APPROVED MANUAL MATERIAL ating speed VLO. § 23.1581 General. (c) For reciprocating multiengine- powered airplanes of more than 6,000 (a) Furnishing information. An Air- pounds maximum weight, and turbine plane Flight Manual must be furnished with each airplane, and it must contain engine-powered airplanes, the max- the following: imum value of the minimum control (1) Information required by §§ 23.1583 speed, V (one-engine-inoperative) de- MC through 23.1589. termined under § 23.149(b). (2) Other information that is nec- [Amdt. 23–7, 34 FR 13097, Aug. 13, 1969, as essary for safe operation because of de- amended by Amdt. 23–45, 58 FR 42166, Aug. 6, sign, operating, or handling character- 1993; Amdt. 23–50, 61 FR 5193, Feb. 9, 1996] istics.

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(3) Further information necessary to § 23.1583 Operating limitations. comply with the relevant operating The Airplane Flight Manual must rules. contain operating limitations deter- (b) Approved information. (1) Except as mined under this part 23, including the provided in paragraph (b)(2) of this sec- following— tion, each part of the Airplane Flight (a) Airspeed limitations. The following Manual containing information pre- information must be furnished: scribed in §§ 23.1583 through 23.1589 (1) Information necessary for the must be approved, segregated, identi- marking of the airspeed limits on the fied and clearly distinguished from indicator as required in § 23.1545, and each unapproved part of that Airplane the significance of each of those limits Flight Manual. and of the color coding used on the in- (2) The requirements of paragraph dicator. (b)(1) of this section do not apply to re- (2) The speeds VMC, VO, VLE, and VLO, ciprocating engine-powered airplanes if established, and their significance. of 6,000 pounds or less maximum (3) In addition, for turbine powered weight, if the following is met: commuter category airplanes— (i) Each part of the Airplane Flight (i) The maximum operating limit Manual containing information pre- speed, VMO/MMO and a statement that scribed in § 23.1583 must be limited to this speed must not be deliberately ex- such information, and must be ap- ceeded in any regime of flight (climb, proved, identified, and clearly distin- cruise or descent) unless a higher speed guished from each other part of the is authorized for flight test or pilot Airplane Flight Manual. training; (ii) The information prescribed in (ii) If an airspeed limitation is based §§ 23.1585 through 23.1589 must be deter- upon compressibility effects, a state- mined in accordance with the applica- ment to this effect and information as ble requirements of this part and pre- to any symptoms, the probable behav- sented in its entirety in a manner ac- ior of the airplane, and the rec- ceptable to the Administrator. ommended recovery procedures; and (3) Each page of the Airplane Flight (iii) The airspeed limits must be Manual containing information pre- shown in terms of VMO/MMO instead of scribed in this section must be of a VNO and VNE. type that is not easily erased, dis- (b) Powerplant limitations. The fol- figured, or misplaced, and is capable of lowing information must be furnished: being inserted in a manual provided by (1) Limitations required by § 23.1521. the applicant, or in a folder, or in any (2) Explanation of the limitations, other permanent binder. when appropriate. (c) The units used in the Airplane (3) Information necessary for mark- Flight Manual must be the same as ing the instruments required by those marked on the appropriate in- § 23.1549 through § 23.1553. struments and placards. (c) Weight. The airplane flight man- (d) All Airplane Flight Manual oper- ual must include— ational airspeeds, unless otherwise (1) The maximum weight; and specified, must be presented as indi- (2) The maximum landing weight, if cated airspeeds. the design landing weight selected by (e) Provision must be made for stow- the applicant is less than the max- ing the Airplane Flight Manual in a imum weight. suitable fixed container which is read- (3) For normal, utility, and acrobatic ily accessible to the pilot. category reciprocating engine-powered (f) Revisions and amendments. Each airplanes of more than 6,000 pounds Airplane Flight Manual (AFM) must maximum weight and for turbine en- contain a means for recording the in- gine-powered airplanes in the normal, corporation of revisions and amend- utility, and acrobatic category, per- ments. formance operating limitations as fol- [Amdt. 23–21, 43 FR 2319, Jan. 16, 1978, as lows— amended by Amdt. 23–34, 52 FR 1834, Jan. 15, (i) The maximum takeoff weight for 1987; Amdt. 23–45, 58 FR 42166, Aug. 6, 1993; each airport altitude and ambient tem- Amdt. 23–50, 61 FR 5193, Feb. 9, 1996] perature within the range selected by

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the applicant at which the airplane associated limitations. No other ma- complies with the climb requirements neuver is authorized. of § 23.63(c)(1). (3) Acrobatic category airplanes. A list (ii) The maximum landing weight for of approved flight maneuvers dem- each airport altitude and ambient tem- onstrated in the type flight tests, to- perature within the range selected by gether with recommended entry speeds the applicant at which the airplane and any other associated limitations. complies with the climb requirements (4) Acrobatic category airplanes and of § 23.63(c)(2). utility category airplanes approved for (4) For commuter category airplanes, spinning. Spin recovery procedure es- the maximum takeoff weight for each tablished to show compliance with airport altitude and ambient tempera- § 23.221(c). ture within the range selected by the (5) Commuter category airplanes. Ma- applicant at which— neuvers are limited to any maneuver (i) The airplane complies with the incident to normal flying, stalls, (ex- climb requirements of § 23.63(d)(1); and cept whip stalls) and steep turns in (ii) The accelerate-stop distance de- which the angle of bank is not more termined under § 23.55 is equal to the than 60 degrees. available runway length plus the (f) Maneuver load factor. The positive length of any stopway, if utilized; and limit load factors in g’s, and, in addi- either: tion, the negative limit load factor for (iii) The takeoff distance determined acrobatic category airplanes. under § 23.59(a) is equal to the available (g) Minimum flight crew. The number runway length; or and functions of the minimum flight (iv) At the option of the applicant, crew determined under § 23.1523. the takeoff distance determined under (h) Kinds of operation. A list of the § 23.59(a) is equal to the available run- kinds of operation to which the air- way length plus the length of any plane is limited or from which it is pro- clearway and the takeoff run deter- hibited under § 23.1525, and also a list of mined under § 23.59(b) is equal to the installed equipment that affects any available runway length. operating limitation and identification (5) For commuter category airplanes, as to the equipment’s required oper- the maximum landing weight for each ational status for the kinds of oper- airport altitude within the range se- ation for which approval has been lected by the applicant at which— given. (i) The airplane complies with the (i) Maximum operating altitude. The climb requirements of § 23.63(d)(2) for maximum altitude established under ambient temperatures within the range § 23.1527. selected by the applicant; and (j) Maximum passenger seating configu- (ii) The landing distance determined ration. The maximum passenger seating under § 23.75 for standard temperatures configuration. is equal to the available runway length. (k) Allowable lateral fuel loading. The (6) The maximum zero wing fuel maximum allowable lateral fuel load- weight, where relevant, as established ing differential, if less than the max- in accordance with § 23.343. imum possible. (d) Center of gravity. The established (l) Baggage and cargo loading. The fol- center of gravity limits. lowing information for each baggage (e) Maneuvers. The following author- and cargo compartment or zone— ized maneuvers, appropriate airspeed (1) The maximum allowable load; and limitations, and unauthorized maneu- (2) The maximum intensity of load- vers, as prescribed in this section. ing. (1) Normal category airplanes. No acro- (m) Systems. Any limitations on the batic maneuvers, including spins, are use of airplane systems and equipment. authorized. (n) Ambient temperatures. Where ap- (2) Utility category airplanes. A list of propriate, maximum and minimum am- authorized maneuvers demonstrated in bient air temperatures for operation. the type flight tests, together with rec- (o) Smoking. Any restrictions on ommended entry speeds and any other smoking in the airplane.

