23.143 General.23.145 Longitudinal control.23.147 Directional and lateral control.23.149 Minimum control speed.23.151 Acrobatic maneuvers.23.153 Control during landings.
23.155 Elevator control force in maneuvers.23.157 Rate of roll.
TRIM
23.161 Trim.
STABILITY
23.171 General.23.173 Static longitudinal stability.23.175 Demonstration of static longitudinal
stability.23.177 Static directional and lateral sta-
bility.23.181 Dynamic stability.
STALLS
23.201 Wings level stall.23.203 Turning flight and accelerated turn-
ing stalls.23.207 Stall warning.
SPINNING
23.221 Spinning.
GROUND AND WATER HANDLINGCHARACTERISTICS
23.231 Longitudinal stability and control.23.233 Directional stability and control.23.235 Operation on unpaved surfaces.23.237 Operation on water.23.239 Spray characteristics.
MISCELLANEOUS FLIGHT REQUIREMENTS
23.251 Vibration and buffeting.23.253 High speed characteristics.
Subpart C—Structure
GENERAL
23.301 Loads.23.302 Canard or tandem wing configura-
tions.23.303 Factor of safety.23.305 Strength and deformation.23.307 Proof of structure.
FLIGHT LOADS
23.321 General.23.331 Symmetrical flight conditions.23.333 Flight envelope.23.335 Design airspeeds.23.337 Limit maneuvering load factors.23.341 Gust loads factors.23.343 Design fuel loads.23.345 High lift devices.23.347 Unsymmetrical flight conditions.23.349 Rolling conditions.23.351 Yawing conditions.23.361 Engine torque.23.363 Side load on engine mount.23.365 Pressurized cabin loads.23.367 Unsymmetrical loads due to engine
failure.23.369 Rear lift truss.23.371 Gyroscopic and aerodynamic loads.
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23.373 Speed control devices.
CONTROL SURFACE AND SYSTEM LOADS
23.391 Control surface loads.23.393 Loads parallel to hinge line.23.395 Control system loads.23.397 Limit control forces and torques.23.399 Dual control system.23.405 Secondary control system.23.407 Trim tab effects.23.409 Tabs.23.415 Ground gust conditions.
23.571 Metallic pressurized cabin structures.23.572 Metallic wing, empennage, and asso-
ciated structures.23.573 Damage tolerance and fatigue evalua-
tion of structure.23.574 Metallic damage tolerance and fa-
tigue evaluation of commuter categoryairplanes.
23.575 Inspections and other procedures.
Subpart D—Design and Construction
23.601 General.23.603 Materials and workmanship.23.605 Fabrication methods.23.607 Fasteners.23.609 Protection of structure.23.611 Accessibility provisions.23.613 Material strength properties and de-
23.751 Main float buoyancy.23.753 Main float design.23.755 Hulls.23.757 Auxiliary floats.
PERSONNEL AND CARGO ACCOMMODATIONS
23.771 Pilot compartment.23.773 Pilot compartment view.23.775 Windshields and windows.23.777 Cockpit controls.23.779 Motion and effect of cockpit controls.23.781 Cockpit control knob shape.23.783 Doors.23.785 Seats, berths, litters, safety belts,
and shoulder harnesses.23.787 Baggage and cargo compartments.23.791 Passenger information signs.23.803 Emergency evacuation.23.805 Flightcrew emergency exits.23.807 Emergency exits.23.811 Emergency exit marking.23.812 Emergency lighting.23.813 Emergency exit access.23.815 Width of aisle.23.831 Ventilation.
23.951 General.23.953 Fuel system independence.23.954 Fuel system lightning protection.23.955 Fuel flow.23.957 Flow between interconnected tanks.23.959 Unusable fuel supply.23.961 Fuel system hot weather operation.23.963 Fuel tanks: General.23.965 Fuel tank tests.23.967 Fuel tank installation.23.969 Fuel tank expansion space.23.971 Fuel tank sump.23.973 Fuel tank filler connection.23.975 Fuel tank vents and carburetor vapor
vents.23.977 Fuel tank outlet.23.979 Pressure fueling systems.
FUEL SYSTEM COMPONENTS
23.991 Fuel pumps.23.993 Fuel system lines and fittings.23.994 Fuel system components.23.995 Fuel valves and controls.23.997 Fuel strainer or filter.23.999 Fuel system drains.23.1001 Fuel jettisoning system.
OIL SYSTEM
23.1011 General.23.1013 Oil tanks.23.1015 Oil tank tests.23.1017 Oil lines and fittings.23.1019 Oil strainer or filter.23.1021 Oil system drains.23.1023 Oil radiators.23.1027 Propeller feathering system.
COOLING
23.1041 General.23.1043 Cooling tests.23.1045 Cooling test procedures for turbine
engine powered airplanes.23.1047 Cooling test procedures for recipro-
cating engine powered airplanes.
LIQUID COOLING
23.1061 Installation.23.1063 Coolant tank tests.
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INDUCTION SYSTEM
23.1091 Air induction system.23.1093 Induction system icing protection.23.1095 Carburetor deicing fluid flow rate.23.1097 Carburetor deicing fluid system ca-
pacity.23.1099 Carburetor deicing fluid system de-
tail design.23.1101 Induction air preheater design.23.1103 Induction system ducts.23.1105 Induction system screens.23.1107 Induction system filters.23.1109 Turbocharger bleed air system.23.1111 Turbine engine bleed air system.
23.1141 Powerplant controls: General.23.1142 Auxiliary power unit controls.23.1143 Engine controls.23.1145 Ignition switches.23.1147 Mixture controls.23.1149 Propeller speed and pitch controls.23.1153 Propeller feathering controls.23.1155 Turbine engine reverse thrust and
propeller pitch settings below the flightregime.
23.1157 Carburetor air temperature controls.23.1163 Powerplant accessories.23.1165 Engine ignition systems.
POWERPLANT FIRE PROTECTION
23.1181 Designated fire zones; regions in-cluded.
23.1182 Nacelle areas behind firewalls.23.1183 Lines, fittings, and components.23.1189 Shutoff means.23.1191 Firewalls.23.1192 Engine accessory compartment dia-
phragm.23.1193 Cowling and nacelle.23.1195 Fire extinguishing systems.23.1197 Fire extinguishing agents.23.1199 Extinguishing agent containers.23.1201 Fire extinguishing systems mate-
rials.23.1203 Fire detector system.
Subpart F—Equipment
GENERAL
23.1301 Function and installation.23.1303 Flight and navigation instruments.23.1305 Powerplant instruments.23.1307 Miscellaneous equipment.23.1309 Equipment, systems, and installa-
tions.
INSTRUMENTS: INSTALLATION
23.1311 Electronic display instrument sys-tems.
23.1321 Arrangement and visibility.23.1322 Warning, caution, and advisory
lights.23.1323 Airspeed indicating system.23.1325 Static pressure system.23.1326 Pitot heat indication systems.23.1327 Magnetic direction indicator.23.1329 Automatic pilot system.23.1331 Instruments using a power source.23.1335 Flight director systems.23.1337 Powerplant instruments installa-
tion.
ELECTRICAL SYSTEMS AND EQUIPMENT
23.1351 General.23.1353 Storage battery design and installa-
tion.23.1357 Circuit protective devices.23.1359 Electrical system fire protection.23.1361 Master switch arrangement.23.1365 Electric cables and equipment.23.1367 Switches.
LIGHTS
23.1381 Instrument lights.23.1383 Taxi and landing lights.23.1385 Position light system installation.23.1387 Position light system dihedral an-
gles.23.1389 Position light distribution and in-
tensities.23.1391 Minimum intensities in the hori-
zontal plane of position lights.23.1393 Minimum intensities in any vertical
plane of position lights.23.1395 Maximum intensities in overlapping
beams of position lights.23.1397 Color specifications.23.1399 Riding light.23.1401 Anticollision light system.
23.1431 Electronic equipment.23.1435 Hydraulic systems.23.1437 Accessories for multiengine air-
planes.23.1438 Pressurization and pneumatic sys-
tems.23.1441 Oxygen equipment and supply.23.1443 Minimum mass flow of supplemental
oxygen.23.1445 Oxygen distribution system.23.1447 Equipment standards for oxygen dis-
pensing units.23.1449 Means for determining use of oxy-
gen.23.1450 Chemical oxygen generators.23.1451 Fire protection for oxygen equip-
ment.
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23.1453 Protection of oxygen equipmentfrom rupture.
23.1457 Cockpit voice recorders.23.1459 Flight recorders.23.1461 Equipment containing high energy
rotors.
Subpart G—Operating Limitations andInformation
23.1501 General.23.1505 Airspeed limitations.23.1507 Operating maneuvering speed.23.1511 Flap extended speed.23.1513 Minimum control speed.23.1519 Weight and center of gravity.23.1521 Powerplant limitations.23.1522 Auxiliary power unit limitations.23.1523 Minimum flight crew.23.1524 Maximum passenger seating configu-
ration.23.1525 Kinds of operation.23.1527 Maximum operating altitude.23.1529 Instructions for Continued Air-
worthiness.
MARKINGS AND PLACARDS
23.1541 General.23.1543 Instrument markings: General.23.1545 Airspeed indicator.23.1547 Magnetic direction indicator.23.1549 Powerplant and auxiliary power unit
instruments.23.1551 Oil quantity indicator.23.1553 Fuel quantity indicator.23.1555 Control markings.23.1557 Miscellaneous markings and plac-
1. Applicability. An applicant is entitled toa type certificate in the normal category fora reciprocating or turbopropeller multien-gine powered small airplane that is to be cer-tificated to carry more than 10 occupantsand that is intended for use in operationsunder Part 135 of the Federal Aviation Regu-lations if he shows compliance with the ap-plicable requirements of Part 23 of the Fed-eral Aviation Regulations, as supplementedor modified by the additional airworthinessrequirements of this regulation.
2. References. Unless otherwise provided, allreferences in this regulation to specific sec-tions of Part 23 of the Federal Aviation Reg-ulations are those sections of Part 23 in ef-fect on March 30, 1967.
FLIGHT REQUIREMENTS
3. General. Compliance must be shown withthe applicable requirements of Subpart B ofPart 23 of the Federal Aviation Regulationsin effect on March 30, 1967, as supplementedor modified in sections 4 through 10 of thisregulation.
PERFORMANCE
4. General. (a) Unless otherwise prescribedin this regulation, compliance with each ap-plicable performance requirement in sections4 through 7 of this regulation must be shownfor ambient atmospheric conditions and stillair.
(b) The performance must correspond tothe propulsive thrust available under theparticular ambient atmospheric conditionsand the particular flight condition. Theavailable propulsive thrust must correspondto engine power or thrust, not exceeding theapproved power or thrust less—
(1) Installation losses; and(2) The power or equivalent thrust ab-
sorbed by the accessories and services appro-priate to the particular ambient atmosphericconditions and the particular flight condi-tion.
(c) Unless otherwise prescribed in this reg-ulation, the applicant must select the take-off, en route, and landing configurations forthe airplane.
(d) The airplane configuration may varywith weight, altitude, and temperature, tothe extent they are compatible with the op-erating procedures required by paragraph (e)of this section.
(e) Unless otherwise prescribed in this reg-ulation, in determining the critical engine
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inoperative takeoff performance, the accel-erate-stop distance, takeoff distance,changes in the airplane’s configuration,speed, power, and thrust, must be made inaccordance with procedures established bythe applicant for operation in service.
(f) Procedures for the execution of balkedlandings must be established by the appli-cant and included in the Airplane FlightManual.
(g) The procedures established under para-graphs (e) and (f) of this section must—
(1) Be able to be consistently executed inservice by a crew of average skill;
(2) Use methods or devices that are safeand reliable; and
(3) Include allowance for any time delays,in the execution of the procedures, that mayreasonably be expected in service.
5. Takeoff—(a) General. The takeoff speedsdescribed in paragraph (b), the accelerate-stop distance described in paragraph (c), andthe takeoff distance described in paragraph(d), must be determined for—
(1) Each weight, altitude, and ambienttemperature within the operational limitsselected by the applicant;
(2) The selected configuration for takeoff;(3) The center of gravity in the most unfa-
vorable position;(4) The operating engine within approved
operating limitation; and(5) Takeoff data based on smooth, dry,
hard-surface runway.(b) Takeoff speeds. (1) The decision speed V1
is the calibrated airspeed on the ground atwhich, as a result of engine failure or otherreasons, the pilot is assumed to have made adecision to continue or discontinue the take-off. The speed V1 must be selected by the ap-plicant but may not be less than—
(i) 1.10 Vs1;(ii) 1.10 VMC;(iii) A speed that permits acceleration to
V1 and stop in accordance with paragraph (c)allowing credit for an overrun distance equalto that required to stop the airplane from aground speed of 35 knots utilizing maximumbraking; or
(iv) A speed at which the airplane can berotated for takeoff and shown to be adequateto safely continue the takeoff, using normalpiloting skill, when the critical engine issuddenly made inoperative.
(2) Other essential takeoff speeds necessaryfor safe operation of the airplane must be de-termined and shown in the Airplane FlightManual.
(c) Accelerate-stop distance. (1) The accel-erate-stop distance is the sum of the dis-tances necessary to—
(i) Accelerate the airplane from a standingstart to V1; and
(ii) Decelerate the airplane from V1 to aspeed not greater than 35 knots, assumingthat in the case of engine failure, failure ofthe critical engine is recognized by the pilot
at the speed V1. The landing gear must re-main in the extended position and maximumbraking may be utilized during deceleration.
(2) Means other than wheel brakes may beused to determine the accelerate-stop dis-tance if that means is available with thecritical engine inoperative and—
(i) Is safe and reliable;(ii) Is used so that consistent results can
be expected under normal operating condi-tions; and
(iii) Is such that exceptional skill is not re-quired to control the airplane.
(d) All engines operating takeoff distance.The all engine operating takeoff distance isthe horizontal distance required to takeoffand climb to a height of 50 feet above thetakeoff surface according to procedures inFAR 23.51(a).
(e) One-engine-inoperative takeoff. The max-imum weight must be determined for eachaltitude and temperature within the oper-ational limits established for the airplane, atwhich the airplane has takeoff capabilityafter failure of the critical engine at orabove V1 determined in accordance withparagraph (b) of this section. This capabilitymay be established—
(1) By demonstrating a measurably posi-tive rate of climb with the airplane in thetakeoff configuration, landing gear extended;or
(2) By demonstrating the capability ofmaintaining flight after engine failure uti-lizing procedures prescribed by the appli-cant.
6. Climb—(a) Landing climb: All-engines-oper-ating. The maximum weight must be deter-mined with the airplane in the landing con-figuration, for each altitude, and ambienttemperature within the operational limitsestablished for the airplane and with themost unfavorable center of gravity and out-of-ground effect in free air, at which thesteady gradient of climb will not be less than3.3 percent, with:
(1) The engines at the power that is avail-able 8 seconds after initiation of movementof the power or thrust controls from themimimum flight idle to the takeoff position.
(2) A climb speed not greater than the ap-proach speed established under section 7 ofthis regulation and not less than the greaterof 1.05MC or 1.10VS1.
(b) En route climb, one-engine-inoperative. (1)the maximum weight must be determinedwith the airplane in the en route configura-tion, the critical engine inoperative, the re-maining engine at not more than maximumcontinuous power or thrust, and the mostunfavorable center of gravity, at which thegradient at climb will be not less than—
(i) 1.2 percent (or a gradient equivalent to0.20 Vso 2, if greater) at 5,000 feet and an am-bient temperature of 41° F. or
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(ii) 0.6 percent (or a gradient equivalent to0.01 Vso 2, if greater) at 5,000 feet and ambienttemperature of 81° F.
(2) The minimum climb gradient specifiedin subdivisions (i) and (ii) of subparagraph (1)of this paragraph must vary linearly between41° F. and 81° F. and must change at the samerate up to the maximum operational tem-perature approved for the airplane.
7. Landing. The landing distance must bedetermined for standard atmosphere at eachweight and altitude in accordance with FAR23.75(a), except that instead of the gliding ap-proach specified in FAR 23.75(a)(1), the land-ing may be preceded by a steady approachdown to the 50-foot height at a gradient ofdescent not greater than 5.2 percent (3°) at acalibrated airspeed not less than 1.3s1.
TRIM
8. Trim—(a) Lateral and directional trim. Theairplane must maintain lateral and direc-tional trim in level flight at a speed of Vh orVMO/MMO, whichever is lower, with landinggear and wing flaps retracted.
(b) Longitudinal trim. The airplane mustmaintain longitudinal trim during the fol-lowing conditions, except that it need notmaintain trim at a speed greater than VMO/MMO:
(1) In the approach conditions specified inFAR 23.161(c)(3) through (5), except that in-stead of the speeds specified therein, trimmust be maintained with a stick force of notmore than 10 pounds down to a speed used inshowing compliance with section 7 of thisregulation or 1.4 Vs
1 whichever is lower.(2) In level flight at any speed from VH or
VMO/MMO, whichever is lower, to either Vx or1.4 Vs1, with the landing gear and wing flapsretracted.
STABILITY
9. Static longitudinal stability. (a) In showingcompliance with the provisions of FAR23.175(b) and with paragraph (b) of this sec-tion, the airspeed must return to within ±71⁄2percent of the trim speed.
(b) Cruise stability. The stick force curvemust have a stable slope for a speed range of±50 knots from the trim speed except thatthe speeds need not exceed VFC/MFC or be lessthan 1.4 Vs1. This speed range will be consid-ered to begin at the outer extremes of thefriction band and the stick force may not ex-ceed 50 pounds with—
(i) Landing gear retracted;(ii) Wing flaps retracted;(iii) The maximum cruising power as se-
lected by the applicant as an operating limi-tation for turbine engines or 75 percent ofmaximum continuous power for recipro-cating engines except that the power neednot exceed that required at VMO/MMO:
(iv) Maximum takeoff weight; and
(v) The airplane trimmed for level flightwith the power specified in subparagraph(iii) of this paragraph.
VFC/MFC may not be less than a speed mid-way between VMO/MMO and VDF/MDF, exceptthat, for altitudes where Mach number is thelimiting factor, MFC need not exceed theMach number at which effective speed warn-ing occurs.
(c) Climb stability. For turbopropeller poweredairplanes only. In showing compliance withFAR 23.175(a), an applicant must in lieu ofthe power specified in FAR 23.175(a)(4), usethe maximum power or thrust selected bythe applicant as an operating limitation foruse during climb at the best rate of climbspeed except that the speed need not be lessthan 1.4 Vs1.
STALLS
10. Stall warning. If artificial stall warningis required to comply with the requirementsof FAR 23.207, the warning device must giveclearly distinguishable indications under ex-pected conditions of flight. The use of a vis-ual warning device that requires the atten-tion of the crew within the cockpit is not ac-ceptable by itself.
CONTROL SYSTEMS
11. Electric trim tabs. The airplane mustmeet the requirements of FAR 23.677 and inaddition it must be shown that the airplaneis safely controllable and that a pilot canperform all the maneuvers and operationsnecessary to effect a safe landing followingany probable electric trim tab runawaywhich might be reasonably expected in serv-ice allowing for appropriate time delay afterpilot recognition of the runaway. This dem-onstration must be conducted at the criticalairplane weights and center of gravity posi-tions.
INSTRUMENTS: INSTALLATION
12. Arrangement and visibility. Each instru-ment must meet the requirements of FAR23.1321 and in addition—
(a) Each flight, navigation, and powerplantinstrument for use by any pilot must beplainly visible to him from his station withthe minimum practicable deviation from hisnormal position and line of vision when he islooking forward along the flight path.
(b) The flight instruments required by FAR23.1303 and by the applicable operating rulesmust be grouped on the instrument paneland centered as nearly as practicable aboutthe vertical plane of each pilot’s forward vi-sion. In addition—
(1) The instrument that most effectivelyindicates the attitude must be on the panelin the top center position;
(2) The instrument that most effectivelyindicates airspeed must be adjacent to and
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directly to the left of the instrument in thetop center position;
(3) The instrument that most effectivelyindicates altitude must be adjacent to anddirectly to the right of the instrument in thetop center position; and
(4) The instrument that most effectivelyindicates direction of flight must be adjacentto and directly below the instrument in thetop center position.
13. Airspeed indicating system. Each airspeedindicating system must meet the require-ments of FAR 23.1323 and in addition—
(a) Airspeed indicating instruments mustbe of an approved type and must be cali-brated to indicate true airspeed at sea levelin the standard atmosphere with amimimum practicable instrument calibra-tion error when the corresponding pilot andstatic pressures are supplied to the instru-ments.
(b) The airspeed indicating system must becalibrated to determine the system error,i.e., the relation between IAS and CAS, inflight and during the accelerate takeoffground run. The ground run calibration mustbe obtained between 0.8 of the mimimumvalue of V1 and 1.2 times the maximum valueof V1, considering the approved ranges of al-titude and weight. The ground run calibra-tion will be determined assuming an enginefailure at the mimimum value of V1.
(c) The airspeed error of the installationexcluding the instrument calibration error,must not exceed 3 percent or 5 knots which-ever is greater, throughout the speed rangefrom VMO to 1.3S1 with flaps retracted andfrom 1.3 VSO to VFE with flaps in the landingposition.
(d) Information showing the relationshipbetween IAS and CAS must be shown in theAirplane Flight Manual.
14. Static air vent system. The static air ventsystem must meet the requirements of FAR23.1325. The altimeter system calibrationmust be determined and shown in the Air-plane Flight Manual.
OPERATING LIMITATIONS AND INFORMATION
15. Maximum operating limit speed VMO/MMO.
Instead of establishing operating limitationsbased on VME and VNO, the applicant mustestablish a maximum operating limit speedVMO/MMO in accordance with the following:
(a) The maximum operating limit speedmust not exceed the design cruising speed Vcand must be sufficiently below VD/MD or VDF/MDF to make it highly improbable that thelatter speeds will be inadvertently exceededin flight.
(b) The speed Vmo must not exceed 0.8 VD/MD or 0.8 VDF/MDF unless flight demonstra-tions involving upsets as specified by the Ad-ministrator indicates a lower speed marginwill not result in speeds exceeding VD/MD orVDF. Atmospheric variations, horizontalgusts, and equipment errors, and airframe
production variations will be taken into ac-count.
16. Minimum flight crew. In addition tomeeting the requirements of FAR 23.1523, theapplicant must establish the minimum num-ber and type of qualified flight crew per-sonnel sufficient for safe operation of theairplane considering—
(a) Each kind of operation for which theapplicant desires approval;
(b) The workload on each crewmember con-sidering the following:
(1) Flight path control.(2) Collision avoidance.(3) Navigation.(4) Communications.(5) Operation and monitoring of all essen-
tial aircraft systems.(6) Command decisions; and(c) The accessibility and ease of operation
of necessary controls by the appropriatecrewmember during all normal and emer-gency operations when at his flight station.
17. Airspeed indicator. The airspeed indi-cator must meet the requirements of FAR23.1545 except that, the airspeed notationsand markings in terms of VNO and VNE mustbe replaced by the VMO/MMO notations. Theairspeed indicator markings must be easilyread and understood by the pilot. A placardadjacent to the airspeed indicator is an ac-ceptable means of showing compliance withthe requirements of FAR 23.1545(c).
AIRPLANE FLIGHT MANUAL
18. General. The Airplane Flight Manualmust be prepared in accordance with the re-quirements of FARs 23.1583 and 23.1587, andin addition the operating limitations andperformance information set forth in sec-tions 19 and 20 must be included.
19. Operating limitations. The AirplaneFlight Manual must include the followinglimitations—
(a) Airspeed limitations. (1) The maximumoperating limit speed VMO/MMO and a state-ment that this speed limit may not be delib-erately exceeded in any regime of flight(climb, cruise, or descent) unless a higherspeed is authorized for flight test or pilottraining;
(2) If an airspeed limitation is based uponcompressibility effects, a statement to thiseffect and information as to any symptoms,the probable behavior of the airplane, andthe recommended recovery procedures; and
(3) The airspeed limits, shown in terms ofVMO/MMO instead of VNO and VNE.
(b) Takeoff weight limitations. The max-imum takeoff weight for each airport ele-vation, ambient temperature, and availabletakeoff runway length within the range se-lected by the applicant. This weight may notexceed the weight at which:
(1) The all-engine operating takeoff dis-tance determined in accordance with section
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5(d) or the accelerate-stop distance deter-mined in accordance with section 5(c), whichever is greater, is equal to the available run-way length;
(2) The airplane complies with the one-en-gine-inoperative takeoff requirements speci-fied in section 5(e); and
(3) The airplane complies with the one-en-gine-inoperative en route climb require-ments specified in section 6(b), assumingthat a standard temperature lapse rate ex-ists from the airport elevation to the alti-tude of 5,000 feet, except that the weight maynot exceed that corresponding to a tempera-ture of 41° F at 5,000 feet.
20. Performance information. The AirplaneFlight Manual must contain the performanceinformation determined in accordance withthe provisions of the performance require-ments of this regulation. The informationmust include the following:
(a) Sufficient information so that the take-off weight limits specified in section 19(b)can be determined for all temperatures andaltitudes within the operation limitationsselected by the applicant.
(b) The conditions under which the per-formance information was obtained, includ-ing the airspeed at the 50-foot height used todetermine landing distances.
(c) The performance information (deter-mined by extrapolation and computed for therange of weights between the maximumlanding and takeoff weights) for—
(1) Climb in the landing configuration; and(2) Landing distance.(d) Procedure established under section 4 of
this regulation related to the limitationsand information required by this section inthe form of guidance material including anyrelevant limitations or information.
(e) An explanation of significant or un-usual flight or ground handling characteris-tics of the airplane.
(f) Airspeeds, as indicated airspeeds, cor-responding to those determined for takeoffin accordance with section 5(b).
21. Maximum operating altitudes. The max-imum operating altitude to which operationis permitted, as limited by flight, structural,powerplant, functional, or equipment char-acteristics, must be specified in the AirplaneFlight Manual.
22. Stowage provision for Airplane FlightManual. Provision must be made for stowingthe Airplane Flight Manual in a suitablefixed container which is readily accessible tothe pilot.
23. Operating procedures. Procedures for re-starting turbine engines in flight (includingthe effects of altitude) must be set forth inthe Airplane Flight Manual.
AIRFRAME REQUIREMENTS
FLIGHT LOADS
24. Engine torque. (a) Each turbopropellerengine mount and its supporting structuremust be designed for the torque effects of—
(1) The conditions set forth in FAR23.361(a).
(2) The limit engine torque correspondingto takeoff power and propeller speed, multi-plied by a factor accounting for propellercontrol system malfunction, including quickfeathering action, simultaneously with 1 glevel flight loads. In the absence of a ration-al analysis, a factor of 1.6 must be used.
(b) The limit torque is obtained by multi-plying the mean torque by a factor of 1.25.
25. Turbine engine gyroscopic loads. Eachturbopropeller engine mount and its sup-porting structure must be designed for thegyroscopic loads that result, with the en-gines at maximum continuous r.p.m., undereither—
(a) The conditions prescribed in FARs23.351 and 23.423; or
(b) All possible combinations of the fol-lowing:
(1) A yaw velocity of 2.5 radius per second.(2) A pitch velocity of 1.0 radians per sec-
ond.(3) A normal load factor of 2.5.(4) Maximum continuous thrust.26. Unsymmetrical loads due to engine failure.
(a) Turbopropeller powered airplanes mustbe designed for the unsymmetrical loads re-sulting from the failure of the critical engineincluding the following conditions in com-bination with a single malfunction of thepropeller drag limiting system, consideringthe probable pilot corrective action on theflight controls.
(1) At speeds between VMC and VD, theloads resulting from power failure because offuel flow interruption are considered to belimit loads.
(2) At speeds between VMC and VC, theloads resulting from the disconnection of theengine compressor from the turbine or fromloss of the turbine blades are considered tobe ultimate loads.
(3) The time history of the thrust decayand drag buildup occurring as a result of theprescribed engine failures must be substan-tiated by test or other data applicable to theparticular engine-propeller combination.
(4) The timing and magnitude of the prob-able pilot corrective action must be conserv-atively estimated, considering the character-istics of the particular engine-propeller-air-plane combination.
(b) Pilot corrective action may be assumedto be initiated at the time maximum yawingvelocity is reached, but not earlier than twoseconds after the engine failure. The mag-nitude of the corrective action may be basedon the control forces specified in FAR 23.397except that lower forces may be assumed
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where it is shown by analysis or test thatthese forces can control the yaw and roll re-sulting from the prescribed engine failureconditions.
GROUND LOADS
27. Dual wheel landing gear units. Each dualwheel landing gear unit and its supportingstructure must be shown to comply with thefollowing:
(a) Pivoting. The airplane must be assumedto pivot about one side of the main gear withthe brakes on that side locked. The limitvertical load factor must be 1.0 and the coef-ficient of friction 0.8. This condition needapply only to the main gear and its sup-porting structure.
(b) Unequal tire inflation. A 60–40 percentdistribution of the loads established in ac-cordance with FAR 23.471 through FAR 23.483must be applied to the dual wheels.
(c) Flat tire. (1) Sixty percent of the loadsspecified in FAR 23.471 through FAR 23.483must be applied to either wheel in a unit.
(2) Sixty percent of the limit drag and sideloads and 100 percent of the limit verticalload established in accordance with FARs23.493 and 23.485 must be applied to eitherwheel in a unit except that the vertical loadneed not exceed the maximum vertical loadin paragraph (c)(1) of this section.
FATIGUE EVALUATION
28. Fatigue evaluation of wing and associatedstructure. Unless it is shown that the struc-ture, operating stress levels, materials, andexpected use are comparable from a fatiguestandpoint to a similar design which has hadsubstantial satisfactory service experience,the strength, detail design, and the fabrica-tion of those parts of the wing, wing carry-through, and attaching structure whose fail-ure would be catastrophic must be evaluatedunder either—
(a) A fatigue strength investigation inwhich the structure is shown by analysis,tests, or both to be able to withstand the re-peated loads of variable magnitude expectedin service; or
(b) A fail-safe strength investigation inwhich it is shown by analysis, tests, or boththat catastrophic failure of the structure isnot probable after fatigue, or obvious partialfailure, of a principal structural element,and that the remaining structure is able towithstand a static ultimate load factor of 75percent of the critical limit load factor atVc. These loads must be multiplied by a fac-tor of 1.15 unless the dynamic effects of fail-ure under static load are otherwise consid-ered.
DESIGN AND CONSTRUCTION
29. Flutter. For Multiengine turbopropellerpowered airplanes, a dynamic evaluationmust be made and must include—
(a) The significant elastic, inertia, and aer-odynamic forces associated with the rota-tions and displacements of the plane of thepropeller; and
(b) Engine-propeller-nacelle stiffness anddamping variations appropriate to the par-ticular configuration.
LANDING GEAR
30. Flap operated landing gear warning de-vice. Airplanes having retractable landinggear and wing flaps must be equipped with awarning device that functions continuouslywhen the wing flaps are extended to a flapposition that activates the warning device togive adequate warning before landing, usingnormal landing procedures, if the landinggear is not fully extended and locked. Theremay not be a manual shut off for this warn-ing device. The flap position sensing unitmay be installed at any suitable location.The system for this device may use any partof the system (including the aural warningdevice) provided for other landing gear warn-ing devices.
PERSONNEL AND CARGO ACCOMMODATIONS
31. Cargo and baggage compartments. Cargoand baggage compartments must be designedto meet the requirements of FAR 23.787 (a)and (b), and in addition means must be pro-vided to protect passengers from injury bythe contents of any cargo or baggage com-partment when the ultimate forward inertiaforce is 9g.
32. Doors and exits. The airplane must meetthe requirements of FAR 23.783 and FAR23.807 (a)(3), (b), and (c), and in addition:
(a) There must be a means to lock andsafeguard each external door and exitagainst opening in flight either inadvert-ently by persons, or as a result of mechan-ical failure. Each external door must be op-erable from both the inside and the outside.
(b) There must be means for direct visualinspection of the locking mechanism bycrewmembers to determine whether externaldoors and exits, for which the initial openingmovement is outward, are fully locked. Inaddition, there must be a visual means tosignal to crewmembers when normally usedexternal doors are closed and fully locked.
(c) The passenger entrance door must qual-ify as a floor level emergency exit. Each ad-ditional required emergency exit except floorlevel exits must be located over the wing ormust be provided with acceptable means toassist the occupants in descending to theground. In addition to the passenger en-trance door:
(1) For a total seating capacity of 15 orless, an emergency exit as defined in FAR23.807(b) is required on each side of the cabin.
(2) For a total seating capacity of 16through 23, three emergency exits as definedin 23.807(b) are required with one on the same
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side as the door and two on the side oppositethe door.
(d) An evacuation demonstration must beconducted utilizing the maximum number ofoccupants for which certification is desired.It must be conducted under simulated nightconditions utilizing only the emergencyexits on the most critical side of the aircraft.The participants must be representative ofaverage airline passengers with no priorpractice or rehearsal for the demonstration.Evacuation must be completed within 90 sec-onds.
(e) Each emergency exit must be markedwith the word ‘‘Exit’’ by a sign which haswhite letters 1 inch high on a red back-ground 2 inches high, be self-illuminated orindependently internally electrically illumi-nated, and have a minimum luminescence(brightness) of at least 160 microlamberts.The colors may be reversed if the passengercompartment illumination is essentially thesame.
(f) Access to window type emergency exitsmust not be obstructed by seats or seatbacks.
(g) The width of the main passenger aisleat any point between seats must equal or ex-ceed the values in the following table.
Total seating capacity
Minimum main passenger aislewidth
Less than 25inches from floor
25 inches andmore from floor
10 through 23 ........... 9 inches ............. 15 inches.
MISCELLANEOUS
33. Lightning strike protection. Parts thatare electrically insulated from the basic air-frame must be connected to it through light-ning arrestors unless a lightning strike onthe insulated part—
(a) Is improbable because of shielding byother parts; or
(b) Is not hazardous.34. Ice protection. If certification with ice
protection provisions is desired, compliancewith the following requirements must beshown:
(a) The recommended procedures for theuse of the ice protection equipment must beset forth in the Airplane Flight Manual.
(b) An analysis must be performed to es-tablish, on the basis of the airplane’s oper-ational needs, the adequacy of the ice protec-tion system for the various components ofthe airplane. In addition, tests of the ice pro-tection system must be conducted to dem-onstrate that the airplane is capable of oper-ating safely in continuous maximum andintermittent maximum icing conditions asdescribed in FAR 25, appendix C.
(c) Compliance with all or portions of thissection may be accomplished by reference,where applicable because of similarity of the
designs, to analysis and tests performed bythe applicant for a type certificated model.
35. Maintenance information. The applicantmust make available to the owner at thetime of delivery of the airplane the informa-tion he considers essential for the propermaintenance of the airplane. That informa-tion must include the following:
(a) Description of systems, including elec-trical, hydraulic, and fuel controls.
(b) Lubrication instructions setting forththe frequency and the lubricants and fluidswhich are to be used in the various systems.
(c) Pressures and electrical loads applica-ble to the various systems.
(d) Tolerances and adjustments necessaryfor proper functioning.
(e) Methods of leveling, raising, and tow-ing.
(f) Methods of balancing control surfaces.(g) Identification of primary and secondary
structures.(h) Frequency and extent of inspections
necessary to the proper operation of the air-plane.
(i) Special repair methods applicable to theairplane.
(j) Special inspection techniques, includingthose that require X-ray, ultrasonic, andmagnetic particle inspection.
(k) List of special tools.
PROPULSION
GENERAL
36. Vibration characteristics. For turbo-propeller powered airplanes, the engine in-stallation must not result in vibration char-acteristics of the engine exceeding those es-tablished during the type certification of theengine.
37. In-flight restarting of engine. If the en-gine on turbopropeller powered airplanescannot be restarted at the maximum cruisealtitude, a determination must be made ofthe altitude below which restarts can be con-sistently accomplished. Restart informationmust be provided in the Airplane FlightManual.
38. Engines—(a) For turbopropeller poweredairplanes. The engine installation must com-ply with the following requirements:
(1) Engine isolation. The powerplants mustbe arranged and isolated from each other toallow operation, in at least one configura-tion, so that the failure or malfunction ofany engine, or of any system that can affectthe engine, will not—
(i) Prevent the continued safe operation ofthe remaining engines; or
(ii) Require immediate action by any crew-member for continued safe operation.
(2) Control of engine rotation. There must bea means to individually stop and restart therotation of any engine in flight except that
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engine rotation need not be stopped if con-tinued rotation could not jeopardize the safe-ty of the airplane. Each component of thestopping and restarting system on the engineside of the firewall, and that might be ex-posed to fire, must be at least fire resistant.If hydraulic propeller feathering systems areused for this purpose, the feathering linesmust be at least fire resistant under the op-erating conditions that may be expected toexist during feathering.
(3) Engine speed and gas temperature controldevices. The powerplant systems associatedwith engine control devices, systems, and in-strumentation must provide reasonable as-surance that those engine operating limita-tions that adversely affect turbine rotorstructural integrity will not be exceeded inservice.
(b) For reciprocating-engine powered air-planes. To provide engine isolation, the pow-erplants must be arranged and isolated fromeach other to allow operation, in at least oneconfiguration, so that the failure or malfunc-tion of any engine, or of any system that canaffect that engine, will not—
(1) Prevent the continued safe operation ofthe remaining engines; or
(2) Require immediate action by any crew-member for continued safe operation.
39. Turbopropeller reversing systems. (a) Tur-bopropeller reversing systems intended forground operation must be designed so thatno single failure or malfunction of the sys-tem will result in unwanted reverse thrustunder any expected operating condition.Failure of structural elements need not beconsidered if the probability of this kind offailure is extremely remote.
(b) Turbopropeller reversing systems in-tended for in-flight use must be designed sothat no unsafe condition will result duringnormal operation of the system, or from anyfailure (or reasonably likely combination offailures) of the reversing system, under anyanticipated condition of operation of the air-plane. Failure of structural elements neednot be considered if the probability of thiskind of failure is extremely remote.
(c) Compliance with this section may beshown by failure analysis, testing, or bothfor propeller systems that allow propellerblades to move from the flight low-pitch po-sition to a position that is substantially lessthan that at the normal flight low-pitch stopposition. The analysis may include or be sup-ported by the analysis made to show compli-ance with the type certification of the pro-peller and associated installation compo-nents. Credit will be given for pertinentanalysis and testing completed by the engineand propeller manufacturers.
40. Turbopropeller drag-limiting systems. Tur-bopropeller drag-limiting systems must bedesigned so that no single failure or malfunc-tion of any of the systems during normal oremergency operation results in propeller
drag in excess of that for which the airplanewas designed. Failure of structural elementsof the drag-limiting systems need not be con-sidered if the probability of this kind of fail-ure is extremely remote.
41. Turbine engine powerplant operatingcharacteristics. For turbopropeller poweredairplanes, the turbine engine powerplant op-erating characteristics must be investigatedin flight to determine that no adverse char-acteristics (such as stall, surge, or flameout)are present to a hazardous degree, duringnormal and emergency operation within therange of operating limitations of the air-plane and of the engine.
42. Fuel flow. (a) For turbopropeller pow-ered airplanes—
(1) The fuel system must provide for con-tinuous supply of fuel to the engines for nor-mal operation without interruption due todepletion of fuel in any tank other than themain tank; and
(2) The fuel flow rate for turbopropeller en-gine fuel pump systems must not be lessthan 125 percent of the fuel flow required todevelop the standard sea level atmosphericconditions takeoff power selected and in-cluded as an operating limitation in the Air-plane Flight Manual.
(b) For reciprocating engine powered air-planes, it is acceptable for the fuel flow ratefor each pump system (main and reserve sup-ply) to be 125 percent of the takeoff fuel con-sumption of the engine.
FUEL SYSTEM COMPONENTS
43. Fuel pumps. For turbopropeller poweredairplanes, a reliable and independent powersource must be provided for each pump usedwith turbine engines which do not have pro-visions for mechanically driving the mainpumps. It must be demonstrated that thepump installations provide a reliability anddurability equivalent to that provided byFAR 23.991(a).
44. Fuel strainer or filter. For turbopropellerpowered airplanes, the following apply:
(a) There must be a fuel strainer or filterbetween the tank outlet and the fuel meter-ing device of the engine. In addition, the fuelstrainer or filter must be—
(1) Between the tank outlet and the en-gine-driven positive displacement pumpinlet, if there is an engine-driven positivedisplacement pump;
(2) Accessible for drainage and cleaningand, for the strainer screen, easily remov-able; and
(3) Mounted so that its weight is not sup-ported by the connecting lines or by theinlet or outlet connections of the strainer orfilter itself.
(b) Unless there are means in the fuel sys-tem to prevent the accumulation of ice onthe filter, there must be means to automati-cally maintain the fuel flow if ice-clogging ofthe filter occurs; and
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(c) The fuel strainer or filter must be ofadequate capacity (with respect to operatinglimitations established to insure proper serv-ice) and of appropriate mesh to insure properengine operation, with the fuel contaminatedto a degree (with respect to particle size anddensity) that can be reasonably expected inservice. The degree of fuel filtering may notbe less than that established for the enginetype certification.
45. Lightning strike protection. Protectionmust be provided against the ignition offlammable vapors in the fuel vent systemdue to lightning strikes.
COOLING
46. Cooling test procedures for turbopropellerpowered airplanes. (a) Turbopropeller poweredairplanes must be shown to comply with therequirements of FAR 23.1041 during takeoff,climb en route, and landing stages of flightthat correspond to the applicable perform-ance requirements. The cooling test must beconducted with the airplane in the configu-ration and operating under the conditionsthat are critical relative to cooling duringeach stage of flight. For the cooling tests atemperature is ‘‘stabilized’’ when its rate ofchange is less than 2° F. per minute.
(b) Temperatures must be stabilized underthe conditions from which entry is made intoeach stage of flight being investigated unlessthe entry condition is not one during whichcomponent and engine fluid temperatureswould stabilize, in which case, operationthrough the full entry condition must beconducted before entry into the stage offlight being investigated in order to allowtemperatures to reach their natural levels atthe time of entry. The takeoff cooling testmust be preceded by a period during whichthe powerplant component and engine fluidtemperatures are stabilized with the enginesat ground idle.
(c) Cooling tests for each stage of flightmust be continued until—
(1) The component and engine fluid tem-peratures stabilize;
(2) The stage of flight is completed; or(3) An operating limitation is reached.
INDUCTION SYSTEM
47. Air induction. For turbopropeller pow-ered airplanes—
(a) There must be means to prevent haz-ardous quantities of fuel leakage or overflowfrom drains, vents, or other components offlammable fluid systems from entering theengine intake system; and
(b) The air inlet ducts must be located orprotected so as to minimize the ingestion offoreign matter during takeoff, landing, andtaxiing.
48. Induction system icing protection. Forturbopropeller powered airplanes, each tur-bine engine must be able to operate through-
out its flight power range without adverseeffect on engine operation or serious loss ofpower or thrust, under the icing conditionsspecified in appendix C of FAR 25. In addi-tion, there must be means to indicate to ap-propriate flight crewmembers the func-tioning of the powerplant ice protection sys-tem.
49. Turbine engine bleed air systems. Turbineengine bleed air systems of turbopropellerpowered airplanes must be investigated todetermine—
(a) That no hazard to the airplane will re-sult if a duct rupture occurs. This conditionmust consider that a failure of the duct canoccur anywhere between the engine port andthe airplane bleed service; and
(b) That if the bleed air system is used fordirect cabin pressurization, it is not possiblefor hazardous contamination of the cabin airsystem to occur in event of lubrication sys-tem failure.
EXHAUST SYSTEM
50. Exhaust system drains. Turbopropellerengine exhaust systems having low spots orpockets must incorporate drains at such lo-cations. These drains must discharge clear ofthe airplane in normal and ground attitudesto prevent the accumulation of fuel after thefailure of an attempted engine start.
POWERPLANT CONTROLS AND ACCESSORIES
51. Engine controls. If throttles or power le-vers for turbopropeller powered airplanes aresuch that any position of these controls willreduce the fuel flow to the engine(s) belowthat necessary for satisfactory and safe idleoperation of the engine while the airplane isin flight, a means must be provided to pre-vent inadvertent movement of the controlinto this position. The means provided mustincorporate a positive lock or stop at thisidle position and must require a separate anddistinct operation by the crew to displacethe control from the normal engine oper-ating range.
52. Reverse thrust controls. For turbo-propeller powered airplanes, the propeller re-verse thrust controls must have a means toprevent their inadvertent operation. Themeans must have a positive lock or stop atthe idle position and must require a separateand distinct operation by the crew to dis-place the control from the flight regime.
53. Engine ignition systems. Each turbo-propeller airplane ignition system must beconsidered an essential electrical load.
54. Powerplant accessories. The powerplantaccessories must meet the requirements ofFAR 23.1163, and if the continued rotation ofany accessory remotely driven by the engineis hazardous when malfunctioning occurs,there must be means to prevent rotationwithout interfering with the continued oper-ation of the engine.
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POWERPLANT FIRE PROTECTION
55. Fire detector system. For turbopropellerpowered airplanes, the following apply:
(a) There must be a means that ensuresprompt detection of fire in the engine com-partment. An overtemperature switch ineach engine cooling air exit is an acceptablemethod of meeting this requirement.