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(p) Types of surface. A statement of (2) Procedures, speeds, and configura- the types of surface on which oper- tion(s) for making a balked landing ations may be conducted. (See § 23.45(g) with one engine inoperative and the and § 23.1587 (a)(4), (c)(2), and (d)(4)). conditions under which a balked land- [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as ing can be performed safely, or a warn- amended by Amdt. 23–7, 34 FR 13097, Aug. 13, ing against attempting a balked land- 1969; Amdt. 23–10, 36 FR 2864, Feb. 11, 1971; ing; Amdt. 23–21, 43 FR 2320, Jan. 16, 1978; Amdt. (3) The VSSE determined in § 23.149; 23–23, 43 FR 50594, Oct. 30, 1978; Amdt. 23–34, and 52 FR 1834, Jan. 15, 1987; Amdt. 23–45, 58 FR 42166, Aug. 6, 1993; Amdt. 23–50, 61 FR 5193, (4) Procedures for restarting any en- Feb. 9, 1996] gine in flight including the effects of altitude. § 23.1585 Operating procedures. (d) In addition to paragraphs (a) and (a) For all airplanes, information either (b) or (c) of this section, as ap- concerning normal, abnormal (if appli- propriate, for all normal, utility, and cable), and emergency procedures and acrobatic category airplanes, the fol- other pertinent information necessary lowing information must be furnished: for safe operation and the achievement (1) Procedures, speeds, and configura- of the scheduled performance must be tion(s) for making a normal takeoff, in furnished, including— accordance with § 23.51 (a) and (b), and (1) An explanation of significant or § 23.53 (a) and (b), and the subsequent unusual flight or ground handling char- climb, in accordance with § 23.65 and acteristics; § 23.69(a). (2) The maximum demonstrated val- (2) Procedures for abandoning a take- ues of crosswind for takeoff and land- off due to engine failure or other cause. ing, and procedures and information (e) In addition to paragraphs (a), (c), pertinent to operations in crosswinds; and (d) of this section, for all normal, (3) A recommended speed for flight in utility, and acrobatic category multi- rough air. This speed must be chosen to protect against the occurrence, as a re- engine airplanes, the information must sult of gusts, of structural damage to include the following: the airplane and loss of control (for ex- (1) Procedures and speeds for con- ample, stalling); tinuing a takeoff following engine fail- (4) Procedures for restarting any tur- ure and the conditions under which bine engine in flight, including the ef- takeoff can safely be continued, or a fects of altitude; and warning against attempting to con- (5) Procedures, speeds, and configura- tinue the takeoff. tion(s) for making a normal approach (2) Procedures, speeds, and configura- and landing, in accordance with §§ 23.73 tions for continuing a climb following and 23.75, and a transition to the engine failure, after takeoff, in accord- balked landing condition. ance with § 23.67, or enroute, in accord- (6) For seaplanes and amphibians, ance with § 23.69(b). water handling procedures and the (f) In addition to paragraphs (a) and demonstrated wave height. (c) of this section, for commuter cat- (b) In addition to paragraph (a) of egory airplanes, the information must this section, for all single-engine air- include the following: planes, the procedures, speeds, and con- (1) Procedures, speeds, and configura- figuration(s) for a glide following en- tion(s) for making a normal takeoff. gine failure, in accordance with § 23.71 and the subsequent forced landing, (2) Procedures and speeds for car- must be furnished. rying out an accelerate-stop in accord- (c) In addition to paragraph (a) of ance with § 23.55. this section, for all multiengine air- (3) Procedures and speeds for con- planes, the following information must tinuing a takeoff following engine fail- be furnished: ure in accordance with § 23.59(a)(1) and (1) Procedures, speeds, and configura- for following the flight path deter- tion(s) for making an approach and mined under § 23.57 and § 23.61(a). landing with one engine inoperative;

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(g) For multiengine airplanes, infor- angle of climb/descent, determined mation identifying each operating con- under § 23.77(a), must be furnished. dition in which the fuel system inde- (c) In addition to paragraphs (a) and pendence prescribed in § 23.953 is nec- (b) of this section, if appropriate, for essary for safety must be furnished, to- normal, utility, and acrobatic category gether with instructions for placing airplanes, the following information the fuel system in a configuration used must be furnished— to show compliance with that section. (1) The takeoff distance, determined (h) For each airplane showing com- under § 23.53 and the type of surface for pliance with § 23.1353 (g)(2) or (g)(3), the which it is valid. operating procedures for disconnecting (2) The effect on takeoff distance of the battery from its charging source operation on other than smooth hard must be furnished. surfaces, when dry, determined under (i) Information on the total quantity § 23.45(g); of usable fuel for each fuel tank, and (3) The effect on takeoff distance of the effect on the usable fuel quantity, runway slope and 50 percent of the as a result of a failure of any pump, headwind component and 150 percent of must be furnished. the tailwind component; (j) Procedures for the safe operation (4) For multiengine reciprocating en- of the airplane’s systems and equip- gine-powered airplanes of more than ment, both in normal use and in the 6,000 pounds maximum weight and mul- event of malfunction, must be fur- tiengine turbine powered airplanes, the nished. one-engine-inoperative takeoff climb/ [Doc. No. 27807, 61 FR 5194, Feb. 9, 1996] descent gradient, determined under § 23.66; § 23.1587 Performance information. (5) For multiengine airplanes, the Unless otherwise prescribed, perform- enroute rate and gradient of climb/de- ance information must be provided scent with one engine inoperative, de- over the altitude and temperature termined under § 23.69(b); and ranges required by § 23.45(b). (6) For single-engine airplanes, the (a) For all airplanes, the following in- glide performance determined under formation must be furnished— § 23.71. (d) In addition to paragraph (a) of (1) The stalling speeds VSO and VS1 with the landing gear and wing flaps this section, for commuter category retracted, determined at maximum airplanes, the following information weight under § 23.49, and the effect on must be furnished— these stalling speeds of angles of bank (1) The accelerate-stop distance de- up to 60 degrees; termined under § 23.55; (2) The steady rate and gradient of (2) The takeoff distance determined climb with all engines operating, deter- under § 23.59(a); mined under § 23.69(a); (3) At the option of the applicant, the (3) The landing distance, determined takeoff run determined under § 23.59(b); under § 23.75 for each airport altitude (4) The effect on accelerate-stop dis- and standard temperature, and the tance, takeoff distance and, if deter- type of surface for which it is valid; mined, takeoff run, of operation on (4) The effect on landing distances of other than smooth hard surfaces, when operation on other than smooth hard dry, determined under § 23.45(g); surfaces, when dry, determined under (5) The effect on accelerate-stop dis- § 23.45(g); and tance, takeoff distance, and if deter- (5) The effect on landing distances of mined, takeoff run, of runway slope runway slope and 50 percent of the and 50 percent of the headwind compo- headwind component and 150 percent of nent and 150 percent of the tailwind the tailwind component. component; (b) In addition to paragraph (a) of (6) The net takeoff flight path deter- this section, for all normal, utility, and mined under § 23.61(b); acrobatic category reciprocating en- (7) The enroute gradient of climb/de- gine-powered airplanes of 6,000 pounds scent with one engine inoperative, de- or less maximum weight, the steady termined under § 23.69(b);

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(8) The effect, on the net takeoff (2) A main wing located closer to the air- flight path and on the enroute gradient plane’s center of gravity than to the aft, fu- of climb/descent with one engine inop- selage-mounted, empennage; erative, of 50 percent of the headwind (3) A main wing that contains a quarter- component and 150 percent of the tail- chord sweep angle of not more than 15 de- wind component; grees fore or aft; (9) Overweight landing performance (4) A main wing that is equipped with trail- ing-edge controls (ailerons or flaps, or both); information (determined by extrapo- (5) A main wing aspect ratio not greater lation and computed for the range of than 7; weights between the maximum landing (6) A horizontal tail aspect ratio not great- and maximum takeoff weights) as fol- er than 4; lows— (7) A horizontal tail volume coefficient not (i) The maximum weight for each air- less than 0.34; port altitude and ambient temperature (8) A vertical tail aspect ratio not greater at which the airplane complies with than 2; the climb requirements of § 23.63(d)(2); (9) A vertical tail platform area not great- and er than 10 percent of the wing platform area; (ii) The landing distance determined and under § 23.75 for each airport altitude (10) Symmetrical airfoils must be used in and standard temperature. both the horizontal and vertical tail designs. (10) The relationship between IAS (b) Appendix A criteria may not be used on and CAS determined in accordance any airplane configuration that contains any with § 23.1323 (b) and (c). of the following design features: (11) The altimeter system calibration (1) Canard, tandem-wing, close-coupled, or required by § 23.1325(e). tailless arrangements of the lifting surfaces; (2) Biplane or multiplane wing arrange- [Doc. No. 27807, 61 FR 5194, Feb. 9, 1996] ments; (3) T-tail, V-tail, or cruciform-tail (+) ar- § 23.1589 Loading information. rangements; The following loading information (4) Highly-swept wing platform (more than must be furnished: 15-degrees of sweep at the quarter-chord), delta planforms, or slatted lifting surfaces; (a) The weight and location of each or item of equipment that can be easily (5) Winglets or other wing tip devices, or removed, relocated, or replaced and outboard fins. that is installed when the airplane was weighed under the requirement of A23.3 Special symbols. § 23.25. n1=Airplane Positive Maneuvering Limit (b) Appropriate loading instructions Load Factor. for each possible loading condition be- n2=Airplane Negative Maneuvering Limit tween the maximum and minimum Load Factor.

weights established under § 23.25, to fa- n3=Airplane Positive Gust Limit Load Fac- cilitate the center of gravity remain- tor at VC. ing within the limits established under n4=Airplane Negative Gust Limit Load Fac- § 23.23. tor at VC. n =Airplane Positive Limit Load Factor [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as flap With Flaps Fully Extended at V amended by Amdt. 23–45, 58 FR 42167, Aug. 6, F. 1993; Amdt. 23–50, 61 FR 5195, Feb. 9, 1996]

APPENDIX A TO PART 23—SIMPLIFIED DESIGN LOAD CRITERIA

A23.1 General. (a) The design load criteria in this appen- dix are an approved equivalent of those in §§ 23.321 through 23.459 of this subchapter for an airplane having a maximum weight of 6,000 pounds or less and the following con- figuration: (1) A single engine excluding turbine pow- erplants;