(b) Each fire detector must be constructedand installed to withstand the vibration, in-ertia, and other loads to which it may besubjected in operation.
(c) No fire detector may be affected by anyoil, water, other fluids, or fumes that mightbe present.
(d) There must be means to allow the flightcrew to check, in flight, the functioning ofeach fire detector electric circuit.
(e) Wiring and other components of eachfire detector system in a fire zone must be atleast fire resistant.
56. Fire protection, cowling and nacelle skin.For reciprocating engine powered airplanes,the engine cowling must be designed andconstructed so that no fire originating in theengine compartment can enter, eitherthrough openings or by burn through, anyother region where it would create addi-tional hazards.
57. Flammable fluid fire protection. If flam-mable fluids or vapors might be liberated bythe leakage of fluid systems in areas otherthan engine compartments, there must bemeans to—
(a) Prevent the ignition of those fluids orvapors by any other equipment; or
(b) Control any fire resulting from that ig-nition.
EQUIPMENT
58. Powerplant instruments. (a) The fol-lowing are required for turbopropeller air-planes:
(1) The instruments required by FAR23.1305 (a)(1) through (4), (b)(2) and (4).
(2) A gas temperature indicator for eachengine.
(3) Free air temperature indicator.(4) A fuel flowmeter indicator for each en-
gine.(5) Oil pressure warning means for each en-
gine.(6) A torque indicator or adequate means
for indicating power output for each engine.(7) Fire warning indicator for each engine.(8) A means to indicate when the propeller
blade angle is below the low-pitch positioncorresponding to idle operation in flight.
(9) A means to indicate the functioning ofthe ice protection system for each engine.
(b) For turbopropeller powered airplanes,the turbopropeller blade position indicatormust begin indicating when the blade hasmoved below the flight low-pitch position.
(c) The following instruments are requiredfor reciprocating-engine powered airplanes:
(1) The instruments required by FAR23.1305.
(2) A cylinder head temperature indicatorfor each engine.
(3) A manifold pressure indicator for eachengine.
SYSTEMS AND EQUIPMENTS
GENERAL
59. Function and installation. The systemsand equipment of the airplane must meet therequirements of FAR 23.1301, and the fol-lowing:
(a) Each item of additional installed equip-ment must—
(1) Be of a kind and design appropriate toits intended function;
(2) Be labeled as to its identification, func-tion, or operating limitations, or any appli-cable combination of these factors, unlessmisuse or inadvertent actuation cannot cre-ate a hazard;
(3) Be installed according to limitationsspecified for that equipment; and
(4) Function properly when installed.(b) Systems and installations must be de-
signed to safeguard against hazards to theaircraft in the event of their malfunction orfailure.
(c) Where an installation, the functioningof which is necessary in showing compliancewith the applicable requirements, requires apower supply, such installation must be con-sidered an essential load on the power sup-ply, and the power sources and the distribu-tion system must be capable of supplying thefollowing power loads in probable operationcombinations and for probable durations:
(1) All essential loads after failure of anyprime mover, power converter, or energystorage device.
(2) All essential loads after failure of anyone engine on two-engine airplanes.
(3) In determining the probable operatingcombinations and durations of essentialloads for the power failure conditions de-scribed in subparagraphs (1) and (2) of thisparagraph, it is permissible to assume thatthe power loads are reduced in accordancewith a monitoring procedure which is con-sistent with safety in the types of operationsauthorized.
60. Ventilation. The ventilation system ofthe airplane must meet the requirements ofFAR 23.831, and in addition, for pressurizedaircraft the ventilating air in flight crew andpassenger compartments must be free ofharmful or hazardous concentrations ofgases and vapors in normal operation and inthe event of reasonably probable failures ormalfunctioning of the ventilating, heating,pressurization, or other systems, and equip-ment. If accumulation of hazardous quan-tities of smoke in the cockpit area is reason-ably probable, smoke evacuation must bereadily accomplished.
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ELECTRICAL SYSTEMS AND EQUIPMENT
61. General. The electrical systems andequipment of the airplane must meet the re-quirements of FAR 23.1351, and the following:
(a) Electrical system capacity. The requiredgenerating capacity, and number and kindsof power sources must—
(1) Be determined by an electrical loadanalysis, and
(2) Meet the requirements of FAR 23.1301.(b) Generating system. The generating sys-
tem includes electrical power sources, mainpower busses, transmission cables, and asso-ciated control, regulation, and protective de-vices. It must be designed so that—
(1) The system voltage and frequency (asapplicable) at the terminals of all essentialload equipment can be maintained withinthe limits for which the equipment is de-signed, during any probable operating condi-tions;
(2) System transients due to switching,fault clearing, or other causes do not makeessential loads inoperative, and do not causea smoke or fire hazard;
(3) There are means, accessible in flight toappropriate crewmembers, for the individualand collective disconnection of the electricalpower sources from the system; and
(4) There are means to indicate to appro-priate crewmembers the generating systemquantities essential for the safe operation ofthe system, including the voltage and cur-rent supplied by each generator.
62. Electrical equipment and installation.Electrical equipment controls, and wiringmust be installed so that operation of anyone unit or system of units will not ad-versely affect the simultaneous operation ofto the safe operation.
63. Distribution system. (a) For the purposeof complying with this section, the distribu-tion system includes the distribution busses,their associated feeders and each control andprotective device.
(b) Each system must be designed so thatessential load circuits can be supplied in theevent of reasonably probable faults or opencircuits, including faults in heavy currentcarrying cables.
(c) If two independent sources of electricalpower for particular equipment or systemsare required by this regulation, their elec-trical energy supply must be insured bymeans such as duplicate electrical equip-ment, throwover switching, or multichannelor loop circuits separately routed.
64. Circuit protective devices. The circuitprotective devices for the electrical circuitsof the airplane must meet the requirementsof FAR 23.1357, and in addition circuits forloads which are essential to safe operation
must have individual and exclusive circuitprotection.
EDITORIAL NOTE: For the text of SFAR No.41, see part 21 of this chapter.
Subpart A—General§ 23.1 Applicability.
(a) This part prescribes airworthinessstandards for the issue of type certifi-cates, and changes to those certifi-cates, for airplanes in the normal, util-ity, acrobatic, and commuter cat-egories.
(b) Each person who applies underPart 21 for such a certificate or changemust show compliance with the appli-cable requirements of this part.
(a) Notwithstanding §§ 21.17 and 21.101of this chapter and irrespective of thetype certification basis, each normal,utility, and acrobatic category air-plane having a passenger seating con-figuration, excluding pilot seats, ofnine or less, manufactured after De-cember 12, 1986, or any such foreign air-plane for entry into the United Statesmust provide a safety belt and shoulderharness for each forward- or aft-facingseat which will protect the occupantfrom serious head injury when sub-jected to the inertia loads resultingfrom the ultimate static load factorsprescribed in § 23.561(b)(2) of this part,or which will provide the occupant pro-tection specified in § 23.562 of this partwhen that section is applicable to theairplane. For other seat orientations,the seat/restraint system must be de-signed to provide a level of occupantprotection equivalent to that providedfor forward- or aft-facing seats with asafety belt and shoulder harness in-stalled.
(b) Each shoulder harness installed ata flight crewmember station, as re-quired by this section, must allow the
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crewmember, when seated with thesafety belt and shoulder harness fas-tened, to perform all functions nec-essary for flight operations.
(c) For the purpose of this section,the date of manufacture is:
(1) The date the inspection accept-ance records, or equivalent, reflectthat the airplane is complete andmeets the FAA approved type designdata; or
(2) In the case of a foreign manufac-tured airplane, the date the foreigncivil airworthiness authority certifiesthe airplane is complete and issues anoriginal standard airworthiness certifi-cate, or the equivalent in that country.
[Amdt. 23–36, 53 FR 30812, Aug. 15, 1988]
§ 23.3 Airplane categories.(a) The normal category is limited to
airplanes that have a seating configu-ration, excluding pilot seats, of nine orless, a maximum certificated takeoffweight of 12,500 pounds or less, and in-tended for nonacrobatic operation.Nonacrobatic operation includes:
turns, in which the angle of bank is notmore than 60 degrees.
(b) The utility category is limited toairplanes that have a seating configu-ration, excluding pilot seats, of nine orless, a maximum certificated takeoffweight of 12,500 pounds or less, and in-tended for limited acrobatic operation.Airplanes certificated in the utilitycategory may be used in any of the op-erations covered under paragraph (a) ofthis section and in limited acrobaticoperations. Limited acrobatic oper-ation includes:
(1) Spins (if approved for the par-ticular type of airplane); and
(2) Lazy eights, chandelles, and steepturns, or similar maneuvers, in whichthe angle of bank is more than 60 de-grees but not more than 90 degrees.
(c) The acrobatic category is limitedto airplanes that have a seating con-figuration, excluding pilot seats, ofnine or less, a maximum certificatedtakeoff weight of 12,500 pounds or less,and intended for use without restric-tions, other than those shown to be
necessary as a result of required flighttests.
(d) The commuter category is limitedto propeller-driven, multiengine air-planes that have a seating configura-tion, excluding pilot seats, of 19 or less,and a maximum certificated takeoffweight of 19,000 pounds or less. Thecommuter category operation is lim-ited to any maneuver incident to nor-mal flying, stalls (except whip stalls),and steep turns, in which the angle ofbank is not more than 60 degrees.
(e) Except for commuter category,airplanes may be type certificated inmore than one category if the require-ments of each requested category aremet.
§ 23.21 Proof of compliance.(a) Each requirement of this subpart
must be met at each appropriate com-bination of weight and center of grav-ity within the range of loading condi-tions for which certification is re-quested. This must be shown—
(1) By tests upon an airplane of thetype for which certification is re-quested, or by calculations based on,and equal in accuracy to, the results oftesting; and
(2) By systematic investigation ofeach probable combination of weightand center of gravity, if compliancecannot be reasonably inferred fromcombinations investigated.
(b) The following general tolerancesare allowed during flight testing. How-ever, greater tolerances may be al-lowed in particular tests:
Item Tolerance
Weight ............................................... +5%, –10%.Critical items affected by weight ....... +5%, –1%.C.G .................................................... ±7% total travel.
§ 23.23 Load distribution limits.(a) Ranges of weights and centers of
gravity within which the airplane maybe safely operated must be established.
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Federal Aviation Administration, DOT § 23.31
If a weight and center of gravity com-bination is allowable only within cer-tain lateral load distribution limitsthat could be inadvertently exceeded,these limits must be established for thecorresponding weight and center ofgravity combinations.
(b) The load distribution limits maynot exceed any of the following:
(1) The selected limits;(2) The limits at which the structure
is proven; or(3) The limits at which compliance
with each applicable flight require-ment of this subpart is shown.
[Doc. No. 26269, 58 FR 42156, Aug. 6, 1993]
§ 23.25 Weight limits.(a) Maximum weight. The maximum
weight is the highest weight at whichcompliance with each applicable re-quirement of this part (other thanthose complied with at the design land-ing weight) is shown. The maximumweight must be established so that itis—
(1) Not more than the least of—(i) The highest weight selected by the
applicant; or(ii) The design maximum weight,
which is the highest weight at whichcompliance with each applicable struc-tural loading condition of this part(other than those complied with at thedesign landing weight) is shown; or
(iii) The highest weight at whichcompliance with each applicable flightrequirement is shown, and
(2) Not less than the weight with—(i) Each seat occupied, assuming a
weight of 170 pounds for each occupantfor normal and commuter category air-planes, and 190 pounds for utility andacrobatic category airplanes, exceptthat seats other than pilot seats maybe placarded for a lesser weight; and
(A) Oil at full capacity, and(B) At least enough fuel for max-
imum continuous power operation of atleast 30 minutes for day-VFR approvedairplanes and at least 45 minutes fornight-VFR and IFR approved airplanes;or
(ii) The required minimum crew, andfuel and oil to full tank capacity.
(b) Minimum weight. The minimumweight (the lowest weight at whichcompliance with each applicable re-quirement of this part is shown) must
be established so that it is not morethan the sum of—
(1) The empty weight determinedunder § 23.29;
(2) The weight of the required min-imum crew (assuming a weight of 170pounds for each crewmember); and
(3) The weight of—(i) For turbojet powered airplanes, 5
percent of the total fuel capacity ofthat particular fuel tank arrangementunder investigation, and
(ii) For other airplanes, the fuel nec-essary for one-half hour of operation atmaximum continuous power.
§ 23.31 Removable ballast.Removable ballast may be used in
showing compliance with the flight re-quirements of this subpart, if—
(a) The place for carrying ballast isproperly designed and installed, and ismarked under § 23.1557; and
(b) Instructions are included in theairplane flight manual, approved man-ual material, or markings and plac-ards, for the proper placement of theremovable ballast under each loading
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14 CFR Ch. I (1–1–01 Edition)§ 23.33
condition for which removable ballastis necessary.
(a) General. The propeller speed andpitch must be limited to values thatwill assure safe operation under normaloperating conditions.
(b) Propellers not controllable in flight.For each propeller whose pitch cannotbe controlled in flight—
(1) During takeoff and initial climbat the all engine(s) operating climbspeed specified in § 23.65, the propellermust limit the engine r.p.m., at fullthrottle or at maximum allowabletakeoff manifold pressure, to a speednot greater than the maximum allow-able takeoff r.p.m.; and
(2) During a closed throttle glide, atVNE, the propeller may not cause an en-gine speed above 110 percent of max-imum continuous speed.
(c) Controllable pitch propellers withoutconstant speed controls. Each propellerthat can be controlled in flight, butthat does not have constant speed con-trols, must have a means to limit thepitch range so that—
(1) The lowest possible pitch allowscompliance with paragraph (b)(1) ofthis section; and
(2) The highest possible pitch allowscompliance with paragraph (b)(2) ofthis section.
(d) Controllable pitch propellers withconstant speed controls. Each control-lable pitch propeller with constantspeed controls must have—
(1) With the governor in operation, ameans at the governor to limit themaximum engine speed to the max-imum allowable takeoff r.p.m.; and
(2) With the governor inoperative,the propeller blades at the lowest pos-sible pitch, with takeoff power, the air-plane stationary, and no wind, either—
(i) A means to limit the maximumengine speed to 103 percent of the max-imum allowable takeoff r.p.m., or
(ii) For an engine with an approvedoverspeed, a means to limit the max-imum engine and propeller speed to not
(a) Unless otherwise prescribed, theperformance requirements of this partmust be met for—
(1) Still air and standard atmosphere;and
(2) Ambient atmospheric conditions,for commuter category airplanes, forreciprocating engine-powered airplanesof more than 6,000 pounds maximumweight, and for turbine engine-poweredairplanes.
(b) Performance data must be deter-mined over not less than the followingranges of conditions—
(1) Airport altitudes from sea level to10,000 feet; and
(2) For reciprocating engine-poweredairplanes of 6,000 pounds, or less, max-imum weight, temperature from stand-ard to 30 °C above standard; or
(3) For reciprocating engine-poweredairplanes of more than 6,000 poundsmaximum weight and turbine engine-powered airplanes, temperature fromstandard to 30 °C above standard, orthe maximum ambient atmospherictemperature at which compliance withthe cooling provisions of § 23.1041 to§ 23.1047 is shown, if lower.
(c) Performance data must be deter-mined with the cowl flaps or othermeans for controlling the engine cool-ing air supply in the position used inthe cooling tests required by § 23.1041 to§ 23.1047.
(d) The available propulsive thrustmust correspond to engine power, notexceeding the approved power, less—
(1) Installation losses; and(2) The power absorbed by the acces-
sories and services appropriate to theparticular ambient atmospheric condi-tions and the particular flight condi-tion.
(e) The performance, as affected byengine power or thrust, must be basedon a relative humidity:
(1) Of 80 percent at and below stand-ard temperature; and
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Federal Aviation Administration, DOT § 23.49
(2) From 80 percent, at the standardtemperature, varying linearly down to34 percent at the standard temperatureplus 50 °F.
(f) Unless otherwise prescribed, in de-termining the takeoff and landing dis-tances, changes in the airplane’s con-figuration, speed, and power must bemade in accordance with procedures es-tablished by the applicant for oper-ation in service. These procedures mustbe able to be executed consistently bypilots of average skill in atmosphericconditions reasonably expected to beencountered in service.
(g) The following, as applicable, mustbe determined on a smooth, dry, hard-surfaced runway—
(1) Takeoff distance of § 23.53(b);(2) Accelerate-stop distance of § 23.55;(3) Takeoff distance and takeoff run
of § 23.59; and(4) Landing distance of § 23.75.NOTE: The effect on these distances of op-
eration on other types of surfaces (for exam-ple, grass, gravel) when dry, may be deter-mined or derived and these surfaces listed inthe Airplane Flight Manual in accordancewith § 23.1583(p).
(h) For commuter category airplanes,the following also apply:
(1) Unless otherwise prescribed, theapplicant must select the takeoff,enroute, approach, and landing con-figurations for the airplane.
(2) The airplane configuration mayvary with weight, altitude, and tem-perature, to the extent that they arecompatible with the operating proce-dures required by paragraph (h)(3) ofthis section.
(3) Unless otherwise prescribed, in de-termining the critical-engine-inoper-ative takeoff performance, takeoffflight path, and accelerate-stop dis-tance, changes in the airplane’s con-figuration, speed, and power must bemade in accordance with procedures es-tablished by the applicant for oper-ation in service.
(4) Procedures for the execution ofdiscontinued approaches and balkedlandings associated with the conditionsprescribed in § 23.67(c)(4) and § 23.77(c)must be established.
(5) The procedures established underparagraphs (h)(3) and (h)(4) of this sec-tion must—
(i) Be able to be consistently exe-cuted by a crew of average skill in at-
mospheric conditions reasonably ex-pected to be encountered in service;
(ii) Use methods or devices that aresafe and reliable; and
(iii) Include allowance for any rea-sonably expected time delays in theexecution of the procedures.
[Doc. No. 27807, 61 FR 5184, Feb. 9, 1996]
§ 23.49 Stalling period.
(a) VSO and VS1 are the stalling speedsor the minimum steady flight speeds,in knots (CAS), at which the airplaneis controllable with—
(1) For reciprocating engine-poweredairplanes, the engine(s) idling, thethrottle(s) closed or at not more thanthe power necessary for zero thrust ata speed not more than 110 percent ofthe stalling speed;
(2) For turbine engine-powered air-planes, the propulsive thrust not great-er than zero at the stalling speed, or, ifthe resultant thrust has no appreciableeffect on the stalling speed, with en-gine(s) idling and throttle(s) closed;
(3) The propeller(s) in the takeoff po-sition;
(4) The airplane in the condition ex-isting in the test, in which VSO and VS1
are being used;(5) The center of gravity in the posi-
tion that results in the highest value ofVSO and VS1; and
(6) The weight used when VSO and VS1
are being used as a factor to determinecompliance with a required perform-ance standard.
(b) VSO and VS1 must be determinedby flight tests, using the procedure andmeeting the flight characteristics spec-ified in § 23.201.
(c) Except as provided in paragraph(d) of this section, VSO and VS1 at max-imum weight must not exceed 61 knotsfor—
(1) Single-engine airplanes; and(2) Multiengine airplanes of 6,000
pounds or less maximum weight thatcannot meet the minimum rate ofclimb specified in § 23.67(a) (1) with thecritical engine inoperative.
(d) All single-engine airplanes, andthose multiengine airplanes of 6,000pounds or less maximum weight with aVSO of more than 61 knots that do not
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14 CFR Ch. I (1–1–01 Edition)§ 23.51
meet the requirements of § 23.67(a)(1),must comply with § 23.562(d).
[Doc. No. 27807, 61 FR 5184, Feb. 9, 1996]
§ 23.51 Takeoff speeds.(a) For normal, utility, and acrobatic
category airplanes, rotation speed, VR,is the speed at which the pilot makes acontrol input, with the intention oflifting the airplane out of contact withthe runway or water surface.
(1) For multiengine landplanes, VR,must not be less than the greater of1.05 VMC; or 1.10 VS1;
(2) For single-engine landplanes, VR,must not be less than VS1; and
(3) For seaplanes and amphibianstaking off from water, VR, may be anyspeed that is shown to be safe under allreasonably expected conditions, includ-ing turbulence and complete failure ofthe critical engine.
(b) For normal, utility, and acrobaticcategory airplanes, the speed at 50 feetabove the takeoff surface level mustnot be less than:
(1) or multiengine airplanes, thehighest of—
(i) A speed that is shown to be safefor continued flight (or emergencylanding, if applicable) under all reason-ably expected conditions, includingturbulence and complete failure of thecritical engine;
(ii) 1.10 VMC; or(iii) 1.20 VS1.(2) For single-engine airplanes, the
higher of—(i) A speed that is shown to be safe
under all reasonably expected condi-tions, including turbulence and com-plete engine failure; or
(ii) 1.20 VS1.(c) For commuter category airplanes,
the following apply:(l) V1 must be established in relation
to VEF as follows:(i) VEF is the calibrated airspeed at
which the critical engine is assumed tofail. VEF must be selected by the appli-cant but must not be less than 1.05 VMC
determined under § 23.149(b) or, at theoption of the applicant, not less thanVMCG determined under § 23.149(f).
(ii) The takeoff decision speed, V1, isthe calibrated airspeed on the groundat which, as a result of engine failureor other reasons, the pilot is assumedto have made a decision to continue or
discontinue the takeoff. The takeoffdecision speed, V1, must be selected bythe applicant but must not be less thanVEF plus the speed gained with the crit-ical engine inoperative during the timeinterval between the instant at whichthe critical engine is failed and the in-stant at which the pilot recognizes andreacts to the engine failure, as indi-cated by the pilot’s application of thefirst retarding means during the accel-erate-stop determination of § 23.55.
(2) The rotation speed, VR, in termsof calibrated airspeed, must be selectedby the applicant and must not be lessthan the greatest of the following:
(i) V1;(ii) 1.05 VMC determined under
§ 23.149(b);(iii) 1.10 VS1; or(iv) The speed that allows attaining
the initial climb-out speed, V2, beforereaching a height of 35 feet above thetakeoff surface in accordance with§ 23.57(c)(2).
(3) For any given set of conditions,such as weight, altitude, temperature,and configuration, a single value of VR
must be used to show compliance withboth the one-engine-inoperative take-off and all-engines-operating takeoffrequirements.
(4) The takeoff safety speed, V2, interms of calibrated airspeed, must beselected by the applicant so as to allowthe gradient of climb required in § 23.67(c)(1) and (c)(2) but mut not be lessthan 1.10 VMC or less than 1.20 VS1.
(5) The one-engine-inoperative take-off distance, using a normal rotationrate at a speed 5 knots less than VR, es-tablished in accordance with paragraph(c)(2) of this section, must be shownnot to exceed the corresponding one-engine-inoperative takeoff distance,determined in accordance with § 23.57and § 23.59(a)(1), using the establishedVR. The takeoff, otherwise performedin accordance with § 23.57, must be con-tinued safely from the point at whichthe airplane is 35 feet above the takeoffsurface and at a speed not less than theestablished V2 minus 5 knots.
(6) The applicant must show, with allengines operating, that marked in-creases in the scheduled takeoff dis-tances, determined in accordance with
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Federal Aviation Administration, DOT § 23.57
§ 23.59(a)(2), do not result from over-ro-tation of the airplane or out-of-trimconditions.
[Doc. No. 27807, 61 FR 5184, Feb. 9, 1996]
§ 23.53 Takeoff performance.(a) For normal, utility, and acrobatic
category airplanes, the takeoff dis-tance must be determined in accord-ance with paragraph (b) of this section,using speeds determined in accordancewith § 23.51 (a) and (b).
(b) For normal, utility, and acrobaticcategory airplanes, the distance re-quired to takeoff and climb to a heightof 50 feet above the takeoff surfacemust be determined for each weight,altitude, and temperature within theoperational limits established for take-off with—
(1) Takeoff power on each engine;(2) Wing flaps in the takeoff posi-
tion(s); and(3) Landing gear extended.(c) For commuter category airplanes,
takeoff performance, as required by§§ 23.55 through 23.59, must be deter-mined with the operating engine(s)within approved operating limitations.
[Doc. No. 27807, 61 FR 5185, Feb. 9, 1996]
§ 23.55 Accelerate-stop distance.For each commuter category air-
plane, the accelerate-stop distancemust be determined as follows:
(a) The accelerate-stop distance isthe sum of the distances necessary to—
(1) Accelerate the airplane from astanding start to VEF with all enginesoperating;
(2) Accelerate the airplane from VEF
to V1, assuming the critical enginefails at VEF; and
(3) Come to a full stop from the pointat which V1 is reached.
(b) Means other than wheel brakesmay be used to determine the accel-erate-stop distances if that means—
(1) Is safe and reliable;(2) Is used so that consistent results
can be expected under normal oper-ating conditions; and
(3) Is such that exceptional skill isnot required to control the airplane.
For each commuter category air-plane, the takeoff path is as follows:
(a) The takeoff path extends from astanding start to a point in the takeoffat which the airplane is 1500 feet abovethe takeoff surface at or below whichheight the transition from the takeoffto the enroute configuration must becompleted; and
(1) The takeoff path must be based onthe procedures prescribed in § 23.45;
(2) The airplane must be acceleratedon the ground to VEF at which point thecritical engine must be made inoper-ative and remain inoperative for therest of the takeoff; and
(3) After reaching VEF, the airplanemust be accelerated to V2.
(b) During the acceleration to speedV2, the nose gear may be raised off theground at a speed not less than VR.However, landing gear retraction mustnot be initiated until the airplane isairborne.
(c) During the takeoff path deter-mination, in accordance with para-graphs (a) and (b) of this section—
(1) The slope of the airborne part ofthe takeoff path must not be negativeat any point;
(2) The airplane must reach V2 beforeit is 35 feet above the takeoff surface,and must continue at a speed as closeas practical to, but not less than V2,until it is 400 feet above the takeoffsurface;
(3) At each point along the takeoffpath, starting at the point at which theairplane reaches 400 feet above thetakeoff surface, the available gradientof climb must not be less than—
(i) 1.2 percent for two-engine air-planes;
(ii) 1.5 percent for three-engine air-planes;
(iii) 1.7 percent for four-engine air-planes; and
(4) Except for gear retraction andautomatic propeller feathering, theairplane configuration must not bechanged, and no change in power thatrequires action by the pilot may bemade, until the airplane is 400 feetabove the takeoff surface.
(d) The takeoff path to 35 feet abovethe takeoff surface must be determinedby a continuous demonstrated takeoff.
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14 CFR Ch. I (1–1–01 Edition)§ 23.59
(e) The takeoff path to 35 feet abovethe takeoff surface must be determinedby synthesis from segments; and
(1) The segments must be clearly de-fined and must be related to distinctchanges in configuration, power, andspeed;
(2) The weight of the airplane, theconfiguration, and the power must beassumed constant throughout each seg-ment and must correspond to the mostcritical condition prevailing in the seg-ment; and
(3) The takeoff flight path must bebased on the airplane’s performancewithout utilizing ground effect.
For each commuter category air-plane, the takeoff distance and, at theoption of the applicant, the takeoffrun, must be determined.
(a) Takeoff distance is the greaterof—
(1) The horizontal distance along thetakeoff path from the start of the take-off to the point at which the airplane is35 feet above the takeoff surface as de-termined under § 23.57; or
(2) With all engines operating, 115percent of the horizontal distance fromthe start of the takeoff to the point atwhich the airplane is 35 feet above thetakeoff surface, determined by a proce-dure consistent with § 23.57.
(b) If the takeoff distance includes aclearway, the takeoff run is the greaterof—
(1) The horizontal distance along thetakeoff path from the start of the take-off to a point equidistant between theliftoff point and the point at which theairplane is 35 feet above the takeoffsurface as determined under § 23.57; or
(2) With all engines operating, 115percent of the horizontal distance fromthe start of the takeoff to a point equi-distant between the liftoff point andthe point at which the airplane is 35feet above the takeoff surface, deter-mined by a procedure consistent with§ 23.57.
For each commuter category air-plane, the takeoff flight path must bedetermined as follows:
(a) The takeoff flight path begins 35feet above the takeoff surface at theend of the takeoff distance determinedin accordance with § 23.59.
(b) The net takeoff flight path datamust be determined so that they rep-resent the actual takeoff flight paths,as determined in accordance with§ 23.57 and with paragraph (a) of thissection, reduced at each point by a gra-dient of climb equal to—
(1) 0.8 percent for two-engine air-planes;
(2) 0.9 percent for three-engine air-planes; and
(3) 1.0 percent for four-engine air-planes.
(c) The prescribed reduction in climbgradient may be applied as an equiva-lent reduction in acceleration alongthat part of the takeoff flight path atwhich the airplane is accelerated inlevel flight.
[Amdt. 23–34, 52 FR 1827, Jan. 15, 1987]
§ 23.63 Climb: General.
(a) Compliance with the require-ments of §§ 23.65, 23.66, 23.67, 23.69, and23.77 must be shown—
(1) Out of ground effect; and(2) At speeds that are not less than
those at which compliance with thepowerplant cooling requirements of§§ 23.1041 to 23.1047 has been dem-onstrated; and
(3) Unless otherwise specified, withone engine inoperative, at a bank anglenot exceeding 5 degrees.
(b) For normal, utility, and acrobaticcategory reciprocating engine-poweredairplanes of 6,000 pounds or less max-imum weight, compliance must beshown with § 23.65(a), § 23.67(a), whereappropriate, and § 23.77(a) at maximumtakeoff or landing weight, as appro-priate, in a standard atmosphere.
(c) For normal, utility, and acrobaticcategory reciprocating engine-poweredairplanes of more than 6,000 poundsmaximum weight, and turbine engine-powered airplanes in the normal, util-ity, and acrobatic category, compli-ance must be shown at weights as a
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Federal Aviation Administration, DOT § 23.67
function of airport altitude and ambi-ent temperature, within the oper-ational limits established for takeoffand landing, respectively, with—
(1) Sections 23.65(b) and 23.67(b) (1)and (2), where appropriate, for takeoff,and
(2) Section 23.67(b)(2), where appro-priate, and § 23.77(b), for landing.
(d) For commuter category airplanes,compliance must be shown at weightsas a function of airport altitude andambient temperature within the oper-ational limits established for takeoffand landing, respectively, with—
(1) Sections 23.67(c)(1), 23.67(c)(2), and23.67(c)(3) for takeoff; and
(2) Sections 23.67(c)(3), 23.67(c)(4), and23.77(c) for landing.
[Doc. No. 27807, 61 FR 5186, Feb. 9, 1996]
§ 23.65 Climb: All engines operating.
(a) Each normal, utility, and acro-batic category reciprocating engine-powered airplane of 6,000 pounds or lessmaximum weight must have a steadyclimb gradient at sea level of at least8.3 percent for landplanes or 6.7 percetfor seaplanes and amphibians with—
(1) Not more than maximum contin-uous power on each engine;
(2) The landing gear retracted;(3) The wing flaps in the takeoff posi-
tion(s); and(4) A climb speed not less than the
greater of 1.1 VMC and 1.2 VS1 for multi-engine airplanes and not less than 1.2VS1 for single—engine airplanes.
(b) Each normal, utility, and acro-batic category reciprocating engine-powered airplane of more than 6,000pounds maximum weight and turbineengine-powered airplanes in the nor-mal, utility, and acrobatic categorymust have a steady gradient of climbafter takeoff of at least 4 percent with
(1) Take off power on each engine;(2) The landing gear extended, except
that if the landing gear can be re-tracted in not more than sven seconds,the test may be conducted with thegear retracted;
(3) The wing flaps in the takeoff posi-tion(s); and
(4) A climb speed as specified in§ 23.65(a)(4).
[Doc. No. 27807, 61 FR 5186, Feb. 9, 1996]
§ 23.66 Takeoff climb: One-engine inop-erative.
For normal, utility, and acrobaticcategory reciprocating engine-poweredairplanes of more than 6,000 poundsmaximum weight, and turbine engine-powered airplanes in the normal, util-ity, and acrobatic category, the steadygradient of climb or descent must bedetermined at each weight, altitude,and ambient temperature within theoperational limits established by theapplicant with—
(a) The critical engine inoperativeand its propeller in the position it rap-idly and automatically assumes;
(b) The remaining engine(s) at take-off power;
(c) The landing gear extended, exceptthat if the landing gear can be re-tracted in not more than seven sec-onds, the test may be conducted withthe gear retracted;
(d) The wing flaps in the takeoff posi-tion(s):
(e) The wings level; and(f) A climb speed equal to that
achieved at 50 feet in the demonstra-tion of § 23.53.
[Doc. No. 27807, 61 FR 5186, Feb. 9, 1996]
§ 23.67 Climb: One engine inoperative.
(a) For normal, utility, and acrobaticcategory reciprocating engine-poweredairplanes of 6,000 pounds or less max-imum weight, the following apply:
(1) Except for those airplanes thatmeet the requirements prescribed in§ 23.562(d), each airplane with a VSO ofmore than 61 knots must be able tomaintain a steady climb gradient of atleast 1.5 percent at a pressure altitudeof 5,000 feet with the—
(i) Critical engine inoperative and itspropeller in the minimum drag posi-tion;
(ii) Remaining engine(s) at not morethan maximum continuous power;
(iii) Landing gear retracted;(iv) Wing flaps retracted; and(v) Climb speed not less than 1.2 VS1.(2) For each airplane that meets the
requirements prescribed in § 23.562(d),or that has a VSO of 61 knots or less,the steady gradient of climb or descentat a pressure altitude of 5,000 feet mustbe determined with the—
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14 CFR Ch. I (1–1–01 Edition)§ 23.67
(i) Critical engine inoperative and itspropeller in the minimum drag posi-tion;
(ii) Remaining engine(s) at not morethan maximum continuous power;
(iii) Landing gear retracted;(iv) Wing flaps retracted; and(v) Climb speed not less than 1.2VS1.(b) For normal, utility, and acrobatic
category reciprocating engine-poweredairplanes of more than 6,000 poundsmaximum weight, and turbine engine-powered airplanes in the normal, util-ity, and acrobatic category—
(1) The steady gradient of climb at analtitude of 400 feet above the takeoffmust be measurably positive with the—
(i) Critical engine inoperative and itspropeller in the minimum drag posi-tion;
(ii) Remaining engine(s) at takeoffpower;
(iii) Landing gear retracted;(iv) Wing flaps in the takeoff posi-
tion(s); and(v) Climb speed equal to that
achieved at 50 feet in the demonstra-tion of § 23.53.
(2) The steady gradient of climb mustnot be less than 0.75 percent at an alti-tude of 1,500 feet above the takeoff sur-face, or landing surface, as appropriate,with the—
(i) Critical engine inoperative and itspropeller in the minimum drag posi-tion;
(ii) Remaining engine(s) at not morethan maximum continuous power;
(iii) Landing gear retracted;(iv) Wing flaps retracted; and(v) Climb speed not less than 1.2 VS1.(c) For commuter category airplanes,
the following apply:(1) Takeoff; landing gear extended. The
steady gradient of climb at the altitudeof the takeoff surface must be measur-ably positive for two-engine airplanes,not less than 0.3 percent for three-en-gine airplanes, or 0.5 percent for four-engine airplanes with—
(i) The critical engine inoperativeand its propeller in the position it rap-idly and automatically assumes;
(ii) The remaining engine(s) at take-off power;
(iii) The landing gear extended, andall landing gear doors open;
(iv) The wing flaps in the takeoff po-sition(s);
(v) The wings level; and(vi) A climb speed equal to V2.(2) Takeoff; landing gear retracted. The
steady gradient of climb at an altitudeof 400 feet above the takeoff surfacemust be not less than 2.0 percent oftwo-engine airplanes, 2.3 percent forthree-engine airplanes, and 2.6 percentfor four-engine airplanes with—
(i) The critical engine inoperativeand its propeller in the position it rap-idly and automatically assumes;
(ii) The remaining engine(s) at take-off power;
(iii) The landing gear retracted;(iv) The wing flaps in the takeoff po-
sition(s);(v) A climb speed equal to V2.(3) Enroute. The steady gradient of
climb at an altitude of 1,500 feet abovethe takeoff or landing surface, as ap-propriate, must be not less than 1.2percent for two-engine airplanes, 1.5percent for three-engine airplanes, and1.7 percent for four-engine airplaneswith—
(i) The critical engine inoperativeand its propeller in the minimum dragposition;
(ii) The remaining engine(s) at notmore than maximum continuouspower;
(iii) The landing gear retracted;(iv) The wing flaps retracted; and(v) A climb speed not less than 1.2
VS1.(4) Discontinued approach. The steady
gradient of climb at an altitude of 400feet above the landing surface must benot less than 2.1 percent for two-engineairplanes, 2.4 percent for three-engineairplanes, and 2.7 percent for four-en-gine airplanes, with—
(i) The critical engine inoperativeand its propeller in the minimum dragposition;
(ii) The remaining engine(s) at take-off power;
(iii) Landing gear retracted;(iv) Wing flaps in the approach posi-
tion(s) in which VS1 for these posi-tion(s) does not exceed 110 percent ofthe VS1 for the related all-engines-oper-ated landing position(s); and
(v) A climb speed established in con-nection with normal landing proce-dures but not exceeding 1.5 VS1.
[Doc. No. 27807, 61 FR 5186, Feb. 9, 1996]
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Federal Aviation Administration, DOT § 23.75
§ 23.69 Enroute climb/descent.(a) All engines operating. The steady
gradient and rate of climb must be de-termined at each weight, altitude, andambient temperature within the oper-ational limits established by the appli-cant with—
(1) Not more than maximum contin-uous power on each engine;
(2) The landing gear retracted;(3) The wing flaps retracted; and(4) A climb speed not less than 1.3
VS1.(b) One engine inoperative. The steady
gradient and rate of climb/descentmust be determined at each weight, al-titude, and ambient temperature with-in the operational limits established bythe applicant with—
(1) The critical engine inoperativeand its propeller in the minimum dragposition;
(2) The remaining engine(s) at notmore than maximum continuouspower;
(3) The landing gear retracted;(4) The wing flaps retracted; and(5) A climb speed not less than 1.2
VS1.
[Doc. No. 27807, 61 FR 5187, Feb. 9, 1996]
§ 23.71 Glide: Single-engine airplanes.The maximum horizontal distance
traveled in still air, in nautical miles,per 1,000 feet of altitude lost in a glide,and the speed necessary to achieve thismust be determined with the engine in-operative, its propeller in the min-imum drag position, and landing gearand wing flaps in the most favorableavailable position.
[Doc. No. 27807, 61 FR 5187, Feb. 9, 1996]
§ 23.73 Reference landing approachspeed.
(a) For normal, utility, and acrobaticcategory reciprocating engine-poweredairplanes of 6,000 pounds or less max-imum weight, the reference landing ap-proach speed, VREF, must not be lessthan the greater of VMC, determined in§ 23.149(b) with the wing flaps in themost extended takeoff position, and 1.3VSO.
(b) For normal, utility, and acrobaticcategory reciprocating engine-poweredairplanes of more than 6,000 poundsmaximum weight, and turbine engine-
powered airplanes in the normal, util-ity, and acrobatic category, the ref-erence landing approach speed, VREF,must not be less than the greater ofVMC, determined in § 23.149(c), and 1.3VSO.
(c) For commuter category airplanes,the reference landing approach speed,VREF, must not be less than the greaterof 1.05 VMC, determined in § 23.149(c),and 1.3 VSO.
[Doc. No. 27807, 61 FR 5187, Feb. 9, 1996]
§ 23.75 Landing distance.The horizontal distance necessary to
land and come to a complete stop froma point 50 feet above the landing sur-face must be determined, for standardtemperatures at each weight and alti-tude within the operational limits es-tablished for landing, as follows:
(a) A steady approach at not lessthan VREF, determined in accordancewith § 23.73 (a), (b), or (c), as appro-priate, must be maintained down to the50 foot height and—
(1) The steady approach must be at agradient of descent not greater than 5.2percent (3 degrees) down to the 50-footheight.
(2) In addition, an applicant maydemonstrate by tests that a maximumsteady approach gradient steeper than5.2 percent, down to the 50-foot height,is safe. The gradient must be estab-lished as an operating limitation andthe information necessary to displaythe gradient must be available to thepilot by an appropriate instrument.
(b) A constant configuration must bemaintained throughout the maneuver.
(c) The landing must be made with-out excessive vertical acceleration ortendency to bounce, nose over, groundloop, porpoise, or water loop.
(d) It must be shown that a safe tran-sition to the balked landing conditionsof § 23.77 can be made from the condi-tions that exist at the 50 foot height, atmaximum landing weight, or at themaximum landing weight for altitudeand temperature of § 23.63 (c)(2) or(d)(2), as appropriate.
(e) The brakes must be used so as tonot cause excessive wear of brakes ortires.
(f) Retardation means other thanwheel brakes may be used if thatmeans—
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14 CFR Ch. I (1–1–01 Edition)§ 23.77
(1) Is safe and reliable; and(2) Is used so that consistent results
can be expected in service.(g) If any device is used that depends
on the operation of any engine, and thelanding distance would be increasedwhen a landing is made with that en-gine inoperative, the landing distancemust be determined with that engineinoperative unless the use of othercompensating means will result in alanding distance not more than thatwith each engine operating.
§ 23.77 Balked landing.(a) Each normal, utility, and acro-
batic category reciprocating engine-powered airplane at 6,000 pounds or lessmaximum weight must be able tomaintain a steady gradient of climb atsea level of at least 3.3 percent with—
(1) Takeoff power on each engine;(2) The landing gear extended;(3) The wing flaps in the landing posi-
tion, except that if the flaps may safelybe retracted in two seconds or lesswithout loss of altitude and withoutsudden changes of angle of attack, theymay be retracted; and
(4) A climb speed equal to VREF, as de-fined in § 23.73(a).
(b) Each normal, utility, and acro-batic category reciprocating engine-powered airplane of more than 6,000pounds maximum weight and each nor-mal, utility, and acrobatic categoryturbine engine-powered airplane mustbe able to maintain a steady gradientof climb of at least 2.5 percent with—
(1) Not more than the power that isavailable on each engine eight secondsafter initiation of movement of thepower controls from minimum flight-idle position;
(2) The landing gear extended;(3) The wing flaps in the landing posi-
tion; and(4) A climb speed equal to VREF, as de-
fined in § 23.73(b).(c) Each commuter category airplane
must be able to maintain a steady gra-dient of climb of at least 3.2 percentwith—
(1) Not more than the power that isavailable on each engine eight seconds
after initiation of movement of thepower controls from the minimumflight idle position;
(2) Landing gear extended;(3) Wing flaps in the landing position;
and(4) A climb speed equal to VREF, as de-
fined in § 23.73(c).
[Doc. No. 27807, 61 FR 5187, Feb. 9, 1996]
FLIGHT CHARACTERISTICS
§ 23.141 General.
The airplane must meet the require-ments of §§ 23.143 through 23.253 at allpractical loading conditions and oper-ating altitudes for which certificationhas been requested, not exceeding themaximum operating altitude estab-lished under § 23.1527, and without re-quiring exceptional piloting skill,alertness, or strength.
[Doc. No. 26269, 58 FR 42156, Aug. 6, 1993]
CONTROLLABILITY ANDMANEUVERABILITY
§ 23.143 General.
(a) The airplane must be safely con-trollable and maneuverable during allflight phases including—
(1) Takeoff;(2) Climb;(3) Level flight;(4) Descent;(5) Go-around; and(6) Landing (power on and power off)
with the wing flaps extended and re-tracted.
(b) It must be possible to make asmooth transition from one flight con-dition to another (including turns andslips) without danger of exceeding thelimit load factor, under any probableoperating condition (including, formultiengine airplanes, those condi-tions normally encountered in the sud-den failure of any engine).
(c) If marginal conditions exist withregard to required pilot strength, thecontrol forces necessary must be deter-mined by quantitative tests. In no casemay the control forces under the condi-tions specified in paragraphs (a) and (b)of this section exceed those prescribedin the following table:
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189
Federal Aviation Administration, DOT § 23.145
Values in pounds force appliedto the relevant control Pitch Roll Yaw
(a) For temporary application:Stick ....................................... 60 30 ............Wheel (Two hands on rim) .... 75 50 ............Wheel (One hand on rim) ...... 50 25 ............Rudder Pedal ......................... ............ ............ 150
§ 23.145 Longitudinal control.(a) With the airplane as nearly as
possible in trim at 1.3 VS1, it must bepossible, at speeds below the trimspeed, to pitch the nose downward sothat the rate of increase in airspeed al-lows prompt acceleration to the trimspeed with—
(1) Maximum continuous power oneach engine;
(2) Power off; and(3) Wing flap and landing gear—(i) retracted, and(ii) extended.(b) Unless otherwise required, it must
be possible to carry out the followingmaneuvers without requiring the appli-cation of single-handed control forcesexceeding those specified in § 23.143(c).The trimming controls must not be ad-justed during the maneuvers:
(1) With the landing gear extended,the flaps retracted, and the airplanesas nearly as possible in trim at 1.4 VS1,extend the flaps as rapidly as possibleand allow the airspeed to transitionfrom 1.4VS1 to 1.4 VSO™
(i) With power off; and(ii) With the power necessary to
maintain level flight in the initial con-dition.