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A23.5 Certification in more than one category. weight loading conditions for the c.g. range selected. The criteria in this appendix may be used (e) The following loads and loading condi- for certification in the normal, utility, and tions are the minimums for which strength acrobatic categories, or in any combination must be provided in the structure: of these categories. If certification in more (1) Airplane equilibrium. The aerodynamic than one category is desired, the design cat- wing loads may be considered to act normal egory weights must be selected to make the to the relative wind, and to have a mag- term n W constant for all categories or 1 nitude of 1.05 times the airplane normal greater for one desired category than for loads (as determined from paragraphs A23.9 others. The wings and control surfaces (in- (b) and (c) of this appendix) for the positive cluding wing flaps and tabs) need only be in- flight conditions and a magnitude equal to vestigated for the maximum value of n W, or 1 the airplane normal loads for the negative for the category corresponding to the max- conditions. Each chordwise and normal com- imum design weight, where n W is constant. 1 ponent of this wing load must be considered. If the acrobatic category is selected, a spe- (2) Minimum design airspeeds. The minimum cial unsymmetrical flight load investigation design airspeeds may be chosen by the appli- in accordance with paragraphs A23.9(c)(2) cant except that they may not be less than and A23.11(c)(2) of this appendix must be the minimum speeds found by using figure 3 completed. The wing, wing carrythrough, of this appendix. In addition, V need not and the horizontal tail structures must be Cmin exceed values of 0.9 V actually obtained at checked for this condition. The basic fuse- H sea level for the lowest design weight cat- lage structure need only be investigated for egory for which certification is desired. In the highest load factor design category se- computing these minimum design airspeeds, lected. The local supporting structure for n1 may not be less than 3.8. dead weight items need only be designed for (3) Flight load factor. The limit flight load the highest load factor imposed when the factors specified in Table 1 of this appendix particular items are installed in the air- represent the ratio of the aerodynamic force plane. The engine mount, however, must be component (acting normal to the assumed designed for a higher side load factor, if cer- longitudinal axis of the airplane) to the tification in the acrobatic category is de- weight of the airplane. A positive flight load sired, than that required for certification in factor is an aerodynamic force acting up- the normal and utility categories. When de- ward, with respect to the airplane. signing for landing loads, the landing gear and the airplane as a whole need only be in- A23.9 Flight conditions. vestigated for the category corresponding to (a) General. Each design condition in para- the maximum design weight. These sim- graphs (b) and (c) of this section must be plifications apply to single-engine aircraft of used to assure sufficient strength for each conventional types for which experience is condition of speed and load factor on or available, and the Administrator may re- within the boundary of a V¥n diagram for quire additional investigations for aircraft the airplane similar to the diagram in figure with unusual design features. 4 of this appendix. This diagram must also be A23.7 Flight loads. used to determine the airplane structural op- erating limitations as specified in (a) Each flight load may be considered §§ 23.1501(c) through 23.1513 and § 23.1519. independent of altitude and, except for the (b) Symmetrical flight conditions. The air- local supporting structure for dead weight plane must be designed for symmetrical items, only the maximum design weight con- flight conditions as follows: ditions must be investigated. (1) The airplane must be designed for at (b) Table 1 and figures 3 and 4 of this ap- least the four basic flight conditions, ‘‘A’’, pendix must be used to determine values of ‘‘D’’, ‘‘E’’, and ‘‘G’’ as noted on the flight enve- n1, n2, n3, and n4, corresponding to the max- lope of figure 4 of this appendix. In addition, imum design weights in the desired cat- the following requirements apply: egories. (i) The design limit flight load factors cor- (c) Figures 1 and 2 of this appendix must be responding to conditions ‘‘D’’ and ‘‘E’’ of figure used to determine values of n3 and n4 cor- 4 must be at least as great as those specified responding to the minimum flying weights in in Table 1 and figure 4 of this appendix, and the desired categories, and, if these load fac- the design speed for these conditions must be tors are greater than the load factors at the at least equal to the value of VD found from design weight, the supporting structure for figure 3 of this appendix. dead weight items must be substantiated for (ii) For conditions ‘‘A’’ and ‘‘G’’ of figure 4, the resulting higher load factors. the load factors must correspond to those (d) Each specified wing and tail loading is specified in Table 1 of this appendix, and the independent of the center of gravity range. design speeds must be computed using these The applicant, however, must select a c.g. load factors with the maximum static lift range, and the basic fuselage structure must coefficient CNA determined by the applicant. be investigated for the most adverse dead However, in the absence of more precise

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computations, these latter conditions may (ii) Compute K from the formula: be based on a value of CNA=±1.35 and the de- sign speed for condition ‘‘A’’ may be less than − δ ()CVmb001. 2 VAmin. = D (iii) Conditions ‘‘C’’ and ‘‘F’’ of figure 4 need K only be investigated when n W/S or n W/S ()CV− 001. δ 3 4 maC2 are greater than n1 W/S or n2 W/S of this ap- pendix, respectively. where da is the down aileron deflection cor- (2) If flaps or other high lift devices in- responding to Da, and db is the down aile- tended for use at the relatively low airspeed ron deflection corresponding to D b as com- of approach, landing, and takeoff, are in- puted in step (i). stalled, the airplane must be designed for the (iii) If K is less than 1.0, Da is D critical and two flight conditions corresponding to the must be used to determine dU and dd. In this values of limit flap-down factors specified in case, VC is the critical speed which must be Table 1 of this appendix with the flaps fully used in computing the wing torsion loads extended at not less than the design flap over the aileron span. speed VFmin from figure 3 of this appendix. (iv) If K is equal to or greater than 1.0, DB (c) Unsymmetrical flight conditions. Each af- is D critical and must be used to determine fected structure must be designed for unsym- dU and dD. In this case, Vd is the critical metrical loadings as follows: speed which must be used in computing the (1) The aft fuselage-to-wing attachment wing torsion loads over the aileron span. must be designed for the critical vertical (d) Supplementary conditions; rear lift truss; surface load determined in accordance with engine torque; side load on engine mount. Each paragraph SA23.11(c)(1) and (2) of this appen- of the following supplementary conditions dix. must be investigated: (2) The wing and wing carry-through struc- (1) In designing the rear lift truss, the spe- tures must be designed for 100 percent of con- cial condition specified in § 23.369 may be in- dition ‘‘A’’ loading on one side of the plane of vestigated instead of condition ‘‘G’’ of figure symmetry and 70 percent on the opposite 4 of this appendix. If this is done, and if cer- side for certification in the normal and util- tification in more than one category is de- ity categories, or 60 percent on the opposite sired, the value of W/S used in the formula side for certification in the acrobatic cat- appearing in § 23.369 must be that for the cat- egory. egory corresponding to the maximum gross (3) The wing and wing carry-through struc- weight. tures must be designed for the loads result- (2) Each engine mount and its supporting ing from a combination of 75 percent of the structures must be designed for the max- positive maneuvering wing loading on both imum limit torque corresponding to METO sides of the plane of symmetry and the max- power and propeller speed acting simulta- imum wing torsion resulting from aileron neously with the limit loads resulting from displacement. The effect of aileron displace- the maximum positive maneuvering flight ment on wing torsion at VC or VA using the load factor n1. The limit torque must be ob- basic airfoil moment coefficient modified tained by multiplying the mean torque by a over the aileron portion of the span, must be factor of 1.33 for engines with five or more computed as follows: cylinders. For 4, 3, and 2 cylinder engines, (i) Cm=Cm +0.01dm (up aileron side) wing the factor must be 2, 3, and 4, respectively. basic airfoil. (3) Each engine mount and its supporting (ii) Cm=Cm ¥0.01dm(down aileron side) wing structure must be designed for the loads re- basic airfoil, where dm is the up aileron de- sulting from a lateral limit load factor of not flection and d d is the down aileron deflec- less than 1.47 for the normal and utility cat- tion. egories, or 2.0 for the acrobatic category. (4) D critical, which is the sum of dm+d d must be computed as follows: A23.11 Control surface loads. (i) Compute Da and DB from the formulas: (a) General. Each control surface load must be determined using the criteria of para- V graph (b) of this section and must lie within ΔΔ=×A a p and the simplified loadings of paragraph (c) of VC this section. (b) Limit pilot forces. In each control surface V loading condition described in paragraphs (c) ΔΔ=×05. A through (e) of this section, the airloads on b p the movable surfaces and the corresponding VD deflections need not exceed those which Where DP=the maximum total deflection could be obtained in flight by employing the (sum of both aileron deflections) at VA maximum limit pilot forces specified in the with VA, VC, and VD described in subpara- table in § 23.397(b). If the surface loads are graph (2) of § 23.7(e) of this appendix. limited by these maximum limit pilot forces,