(2) With landing gear and flaps ex-tended, power off, and the airplane asnearly as possible in trim at 1.3 VSO,quickly apply takeoff power and re-tract the flaps as rapidly as possible tothe recommended go around settingand allow the airspeed to transitionfrom 1.3 VSO to 1.3 VS1. Retract the gearwhen a positive rate of climb is estab-lished.
(3) With landing gear and flaps ex-tended, in level flight, power necessaryto attain level flight at 1.1 VSO, and theairplane as nearly as possible in trim,
it must be possible to maintain ap-proximately level flight while retract-ing the flaps as rapidly as possible withsimultaneous application of not morethan maximum continuous power. Ifgated flat positions are provided, theflap retraction may be demonstrated instages with power and trim reset forlevel flight at 1.1 VS1, in the initial con-figuration for each stage—
(i) From the fully extended positionto the most extended gated position;
(ii) Between intermediate gated posi-tions, if applicable; and
(iii) From the least extended gatedposition to the fully retracted position.
(4) With power off, flaps and landinggear retracted and the airplane asnearly as possible in trim at 1.4 VS1,apply takeoff power rapidly whilemaintaining the same airspeed.
(5) With power off, landing gear andflaps extended, and the airplane asnearly as possible in trim at VREF, ob-tain and maintain airspeeds between1.1 VSO, and either 1.7 VSO or VFE,whichever is lower without requiringthe application of two-handed controlforces exceeding those specified in§ 23.143(c).
(6) With maximum takeoff power,landing gear retracted, flaps in thetakeoff position, and the airplane asnearly as possible in trim at VFE appro-priate to the takeoff flap position, re-tract the flaps as rapidly as possiblewhile maintaining constant speed.
(c) At speeds above VMO/MMO, and upto the maximum speed shown under§ 23.251, a maneuvering capability of 1.5g must be demonstrated to provide amargin to recover from upset or inad-vertent speed increase.
(d) It must be possible, with a pilotcontrol force of not more than 10pounds, to maintain a speed of notmore than VREF during a power-off glidewith landing gear and wing flaps ex-tended, for any weight of the airplane,up to and including the maximumweight.
(e) By using normal flight and powercontrols, except as otherwise noted inparagraphs (e)(1) and (e)(2) of this sec-tion, it must be possible to establish azero rate of descent at an attitude suit-able for a controlled landing without
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190
14 CFR Ch. I (1–1–01 Edition)§ 23.147
exceeding the operational and struc-tural limitations of the airplane, asfollows:
(1) For single-engine and multiengineairplanes, without the use of the pri-mary longitudinal control system.
(2) For multiengine airplanes—(i) Without the use of the primary di-
rectional control; and(ii) If a single failure of any one con-
necting or transmitting link would af-fect both the longitudinal and direc-tional primary control system, withoutthe primary longitudinal and direc-tional control system.
(a) For each multiengine airplane, itmust be possible, while holding thewings level within five degrees, tomake sudden changes in heading safelyin both directions. This ability must beshown at 1.4 VS1 with heading changesup to 15 degrees, except that the head-ing change at which the rudder forcecorresponds to the limits specified in§ 23.143 need not be exceeded, with the—
(1) Critical engine inoperative and itspropeller in the minimum drag posi-tion;
(2) Remaining engines at maximumcontinuous power;
(3) Landing gear—(i) Retracted; and(ii) Extended; and(4) Flaps retracted.(b) For each multiengine airplane, it
must be possible to regain full controlof the airplane without exceeding abank angle of 45 degrees, reaching adangerous attitude or encounteringdangerous characteristics, in the eventof a sudden and complete failure of thecritical engine, making allowance for adelay of two seconds in the initiationof recovery action appropriate to thesituation, with the airplane initially intrim, in the following condition:
(1) Maximum continuous power oneach engine;
(2) The wing flaps retracted;(3) The landing gear retracted;
(4) A speed equal to that at whichcompliance with § 23.69(a) has beenshown; and
(5) All propeller controls in the posi-tion at which compliance with § 23.69(a)has been shown.
(c) For all airplanes, it must beshown that the airplane is safely con-trollable without the use of the pri-mary lateral control system in any all-engine configuration(s) and at anyspeed or altitude within the approvedoperating envelope. It must also beshown that the airplane’s flight char-acteristics are not impaired below alevel needed to permit continued safeflight and the ability to maintain atti-tudes suitable for a controlled landingwithout exceeding the operational andstructural limitations of the airplane.If a single failure of any one con-necting or transmitting link in the lat-eral control system would also causethe loss of additional control sys-tem(s), compliance with the above re-quirement must be shown with thoseadditional systems also assumed to beinoperative.
[Doc. No. 27807, 61 FR 5188, Feb. 9, 1996]
§ 23.149 Minimum control speed.(a) VMC is the calibrated airspeed at
which, when the critical engine is sud-denly made inoperative, it is possibleto maintain control of the airplanewith that engine still inoperative, andthereafter maintain straight flight atthe same speed with an angle of bankof not more than 5 degrees. The methodused to simulate critical engine failuremust represent the most critical modeof powerplant failure expected in serv-ice with respect to controllability.
(b) VMC for takeoff must not exceed1.2 VS1, where VS1 is determined at themaximum takeoff weight. VMC must bedetermined with the most unfavorableweight and center of gravity positionand with the airplane airborne and theground effect negligible, for the takeoffconfiguration(s) with—
(1) Maximum available takeoff powerinitially on each engine;
(2) The airplane trimmed for takeoff;(3) Flaps in the takeoff position(s);(4) Landing gear retracted; and(5) All propeller controls in the rec-
ommended takeoff position through-out.
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191
Federal Aviation Administration, DOT § 23.155
(c) For all airplanes except recipro-cating engine-powered airplanes of6,000 pounds or less maximum weight,the conditions of paragraph (a) of thissection must also be met for the land-ing configuration with—
(1) Maximum available takeoff powerinitially on each engine;
(2) The airplane trimmed for an ap-proach, with all engines operating, atVREF, at an approach gradient equal tothe steepest used in the landing dis-tance demonstration of § 23.75;
(3) Flaps in the landing position;(4) Landing gear extended; and(5) All propeller controls in the posi-
tion recommended for approach withall engines operating.
(d) A minimum speed to inten-tionally render the critical engine in-operative must be established and des-ignated as the safe, intentional, one-engine-inoperative speed, VSSE.
(e) At VMC, the rudder pedal force re-quired to maintain control must notexceed 150 pounds and it must not benecessary to reduce power of the opera-tive engine(s). During the maneuver,the airplane must not assume any dan-gerous attitude and it must be possibleto prevent a heading change of morethan 20 degrees.
(f) At the option of the applicant, tocomply with the requirements of§ 23.51(c)(1), VMCG may be determined.VMCG is the minimum control speed onthe ground, and is the calibrated air-speed during the takeoff run at which,when the critical engine is suddenlymade inoperative, it is possible tomaintain control of the airplane usingthe rudder control alone (without theuse of nosewheel steering), as limitedby 150 pounds of force, and using thelateral control to the extent of keepingthe wings level to enable the takeoff tobe safely continued. In the determina-tion of VMCG, assuming that the path ofthe airplane accelerating with all en-gines operating is along the centerlineof the runway, its path from the pointat which the critical engine is made in-operative to the point at which recov-ery to a direction parallel to the cen-terline is completed may not deviatemore than 30 feet laterally from thecenterline at any point. VMCG must beestablished with—
(1) The airplane in each takeoff con-figuration or, at the option of the ap-plicant, in the most critical takeoffconfiguration;
(2) Maximum available takeoff poweron the operating engines;
(3) The most unfavorable center ofgravity;
(4) The airplane trimmed for takeoff;and
(5) The most unfavorable weight inthe range of takeoff weights.
[Doc. No. 27807, 61 FR 5189, Feb. 9, 1996]
§ 23.151 Acrobatic maneuvers.
Each acrobatic and utility categoryairplane must be able to perform safelythe acrobatic maneuvers for which cer-tification is requested. Safe entryspeeds for these maneuvers must be de-termined.
§ 23.153 Control during landings.
It must be possible, while in the land-ing configuration, to safely complete alanding without exceeding the one-hand control force limits specified in§ 23.143(c) following an approach toland—
(a) At a speed of VREF minus 5 knots;(b) With the airplane in trim, or as
nearly as possible in trim and withoutthe trimming control being movedthroughout the maneuver;
(c) At an approach gradient equal tothe steepest used in the landing dis-tance demonstration of § 23.75; and
(d) With only those power changes, ifany, that would be made when landingnormally from an approach at VREF.
[Doc. No. 27807, 61 FR 5189, Feb. 9, 1996]
§ 23.155 Elevator control force in ma-neuvers.
(a) The elevator control force neededto achieve the positive limit maneu-vering load factor may not be lessthan:
(1) For wheel controls, W/100 (whereW is the maximum weight) or 20pounds, whichever is greater, exceptthat it need not be greater than 50pounds; or
(2) For stick controls, W/140 (where Wis the maximum weight) or 15 pounds,whichever is greater, except that itneed not be greater than 35 pounds.
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14 CFR Ch. I (1–1–01 Edition)§ 23.157
(b) The requirement of paragraph (a)of this section must be met at 75 per-cent of maximum continuous power forreciprocating engines, or the maximumcontinuous power for turbine engines,and with the wing flaps and landinggear retracted—
(1) In a turn, with the trim settingused for wings level flight at VO; and
(2) In a turn with the trim settingused for the maximum wings levelflight speed, except that the speed maynot exceed VNE or VMO/MMO, whicheveris appropriate.
(c) There must be no excessive de-crease in the gradient of the curve ofstick force versus maneuvering loadfactor with increasing load factor.
§ 23.157 Rate of roll.(a) Takeoff. It must be possible, using
a favorable combination of controls, toroll the airplane from a steady 30-de-gree banked turn through an angle of60 degrees, so as to reverse the direc-tion of the turn within:
(1) For an airplane of 6,000 pounds orless maximum weight, 5 seconds frominitiation of roll; and
(2) For an airplane of over 6,000pounds maximum weight,
(W+500)/1,300
seconds, but not more than 10 seconds,where W is the weight in pounds.
(b) The requirement of paragraph (a)of this section must be met when roll-ing the airplane in each directionwith—
(1) Flaps in the takeoff position;(2) Landing gear retracted;(3) For a single-engine airplane, at
maximum takeoff power; and for amultiengine airplane with the criticalengine inoperative and the propeller inthe minimum drag position, and theother engines at maximum takeoffpower; and
(4) The airplane trimmed at a speedequal to the greater of 1.2 VS1 or 1.1VMC, or as nearly as possible in trim forstraight flight.
(c) Approach. It must be possible,using a favorable combination of con-trols, to roll the airplane from a steady
30-degree banked turn through an angleof 60 degrees, so as to reverse the direc-tion of the turn within:
(1) For an airplane of 6,000 pounds orless maximum weight, 4 seconds frominitiation of roll; and
(2) For an airplane of over 6,000pounds maximum weight,
(W+2,800)/2,200
seconds, but not more than 7 seconds,where W is the weight in pounds.
(d) The requirement of paragraph (c)of this section must be met when roll-ing the airplane in each direction inthe following conditions—
(1) Flaps in the landing position(s);(2) Landing gear extended;(3) All engines operating at the power
for a 3 degree approach; and(4) The airplane trimmed at VREF.
§ 23.161 Trim.(a) General. Each airplane must meet
the trim requirements of this sectionafter being trimmed and without fur-ther pressure upon, or movement of,the primary controls or their cor-responding trim controls by the pilotor the automatic pilot. In addition, itmust be possible, in other conditions ofloading, configuration, speed and powerto ensure that the pilot will not be un-duly fatigued or distracted by the needto apply residual control forces exceed-ing those for prolonged application of§ 23.143(c). This applies in normal oper-ation of the airplane and, if applicable,to those conditions associated with thefailure of one engine for which per-formance characteristics are estab-lished.
(b) Lateral and directional trim. Theairplane must maintain lateral and di-rectional trim in level flight with thelanding gear and wing flaps retractedas follows:
(1) For normal, utility, and acrobaticcategory airplanes, at a speed of 0.9 VH,VC, or VMO/MO, whichever is lowest; and
(2) For commuter category airplanes,at all speeds from 1.4 VS1 to the lesserof VH or VMO/MMO.
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Federal Aviation Administration, DOT § 23.173
(c) Longitudinal trim. The airplanemust maintain longitudinal trim undereach of the following conditions:
(1) A climb with—(i) Takeoff power, landing gear re-
tracted, wing flaps in the takeoff posi-tion(s), at the speeds used in deter-mining the climb performance requiredby § 23.65; and
(ii) Maximum continuous power atthe speeds and in the configurationused in determining the climb perform-ance required by § 23.69(a).
(2) Level flight at all speeds from thelesser of VH and either VNO or VMO/MMO
(as appropriate), to 1.4 VS1, with thelanding gear and flaps retracted.
(3) A descent at VNO or VMO/MMO,whichever is applicable, with power offand with the landing gear and flaps re-tracted.
(4) Approach with landing gear ex-tended and with—
(i) A 3 degree angle of descent, withflaps retracted and at a speed of 1.4 VS1;
(ii) A 3 degree angle of descent, flapsin the landing position(s) at VREF; and
(iii) An approach gradient equal tothe steepest used in the landing dis-tance demonstrations of § 23.75, flaps inthe landing position(s) at VREF.
(d) In addition, each multiple air-plane must maintain longitudinal anddirectional trim, and the lateral con-trol force must not exceed 5 pounds atthe speed used in complying with§ 23.67(a), (b)(2), or (c)(3), as appro-priate, with—
(1) The critical engine inoperative,and if applicable, its propeller in theminimum drag position;
(2) The remaining engines at max-imum continuous power;
(3) The landing gear retracted;(4) Wing flaps retracted; and(5) An angle of bank of not more than
five degrees.(e) In addition, each commuter cat-
egory airplane for which, in the deter-mination of the takeoff path in accord-ance with § 23.57, the climb in the take-off configuration at V2 extends beyond400 feet above the takeoff surface, itmust be possible to reduce the longitu-dinal and lateral control forces to 10pounds and 5 pounds, respectively, andthe directional control force must notexceed 50 pounds at V2 with—
(1) The critical engine inoperativeand its propeller in the minimum dragposition;
(2) The remaining engine(s) at take-off power;
(3) Landing gear retracted;(4) Wing flaps in the takeoff posi-
The airplane must be longitudinally,directionally, and laterally stableunder §§ 23.173 through 23.181. In addi-tion, the airplane must show suitablestability and control ‘‘feel’’ (static sta-bility) in any condition normally en-countered in service, if flight testsshow it is necessary for safe operation.
§ 23.173 Static longitudinal stability.
Under the conditions specified in§ 23.175 and with the airplane trimmedas indicated, the characteristics of theelevator control forces and the frictionwithin the control system must be asfollows:
(a) A pull must be required to obtainand maintain speeds below the speci-fied trim speed and a push required toobtain and maintain speeds above thespecified trim speed. This must beshown at any speed that can be ob-tained, except that speeds requiring acontrol force in excess of 40 pounds orspeeds above the maximum allowablespeed or below the minimum speed forsteady unstalled flight, need not beconsidered.
(b) The airspeed must return to with-in the tolerances specified for applica-ble categories of airplanes when thecontrol force is slowly released at anyspeed within the speed range specifiedin paragraph (a) of this section. The ap-plicable tolerances are—
(1) The airspeed must return to with-in plus or minus 10 percent of the origi-nal trim airspeed; and
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14 CFR Ch. I (1–1–01 Edition)§ 23.175
(2) For commuter category airplanes,the airspeed must return to within plusor minus 7.5 percent of the originaltrim airspeed for the cruising conditionspecified in § 23.175(b).
(c) The stick force must vary withspeed so that any substantial speedchange results in a stick force clearlyperceptible to the pilot.
§ 23.175 Demonstration of static longi-tudinal stability.
Static longitudinal stability must beshown as follows:
(a) Climb. The stick force curve musthave a stable slope at speeds between85 and 115 percent of the trim speed,with—
(1) Flaps retracted;(2) Landing gear retracted;(3) Maximum continuous power; and(4) The airplane trimmed at the speed
used in determining the climb perform-ance required by § 23.69(a).
(b) Cruise. With flaps and landinggear retracted and the airplane in trimwith power for level flight at represent-ative cruising speeds at high and lowaltitudes, including speeds up to VNO orVMO/MMO, as appropriate, except thatthe speed need not exceed VH—
(1) For normal, utility, and acrobaticcategory airplanes, the stick forcecurve must have a stable slope at allspeeds within a range that is the great-er of 15 percent of the trim speed plusthe resulting free return speed range,or 40 knots plus the resulting free re-turn speed range, above and below thetrim speed, except that the slope neednot be stable—
(i) At speeds less than 1.3 VS1; or(ii) For airplanes with VNE estab-
lished under § 23.1505(a), at speedsgreater than VNE; or
(iii) For airplanes with VMO/MMO es-tablished under § 23.1505(c), at speedsgreater than VFC/MFC.
(2) For commuter category airplanes,the stick force curve must have a sta-ble slope at all speeds within a range of50 knots plus the resulting free returnspeed range, above and below the trimspeed, except that the slope need not bestable—
(i) At speeds less than 1.4 VS1; or
(ii) At speeds greater than VFC/MFC;or
(iii) At speeds that require a stickforce greater than 50 pounds.
(c) Landing. The stick force curvemust have a stable slope at speeds be-tween 1.1 VS1 and 1.8 VS1 with—
(1) Flaps in the landing position;(2) Landing gear extended; and(3) The airplane trimmed at—(i) VREF, or the minimum trim speed
if higher, with power off; and(ii) VREF with enough power to main-
tain a 3 degree angle of descent.
[Doc. No. 27807, 61 FR 5190, Feb. 9, 1996]
§ 23.177 Static directional and lateralstability.
(a) The static directional stability, asshown by the tendency to recover froma wings level sideslip with the rudderfree, must be positive for any landinggear and flap position appropriate tothe takeoff, climb, cruise, approach,and landing configurations. This mustbe shown with symmetrical power upto maximum continuous power, and atspeeds from 1.2 VS1 up to the maximumallowable speed for the condition beinginvestigated. The angel of sideslip forthese tests must be appropriate to thetype of airplane. At larger angles ofsideslip, up to that at which full rudderis used or a control force limit in§ 23.143 is reached, whichever occursfirst, and at speeds from 1.2 VS1 to VO,the rudder pedal force must not re-verse.
(b) The static lateral stability, asshown by the tendency to raise the lowwing in a sideslip, must be positive forall landing gear and flap positions.This must be shown with symmetricalpower up to 75 percent of maximumcontinuous power at speeds above 1.2VS1 in the take off configuration(s) andat speeds above 1.3 VS1 in other con-figurations, up to the maximum allow-able speed for the configuration beinginvestigated, in the takeoff, climb,cruise, and approach configurations.For the landing configuration, thepower must be that necessary to main-tain a 3 degree angle of descent in co-ordinated flight. The static lateral sta-bility must not be negative at 1.2 VS1 inthe takeoff configuration, or at 1.3 VS1
in other configurations. The angle of
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Federal Aviation Administration, DOT § 23.201
sideslip for these tests must be appro-priate to the type of airplane, but in nocase may the constant heading sideslipangle be less than that obtainable witha 10 degree bank, or if less, the max-imum bank angle obtainable with fullrudder deflection or 150 pound rudderforce.
(c) Paragraph (b) of this section doesnot apply to acrobatic category air-planes certificated for inverted flight.
(d) In straight, steady slips at 1.2 VS1
for any landing gear and flap positions,and for any symmetrical power condi-tions up to 50 percent of maximum con-tinuous power, the aileron and ruddercontrol movements and forces must in-crease steadily, but not necessarily inconstant proportion, as the angle ofsideslip is increased up to the max-imum appropriate to the type of air-plane. At larger slip angles, up to theangle at which full rudder or aileroncontrol is used or a control force limitcontained in § 23.143 is reached, the ai-leron and rudder control movementsand forces must not reverse as theangle of sideslip is increased. Rapidentry into, and recovery from, a max-imum sideslip considered appropriatefor the airplane must not result in un-controllable flight characteristics.
[Doc. No. 27807, 61 FR 5190, Feb. 9, 1996]
§ 23.181 Dynamic stability.(a) Any short period oscillation not
including combined lateral-directionaloscillations occurring between thestalling speed and the maximum allow-able speed appropriate to the configu-ration of the airplane must be heavilydamped with the primary controls—
(1) Free; and(2) In a fixed position.(b) Any combined lateral-directional
oscillations (‘‘Dutch roll’’) occurringbetween the stalling speed and themaximum allowable speed appropriateto the configuration of the airplanemust be damped to 1/10 amplitude in 7cycles with the primary controls—
(1) Free; and(2) In a fixed position.(c) If it is determined that the func-
tion of a stability augmentation sys-tem, reference § 23.672, is needed tomeet the flight characteristic require-ments of this part, the primary controlrequirements of paragraphs (a)(2) and
(b)(2) of this section are not applicableto the tests needed to verify the ac-ceptability of that system.
(d) During the conditions as specifiedin § 23.175, when the longitudinal con-trol force required to maintain speedsdiffering from the trim speed by atleast plus and minus 15 percent is sud-denly released, the response of the air-plane must not exhibit any dangerouscharacteristics nor be excessive in rela-tion to the magnitude of the controlforce released. Any long-period oscilla-tion of flight path, phugoid oscillation,that results must not be so unstable asto increase the pilot’s workload or oth-erwise endanger the airplane.
[Amdt. 23–21, 43 FR 2318, Jan. 16, 1978, asamended by Amdt. 23–45, 58 FR 42158, Aug. 6,1993]
STALLS
§ 23.201 Wings level stall.(a) It must be possible to produce and
to correct roll by unreversed use of therolling control and to produce and tocorrect yaw by unreversed use of thedirectional control, up to the time theairplane stalls.
(b) The wings level stall characteris-tics must be demonstrated in flight asfollows. Starting from a speed at least10 knots above the stall speed, the ele-vator control must be pulled back sothat the rate of speed reduction willnot exceed one knot per second until astall is produced, as shown by either:
(1) An uncontrollable downwardpitching motion of the airplane;
(2) A downward pitching motion ofthe airplane that results from the acti-vation of a stall avoidance device (forexample, stick pusher); or
(3) The control reaching the stop.(c) Normal use of elevator control for
recovery is allowed after the downwardpitching motion of paragraphs (b)(1) or(b)(2) of this section has unmistakablybeen produced, or after the control hasbeen held against the stop for not lessthan the longer of two seconds or thetime employed in the minimum steadyslight speed determination of § 23.49.
(d) During the entry into and the re-covery from the maneuver, it must bepossible to prevent more than 15 de-grees of roll or yaw by the normal useof controls.
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14 CFR Ch. I (1–1–01 Edition)§ 23.203
(e) Compliance with the require-ments of this section must be shownunder the following conditions:
(1) Wing flaps. Retracted, fully ex-tended, and each intermediate normaloperating position.
(2) Landing gear. Retracted and ex-tended.
(3) Cowl flaps. Appropriate to configu-ration.
(4) Power:(i) Power off; and(ii) 75 percent of maximum contin-
uous power. However, if the power-to-weight ratio at 75 percent of maximumcontinuous power result in extremenose-up attitudes, the test may be car-ried out with the power required forlevel flight in the landing configura-tion at maximum landing weight and aspeed of 1.4 VSO, except that the powermay not be less than 50 percent of max-imum continuous power.
(5) Trim. The airplane trimmed at aspeed as near 1.5 VS1 as practicable.
(6) Propeller. Full increase r.p.m. posi-tion for the power off condition.
[Doc. No. 27807, 61 FR 5191, Feb. 9, 1996]
§ 23.203 Turning flight and acceleratedturning stalls.
Turning flight and accelerated turn-ing stalls must be demonstrated intests as follows:
(a) Establish and maintain a coordi-nated turn in a 30 degree bank. Reducespeed by steadily and progressivelytightening the turn with the elevatoruntil the airplane is stalled, as definedin § 23.201(b). The rate of speed reduc-tion must be constant, and—
(1) For a turning flight stall, may notexceed one knot per second; and
(2) For an accelerated turning stall,be 3 to 5 knots per second with steadilyincreasing normal acceleration.
(b) After the airplane has stalled, asdefined in § 23.201(b), it must be possibleto regain wings level flight by normaluse of the flight controls, but withoutincreasing power and without—
(1) Excessive loss of altitude;(2) Undue pitchup;(3) Uncontrollable tendency to spin;(4) Exceeding a bank angle of 60 de-
grees in the original direction of theturn or 30 degrees in the opposite direc-tion in the case of turning flight stalls;
(5) Exceeding a bank angle of 90 de-grees in the original direction of theturn or 60 degrees in the opposite direc-tion in the case of accelerated turningstalls; and
(6) Exceeding the maximum permis-sible speed or allowable limit load fac-tor.
(c) Compliance with the require-ments of this section must be shownunder the following conditions:
(1) Wing flaps: Retracted, fully ex-tended, and each intermediate normaloperating position;
(2) Landing gear: Retracted and ex-tended;
(3) Cowl flaps: Appropriate to configu-ration;
(4) Power:(i) Power off; and(ii) 75 percent of maximum contin-
uous power. However, if the power-to-weight ratio at 75 percent of maximumcontinuous power results in extremenose-up attitudes, the test may be car-ried out with the power required forlevel flight in the landing configura-tion at maximum landing weight and aspeed of 1.4 VSO, except that the powermay not be less than 50 percent of max-imum continuous power.
(5) Trim: The airplane trimmed at aspeed as near 1.5 VS1 as practicable.
(6) Propeller. Full increase rpm posi-tion for the power off condition.
§ 23.207 Stall warning.(a) There must be a clear and distinc-
tive stall warning, with the flaps andlanding gear in any normal position, instraight and turning flight.
(b) The stall warning may be fur-nished either through the inherent aer-odynamic qualities of the airplane orby a device that will give clearly dis-tinguishable indications under ex-pected conditions of flight. However, avisual stall warning device that re-quires the attention of the crew withinthe cockpit is not acceptable by itself.
(c) During the stall tests required by§ 23.201(b) and § 23.203(a)(1), the stallwarning must begin at a speed exceed-ing the stalling speed by a margin ofnot less than 5 knots and must con-tinue until the stall occurs.
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Federal Aviation Administration, DOT § 23.221
(d) When following procedures fur-nished in accordance with § 23.1585, thestall warning must not occur during atakeoff with all engines operating, atakeoff continued with one engine in-operative, or during an approach tolanding.
(e) During the stall tests required by§ 23.203(a)(2), the stall warning mustbegin sufficiently in advance of thestall for the stall to be averted by pilotaction taken after the stall warningfirst occurs.
(f) For acrobatic category airplanes,an artificial stall warning may be mu-table, provided that it is armed auto-matically during takeoff and rearmedautomatically in the approach configu-ration.
[Amdt. 23–7, 34 FR 13087, Aug. 13, 1969, asamended by Amdt. 23–45, 58 FR 42159, Aug. 6,1993; Amdt. 23–50, 61 FR 5191, Feb. 9, 1996]
SPINNING
§ 23.221 Spinning.(a) Normal category airplanes. A sin-
gle-engine, normal category airplanemust be able to recover from a one-turn spin or a three-second spin, which-ever takes longer, in not more than oneadditional turn after initiation of thefirst control action for recovery, ordemonstrate compliance with the op-tional spin resistant requirements ofthis section.
(1) The following apply to one turn orthree second spins:
(i) For both the flaps-retracted andflaps-extended conditions, the applica-ble airspeed limit and positive limitmaneuvering load factor must not beexceeded;
(ii) No control forces or char-acteristic encountered during the spinor recovery may adversely affectprompt recovery;
(iii) It must be impossible to obtainunrecoverable spins with any use of theflight or engine power controls eitherat the entry into or during the spin;and
(iv) For the flaps-extended condition,the flaps may be retracted during therecovery but not before rotation hasceased.
(2) At the applicant’s option, the air-plane may be demonstrated to be spinresistant by the following:
(i) During the stall maneuver con-tained in § 23.201, the pitch controlmust be pulled back and held againstthe stop. Then, using ailerons and rud-ders in the proper direction, it must bepossible to maintain wings-level flightwithin 15 degrees of bank and to rollthe airplane from a 30 degree bank inone direction to a 30 degree bank in theother direction;
(ii) Reduce the airplane speed usingpitch control at a rate of approxi-mately one knot per second until thepitch control reaches the stop; then,with the pitch control pulled back andheld against the stop, apply full ruddercontrol in a manner to promote spinentry for a period of seven seconds orthrough a 360 degree heading change,whichever occurs first. If the 360 degreeheading change is reached first, it musthave taken no fewer than four seconds.This maneuver must be performed firstwith the ailerons in the neutral posi-tion, and then with the ailerons de-flected opposite the direction of turn inthe most adverse manner. Power andairplane configuration must be set inaccordance with § 23.201(e) withoutchange during the maneuver. At theend of seven seconds or a 360 degreeheading change, the airplane must re-spond immediately and normally toprimary flight controls applied to re-gain coordinated, unstalled flight with-out reversal of control effect and with-out exceeding the temporary controlforces specified by § 23.143(c); and
(iii) Compliance with §§ 23.201 and23.203 must be demonstrated with theairplane in uncoordinated flight, cor-responding to one ball width displace-ment on a slip-skid indicator, unlessone ball width displacement cannot beobtained with full rudder, in whichcase the demonstration must be withfull rudder applied.
(b) Utility category airplanes. A utilitycategory airplane must meet the re-quirements of paragraph (a) of this sec-tion. In addition, the requirements ofparagraph (c) of this section and§ 23.807(b)(7) must be met if approval forspinning is requested.
(c) Acrobatic category airplanes. An ac-robatic category airplane must meetthe spin requirements of paragraph (a)of this section and § 23.807(b)(6). In addi-tion, the following requirements must
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14 CFR Ch. I (1–1–01 Edition)§ 23.231
be met in each configuration for whichapproval for spinning is requested:
(1) The airplane must recover fromany point in a spin up to and includingsix turns, or any greater number ofturns for which certification is re-quested, in not more than one and one-half additional turns after initiation ofthe first control action for recovery.However, beyond three turns, the spinmay be discontinued if spiral charac-teristics appear.
(2) The applicable airspeed limits andlimit maneuvering load factors mustnot be exceeded. For flaps-extendedconfigurations for which approval is re-quested, the flaps must not be re-tracted during the recovery.
(3) It must be impossible to obtainunrecoverable spins with any use of theflight or engine power controls eitherat the entry into or during the spin.
(4) There must be no characteristicsduring the spin (such as excessive ratesof rotation or extreme oscillatory mo-tion) that might prevent a successfulrecovery due to disorientation or inca-pacitation of the pilot.
[Doc. No. 27807, 61 FR 5191, Feb. 9, 1996]
GROUND AND WATER HANDLINGCHARACTERISTICS
§ 23.231 Longitudinal stability andcontrol.
(a) A landplane may have no uncon-trollable tendency to nose over in anyreasonably expected operating condi-tion, including rebound during landingor takeoff. Wheel brakes must operatesmoothly and may not induce anyundue tendency to nose over.
(b) A seaplane or amphibian may nothave dangerous or uncontrollableporpoising characteristics at any nor-mal operating speed on the water.
§ 23.233 Directional stability and con-trol.
(a) A 90 degree cross-component ofwind velocity, demonstrated to be safefor taxiing, takeoff, and landing mustbe established and must be not lessthan 0.2 VSO.
(b) The airplane must be satisfac-torily controllable in power-off land-ings at normal landing speed, withoutusing brakes or engine power to main-tain a straight path until the speed has
decreased to at least 50 percent of thespeed at touchdown.
(c) The airplane must have adequatedirectional control during taxiing.
(d) Seaplanes must demonstrate sat-isfactory directional stability and con-trol for water operations up to themaximum wind velocity specified inparagraph (a) of this section.
The airplane must be demonstratedto have satisfactory characteristicsand the shock-absorbing mechanismmust not damage the structure of theairplane when the airplane is taxied onthe roughest ground that may reason-ably be expected in normal operationand when takeoffs and landings areperformed on unpaved runways havingthe roughest surface that may reason-ably be expected in normal operation.
[Doc. No. 27807, 61 FR 5192, Feb. 9, 1996]
§ 23.237 Operation on water.
A wave height, demonstrated to besafe for operation, and any necessarywater handling procedures for sea-planes and amphibians must be estab-lished.
[Doc. No. 27807, 61 FR 5192, Feb. 9, 1996]
§ 23.239 Spray characteristics.
Spray may not dangerously obscurethe vision of the pilots or damage thepropellers or other parts of a seaplaneor amphibian at any time during tax-iing, takeoff, and landing.
MISCELLANEOUS FLIGHT REQUIREMENTS
§ 23.251 Vibration and buffeting.
There must be no vibration or buf-feting severe enough to result in struc-tural damage, and each part of the air-plane must be free from excessive vi-bration, under any appropriate speedand power conditions up to VD/MD. Inaddition, there must be no buffeting inany normal flight condition severeenough to interfere with the satisfac-tory control of the airplane or causeexcessive fatigue to the flight crew.
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Federal Aviation Administration, DOT § 23.305
Stall warning buffeting within theselimits is allowable.
[Doc. No. 26269, 58 FR 42159, Aug. 6, 1993]
§ 23.253 High speed characteristics.
If a maximum operating speed VMO/MMO is established under § 23.1505(c),the following speed increase and recov-ery characteristics must be met:
(a) Operating conditions and charac-teristics likely to cause inadvertentspeed increases (including upsets inpitch and roll) must be simulated withthe airplane trimmed at any likelyspeed up to VMO/MMO. These conditionsand characteristics include gust upsets,inadvertent control movements, lowstick force gradients in relation to con-trol friction, passenger movement, lev-eling off from climb, and descent fromMach to airspeed limit altitude.
(b) Allowing for pilot reaction timeafter occurrence of the effective inher-ent or artificial speed warning speci-fied in § 23.1303, it must be shown thatthe airplane can be recovered to a nor-mal attitude and its speed reduced toVMO/MMO, without—
(1) Exceeding VD/MD, the maximumspeed shown under § 23.251, or the struc-tural limitations; or
(2) Buffeting that would impair thepilot’s ability to read the instrumentsor to control the airplane for recovery.
(c) There may be no control reversalabout any axis at any speed up to themaximum speed shown under § 23.251.Any reversal of elevator control forceor tendency of the airplane to pitch,roll, or yaw must be mild and readilycontrollable, using normal pilotingtechniques.
(a) Strength requirements are speci-fied in terms of limit loads (the max-imum loads to be expected in service)and ultimate loads (limit loads multi-plied by prescribed factors of safety).
Unless otherwise provided, prescribedloads are limit loads.
(b) Unless otherwise provided, theair, ground, and water loads must beplaced in equilibrium with inertiaforces, considering each item of massin the airplane. These loads must bedistributed to conservatively approxi-mate or closely represent actual condi-tions. Methods used to determine loadintensities and distribution on canardand tandem wing configurations mustbe validated by flight test measure-ment unless the methods used for de-termining those loading conditions areshown to be reliable or conservative onthe configuration under consideration.
(c) If deflections under load wouldsignificantly change the distribution ofexternal or internal loads, this redis-tribution must be taken into account.
(d) Simplified structural design cri-teria may be used if they result in de-sign loads not less than those pre-scribed in §§ 23.331 through 23.521. Forairplane configurations described inappendix A, § 23.1, the design criteria ofappendix A of this part are an approvedequivalent of §§ 23.321 through 23.459. Ifappendix A of this part is used, the en-tire appendix must be substituted forthe corresponding sections of this part.
The forward structure of a canard ortandem wing configuration must:
(a) Meet all requirements of subpartC and subpart D of this part applicableto a wing; and
(b) Meet all requirements applicableto the function performed by these sur-faces.
[Amdt. 23–42, 56 FR 352, Jan. 3, 1991]
§ 23.303 Factor of safety.Unless otherwise provided, a factor of
safety of 1.5 must be used.
§ 23.305 Strength and deformation.(a) The structure must be able to
support limit loads without detri-mental, permanent deformation. At
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14 CFR Ch. I (1–1–01 Edition)§ 23.307
any load up to limit loads, the defor-mation may not interfere with safe op-eration.
(b) The structure must be able tosupport ultimate loads without failurefor at least three seconds, except localfailures or structural instabilities be-tween limit and ultimate load are ac-ceptable only if the structure can sus-tain the required ultimate load for atleast three seconds. However whenproof of strength is shown by dynamictests simulating actual load condi-tions, the three second limit does notapply.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–45, 58 FR 42160, Aug. 6,1993]
§ 23.307 Proof of structure.(a) Compliance with the strength and
deformation requirements of § 23.305must be shown for each critical loadcondition. Structural analysis may beused only if the structure conforms tothose for which experience has shownthis method to be reliable. In othercases, substantiating load tests mustbe made. Dynamic tests, includingstructural flight tests, are acceptable ifthe design load conditions have beensimulated.
(b) Certain parts of the structuremust be tested as specified in SubpartD of this part.
FLIGHT LOADS
§ 23.321 General.(a) Flight load factors represent the
ratio of the aerodynamic force compo-nent (acting normal to the assumedlongitudinal axis of the airplane) to theweight of the airplane. A positive flightload factor is one in which the aero-dynamic force acts upward, with re-spect to the airplane.
(b) Compliance with the flight loadrequirements of this subpart must beshown—
(1) At each critical altitude withinthe range in which the airplane may beexpected to operate;
(2) At each weight from the designminimum weight to the design max-imum weight; and
(3) For each required altitude andweight, for any practicable distributionof disposable load within the operating
limitations specified in §§ 23.1583through 23.1589.
(c) When significant, the effects ofcompressibility must be taken into ac-count.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–45, 58 FR 42160, Aug. 6,1993]
§ 23.331 Symmetrical flight conditions.(a) The appropriate balancing hori-
zontal tail load must be accounted forin a rational or conservative mannerwhen determining the wing loads andlinear inertia loads corresponding toany of the symmetrical flight condi-tions specified in §§ 23.333 through23.341.
(b) The incremental horizontal tailloads due to maneuvering and gustsmust be reacted by the angular inertiaof the airplane in a rational or conserv-ative manner.
(c) Mutual influence of the aero-dynamic surfaces must be taken intoaccount when determining flight loads.
§ 23.333 Flight envelope.(a) General. Compliance with the
strength requirements of this subpartmust be shown at any combination ofairspeed and load factor on and withinthe boundaries of a flight envelope(similar to the one in paragraph (d) ofthis section) that represents the enve-lope of the flight loading conditionsspecified by the maneuvering and gustcriteria of paragraphs (b) and (c) of thissection respectively.
(b) Maneuvering envelope. Exceptwhere limited by maximum (static) liftcoefficients, the airplane is assumed tobe subjected to symmetrical maneu-vers resulting in the following limitload factors:
(1) The positive maneuvering loadfactor specified in § 23.337 at speeds upto VD;
(2) The negative maneuvering loadfactor specified in § 23.337 at VC; and
(3) Factors varying linearly withspeed from the specified value at VC to0.0 at VD for the normal and commutercategory, and —1.0 at VD for the acro-batic and utility categories.
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Federal Aviation Administration, DOT § 23.335
(c) Gust envelope. (1) The airplane isassumed to be subjected to symmet-rical vertical gusts in level flight. Theresulting limit load factors must cor-respond to the conditions determinedas follows:
(i) Positive (up) and negative (down)gusts of 50 f.p.s. at VC must be consid-ered at altitudes between sea level and20,000 feet. The gust velocity may bereduced linearly from 50 f.p.s. at 20,000feet to 25 f.p.s. at 50,000 feet.
(ii) Positive and negative gusts of 25f.p.s. at VD must be considered at alti-tudes between sea level and 20,000 feet.The gust velocity may be reduced lin-early from 25 f.p.s. at 20,000 feet to 12.5f.p.s. at 50,000 feet.
(iii) In addition, for commuter cat-egory airplanes, positive (up) and nega-tive (down) rough air gusts of 66 f.p.s.at VΒ must be considered at altitudes
between sea level and 20,000 feet. Thegust velocity may be reduced linearlyfrom 66 f.p.s. at 20,000 feet to 38 f.p.s. at50,000 feet.
(2) The following assumptions mustbe made:
(i) The shape of the gust is—
UU s
C
de= −
2
12
25COS
π
Where—
s =Distance penetrated into gust (ft.);C =Mean geometric chord of wing (ft.);
andUde =Derived gust velocity referred to
in subparagraph (1) of this section.
(ii) Gust load factors vary linearlywith speed between VC and VD .
(d) Flight envelope.
[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.
Except as provided in paragraph(a)(4) of this section, the selected de-sign airspeeds are equivalent airspeeds(EAS).
(a) Design cruising speed, VC. For VC
the following apply:
(1) Where W/S′=wing loading at thedesign maximum takeoff weight, Vc (inknots) may not be less than—
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14 CFR Ch. I (1–1–01 Edition)§ 23.337
(2) For values of W/S more than 20,the multiplying factors may be de-creased linearly with W/S to a value of28.6 where W/S =100.
(3) VC need not be more than 0.9 VH atsea level.
(4) At altitudes where an MD is estab-lished, a cruising speed MC limited bycompressibility may be selected.
(b) Design dive speed VD. For VD, thefollowing apply:
(1) VD/MD may not be less than 1.25VC/MC; and
(2) With VC min, the required min-imum design cruising speed, VD (inknots) may not be less than—
(i) 1.40 Vc min (for normal and com-muter category airplanes);
(ii) 1.50 VC min (for utility categoryairplanes); and
(iii) 1.55 VC min (for acrobatic cat-egory airplanes).
(3) For values of W/S more than 20,the multiplying factors in paragraph(b)(2) of this section may be decreasedlinearly with W/S to a value of 1.35where W/S=100.
(4) Compliance with paragraphs (b)(1)and (2) of this section need not beshown if VD/MD is selected so that theminimum speed margin between VC/MC
and VD/MD is the greater of the fol-lowing:
(i) The speed increase resulting when,from the initial condition of stabilizedflight at VC/MC, the airplane is as-sumed to be upset, flown for 20 secondsalong a flight path 7.5° below the ini-tial path, and then pulled up with aload factor of 1.5 (0.5 g. acceleration in-crement). At least 75 percent maximumcontinuous power for reciprocating en-gines, and maximum cruising power forturbines, or, if less, the power requiredfor VC/MC for both kinds of engines,must be assumed until the pullup isinitiated, at which point power reduc-tion and pilot-controlled drag devicesmay be used; and either—
(ii) Mach 0.05 for normal, utility, andacrobatic category airplanes (at alti-tudes where MD is established); or
(iii) Mach 0.07 for commuter categoryairplanes (at altitudes where MD is es-tablished) unless a rational analysis,including the effects of automatic sys-tems, is used to determine a lower mar-gin. If a rational analysis is used, theminimum speed margin must be
enough to provide for atmosphericvariations (such as horizontal gusts),and the penetration of jet streams orcold fronts), instrument errors, air-frame production variations, and mustnot be less than Mach 0.05.
(c) Design maneuvering speed VA. ForVA, the following applies:
(1) VA may not be less than VS√nwhere—
(i) VS is a computed stalling speedwith flaps retracted at the designweight, normally based on the max-imum airplane normal force coeffi-cients, CNA; and
(ii) n is the limit maneuvering loadfactor used in design
(2) The value of VA need not exceedthe value of VC used in design.
(d) Design speed for maximum gust in-tensity, VB. For VB, the following apply:
(1) VB may not be less than the speeddetermined by the intersection of theline representing the maximum posi-tive lift, CNMAX, and the line rep-resenting the rough air gust velocityon the gust V-n diagram, or VS1√ng,
whichever is less, where:(i) ng the positive airplane gust load
factor due to gust, at speed VC (in ac-cordance with § 23.341), and at the par-ticular weight under consideration; and
(ii) VS1 is the stalling speed with theflaps retracted at the particular weightunder consideration.
(a) The positive limit maneuveringload factor n may not be less than—
(1) 2.1+(24,000÷(W+10,000)) for normaland commuter category airplanes,where W=design maximum takeoffweight, except that n need not be morethan 3.8;
(2) 4.4 for utility category airplanes;or
(3) 6.0 for acrobatic category air-planes.
(b) The negative limit maneuveringload factor may not be less than—
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Federal Aviation Administration, DOT § 23.345
(1) 0.4 times the positive load factorfor the normal utility and commutercategories; or
(2) 0.5 times the positive load factorfor the acrobatic category.
(c) Maneuvering load factors lowerthan those specified in this sectionmay be used if the airplane has designfeatures that make it impossible to ex-ceed these values in flight.
§ 23.341 Gust loads factors.(a) Each airplane must be designed to
withstand loads on each lifting surfaceresulting from gusts specified in§ 23.333(c).
(b) The gust load for a canard or tan-dem wing configuration must be com-puted using a rational analysis, or maybe computed in accordance with para-graph (c) of this section, provided thatthe resulting net loads are shown to beconservative with respect to the gustcriteria of § 23.333(c).