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the tabs must either be considered to be de- gusts, the most severe loads must be consid- flected to their maximum travel in the direc- ered in association with every center of pres- tion which would assist the pilot or the de- sure position between the leading edge and flection must correspond to the maximum the half chord of the mean chord of the sur- degree of ‘‘out of trim’’ expected at the speed face (stabilizer and elevator, or fin and rud- for the condition under consideration. The der). tab load, however, need not exceed the value (iv) To ensure adequate strength under specified in Table 2 of this appendix. high leading edge loads, the most severe sta- (c) Surface loading conditions. Each surface bilizer and fin loads must be further consid- loading condition must be investigated as ered as being increased by 50 percent over follows: the leading 10 percent of the chord with the (1) Simplified limit surface loadings for the loads aft of this appropriately decreased to horizontal tail, vertical tail, aileron, wing retain the same total load. flaps, and trim tabs are specified in figures 5 (v) The most severe elevator and rudder and 6 of this appendix. loads should be further considered as being (i) The distribution of load along the span distributed parabolically from three times of the surface, irrespective of the chordwise the mean loading of the surface (stabilizer load distribution, must be assumed propor- and elevator, or fin and rudder) at the lead- tional to the total chord, except on horn bal- ing edge of the elevator and rudder, respec- ance surfaces. tively, to zero at the trailing edge according (ii) The load on the stabilizer and elevator, to the equation: and the load on fin and rudder, must be dis- tributed chordwise as shown in figure 7 of ()cx− 2 this appendix. Px()= 3 ( w ) (iii) In order to ensure adequate torsional 2 strength and to account for maneuvers and cf

Where— (vi) The chordwise loading distribution for P(x)=local pressure at the chordwise stations ailerons, wing flaps, and trim tabs are speci- x, fied in Table 2 of this appendix. (2) If certification in the acrobatic cat- c=chord length of the tail surface, egory is desired, the horizontal tail must be cf=chord length of the elevator and rudder investigated for an unsymmetrical load of respectively, and 100 percent w on one side of the airplane cen- w˘ ≤=average surface loading as specified in terline and 50 percent on the other side of Figure A5. the airplane centerline. (d) Outboard fins. Outboard fins must meet the requirements of § 23.445.

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(e) Special devices. Special devices must (b) Dual controls. If there are dual controls, meet the requirements of § 23.459. the systems must be designed for pilots oper- ating in opposition, using individual pilot A23.13 Control system loads. loads equal to 75 percent of those obtained in accordance with paragraph (a) of this sec- (a) Primary flight controls and systems. Each tion, except that individual pilot loads may primary flight control and system must be not be less than the minimum limit pilot designed as follows: forces shown in the table in § 23.397(b). (1) The flight control system and its sup- (c) Ground gust conditions. Ground gust con- porting structure must be designed for loads ditions must meet the requirements of corresponding to 125 percent of the computed § 23.415. hinge moments of the movable control sur- (d) Secondary controls and systems. Sec- face in the conditions prescribed in A23.11 of ondary controls and systems must meet the this appendix. In addition— requirements of § 23.405. (i) The system limit loads need not exceed those that could be produced by the pilot and TABLE 1—LIMIT FLIGHT LOAD FACTORS automatic devices operating the controls; [Limit flight load factors] and (ii) The design must provide a rugged sys- Flight load factors Normal Utility cat- Acrobatic tem for service use, including jamming, category egory category ground gusts, taxiing downwind, control in- Flaps up: ertia, and friction. n1 ...... 3.8 4.4 6.0 (2) Acceptable maximum and minimum n2 ...... ¥0.5 n1 ...... 1 limit pilot forces for elevator, aileron, and n3 ...... ( ) ...... 2 rudder controls are shown in the table in n4 ...... ( ) ...... § 23.397(b). These pilots loads must be as- Flaps down: n flap ...... 0.5 n ...... sumed to act at the appropriate control grips 1 n flap ...... 3 Zero ...... or pads as they would under flight condi- 1 tions, and to be reacted at the attachments Find n3 from Fig. 1 2 Find n4 from Fig. 2 of the control system to the control surface 3 Vertical wing load may be assumed equal to zero and only horn. the flap part of the wing need be checked for this condition.

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FIGURE A7—CHORDWISE LOAD DISTRIBUTION FOR STABILIZER AND ELEVATOR OR FIN AND RUDDER

−− from stabilizer (or fin) leading edge to the = (')23Ed local chord. Sign convention is positive Pw1 2 () when center of pressure is behind leading ()1− E edge. c=local chord. =+− PwdE2 23()(' 1 ) NOTE: Positive values of w¯ , P1 and P2 are where: all measured in the same direction. w¯ =average surface loading (as specified in [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as figure A.5) amended by Amdt. 23–7, 34 FR 13097, Aug. 13, E=ratio of elevator (or rudder) chord to total 1969; 34 FR 14727, Sept. 24, 1969; Amdt. 23–16, stabilizer and elevator (or fin and rudder) 40 FR 2577, Jan. 14, 1975; Amdt. 23–28, 47 FR chord. 13315, Mar. 29, 1982; Amdt. 23–48, 61 FR 5149, d′=ratio of distance of center of pressure of a Feb. 9, 1996] unit spanwise length of combined stabilizer and elevator (or fin and rudder) measured APPENDIX B TO PART 23 [RESERVED]

APPENDIX C TO PART 23—BASIC LANDING CONDITIONS [C23.1 Basic landing conditions]

Tail wheel type Nose wheel type Level landing Condition Tail-down land- Level landing with nose wheel Tail-down land- Level landing ing with inclined just clear of ing reactions ground

Reference section ...... 23.479(a)(1) 23.481(a)(1) .... 23.479(a)(2)(i) 23.479(a)(2)(ii) ... 23.481(a)(2) and (b).

Vertical component at c. g ...... nW ...... nW ...... nW ...... nW ...... nW. Fore and aft component at c. g ...... KnW ...... 0 ...... KnW ...... KnW ...... 0. Lateral component in either direction 0 ...... 0 ...... 0 ...... 0 ...... 0. at c. g. Shock absorber extension (hydraulic Note (2) ...... Note (2) ...... Note (2) ...... Note (2) ...... Note (2). shock absorber). Shock absorber deflection (rubber or 100 ...... 100 ...... 100 ...... 100 ...... 100. spring shock absorber), percent. Tire deflection ...... Static ...... Static ...... Static ...... Static ...... Static. Main wheel loads (both wheels) (Vr) (n-L)W ...... (n-L)W b/d ...... (n-L)W a′/d′ ..... (n-L)W ...... (n-L)W. Main wheel loads (both wheels) (Dr) KnW ...... 0 ...... KnW a′/d′ ...... KnW ...... 0.

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[C23.1 Basic landing conditions]

Tail wheel type Nose wheel type Level landing Condition Tail-down land- Level landing with nose wheel Tail-down land- Level landing ing with inclined just clear of ing reactions ground

Tail (nose) wheel loads (Vf) ...... 0 ...... (n-L)W a/d ...... (n-L)W b′/d′ ..... 0 ...... 0. Tail (nose) wheel loads (Df) ...... 0 ...... 0 ...... KnW b′/d′ ...... 0 ...... 0. Notes ...... (1), (3), and (4) ...... (1) ...... (1), (3), and (4) .. (3) and (4). (4).