(c) In the absence of a more rationalanalysis, the gust load factors must becomputed as follows—
nK U V a
W S
g de= +1498 ( / )
Where—
Kg=0.88µg/5.3+µg=gust alleviation factor;µg=2(W/S)/ρ Cag=airplane mass ratio;Ude=Derived gust velocities referred to in
§ 23.333(c) (f.p.s.);ρ=Density of air (slugs/cu.ft.);W/S =Wing loading (p.s.f.) due to the applica-
ble weight of the airplane in the par-ticular load case.
W/S =Wing loading (p.s.f.);C =Mean geometric chord (ft.);g =Acceleration due to gravity (ft./sec. 2)V =Airplane equivalent speed (knots); anda =Slope of the airplane normal force coeffi-
cient curve CNA per radian if the gustloads are applied to the wings and hori-zontal tail surfaces simultaneously by arational method. The wing lift curveslope CL per radian may be used when thegust load is applied to the wings only andthe horizontal tail gust loads are treatedas a separate condition.
(a) The disposable load combinationsmust include each fuel load in therange from zero fuel to the selectedmaximum fuel load.
(b) If fuel is carried in the wings, themaximum allowable weight of the air-plane without any fuel in the wingtank(s) must be established as ‘‘max-imum zero wing fuel weight,’’ if it isless than the maximum weight.
(c) For commuter category airplanes,a structural reserve fuel condition, notexceeding fuel necessary for 45 minutesof operation at maximum continuouspower, may be selected. If a structuralreserve fuel condition is selected, itmust be used as the minimum fuelweight condition for showing compli-ance with the flight load requirementsprescribed in this part and—
(1) The structure must be designed towithstand a condition of zero fuel inthe wing at limit loads correspondingto:
(i) Ninety percent of the maneu-vering load factors defined in § 23.337,and
(ii) Gust velocities equal to 85 per-cent of the values prescribed in§ 23.333(c).
(2) The fatigue evaluation of thestructure must account for any in-crease in operating stresses resultingfrom the design condition of paragraph(c)(1) of this section.
(3) The flutter, deformation, and vi-bration requirements must also be metwith zero fuel in the wings.
[Doc. No. 27805, 61 FR 5144, Feb. 9, 1996]
§ 23.345 High lift devices.
(a) If flaps or similar high lift devicesare to be used for takeoff, approach orlanding, the airplane, with the flapsfully extended at VF, is assumed to besubjected to symmetrical maneuversand gusts within the range determinedby—
(1) Maneuvering, to a positive limitload factor of 2.0; and
(2) Positive and negative gust of 25feet per second acting normal to theflight path in level flight.
(b) VF must be assumed to be not lessthan 1.4 VS or 1.8 VSF, whichever isgreater, where—
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14 CFR Ch. I (1–1–01 Edition)§ 23.347
(1) VS is the computed stalling speedwith flaps retracted at the designweight; and
(2) VSF is the computed stalling speedwith flaps fully extended at the designweight.
(3) If an automatic flap load limitingdevice is used, the airplane may be de-signed for the critical combinations ofairspeed and flap position allowed bythat device.
(c) In determining external loads onthe airplane as a whole, thrust, slip-stream, and pitching acceleration maybe assumed to be zero.
(d) The flaps, their operating mecha-nism, and their supporting structures,must be designed to withstand the con-ditions prescribed in paragraph (a) ofthis section. In addition, with the flapsfully extended at VF, the following con-ditions, taken separately, must be ac-counted for:
(1) A head-on gust having a velocityof 25 feet per second (EAS), combinedwith propeller slipstream cor-responding to 75 percent of maximumcontinuous power; and
(2) The effects of propeller slipstreamcorresponding to maximum takeoffpower.
[Doc. No. 27805, 61 FR 5144, Feb. 9, 1996]
§ 23.347 Unsymmetrical flight condi-tions.
(a) The airplane is assumed to be sub-jected to the unsymmetrical flight con-ditions of §§ 23.349 and 23.351. Unbal-anced aerodynamic moments about thecenter of gravity must be reacted in arational or conservative manner, con-sidering the principal masses fur-nishing the reacting inertia forces.
(b) Acrobatic category airplanes cer-tified for flick maneuvers (snap roll)must be designed for additional asym-metric loads acting on the wing andthe horizontal tail.
§ 23.349 Rolling conditions.The wing and wing bracing must be
designed for the following loading con-ditions:
(a) Unsymmetrical wing loads appro-priate to the category. Unless the fol-lowing values result in unrealistic
loads, the rolling accelerations may beobtained by modifying the symmet-rical flight conditions in § 23.333(d) asfollows:
(1) For the acrobatic category, inconditions A and F, assume that 100percent of the semispan wing airloadacts on one side of the plane of sym-metry and 60 percent of this load actson the other side.
(2) For normal, utility, and com-muter categories, in Condition A, as-sume that 100 percent of the semispanwing airload acts on one side of the air-plane and 75 percent of this load actson the other side.
(b) The loads resulting from the aile-ron deflections and speeds specified in§ 23.455, in combination with an air-plane load factor of at least two thirdsof the positive maneuvering load factorused for design. Unless the followingvalues result in unrealistic loads, theeffect of aileron displacement on wingtorsion may be accounted for by addingthe following increment to the basicairfoil moment coefficient over the ai-leron portion of the span in the criticalcondition determined in § 23.333(d):
∆cm=—0.01δwhere—
∆cm is the moment coefficient increment; andδ is the down aileron deflection in degrees in
(a) Each engine mount and its sup-porting structure must be designed forthe effects of—
(1) A limit engine torque cor-responding to takeoff power and pro-peller speed acting simultaneouslywith 75 percent of the limit loads fromflight condition A of § 23.333(d);
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Federal Aviation Administration, DOT § 23.367
(2) A limit engine torque cor-responding to maximum continuouspower and propeller speed acting si-multaneously with the limit loads fromflight condition A of § 23.333(d); and
(3) For turbopropeller installations,in addition to the conditions specifiedin paragraphs (a)(1) and (a)(2) of thissection, a limit engine torque cor-responding to takeoff power and pro-peller speed, multiplied by a factor ac-counting for propeller control systemmalfunction, including quick feath-ering, acting simultaneously with lglevel flight loads. In the absence of arational analysis, a factor of 1.6 mustbe used.
(b) For turbine engine installations,the engine mounts and supportingstructure must be designed to with-stand each of the following:
(1) A limit engine torque load im-posed by sudden engine stoppage due tomalfunction or structural failure (suchas compressor jamming).
(2) A limit engine torque load im-posed by the maximum acceleration ofthe engine.
(c) The limit engine torque to be con-sidered under paragraph (a) of this sec-tion must be obtained by multiplyingthe mean torque by a factor of—
(1) 1.25 for turbopropeller installa-tions;
(2) 1.33 for engines with five or morecylinders; and
(3) Two, three, or four, for engineswith four, three, or two cylinders, re-spectively.
[Amdt. 23–26, 45 FR 60171, Sept. 11, 1980, asamended by Amdt. 23–45, 58 FR 42160, Aug. 6,1993]
§ 23.363 Side load on engine mount.
(a) Each engine mount and its sup-porting structure must be designed fora limit load factor in a lateral direc-tion, for the side load on the enginemount, of not less than—
(1) 1.33, or(2) One-third of the limit load factor
for flight condition A.(b) The side load prescribed in para-
graph (a) of this section may be as-sumed to be independent of other flightconditions.
§ 23.365 Pressurized cabin loads.For each pressurized compartment,
the following apply:(a) The airplane structure must be
strong enough to withstand the flightloads combined with pressure differen-tial loads from zero up to the max-imum relief valve setting.
(b) The external pressure distributionin flight, and any stress concentra-tions, must be accounted for.
(c) If landings may be made with thecabin pressurized, landing loads mustbe combined with pressure differentialloads from zero up to the maximum al-lowed during landing.
(d) The airplane structure must bestrong enough to withstand the pres-sure differential loads corresponding tothe maximum relief valve setting mul-tiplied by a factor of 1.33, omittingother loads.
(e) If a pressurized cabin has two ormore compartments separated by bulk-heads or a floor, the primary structuremust be designed for the effects of sud-den release of pressure in any compart-ment with external doors or windows.This condition must be investigated forthe effects of failure of the largestopening in the compartment. The ef-fects of intercompartmental ventingmay be considered.
§ 23.367 Unsymmetrical loads due toengine failure.
(a) Turbopropeller airplanes must bedesigned for the unsymmetrical loadsresulting from the failure of the crit-ical engine including the following con-ditions in combination with a singlemalfunction of the propeller drag lim-iting system, considering the probablepilot corrective action on the flightcontrols:
(1) At speeds between VMC and VD,
the loads resulting from power failurebecause of fuel flow interruption areconsidered to be limit loads.
(2) At speeds between VMC and VC,the loads resulting from the disconnec-tion of the engine compressor from theturbine or from loss of the turbineblades are considered to be ultimateloads.
(3) The time history of the thrustdecay and drag buildup occurring as aresult of the prescribed engine failuresmust be substantiated by test or other
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14 CFR Ch. I (1–1–01 Edition)§ 23.369
data applicable to the particular en-gine-propeller combination.
(4) The timing and magnitude of theprobable pilot corrective action mustbe conservatively estimated, consid-ering the characteristics of the par-ticular engine-propeller-airplane com-bination.
(b) Pilot corrective action may be as-sumed to be initiated at the time max-imum yawing velocity is reached, butnot earlier than 2 seconds after the en-gine failure. The magnitude of the cor-rective action may be based on thelimit pilot forces specified in § 23.397except that lower forces may be as-sumed where it is shown by analysis ortest that these forces can control theyaw and roll resulting from the pre-scribed engine failure conditions.
[Amdt. 23–7, 34 FR 13089, Aug. 13, 1969]
§ 23.369 Rear lift truss.(a) If a rear lift truss is used, it must
be designed to withstand conditions ofreversed airflow at a design speed of—
V = 8.7 √(W/S) + 8.7 (knots), where W/S = wing loading at design maximumtakeoff weight.
(b) Either aerodynamic data for theparticular wing section used, or a valueof CL equalling ¥0.8 with a chordwisedistribution that is triangular betweena peak at the trailing edge and zero atthe leading edge, must be used.
(a) Each engine mount and its sup-porting structure must be designed forthe gyroscopic, inertial, and aero-dynamic loads that result, with the en-gine(s) and propeller(s), if applicable,at maximum continuous r.p.m., undereither:
(1) The conditions prescribed in§ 23.351 and § 23.423; or
(2) All possible combinations of thefollowing—
(i) A yaw velocity of 2.5 radians persecond;
(ii) A pitch velocity of 1.0 radian persecond;
(iii) A normal load factor of 2.5; and
(iv) Maximum continuous thrust.(b) For airplanes approved for aero-
batic maneuvers, each engine mountand its supporting structure must meetthe requirements of paragraph (a) ofthis section and be designed to with-stand the load factors expected duringcombined maximum yaw and pitch ve-locities.
(c) For airplanes certificated in thecommuter category, each enginemount and its supporting structuremust meet the requirements of para-graph (a) of this section and the gustconditions specified in § 23.341 of thispart.
[Doc. No. 27805, 61 FR 5145, Feb. 9, 1996]
§ 23.373 Speed control devices.If speed control devices (such as
spoilers and drag flaps) are incor-porated for use in enroute conditions—
(a) The airplane must be designed forthe symmetrical maneuvers and gustsprescribed in §§ 23.333, 23.337, and 23.341,and the yawing maneuvers and lateralgusts in §§ 23.441 and 23.443, with the de-vice extended at speeds up to theplacard device extended speed; and
(b) If the device has automatic oper-ating or load limiting features, the air-plane must be designed for the maneu-ver and gust conditions prescribed inparagraph (a) of this section at thespeeds and corresponding device posi-tions that the mechanism allows.
[Amdt. 23–7, 34 FR 13089, Aug. 13, 1969]
CONTROL SURFACE AND SYSTEM LOADS
§ 23.391 Control surface loads.The control surface loads specified in
§§ 23.397 through 23.459 are assumed tooccur in the conditions described in§§ 23.331 through 23.351.
§ 23.393 Loads parallel to hinge line.(a) Control surfaces and supporting
hinge brackets must be designed towithstand inertial loads acting parallelto the hinge line.
(b) In the absence of more rationaldata, the inertial loads may be as-sumed to be equal to KW, where—
(1) K = 24 for vertical surfaces;
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Federal Aviation Administration, DOT § 23.399
(2) K = 12 for horizontal surfaces; and(3) W = weight of the movable sur-
faces.
[Doc. No. 27805, 61 FR 5145, Feb. 9, 1996]
§ 23.395 Control system loads.
(a) Each flight control system and itssupporting structure must be designedfor loads corresponding to at least 125percent of the computed hinge mo-ments of the movable control surfacein the conditions prescribed in §§ 23.391through 23.459. In addition, the fol-lowing apply:
(1) The system limit loads need notexceed the higher of the loads that canbe produced by the pilot and automaticdevices operating the controls. How-ever, autopilot forces need not be addedto pilot forces. The system must be de-signed for the maximum effort of thepilot or autopilot, whichever is higher.In addition, if the pilot and the auto-pilot act in opposition, the part of thesystem between them may be designedfor the maximum effort of the one thatimposes the lesser load. Pilot forcesused for design need not exceed themaximum forces prescribed in§ 23.397(b).
(2) The design must, in any case, pro-vide a rugged system for service use,considering jamming, ground gusts,taxiing downwind, control inertia, andfriction. Compliance with this subpara-graph may be shown by designing forloads resulting from application of theminimum forces prescribed in§ 23.397(b).
(b) A 125 percent factor on computedhinge moments must be used to designelevator, aileron, and rudder systems.However, a factor as low as 1.0 may beused if hinge moments are based on ac-curate flight test data, the exact reduc-tion depending upon the accuracy andreliability of the data.
(c) Pilot forces used for design are as-sumed to act at the appropriate controlgrips or pads as they would in flight,and to react at the attachments of thecontrol system to the control surfacehorns.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–7, 34 FR 13089, Aug. 13,1969]
§ 23.397 Limit control forces andtorques.
(a) In the control surface flight load-ing condition, the airloads on movablesurfaces and the corresponding deflec-tions need not exceed those that wouldresult in flight from the application ofany pilot force within the ranges speci-fied in paragraph (b) of this section. Inapplying this criterion, the effects ofcontrol system boost and servo-mecha-nisms, and the effects of tabs must beconsidered. The automatic pilot effortmust be used for design if it alone canproduce higher control surface loadsthan the human pilot.
(b) The limit pilot forces and torquesare as follows:
Control
Maximum forcesor torques fordesign weight,weight equal to
or less than5,000 pounds 1
Minimumforces ortorques 2
Aileron:Stick ................................ 67 lbs ................ 40 lbs.Wheel 3 ............................ 50 D in.-lbs 4 ..... 40 D in.-
1 For design weight (W) more than 5,000 pounds, the speci-fied maximum values must be increased linearly with weightto 1.18 times the specified values at a design weight of12,500 pounds and for commuter category airplanes, thespecified values must be increased linearly with weight to1.35 times the specified values at a design weight of 19,000pounds.
2 If the design of any individual set of control systems orsurfaces 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 thespecified minimum forces or torques, may be used.
3 The critical parts of the aileron control system must alsobe designed for a single tangential force with a limit value of1.25 times the couple force determined from the above cri-teria.
4 D=wheel diameter (inches).5 The unsymmetrical force must be applied at one of the
§ 23.399 Dual control system.(a) Each dual control system must be
designed to withstand the force of thepilots operating in opposition, using in-dividual pilot forces not less than thegreater of—
(1) 0.75 times those obtained under§ 23.395; or
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14 CFR Ch. I (1–1–01 Edition)§ 23.405
(2) The minimum forces specified in§ 23.397(b).
(b) Each dual control system must bedesigned to withstand the force of thepilots applied together, in the same di-rection, using individual pilot forcesnot less than 0.75 times those obtainedunder § 23.395.
[Doc. No. 27805, 61 FR 5145, Feb. 9, 1996]
§ 23.405 Secondary control system.Secondary controls, such as wheel
brakes, spoilers, and tab controls, mustbe designed for the maximum forcesthat a pilot is likely to apply to thosecontrols.
§ 23.407 Trim tab effects.The effects of trim tabs on the con-
trol surface design conditions must beaccounted for only where the surfaceloads are limited by maximum pilot ef-fort. In these cases, the tabs are con-sidered to be deflected in the directionthat would assist the pilot. These de-flections must correspond to the max-imum degree of ‘‘out of trim’’ expectedat the speed for the condition underconsideration.
§ 23.409 Tabs.Control surface tabs must be de-
signed for the most severe combinationof airspeed and tab deflection likely tobe obtained within the flight envelopefor any usable loading condition.
§ 23.415 Ground gust conditions.(a) The control system must be inves-
tigated as follows for control surfaceloads due to ground gusts and taxiingdownwind:
(1) If an investigation of the controlsystem for ground gust loads is not re-quired by paragraph (a)(2) of this sec-tion, but the applicant elects to designa part of the control system of theseloads, these loads need only be carriedfrom control surface horns through thenearest stops or gust locks and theirsupporting structures.
(2) If pilot forces less than the mini-mums specified in § 23.397(b) are usedfor design, the effects of surface loadsdue to ground gusts and taxiing down-wind must be investigated for the en-tire control system according to theformula:
H = K c S qwhere—H = limit hinge moment (ft.-lbs.);c = mean chord of the control surface
aft of the hinge line (ft.);S = area of control surface aft of the
hinge line (sq. ft.);q = dynamic pressure (p.s.f.) based on a
design speed not less than 14.6 √(W/S) + 14.6 (f.p.s.) where W/S = wingloading at design maximum weight,except that the design speed neednot exceed 88 (f.p.s.);
K = limit hinge moment factor forground gusts derived in paragraph(b) of this section. (For aileronsand elevators, a positive value of Kindicates a moment tending to de-press the surface and a negativevalue of K indicates a momenttending to raise the surface).
(b) The limit hinge moment factor Kfor ground gusts must be derived as fol-lows:
Surface K Position of controls
(a) Aileron ......... 0.75 Control column locked lashed inmid-position.
(b) Aileron ......... ±0.50 Ailerons at full throw; + momenton one aileron, ¥ moment onthe other.
(c) Elevator ....... ±0.75 (c) Elevator full up (¥).(d) Elevator ....... ............ (d) Elevator full down (+).(e) Rudder ......... ±0.75 (e) Rudder in neutral.(f) Rudder .......... ............ (f) Rudder at full throw.
(c) At all weights between the emptyweight and the maximum weight de-clared for tie-down stated in the appro-priate manual, any declared tie-downpoints and surrounding structure, con-trol system, surfaces and associatedgust locks, must be designed to with-stand the limit load conditions thatexist when the airplane is tied downand that result from wind speeds of upto 65 knots horizontally from any di-rection.
§ 23.421 Balancing loads.(a) A horizontal surface balancing
load is a load necessary to maintainequilibrium in any specified flight con-dition with no pitching acceleration.
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Federal Aviation Administration, DOT § 23.427
(b) Horizontal balancing surfacesmust be designed for the balancingloads occurring at any point on thelimit maneuvering envelope and in theflap conditions specified in § 23.345.
Each horizontal surface and its sup-porting structure, and the main wingof a canard or tandem wing configura-tion, if that surface has pitch control,must be designed for the maneuveringloads imposed by the following condi-tions:
(a) A sudden movement of the pitch-ing control, at the speed VA, to themaximum aft movement, and the max-imum forward movement, as limited bythe control stops, or pilot effort,whichever is critical.
(b) A sudden aft movement of thepitching control at speeds above VA,followed by a forward movement of thepitching control resulting in the fol-lowing combinations of normal and an-gular acceleration:
load factor used in the design of theairplane; and
(2) V=initial speed in knots.The conditions in this paragraph in-
volve loads corresponding to the loadsthat may occur in a ‘‘checked maneu-ver’’ (a maneuver in which the pitchingcontrol is suddenly displaced in one di-rection and then suddenly moved in theopposite direction). The deflections andtiming of the ‘‘checked maneuver’’must avoid exceeding the limit maneu-vering load factor. The total horizontalsurface load for both nose-up and nose-down pitching conditions is the sum ofthe balancing loads at V and the speci-fied value of the normal load factor n,plus the maneuvering load increment
due to the specified value of the angu-lar acceleration.
§ 23.425 Gust loads.(a) Each horizontal surface, other
than a main wing, must be designed forloads resulting from—
(1) Gust velocities specified in§ 23.333(c) with flaps retracted; and
(2) Positive and negative gusts of 25f.p.s. nominal intensity at VF cor-responding to the flight conditionsspecified in § 23.345(a)(2).
(b) [Reserved](c) When determining the total load
on the horizontal surfaces for the con-ditions specified in paragraph (a) ofthis section, the initial balancing loadsfor steady unaccelerated flight at thepertinent design speeds VF, VC, and VD
must first be determined. The incre-mental load resulting from the gustsmust be added to the initial balancingload to obtain the total load.
(d) In the absence of a more rationalanalysis, the incremental load due tothe gust must be computed as followsonly on airplane configurations withaft-mounted, horizontal surfaces, un-less its use elsewhere is shown to beconservative:
∆ LK U Va S
498ht
g de ht ht= −
1d
d
ε
αwhere—∆Lht=Incremental horizontal tailload (lbs.);Kg=Gust alleviation factor defined in § 23.341;Ude=Derived gust velocity (f.p.s.);V=Airplane equivalent speed (knots);aht=Slope of aft horizontal lift curve (per ra-
dian)Sht=Area of aft horizontal lift surface (ft2);
§ 23.427 Unsymmetrical loads.(a) Horizontal surfaces other than
main wing and their supporting struc-ture must be designed for unsymmet-rical loads arising from yawing and
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14 CFR Ch. I (1–1–01 Edition)§ 23.441
slipstream effects, in combination withthe loads prescribed for the flight con-ditions set forth in §§ 23.421 through23.425.
(b) In the absence of more rationaldata for airplanes that are conven-tional in regard to location of engines,wings, horizontal surfaces other thanmain wing, and fuselage shape:
(1) 100 percent of the maximum load-ing from the symmetrical flight condi-tions may be assumed on the surfaceon one side of the plane of symmetry;and
(2) The following percentage of thatloading must be applied to the oppositeside:
Percent=100¥10 (n¥1), where n is the spec-ified positive maneuvering load factor, butthis value may not be more than 80 percent.
(c) For airplanes that are not conven-tional (such as airplanes with hori-zontal surfaces other than main winghaving appreciable dihedral or sup-ported by the vertical tail surfaces) thesurfaces and supporting structuresmust be designed for combined verticaland horizontal surface loads resultingfrom each prescribed flight conditiontaken separately.
§ 23.441 Maneuvering loads.(a) At speeds up to VA, the vertical
surfaces must be designed to withstand
the following conditions. In computingthe loads, the yawing velocity may beassumed to be zero:
(1) With the airplane in unacceler-ated flight at zero yaw, it is assumedthat the rudder control is suddenly dis-placed to the maximum deflection, aslimited by the control stops or by limitpilot forces.
(2) With the rudder deflected as speci-fied in paragraph (a)(1) of this section,it is assumed that the airplane yaws tothe overswing sideslip angle. In lieu ofa rational analysis, an overswing angleequal to 1.5 times the static sideslipangle of paragraph (a)(3) of this sectionmay be assumed.
(3) A yaw angle of 15 degrees with therudder control maintained in the neu-tral position (except as limited by pilotstrength).
(b) For commuter category airplanes,the loads imposed by the following ad-ditional maneuver must be substan-tiated at speeds from VA to VD/MD.When computing the tail loads—
(1) The airplane must be yawed to thelargest attainable steady state sideslipangle, with the rudder at maximum de-flection caused by any one of the fol-lowing:
(i) Control surface stops;(ii) Maximum available booster ef-
fort;(iii) Maximum pilot rudder force as
shown below:
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211
Federal Aviation Administration, DOT § 23.441
(2) The rudder must be suddenly dis-placed from the maximum deflection tothe neutral position.
(c) The yaw angles specified in para-graph (a)(3) of this section may be re-duced if the yaw angle chosen for a par-ticular speed cannot be exceeded in—
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212
14 CFR Ch. I (1–1–01 Edition)§ 23.443
(1) Steady slip conditions;(2) Uncoordinated rolls from steep
§ 23.443 Gust loads.(a) Vertical surfaces must be de-
signed to withstand, in unacceleratedflight at speed VC, lateral gusts of thevalues prescribed for VC in § 23.333(c).
(b) In addition, for commuter cat-egory airplanes, the airplane is as-sumed to encounter derived gusts nor-mal to the plane of symmetry while inunaccelerated flight at VB, VC, VD, andVF. The derived gusts and airplanespeeds corresponding to these condi-tions, as determined by §§ 23.341 and23.345, must be investigated. The shapeof the gust must be as specified in§ 23.333(c)(2)(i).
(c) In the absence of a more rationalanalysis, the gust load must be com-puted as follows:
LK U V a S
vtgt de vt vt=
498Where—Lvt=Vertical surface loads (lbs.);
kgtgt
gt
=+
=0 88
5 3
.
.
µµ
gust alleviation factor;
µρ
gt
t vt vt vt
W
c g a S
K
llateral= =
2 2
massratio;
Ude=Derived gust velocity (f.p.s.);ρ=Air density (slugs/cu.ft.);W=the applicable weight of the air-
plane in the particular load case(lbs.);
Svt=Area of vertical surface (ft.2);c̆≤t=Mean geometric chord of vertical
surface (ft.);avt=Lift curve slope of vertical surface
(per radian);K=Radius of gyration in yaw (ft.);lvt=Distance from airplane c.g. to lift
(a) If outboard fins or winglets are in-cluded on the horizontal surfaces orwings, the horizontal surfaces or wingsmust be designed for their maximumload in combination with loads inducedby the fins or winglets and moments orforces exerted on the horizontal sur-faces or wings by the fins or winglets.
(b) If outboard fins or winglets ex-tend above and below the horizontalsurface, the critical vertical surfaceloading (the load per unit area as de-termined under §§ 23.441 and 23.443)must be applied to—
(1) The part of the vertical surfacesabove the horizontal surface with 80percent of that loading applied to thepart below the horizontal surface; and
(2) The part of the vertical surfacesbelow the horizontal surface with 80percent of that loading applied to thepart above the horizontal surface.
(c) The end plate effects of outboardfins or winglets must be taken into ac-count in applying the yawing condi-tions of §§ 23.441 and 23.443 to thevertical surfaces in paragraph (b) ofthis section.
(d) When rational methods are usedfor computing loads, the maneuveringloads of § 23.441 on the vertical surfacesand the one-g horizontal surface load,including induced loads on the hori-zontal surface and moments or forcesexerted on the horizontal surfaces bythe vertical surfaces, must be appliedsimultaneously for the structural load-ing condition.
§ 23.459 Special devices.The loading for special devices using
aerodynamic surfaces (such as slotsand spoilers) must be determined fromtest data.
GROUND LOADS
§ 23.471 General.The limit ground loads specified in
this subpart are considered to be exter-nal loads and inertia forces that actupon an airplane structure. In eachspecified ground load condition, the ex-ternal reactions must be placed inequilibrium with the linear and angu-lar inertia forces in a rational or con-servative manner.
§ 23.473 Ground load conditions andassumptions.
(a) The ground load requirements ofthis subpart must be complied with atthe design maximum weight exceptthat §§ 23.479, 23.481, and 23.483 may becomplied with at a design landingweight (the highest weight for landingconditions at the maximum descent ve-locity) allowed under paragraphs (b)and (c) of this section.
(b) The design landing weight may beas low as—
(1) 95 percent of the maximum weightif the minimum fuel capacity is enoughfor at least one-half hour of operationat maximum continuous power plus acapacity equal to a fuel weight which
is the difference between the designmaximum weight and the design land-ing weight; or
(2) The design maximum weight lessthe weight of 25 percent of the totalfuel capacity.
(c) The design landing weight of amultiengine airplane may be less thanthat allowed under paragraph (b) ofthis section if—
(1) The airplane meets the one-en-gine-inoperative climb requirements of§ 23.67(b)(1) or (c); and
(2) Compliance is shown with the fueljettisoning system requirements of§ 23.1001.
(d) The selected limit vertical inertiaload factor at the center of gravity ofthe airplane for the ground load condi-tions prescribed in this subpart maynot be less than that which would beobtained when landing with a descentvelocity (V), in feet per second, equalto 4.4 (W/S)1⁄4, except that this velocityneed not be more than 10 feet per sec-ond and may not be less than sevenfeet per second.
(e) Wing lift not exceeding two-thirdsof the weight of the airplane may beassumed to exist throughout the land-ing impact and to act through the cen-ter of gravity. The ground reactionload factor may be equal to the inertiaload factor minus the ratio of theabove assumed wing lift to the airplaneweight.
(f) If energy absorption tests aremade to determine the limit load fac-tor corresponding to the required limitdescent velocities, these tests must bemade under § 23.723(a).
(g) No inertia load factor used for de-sign purposes may be less than 2.67, normay the limit ground reaction load fac-tor be less than 2.0 at design maximumweight, unless these lower values willnot be exceeded in taxiing at speeds upto takeoff speed over terrain as roughas that expected in service.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–7, 34 FR 13090, Aug. 13,1969; Amdt. 23–28, 47 FR 13315, Mar. 29, 1982;Amdt. 23–45, 58 FR 42160, Aug. 6, 1993; Amdt.23–48, 61 FR 5147, Feb. 9, 1996]
§ 23.477 Landing gear arrangement.Sections 23.479 through 23.483, or the
conditions in appendix C, apply to air-planes with conventional arrangements
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214
14 CFR Ch. I (1–1–01 Edition)§ 23.479
of main and nose gear, or main and tailgear.
§ 23.479 Level landing conditions.
(a) For a level landing, the airplaneis assumed to be in the following atti-tudes:
(1) For airplanes with tail wheels, anormal level flight attitude.
(2) For airplanes with nose wheels,attitudes in which—
(i) The nose and main wheels contactthe ground simultaneously; and
(ii) The main wheels contact theground and the nose wheel is just clearof the ground.
The attitude used in paragraph (a)(2)(i)of this section may be used in the anal-ysis required under paragraph (a)(2)(ii)of this section.
(b) When investigating landing condi-tions, the drag components simulatingthe forces required to accelerate thetires and wheels up to the landingspeed (spin-up) must be properly com-bined with the corresponding instanta-neous vertical ground reactions, andthe forward-acting horizontal loads re-sulting from rapid reduction of thespin-up drag loads (spring-back) mustbe combined with vertical ground reac-tions at the instant of the peak for-ward load, assuming wing lift and atire-sliding coefficient of friction of 0.8.However, the drag loads may not beless than 25 percent of the maximumvertical ground reactions (neglectingwing lift).
(c) In the absence of specific tests ora more rational analysis for deter-mining the wheel spin-up and spring-back loads for landing conditions, themethod set forth in appendix D of thispart must be used. If appendix D of thispart is used, the drag components usedfor design must not be less than thosegiven by appendix C of this part.
(d) For airplanes with tip tanks orlarge overhung masses (such as turbo-propeller or jet engines) supported bythe wing, the tip tanks and the struc-ture supporting the tanks or overhungmasses must be designed for the effectsof dynamic responses under the levellanding conditions of either paragraph(a)(1) or (a)(2)(ii) of this section. Inevaluating the effects of dynamic re-
sponse, an airplane lift equal to theweight of the airplane may be assumed.
(a) For a tail down landing, the air-plane is assumed to be in the followingattitudes:
(1) For airplanes with tail wheels, anattitude in which the main and tailwheels contact the ground simulta-neously.
(2) For airplanes with nose wheels, astalling attitude, or the maximumangle allowing ground clearance byeach part of the airplane, whichever isless.
(b) For airplanes with either tail ornose wheels, ground reactions are as-sumed to be vertical, with the wheelsup to speed before the maximumvertical load is attained.
§ 23.483 One-wheel landing conditions.
For the one-wheel landing condition,the airplane is assumed to be in thelevel attitude and to contact theground on one side of the main landinggear. In this attitude, the ground reac-tions must be the same as those ob-tained on that side under § 23.479.
§ 23.485 Side load conditions.
(a) For the side load condition, theairplane is assumed to be in a level at-titude with only the main wheels con-tacting the ground and with the shockabsorbers and tires in their static posi-tions.
(b) The limit vertical load factormust be 1.33, with the vertical groundreaction divided equally between themain wheels.
(c) The limit side inertia factor mustbe 0.83, with the side ground reactiondivided between the main wheels sothat—
(1) 0.5 (W) is acting inboard on oneside; and
(2) 0.33 (W) is acting outboard on theother side.
(d) The side loads prescribed in para-graph (c) of this section are assumed tobe applied at the ground contact point
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215
Federal Aviation Administration, DOT § 23.499
and the drag loads may be assumed tobe zero.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–45, 58 FR 42160, Aug. 6,1993]
§ 23.493 Braked roll conditions.
Under braked roll conditions, withthe shock absorbers and tires in theirstatic positions, the following apply:
(a) The limit vertical load factormust be 1.33.
(b) The attitudes and ground con-tacts must be those described in § 23.479for level landings.
(c) A drag reaction equal to thevertical reaction at the wheel multi-plied by a coefficient of friction of 0.8must be applied at the ground contactpoint of each wheel with brakes, exceptthat the drag reaction need not exceedthe maximum value based on limitingbrake torque.
§ 23.497 Supplementary conditions fortail wheels.
In determining the ground loads onthe tail wheel and affected supportingstructures, the following apply:
(a) For the obstruction load, thelimit ground reaction obtained in thetail down landing condition is assumedto act up and aft through the axle at 45degrees. The shock absorber and tiremay be assumed to be in their staticpositions.
(b) For the side load, a limit verticalground reaction equal to the staticload on the tail wheel, in combinationwith a side component of equal mag-nitude, is assumed. In addition—
(1) If a swivel is used, the tail wheelis assumed to be swiveled 90 degrees tothe airplane longitudinal axis with theresultant ground load passing throughthe axle;
(2) If a lock, steering device, or shim-my damper is used, the tail wheel isalso assumed to be in the trailing posi-tion with the side load acting at theground contact point; and
(3) The shock absorber and tire areassumed to be in their static positions.
(c) If a tail wheel, bumper, or an en-ergy absorption device is provided toshow compliance with § 23.925(b), thefollowing apply:
(1) Suitable design loads must be es-tablished for the tail wheel, bumper, orenergy absorption device; and
(2) The supporting structure of thetail wheel, bumper, or energy absorp-tion device must be designed to with-stand the loads established in para-graph (c)(1) of this section.
In determining the ground loads onnose wheels and affected supportingstructures, and assuming that theshock absorbers and tires are in theirstatic positions, the following condi-tions must be met:
(a) For aft loads, the limit force com-ponents at the axle must be—
(1) A vertical component of 2.25 timesthe static load on the wheel; and
(2) A drag component of 0.8 times thevertical load.
(b) For forward loads, the limit forcecomponents at the axle must be—
(1) A vertical component of 2.25 timesthe static load on the wheel; and
(2) A forward component of 0.4 timesthe vertical load.
(c) For side loads, the limit forcecomponents at ground contact mustbe—
(1) A vertical component of 2.25 timesthe static load on the wheel; and
(2) A side component of 0.7 times thevertical load.
(d) For airplanes with a steerablenose wheel that is controlled by hy-draulic or other power, at design take-off weight with the nose wheel in anysteerable position, the application of1.33 times the full steering torque com-bined with a vertical reaction equal to1.33 times the maximum static reactionon the nose gear must be assumed.However, if a torque limiting device isinstalled, the steering torque can be re-duced to the maximum value allowedby that device.
(e) For airplanes with a steerablenose wheel that has a direct mechan-ical connection to the rudder pedals,the mechanism must be designed towithstand the steering torque for the
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In determining ground loads for ski-planes, and assuming that the airplaneis resting on the ground with one mainski frozen at rest and the other skisfree to slide, a limit side force equal to0.036 times the design maximum weightmust be applied near the tail assembly,with a factor of safety of 1.
[Amdt. 23–7, 34 FR 13090, Aug. 13, 1969]
§ 23.507 Jacking loads.
(a) The airplane must be designed forthe loads developed when the aircraftis supported on jacks at the designmaximum weight assuming the fol-lowing load factors for landing gearjacking points at a three-point attitudeand for primary flight structure jack-ing points in the level attitude:
(1) Vertical-load factor of 1.35 timesthe static reactions.
(2) Fore, aft, and lateral load factorsof 0.4 times the vertical static reac-tions.
(b) The horizontal loads at the jackpoints must be reacted by inertiaforces so as to result in no change inthe direction of the resultant loads atthe jack points.
(c) The horizontal loads must be con-sidered in all combinations with thevertical load.
[Amdt. 23–14, 38 FR 31821, Nov. 19, 1973]
§ 23.509 Towing loads.
The towing loads of this section mustbe applied to the design of tow fittings
and their immediate attaching struc-ture.
(a) The towing loads specified inparagraph (d) of this section must beconsidered separately. These loadsmust be applied at the towing fittingsand must act parallel to the ground. Inaddition:
(1) A vertical load factor equal to 1.0must be considered acting at the centerof gravity; and
(2) The shock struts and tires mustbe in there static positions.
(b) For towing points not on thelanding gear but near the plane of sym-metry of the airplane, the drag andside tow load components specified forthe auxiliary gear apply. For towingpoints located outboard of the maingear, the drag and side tow load compo-nents specified for the main gear apply.Where the specified angle of swivelcannot be reached, the maximum ob-tainable angle must be used.
(c) The towing loads specified inparagraph (d) of this section must bereacted as follows:
(1) The side component of the towingload at the main gear must be reactedby a side force at the static ground lineof the wheel to which the load is ap-plied.
(2) The towing loads at the auxiliarygear and the drag components of thetowing loads at the main gear must bereacted as follows:
(i) A reaction with a maximum valueequal to the vertical reaction must beapplied at the axle of the wheel towhich the load is applied. Enough air-plane inertia to achieve equilibriummust be applied.
(ii) The loads must be reacted by air-plane inertia.
(d) The prescribed towing loads are asfollows, where W is the design max-imum weight:
Tow point PositionLoad
Magnitude No. Direction
Main gear ............................... .......................................................... 0.225W 1234
Forward, parallel to drag axis.Forward, at 30° to drag axis.Aft, parallel to drag axis.Aft, at 30° to drag axis.
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217
Federal Aviation Administration, DOT § 23.525
Tow point PositionLoad
Magnitude No. Direction
Swiveled 45° from forward .............. 0.15W 910
Forward, in plane of wheel.Aft, in plane of wheel.
Swiveled 45° from aft ...................... 0.15W 1112
Forward, in plane of wheel.Aft, in plane of wheel.
[Amdt. 23–14, 38 FR 31821, Nov. 19, 1973]
§ 23.511 Ground load; unsymmetricalloads on multiple-wheel units.
(a) Pivoting loads. The airplane is as-sumed to pivot about on side of themain gear with—
(1) The brakes on the pivoting unitlocked; and
(2) Loads corresponding to a limitvertical load factor of 1, and coefficientof friction of 0.8 applied to the maingear and its supporting structure.
(b) Unequal tire loads. The loads es-tablished under §§ 23.471 through 23.483must be applied in turn, in a 60/40 per-cent distribution, to the dual wheelsand tires in each dual wheel landinggear unit.
(c) Deflated tire loads. For the deflatedtire condition—
(1) 60 percent of the loads establishedunder §§ 23.471 through 23.483 must beapplied in turn to each wheel in a land-ing gear unit; and
(2) 60 percent of the limit drag andside loads, and 100 percent of the limitvertical load established under §§ 23.485and 23.493 or lesser vertical load ob-tained under paragraph (c)(1) of thissection, must be applied in turn toeach wheel in the dual wheel landinggear unit.
[Amdt. 23–7, 34 FR 13090, Aug. 13, 1969]
WATER LOADS
§ 23.521 Water load conditions.
(a) The structure of seaplanes andamphibians must be designed for waterloads developed during takeoff andlanding with the seaplane in any atti-tude likely to occur in normal oper-ation at appropriate forward and sink-ing velocities under the most severesea conditions likely to be encoun-tered.
(b) Unless the applicant makes a ra-tional analysis of the water loads,§§ 23.523 through 23.537 apply.
§ 23.523 Design weights and center ofgravity positions.
(a) Design weights. The water load re-quirements must be met at each oper-ating weight up to the design landingweight except that, for the takeoff con-dition prescribed in § 23.531, the designwater takeoff weight (the maximumweight for water taxi and takeoff run)must be used.
(b) Center of gravity positions. Thecritical centers of gravity within thelimits for which certification is re-quested must be considered to reachmaximum design loads for each part ofthe seaplane structure.
[Doc. No. 26269, 58 FR 42160, Aug. 6, 1993]
§ 23.525 Application of loads.
(a) Unless otherwise prescribed, theseaplane as a whole is assumed to besubjected to the loads corresponding tothe load factors specified in § 23.527.
(b) In applying the loads resultingfrom the load factors prescribed in§ 23.527, the loads may be distributedover the hull or main float bottom (inorder to avoid excessive local shearloads and bending moments at the lo-cation of water load application) usingpressures not less than those pre-scribed in § 23.533(c).
(c) For twin float seaplanes, eachfloat must be treated as an equivalenthull on a fictitious seaplane with aweight equal to one-half the weight ofthe twin float seaplane.
(d) Except in the takeoff condition of§ 23.531, the aerodynamic lift on the
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218
14 CFR Ch. I (1–1–01 Edition)§ 23.527
seaplane during the impact is assumedto be 2⁄3 of the weight of the seaplane.
(b) The following values are used:(1) nw=water reaction load factor
(that is, the water reaction divided byseaplane weight).
(2) C1=empirical seaplane operationsfactor equal to 0.012 (except that thisfactor may not be less than that nec-essary to obtain the minimum value ofstep load factor of 2.33).
(3) VSO=seaplane stalling speed inknots with flaps extended in the appro-priate landing position and with noslipstream effect.
(4) β=Angle of dead rise at the longi-tudinal station at which the load fac-tor is being determined in accordancewith figure 1 of appendix I of this part.
(5) W=seaplane landing weight inpounds.
(6) K1=empirical hull station weigh-ing factor, in accordance with figure 2of appendix I of this part.
(7) rx=ratio of distance, measuredparallel to hull reference axis, from thecenter of gravity of the seaplane to thehull longitudinal station at which theload factor is being computed to the ra-dius of gyration in pitch of the sea-plane, the hull reference axis being astraight line, in the plane of sym-metry, tangential to the keel at themain step.
(c) For a twin float seaplane, becauseof the effect of flexibility of the attach-ment of the floats to the seaplane, the
factor K1 may be reduced at the bowand stern to 0.8 of the value shown infigure 2 of appendix I of this part. Thisreduction applies only to the design ofthe carrythrough and seaplane struc-ture.
(a) Symmetrical step, bow, and sternlanding. For symmetrical step, bow,and stern landings, the limit water re-action load factors are those computedunder § 23.527. In addition—
(1) For symmetrical step landings,the resultant water load must be ap-plied at the keel, through the center ofgravity, and must be directed per-pendicularly to the keel line;
(2) For symmetrical bow landings,the resultant water load must be ap-plied at the keel, one-fifth of the longi-tudinal distance from the bow to thestep, and must be directed perpendicu-larly to the keel line; and
(3) For symmetrical stern landings,the resultant water load must be ap-plied at the keel, at a point 85 percentof the longitudinal distance from thestep to the stern post, and must be di-rected perpendicularly to the keel line.
(b) Unsymmetrical landing for hull andsingle float seaplanes. Unsymmetricalstep, bow, and stern landing conditionsmust be investigated. In addition—
(1) The loading for each conditionconsists of an upward component and aside component equal, respectively, to0.75 and 0.25 tan β times the resultantload in the corresponding symmetricallanding condition; and
(2) The point of application and di-rection of the upward component of theload is the same as that in the sym-metrical condition, and the point of ap-plication of the side component is atthe same longitudinal station as theupward component but is directed in-ward perpendicularly to the plane ofsymmetry at a point midway betweenthe keel and chine lines.
(c) Unsymmetrical landing; twin floatseaplanes. The unsymmetrical loadingconsists of an upward load at the stepof each float of 0.75 and a side load of0.25 tan β at one float times the steplanding load reached under § 23.527. The
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219
Federal Aviation Administration, DOT § 23.533
side load is directed inboard, per-pendicularly to the plane of symmetrymidway between the keel and chinelines of the float, at the same longitu-dinal station as the upward load.
[Doc. No. 26269, 58 FR 42161, Aug. 6, 1993]
§ 23.531 Hull and main float takeoffcondition.
For the wing and its attachment tothe hull or main float—
(a) The aerodynamic wing lift is as-sumed to be zero; and
(b) A downward inertia load, cor-responding to a load factor computedfrom the following formula, must beapplied:
tor equal to 0.004;VS1=seaplane stalling speed (knots) at
the design takeoff weight with theflaps extended in the appropriatetakeoff position;
β=angle of dead rise at the main step(degrees); and
W=design water takeoff weight inpounds.
[Doc. No. 26269, 58 FR 42161, Aug. 6, 1993]
§ 23.533 Hull and main float bottompressures.
(a) General. The hull and main floatstructure, including frames and bulk-heads, stringers, and bottom plating,must be designed under this section.
(b) Local pressures. For the design ofthe bottom plating and stringers andtheir attachments to the supportingstructure, the following pressure dis-tributions must be applied:
(1) For an unflared bottom, the pres-sure at the chine is 0.75 times the pres-sure at the keel, and the pressures be-tween the keel and chine vary linearly,in accordance with figure 3 of appendixI of this part. The pressure at the keel(p.s.i.) is computed as follows:
PC K V
TanK
2 2 S1
k
=2
βwhere—Pk=pressure (p.s.i.) at the keel;C2=0.00213;K2=hull station weighing factor, in ac-
cordance with figure 2 of appendix Iof this part;
VS1=seaplane stalling speed (knots) atthe design water takeoff weight withflaps extended in the appropriatetakeoff position; and
βK=angle of dead rise at keel, in ac-cordance with figure 1 of appendix Iof this part.
(2) For a flared bottom, the pressureat the beginning of the flare is thesame as that for an unflared bottom,and the pressure between the chine andthe beginning of the flare varies lin-early, in accordance with figure 3 of ap-pendix I of this part. The pressure dis-tribution is the same as that prescribedin paragraph (b)(1) of this section foran unflared bottom except that thepressure at the chine is computed asfollows:
PC K V
Tanch
3 2 S1=2
βwhere—Pch=pressure (p.s.i.) at the chine;C3=0.0016;K2=hull station weighing factor, in ac-
cordance with figure 2 of appendix Iof this part;
VS1=seaplane stalling speed (knots) atthe design water takeoff weight withflaps extended in the appropriatetakeoff position; and
β=angle of dead rise at appropriate sta-tion.
The area over which these pressuresare applied must simulate pressures oc-curring during high localized impactson the hull or float, but need not ex-tend over an area that would inducecritical stresses in the frames or in theoverall structure.
(c) Distributed pressures. For the de-sign of the frames, keel, and chinestructure, the following pressure dis-tributions apply:
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14 CFR Ch. I (1–1–01 Edition)§ 23.535
(1) Symmetrical pressures are com-puted as follows:
PC K V
Tan4 2 SO=
2
βwhere—P=pressure (p.s.i.);C4=0.078 C1 (with C1 computed under
§ 23.527);K2=hull station weighing factor, deter-
mined in accordance with figure 2 ofappendix I of this part;
VS0=seaplane stalling speed (knots)with landing flaps extended in the ap-propriate position and with no slip-stream effect; and
β=angle of dead rise at appropriate sta-tion.(2) The unsymmetrical pressure dis-
tribution consists of the pressures pre-scribed in paragraph (c)(1) of this sec-tion on one side of the hull or mainfloat centerline and one-half of thatpressure on the other side of the hull ormain float centerline, in accordancewith figure 3 of appendix I of this part.
(3) These pressures are uniform andmust be applied simultaneously overthe entire hull or main float bottom.The loads obtained must be carriedinto the sidewall structure of the hullproper, but need not be transmitted ina fore and aft direction as shear andbending loads.
§ 23.535 Auxiliary float loads.(a) General. Auxiliary floats and their
attachments and supporting structuresmust be designed for the conditionsprescribed in this section. In the casesspecified in paragraphs (b) through (e)of this section, the prescribed waterloads may be distributed over the floatbottom to avoid excessive local loads,using bottom pressures not less thanthose prescribed in paragraph (g) ofthis section.
(b) Step loading. The resultant waterload must be applied in the plane ofsymmetry of the float at a point three-fourths of the distance from the bow tothe step and must be perpendicular tothe keel. The resultant limit load iscomputed as follows, except that thevalue of L need not exceed three times
the weight of the displaced water whenthe float is completely submerged:
with landing flaps extended in the ap-propriate position and with no slip-stream effect;
W=seaplane design landing weight inpounds;
βs=angle of dead rise at a station 3⁄4 ofthe distance from the bow to thestep, but need not be less than 15 de-grees; and
ry=ratio of the lateral distance betweenthe center of gravity and the plane ofsymmetry of the float to the radiusof gyration in roll.
(c) Bow loading. The resultant limitload must be applied in the plane ofsymmetry of the float at a point one-fourth of the distance from the bow tothe step and must be perpendicular tothe tangent to the keel line at thatpoint. The magnitude of the resultantload is that specified in paragraph (b)of this section.
(d) Unsymmetrical step loading. The re-sultant water load consists of a compo-nent equal to 0.75 times the load speci-fied in paragraph (a) of this section anda side component equal to 0.025 tan βtimes the load specified in paragraph(b) of this section. The side load mustbe applied perpendicularly to the planeof symmetry of the float at a pointmidway between the keel and thechine.
(e) Unsymmetrical bow loading. The re-sultant water load consists of a compo-nent equal to 0.75 times the load speci-fied in paragraph (b) of this section anda side component equal to 0.25 tan βtimes the load specified in paragraph(c) of this section. The side load mustbe applied perpendicularly to the planeof symmetry at a point midway be-tween the keel and the chine.
(f) Immersed float condition. The re-sultant load must be applied at thecentroid of the cross section of the
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Federal Aviation Administration, DOT § 23.561
float at a point one-third of the dis-tance from the bow to the step. Thelimit load components are as follows:
vertical PgV
aftC PV (KV )
2
sideC PV (KV )
2
X SO2
Y SO2
23
23
=
=
=
where—P=mass density of water (slugs/ft.3)V=volume of float (ft.3);CX=coefficient of drag force, equal to
0.133;Cy=coefficient of side force, equal to
0.106;K=0.8, except that lower values may be
used if it is shown that the floats areincapable of submerging at a speed of0.8 Vso in normal operations;
Vso=seaplane stalling speed (knots)with landing flaps extended in the ap-propriate position and with no slip-stream effect; and
g=acceleration due to gravity (ft/sec2).(g) Float bottom pressures. The float
bottom pressures must be establishedunder § 23.533, except that the value ofK2 in the formulae may be taken as 1.0.The angle of dead rise to be used in de-termining the float bottom pressures isset forth in paragraph (b) of this sec-tion.
§ 23.537 Seawing loads.Seawing design loads must be based
on applicable test data.
[Doc. No. 26269, 58 FR 42163, Aug. 6, 1993]
EMERGENCY LANDING CONDITIONS
§ 23.561 General.(a) The airplane, although it may be
damaged in emergency landing condi-tions, must be designed as prescribed inthis section to protect each occupantunder those conditions.
(b) The structure must be designed togive each occupant every reasonablechance of escaping serious injurywhen—
(1) Proper use is made of the seats,safety belts, and shoulder harnessesprovided for in the design;
(2) The occupant experiences thestatic inertia loads corresponding tothe following ultimate load factors—
(i) Upward, 3.0g for normal, utility,and commuter category airplanes, or4.5g for acrobatic category airplanes;
(ii) Forward, 9.0g;(iii) Sideward, 1.5g; and(iv) Downward, 6.0g when certifi-
cation to the emergency exit provi-sions of § 23.807(d)(4) is requested; and
(3) The items of mass within thecabin, that could injure an occupant,experience the static inertia loads cor-responding to the following ultimateload factors—
(i) Upward, 3.0g;(ii) Forward, 18.0g; and(iii) Sideward, 4.5g.(c) Each airplane with retractable
landing gear must be designed to pro-tect each occupant in a landing—
(1) With the wheels retracted;(2) With moderate descent velocity;
and(3) Assuming, in the absence of a
more rational analysis—(i) A downward ultimate inertia force
of 3 g; and(ii) A coefficient of friction of 0.5 at
the ground.(d) If it is not established that a
turnover is unlikely during an emer-gency landing, the structure must bedesigned to protect the occupants in acomplete turnover as follows:
(1) The likelihood of a turnover maybe shown by an analysis assuming thefollowing conditions—
(i) The most adverse combination ofweight and center of gravity position;
(ii) Longitudinal load factor of 9.0g;(iii) Vertical load factor of 1.0g; and(iv) For airplanes with tricycle land-
ing gear, the nose wheel strut failedwith the nose contacting the ground.
(i) Maximum weight;(ii) Most forward center of gravity
position;(iii) Longitudinal load factor of 9.0g;(iv) Vertical load factor of 1.0g; and(v) For airplanes with tricycle land-
ing gear, the nose wheel strut failedwith the nose contacting the ground.
(2) For determining the loads to beapplied to the inverted airplane after a
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14 CFR Ch. I (1–1–01 Edition)§ 23.562
turnover, an upward ultimate inertiaload factor of 3.0g and a coefficient offriction with the ground of 0.5 must beused.
(e) Except as provided in § 23.787(c),the supporting structure must be de-signed to restrain, under loads up tothose specified in paragraph (b)(3) ofthis section, each item of mass thatcould injure an occupant if it cameloose in a minor crash landing.
(a) Each seat/restraint system for usein a normal, utility, or acrobatic cat-egory airplane must be designed to pro-tect each occupant during an emer-gency landing when—
(1) Proper use is made of seats, safetybelts, and shoulder harnesses providedfor in the design; and
(2) The occupant is exposed to theloads resulting from the conditionsprescribed in this section.
(b) Except for those seat/restraintsystems that are required to meetparagraph (d) of this section, each seat/restraint system for crew or passengeroccupancy in a normal, utility, or acro-batic category airplane, must success-fully complete dynamic tests or bedemonstrated by rational analysis sup-ported by dynamic tests, in accordancewith each of the following conditions.These tests must be conducted with anoccupant simulated by ananthropomorphic test dummy (ATD)defined by 49 CFR Part 572, Subpart B,or an FAA-approved equivalent, with anominal weight of 170 pounds and seat-ed in the normal upright position.
(1) For the first test, the change invelocity may not be less than 31 feetper second. The seat/restraint systemmust be oriented in its nominal posi-tion with respect to the airplane andwith the horizontal plane of the air-plane pitched up 60 degrees, with noyaw, relative to the impact vector. Forseat/restraint systems to be installedin the first row of the airplane, peakdeceleration must occur in not more
than 0.05 seconds after impact andmust reach a minimum of 19g. For allother seat/restraint systems, peak de-celeration must occur in not more than0.06 seconds after impact and mustreach a minimum of 15g.
(2) For the second test, the change invelocity may not be less than 42 feetper second. The seat/restraint systemmust be oriented in its nominal posi-tion with respect to the airplane andwith the vertical plane of the airplaneyawed 10 degrees, with no pitch, rel-ative to the impact vector in a direc-tion that results in the greatest loadon the shoulder harness. For seat/re-straint systems to be installed in thefirst row of the airplane, peak decelera-tion must occur in not more than 0.05seconds after impact and must reach aminimum of 26g. For all other seat/re-straint systems, peak decelerationmust occur in not more than 0.06 sec-onds after impact and must reach aminimum of 21g.
(3) To account for floor warpage, thefloor rails or attachment devices usedto attach the seat/restraint system tothe airframe structure must be pre-loaded to misalign with respect to eachother by at least 10 degrees vertically(i.e., pitch out of parallel) and one ofthe rails or attachment devices mustbe preloaded to misalign by 10 degreesin roll prior to conducting the test de-fined by paragraph (b)(2) of this sec-tion.
(c) Compliance with the following re-quirements must be shown during thedynamic tests conducted in accordancewith paragraph (b) of this section:
(1) The seat/restraint system mustrestrain the ATD although seat/re-straint system components may experi-ence deformation, elongation, displace-ment, or crushing intended as part ofthe design.
(2) The attachment between the seat/restraint system and the test fixturemust remain intact, although the seatstructure may have deformed.
(3) Each shoulder harness strap mustremain on the ATD’s shoulder duringthe impact.
(4) The safety belt must remain onthe ATD’s pelvis during the impact.
(5) The results of the dynamic testsmust show that the occupant is pro-tected from serious head injury.
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Federal Aviation Administration, DOT § 23.571
(i) When contact with adjacent seats,structure, or other items in the cabincan occur, protection must be providedso that the head impact does not ex-ceed a head injury criteria (HIC) of1,000.
(ii) The value of HIC is defined as—
HIC t t1
t ta(t)dt2 1
2 1 t
t2.5
Max1
2
= −( )−( )
∫
Where: t1 is the initial integration time, ex-pressed in seconds, t2 is the final integra-tion time, expressed in seconds, (t2¥ t1) isthe time duration of the major head im-pact, expressed in seconds, and a(t) is theresultant deceleration at the center ofgravity of the head form expressed as amultiple of g (units of gravity).
(iii) Compliance with the HIC limitmust be demonstrated by measuringthe head impact during dynamic test-ing as prescribed in paragraphs (b)(1)and (b)(2) of this section or by a sepa-rate showing of compliance with thehead injury criteria using test or anal-ysis procedures.
(6) Loads in individual shoulder har-ness straps may not exceed 1,750pounds. If dual straps are used for re-taining the upper torso, the total straploads may not exceed 2,000 pounds.
(7) The compression load measuredbetween the pelvis and the lumbarspine of the ATD may not exceed 1,500pounds.
(d) For all single-engine airplaneswith a VSO of more than 61 knots atmaximum weight, and those multien-gine airplanes of 6,000 pounds or lessmaximum weight with a VSO of morethan 61 knots at maximum weight thatdo not comply with § 23.67(a)(1);
(1) The ultimate load factors of§ 23.561(b) must be increased by multi-plying the load factors by the square ofthe ratio of the increased stall speed to61 knots. The increased ultimate loadfactors need not exceed the valuesreached at a VS0 of 79 knots. The up-ward ultimate load factor for acrobaticcategory airplanes need not exceed5.0g.
(2) The seat/restraint system test re-quired by paragraph (b)(1) of this sec-tion must be conducted in accordancewith the following criteria:
(i) The change in velocity may not beless than 31 feet per second.
(ii)(A) The peak deceleration (gp) of19g and 15g must be increased and mul-tiplied by the square of the ratio of theincreased stall speed to 61 knots:
gp=19.0 (VS0/61)2 or gp=15.0 (VS0/61)2
(B) The peak deceleration need notexceed the value reached at a VS0 of 79knots.
(iii) The peak deceleration mustoccur in not more than time (tr), whichmust be computed as follows:
t31
32.2 g
.96
gr
p p
= ( ) =
where—gp=The peak deceleration calculated in ac-
cordance with paragraph (d)(2)(ii) of thissection
tr=The rise time (in seconds) to the peak de-celeration.
(e) An alternate approach thatachieves an equivalent, or greater,level of occupant protection to that re-quired by this section may be used ifsubstantiated on a rational basis.
[Amdt. 23–36, 53 FR 30812, Aug. 15, 1988, asamended by Amdt. 23–44, 58 FR 38639, July 19,1993; Amdt. 23–50, 61 FR 5192, Feb. 9, 1996]
FATIGUE EVALUATION
§ 23.571 Metallic pressurized cabinstructures.
For normal, utility, and acrobaticcategory airplanes, the strength, detaildesign, and fabrication of the metallicstructure of the pressure cabin must beevaluated under one of the following:
(a) A fatigue strength investigationin which the structure is shown bytests, or by analysis supported by testevidence, to be able to withstand therepeated loads of variable magnitudeexpected in service; or
(b) A fail safe strength investigation,in which it is shown by analysis, tests,or both that catastrophic failure of thestructure is not probable after fatiguefailure, or obvious partial failure, of aprincipal structural element, and thatthe remaining structures are able towithstand a static ultimate load factorof 75 percent of the limit load factor atVC, considering the combined effects ofnormal operating pressures, expected
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14 CFR Ch. I (1–1–01 Edition)§ 23.572
external aerodynamic pressures, andflight loads. These loads must be mul-tiplied by a factor of 1.15 unless the dy-namic effects of failure under staticload are otherwise considered.
(c) The damage tolerance evaluationof § 23.573(b).
(a) For normal, utility, and acrobaticcategory airplanes, the strength, detaildesign, and fabrication of those partsof the airframe structure whose failurewould be catastrophic must be evalu-ated under one of the following unlessit is shown that the structure, oper-ating stress level, materials and ex-pected uses are comparable, from a fa-tigue standpoint, to a similar designthat has had extensive satisfactoryservice experience:
(1) A fatigue strength investigationin which the structure is shown bytests, or by analysis supported by testevidence, to be able to withstand therepeated loads of variable magnitudeexpected in service; or
(2) A fail-safe strength investigationin which it is shown by analysis, tests,or both, that catastrophic failure ofthe structure is not probable after fa-tigue failure, or obvious partial failure,of a principal structural element, andthat the remaining structure is able towithstand a static ultimate load factorof 75 percent of the critical limit loadfactor at Vc. These loads must be mul-tiplied by a factor of 1.15 unless the dy-namic effects of failure under staticload are otherwise considered.
(3) The damage tolerance evaluationof § 23.573(b).
(b) Each evaluation required by thissection must—
(1) Include typical loading spectra(e.g. taxi, ground-air-ground cycles,maneuver, gust);
(2) Account for any significant effectsdue to the mutual influence of aero-dynamic surfaces; and
(3) Consider any significant effectsfrom propeller slipstream loading, andbuffet from vortex impingements.
§ 23.573 Damage tolerance and fatigueevaluation of structure.
(a) Composite airframe structure. Com-posite airframe structure must be eval-uated under this paragraph instead of§§ 23.571 and 23.572. The applicant mustevaluate the composite airframe struc-ture, the failure of which would resultin catastrophic loss of the airplane, ineach wing (including canards, tandemwings, and winglets), empennage, theircarrythrough and attaching structure,moveable control surfaces and their at-taching structure fuselage, and pres-sure cabin using the damage-tolerancecriteria prescribed in paragraphs (a)(1)through (a)(4) of this section unlessshown to be impractical. If the appli-cant establishes that damage-tolerancecriteria is impractical for a particularstructure, the structure must be evalu-ated in accordance with paragraphs(a)(1) and (a)(6) of this section. Wherebonded joints are used, the structuremust also be evaluated in accordancewith paragraph (a)(5) of this section.The effects of material variability andenvironmental conditions on thestrength and durability properties ofthe composite materials must be ac-counted for in the evaluations requiredby this section.
(1) It must be demonstrated by tests,or by analysis supported by tests, thatthe structure is capable of carrying ul-timate load with damage up to thethreshold of detectability consideringthe inspection procedures employed.
(2) The growth rate or no-growth ofdamage that may occur from fatigue,corrosion, manufacturing flaws or im-pact damage, under repeated loads ex-pected in service, must be establishedby tests or analysis supported by tests.
(3) The structure must be shown byresidual strength tests, or analysis sup-ported by residual strength tests, to beable to withstand critical limit flightloads, considered as ultimate loads,
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Federal Aviation Administration, DOT § 23.574
with the extent of detectable damageconsistent with the results of the dam-age tolerance evaluations. For pressur-ized cabins, the following loads must bewithstood:
(i) Critical limit flight loads with thecombined effects of normal operatingpressure and expected external aero-dynamic pressures.
(ii) The expected external aero-dynamic pressures in 1g flight com-bined with a cabin differential pressureequal to 1.1 times the normal operatingdifferential pressure without any otherload.
(4) The damage growth, between ini-tial detectability and the value se-lected for residual strength demonstra-tions, factored to obtain inspection in-tervals, must allow development of aninspection program suitable for appli-cation by operation and maintenancepersonnel.
(5) For any bonded joint, the failureof which would result in catastrophicloss of the airplane, the limit load ca-pacity must be substantiated by one ofthe following methods—
(i) The maximum disbonds of eachbonded joint consistent with the capa-bility to withstand the loads in para-graph (a)(3) of this section must be de-termined by analysis, tests, or both.Disbonds of each bonded joint greaterthan this must be prevented by designfeatures; or
(ii) Proof testing must be conductedon each production article that willapply the critical limit design load toeach critical bonded joint; or
(iii) Repeatable and reliable non-de-structive inspection techniques mustbe established that ensure the strengthof each joint.
(6) Structural components for whichthe damage tolerance method is shownto be impractical must be shown bycomponent fatigue tests, or analysissupported by tests, to be able to with-stand the repeated loads of variablemagnitude expected in service. Suffi-cient component, subcomponent, ele-ment, or coupon tests must be done toestablish the fatigue scatter factor andthe environmental effects. Damage upto the threshold of detectability andultimate load residual strength capa-bility must be considered in the dem-onstration.
(b) Metallic airframe structure. If theapplicant elects to use § 23.571(a)(3) or§ 23.572(a)(3), then the damage toleranceevaluation must include a determina-tion of the probable locations andmodes of damage due to fatigue, corro-sion, or accidental damage. The deter-mination must be by analysis sup-ported by test evidence and, if avail-able, service experience. Damage atmultiple sites due to fatigue must beincluded where the design is such thatthis type of damage can be expected tooccur. The evaluation must incor-porate repeated load and static anal-yses supported by test evidence. Theextent of damage for residual strengthevaluation at any time within theoperational life of the airplane must beconsistent with the initial detect-ability and subsequent growth underrepeated loads. The residual strengthevaluation must show that the remain-ing structure is able to withstand crit-ical limit flight loads, considered as ul-timate, with the extent of detectabledamage consistent with the results ofthe damage tolerance evaluations. Forpressurized cabins, the following loadmust be withstood:
(1) The normal operating differentialpressure combined with the expectedexternal aerodynamic pressures appliedsimultaneously with the flight loadingconditions specified in this part, and
(2) The expected external aero-dynamic pressures in 1g flight com-bined with a cabin differential pressureequal to 1.1 times the normal operatingdifferential pressure without any otherload.
[Doc. No. 26269, 58 FR 42163, Aug. 6, 1993; 58FR 51970, Oct. 5, 1993, as amended by Amdt.23–48, 61 FR 5147, Feb. 9, 1996]
For commuter category airplanes—(a) Metallic damage tolerance. An eval-
uation of the strength, detail design,and fabrication must show that cata-strophic failure due to fatigue, corro-sion, defects, or damage will be avoidedthroughout the operational life of theairplane. This evaluation must be con-ducted in accordance with the provi-sions of § 23.573, except as specified inparagraph (b) of this section, for each
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14 CFR Ch. I (1–1–01 Edition)§ 23.575
part of the structure that could con-tribute to a catastrophic failure.
(b) Fatigue (safe-life) evaluation. Com-pliance with the damage tolerance re-quirements of paragraph (a) of this sec-tion is not required if the applicant es-tablishes that the application of thoserequirements is impractical for a par-ticular structure. This structure mustbe shown, by analysis supported by testevidence, to be able to withstand therepeated loads of variable magnitudeexpected during its service life withoutdetectable cracks. Appropriate safe-lifescatter factors must be applied.
[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]
§ 23.575 Inspections and other proce-dures.
Each inspection or other procedure,based on an evaluation required by§§ 23.571, 23.572, 23.573 or 23.574, must beestablished to prevent catastrophicfailure and must be included in theLimitations Section of the Instructionsfor Continued Airworthiness requiredby § 23.1529.
[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]
Subpart D—Design andConstruction
§ 23.601 General.The suitability of each questionable
design detail and part having an impor-tant bearing on safety in operations,must be established by tests.
§ 23.603 Materials and workmanship.(a) The suitability and durability of
materials used for parts, the failure ofwhich could adversely affect safety,must—
(1) Be established by experience ortests;
(2) Meet approved specifications thatensure their having the strength andother properties assumed in the designdata; and
(3) Take into account the effects ofenvironmental conditions, such as tem-perature and humidity, expected inservice.
(a) The methods of fabrication usedmust produce consistently sound struc-tures. If a fabrication process (such asgluing, spot welding, or heat-treating)requires close control to reach this ob-jective, the process must be performedunder an approved process specifica-tion.
(b) Each new aircraft fabricationmethod must be substantiated by atest program.
(a) Each removable fastener must in-corporate two retaining devices if theloss of such fastener would precludecontinued safe flight and landing.
(b) Fasteners and their locking de-vices must not be adversely affected bythe environmental conditions associ-ated with the particular installation.
(c) No self-locking nut may be usedon any bolt subject to rotation in oper-ation unless a non-friction locking de-vice is used in addition to the self-lock-ing device.
[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]
§ 23.609 Protection of structure.
Each part of the structure must—(a) Be suitably protected against de-
terioration or loss of strength in serv-ice due to any cause, including—
(1) Weathering;(2) Corrosion; and(3) Abrasion; and(b) Have adequate provisions for ven-
tilation and drainage.
§ 23.611 Accessibility provisions.
For each part that requires mainte-nance, inspection, or other servicing,appropriate means must be incor-porated into the aircraft design toallow such servicing to be accom-plished.
[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]
§ 23.613 Material strength propertiesand design values.
(a) Material strength properties mustbe based on enough tests of material
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Federal Aviation Administration, DOT § 23.621
meeting specifications to establish de-sign values on a statistical basis.
(b) Design values must be chosen tominimize the probability of structuralfailure due to material variability. Ex-cept as provided in paragraph (e) ofthis section, compliance with thisparagraph must be shown by selectingdesign values that ensure materialstrength with the following prob-ability:
(1) Where applied loads are eventu-ally distributed through a single mem-ber within an assembly, the failure ofwhich would result in loss of structuralintegrity of the component; 99 percentprobability with 95 percent confidence.
(2) For redundant structure, in whichthe failure of individual elementswould result in applied loads beingsafely distributed to other load car-rying members; 90 percent probabilitywith 95 percent confidence.
(c) The effects of temperature on al-lowable stresses used for design in anessential component or structure mustbe considered where thermal effects aresignificant under normal operatingconditions.
(d) The design of the structure mustminimize the probability of cata-strophic fatigue failure, particularly atpoints of stress concentration.
(e) Design values greater than theguaranteed minimums required by thissection may be used where only guar-anteed minimum values are normallyallowed if a ‘‘premium selection’’ ofthe material is made in which a speci-men of each individual item is testedbefore use to determine that the actualstrength properties of that particularitem will equal or exceed those used indesign.
§ 23.619 Special factors.The factor of safety prescribed in
§ 23.303 must be multiplied by the high-est pertinent special factors of safetyprescribed in §§ 23.621 through 23.625 foreach part of the structure whosestrength is—
(a) Uncertain;(b) Likely to deteriorate in service
before normal replacement; or
(c) Subject to appreciable variabilitybecause of uncertainties in manufac-turing processes or inspection methods.
[Amdt. 23–7, 34 FR 13091, Aug. 13, 1969]
§ 23.621 Casting factors.(a) General. The factors, tests, and in-
spections specified in paragraphs (b)through (d) of this section must be ap-plied in addition to those necessary toestablish foundry quality control. Theinspections must meet approved speci-fications. Paragraphs (c) and (d) of thissection apply to any structural cast-ings except castings that are pressuretested as parts of hydraulic or otherfluid systems and do not support struc-tural loads.
(b) Bearing stresses and surfaces. Thecasting factors specified in paragraphs(c) and (d) of this section—
(1) Need not exceed 1.25 with respectto bearing stresses regardless of themethod of inspection used; and
(2) Need not be used with respect tothe bearing surfaces of a part whosebearing factor is larger than the appli-cable casting factor.
(c) Critical castings. For each castingwhose failure would preclude continuedsafe flight and landing of the airplaneor result in serious injury to occu-pants, the following apply:
(1) Each critical casting must ei-ther—
(i) Have a casting factor of not lessthan 1.25 and receive 100 percent in-spection by visual, radiographic, andeither magnetic particle, penetrant orother approved equivalent non-destruc-tive inspection method; or
(ii) Have a casting factor of not lessthan 2.0 and receive 100 percent visualinspection and 100 percent approvednon-destructive inspection. When anapproved quality control procedure isestablished and an acceptable statis-tical analysis supports reduction, non-destructive inspection may be reducedfrom 100 percent, and applied on a sam-pling basis.
(2) For each critical casting with acasting factor less than 1.50, three sam-ple castings must be static tested andshown to meet—
(i) The strength requirements of§ 23.305 at an ultimate load cor-responding to a casting factor of 1.25;and
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14 CFR Ch. I (1–1–01 Edition)§ 23.623
(ii) The deformation requirements of§ 23.305 at a load of 1.15 times the limitload.
(3) Examples of these castings arestructural attachment fittings, parts offlight control systems, control surfacehinges and balance weight attach-ments, seat, berth, safety belt, and fueland oil tank supports and attachments,and cabin pressure valves.
(d) Non-critical castings. For eachcasting other than those specified inparagraph (c) or (e) of this section, thefollowing apply:
(1) Except as provided in paragraphs(d)(2) and (3) of this section, the castingfactors and corresponding inspectionsmust meet the following table:
Casting factor Inspection
2.0 or more .................... 100 percent visual.Less than 2.0 but more
than 1.5.100 percent visual, and magnetic
particle or penetrant or equiva-lent nondestructive inspectionmethods.
1.25 through 1.50 .......... 100 percent visual, magnetic par-ticle or penetrant, and radio-graphic, or approved equivalentnondestructive inspection meth-ods.
(2) The percentage of castings in-spected by nonvisual methods may bereduced below that specified in sub-paragraph (d)(1) of this section when anapproved quality control procedure isestablished.
(3) For castings procured to a speci-fication that guarantees the mechan-ical properties of the material in thecasting and provides for demonstrationof these properties by test of couponscut from the castings on a samplingbasis—
(i) A casting factor of 1.0 may beused; and
(ii) The castings must be inspected asprovided in paragraph (d)(1) of this sec-tion for casting factors of ‘‘1.25 through1.50’’ and tested under paragraph (c)(2)of this section.
(e) Non-structural castings. Castingsused for non-structural purposes do notrequire evaluation, testing or close in-spection.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–45, 58 FR 42164, Aug. 6,1993]
§ 23.623 Bearing factors.(a) Each part that has clearance (free
fit), and that is subject to pounding orvibration, must have a bearing factorlarge enough to provide for the effectsof normal relative motion.
(b) For control surface hinges andcontrol system joints, compliance withthe factors prescribed in §§ 23.657 and23.693, respectively, meets paragraph(a) of this section.
[Amdt. 23–7, 34 FR 13091, Aug. 13, 1969]
§ 23.625 Fitting factors.For each fitting (a part or terminal
used to join one structural member toanother), the following apply:
(a) For each fitting whose strength isnot proven by limit and ultimate loadtests in which actual stress conditionsare simulated in the fitting and sur-rounding structures, a fitting factor ofat least 1.15 must be applied to eachpart of—
(1) The fitting;(2) The means of attachment; and(3) The bearing on the joined mem-
bers.(b) No fitting factor need be used for
joint designs based on comprehensivetest data (such as continuous joints inmetal plating, welded joints, and scarfjoints in wood).
(c) For each integral fitting, the partmust be treated as a fitting up to thepoint at which the section propertiesbecome typical of the member.
(d) For each seat, berth, safety belt,and harness, its attachment to thestructure must be shown, by analysis,tests, or both, to be able to withstandthe inertia forces prescribed in § 23.561multiplied by a fitting factor of 1.33.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–7, 34 FR 13091, Aug. 13,1969]
§ 23.627 Fatigue strength.The structure must be designed, as
far as practicable, to avoid points ofstress concentration where variablestresses above the fatigue limit arelikely to occur in normal service.
§ 23.629 Flutter.(a) It must be shown by the methods
of paragraph (b) and either paragraph
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Federal Aviation Administration, DOT § 23.629
(c) or (d) of this section, that the air-plane is free from flutter, control re-versal, and divergence for any condi-tion of operation within the limit V-nenvelope and at all speeds up to thespeed specified for the selected method.In addition—
(1) Adequate tolerances must be es-tablished for quantities which affectflutter, including speed, damping, massbalance, and control system stiffness;and
(2) The natural frequencies of mainstructural components must be deter-mined by vibration tests or other ap-proved methods.
(b) Flight flutter tests must be madeto show that the airplane is free fromflutter, control reversal and divergenceand to show that—
(1) Proper and adequate attempts toinduce flutter have been made withinthe speed range up to VD;
(2) The vibratory response of thestructure during the test indicatesfreedom from flutter;
(3) A proper margin of damping existsat VD; and
(4) There is no large and rapid reduc-tion in damping as VD is approached.
(c) Any rational analysis used to pre-dict freedom from flutter, control re-versal and divergence must cover allspeeds up to 1.2 VD.
(d) Compliance with the rigidity andmass balance criteria (pages 4–12), inAirframe and Equipment EngineeringReport No. 45 (as corrected) ‘‘Sim-plified Flutter Prevention Criteria’’(published by the Federal Aviation Ad-ministration) may be accomplished toshow that the airplane is free fromflutter, control reversal, or divergenceif—
(1) VD/MD for the airplane is less than260 knots (EAS) and less than Mach 0.5,
(2) The wing and aileron flutter pre-vention criteria, as represented by thewing torsional stiffness and aileronbalance criteria, are limited in use toairplanes without large mass con-centrations (such as engines, floats, orfuel tanks in outer wing panels) alongthe wing span, and
(3) The airplane—(i) Does not have a T-tail or other un-
conventional tail configurations;
(ii) Does not have unusual mass dis-tributions or other unconventional de-sign features that affect the applica-bility of the criteria, and
(iii) Has fixed-fin and fixed-stabilizersurfaces.
(e) For turbopropeller-powered air-planes, the dynamic evaluation mustinclude—
(1) Whirl mode degree of freedomwhich takes into account the stabilityof the plane of rotation of the propellerand significant elastic, inertial, andaerodynamic forces, and
(f) Freedom from flutter, control re-versal, and divergence up to VD/MD
must be shown as follows:(1) For airplanes that meet the cri-
teria of paragraphs (d)(1) through (d)(3)of this section, after the failure, mal-function, or disconnection of any singleelement in any tab control system.
(2) For airplanes other than those de-scribed in paragraph (f)(1) of this sec-tion, after the failure, malfunction, ordisconnection of any single element inthe primary flight control system, anytab control system, or any flutterdamper.
(g) For airplanes showing compliancewith the fail-safe criteria of §§ 23.571and 23.572, the airplane must be shownby analysis to be free from flutter upto VD/MD after fatigue failure, or obvi-ous partial failure, of a principal struc-tural element.
(h) For airplanes showing compliancewith the damage tolerance criteria of§ 23.573, the airplane must be shown byanalysis to be free from flutter up toVD/MD with the extent of damage forwhich residual strength is dem-onstrated.
(i) For modifications to the type de-sign that could affect the flutter char-acteristics, compliance with paragraph(a) of this section must be shown, ex-cept that analysis based on previouslyapproved data may be used alone toshow freedom from flutter, control re-versal and divergence, for all speeds up
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The strength of stressed-skin wingsmust be proven by load tests or bycombined structural analysis and loadtests.
CONTROL SURFACES
§ 23.651 Proof of strength.
(a) Limit load tests of control sur-faces are required. These tests must in-clude the horn or fitting to which thecontrol system is attached.
(b) In structural analyses, riggingloads due to wire bracing must be ac-counted for in a rational or conserv-ative manner.
§ 23.655 Installation.
(a) Movable surfaces must be in-stalled so that there is no interferencebetween any surfaces, their bracing, oradjacent fixed structure, when one sur-face is held in its most critical clear-ance positions and the others are oper-ated through their full movement.
(b) If an adjustable stabilizer is used,it must have stops that will limit itsrange of travel to that allowing safeflight and landing.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–45, 58 FR 42164, Aug. 6,1993]
§ 23.657 Hinges.
(a) Control surface hinges, exceptball and roller bearing hinges, musthave a factor of safety of not less than6.67 with respect to the ultimate bear-ing strength of the softest materialused as a bearing.
(b) For ball or roller bearing hinges,the approved rating of the bearing maynot be exceeded.
The supporting structure and the at-tachment of concentrated mass bal-ance weights used on control surfacesmust be designed for—
(a) 24 g normal to the plane of thecontrol surface;
(b) 12 g fore and aft; and(c) 12 g parallel to the hinge line.
CONTROL SYSTEMS
§ 23.671 General.
(a) Each control must operate easily,smoothly, and positively enough toallow proper performance of its func-tions.
(b) Controls must be arranged andidentified to provide for convenience inoperation and to prevent the possi-bility of confusion and subsequent in-advertent operation.
§ 23.672 Stability augmentation andautomatic and power-operated sys-tems.
If the functioning of stability aug-mentation or other automatic orpower-operated systems is necessary toshow compliance with the flight char-acteristics requirements of this part,such systems must comply with § 23.671and the following:
(a) A warning, which is clearly dis-tinguishable to the pilot under ex-pected flight conditions without re-quiring the pilot’s attention, must beprovided for any failure in the stabilityaugmentation system or in any otherautomatic or power-operated systemthat could result in an unsafe condi-tion if the pilot was not aware of thefailure. Warning systems must not ac-tivate the control system.
(b) The design of the stability aug-mentation system or of any other auto-matic or power-operated system mustpermit initial counteraction of failureswithout requiring exceptional pilotskill or strength, by either the deacti-vation of the system or a failed portionthereof, or by overriding the failure bymovement of the flight controls in thenormal sense.
(c) It must be shown that, after anysingle failure of the stability aug-mentation system or any other auto-matic or power-operated system—
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Federal Aviation Administration, DOT § 23.679
(1) The airplane is safely controllablewhen the failure or malfunction occursat any speed or altitude within the ap-proved operating limitations that iscritical for the type of failure beingconsidered;
(2) The controllability and maneuver-ability requirements of this part aremet within a practical operationalflight envelope (for example, speed, al-titude, normal acceleration, and air-plane configuration) that is describedin the Airplane Flight Manual (AFM);and
(3) The trim, stability, and stall char-acteristics are not impaired below alevel needed to permit continued safeflight and landing.
[Doc. No. 26269, 58 FR 42164, Aug. 6, 1993]
§ 23.673 Primary flight controls.Primary flight controls are those
used by the pilot for the immediatecontrol of pitch, roll, and yaw.
stops that positively limit the range ofmotion of each movable aerodynamicsurface controlled by the system.
(b) Each stop must be located so thatwear, slackness, or takeup adjustmentswill not adversely affect the controlcharacteristics of the airplane becauseof a change in the range of surfacetravel.
(c) Each stop must be able to with-stand any loads corresponding to thedesign conditions for the control sys-tem.
[Amdt. 23–17, 41 FR 55464, Dec. 20, 1976]
§ 23.677 Trim systems.(a) Proper precautions must be taken
to prevent inadvertent, improper, orabrupt trim tab operation. There mustbe means near the trim control to indi-cate to the pilot the direction of trimcontrol movement relative to airplanemotion. In addition, there must bemeans to indicate to the pilot the posi-tion of the trim device with respect toboth the range of adjustment and, inthe case of lateral and directionaltrim, the neutral position. This means
must be visible to the pilot and mustbe located and designed to prevent con-fusion. The pitch trim indicator mustbe clearly marked with a position orrange within which it has been dem-onstrated that take-off is safe for allcenter of gravity positions and eachflap position approved for takeoff.
(b) Trimming devices must be de-signed so that, when any one con-necting or transmitting element in theprimary flight control system fails,adequate control for safe flight andlanding is available with—
(1) For single-engine airplanes, thelongitudinal trimming devices; or
(2) For multiengine airplanes, thelongitudinal and directional trimmingdevices.
(c) Tab controls must be irreversibleunless the tab is properly balanced andhas no unsafe flutter characteristics.Irreversible tab systems must haveadequate rigidity and reliability in theportion of the system from the tab tothe attachment of the irreversible unitto the airplane structure.
(d) It must be demonstrated that theairplane is safely controllable and thatthe pilot can perform all maneuversand operations necessary to effect asafe landing following any probablepowered trim system runaway thatreasonably might be expected in serv-ice, allowing for appropriate timedelay after pilot recognition of thetrim system runaway. The demonstra-tion must be conducted at critical air-plane weights and center of gravity po-sitions.
§ 23.679 Control system locks.If there is a device to lock the con-
trol system on the ground or water:(a) There must be a means to—(1) Give unmistakable warning to the
pilot when lock is engaged; or(2) Automatically disengage the de-
vice when the pilot operates the pri-mary flight controls in a normal man-ner.
(b) The device must be installed tolimit the operation of the airplane sothat, when the device is engaged, the
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14 CFR Ch. I (1–1–01 Edition)§ 23.681
pilot receives unmistakable warning atthe start of the takeoff.
(c) The device must have a means topreclude the possibility of it becominginadvertently engaged in flight.
[Doc. No. 26269, 58 FR 42164, Aug. 6, 1993]
§ 23.681 Limit load static tests.
(a) Compliance with the limit loadrequirements of this part must beshown by tests in which—
(1) The direction of the test loadsproduces the most severe loading in thecontrol system; and
(2) Each fitting, pulley, and bracketused in attaching the system to themain structure is included.
(b) Compliance must be shown (byanalyses or individual load tests) withthe special factor requirements forcontrol system joints subject to angu-lar motion.
§ 23.683 Operation tests.
(a) It must be shown by operationtests that, when the controls are oper-ated from the pilot compartment withthe system loaded as prescribed inparagraph (b) of this section, the sys-tem is free from—
(1) Jamming;(2) Excessive friction; and(3) Excessive deflection.(b) The prescribed test loads are—(1) For the entire system, loads cor-
responding to the limit airloads on theappropriate surface, or the limit pilotforces in § 23.397(b), whichever are less;and
(2) For secondary controls, loads notless than those corresponding to themaximum pilot effort establishedunder § 23.405.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–7, 34 FR 13091, Aug. 13,1969]
§ 23.685 Control system details.
(a) Each detail of each control sys-tem must be designed and installed toprevent jamming, chafing, and inter-ference from cargo, passengers, looseobjects, or the freezing of moisture.
(b) There must be means in the cock-pit to prevent the entry of foreign ob-jects into places where they would jamthe system.
(c) There must be means to preventthe slapping of cables or tubes againstother parts.
(d) Each element of the flight controlsystem must have design features, ormust be distinctively and permanentlymarked, to minimize the possibility ofincorrect assembly that could result inmalfunctioning of the control system.
The reliability of any spring deviceused in the control system must be es-tablished by tests simulating serviceconditions unless failure of the springwill not cause flutter or unsafe flightcharacteristics.
§ 23.689 Cable systems.
(a) Each cable, cable fitting, turn-buckle, splice, and pulley used mustmeet approved specifications. In addi-tion—
(1) No cable smaller than 1⁄8 inch di-ameter may be used in primary controlsystems;
(2) Each cable system must be de-signed so that there will be no haz-ardous change in cable tensionthroughout the range of travel underoperating conditions and temperaturevariations; and
(3) There must be means for visualinspection at each fairlead, pulley, ter-minal, and turnbuckle.
(b) Each kind and size of pulley mustcorrespond to the cable with which it isused. Each pulley must have closelyfitted guards to prevent the cablesfrom being misplaced or fouled, evenwhen slack. Each pulley must lie in theplane passing through the cable so thatthe cable does not rub against the pul-ley flange.
(c) Fairleads must be installed sothat they do not cause a change incable direction of more than three de-grees.
(d) Clevis pins subject to load or mo-tion and retained only by cotter pinsmay not be used in the control system.
(e) Turnbuckles must be attached toparts having angular motion in a man-ner that will positively prevent bindingthroughout the range of travel.
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Federal Aviation Administration, DOT § 23.697
(f) Tab control cables are not part ofthe primary control system and may beless than 1⁄8 inch diameter in airplanesthat are safely controllable with thetabs in the most adverse positions.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–7, 34 FR 13091, Aug. 13,1969]
§ 23.691 Artificial stall barrier system.
If the function of an artificial stallbarrier, for example, stick pusher, isused to show compliance with§ 23.201(c), the system must complywith the following:
(a) With the system adjusted for op-eration, the plus and minus airspeedsat which downward pitching controlwill be provided must be established.
(b) Considering the plus and minusairspeed tolerances established byparagraph (a) of this section, an air-speed must be selected for the activa-tion of the downward pitching controlthat provides a safe margin above anyairspeed at which any unsatisfactorystall characteristics occur.
(c) In addition to the stall warningrequired § 23.07, a warning that is clear-ly distinguishable to the pilot under allexpected flight conditions without re-quiring the pilot’s attention, must beprovided for faults that would preventthe system from providing the requiredpitching motion.
(d) Each system must be designed sothat the artificial stall barrier can bequickly and positively disengaged bythe pilots to prevent unwanted down-ward pitching of the airplane by aquick release (emergency) control thatmeets the requirements of § 23.1329(b).
(e) A preflight check of the completesystem must be established and theprocedure for this check made avail-able in the Airplane Flight Manual(AFM). Preflight checks that are crit-ical to the safety of the airplane mustbe included in the limitations sectionof the AFM.
(f) For those airplanes whose designincludes an autopilot system:
(1) A quick release (emergency) con-trol installed in accordance with§ 23.1329(b) may be used to meet the re-quirements of paragraph (d), of thissection, and
(2) The pitch servo for that systemmay be used to provide the stall down-ward pitching motion.
(g) In showing compliance with§ 23.1309, the system must be evaluatedto determine the effect that any an-nounced or unannounced failure mayhave on the continued safe flight andlanding of the airplane or the ability ofthe crew to cope with any adverse con-ditions that may result from such fail-ures. This evaluation must considerthe hazards that would result from theairplane’s flight characteristics if thesystem was not provided, and the haz-ard that may result from unwanteddownward pitching motion, whichcould result from a failure at airspeedsabove the selected stall speed.
[Doc. No. 27806, 61 FR 5165, Feb. 9, 1996]
§ 23.693 Joints.