NOTE (1). K may be determined as follows: K=0.25 for W=3,000 pounds or less; K=0.33 for W=6,000 pounds or greater, with linear variation of K between these weights. NOTE (2). For the purpose of design, the maximum load factor is assumed to occur throughout the shock absorber stroke from 25 percent deflection to 100 percent deflection unless otherwise shown and the load factor must be used with whatever shock absorber extension is most critical for each element of the landing gear. NOTE (3). Unbalanced moments must be balanced by a rational or conservative method. NOTE (4). L is defined in § 23.735(b). NOTE (5). n is the limit inertia load factor, at the c.g. of the airplane, selected under § 23.473 (d), (f), and (g).

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13099, Aug. 13, 1969]

APPENDIX D TO PART 23—WHEEL SPIN- FHmax=1/re √ 2Iw(VH—Vc)nFVmax/tS UP AND SPRING-BACK LOADS where—

FHmax=maximum rearward horizontal force D23.1 Wheel spin-up loads. acting on the wheel (in pounds); (a) The following method for determining re=effective rolling radius of wheel under im- wheel spin-up loads for landing conditions is pact based on recommended operating tire based on NACA T.N. 863. However, the drag pressure (which may be assumed to be component used for design may not be less equal to the rolling radius under a static than the drag load prescribed in § 23.479(b). load of njWe) in feet; 351

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Iw=rotational mass moment of inertia of (a) Conditioning. Specimens must be condi- rolling assembly (in slug feet); tioned to 70 degrees F, plus or minus 5 de- VH=linear velocity of airplane parallel to grees, and at 50 percent plus or minus 5 per- ground at instant of contact (assumed to cent relative humidity until moisture equi- be 1.2 VS0, in feet per second); librium is reached or for 24 hours. Only one Vc=peripheral speed of tire, if prerotation is specimen at a time may be removed from the used (in feet per second) (there must be a conditioning environment immediately be- positive means of pre-rotation before pre- fore subjecting it to the flame. rotation may be considered); (b) Specimen configuration. Except as pro- n=equals effective coefficient of friction (0.80 vided for materials used in electrical wire may be used); and cable insulation and in small parts, ma- terials must be tested either as a section cut F =maximum vertical force on wheel Vmax from a fabricated part as installed in the air- (pounds)=n W where W and n are defined j e, e j plane or as a specimen simulating a cut sec- in § 23.725; tion, such as: a specimen cut from a flat ts=time interval between ground contact and sheet of the material or a model of the fab- attainment of maximum vertical force on ricated part. The specimen may be cut from wheel (seconds). (However, if the value of any location in a fabricated part; however, FVmax, from the above equation exceeds 0.8 fabricated units, such as sandwich panels, FVmax, the latter value must be used for may not be separated for a test. The speci- FHmax.) men thickness must be no thicker than the (b) The equation assumes a linear vari- minimum thickness to be qualified for use in ation of load factor with time until the peak the airplane, except that: (1) Thick foam load is reached and under this assumption, parts, such as seat cushions, must be tested the equation determines the drag force at in 1⁄2 inch thickness; (2) when showing com- the time that the wheel peripheral velocity pliance with § 23.853(d)(3)(v) for materials used in small parts that must be tested, the at radius re equals the airplane velocity. Most shock absorbers do not exactly follow a materials must be tested in no more than 1⁄8 linear variation of load factor with time. inch thickness; (3) when showing compliance Therefore, rational or conservative allow- with § 23.1359(c) for materials used in elec- ances must be made to compensate for these trical wire and cable insulation, the wire and variations. On most landing gears, the time cable specimens must be the same size as for wheel spin-up will be less than the time used in the airplane. In the case of fabrics, required to develop maximum vertical load both the warp and fill direction of the weave factor for the specified rate of descent and must be tested to determine the most crit- forward velocity. For exceptionally large ical flammability conditions. When per- wheels, a wheel peripheral velocity equal to forming the tests prescribed in paragraphs the ground speed may not have been attained (d) and (e) of this appendix, the specimen at the time of maximum vertical gear load. must be mounted in a metal frame so that (1) However, as stated above, the drag spin-up in the vertical tests of paragraph (d) of this load need not exceed 0.8 of the maximum appendix, the two long edges and the upper vertical loads. edge are held securely; (2) in the horizontal test of paragraph (e) of this appendix, the (c) Dynamic spring-back of the landing two long edges and the edge away from the gear and adjacent structure at the instant flame are held securely; (3) the exposed area just after the wheels come up to speed may of the specimen is at least 2 inches wide and result in dynamic forward acting loads of 12 inches long, unless the actual size used in considerable magnitude. This effect must be the airplane is smaller; and (4) the edge to determined, in the level landing condition, which the burner flame is applied must not by assuming that the wheel spin-up loads consist of the finished or protected edge of calculated by the methods of this appendix the specimen but must be representative of are reversed. Dynamic spring-back is likely the actual cross section of the material or to become critical for landing gear units part installed in the airplane. When per- having wheels of large mass or high landing forming the test prescribed in paragraph (f) speeds. of this appendix, the specimen must be [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as mounted in metal frame so that all four amended by Amdt. 23–45, 58 FR 42167, Aug. 6, edges are held securely and the exposed area 1993] of the specimen is at least 8 inches by 8 inches. APPENDIX E TO PART 23 [RESERVED] (c) Apparatus. Except as provided in para- graph (g) of this appendix, tests must be con- APPENDIX F TO PART 23—TEST ducted in a draft-free cabinet in accordance PROCEDURE with Federal Test Method Standard 191 Method 5903 (revised Method 5902) which is Acceptable test procedure for self-extin- available from the General Services Admin- guishing materials for showing compliance istration, Business Service Center, Region 3, with §§ 23.853, 23.855 and 23.1359. Seventh and D Streets SW., Washington,