Control system joints (in push-pullsystems) that are subject to angularmotion, except those in ball and rollerbearing systems, must have a specialfactor of safety of not less than 3.33with respect to the ultimate bearingstrength of the softest material used asa bearing. This factor may be reducedto 2.0 for joints in cable control sys-tems. For ball or roller bearings, theapproved ratings may not be exceeded.
§ 23.697 Wing flap controls.
(a) Each wing flap control must bedesigned so that, when the flap hasbeen placed in any position upon whichcompliance with the performance re-quirements of this part is based, theflap will not move from that positionunless the control is adjusted or ismoved by the automatic operation of aflap load limiting device.
(b) The rate of movement of the flapsin response to the operation of the pi-lot’s control or automatic device mustgive satisfactory flight and perform-ance characteristics under steady orchanging conditions of airspeed, enginepower, and attitude.
(c) If compliance with § 23.145(b)(3)necessitates wing flap retraction to po-sitions that are not fully retracted, thewing flap control lever settings cor-responding to those positions must bepositively located such that a definitechange of direction of movement of the
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14 CFR Ch. I (1–1–01 Edition)§ 23.699
lever is necessary to select settings be-yond those settings.
§ 23.699 Wing flap position indicator.There must be a wing flap position
indicator for—(a) Flap installations with only the
retracted and fully extended position,unless—
(1) A direct operating mechanismprovides a sense of ‘‘feel’’ and position(such as when a mechanical linkage isemployed); or
(2) The flap position is readily deter-mined without seriously detractingfrom other piloting duties under anyflight condition, day or night; and
(b) Flap installation with inter-mediate flap positions if—
(1) Any flap position other than re-tracted or fully extended is used toshow compliance with the performancerequirements of this part; and
(2) The flap installation does notmeet the requirements of paragraph(a)(1) of this section.
§ 23.701 Flap interconnection.(a) The main wing flaps and related
movable surfaces as a system must—(1) Be synchronized by a mechanical
interconnection between the movableflap surfaces that is independent of theflap drive system; or by an approvedequivalent means; or
(2) Be designed so that the occur-rence of any failure of the flap systemthat would result in an unsafe flightcharacteristic of the airplane is ex-tremely improbable; or
(b) The airplane must be shown tohave safe flight characteristics withany combination of extreme positionsof individual movable surfaces (me-chanically interconnected surfaces areto be considered as a single surface).
(c) If an interconnection is used inmultiengine airplanes, it must be de-signed to account for theunsummetrical loads resulting fromflight with the engines on one side ofthe plane of symmetry inoperative andthe remaining engines at takeoffpower. For single-engine airplanes, andmultiengine airplanes with no slip-stream effects on the flaps, it may be
assumed that 100 percent of the criticalair load acts on one side and 70 percenton the other.
unless it can be shown that a lift orlongitudinal trim device that affectsthe takeoff performance of the aircraftwould not give an unsafe takeoff con-figuration when selection out of an ap-proved takeoff position, a takeoffwarning system must be installed andmeet the following requirements:
(a) The system must provide to thepilots an aural warning that is auto-matically activated during the initialportion of the takeoff role if the air-plane is in a configuration that wouldnot allow a safe takeoff. The warningmust continue until—
(1) The configuration is changed toallow safe takeoff, or
(2) Action is taken by the pilot toabandon the takeoff roll.
(b) The means used to activate thesystem must function properly for allauthorized takeoff power settings andprocedures and throughout the rangesof takeoff weights, altitudes, and tem-peratures for which certification is re-quested.
[Doc. No. 27806, 61 FR 5166, Feb. 9, 1996]
LANDING GEAR
§ 23.721 General.For commuter category airplanes
that have a passenger seating configu-ration, excluding pilot seats, of 10 ormore, the following general require-ments for the landing gear apply:
(a) The main landing-gear systemmust be designed so that if it fails dueto overloads during takeoff and landing(assuming the overloads to act in theupward and aft directions), the failuremode is not likely to cause the spillageof enough fuel from any part of the fuelsystem to consitute a fire hazard.
(b) Each airplane must be designed sothat, with the airplane under control,it can be landed on a paved runwaywith any one or more landing-gear legs
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Federal Aviation Administration, DOT § 23.726
not extended without sustaining astructural component failure that islikely to cause the spillage of enoughfuel to consitute a fire hazard.
(c) Compliance with the provisions ofthis section may be shown by analysisor tests, or both.
[Amdt. 23–34, 52 FR 1830, Jan. 15, 1987]
§ 23.723 Shock absorption tests.(a) It must be shown that the limit
load factors selected for design in ac-cordance with § 23.473 for takeoff andlanding weights, respectively, will notbe exceeded. This must be shown by en-ergy absorption tests except that anal-ysis based on tests conducted on alanding gear system with identical en-ergy absorption characteristics may beused for increases in previously ap-proved takeoff and landing weights.
(b) The landing gear may not fail, butmay yield, in a test showing its reserveenergy absorption capacity, simulatinga descent velocity of 1.2 times the limitdescent velocity, assuming wing liftequal to the weight of the airplane.
§ 23.725 Limit drop tests.(a) If compliance with § 23.723(a) is
shown by free drop tests, these testsmust be made on the complete air-plane, or on units consisting of wheel,tire, and shock absorber, in their prop-er relation, from free drop heights notless than those determined by the fol-lowing formula:
h (inches) = 3.6 (W/S) 1⁄2
However, the free drop height may notbe less than 9.2 inches and need not bemore than 18.7 inches.
(b) If the effect of wing lift is pro-vided for in free drop tests, the landinggear must be dropped with an effectiveweight equal to
W Wh L d
h de =
+ −( )[ ]+( )
1
where—
We =the effective weight to be used in thedrop test (lbs.);
h = specified free drop height (inches);
d = deflection under impact of the tire (atthe approved inflation pressure) plus thevertical component of the axle travel rel-ative to the drop mass (inches);
W=WM for main gear units (lbs), equal to thestatic weight on that unit with the air-plane in the level attitude (with the nosewheel clear in the case of nose wheeltype airplanes);
W=WT for tail gear units (lbs.), equal to thestatic weight on the tail unit with theairplane in the tail-down attitude;
W=WN for nose wheel units lbs.), equal to thevertical component of the static reactionthat would exist at the nose wheel, as-suming that the mass of the airplaneacts at the center of gravity and exerts aforce of 1.0 g downward and 0.33 g for-ward; and
L= the ratio of the assumed wing lift to theairplane weight, but not more than 0.667.
(c) The limit inertia load factor mustbe determined in a rational or conserv-ative manner, during the drop test,using a landing gear unit attitude, andapplied drag loads, that represent thelanding conditions.
(d) The value of d used in the com-putation of We in paragraph (b) of thissection may not exceed the value actu-ally obtained in the drop test.
(e) The limit inertia load factor mustbe determined from the drop test inparagraph (b) of this section accordingto the following formula:
n nW
WLj
e= +
where—
nj=the load factor developed in the drop test(that is, the acceleration (dv/dt) in g’ s re-corded in the drop test) plus 1.0; and
We, W, and L are the same as in the drop testcomputation.
(f) The value of n determined in ac-cordance with paragraph (e) may notbe more than the limit inertia load fac-tor used in the landing conditions in§ 23.473.
§ 23.726 Ground load dynamic tests.(a) If compliance with the ground
load requirements of §§ 23.479 through23.483 is shown dynamically by droptest, one drop test must be conductedthat meets § 23.725 except that the dropheight must be—
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14 CFR Ch. I (1–1–01 Edition)§ 23.727
(1) 2.25 times the drop height pre-scribed in § 23.725(a); or
(2) Sufficient to develop 1.5 times thelimit load factor.
(b) The critical landing condition foreach of the design conditions specifiedin §§ 23.479 through 23.483 must be usedfor proof of strength.
[Amdt. 23–7, 34 FR 13091, Aug. 13, 1969]
§ 23.727 Reserve energy absorptiondrop test.
(a) If compliance with the reserve en-ergy absorption requirement in§ 23.723(b) is shown by free drop tests,the drop height may not be less than1.44 times that specified in § 23.725.
(b) If the effect of wing lift is pro-vided for, the units must be droppedwith an effective mass equal to We=Wh/(h+d), when the symbols and other de-tails are the same as in § 23.725.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–7, 34 FR 13091, Aug. 13,1969]
(a) General. For airplanes with re-tractable landing gear, the followingapply:
(1) Each landing gear retractingmechanism and its supporting struc-ture must be designed for maximumflight load factors with the gear re-tracted and must be designed for thecombination of friction, inertia, braketorque, and air loads, occurring duringretraction at any airspeed up to 1.6 VS1
with flaps retracted, and for any loadfactor up to those specified in § 23.345for the flaps-extended condition.
(2) The landing gear and retractingmechanism, including the wheel welldoors, must withstand flight loads, in-cluding loads resulting from all yawingconditions specified in § 23.351, with thelanding gear extended at any speed upto at least 1.6 VS1 with the flaps re-tracted.
(b) Landing gear lock. There must bepositive means (other than the use ofhydraulic pressure) to keep the landinggear extended.
(c) Emergency operation. For a land-plane having retractable landing gearthat cannot be extended manually,there must be means to extend thelanding gear in the event of either—
(1) Any reasonably probable failure inthe normal landing gear operation sys-tem; or
(2) Any reasonably probable failure ina power source that would prevent theoperation of the normal landing gearoperation system.
(d) Operation test. The proper func-tioning of the retracting mechanismmust be shown by operation tests.
(e) Position indicator. If a retractablelanding gear is used, there must be alanding gear position indicator (as wellas necessary switches to actuate theindicator) or other means to inform thepilot that each gear is secured in theextended (or retracted) position. Ifswitches are used, they must be locatedand coupled to the landing gear me-chanical system in a manner that pre-vents an erroneous indication of either‘‘down and locked’’ if each gear is notin the fully extended position, or ‘‘upand locked’’ if each landing gear is notin the fully retracted position.
(f) Landing gear warning. For land-planes, the following aural or equallyeffective landing gear warning devicesmust be provided:
(1) A device that functions continu-ously when one or more throttles areclosed beyond the power settings nor-mally used for landing approach if thelanding gear is not fully extended andlocked. A throttle stop may not beused in place of an aural device. Ifthere is a manual shutoff for the warn-ing device prescribed in this paragraph,the warning system must be designedso that when the warning has been sus-pended after one or more throttles areclosed, subsequent retardation of anythrottle to, or beyond, the position fornormal landing approach will activatethe warning device.
(2) A device that functions continu-ously when the wing flaps are extendedbeyond the maximum approach flap po-sition, using a normal landing proce-dure, if the landing gear is not fully ex-tended and locked. There may not be amanual shutoff for this warning device.The flap position sensing unit may beinstalled at any suitable location. Thesystem for this device may use anypart of the system (including the auralwarning device) for the device requiredin paragraph (f)(1) of this section.
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Federal Aviation Administration, DOT § 23.735
(g) Equipment located in the landinggear bay. If the landing gear bay is usedas the location for equipment otherthan the landing gear, that equipmentmust be designed and installed to mini-mize damage from items such as a tireburst, or rocks, water, and slush thatmay enter the landing gear bay.
§ 23.731 Wheels.(a) The maximum static load rating
of each wheel may not be less than thecorresponding static ground reactionwith—
(1) Design maximum weight; and(2) Critical center of gravity.(b) The maximum limit load rating of
each wheel must equal or exceed themaximum radial limit load determinedunder the applicable ground load re-quirements of this part.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–45, 58 FR 42165, Aug. 6,1993]
§ 23.733 Tires.(a) Each landing gear wheel must
have a tire whose approved tire ratings(static and dynamic) are not exceed-ed—
(1) By a load on each main wheel tire)to be compared to the static rating ap-proved for such tires) equal to the cor-responding static ground reactionunder the design maximum weight andcritical center of gravity; and
(2) By a load on nose wheel tires (tobe compared with the dynamic ratingapproved for such tires) equal to the re-action obtained at the nose wheel, as-suming the mass of the airplane to beconcentrated at the most critical cen-ter of gravity and exerting a force of1.0 W downward and 0.31 W forward(where W is the design maximumweight), with the reactions distributedto the nose and main wheels by theprinciples of statics and with the dragreaction at the ground applied only atwheels with brakes.
(b) If specially constructed tires areused, the wheels must be plainly andconspicuously marked to that effect.
The markings must include the make,size, number of plies, and identificationmarking of the proper tire.
(c) Each tire installed on a retract-able landing gear system must, at themaximum size of the tire type expectedin service, have a clearance to sur-rounding structure and systems that isadequate to prevent contact betweenthe tire and any part of the structureof systems.
(a) Brakes must be provided. Thelanding brake kinetic energy capacityrating of each main wheel brake assem-bly must not be less than the kineticenergy absorption requirements deter-mined under either of the followingmethods:
(1) The brake kinetic energy absorp-tion requirements must be based on aconservative rational analysis of thesequence of events expected duringlanding at the design landing weight.
(2) Instead of a rational analysis, thekinetic energy absorption require-ments for each main wheel brake as-sembly may be derived from the fol-lowing formula:
KE=0.0443 WV2/N
where—
KE=Kinetic energy per wheel (ft.-lb.);W=Design landing weight (lb.);V=Airplane speed in knots. V must be not
less than VS√, the poweroff stalling speedof the airplane at sea level, at the designlanding weight, and in the landing con-figuration; and
N=Number of main wheels with brakes.
(b) Brakes must be able to preventthe wheels from rolling on a paved run-way with takeoff power on the criticalengine, but need not prevent movementof the airplane with wheels locked.
(c) During the landing distance deter-mination required by § 23.75, the pres-sure on the wheel braking system mustnot exceed the pressure specified by thebrake manufacturer.
(d) If antiskid devices are installed,the devices and associated systemsmust be designed so that no single
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14 CFR Ch. I (1–1–01 Edition)§ 23.737
probable malfunction or failure will re-sult in a hazardous loss of braking abil-ity or directional control of the air-plane.
(e) In addition, for commuter cat-egory airplanes, the rejected takeoffbrake kinetic energy capacity rating ofeach main wheel brake assembly mustnot be less than the kinetic energy ab-sorption requirements determinedunder either of the following methods—
(1) The brake kinetic energy absorp-tion requirements must be based on aconservative rational analysis of thesequence of events expected during arejected takeoff at the design takeoffweight.
(2) Instead of a rational analysis, thekinetic energy absorption require-ments for each main wheel brake as-sembly may be derived from the fol-lowing formula—KE=0.0443 WV2Nwhere,KE=Kinetic energy per wheel (ft.-lbs.);W=Design takeoff weight (lbs.);V=Ground speed, in knots, associated
with the maximum value of V1 se-lected in accordance with§ 23.51(c)(1);
each ski must equal or exceed the max-imum limit load determined under theapplicable ground load requirements ofthis part.
[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]
§ 23.745 Nose/tail wheel steering.(a) If nose/tail wheel steering is in-
stalled, it must be demonstrated thatits use does not require exceptionalpilot skill during takeoff and landing,in crosswinds, or in the event of an en-gine failure; or its use must be limitedto low speed maneuvering.
(b) Movement of the pilot’s steeringcontrol must not interfere with the re-traction or extension of the landinggear.
[Doc. No. 27806, 61 FR 5166, Feb. 9, 1996]
FLOATS AND HULLS
§ 23.751 Main float buoyancy.
(a) Each main float must have—(1) A buoyancy of 80 percent in excess
of the buoyancy required by that floatto support its portion of the maximumweight of the seaplane or amphibian infresh water; and
(2) Enough watertight compartmentsto provide reasonable assurance thatthe seaplane or amphibian will stayafloat without capsizing if any twocompartments of any main float areflooded.
(b) Each main float must contain atleast four watertight compartmentsapproximately equal in volume.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–45, 58 FR 42165, Aug. 6,1993]
§ 23.753 Main float design.
Each seaplane main float must meetthe requirements of § 23.521.
[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]
§ 23.755 Hulls.(a) The hull of a hull seaplane or am-
phibian of 1,500 pounds or more max-imum weight must have watertightcompartments designed and arrangedso that the hull auxiliary floats, andtires (if used), will keep the airplaneafloat without capsizing in fresh waterwhen—
(1) For airplanes of 5,000 pounds ormore maximum weight, any two adja-cent compartments are flooded; and
(2) For airplanes of 1,500 pounds upto, but not including, 5,000 pounds max-imum weight, any single compartmentis flooded.
(b) Watertight doors in bulkheadsmay be used for communication be-tween compartments.
§ 23.757 Auxiliary floats.Auxiliary floats must be arranged so
that, when completely submerged infresh water, they provide a rightingmoment of at least 1.5 times the upset-ting moment caused by the seaplane oramphibian being tilted.
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Federal Aviation Administration, DOT § 23.775
PERSONNEL AND CARGOACCOMMODATIONS
§ 23.771 Pilot compartment.For each pilot compartment—(a) The compartment and its equip-
ment must allow each pilot to performhis duties without unreasonable con-centration or fatigue;
(b) Where the flight crew are sepa-rated from the passengers by a parti-tion, an opening or openable window ordoor must be provided to facilitatecommunication between flight crewand the passengers; and
(c) The aerodynamic controls listedin § 23.779, excluding cables and controlrods, must be located with respect tothe propellers so that no part of thepilot or the controls lies in the regionbetween the plane of rotation of anyinboard propeller and the surface gen-erated by a line passing through thecenter of the propeller hub making anangle of 5 degrees forward or aft of theplane of rotation of the propeller.
§ 23.773 Pilot compartment view.(a) Each pilot compartment must
be—(1) Arranged with sufficiently exten-
sive, clear and undistorted view to en-able the pilot to safely taxi, takeoff,approach, land, and perform any ma-neuvers within the operating limita-tions of the airplane.
(2) Free from glare and reflectionsthat could interfere with the pilot’s vi-sion. Compliance must be shown in alloperations for which certification is re-quested; and
(3) Designed so that each pilot is pro-tected from the elements so that mod-erate rain conditions do not unduly im-pair the pilot’s view of the flight pathin normal flight and while landing.
(b) Each pilot compartment musthave a means to either remove or pre-vent the formation of fog or frost on anarea of the internal portion of thewindshield and side windows suffi-ciently large to provide the view speci-fied in paragraph (a)(1) of this section.Compliance must be shown under allexpected external and internal ambientoperating conditions, unless it can be
shown that the windshield and sidewindows can be easily cleared by thepilot without interruption of moralpilot duties.
[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]
§ 23.775 Windshields and windows.
(a) The internal panels of windshieldsand windows must be constructed of anonsplintering material, such as non-splintering safety glass.
(b) The design of windshields, win-dows, and canopies in pressurized air-planes must be based on factors pecu-liar to high altitude operation, includ-ing—
(1) The effects of continuous and cy-clic pressurization loadings;
(2) The inherent characteristics ofthe material used; and
(3) The effects of temperatures andtemperature gradients.
(c) On pressurized airplanes, if cer-tification for operation up to and in-cluding 25,000 feet is requested, an en-closure canopy including a representa-tive part of the installation must besubjected to special tests to accountfor the combined effects of continuousand cyclic pressurization loadings andflight loads, or compliance with thefail-safe requirements of paragraph (d)of this section must be shown.
(d) If certification for operationabove 25,000 feet is requested the wind-shields, window panels, and canopiesmust be strong enough to withstandthe maximum cabin pressure differen-tial loads combined with critical aero-dynamic pressure and temperature ef-fects, after failure of any load-carryingelement of the windshield, windowpanel, or canopy.
(e) The windshield and side windowsforward of the pilot’s back when thepilot is seated in the normal flight po-sition must have a luminous transmit-tance value of not less than 70 percent.
(f) Unless operation in known or fore-cast icing conditions is prohibited byoperating limitations, a means must beprovided to prevent or to clear accumu-lations of ice from the windshield sothat the pilot has adequate view fortaxi, takeoff, approach, landing, and toperform any maneuvers within the op-erating limitations of the airplane.
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14 CFR Ch. I (1–1–01 Edition)§ 23.777
(g) In the event of any probable sin-gle failure, a transparency heating sys-tem must be incapable of raising thetemperature of any windshield or win-dow to a point where there would be—
(1) Structural failure that adverselyaffects the integrity of the cabin; or
(2) There would be a danger of fire.(h) In addition, for commuter cat-
egory airplanes, the following applies:(1) Windshield panes directly in front
of the pilots in the normal conduct oftheir duties, and the supporting struc-tures for these panes, must withstand,without penetration, the impact of atwo-pound bird when the velocity ofthe airplane (relative to the bird alongthe airplane’s flight path) is equal tothe airplane’s maximum approach flapspeed.
(2) The windshield panels in front ofthe pilots must be arranged so that, as-suming the loss of vision through anyone panel, one or more panels remainavailable for use by a pilot seated at apilot station to permit continued safeflight and landing.
§ 23.777 Cockpit controls.(a) Each cockpit control must be lo-
cated and (except where its function isobvious) identified to provide conven-ient operation and to prevent confusionand inadvertent operation.
(b) The controls must be located andarranged so that the pilot, when seat-ed, has full and unrestricted movementof each control without interferencefrom either his clothing or the cockpitstructure.
(c) Powerplant controls must be lo-cated—
(1) For multiengine airplanes, on thepedestal or overhead at or near thecenter of the cockpit;
(2) For single and tandem seated sin-gle-engine airplanes, on the left sideconsole or instrument panel;
(3) For other single-engine airplanesat or near the center of the cockpit, onthe pedestal, instrument panel, oroverhead; and
(4) For airplanes, with side-by-sidepilot seats and with two sets of power-
plant controls, on left and right con-soles.
(d) The control location order fromleft to right must be power (thrust)lever, propeller (rpm control), and mix-ture control (condition lever and fuelcutoff for turbine-powered airplanes).Power (thrust) levers must be at leastone inch higher or longer to makethem more prominent than propeller(rpm control) or mixture controls. Car-buretor heat or alternate air controlmust be to the left of the throttle or atleast eight inches from the mixturecontrol when located other than on apedestal. Carburetor heat or alternateair control, when located on a pedestalmust be aft or below the power (thrust)lever. Supercharger controls must belocated below or aft of the propellercontrols. Airplanes with tandem seat-ing or single-place airplanes may uti-lize control locations on the left side ofthe cabin compartment; however, loca-tion order from left to right must bepower (thrust) lever, propeller (rpmcontrol) and mixture control.
(e) Identical powerplant controls foreach engine must be located to preventconfusion as to the engines they con-trol.
(1) Conventional multiengine power-plant controls must be located so thatthe left control(s) operates the left en-gines(s) and the right control(s) oper-ates the right engine(s).
(2) On twin-engine airplanes withfront and rear engine locations (tan-dem), the left powerplant controlsmust operate the front engine and theright powerplant controls must operatethe rear engine.
(f) Wing flap and auxiliary lift devicecontrols must be located—
(1) Centrally, or to the right of thepedestal or powerplant throttle controlcenterline; and
(2) Far enough away from the landinggear control to avoid confusion.
(g) The landing gear control must belocated to the left of the throttle cen-terline or pedestal centerline.
(h) Each fuel feed selector controlmust comply with § 23.995 and be lo-cated and arranged so that the pilotcan see and reach it without movingany seat or primary flight controlwhen his seat is at any position inwhich it can be placed.
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Federal Aviation Administration, DOT § 23.781
(1) For a mechanical fuel selector:(i) The indication of the selected fuel
valve position must be by means of apointer and must provide positive iden-tification and feel (detent, etc.) of theselected position.
(ii) The position indicator pointermust be located at the part of the han-dle that is the maximum dimension ofthe handle measured from the center ofrotation.
(2) For electrical or electronic fuelselector:
(i) Digital controls or electricalswitches must be properly labelled.
(ii) Means must be provided to indi-cate to the flight crew the tank orfunction selected. Selector switch posi-tion is not acceptable as a means of in-dication. The ‘‘off’’ or ‘‘closed’’ posi-tion must be indicated in red.
(3) If the fuel valve selector handle orelectrical or digital selection is also afuel shut-off selector, the off positionmarking must be colored red. If a sepa-rate emergency shut-off means is pro-vided, it also must be colored red.
Cockpit controls must be designed sothat they operate in accordance withthe following movement and actuation:
(a) Aerodynamic controls:
Motion and effect
(1) Primary con-trols:
Aileron ...... Right (clockwise) for rightwing down.
Elevator ..... Rearward for nose up.Rudder ....... Right pedal forward for
nose right.(2) Secondary
controls:Flaps (or
auxiliarylift de-vices).
Forward or up for flaps upor auxiliary devicestowed; rearward ordown for flaps down orauxiliary device de-ployed.
Motion and effect
Trim tabs(or equiv-alent).
Switch motion or mechan-ical rotation of controlto produce similar rota-tion of the airplaneabout an axis parallel tothe axis control. Axis ofroll trim control may bedisplaced to accommo-date comfortable actu-ation by the pilot. Forsingle-engine airplanes,direction of pilot’s handmovement must be inthe same sense as air-plane response for rud-der trim if only a por-tion of a rotational ele-ment is accessible.
(b) Powerplant and auxiliary con-trols:
Motion and effect
(1) Powerplantcontrols:
Power(thrust)lever.
Forward to increase for-ward thrust and rear-ward to increase rear-ward thrust.
Propellers .. Forward to increase rpm.Mixture ...... Forward or upward for
rich.Fuel ........... Forward for open.Carburetor,
air heator alter-nate air.
Forward or upward forcold.
Super-charger.
Forward or upward for lowblower.
Turbosuper-chargers.
Forward, upward, orclockwise to increasepressure.
Rotary con-trols.
Clockwise from off to fullon.
(2) Auxiliarycontrols:
Fuel tankselector.
Right for right tanks, leftfor left tanks.
Landinggear.
Down to extend.
Speedbrakes.
Aft to extend.
[Amdt. 23–33, 51 FR 26656, July 24, 1986, asamended by Amdt. 23–51, 61 FR 5136, Feb. 9,1996]
§ 23.781 Cockpit control knob shape.(a) Flap and landing gear control
knobs must conform to the generalshapes (but not necessarily the exactsizes or specific proportions) in the fol-lowing figure:
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14 CFR Ch. I (1–1–01 Edition)§ 23.781
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Federal Aviation Administration, DOT § 23.783
(b) Powerplant control knobs mustconform to the general shapes (but notnecessarily the exact sizes or specificproportions) in the following figure:
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30FR 258, Jan. 9, 1965, as amended by Amdt. 23–33, 51 FR 26657, July 24, 1986]
§ 23.783 Doors.(a) Each closed cabin with passenger
accommodations must have at leastone adequate and easily accessible ex-ternal door.
(b) Passenger doors must not be lo-cated with respect to any propeller
disk or any other potential hazard soas to endanger persons using the door.
(c) Each external passenger or crewdoor must comply with the followingrequirements:
(1) There must be a means to lockand safeguard the door against inad-vertent opening during flight by per-sons, by cargo, or as a result of me-chanical failure.
(2) The door must be openable fromthe inside and the outside when the in-ternal locking mechanism is in thelocked position.
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14 CFR Ch. I (1–1–01 Edition)§ 23.785
(3) There must be a means of openingwhich is simple and obvious and is ar-ranged and marked inside and outsideso that the door can be readily located,unlocked, and opened, even in dark-ness.
(4) The door must meet the markingrequirements of § 23.811 of this part.
(5) The door must be reasonably freefrom jamming as a result of fuselagedeformation in an emergency landing.
(6) Auxiliary locking devices that areactuated externally to the airplanemay be used but such devices must beoverridden by the normal internalopening means.
(d) In addition, each external pas-senger or crew door, for a commutercategory airplane, must comply withthe following requirements:
(1) Each door must be openable fromboth the inside and outside, eventhough persons may be crowdedagainst the door on the inside of theairplane.
(2) If inward opening doors are used,there must be a means to prevent occu-pants from crowding against the doorto the extent that would interfere withopening the door.
(3) Auxiliary locking devices may beused.
(e) Each external door on a com-muter category airplane, each externaldoor forward of any engine or propelleron a normal, utility, or acrobatic cat-egory airplane, and each door of thepressure vessel on a pressurized air-plane must comply with the followingrequirements:
(1) There must be a means to lockand safeguard each external door, in-cluding cargo and service type doors,against inadvertent opening in flight,by persons, by cargo, or as a result ofmechanical failure or failure of a singlestructural element, either during orafter closure.
(2) There must be a provision for di-rect visual inspection of the lockingmechanism to determine if the exter-nal door, for which the initial openingmovement is not inward, is fully closedand locked. The provisions must be dis-cernible, under operating lighting con-ditions, by a crewmember using aflashlight or an equivalent lightingsource.
(3) There must be a visual warningmeans to signal a flight crewmember ifthe external door is not fully closedand locked. The means must be de-signed so that any failure, or combina-tion of failures, that would result in anerroneous closed and locked indicationis improbable for doors for which theinitial opening movement is not in-ward.
(f) In addition, for commuter cat-egory airplanes, the following require-ments apply:
(1) Each passenger entry door mustqualify as a floor level emergency exit.This exit must have a rectangularopening of not less than 24 inches wideby 48 inches high, with corner radii notgreater than one-third the width of theexit.
(2) If an integral stair is installed ata passenger entry door, the stair mustbe designed so that, when subjected tothe inertia loads resulting from the ul-timate static load factors in§ 23.561(b)(2) and following the collapseof one or more legs of the landing gear,it will not reduce the effectiveness ofemergency egress through the pas-senger entry door.
(g) If lavatory doors are installed,they must be designed to preclude anoccupant from becoming trapped insidethe lavatory. If a locking mechanism isinstalled, it must be capable of beingunlocked from outside of the lavatory.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30FR 258, Jan. 9, 1965, as amended by Amdt. 23–36, 53 FR 30813, Aug. 15, 1988; Amdt. 23–46, 59FR 25772, May 17, 1994; Amdt. 23–49, 61 FR5166, Feb. 9, 1996]
§ 23.785 Seats, berths, litters, safetybelts, and shoulder harnesses.
There must be a seat or berth foreach occupant that meets the fol-lowing:
(a) Each seat/restraint system andthe supporting structure must be de-signed to support occupants weighingat least 215 pounds when subjected tothe maximum load factors cor-responding to the specified flight andground load conditions, as defined inthe approved operating envelope of theairplane. In addition, these loads mustbe multiplied by a factor of 1.33 in de-termining the strength of all fittingsand the attachment of—
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Federal Aviation Administration, DOT § 23.785
(1) Each seat to the structure; and(2) Each safety belt and shoulder har-
ness to the seat or structure.(b) Each forward-facing or aft-facing
seat/restraint system in normal, util-ity, or acrobatic category airplanesmust consist of a seat, a safety belt,and a shoulder harness, with a metal-to-metal latching device, that are de-signed to provide the occupant protec-tion provisions required in § 23.562.Other seat orientations must providethe same level of occupant protectionas a forward-facing or aft-facing seatwith a safety belt and a shoulder har-ness, and must provide the protectionprovisions of § 23.562.
(c) For commuter category airplanes,each seat and the supporting structuremust be designed for occupants weigh-ing at least 170 pounds when subjectedto the inertia loads resulting from theultimate static load factors prescribedin § 23.561(b)(2) of this part. Each occu-pant must be protected from serioushead injury when subjected to the iner-tia loads resulting from these load fac-tors by a safety belt and shoulder har-ness, with a metal-to-metal latchingdevice, for the front seats and a safetybelt, or a safety belt and shoulder har-ness, with a metal-to-metal latchingdevice, for each seat other than thefront seats.
(d) Each restraint system must havea single-point release for occupantevacuation.
(e) The restraint system for eachcrewmember must allow the crew-member, when seated with the safetybelt and shoulder harness fastened, toperform all functions necessary forflight operations.
(f) Each pilot seat must be designedfor the reactions resulting from the ap-plication of pilot forces to the primaryflight controls as prescribed in § 23.395of this part.
(g) There must be a means to secureeach safety belt and shoulder harness,when not in use, to prevent inter-ference with the operation of the air-plane and with rapid occupant egress inan emergency.
(h) Unless otherwise placarded, eachseat in a utility or acrobatic categoryairplane must be designed to accommo-date an occupant wearing a parachute.
(i) The cabin area surrounding eachseat, including the structure, interiorwalls, instrument panel, control wheel,pedals, and seats within striking dis-tance of the occupant’s head or torso(with the restraint system fastened)must be free of potentially injuriousobjects, sharp edges, protuberances,and hard surfaces. If energy absorbingdesigns or devices are used to meet thisrequirement, they must protect the oc-cupant from serious injury when theoccupant is subjected to the inertialoads resulting from the ultimate stat-ic load factors prescribed in§ 23.561(b)(2) of this part, or they mustcomply with the occupant protectionprovisions of § 23.562 of this part, as re-quired in paragraphs (b) and (c) of thissection.
(j) Each seat track must be fittedwith stops to prevent the seat fromsliding off the track.
(k) Each seat/restraint system mayuse design features, such as crushing orseparation of certain components, toreduce occupant loads when showingcompliance with the requirements of§ 23.562 of this part; otherwise, the sys-tem must remain intact.
(l) For the purposes of this section, afront seat is a seat located at a flightcrewmember station or any seat lo-cated alongside such a seat.
(m) Each berth, or provisions for alitter, installed parallel to the longitu-dinal axis of the airplane, must be de-signed so that the forward part has apadded end-board, canvas diaphragm,or equivalent means that can with-stand the load reactions from a 215-pound occupant when subjected to theinertia loads resulting from the ulti-mate static load factors of § 23.561(b)(2)of this part. In addition—
(1) Each berth or litter must have anoccupant restraint system and may nothave corners or other parts likely tocause serious injury to a person occu-pying it during emergency landing con-ditions; and
(2) Occupant restraint system attach-ments for the berth or litter mustwithstand the inertia loads resultingfrom the ultimate static load factors of§ 23.561(b)(2) of this part.
(n) Proof of compliance with the stat-ic strength requirements of this sec-tion for seats and berths approved as
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14 CFR Ch. I (1–1–01 Edition)§ 23.787
part of the type design and for seat andberth installations may be shown by—
(1) Structural analysis, if the struc-ture conforms to conventional airplanetypes for which existing methods ofanalysis are known to be reliable;
(2) A combination of structural anal-ysis and static load tests to limit load;or
(1) Be designed for its placarded max-imum weight of contents and for thecritical load distributions at the appro-priate maximum load factors cor-responding to the flight and groundload conditions of this part.
(2) Have means to prevent the con-tents of any compartment from becom-ing a hazard by shifting, and to protectany controls, wiring, lines, equipmentor accessories whose damage or failurewould affect safe operations.
(3) Have a means to protect occu-pants from injury by the contents ofany compartment, located aft of theoccupants and separated by structure,when the ultimate forward inertialload factor is 9g and assuming the max-imum allowed baggage or cargo weightfor the compartment.
(b) Designs that provide for baggageor cargo to be carried in the same com-partment as passengers must have ameans to protect the occupants frominjury when the baggage or cargo issubjected to the inertial loads result-ing from the ultimate static load fac-tors of § 23.561(b)(3), assuming the max-imum allowed baggage or cargo weightfor the compartment.
(c) For airplanes that are used onlyfor the carriage of cargo, the flightcrewemergency exits must meet the re-quirements of § 23.807 under any cargoloading conditions.
[Doc. No. 27806, 61 FR 5167, Feb. 9, 1996]
§ 23.791 Passenger information signs.For those airplanes in which the
flightcrew members cannot observe the
other occupants’ seats or where theflightcrew members’ compartment isseparated from the passenger compart-ment, there must be at least one illu-minated sign (using either letters orsymbols) notifying all passengers whenseat belts should be fastened. Signsthat notify when seat belts should befastened must:
(a) When illuminated, be legible toeach person seated in the passengercompartment under all probable light-ing conditions; and
(b) Be installed so that a flightcrewmember can, when seated at theflightcrew member’s station, turn theillumination on and off.
[Doc. No. 27806, 61 FR 5167, Feb. 9, 1996]
§ 23.803 Emergency evacuation.
(a) For commuter category airplanes,an evacuation demonstration must beconducted utilizing the maximumnumber of occupants for which certifi-cation is desired. The demonstrationmust be conducted under simulatednight conditions using only the emer-gency exits on the most critical side ofthe airplane. The participants must berepresentative of average airline pas-sengers with no prior practice or re-hearsal for the demonstration. Evacu-ation must be completed within 90 sec-onds.
(b) In addition, when certification tothe emergency exit provisions of§ 23.807(d)(4) is requested, only theemergency lighting system required by§ 23.812 may be used to provide cabin in-terior illumination during the evacu-ation demonstration required in para-graph (a) of this section.
[Amdt. 23–34, 52 FR 1831, Jan. 15, 1987, asamended by Amdt. 23–46, 59 FR 25773, May 17,1994]
§ 23.805 Flightcrew emergency exits.
For airplanes where the proximity ofthe passenger emergency exits to theflightcrew area does not offer a conven-ient and readily accessible means ofevacuation for the flightcrew, the fol-lowing apply:
(a) There must be either one emer-gency exit on each side of the airplane,or a top hatch emergency exit, in theflightcrew area;
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Federal Aviation Administration, DOT § 23.807
(b) Each emergency exit must be lo-cated to allow rapid evacuation of thecrew and have a size and shape of atleast a 19- by 20-inch unobstructed rec-tangular opening; and
(c) For each emergency exit that isnot less than six feet from the ground,an assisting means must be provided.The assisting means may be a rope orany other means demonstrated to besuitable for the purpose. If the assist-ing means is a rope, or an approved de-vice equivalent to a rope, it must be—
(1) Attached to the fuselage structureat or above the top of the emergencyexit opening or, for a device at a pilot’semergency exit window, at another ap-proved location if the stowed device, orits attachment, would reduce the pi-lot’s view; and
(2) Able (with its attachment) towithstand a 400-pound static load.
[Doc. No. 26324, 59 FR 25773, May 17, 1994]
§ 23.807 Emergency exits.(a) Number and location. Emergency
exits must be located to allow escapewithout crowding in any probablecrash attitude. The airplane must haveat least the following emergency exits:
(1) For all airplanes with a seatingcapacity of two or more, excluding air-planes with canopies, at least oneemergency exit on the opposite side ofthe cabin from the main door specifiedin § 23.783 of this part.
(2) [Reserved](3) If the pilot compartment is sepa-
rated from the cabin by a door that islikely to block the pilot’s escape in aminor crash, there must be an exit inthe pilot’s compartment. The numberof exits required by paragraph (a)(1) ofthis section must then be separatelydetermined for the passenger compart-ment, using the seating capacity ofthat compartment.
(4) Emergency exits must not be lo-cated with respect to any propellerdisk or any other potential hazard soas to endanger persons using that exit.
(b) Type and operation. Emergencyexits must be movable windows, panels,canopies, or external doors, openablefrom both inside and outside the air-plane, that provide a clear and unob-structed opening large enough to admita 19-by-26-inch ellipse. Auxiliary lock-ing devices used to secure the airplane
must be designed to be overridden bythe normal internal opening means.The inside handles of emergency exitsthat open outward must be adequatelyprotected against inadvertent oper-ation. In addition, each emergency exitmust—
(1) Be readily accessible, requiring noexceptional agility to be used in emer-gencies;
(2) Have a method of opening that issimple and obvious;
(3) Be arranged and marked for easylocation and operation, even in dark-ness;
(4) Have reasonable provisionsagainst jamming by fuselage deforma-tion; and
(5) In the case of acrobatic categoryairplanes, allow each occupant to aban-don the airplane at any speed betweenVSO and VD; and
(6) In the case of utility category air-planes certificated for spinning, alloweach occupant to abandon the airplaneat the highest speed likely to beachieved in the maneuver for which theairplane is certificated.
(c) Tests. The proper functioning ofeach emergency exit must be shown bytests.
(d) Doors and exits. In addition, forcommuter category airplanes, the fol-lowing requirements apply:
(1) In addition to the passenger entrydoor—
(i) For an airplane with a total pas-senger seating capacity of 15 or fewer,an emergency exit, as defined in para-graph (b) of this section, is required oneach side of the cabin; and
(ii) For an airplane with a total pas-senger seating capacity of 16 through19, three emergency exits, as defined inparagraph (b) of this section, are re-quired with one on the same side as thepassenger entry door and two on theside opposite the door.
(2) A means must be provided to lockeach emergency exit and to safeguardagainst its opening in flight, either in-advertently by persons or as a result ofmechanical failure. In addition, ameans for direct visual inspection ofthe locking mechanism must be pro-vided to determine that each emer-gency exit for which the initial openingmovement is outward is fully locked.
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14 CFR Ch. I (1–1–01 Edition)§ 23.811
(3) Each required emergency exit, ex-cept floor level exits, must be locatedover the wing or, if not less than sixfeet from the ground, must be providedwith an acceptable means to assist theoccupants to descend to the ground.Emergency exits must be distributed asuniformly as practical, taking into ac-count passenger seating configuration.
(4) Unless the applicant has compliedwith paragraph (d)(1) of this section,there must be an emergency exit onthe side of the cabin opposite the pas-senger entry door, provided that—
(i) For an airplane having a pas-senger seating configuration of nine orfewer, the emergency exit has a rectan-gular opening measuring not less than19 inches by 26 inches high with cornerradii not greater than one-third thewidth of the exit, located over thewing, with a step up inside the airplaneof not more than 29 inches and a stepdown outside the airplane of not morethan 36 inches;
(ii) For an airplane having a pas-senger seating configuration of 10 to 19passengers, the emergency exit has arectangular opening measuring not lessthan 20 inches wide by 36 inches high,with corner radii not greater than one-third the width of the exit, and with astep up inside the airplane of not morethan 20 inches. If the exit is locatedover the wing, the step down outsidethe airplane may not exceed 27 inches;and
(iii) The airplane complies with theadditional requirements of§§ 23.561(b)(2)(iv), 23.803(b), 23.811(c),23.812, 23.813(b), and 23.815.
(e) For multiengine airplanes, ditch-ing emergency exits must be providedin accordance with the following re-quirements, unless the emergency exitsrequired by paragraph (a) or (d) of thissection already comply with them:
(1) One exit above the waterline oneach side of the airplane having the di-mensions specified in paragraph (b) or(d) of this section, as applicable; and
(2) If side exits cannot be above thewaterline, there must be a readily ac-cessible overhead hatch emergency exitthat has a rectangular opening meas-uring not less than 20 inches wide by 36inches long, with corner radii not
§ 23.811 Emergency exit marking.(a) Each emergency exit and external
door in the passenger compartmentmust be externally marked and readilyidentifiable from outside the airplaneby—
(1) A conspicuous visual identifica-tion scheme; and
(2) A permanent decal or placard onor adjacent to the emergency exitwhich shows the means of opening theemergency exit, including any specialinstructions, if applicable.
(b) In addition, for commuter cat-egory airplanes, these exits and doorsmust be internally marked with theword ‘‘exit’’ by a sign which has whiteletters 1 inch high on a red background2 inches high, be self-illuminated orindependently, internally electricallyilluminated, and have a minimumbrightness of at least 160 micro-lamberts. The color may be reversed ifthe passenger compartment illumina-tion is essentially the same.
(c) In addition, when certification tothe emergency exit provisions of§ 23.807(d)(4) is requested, the followingapply:
(1) Each emergency exit, its means ofaccess, and its means of opening, mustbe conspicuously marked;
(2) The identity and location of eachemergency exit must be recognizablefrom a distance equal to the width ofthe cabin;
(3) Means must be provided to assistoccupants in locating the emergencyexits in conditions of dense smoke;
(4) The location of the operating han-dle and instructions for opening eachemergency exit from inside the air-plane must be shown by marking thatis readable from a distance of 30 inches;
(5) Each passenger entry door oper-ating handle must—
(i) Be self-illuminated with an initialbrightness of at least 160 micro-lamberts; or
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Federal Aviation Administration, DOT § 23.812
(ii) Be conspicuously located and wellilluminated by the emergency lightingeven in conditions of occupant crowd-ing at the door;
(6) Each passenger entry door with alocking mechanism that is released byrotary motion of the handle must bemarked—
(i) With a red arrow, with a shaft ofat least three-fourths of an inch wideand a head twice the width of the shaft,extending along at least 70 degrees ofarc at a radius approximately equal tothree-fourths of the handle length;
(ii) So that the center line of the exithandle is within ± one inch of the pro-jected point of the arrow when the han-dle has reached full travel and has re-leased the locking mechanism;
(iii) With the word ‘‘open’’ in red let-ters, one inch high, placed horizontallynear the head of the arrow; and
(7) In addition to the requirements ofparagraph (a) of this section, the exter-nal marking of each emergency exitmust—
(i) Include a 2-inch colorband out-lining the exit; and
(ii) Have a color contrast that isreadily distinguishable from the sur-rounding fuselage surface. The contrastmust be such that if the reflectance ofthe darker color is 15 percent or less,the reflectance of the lighter colormust be at least 45 percent. ‘‘Reflec-tance’’ is the ratio of the luminous fluxreflected by a body to the luminousflux it receives. When the reflectanceof the darker color is greater than 15percent, at least a 30 percent differencebetween its reflectance and the reflec-tance of the lighter color must be pro-vided.
[Amdt. 23–36, 53 FR 30814, Aug. 15, 1988; 53 FR34194, Sept. 2, 1988, as amended by Amdt. 23–46, 59 FR 25773, May 17, 1994]
§ 23.812 Emergency lighting.
When certification to the emergencyexit provisions of § 23.807(d)(4) is re-quested, the following apply:
(a) An emergency lighting system,independent of the main cabin lightingsystem, must be installed. However,the source of general cabin illumina-tion may be common to both the emer-gency and main lighting systems if thepower supply to the emergency light-
ing system is independent of the powersupply to the main lighting system.
(b) There must be a crew warninglight that illuminates in the cockpitwhen power is on in the airplane andthe emergency lighting control deviceis not armed.
(c) The emergency lights must be op-erable manually from the flightcrewstation and be provided with automaticactivation. The cockpit control devicemust have ‘‘on,’’ ‘‘off,’’ and ‘‘armed’’positions so that, when armed in thecockpit, the lights will operate byautomatic activation.
(d) There must be a means to safe-guard against inadvertent operation ofthe cockpit control device from the‘‘armed’’ or ‘‘on’’ positions.
(e) The cockpit control device musthave provisions to allow the emergencylighting system to be armed or acti-vated at any time that it may be need-ed.
(f) When armed, the emergency light-ing system must activate and remainlighted when—
(1) The normal electrical power ofthe airplane is lost; or
(2) The airplane is subjected to animpact that results in a deceleration inexcess of 2g and a velocity change inexcess of 3.5 feet-per-second, actingalong the longitudinal axis of the air-plane; or
(3) Any other emergency conditionexists where automatic activation ofthe emergency lighting is necessary toaid with occupant evacuation.
(g) The emergency lighting systemmust be capable of being turned off andreset by the flightcrew after automaticactivation.
(h) The emergency lighting systemmust provide internal lighting, includ-ing—
(1) Illuminated emergency exit mark-ing and locating signs, including thoserequired in § 23.811(b);
(2) Sources of general illumination inthe cabin that provide an average illu-mination of not less than 0.05 foot-can-dle and an illumination at any point ofnot less than 0.01 foot-candle whenmeasured along the center line of themain passenger aisle(s) and at the seatarmrest height; and
(3) Floor proximity emergency escapepath marking that provides emergency
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14 CFR Ch. I (1–1–01 Edition)§23.813
evacuation guidance for the airplaneoccupants when all sources of illumina-tion more than 4 feet above the cabinaisle floor are totally obscured.
(i) The energy supply to each emer-gency lighting unit must provide therequired level of illumination for atleast 10 minutes at the critical ambientconditions after activation of theemergency lighting system.
(j) If rechargeable batteries are usedas the energy supply for the emergencylighting system, they may be re-charged from the main electrical powersystem of the airplane provided thecharging circuit is designed to precludeinadvertent battery discharge into thecharging circuit faults. If the emer-gency lighting system does not includea charging circuit, battery conditionmonitors are required.
(k) Components of the emergencylighting system, including batteries,wiring, relays, lamps, and switches,must be capable of normal operationafter being subjected to the inertiaforces resulting from the ultimate loadfactors prescribed in § 23.561(b)(2).
(l) The emergency lighting systemmust be designed so that after any sin-gle transverse vertical separation ofthe fuselage during a crash landing:
(1) At least 75 percent of all elec-trically illuminated emergency lightsrequired by this section remain opera-tive; and
(2) Each electrically illuminated exitsign required by § 23.811 (b) and (c) re-mains operative, except those that aredirectly damaged by the fuselage sepa-ration.
[Doc. No. 26324, 59 FR 25774, May 17, 1994]
§23.813 Emergency exit access.(a) For commuter category airplanes,
access to window-type emergency exitsmay not be obstructed by seats or seatbacks.
(b) In addition, when certification tothe emergency exit provisions of§ 23.807(d)(4) is requested, the followingemergency exit access must be pro-vided:
(1) The passageway leading from theaisle to the passenger entry door mustbe unobstructed and at least 20 incheswide.
(2) There must be enough space nextto the passenger entry door to allow
assistance in evacuation of passengerswithout reducing the unobstructedwidth of the passageway below 20inches.
(3) If it is necessary to pass througha passageway between passenger com-partments to reach a required emer-gency exit from any seat in the pas-senger cabin, the passageway must beunobstructed; however, curtains maybe used if they allow free entrythrough the passageway.
(4) No door may be installed in anypartition between passenger compart-ments unless that door has a means tolatch it in the open position. The latch-ing means must be able to withstandthe loads imposed upon it by the doorwhen the door is subjected to the iner-tia loads resulting from the ultimatestatic load factors prescribed in§ 23.561(b)(2).
(5) If it is necessary to pass througha doorway separating the passengercabin from other areas to reach a re-quired emergency exit from any pas-senger seat, the door must have ameans to latch it in the open position.The latching means must be able towithstand the loads imposed upon it bythe door when the door is subjected tothe inertia loads resulting from the ul-timate static load factors prescribed in§ 23.561(b)(2).
[Amdt. 23–36, 53 FR 30815, Aug. 15, 1988, asamended by Amdt. 23–46, 59 FR 25774, May 17,1994]
§ 23.815 Width of aisle.
(a) Except as provided in paragraph(b) of this section, for commuter cat-egory airplanes, the width of the mainpassenger aisle at any point betweenseats must equal or exceed the valuesin the following table:
Number of pas-senger seats
Minimum main passenger aisle width
Less than 25inches from floor
25 inches andmore from floor
10 through 19 ....... 9 inches ................ 15 inches.
(b) When certification to the emer-gency exist provisions of § 23.807(d)(4) isrequested, the main passenger aislewidth at any point between the seatsmust equal or exceed the following val-ues:
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Federal Aviation Administration, DOT § 23.843
Number of passenger seats
Minimum main passengeraisle width (inches)
Less than25 inchesfrom floor
25 inchesand morefrom floor
10 or fewer ................................ 1 12 1511 through 19 ............................ 12 20
1 A narrower width not less than 9 inches may be approvedwhen substantiated by tests found necessary by theAdministrator.
[Amdt. 23–34, 52 FR 1831, Jan. 15, 1987, asamended by Amdt. 23–46, 59 FR 25774, May 17,1994]
§ 23.831 Ventilation.(a) Each passenger and crew compart-
ment must be suitably ventilated. Car-bon monoxide concentration may notexceed one part in 20,000 parts of air.
(b) For pressurized airplanes, theventilating air in the flightcrew andpassenger compartments must be freeof harmful or hazardous concentrationsof gases and vapors in normal oper-ations and in the event of reasonablyprobable failures or malfunctioning ofthe ventilating, heating, pressuriza-tion, or other systems and equipment.If accumulation of hazardous quan-tities of smoke in the cockpit area isreasonably probable, smoke evacuationmust be readily accomplished startingwith full pressurization and withoutdepressurizing beyond safe limits.
§ 23.841 Pressurized cabins.(a) If certification for operation over
25,000 feet is requested, the airplanemust be able to maintain a cabin pres-sure altitude of not more than 15,000feet in event of any probable failure ormalfunction in the pressurization sys-tem.
(b) Pressurized cabins must have atleast the following valves, controls,and indicators, for controlling cabinpressure:
(1) Two pressure relief valves to auto-matically limit the positive pressuredifferential to a predetermined valueat the maximum rate of flow deliveredby the pressure source. The combinedcapacity of the relief valves must be
large enough so that the failure of anyone valve would not cause an appre-ciable rise in the pressure differential.The pressure differential is positivewhen the internal pressure is greaterthan the external.
(2) Two reverse pressure differentialrelief valves (or their equivalent) toautomatically prevent a negative pres-sure differential that would damagethe structure. However, one valve isenough if it is of a design that reason-ably precludes its malfunctioning.
(3) A means by which the pressuredifferential can be rapidly equalized.
(4) An automatic or manual regulatorfor controlling the intake or exhaustairflow, or both, for maintaining therequired internal pressures and airflowrates.
(5) Instruments to indicate to thepilot the pressure differential, thecabin pressure altitude, and the rate ofchange of cabin pressure altitude.
(6) Warning indication at the pilotstation to indicate when the safe orpreset pressure differential is exceededand when a cabin pressure altitude of10,000 feet is exceeded.
(7) A warning placard for the pilot ifthe structure is not designed for pres-sure differentials up to the maximumrelief valve setting in combinationwith landing loads.
(8) A means to stop rotation of thecompressor or to divert airflow fromthe cabin if continued rotation of anengine-driven cabin compressor or con-tinued flow of any compressor bleed airwill create a hazard if a malfunctionoccurs.
(a) Strength test. The complete pres-surized cabin, including doors, win-dows, canopy, and valves, must be test-ed as a pressure vessel for the pressuredifferential specified in § 23.365(d).
(b) Functional tests. The followingfunctional tests must be performed:
(1) Tests of the functioning and ca-pacity of the positive and negativepressure differential valves, and of theemergency release valve, to simulatethe effects of closed regulator valves.
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14 CFR Ch. I (1–1–01 Edition)§ 23.851
(2) Tests of the pressurization systemto show proper functioning under eachpossible condition of pressure, tem-perature, and moisture, up to the max-imum altitude for which certificationis requested.
(3) Flight tests, to show the perform-ance of the pressure supply, pressureand flow regulators, indicators, andwarning signals, in steady and steppedclimbs and descents at rates cor-responding to the maximum attainablewithin the operating limitations of theairplane, up to the maximum altitudefor which certification is requested.
(4) Tests of each door and emergencyexit, to show that they operate prop-erly after being subjected to the flighttests prescribed in paragraph (b)(3) ofthis section.
FIRE PROTECTION
§ 23.851 Fire extinguishers.
(a) There must be at least one handfire extinguisher for use in the pilotcompartment that is located withineasy access of the pilot while seated.
(b) There must be at least one handfire extinguisher located convenientlyin the passenger compartment—
(1) Of each airplane accommodatingmore than 6 passengers; and
(2) Of each commuter category air-plane.
(c) For hand fire extinguishers, thefollowing apply:
(1) The type and quantity of each ex-tinguishing agent used must be appro-priate to the kinds of fire likely tooccur where that agent is to be used.
(2) Each extinguisher for use in a per-sonnel compartment must be designedto minimize the hazard of toxic gasconcentrations.
[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]
§ 23.853 Passenger and crew compart-ment interiors.
For each compartment to be used bythe crew or passengers:
(a) The materials must be at leastflame-resistant;
(b) [Reserved](c) If smoking is to be prohibited,
there must be a placard so stating, andif smoking is to be allowed—
(1) There must be an adequate num-ber of self-contained, removable ash-trays; and
(2) Where the crew compartment isseparated from the passenger compart-ment, there must be at least one illu-minated sign (using either letters orsymbols) notifying all passengers whensmoking is prohibited. Signs which no-tify when smoking is prohibited must—
(i) When illuminated, be legible toeach passenger seated in the passengercabin under all probable lighting condi-tions; and
(ii) Be so constructed that the crewcan turn the illumination on and off;and
(d) In addition, for commuter cat-egory airplanes the following require-ments apply:
(1) Each disposal receptacle for tow-els, paper, or waste must be fully en-closed and constructed of at least fireresistant materials and must containfires likely to occur in it under normaluse. The ability of the disposal recep-tacle to contain those fires under allprobable conditions of wear, misalign-ment, and ventilation expected in serv-ice must be demonstrated by test. Aplacard containing the legible words‘‘No Cigarette Disposal’’ must be lo-cated on or near each disposal recep-tacle door.
(2) Lavatories must have ‘‘No Smok-ing’’ or ‘‘No Smoking in Lavatory’’placards located conspicuously on eachside of the entry door and self-con-tained, removable ashtrays locatedconspicuously on or near the entry sideof each lavatory door, except that oneashtray may serve more than one lava-tory door if it can be seen from thecabin side of each lavatory door served.The placards must have red letters atleast 1⁄2 inch high on a white back-ground at least 1 inch high (a ‘‘NoSmoking’’ symbol may be included onthe placard).
(3) Materials (including finishes ordecorative surfaces applied to the ma-terials) used in each compartment oc-cupied by the crew or passengers mustmeet the following test criteria as ap-plicable:
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flooring, and materials used in the con-struction of stowage compartments(other than underseat stowage com-partments and compartments for stow-ing small items such as magazines andmaps) must be self-extinguishing whentested vertically in accordance withthe applicable portions of appendix Fof this part or by other equivalentmethods. The average burn length maynot exceed 6 inches and the averageflame time after removal of the flamesource may not exceed 15 seconds.Drippings from the test specimen maynot continue to flame for more than anaverage of 3 seconds after falling.
(ii) Floor covering, textiles (includ-ing draperies and upholstery), seatcushions, padding, decorative and non-decorative coated fabrics, leather,trays and galley furnishings, electricalconduit, thermal and acoustical insula-tion and insulation covering, air duct-ing, joint and edge covering, cargocompartment liners, insulation blan-kets, cargo covers and transparencies,molded and thermoformed parts, airducting joints, and trim strips (decora-tive and chafing), that are constructedof materials not covered in paragraph(d)(3)(iv) of this section must be self ex-tinguishing when tested vertically inaccordance with the applicable por-tions of appendix F of this part orother approved equivalent methods.The average burn length may not ex-ceed 8 inches and the average flametime after removal of the flame sourcemay not exceed 15 seconds. Drippingsfrom the test specimen may not con-tinue to flame for more than an aver-age of 5 seconds after falling.
(iii) Motion picture film must besafety film meeting the Standard Spec-ifications for Safety PhotographicFilm PH1.25 (available from the Amer-ican National Standards Institute, 1430Broadway, New York, N.Y. 10018) or anFAA approved equivalent. If the filmtravels through ducts, the ducts mustmeet the requirements of paragraph(d)(3)(ii) of this section.
(iv) Acrylic windows and signs, partsconstructed in whole or in part of elas-tomeric materials, edge-lighted instru-ment assemblies consisting of two ormore instruments in a common hous-ing, seatbelts, shoulder harnesses, and
cargo and baggage tiedown equipment,including containers, bins, pallets, etc.,used in passenger or crew compart-ments, may not have an average burnrate greater than 2.5 inches per minutewhen tested horizontally in accordancewith the applicable portions of appen-dix F of this part or by other approvedequivalent methods.
(v) Except for electrical wire cableinsulation, and for small parts (such asknobs, handles, rollers, fasteners, clips,grommets, rub strips, pulleys, andsmall electrical parts) that the Admin-istrator finds would not contribute sig-nificantly to the propagation of a fire,materials in items not specified inparagraphs (d)(3)(i), (ii), (iii), or (iv) ofthis section may not have a burn rategreater than 4.0 inches per minutewhen tested horizontally in accordancewith the applicable portions of appen-dix F of this part or by other approvedequivalent methods.
(e) Lines, tanks, or equipment con-taining fuel, oil, or other flammablefluids may not be installed in suchcompartments unless adequatelyshielded, isolated, or otherwise pro-tected so that any breakage or failureof such an item would not create a haz-ard.
(f) Airplane materials located on thecabin side of the firewall must be self-extinguishing or be located at such adistance from the firewall, or otherwiseprotected, so that ignition will notoccur if the firewall is subjected to aflame temperature of not less than2,000 degrees F for 15 minutes. For self-extinguishing materials (except elec-trical wire and cable insulation andsmall parts that the Administratorfinds would not contribute signifi-cantly to the propagation of a fire), avertifical self-extinguishing test mustbe conducted in accordance with appen-dix F of this part or an equivalentmethod approved by the Adminis-trator. The average burn length of thematerial may not exceed 6 inches andthe average flame time after removalof the flame source may not exceed 15seconds. Drippings from the materialtest specimen may not continue to
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14 CFR Ch. I (1–1–01 Edition)§ 23.855
flame for more than an average of 3seconds after falling.
§ 23.855 Cargo and baggage compart-ment fire protection.
(a) Sources of heat within each cargoand baggage compartment that are ca-pable of igniting the compartment con-tents must be shielded and insulated toprevent such ignition.
(b) Each cargo and baggage compart-ment must be constructed of materialsthat meet the appropriate provisions of§ 23.853(d)(3).
(c) In addition, for commuter cat-egory airplanes, each cargo and bag-gage compartment must:
(1) Be located where the presence of afire would be easily discovered by thepilots when seated at their duty sta-tion, or it must be equipped with asmoke or fire detector system to give awarning at the pilots’ station, and pro-vide sufficient access to enable a pilotto effectively reach any part of thecompartment with the contents of ahand held fire extinguisher, or
(2) Be equipped with a smoke or firedetector system to give a warning atthe pilots’ station and have ceiling andsidewall liners and floor panels con-structed of materials that have beensubjected to and meet the 45 degreeangle test of appendix F of this part.The flame may not penetrate (passthrough) the material during applica-tion of the flame or subsequent to itsremoval. The average flame time afterremoval of the flame source may notexceed 15 seconds, and the average glowtime may not exceed 10 seconds. Thecompartment must be constructed toprovide fire protection that is not lessthan that required of its individualpanels; or
(3) Be constructed and sealed to con-tain any fire within the compartment.
[Doc. No. 27806, 61 FR 5167, Feb. 9, 1996]
§ 23.859 Combustion heater fire pro-tection.
(a) Combustion heater fire regions. Thefollowing combustion heater fire re-gions must be protected from fire in ac-
cordance with the applicable provisionsof §§23.1182 through 23.1191 and 23.1203:
(1) The region surrounding the heat-er, if this region contains any flam-mable fluid system components (ex-cluding the heater fuel system) thatcould—
(i) Be damaged by heater malfunc-tioning; or
(ii) Allow flammable fluids or vaporsto reach the heater in case of leakage.
(2) The region surrounding the heat-er, if the heater fuel system has fit-tings that, if they leaked, would allowfuel vapor to enter this region.
(3) The part of the ventilating airpassage that surrounds the combustionchamber.
(b) Ventilating air ducts. Each ven-tilating air duct passing through anyfire region must be fireproof. In addi-tion—
(1) Unless isolation is provided byfireproof valves or by equally effectivemeans, the ventilating air duct down-stream of each heater must be fireprooffor a distance great enough to ensurethat any fire originating in the heatercan be contained in the duct; and
(2) Each part of any ventilating ductpassing through any region having aflammable fluid system must be con-structed or isolated from that systemso that the malfunctioning of any com-ponent of that system cannot intro-duce flammable fluids or vapors intothe ventilating airstream.
(c) Combustion air ducts. Each com-bustion air duct must be fireproof for adistance great enough to prevent dam-age from backfiring or reverse flamepropagation. In addition—
(1) No combustion air duct may havea common opening with the ventilatingairstream unless flames from backfiresor reverse burning cannot enter theventilating airstream under any oper-ating condition, including reverse flowor malfunctioning of the heater or itsassociated components; and
(2) No combustion air duct may re-strict the prompt relief of any backfirethat, if so restricted, could cause heat-er failure.
(d) Heater controls: general. Provisionmust be made to prevent the hazardousaccumulation of water or ice on or inany heater control component, controlsystem tubing, or safety control.
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(e) Heater safety controls. (1) Eachcombustion heater must have the fol-lowing safety controls:
(i) Means independent of the compo-nents for the normal continuous con-trol of air temperature, airflow, andfuel flow must be provided to auto-matically shut off the ignition and fuelsupply to that heater at a point remotefrom that heater when any of the fol-lowing occurs:
(A) The heater exchanger tempera-ture exceeds safe limits.
(B) The ventilating air temperatureexceeds safe limits.
(C) The combustion airflow becomesinadequate for safe operation.
(D) The ventilating airflow becomesinadequate for safe operation.
(ii) Means to warn the crew when anyheater whose heat output is essentialfor safe operation has been shut off bythe automatic means prescribed inparagraph (e)(1)(i) of this section.
(2) The means for complying withparagraph (e)(1)(i) of this section forany individual heater must—
(i) Be independent of componentsserving any other heater whose heatoutput is essential for safe operations;and
(ii) Keep the heater off until re-started by the crew.
(f) Air intakes. Each combustion andventilating air intake must be locatedso that no flammable fluids or vaporscan enter the heater system under anyoperating condition—
(1) During normal operation; or(2) As a result of the malfunctioning
of any other component.(g) Heater exhaust. Heater exhaust
systems must meet the provisions of§§ 23.1121 and 23.1123. In addition, theremust be provisions in the design of theheater exhaust system to safely expelthe products of combustion to preventthe occurrence of—
(1) Fuel leakage from the exhaust tosurrounding compartments;
(2) Exhaust gas impingement on sur-rounding equipment or structure;
(3) Ignition of flammable fluids bythe exhaust, if the exhaust is in a com-partment containing flammable fluidlines; and
(4) Restrictions in the exhaust sys-tem to relieve backfires that, if so re-stricted, could cause heater failure.
(h) Heater fuel systems. Each heaterfuel system must meet each power-plant fuel system requirement affect-ing safe heater operation. Each heaterfuel system component within the ven-tilating airstream must be protectedby shrouds so that no leakage fromthose components can enter the ven-tilating airstream.
(i) Drains. There must be means tosafely drain fuel that might accumu-late within the combustion chamber orthe heater exchanger. In addition—
(1) Each part of any drain that oper-ates at high temperatures must be pro-tected in the same manner as heaterexhausts; and
(2) Each drain must be protectedfrom hazardous ice accumulation underany operating condition.
[Amdt. 23–27, 45 FR 70387, Oct. 23, 1980]
§ 23.863 Flammable fluid fire protec-tion.
(a) In each area where flammablefluids or vapors might escape by leak-age of a fluid system, there must bemeans to minimize the probability ofignition of the fluids and vapors, andthe resultant hazard if ignition doesoccur.
(b) Compliance with paragraph (a) ofthis section must be shown by analysisor tests, and the following factors mustbe considered:
(1) Possible sources and paths of fluidleakage, and means of detecting leak-age.
(2) Flammability characteristics offluids, including effects of any combus-tible or absorbing materials.
(3) Possible ignition sources, includ-ing electrical faults, overheating ofequipment, and malfunctioning of pro-tective devices.
(4) Means available for controlling orextinguishing a fire, such as stoppingflow of fluids, shutting down equip-ment, fireproof containment, or use ofextinguishing agents.
(5) Ability of airplane componentsthat are critical to safety of flight towithstand fire and heat.
(c) If action by the flight crew is re-quired to prevent or counteract a fluidfire (e.g. equipment shutdown or actu-ation of a fire extinguisher), quick act-ing means must be provided to alertthe crew.
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(d) Each area where flammable fluidsor vapors might escape by leakage of afluid system must be identified and de-fined.
[Amdt. 23–23, 43 FR 50593, Oct. 30, 1978]
§ 23.865 Fire protection of flight con-trols, engine mounts, and otherflight structure.
Flight controls, engine mounts, andother flight structure located in des-ignated fire zones, or in adjacent areasthat would be subjected to the effectsof fire in the designated fire zones,must be constructed of fireproof mate-rial or be shielded so that they are ca-pable of withstanding the effects of afire. Engine vibration isolators mustincorporate suitable features to ensurethat the engine is retained if the non-fireproof portions of the isolators dete-riorate from the effects of a fire.
[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]
ELECTRICAL BONDING AND LIGHTNINGPROTECTION
§ 23.867 Electrical bonding and protec-tion against lightning and staticelectricity.
(a) The airplane must be protectedagainst catastrophic effects from light-ning.
(b) For metallic components, compli-ance with paragraph (a) of this sectionmay be shown by—
(1) Bonding the components properlyto the airframe; or
(2) Designing the components so thata strike will not endanger the airplane.
(c) For nonmetallic components,compliance with paragraph (a) of thissection may be shown by—
(1) Designing the components to min-imize the effect of a strike; or
(2) Incorporating acceptable means ofdiverting the resulting electrical cur-rent so as not to endanger the airplane.
[Amdt. 23–7, 34 FR 13092, Aug. 13, 1969]
MISCELLANEOUS
§ 23.871 Leveling means.There must be means for determining
when the airplane is in a level positionon the ground.
[Amdt. 23–7, 34 FR 13092, Aug. 13, 1969]
Subpart E—PowerplantGENERAL
§ 23.901 Installation.(a) For the purpose of this part, the
airplane powerplant installation in-cludes each component that—
(1) Is necessary for propulsion; and(2) Affects the safety of the major
propulsive units.(b) Each powerplant installation
must be constructed and arranged to—(1) Ensure safe operation to the max-
imum altitude for which approval is re-quested.
(2) Be accessible for necessary inspec-tions and maintenance.
(c) Engine cowls and nacelles must beeasily removable or openable by thepilot to provide adequate access to andexposure of the engine compartmentfor preflight checks.
(d) Each turbine engine installationmust be constructed and arranged to—
(1) Result in carcass vibration char-acteristics that do not exceed those es-tablished during the type certificationof the engine.
(2) Ensure that the capability of theinstalled engine to withstand the in-gestion of rain, hail, ice, and birds intothe engine inlet is not less than the ca-pability established for the engineitself under § 23.903(a)(2).
(e) The installation must complywith—
(1) The instructions provided underthe engine type certificate and the pro-peller type certificate.
(2) The applicable provisions of thissubpart.
(f) Each auxiliary power unit instal-lation must meet the applicable por-tions of this part.
§ 23.903 Engines.(a) Engine type certificate. (1) Each en-
gine must have a type certificate andmust meet the applicable requirementsof part 34 of this chapter.
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(2) Each turbine engine and its in-stallation must comply with one of thefollowing:
(i) Sections 33.76, 33.77 and 33.78 ofthis chapter in effect on December 13,2000.
(ii) Sections 33.77 and 33.78 of thischapter in effect on April 30, 1998, or assubsequently amended before Decem-ber 13, 2000; or
(iii) Section 33.77 of this chapter ineffect on October 31, 1974, or as subse-quently amended before April 30, 1998,unless that engine’s foreign object in-gestion service history has resulted inan unsafe condition; or
(iv) Be shown to have a foreign objectingestion service history in similar in-stallation locations which has not re-sulted in any unsafe condition.
NOTE: § 33.77 of this chapter in effect on Oc-tober 31, 1974, was published in 14 CFR parts1 to 59, Revised as of January 1, 1975. See 39FR 35467, October 1, 1974.
(1) Design precautions must be takento minimize the hazards to the airplanein the event of an engine rotor failureor of a fire originating inside the en-gine which burns through the enginecase.
(2) The powerplant systems associ-ated with engine control devices, sys-tems, and instrumentation must be de-signed to give reasonable assurancethat those operating limitations thatadversely affect turbine rotor struc-tural integrity will not be exceeded inservice.
(c) Engine isolation. The powerplantsmust be arranged and isolated fromeach other to allow operation, in atleast one configuration, so that thefailure or malfunction of any engine, orthe failure or malfunction (includingdestruction by fire in the engine com-partment) of any system that can af-fect an engine (other than a fuel tankif only one fuel tank is installed), willnot:
(1) Prevent the continued safe oper-ation of the remaining engines; or
(2) Require immediate action by anycrewmember for continued safe oper-ation of the remaining engines.
(d) Starting and stopping (piston en-gine). (1) The design of the installationmust be such that risk of fire or me-
chanical damage to the engine or air-plane, as a result of starting the enginein any conditions in which starting isto be permitted, is reduced to a min-imum. Any techniques and associatedlimitations for engine starting must beestablished and included in the Air-plane Flight Manual, approved manualmaterial, or applicable operating plac-ards. Means must be provided for—
(i) Restarting any engine of a multi-engine airplane in flight, and
(ii) Stopping any engine in flight,after engine failure, if continued en-gine rotation would cause a hazard tothe airplane.
(2) In addition, for commuter cat-egory airplanes, the following apply:
(i) Each component of the stoppingsystem on the engine side of the fire-wall that might be exposed to fire mustbe at least fire resistant.
(ii) If hydraulic propeller featheringsystems are used for this purpose, thefeathering lines must be at least fireresistant under the operating condi-tions that may be expected to existduring feathering.
(e) Starting and stopping (turbine en-gine). Turbine engine installationsmust comply with the following:
(1) The design of the installationmust be such that risk of fire or me-chanical damage to the engine or theairplane, as a result of starting the en-gine in any conditions in which start-ing is to be permitted, is reduced to aminimum. Any techniques and associ-ated limitations must be establishedand included in the Airplane FlightManual, approved manual material, orapplicable operating placards.
(2) There must be means for stoppingcombustion within any engine and forstopping the rotation of any engine ifcontinued rotation would cause a haz-ard to the airplane. Each component ofthe engine stopping system located inany fire zone must be fire resistant. Ifhydraulic propeller feathering systemsare used for stopping the engine, thehydraulic feathering lines or hosesmust be fire resistant.
(3) It must be possible to restart anengine in flight. Any techniques andassociated limitations must be estab-lished and included in the AirplaneFlight Manual, approved manual mate-rial, or applicable operating placards.
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(4) It must be demonstrated in flightthat when restarting engines followinga false start, all fuel or vapor is dis-charged in such a way that it does notconstitute a fire hazard.
(f) Restart envelope. An altitude andairspeed envelope must be establishedfor the airplane for in-flight engine re-starting and each installed enginemust have a restart capability withinthat envelope.
(g) Restart capability. For turbine en-gine powered airplanes, if the min-imum windmilling speed of the en-gines, following the in-flight shutdownof all engines, is insufficient to providethe necessary electrical power for en-gine ignition, a power source inde-pendent of the engine-driven electricalpower generating system must be pro-vided to permit in-flight engine igni-tion for restarting.
If installed, an automatic power re-serve (APR) system that automaticallyadvances the power or thrust on the op-erating engine(s), when any enginefails during takeoff, must comply withappendix H of this part.
[Doc. No. 26344, 58 FR 18970, Apr. 9, 1993]
§ 23.905 Propellers.(a) Each propeller must have a type
certificate.(b) Engine power and propeller shaft
rotational speed may not exceed thelimits for which the propeller is certifi-cated.
(c) Each featherable propeller musthave a means to unfeather it in flight.
(d) Each component of the propellerblade pitch control system must meetthe requirements of § 35.42 of this chap-ter.
(e) All areas of the airplane forwardof the pusher propeller that are likelyto accumulate and shed ice into thepropeller disc during any operating
condition must be suitably protectedto prevent ice formation, or it must beshown that any ice shed into the pro-peller disc will not create a hazardouscondition.
(f) Each pusher propeller must bemarked so that the disc is conspicuousunder normal daylight ground condi-tions.
(g) If the engine exhaust gases aredischarged into the pusher propellerdisc, it must be shown by tests, oranalysis supported by tests, that thepropeller is capable of continuous safeoperation.
(h) All engine cowling, access doors,and other removable items must be de-signed to ensure that they will not sep-arate from the airplane and contactthe pusher propeller.
§ 23.907 Propeller vibration.(a) Each propeller other than a con-
ventional fixed-pitch wooden propellermust be shown to have vibrationstresses, in normal operating condi-tions, that do not exceed values thathave been shown by the propeller man-ufacturer to be safe for continuous op-eration. This must be shown by—
(1) Measurement of stresses throughdirect testing of the propeller;
(2) Comparison with similar installa-tions for which these measurementshave been made; or
(3) Any other acceptable test methodor service experience that proves thesafety of the installation.
(b) Proof of safe vibration character-istics for any type of propeller, exceptfor conventional, fixed-pitch, wood pro-pellers must be shown where necessary.
§ 23.909 Turbocharger systems.(a) Each turbocharger must be ap-
proved under the engine type certifi-cate or it must be shown that the tur-bocharger system, while in its normalengine installation and operating inthe engine environment—
(1) Can withstand, without defect, anendurance test of 150 hours that meets
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Federal Aviation Administration, DOT § 23.929
the applicable requirements of § 33.49 ofthis subchapter; and
(2) Will have no adverse effect uponthe engine.
(b) Control system malfunctions, vi-brations, and abnormal speeds andtemperatures expected in service maynot damage the turbocharger com-pressor or turbine.
(c) Each turbocharger case must beable to contain fragments of a com-pressor or turbine that fails at thehighest speed that is obtainable withnormal speed control devices inoper-ative.
(d) Each intercooler installation,where provided, must comply with thefollowing—
(1) The mounting provisions of theintercooler must be designed to with-stand the loads imposed on the system;
(2) It must be shown that, under theinstalled vibration environment, theintercooler will not fail in a manner al-lowing portions of the intercooler to beingested by the engine; and
(3) Airflow through the intercoolermust not discharge directly on any air-plane component (e.g., windshield) un-less such discharge is shown to causeno hazard to the airplane under all op-erating conditions.
(e) Engine power, cooling character-istics, operating limits, and proceduresaffected by the turbocharger system in-stallations must be evaluated. Turbo-charger operating procedures and limi-tations must be included in the Air-plane Flight Manual in accordancewith § 23.1581.
[Amdt. 23–7, 34 FR 13092, Aug. 13, 1969, asamended by Amdt. 23–43, 58 FR 18970, Apr. 9,1993]
§ 23.925 Propeller clearance.Unless smaller clearances are sub-
stantiated, propeller clearances, withthe airplane at the most adverse com-bination of weight and center of grav-ity, and with the propeller in the mostadverse pitch position, may not be lessthan the following:
(a) Ground clearance. There must be aclearance of at least seven inches (foreach airplane with nose wheel landinggear) or nine inches (for each airplanewith tail wheel landing gear) betweeneach propeller and the ground with thelanding gear statically deflected and in
the level, normal takeoff, or taxing at-titude, whichever is most critical. Inaddition, for each airplane with con-ventional landing gear struts usingfluid or mechanical means for absorb-ing landing shocks, there must be posi-tive clearance between the propellerand the ground in the level takeoff at-titude with the critical tire completelydeflated and the corresponding landinggear strut bottomed. Positive clear-ance for airplanes using leaf springstruts is shown with a deflection cor-responding to 1.5g.
(b) Aft-mounted propellers. In additionto the clearances specified in para-graph (a) of this section, an airplanewith an aft mounted propeller must bedesigned such that the propeller willnot contact the runway surface whenthe airplane is in the maximum pitchattitude attainable during normaltakeoffs and landings.
(c) Water clearance. There must be aclearance of at least 18 inches betweeneach propeller and the water, unlesscompliance with § 23.239 can be shownwith a lesser clearance.
(d) Structural clearance. There mustbe—
(1) At least one inch radial clearancebetween the blade tips and the airplanestructure, plus any additional radialclearance necessary to prevent harmfulvibration;
(2) At least one-half inch longitudinalclearance between the propeller bladesor cuffs and stationary parts of the air-plane; and
(3) Positive clearance between otherrotating parts of the propeller or spin-ner and stationary parts of the air-plane.
Propellers (except wooden propellers)and other components of complete en-gine installations must be protectedagainst the accumulation of ice as nec-essary to enable satisfactory func-tioning without appreciable loss of
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thrust when operated in the icing con-ditions for which certification is re-quested.
§ 23.933 Reversing systems.(a) For turbojet and turbofan reversing
systems. (1) Each system intended forground operation only must be de-signed so that, during any reversal inflight, the engine will produce no morethan flight idle thrust. In addition, itmust be shown by analysis or test, orboth, that—
(i) Each operable reverser can be re-stored to the forward thrust position;or
(ii) The airplane is capable of contin-ued safe flight and landing under anypossible position of the thrust reverser.
(2) Each system intended for in-flightuse must be designed so that no unsafecondition will result during normal op-eration of the system, or from any fail-ure, or likely combination of failures,of the reversing system under any op-erating condition including ground op-eration. Failure of structural elementsneed not be considered if the prob-ability of this type of failure is ex-tremely remote.
(3) Each system must have a meansto prevent the engine from producingmore than idle thrust when the revers-ing system malfunctions; except that itmay produce any greater thrust that isshown to allow directional control tobe maintained, with aerodynamicmeans alone, under the most criticalreversing condition expected in oper-ation.
(b) For propeller reversing systems. (1)Each system must be designed so thatno single failure, likely combination offailures or malfunction of the systemwill result in unwanted reverse thrustunder any operating condition. Failureof structural elements need not be con-sidered if the probability of this type offailure is extremely remote.
(2) Compliance with paragraph (b)(1)of this section must be shown by fail-ure analysis, or testing, or both, forpropeller systems that allow the pro-peller blades to move from the flightlow-pitch position to a position that issubstantially less than the normal
flight, low-pitch position. The analysismay include or be supported by theanalysis made to show compliance with§ 35.21 for the type certification of thepropeller and associated installationcomponents. Credit will be given forpertinent analysis and testing com-pleted by the engine and propellermanufacturers.
§ 23.934 Turbojet and turbofan enginethrust reverser systems tests.
Thrust reverser systems of turbojetor turbofan engines must meet the re-quirements of § 33.97 of this chapter orit must be demonstrated by tests thatengine operation and vibratory levelsare not affected.
[Doc. No. 26344, 58 FR 18971, Apr. 9, 1993]
§ 23.937 Turbopropeller-drag limitingsystems.
(a) Turbopropeller-powered airplanepropeller-drag limiting systems mustbe designed so that no single failure ormalfunction of any of the systems dur-ing normal or emergency operation re-sults in propeller drag in excess of thatfor which the airplane was designedunder the structural requirements ofthis part. Failure of structural ele-ments of the drag limiting systemsneed not be considered if the prob-ability of this kind of failure is ex-tremely remote.
(b) As used in this section, drag lim-iting systems include manual or auto-matic devices that, when actuatedafter engine power loss, can move thepropeller blades toward the feather po-sition to reduce windmilling drag to asafe level.
[Amdt. 23–7, 34 FR 13093, Aug. 13, 1969, asamended by Amdt. 23–43, 58 FR 18971, Apr. 9,1993]
§ 23.939 Powerplant operating charac-teristics.
(a) Turbine engine powerplant oper-ating characteristics must be inves-tigated in flight to determine that noadverse characteristics (such as stall,surge, or flameout) are present, to ahazardous degree, during normal andemergency operation within the range
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Federal Aviation Administration, DOT § 23.954
of operating limitations of the airplaneand of the engine.
(b) Turbocharged reciprocating en-gine operating characteristics must beinvestigated in flight to assure that noadverse characteristics, as a result ofan inadvertent overboost, surge, flood-ing, or vapor lock, are present duringnormal or emergency operation of theengine(s) throughout the range of oper-ating limitations of both airplane andengine.
(c) For turbine engines, the air inletsystem must not, as a result of airflowdistortion during normal operation,cause vibration harmful to the engine.
§ 23.943 Negative acceleration.No hazardous malfunction of an en-
gine, an auxiliary power unit approvedfor use in flight, or any component orsystem associated with the powerplantor auxiliary power unit may occurwhen the airplane is operated at thenegative accelerations within theflight envelopes prescribed in § 23.333.This must be shown for the greatestvalue and duration of the accelerationexpected in service.
[Amdt. 23–18, 42 FR 15041, Mar. 17, 1977, asamended by Amdt. 23–43, 58 FR 18971, Apr. 9,1993]
FUEL SYSTEM
§ 23.951 General.(a) Each fuel system must be con-
structed and arranged to ensure fuelflow at a rate and pressure establishedfor proper engine and auxiliary powerunit functioning under each likely op-erating condition, including any ma-neuver for which certification is re-quested and during which the engine orauxiliary power unit is permitted to bein operation.
(b) Each fuel system must be ar-ranged so that—
(1) No fuel pump can draw fuel frommore than one tank at a time; or
(2) There are means to prevent intro-ducing air into the system.
(c) Each fuel system for a turbine en-gine must be capable of sustained oper-ation throughout its flow and pressure
range with fuel initially saturated withwater at 80° F and having 0.75cc of freewater per gallon added and cooled tothe most critical condition for icinglikely to be encountered in operation.
(d) Each fuel system for a turbine en-gine powered airplane must meet theapplicable fuel venting requirements ofpart 34 of this chapter.
(a) Each fuel system for a multien-gine airplane must be arranged so that,in at least one system configuration,the failure of any one component(other than a fuel tank) will not resultin the loss of power of more than oneengine or require immediate action bythe pilot to prevent the loss of power ofmore than one engine.
(b) If a single fuel tank (or series offuel tanks interconnected to functionas a single fuel tank) is used on amultiengine airplane, the followingmust be provided:
(1) Independent tank outlets for eachengine, each incorporating a shut-offvalve at the tank. This shutoff valvemay also serve as the fire wall shutoffvalve required if the line between thevalve and the engine compartment doesnot contain more than one quart offuel (or any greater amount shown tobe safe) that can escape into the enginecompartment.
(2) At least two vents arranged tominimize the probability of both ventsbecoming obstructed simultaneously.
(3) Filler caps designed to minimizethe probability of incorrect installa-tion or inflight loss.
(4) A fuel system in which those partsof the system from each tank outlet toany engine are independent of eachpart of the system supplying fuel toany other engine.
The fuel system must be designedand arranged to prevent the ignition offuel vapor within the system by—
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(a) Direct lightning strikes to areashaving a high probability of stroke at-tachment;
(b) Swept lightning strokes on areaswhere swept strokes are highly prob-able; and
(c) Corona or streamering at fuelvent outlets.
[Amdt. 23–7, 34 FR 13093, Aug. 13, 1969]
§ 23.955 Fuel flow.(a) General. The ability of the fuel
system to provide fuel at the ratesspecified in this section and at a pres-sure sufficient for proper engine oper-ation must be shown in the attitudethat is most critical with respect tofuel feed and quantity of unusable fuel.These conditions may be simulated in asuitable mockup. In addition—
(1) The quantity of fuel in the tankmay not exceed the amount establishedas the unusable fuel supply for thattank under § 23.959(a) plus that quan-tity necessary to show compliance withthis section.
(2) If there is a fuel flowmeter, itmust be blocked during the flow testand the fuel must flow through themeter or its bypass.
(3) If there is a flowmeter without abypass, it must not have any probablefailure mode that would restrict fuelflow below the level required for thisfuel demonstration.
(4) The fuel flow must include thatflow necessary for vapor return flow,jet pump drive flow, and for all otherpurposes for which fuel is used.
(b) Gravity systems. The fuel flow ratefor gravity systems (main and reservesupply) must be 150 percent of thetakeoff fuel consumption of the engine.
(c) Pump systems. The fuel flow ratefor each pump system (main and re-serve supply) for each reciprocating en-gine must be 125 percent of the fuelflow required by the engine at the max-imum takeoff power approved underthis part.
(1) This flow rate is required for eachmain pump and each emergency pump,and must be available when the pumpis operating as it would during takeoff;
(2) For each hand-operated pump,this rate must occur at not more than60 complete cycles (120 single strokes)per minute.
(3) The fuel pressure, with main andemergency pumps operating simulta-neously, must not exceed the fuel inletpressure limits of the engine unless itcan be shown that no adverse effect oc-curs.
(d) Auxiliary fuel systems and fueltransfer systems. Paragraphs (b), (c), and(f) of this section apply to each auxil-iary and transfer system, except that—
(1) The required fuel flow rate mustbe established upon the basis of max-imum continuous power and engine ro-tational speed, instead of takeoff powerand fuel consumption; and
(2) If there is a placard providing op-erating instructions, a lesser flow ratemay be used for transferring fuel fromany auxiliary tank into a larger maintank. This lesser flow rate must be ade-quate to maintain engine maximumcontinuous power but the flow ratemust not overfill the main tank atlower engine powers.
(e) Multiple fuel tanks. For recipro-cating engines that are supplied withfuel from more than one tank, if enginepower loss becomes apparent due tofuel depletion from the tank selected,it must be possible after switching toany full tank, in level flight, to obtain75 percent maximum continuous poweron that engine in not more than—
(1) 10 seconds for naturally aspiratedsingle-engine airplanes;
(2) 20 seconds for turbocharged sin-gle-engine airplanes, provided that 75percent maximum continuous natu-rally aspirated power is regained with-in 10 seconds; or
(3) 20 seconds for multiengine air-planes.
(f) Turbine engine fuel systems. Eachturbine engine fuel system must pro-vide at least 100 percent of the fuel flowrequired by the engine under each in-tended operation condition and maneu-ver. The conditions may be simulatedin a suitable mockup. This flow must—
(1) Be shown with the airplane in themost adverse fuel feed condition (withrespect to altitudes, attitudes, andother conditions) that is expected inoperation; and
(2) For multiengine airplanes, not-withstanding the lower flow rate al-lowed by paragraph (d) of this section,be automatically uninterrupted withrespect to any engine until all the fuel
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Federal Aviation Administration, DOT § 23.965
scheduled for use by that engine hasbeen consumed. In addition—
(i) For the purposes of this section,‘‘fuel scheduled for use by that engine’’means all fuel in any tank intended foruse by a specific engine.
(ii) The fuel system design mustclearly indicate the engine for whichfuel in any tank is scheduled.
(iii) Compliance with this paragraphmust require no pilot action after com-pletion of the engine starting phase ofoperations.
(3) For single-engine airplanes, re-quire no pilot action after completionof the engine starting phase of oper-ations unless means are provided thatunmistakenly alert the pilot to takeany needed action at least five minutesprior to the needed action; such pilotaction must not cause any change inengine operation; and such pilot actionmust not distract pilot attention fromessential flight duties during any phaseof operations for which the airplane isapproved.
(a) It must be impossible, in a grav-ity feed system with interconnectedtank outlets, for enough fuel to flowbetween the tanks to cause an overflowof fuel from any tank vent under theconditions in § 23.959, except that fulltanks must be used.