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D.C. 20407, or with some other approved tacting the material at the center of the equivalent method. Specimens which are too specimen and then removed. Flame time, large for the cabinet must be tested in simi- glow time, and whether the flame penetrates lar draft-free conditions. (passes through) the specimen must be re- (d) Vertical test. A minimum of three speci- corded. mens must be tested and the results aver- (g) Sixty-degree test. A minimum of three aged. For fabrics, the direction of weave cor- specimens of each wire specification (make responding to the most critical flammability and size) must be tested. The specimen of conditions must be parallel to the longest di- wire or cable (including insulation) must be mension. Each specimen must be supported placed at an angle of 60 degrees with the hor- vertically. The specimen must be exposed to izontal in the cabinet specified in paragraph 3 a Bunsen or Tirrill burner with a nominal ⁄8- (c) of this appendix, with the cabinet door 1 inch I.D. tube adjusted to give a flame of 1 ⁄2 open during the test or placed within a inches in height. The minimum flame tem- chamber approximately 2 feet high × 1 foot × perature measured by a calibrated thermo- 1 foot, open at the top and at one vertical couple pryometer in the center of the flame side (front), that allows sufficient flow of air must be 1550 °F. The lower edge of the speci- for complete combustion but is free from men must be three-fourths inch above the drafts. The specimen must be parallel to and top edge of the burner. The flame must be approximately 6 inches from the front of the applied to the center line of the lower edge of the specimen. For materials covered by chamber. The lower end of the specimen §§ 23.853(d)(3)(i) and 23.853(f), the flame must must be held rigidly clamped. The upper end be applied for 60 seconds and then removed. of the specimen must pass over a pulley or For materials covered by § 23.853(d)(3)(ii), the rod and must have an appropriate weight at- flame must be applied for 12 seconds and tached to it so that the specimen is held then removed. Flame time, burn length, and tautly throughout the flammability test. flaming time of drippings, if any, must be re- The test specimen span between lower clamp corded. The burn length determined in ac- and upper pulley or rod must be 24 inches cordance with paragraph (h) of this appendix and must be marked 8 inches from the lower must be measured to the nearest one-tenth end to indicate the central point for flame inch. application. A flame from a Bunsen or Tirrill (e) Horizontal test. A minimum of three burner must be applied for 30 seconds at the specimens must be tested and the results test mark. The burner must be mounted un- averaged. Each specimen must be supported derneath the test mark on the specimen, per- horizontally. The exposed surface when in- pendicular to the specimen and at an angle stalled in the airplane must be face down for of 30 degrees to the vertical plane of the the test. The specimen must be exposed to a specimen. The burner must have a nominal Bunsen burner or Tirrill burner with a nomi- bore of three-eighths inch, and must be ad- nal 3⁄8-inch I.D. tube adjusted to give a flame justed to provide a three-inch-high flame of 11⁄2 inches in height. The minimum flame with an inner cone approximately one-third temperature measured by a calibrated ther- of the flame height. The minimum tempera- mocouple pyrometer in the center of the ture of the hottest portion of the flame, as flame must be 1550 °F. The specimen must be measured with a calibrated thermocouple positioned so that the edge being tested is pyrometer, may not be less than 1,750 °F. The three-fourths of an inch above the top of, and burner must be positioned so that the hot- on the center line of, the burner. The flame test portion of the flame is applied to the must be applied for 15 seconds and then re- test mark on the wire. Flame time, burn moved. A minimum of 10 inches of the speci- length, and flaming time drippings, if any, men must be used for timing purposes, ap- must be recorded. The burn length deter- proximately 11⁄2 inches must burn before the mined in accordance with paragraph (h) of burning front reaches the timing zone, and this appendix must be measured to the near- the average burn rate must be recorded. est one-tenth inch. Breaking of the wire (f) Forty-five degree test. A minimum of specimen is not considered a failure. three specimens must be tested and the re- (h) Burn length. Burn length is the distance sults averaged. The specimens must be sup- from the original edge to the farthest evi- ported at an angle of 45 degrees to a hori- dence of damage to the test specimen due to zontal surface. The exposed surface when in- flame impingement, including areas of par- stalled in the aircraft must be face down for tial or complete consumption, charring, or the test. The specimens must be exposed to embrittlement, but not including areas soot- 3 a Bunsen or Tirrill burner with a nominal ⁄8 ed, stained, warped, or discolored, nor areas 1 inch I.D. tube adjusted to give a flame of 1 ⁄2 where material has shrunk or melted away inches in height. The minimum flame tem- from the heat source. perature measured by a calibrated thermo- couple pyrometer in the center of the flame [Amdt. 23–23, 43 FR 50594, Oct. 30, 1978, as must be 1550 °F. Suitable precautions must amended by Amdt. 23–34, 52 FR 1835, Jan. 15, be taken to avoid drafts. The flame must be 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23–49, applied for 30 seconds with one-third con- 61 FR 5170, Feb. 9, 1996]

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APPENDIX G TO PART 23—INSTRUCTIONS which they should be cleaned, inspected, ad- FOR CONTINUED AIRWORTHINESS justed, tested, and lubricated, and the degree of inspection, the applicable wear tolerances, G23.1 General. (a) This appendix specifies and work recommended at these periods. requirements for the preparation of Instruc- However, the applicant may refer to an ac- tions for Continued Airworthiness as re- cessory, instrument, or equipment manufac- quired by § 23.1529. turer as the source of this information if the (b) The Instructions for Continued Air- applicant shows that the item has an excep- worthiness for each airplane must include tionally high degree of complexity requiring the Instructions for Continued Airworthiness for each engine and propeller (hereinafter specialized maintenance techniques, test designated ‘products’), for each appliance re- equipment, or expertise. The recommended quired by this chapter, and any required in- overhaul periods and necessary cross ref- formation relating to the interface of those erence to the Airworthiness Limitations sec- appliances and products with the airplane. If tion of the manual must also be included. In Instructions for Continued Airworthiness are addition, the applicant must include an in- not supplied by the manufacturer of an ap- spection program that includes the fre- pliance or product installed in the airplane, quency and extent of the inspections nec- the Instructions for Continued Airworthiness essary to provide for the continued air- for the airplane must include the informa- worthiness of the airplane. tion essential to the continued airworthiness (2) Troubleshooting information describing of the airplane. probable malfunctions, how to recognize (c) The applicant must submit to the FAA those malfunctions, and the remedial action a program to show how changes to the In- for those malfunctions. structions for Continued Airworthiness made by the applicant or by the manufacturers of (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 G23.2 Format. (a) The Instructions for be taken. Continued Airworthiness must be in the (4) Other general procedural instructions form of a manual or manuals as appropriate including procedures for system testing dur- for the quantity of data to be provided. ing ground running, symmetry checks, (b) The format of the manual or manuals weighing and determining the center of grav- must provide for a practical arrangement. ity, lifting and shoring, and storage limita- G23.3 Content. The contents of the manual tions. or manuals must be prepared in the English (c) Diagrams of structural access plates language. The Instructions for Continued and information needed to gain access for in- Airworthiness must contain the following spections when access plates are not pro- manuals or sections, as appropriate, and in- vided. formation: (a) Airplane maintenance manual or section. (d) Details for the application of special in- (1) Introduction information that includes an spection techniques including radiographic explanation of the airplane’s features and and ultrasonic testing where such processes data to the extent necessary for mainte- are specified. nance or preventive maintenance. (e) Information needed to apply protective (2) A description of the airplane and its treatments to the structure after inspection. systems and installations including its en- (f) All data relative to structural fasteners gines, propellers, and appliances. such as identification, discard recommenda- (3) Basic control and operation information tions, and torque values. describing how the airplane components and (g) A list of special tools needed. systems are controlled and how they oper- (h) In addition, for commuter category air- ate, including any special procedures and planes, the following information must be limitations that apply. furnished: (4) Servicing information that covers de- tails regarding servicing points, capacities of (1) Electrical loads applicable to the var- tanks, reservoirs, types of fluids to be used, ious systems; pressures applicable to the various systems, (2) Methods of balancing control surfaces; location of access panels for inspection and (3) Identification of primary and secondary servicing, locations of lubrication points, lu- structures; and bricants to be used, equipment required for (4) Special repair methods applicable to servicing, tow instructions and limitations, the airplane. mooring, jacking, and leveling information. G23.4 Airworthiness Limitations section. The (b) Maintenance instructions. (1) Scheduling Instructions for Continued Airworthiness information for each part of the airplane and must contain a section titled Airworthiness its engines, auxiliary power units, propellers, Limitations that is segregated and clearly accessories, instruments, and equipment that provides the recommended periods at

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distinguishable from the rest of the docu- requirements (except as provided in this ap- ment. This section must set forth each man- pendix) must be met without requiring any datory replacement time, structural inspec- action by the crew to increase power or tion interval, and related structural inspec- thrust. tion procedure required for type certifi- H23.2, Definitions. cation. If the Instructions for Continued Air- (a) Automatic power reserve system means worthiness consist of multiple documents, the entire automatic system used only dur- the section required by this paragraph must ing takeoff, including all devices both me- be included in the principal manual. This chanical and electrical that sense engine section must contain a legible statement in failure, transmit signals, actuate fuel con- a prominent location that reads: ‘‘The Air- trols or power levers on operating engines, worthiness Limitations section is FAA ap- proved and specifies maintenance required including power sources, to achieve the under §§ 43.16 and 91.403 of the Federal Avia- scheduled power increase and furnish cockpit tion Regulations unless an alternative pro- information on system operation. gram has been FAA approved.’’ (b) Selected takeoff power, notwithstanding the definition of ‘‘Takeoff Power’’ in part 1 of [Amdt. 23–26, 45 FR 60171, Sept. 11, 1980, as the Federal Aviation Regulations, means the amended by Amdt. 23–34, 52 FR 1835, Jan. 15, power obtained from each initial power set- 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23–37, ting approved for takeoff. 54 FR 34329, Aug. 18, 1989] (c) Critical Time Interval, as illustrated in figure H1, means that period starting at V APPENDIX H TO PART 23—INSTALLATION 1 minus one second and ending at the intersec- OF AN AUTOMATIC POWER RESERVE tion of the engine and APR failure flight (APR) SYSTEM path line with the minimum performance all H23.1, General. engine flight path line. The engine and APR (a) This appendix specifies requirements failure flight path line intersects the one-en- for installation of an APR engine power con- gine-inoperative flight path line at 400 feet trol system that automatically advances above the takeoff surface. The engine and power or thrust on the operating engine(s) in APR failure flight path is based on the air- the event any engine fails during takeoff. plane’s performance and must have a posi- (b) With the APR system and associated tive gradient of at least 0.5 percent at 400 systems functioning normally, all applicable feet above the takeoff surface.

H23.3, Reliability and performance require- gine will not create a hazard to the airplane, ments. or it must be shown that such failures are (a) It must be shown that, during the crit- improbable. ical time interval, an APR failure that in- (b) It must be shown that, during the crit- creases or does not affect power on either en- ical time interval, there are no failure modes

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of the APR system that would result in a tainable takeoff power without exceeding en- failure that will decrease the power on either gine operating limits; engine or it must be shown that such failures (3) Prevent deactivation of the APR by are extremely improbable. manual adjustment of the power levers fol- (c) It must be shown that, during the crit- lowing an engine failure; ical time interval, there will be no failure of (4) Provide a means for the flight crew to the APR system in combination with an en- deactivate the automatic function. This gine failure or it must be shown that such means must be designed to prevent inad- failures are extremely improbable. vertent deactivation; and (d) All applicable performance require- (5) Allow normal manual decrease or in- ments must be met with an engine failure crease in power up to the maximum takeoff occurring at the most critical point during power approved for the airplane under the takeoff with the APR system functioning existing conditions through the use of power normally. levers, as stated in § 23.1141(c), except as pro- H23.4, Power setting. vided under paragraph (c) of H23.5 of this ap- The selected takeoff power set on each en- pendix. gine at the beginning of the takeoff roll may (c) For airplanes equipped with limiters not be less than— that automatically prevent engine operating (a) The power necessary to attain, at V1, 90 limits from being exceeded, other means percent of the maximum takeoff power ap- may be used to increase the maximum level proved for the airplane for the existing con- of power controlled by the power levers in ditions; the event of an APR failure. The means must (b) That required to permit normal oper- be located on or forward of the power levers, ation of all safety-related systems and equip- must be easily identified and operated under ment that are dependent upon engine power all operating conditions by a single action of or power lever position; and any pilot with the hand that is normally (c) That shown to be free of hazardous en- used to actuate the power levers, and must gine response characteristics when power is meet the requirements of § 23.777 (a), (b), and advanced from the selected takeoff power (c). level to the maximum approved takeoff H23.6, Powerplant instruments. power. In addition to the requirements of § 23.1305: H23.5, Powerplant controls—general. (a) A means must be provided to indicate (a) In addition to the requirements of when the APR is in the armed or ready con- § 23.1141, no single failure or malfunction (or dition. probable combination thereof) of the APR, (b) If the inherent flight characteristics of including associated systems, may cause the the airplane do not provide warning that an failure of any powerplant function necessary engine has failed, a warning system inde- for safety. pendent of the APR must be provided to give (b) The APR must be designed to— the pilot a clear warning of any engine fail- (1) Provide a means to verify to the flight ure during takeoff. crew before takeoff that the APR is in an op- (c) Following an engine failure at V or erating condition to perform its intended 1 above, there must be means for the crew to function; readily and quickly verify that the APR has (2) Automatically advance power on the operated satisfactorily. operating engines following an engine failure during takeoff to achieve the maximum at- [Doc. 26344, 58 FR 18979, Apr. 9, 1993]

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APPENDIX I TO PART 23—SEAPLANE LOADS

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[Amdt. 23–45, 58 FR 42167, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

APPENDIX J TO PART 23—HIRF ENVI- test levels are expressed in root-mean-square RONMENTS AND EQUIPMENT HIRF units measured during the peak of the modu- TEST LEVELS lation cycle. (a) HIRF environment I is specified in the This appendix specifies the HIRF environ- following table: ments and equipment HIRF test levels for electrical and electronic systems under § 23.1308. The field strength values for the HIRF environments and equipment HIRF

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

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