(b) If fuel can be pumped from onetank to another in flight, the fuel tankvents and the fuel transfer systemmust be designed so that no structuraldamage to any airplane component canoccur because of overfilling of anytank.
§ 23.959 Unusable fuel supply.(a) The unusable fuel supply for each
tank must be established as not lessthan that quantity at which the firstevidence of malfunctioning occursunder the most adverse fuel feed condi-tion occurring under each intended op-eration and flight maneuver involving
that tank. Fuel system component fail-ures need not be considered.
(b) The effect on the usable fuelquantity as a result of a failure of anypump shall be determined.
[Amdt. 23–7, 34 FR 13093, Aug. 13, 1969, asamended by Amdt. 23–18, 42 FR 15041, Mar. 17,1977; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996]
§ 23.961 Fuel system hot weather oper-ation.
Each fuel system must be free fromvapor lock when using fuel at its crit-ical temperature, with respect to vaporformation, when operating the airplanein all critical operating and environ-mental conditions for which approvalis requested. For turbine fuel, the ini-tial temperature must be 110 °F, ¥0 °,+5 °F or the maximum outside air tem-perature for which approval is re-quested, whichever is more critical.
§ 23.965 Fuel tank tests.(a) Each fuel tank must be able to
withstand the following pressures with-out failure or leakage:
(1) For each conventional metal tankand nonmetallic tank with walls notsupported by the airplane structure, apressure of 3.5 p.s.i., or that pressure
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14 CFR Ch. I (1–1–01 Edition)§ 23.967
developed during maximum ultimateacceleration with a full tank, which-ever is greater.
(2) For each integral tank, the pres-sure developed during the maximumlimit acceleration of the airplane witha full tank, with simultaneous applica-tion of the critical limit structuralloads.
(3) For each nonmetallic tank withwalls supported by the airplane struc-ture and constructed in an acceptablemanner using acceptable basic tankmaterial, and with actual or simulatedsupport conditions, a pressure of 2 p.s.i.for the first tank of a specific design.The supporting structure must be de-signed for the critical loads occurringin the flight or landing strength condi-tions combined with the fuel pressureloads resulting from the correspondingaccelerations.
(b) Each fuel tank with large, unsup-ported, or unstiffened flat sur-faces,whose failure or deformationcould cause fuel leakage, must be ableto withstand the following test withoutleakage, failure, or excessive deforma-tion of the tank walls:
(1) Each complete tank assembly andits support must be vibration testedwhile mounted to simulate the actualinstallation.
(2) Except as specified in paragraph(b)(4) of this section, the tank assemblymust be vibrated for 25 hours at a totaldisplacement of not less than 1⁄32 of aninch (unless another displacement issubstantiated) while 2⁄3 filled withwater or other suitable test fluid.
(3) The test frequency of vibrationmust be as follows:
(i) If no frequency of vibration result-ing from any rpm within the normaloperating range of engine or propellerspeeds is critical, the test frequency ofvibration is:
(A) The number of cycles per minuteobtained by multiplying the maximumcontinuous propeller speed in rpm by0.9 for propeller-driven airplanes, and
(B) For non-propeller driven air-planes the test frequency of vibrationis 2,000 cycles per minute.
(ii) If only one frequency of vibrationresulting from any rpm within the nor-mal operating range of engine or pro-peller speeds is critical, that frequency
of vibration must be the test fre-quency.
(iii) If more than one frequency of vi-bration resulting from any rpm withinthe normal operating range of engineor propeller speeds is critical, the mostcritical of these frequencies must bethe test frequency.
(4) Under paragraph (b)(3) (ii) and (iii)of this section, the time of test must beadjusted to accomplish the same num-ber of vibration cycles that would beaccomplished in 25 hours at the fre-quency specified in paragraph (b)(3)(i)of this section.
(5) During the test, the tank assem-bly must be rocked at a rate of 16 to 20complete cycles per minute, throughan angle of 15° on either side of the hor-izontal (30° total), about an axis par-allel to the axis of the fuselage, for 25hours.
(c) Each integral tank using methodsof construction and sealing not pre-viously proven to be adequate by testdata or service experience must be ableto withstand the vibration test speci-fied in paragraphs (b)(1) through (4) ofthis section.
(d) Each tank with a nonmetallicliner must be subjected to the sloshingtest outlined in paragraph (b)(5) of thissection, with the fuel at room tempera-ture. In addition, a specimen liner ofthe same basic construction as that tobe used in the airplane must, when in-stalled in a suitable test tank, with-stand the sloshing test with fuel at atemperature of 110° F.
§ 23.967 Fuel tank installation.(a) Each fuel tank must be supported
so that tank loads are not con-centrated. In addition—
(1) There must be pads, if necessary,to prevent chafing between each tankand its supports;
(2) Padding must be nonabsorbent ortreated to prevent the absorption offuel;
(3) If a flexible tank liner is used, itmust be supported so that it is not re-quired to withstand fluid loads;
(4) Interior surfaces adjacent to theliner must be smooth and free from
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Federal Aviation Administration, DOT § 23.971
projections that could cause wear, un-less—
(i) Provisions are made for protectionof the liner at those points; or
(ii) The construction of the lineritself provides such protection; and
(5) A positive pressure must be main-tained within the vapor space of eachbladder cell under any condition of op-eration, except for a particular condi-tion for which it is shown that a zeroor negative pressure will not cause thebladder cell to collapse; and
(6) Syphoning of fuel (other thanminor spillage) or collapse of bladderfuel cells may not result from impropersecuring or loss of the fuel filler cap.
(b) Each tank compartment must beventilated and drained to prevent theaccumulation of flammable fluids orvapors. Each compartment adjacent toa tank that is an integral part of theairplane structure must also be venti-lated and drained.
(c) No fuel tank may be on the engineside of the firewall. There must be atleast one-half inch of clearance be-tween the fuel tank and the firewall.No part of the engine nacelle skin thatlies immediately behind a major airopening from the engine compartmentmay act as the wall of an integraltank.
(d) Each fuel tank must be isolatedfrom personnel compartments by afume-proof and fuel-proof enclosurethat is vented and drained to the exte-rior of the airplane. The required en-closure must sustain any personnelcompartment pressurization loadswithout permanent deformation or fail-ure under the conditions of §§ 23.365 and23.843 of this part. A bladder-type fuelcell, if used, must have a retainingshell at least equivalent to a metal fueltank in structural integrity.
(e) Fuel tanks must be designed, lo-cated, and installed so as to retain fuel:
(1) When subjected to the inertialoads resulting from the ultimate stat-ic load factors prescribed in§ 23.561(b)(2) of this part; and
(2) Under conditions likely to occurwhen the airplane lands on a pavedrunway at a normal landing speedunder each of the following conditions:
(i) The airplane in a normal landingattitude and its landing gear retracted.
(ii) The most critical landing gear legcollapsed and the other landing gearlegs extended.
In showing compliance with paragraph(e)(2) of this section, the tearing awayof an engine mount must be consideredunless all the engines are installedabove the wing or on the tail or fuse-lage of the airplane.
Each fuel tank must have an expan-sion space of not less than two percentof the tank capacity, unless the tankvent discharges clear of the airplane(in which case no expansion space is re-quired). It must be impossible to fillthe expansion space inadvertently withthe airplane in the normal ground atti-tude.
§ 23.971 Fuel tank sump.
(a) Each fuel tank must have a drain-able sump with an effective capacity,in the normal ground and flight atti-tudes, of 0.25 percent of the tank capac-ity, or 1⁄16 gallon, whichever is greater.
(b) Each fuel tank must allow drain-age of any hazardous quantity of waterfrom any part of the tank to its sumpwith the airplane in the normal groundattitude.
(c) Each reciprocating engine fuelsystem must have a sediment bowl orchamber that is accessible for drain-age; has a capacity of 1 ounce for every20 gallons of fuel tank capacity; andeach fuel tank outlet is located so that,in the normal flight attitude, waterwill drain from all parts of the tank ex-cept the sump to the sediment bowl orchamber.
(d) Each sump, sediment bowl, andsediment chamber drain required byparagraphs (a), (b), and (c) of this sec-tion must comply with the drain provi-sions of § 23.999(b)(1) and (b)(2).
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14 CFR Ch. I (1–1–01 Edition)§ 23.973
§ 23.973 Fuel tank filler connection.(a) Each fuel tank filler connection
must be marked as prescribed in§ 23.1557(c).
(b) Spilled fuel must be preventedfrom entering the fuel tank compart-ment or any part of the airplane otherthan the tank itself.
(c) Each filler cap must provide afuel-tight seal for the main filler open-ing. However, there may be small open-ings in the fuel tank cap for ventingpurposes or for the purpose of allowingpassage of a fuel gauge through the capprovided such openings comply withthe requirements of § 23.975(a).
(d) Each fuel filling point, exceptpressure fueling connection points,must have a provision for electricallybonding the airplane to ground fuelingequipment.
(e) For airplanes with engines requir-ing gasoline as the only permissiblefuel, the inside diameter of the fuelfiller opening must be no larger than2.36 inches.
(f) For airplanes with turbine en-gines, the inside diameter of the fuelfiller opening must be no smaller than2.95 inches.
§ 23.975 Fuel tank vents and carbu-retor vapor vents.
(a) Each fuel tank must be ventedfrom the top part of the expansionspace. In addition—
(1) Each vent outlet must be locatedand constructed in a manner that mini-mizes the possibility of its being ob-structed by ice or other foreign matter;
(2) Each vent must be constructed toprevent siphoning of fuel during nor-mal operation;
(3) The venting capacity must allowthe rapid relief of excessive differencesof pressure between the interior andexterior of the tank;
(4) Airspaces of tanks with inter-connected outlets must be inter-connected;
(5) There may be no point in any ventline where moisture can accumulatewith the airplane in either the groundor level flight attitudes, unless drain-
age is provided. Any drain valve in-stalled must be accessible for drainage;
(6) No vent may terminate at a pointwhere the discharge of fuel from thevent outlet will constitute a fire haz-ard or from which fumes may enterpersonnel compartments; and
(7) Vents must be arranged to pre-vent the loss of fuel, except fuel dis-charged because of thermal expansion,when the airplane is parked in any di-rection on a ramp having a one-percentslope.
(b) Each carburetor with vapor elimi-nation connections and each fuel injec-tion engine employing vapor returnprovisions must have a separate ventline to lead vapors back to the top ofone of the fuel tanks. If there is morethan one tank and it is necessary touse these tanks in a definite sequencefor any reason, the vapor vent linemust lead back to the fuel tank to beused first, unless the relative capac-ities of the tanks are such that returnto another tank is preferable.
(c) For acrobatic category airplanes,excessive loss of fuel during acrobaticmaneuvers, including short periods ofinverted flight, must be prevented. Itmust be impossible for fuel to siphonfrom the vent when normal flight hasbeen resumed after any acrobatic ma-neuver for which certification is re-quested.
For pressure fueling systems, the fol-lowing apply:
(a) Each pressure fueling system fuelmanifold connection must have meansto prevent the escape of hazardousquantities of fuel from the system ifthe fuel entry valve fails.
(b) An automatic shutoff means mustbe provided to prevent the quantity offuel in each tank from exceeding themaximum quantity approved for thattank. This means must—
(1) Allow checking for proper shutoffoperation before each fueling of thetank; and
(2) For commuter category airplanes,indicate at each fueling station, a fail-ure of the shutoff means to stop thefuel flow at the maximum quantity ap-proved for that tank.
(c) A means must be provided to pre-vent damage to the fuel system in theevent of failure of the automatic shut-off means prescribed in paragraph (b)of this section.
(d) All parts of the fuel system up tothe tank which are subjected to fuelingpressures must have a proof pressure of1.33 times, and an ultimate pressure ofat least 2.0 times, the surge pressurelikely to occur during fueling.
§ 23.991 Fuel pumps.(a) Main pumps. For main pumps, the
following apply:(1) For reciprocating engine installa-
tions having fuel pumps to supply fuelto the engine, at least one pump foreach engine must be directly driven bythe engine and must meet § 23.955. Thispump is a main pump.
(2) For turbine engine installations,each fuel pump required for proper en-gine operation, or required to meet thefuel system requirements of this sub-part (other than those in paragraph (b)
of this section), is a main pump. In ad-dition—
(i) There must be at least one mainpump for each turbine engine;
(ii) The power supply for the mainpump for each engine must be inde-pendent of the power supply for eachmain pump for any other engine; and
(iii) For each main pump, provisionmust be made to allow the bypass ofeach positive displacement fuel pumpother than a fuel injection pump ap-proved as part of the engine.
(b) Emergency pumps. There must bean emergency pump immediately avail-able to supply fuel to the engine if anymain pump (other than a fuel injectionpump approved as part of an engine)fails. The power supply for each emer-gency pump must be independent of thepower supply for each correspondingmain pump.
(c) Warning means. If both the mainpump and emergency pump operatecontinuously, there must be a means toindicate to the appropriate flight crew-members a malfunction of either pump.
(d) Operation of any fuel pump maynot affect engine operation so as tocreate a hazard, regardless of the en-gine power or thrust setting or thefunctional status of any other fuelpump.
(a) Each fuel line must be installedand supported to prevent excessive vi-bration and to withstand loads due tofuel pressure and accelerated flightconditions.
(b) Each fuel line connected to com-ponents of the airplane between whichrelative motion could exist must haveprovisions for flexibility.
(c) Each flexible connection in fuellines that may be under pressure andsubjected to axial loading must useflexible hose assemblies.
(d) Each flexible hose must be shownto be suitable for the particular appli-cation.
(e) No flexible hose that might be ad-versely affected by exposure to high
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14 CFR Ch. I (1–1–01 Edition)§ 23.994
temperatures may be used where exces-sive temperatures will exist during op-eration or after engine shutdown.
Fuel system components in an enginenacelle or in the fuselage must be pro-tected from damage which could resultin spillage of enough fuel to constitutea fire hazard as a result of a wheels-uplanding on a paved runway.
[Amdt. 23–29, 49 FR 6847, Feb. 23, 1984]
§ 23.995 Fuel valves and controls.
(a) There must be a means to allowappropriate flight crew members torapidly shut off, in flight, the fuel toeach engine individually.
(b) No shutoff valve may be on theengine side of any firewall. In addition,there must be means to—
(1) Guard against inadvertent oper-ation of each shutoff valve; and
(2) Allow appropriate flight crewmembers to reopen each valve rapidlyafter it has been closed.
(c) Each valve and fuel system con-trol must be supported so that loads re-sulting from its operation or from ac-celerated flight conditions are nottransmitted to the lines connected tothe valve.
(d) Each valve and fuel system con-trol must be installed so that gravityand vibration will not affect the se-lected position.
(e) Each fuel valve handle and itsconnections to the valve mechanismmust have design features that mini-mize the possibility of incorrect instal-lation.
(f) Each check valve must be con-structed, or otherwise incorporate pro-visions, to preclude incorrect assemblyor connection of the valve.
(g) Fuel tank selector valves must—(1) Require a separate and distinct
action to place the selector in the‘‘OFF’’ position; and
(2) Have the tank selector positionslocated in such a manner that it is im-possible for the selector to pass
through the ‘‘OFF’’ position whenchanging from one tank to another.
There must be a fuel strainer or filterbetween the fuel tank outlet and theinlet of either the fuel metering deviceor an engine driven positive displace-ment pump, whichever is nearer thefuel tank outlet. This fuel strainer orfilter must—
(a) Be accessible for draining andcleaning and must incorporate a screenor element which is easily removable;
(b) Have a sediment trap and drainexcept that it need not have a drain ifthe strainer or filter is easily remov-able for drain purposes;
(c) Be mounted so that its weight isnot supported by the connecting linesor by the inlet or outlet connections ofthe strainer or filter itself, unless ade-quate strength margins under all load-ing conditions are provided in the linesand connections; and
(d) Have the capacity (with respect tooperating limitations established forthe engine) to ensure that engine fuelsystem functioning is not impaired,with the fuel contaminated to a degree(with respect to particle size and den-sity) that is greater than that estab-lished for the engine during its typecertification.
(e) In addition, for commuter cat-egory airplanes, unless means are pro-vided in the fuel system to prevent theaccumulation of ice on the filter, ameans must be provided to automati-cally maintain the fuel flow if ice clog-ging of the filter occurs.
§ 23.1001 Fuel jettisoning system.(a) If the design landing weight is
less than that permitted under the re-quirements of § 23.473(b), the airplanemust have a fuel jettisoning system in-stalled that is able to jettison enoughfuel to bring the maximum weightdown to the design landing weight. Theaverage rate of fuel jettisoning must beat least 1 percent of the maximumweight per minute, except that thetime required to jettison the fuel neednot be less than 10 minutes.
(b) Fuel jettisoning must be dem-onstrated at maximum weight withflaps and landing gear up and in—
(1) A power-off glide at 1.4 VS1;(2) A climb, at the speed at which the
one-engine-inoperative enroute climbdata have been established in accord-ance with § 23.69(b), with the criticalengine inoperative and the remainingengines at maximum continuouspower; and
(3) Level flight at 1.4 VS1, if the re-sults of the tests in the conditionsspecified in paragraphs (b)(1) and (2) ofthis section show that this conditioncould be critical.
(c) During the flight tests prescribedin paragraph (b) of this section, it mustbe shown that—
(1) The fuel jettisoning system andits operation are free from fire hazard;
(2) The fuel discharges clear of anypart of the airplane;
(3) Fuel or fumes do not enter anyparts of the airplane; and
(4) The jettisoning operation does notadversely affect the controllability ofthe airplane.
(d) For reciprocating engine poweredairplanes, the jettisoning system mustbe designed so that it is not possible tojettison the fuel in the tanks used fortakeoff and landing below the level al-lowing 45 minutes flight at 75 percentmaximum continuous power. However,if there is an auxiliary control inde-pendent of the main jettisoning con-trol, the system may be designed tojettison all the fuel.
(e) For turbine engine powered air-planes, the jettisoning system must bedesigned so that it is not possible tojettison fuel in the tanks used fortakeoff and landing below the level al-lowing climb from sea level to 10,000feet and thereafter allowing 45 minutescruise at a speed for maximum range.
(f) The fuel jettisoning valve must bedesigned to allow flight crewmembersto close the valve during any part ofthe jettisoning operation.
(g) Unless it is shown that using anymeans (including flaps, slots, and slats)for changing the airflow across oraround the wings does not adversely af-fect fuel jettisoning, there must be aplacard, adjacent to the jettisoningcontrol, to warn flight crewmembersagainst jettisoning fuel while themeans that change the airflow arebeing used.
(h) The fuel jettisoning system mustbe designed so that any reasonablyprobable single malfunction in the sys-tem will not result in a hazardous con-dition due to unsymmetrical jetti-soning of, or inability to jettison, fuel.
(a) For oil systems and componentsthat have been approved under the en-gine airworthiness requirements andwhere those requirements are equal toor more severe than the correspondingrequirements of subpart E of this part,that approval need not be duplicated.Where the requirements of subpart E of
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14 CFR Ch. I (1–1–01 Edition)§ 23.1013
this part are more severe, substan-tiation must be shown to the require-ments of subpart E of this part.
(b) Each engine must have an inde-pendent oil system that can supply itwith an appropriate quantity of oil at atemperature not above that safe forcontinuous operation.
(c) The usable oil tank capacity maynot be less than the product of the en-durance of the airplane under criticaloperating conditions and the maximumoil consumption of the engine underthe same conditions, plus a suitablemargin to ensure adequate circulationand cooling.
(d) For an oil system without an oiltransfer system, only the usable oiltank capacity may be considered. Theamount of oil in the engine oil lines,the oil radiator, and the feathering re-serve, may not be considered.
(e) If an oil transfer system is used,and the transfer pump can pump someof the oil in the transfer lines into themain engine oil tanks, the amount ofoil in these lines that can be pumpedby the transfer pump may be includedin the oil capacity.
§ 23.1013 Oil tanks.(a) Installation. Each oil tank must be
installed to—(1) Meet the requirements of § 23.967
(a) and (b); and(2) Withstand any vibration, inertia,
and fluid loads expected in operation.(b) Expansion space. Oil tank expan-
sion space must be provided so that—(1) Each oil tank used with a recipro-
cating engine has an expansion space ofnot less than the greater of 10 percentof the tank capacity or 0.5 gallon, andeach oil tank used with a turbine en-gine has an expansion space of not lessthan 10 percent of the tank capacity;and
(2) It is impossible to fill the expan-sion space inadvertently with the air-plane in the normal ground attitude.
(c) Filler connection. Each oil tankfiller connection must be marked asspecified in § 23.1557(c). Each recessedoil tank filler connection of an oil tankused with a turbine engine, that can re-tain any appreciable quantity of oil,
must have provisions for fitting adrain.
(d) Vent. Oil tanks must be vented asfollows:
(1) Each oil tank must be vented tothe engine from the top part of the ex-pansion space so that the vent connec-tion is not covered by oil under anynormal flight condition.
(2) Oil tank vents must be arrangedso that condensed water vapor thatmight freeze and obstruct the line can-not accumulate at any point.
(3) For acrobatic category airplanes,there must be means to prevent haz-ardous loss of oil during acrobatic ma-neuvers, including short periods of in-verted flight.
(e) Outlet. No oil tank outlet may beenclosed by any screen or guard thatwould reduce the flow of oil below asafe value at any operating tempera-ture. No oil tank outlet diameter maybe less than the diameter of the engineoil pump inlet. Each oil tank used witha turbine engine must have means toprevent entrance into the tank itself,or into the tank outlet, of any objectthat might obstruct the flow of oilthrough the system. There must be ashutoff valve at the outlet of each oiltank used with a turbine engine, unlessthe external portion of the oil system(including oil tank supports) is fire-proof.
(f) Flexible liners. Each flexible oiltank liner must be of an acceptablekind.
(g) Each oil tank filler cap of an oiltank that is used with an engine mustprovide an oiltight seal.
Each oil tank must be tested under§ 23.965, except that—
(a) The applied pressure must be fivep.s.i. for the tank construction insteadof the pressures specified in § 23.965(a);
(b) For a tank with a nonmetallicliner the test fluid must be oil ratherthan fuel as specified in § 23.965(d), andthe slosh test on a specimen liner mustbe conducted with the oil at 250° F.;and
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Federal Aviation Administration, DOT § 23.1027
(c) For pressurized tanks used with aturbine engine, the test pressure maynot be less than 5 p.s.i. plus the max-imum operating pressure of the tank.
§ 23.1017 Oil lines and fittings.(a) Oil lines. Oil lines must meet
§ 23.993 and must accommodate a flowof oil at a rate and pressure adequatefor proper engine functioning underany normal operating condition.
(b) Breather lines. Breather lines mustbe arranged so that—
(1) Condensed water vapor or oil thatmight freeze and obstruct the line can-not accumulate at any point;
(2) The breather discharge will notconstitute a fire hazard if foaming oc-curs, or cause emitted oil to strike thepilot’s windshield;
(3) The breather does not dischargeinto the engine air induction system;and
(4) For acrobatic category airplanes,there is no excessive loss of oil fromthe breather during acrobatic maneu-vers, including short periods of in-verted flight.
(5) The breather outlet is protectedagainst blockage by ice or foreign mat-ter.
§ 23.1019 Oil strainer or filter.(a) Each turbine engine installation
must incorporate an oil strainer or fil-ter through which all of the engine oilflows and which meets the following re-quirements:
(1) Each oil strainer or filter that hasa bypass, must be constructed and in-stalled so that oil will flow at the nor-mal rate through the rest of the sys-tem with the strainer or filter com-pletely blocked.
(2) The oil strainer or filter musthave the capacity (with respect to op-erating limitations established for theengine) to ensure that engine oil sys-tem functioning is not impaired whenthe oil is contaminated to a degree(with respect to particle size and den-sity) that is greater than that estab-
lished for the engine for its type cer-tification.
(3) The oil strainer or filter, unless itis installed at an oil tank outlet, mustincorporate a means to indicate con-tamination before it reaches the capac-ity established in accordance withparagraph (a)(2) of this section.
(4) The bypass of a strainer or filtermust be constructed and installed sothat the release of collected contami-nants is minimized by appropriate lo-cation of the bypass to ensure that col-lected contaminants are not in the by-pass flow path.
(5) An oil strainer or filter that hasno bypass, except one that is installedat an oil tank outlet, must have ameans to connect it to the warningsystem required in § 23.1305(c)(9).
(b) Each oil strainer or filter in apowerplant installation using recipro-cating engines must be constructed andinstalled so that oil will flow at thenormal rate through the rest of thesystem with the strainer or filter ele-ment completely blocked.
The powerplant and auxiliary powerunit cooling provisions must maintainthe temperatures of powerplant compo-nents and engine fluids, and auxiliarypower unit components and fluids with-in the limits established for those com-ponents and fluids under the most ad-verse ground, water, and flight oper-ations to the maximum altitude andmaximum ambient atmospheric tem-perature conditions for which approvalis requested, and after normal engineand auxiliary power unit shutdown.
§ 23.1043 Cooling tests.(a) General. Compliance with § 23.1041
must be shown on the basis of tests, forwhich the following apply:
(1) If the tests are conducted underambient atmospheric temperature con-ditions deviating from the maximumfor which approval is requested, the re-corded powerplant temperatures mustbe corrected under paragraphs (c) and(d) of this section, unless a more ra-tional correction method is applicable.
(2) No corrected temperature deter-mined under paragraph (a)(1) of thissection may exceed established limits.
(3) The fuel used during the coolingtests must be of the minimum gradeapproved for the engine.
(4) For turbocharged engines, eachturbocharger must be operated throughthat part of the climb profile for which
operation with the turbocharger is re-quested.
(5) For a reciprocating engine, themixture settings must be the leanestrecommended for climb.
(b) Maximum ambient atmospheric tem-perature. A maximum ambient atmos-pheric temperature corresponding tosea level conditions of at least 100 de-grees F must be established. The as-sumed temperature lapse rate is 3.6 de-grees F per thousand feet of altitudeabove sea level until a temperature of¥69.7 degrees F is reached, above whichaltitude the temperature is consideredconstant at ¥69.7 degrees F. However,for winterization installations, the ap-plicant may select a maximum ambi-ent atmospheric temperature cor-responding to sea level conditions ofless than 100 degrees F.
(c) Correction factor (except cylinderbarrels). Temperatures of engine fluidsand powerplant components (exceptcylinder barrels) for which temperaturelimits are established, must be cor-rected by adding to them the differencebetween the maximum ambient atmos-pheric temperature for the relevant al-titude for which approval has been re-quested and the temperature of the am-bient air at the time of the first occur-rence of the maximum fluid or compo-nent temperature recorded during thecooling test.
(d) Correction factor for cylinder barreltemperatures. Cylinder barrel tempera-tures must be corrected by adding tothem 0.7 times the difference betweenthe maximum ambient atmospherictemperature for the relevant altitudefor which approval has been requestedand the temperature of the ambient airat the time of the first occurrence ofthe maximum cylinder barrel tempera-ture recorded during the cooling test.
§ 23.1045 Cooling test procedures forturbine engine powered airplanes.
(a) Compliance with § 23.1041 must beshown for all phases of operation. Theairplane must be flown in the configu-rations, at the speeds, and followingthe procedures recommended in the
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Federal Aviation Administration, DOT § 23.1061
Airplane Flight Manual for the rel-evant stage of flight, that correspondto the applicable performance require-ments that are critical to cooling.
(b) Temperatures must be stabilizedunder the conditions from which entryis made into each stage of flight beinginvestigated, unless the entry condi-tion normally is not one during whichcomponent and engine fluid tempera-tures would stabilize (in which case,operation through the full entry condi-tion must be conducted before entryinto the stage of flight being inves-tigated in order to allow temperaturesto reach their natural levels at thetime of entry). The takeoff cooling testmust be preceded by a period duringwhich the powerplant component andengine fluid temperatures are sta-bilized with the engines at ground idle.
(c) Cooling tests for each stage offlight must be continued until—
(1) The component and engine fluidtemperatures stabilize;
(2) The stage of flight is completed;or
(3) An operating limitation isreached.
[Amdt. 23–7, 34 FR 13094, Aug. 13, 1969, asamended by Amdt. 23–51, 61 FR 5137, Feb. 9,1996]
§ 23.1047 Cooling test procedures forreciprocating engine powered air-planes.
Compliance with § 23.1041 must beshown for the climb (or, for multien-gine airplanes with negative one-en-gine-inoperative rates of climb, the de-scent) stage of flight. The airplanemust be flown in the configurations, atthe speeds and following the proceduresrecommended in the Airplane FlightManual, that correspond to the appli-cable performance requirements thatare critical to cooling.
[Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
LIQUID COOLING
§ 23.1061 Installation.(a) General. Each liquid-cooled engine
must have an independent cooling sys-tem (including coolant tank) installedso that—
(1) Each coolant tank is supported sothat tank loads are distributed over alarge part of the tank surface;
(2) There are pads or other isolationmeans between the tank and its sup-ports to prevent chafing.
(3) Pads or any other isolation meansthat is used must be nonabsorbent ormust be treated to prevent absorptionof flammable fluids; and
(4) No air or vapor can be trapped inany part of the system, except thecoolant tank expansion space, duringfilling or during operation.
(b) Coolant tank. The tank capacitymust be at least one gallon, plus 10 per-cent of the cooling system capacity. Inaddition—
(1) Each coolant tank must be able towithstand the vibration, inertia, andfluid loads to which it may be sub-jected in operation;
(2) Each coolant tank must have anexpansion space of at least 10 percentof the total cooling system capacity;and
(3) It must be impossible to fill theexpansion space inadvertently with theairplane in the normal ground attitude.
(c) Filler connection. Each coolanttank filler connection must be markedas specified in § 23.1557(c). In addition—
(1) Spilled coolant must be preventedfrom entering the coolant tank com-partment or any part of the airplaneother than the tank itself; and
(2) Each recessed coolant filler con-nection must have a drain that dis-charges clear of the entire airplane.
(d) Lines and fittings. Each coolantsystem line and fitting must meet therequirements of § 23.993, except that theinside diameter of the engine coolantinlet and outlet lines may not be lessthan the diameter of the correspondingengine inlet and outlet connections.
(e) Radiators. Each coolant radiatormust be able to withstand any vibra-tion, inertia, and coolant pressure loadto which it may normally be subjected.In addition—
(1) Each radiator must be supportedto allow expansion due to operatingtemperatures and prevent the trans-mittal of harmful vibration to the radi-ator; and
(2) If flammable coolant is used, theair intake duct to the coolant radiatormust be located so that (in case of fire)flames from the nacelle cannot strikethe radiator.
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14 CFR Ch. I (1–1–01 Edition)§ 23.1063
(f) Drains. There must be an acces-sible drain that—
(1) Drains the entire cooling system(including the coolant tank, radiator,and the engine) when the airplane is inthe normal ground altitude;
Each coolant tank must be testedunder § 23.965, except that—
(a) The test required by § 23.965(a)(1)must be replaced with a similar testusing the sum of the pressure devel-oped during the maximum ultimate ac-celeration with a full tank or a pres-sure of 3.5 pounds per square inch,whichever is greater, plus the max-imum working pressure of the system;and
(b) For a tank with a nonmetallicliner the test fluid must be coolantrather than fuel as specified in§ 23.965(d), and the slosh test on a speci-men liner must be conducted with thecoolant at operating temperature.
INDUCTION SYSTEM
§ 23.1091 Air induction system.
(a) The air induction system for eachengine and auxiliary power unit andtheir accessories must supply the airrequired by that engine and auxiliarypower unit and their accessories underthe operating conditions for which cer-tification is requested.
(b) Each reciprocating engine instal-lation must have at least two separateair intake sources and must meet thefollowing:
(1) Primary air intakes may openwithin the cowling if that part of thecowling is isolated from the engine ac-cessory section by a fire-resistant dia-phragm or if there are means to pre-vent the emergence of backfire flames.
(2) Each alternate air intake must belocated in a sheltered position and maynot open within the cowling if theemergence of backfire flames will re-sult in a hazard.
(3) The supplying of air to the enginethrough the alternate air intake sys-tem may not result in a loss of exces-sive power in addition to the power lossdue to the rise in air temperature.
(4) Each automatic alternate air doormust have an override means acces-sible to the flight crew.
(5) Each automatic alternate air doormust have a means to indicate to theflight crew when it is not closed.
(c) For turbine engine powered air-planes—
(1) There must be means to preventhazardous quantities of fuel leakage oroverflow from drains, vents, or othercomponents of flammable fluid systemsfrom entering the engine intake sys-tem; and
(2) The airplane must be designed toprevent water or slush on the runway,taxiway, or other airport operatingsurfaces from being directed into theengine or auxiliary power unit air in-take ducts in hazardous quantities.The air intake ducts must be located orprotected so as to minimize the hazardof ingestion of foreign matter duringtakeoff, landing, and taxiing.
(a) Reciprocating engines. Each recip-rocating engine air induction systemmust have means to prevent and elimi-nate icing. Unless this is done by othermeans, it must be shown that, in airfree of visible moisture at a tempera-ture of 30° F.—
(1) Each airplane with sea level en-gines using conventional venturi car-buretors has a preheater that can pro-vide a heat rise of 90° F. with the en-gines at 75 percent of maximum contin-uous power;
(2) Each airplane with altitude en-gines using conventional venturi car-buretors has a preheater that can pro-vide a heat rise of 120° F. with the en-gines at 75 percent of maximum contin-uous power;
(3) Each airplane with altitude en-gines using fuel metering device tend-ing to prevent icing has a preheater
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Federal Aviation Administration, DOT § 23.1097
that, with the engines at 60 percent ofmaximum continuous power, can pro-vide a heat rise of—
(i) 100° F.; or(ii) 40° F., if a fluid deicing system
meeting the requirements of §§ 23.1095through 23.1099 is installed;
(4) Each airplane with sea level en-gine(s) using fuel metering device tend-ing to prevent icing has a sheltered al-ternate source of air with a preheat ofnot less than 60 °F with the engines at75 percent of maximum continuouspower;
(5) Each airplane with sea level or al-titude engine(s) using fuel injectionsystems having metering componentson which impact ice may accumulatehas a preheater capable of providing aheat rise of 75 °F when the engine is op-erating at 75 percent of its maximumcontinuous power; and
(6) Each airplane with sea level or al-titude engine(s) using fuel injectionsystems not having fuel metering com-ponents projecting into the airstreamon which ice may form, and intro-ducing fuel into the air induction sys-tem downstream of any components orother obstruction on which ice pro-duced by fuel evaporation may form,has a sheltered alternate source of airwith a preheat of not less than 60 °Fwith the engines at 75 percent of itsmaximum continuous power.
(b) Turbine engines. (1) Each turbineengine and its air inlet system mustoperate throughout the flight powerrange of the engine (including idling),without the accumulation of ice on en-gine or inlet system components thatwould adversely affect engine oper-ation or cause a serious loss of poweror thrust—
(i) Under the icing conditions speci-fied in appendix C of part 25 of thischapter; and
(ii) In snow, both falling and blowing,within the limitations established forthe airplane for such operation.
(2) Each turbine engine must idle for30 minutes on the ground, with the airbleed available for engine icing protec-tion at its critical condition, withoutadverse effect, in an atmosphere that isat a temperature between 15° and 30° F(between ¥9° and ¥1° C) and has a liq-uid water content not less than 0.3grams per cubic meter in the form of
drops having a mean effective diameternot less than 20 microns, followed bymomentary operation at takeoff poweror thrust. During the 30 minutes of idleoperation, the engine may be run upperiodically to a moderate power orthrust setting in a manner acceptableto the Administrator.
(c) Reciprocating engines with Super-chargers. For airplanes with recipro-cating engines having superchargers topressurize the air before it enters thefuel metering device, the heat rise inthe air caused by that supercharging atany altitude may be utilized in deter-mining compliance with paragraph (a)of this section if the heat rise utilizedis that which will be available, auto-matically, for the applicable altitudesand operating condition because of su-percharging.
(a) If a carburetor deicing fluid sys-tem is used, it must be able to simulta-neously supply each engine with a rateof fluid flow, expressed in pounds perhour, of not less than 2.5 times thesquare root of the maximum contin-uous power of the engine.
(b) The fluid must be introduced intothe air induction system—
(1) Close to, and upstream of, the car-buretor; and
(2) So that it is equally distributedover the entire cross section of the in-duction system air passages.
(a) The capacity of each carburetordeicing fluid system—
(1) May not be less than the greaterof—
(i) That required to provide fluid atthe rate specified in § 23.1095 for a timeequal to three percent of the maximumendurance of the airplane; or
(ii) 20 minutes at that flow rate; and(2) Need not exceed that required for
two hours of operation.
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14 CFR Ch. I (1–1–01 Edition)§ 23.1099
(b) If the available preheat exceeds50° F. but is less than 100° F., the ca-pacity of the system may be decreasedin proportion to the heat rise availablein excess of 50° F.
Each carburetor deicing fluid systemmust meet the applicable requirementsfor the design of a fuel system, exceptas specified in §§ 23.1095 and 23.1097.
§ 23.1101 Induction air preheater de-sign.
Each exhaust-heated, induction airpreheater must be designed and con-structed to—
(a) Ensure ventilation of the pre-heater when the induction air pre-heater is not being used during engineoperation;
(b) Allow inspection of the exhaustmanifold parts that it surrounds; and
(c) Allow inspection of critical partsof the preheater itself.
§ 23.1103 Induction system ducts.(a) Each induction system duct must
have a drain to prevent the accumula-tion of fuel or moisture in the normalground and flight attitudes. No drainmay discharge where it will cause afire hazard.
(b) Each duct connected to compo-nents between which relative motioncould exist must have means for flexi-bility.
(c) Each flexible induction systemduct must be capable of withstandingthe effects of temperature extremes,fuel, oil, water, and solvents to whichit is expected to be exposed in serviceand maintenance without hazardousdeterioration or delamination.
(d) For reciprocating engine installa-tions, each induction system duct mustbe—
(1) Strong enough to prevent induc-tion system failures resulting fromnormal backfire conditions; and
(2) Fire resistant in any compart-ment for which a fire extinguishingsystem is required.
(e) Each inlet system duct for an aux-iliary power unit must be—
(1) Fireproof within the auxiliarypower unit compartment;
(2) Fireproof for a sufficient distanceupstream of the auxiliary power unitcompartment to prevent hot gas re-verse flow from burning through theduct and entering any other compart-ment of the airplane in which a hazardwould be created by the entry of thehot gases;
(3) Constructed of materials suitableto the environmental conditions ex-pected in service, except in those areasrequiring fireproof or fire resistant ma-terials; and
(4) Constructed of materials that willnot absorb or trap hazardous quantitiesof flammable fluids that could be ig-nited by a surge or reverse-flow condi-tion.
(f) Induction system ducts that sup-ply air to a cabin pressurization sys-tem must be suitably constructed ofmaterial that will not produce haz-ardous quantities of toxic gases or iso-lated to prevent hazardous quantitiesof toxic gases from entering the cabinduring a powerplant fire.
If an air filter is used to protect theengine against foreign material par-ticles in the induction air supply—
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Federal Aviation Administration, DOT § 23.1123
(a) Each air filter must be capable ofwithstanding the effects of tempera-ture extremes, rain, fuel, oil, and sol-vents to which it is expected to be ex-posed in service and maintenance; and
(b) Each air filter shall have a designfeature to prevent material separatedfrom the filter media from interferingwith proper fuel metering operation.
The following applies toturbocharged bleed air systems usedfor cabin pressurization:
(a) The cabin air system may not besubject to hazardous contaminationfollowing any probable failure of theturbocharger or its lubrication system.
(b) The turbocharger supply air mustbe taken from a source where it cannotbe contaminated by harmful or haz-ardous gases or vapors following anyprobable failure or malfunction of theengine exhaust, hydraulic, fuel, or oilsystem.
[Amdt. 23–42, 56 FR 354, Jan. 3, 1991]
§ 23.1111 Turbine engine bleed air sys-tem.
For turbine engine bleed air systems,the following apply:
(a) No hazard may result if duct rup-ture or failure occurs anywhere be-tween the engine port and the airplaneunit served by the bleed air.
(b) The effect on airplane and engineperformance of using maximum bleedair must be established.
(c) Hazardous contamination of cabinair systems may not result from fail-ures of the engine lubricating system.
[Amdt. 23–7, 34 FR 13095, Aug. 13, 1969, asamended by Amdt. 23–17, 41 FR 55465, Dec. 20,1976]
EXHAUST SYSTEM
§ 23.1121 General.For powerplant and auxiliary power
unit installations, the followingapply—
(a) Each exhaust system must ensuresafe disposal of exhaust gases withoutfire hazard or carbon monoxide con-
tamination in any personnel compart-ment.
(b) Each exhaust system part with asurface hot enough to ignite flammablefluids or vapors must be located orshielded so that leakage from any sys-tem carrying flammable fluids or va-pors will not result in a fire caused byimpingement of the fluids or vapors onany part of the exhaust system includ-ing shields for the exhaust system.
(c) Each exhaust system must be sep-arated by fireproof shields from adja-cent flammable parts of the airplanethat are outside of the engine and aux-iliary power unit compartments.
(d) No exhaust gases may dischargedangerously near any fuel or oil systemdrain.
(e) No exhaust gases may be dis-charged where they will cause a glareseriously affecting pilot vision atnight.
(f) Each exhaust system componentmust be ventilated to prevent points ofexcessively high temperature.
(g) If significant traps exist, eachturbine engine and auxiliary powerunit exhaust system must have drainsdischarging clear of the airplane, inany normal ground and flight attitude,to prevent fuel accumulation after thefailure of an attempted engine or auxil-iary power unit start.
(h) Each exhaust heat exchangermust incorporate means to preventblockage of the exhaust port after anyinternal heat exchanger failure.
(i) For the purpose of compliancewith § 23.603, the failure of any part ofthe exhaust system will be consideredto adversely affect safety.
(a) Powerplant controls must be lo-cated and arranged under § 23.777 andmarked under § 23.1555(a).
(b) Each flexible control must beshown to be suitable for the particularapplication.
(c) Each control must be able tomaintain any necessary position with-out—
(1) Constant attention by flight crewmembers; or
(2) Tendency to creep due to controlloads or vibration.
(d) Each control must be able towithstand operating loads without fail-ure or excessive deflection.
(e) For turbine engine powered air-planes, no single failure or malfunc-tion, or probable combination thereof,in any powerplant control system maycause the failure of any powerplantfunction necessary for safety.
(f) The portion of each powerplantcontrol located in the engine compart-ment that is required to be operated inthe event of fire must be at least fireresistant.
(g) Powerplant valve controls locatedin the cockpit must have—
(1) For manual valves, positive stopsor in the case of fuel valves suitableindex provisions, in the open and closedposition; and
(2) For power-assisted valves, ameans to indicate to the flight crewwhen the valve—
(i) Is in the fully open or fully closedposition; or
(ii) Is moving between the fully openand fully closed position.
Means must be provided on the flightdeck for the starting, stopping, moni-toring, and emergency shutdown ofeach installed auxiliary power unit.
[Doc. No. 26344, 58 FR 18974, Apr. 9, 1993]
§ 23.1143 Engine controls.(a) There must be a separate power or
thrust control for each engine and aseparate control for each superchargerthat requires a control.
(b) Power, thrust, and superchargercontrols must be arranged to allow—
(1) Separate control of each engineand each supercharger; and
(2) Simultaneous control of all en-gines and all superchargers.
(c) Each power, thrust, or super-charger control must give a positiveand immediate responsive means ofcontrolling its engine or supercharger.
(d) The power, thrust, or super-charger controls for each engine or su-percharger must be independent ofthose for every other engine or super-charger.
(e) For each fluid injection (otherthan fuel) system and its controls notprovided and approved as part of theengine, the applicant must show thatthe flow of the injection fluid is ade-quately controlled.
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Federal Aviation Administration, DOT § 23.1157
(f) If a power, thrust, or a fuel con-trol (other than a mixture control) in-corporates a fuel shutoff feature, thecontrol must have a means to preventthe inadvertent movement of the con-trol into the off position. The meansmust—
(1) Have a positive lock or stop at theidle position; and
(2) Require a separate and distinctoperation to place the control in theshutoff position.
(g) For reciprocating single-engineairplanes, each power or thrust controlmust be designed so that if the controlseparates at the engine fuel meteringdevice, the airplane is capable of con-tinued safe flight and landing.
(a) Ignition switches must controland shut off each ignition circuit oneach engine.
(b) There must be means to quicklyshut off all ignition on multiengine air-planes by the grouping of switches orby a master ignition control.
(c) Each group of ignition switches,except ignition switches for turbine en-gines for which continuous ignition isnot required, and each master ignitioncontrol must have a means to preventits inadvertent operation.
(a) If there are mixture controls,each engine must have a separate con-trol, and each mixture control musthave guards or must be shaped or ar-ranged to prevent confusion by feelwith other controls.
(1) The controls must be grouped andarranged to allow—
(i) Separate control of each engine;and
(ii) Simultaneous control of all en-gines.
(2) The controls must require a sepa-rate and distinct operation to move the
control toward lean or shut-off posi-tion.
(b) For reciprocating single-engineairplanes, each manual engine mixturecontrol must be designed so that, if thecontrol separates at the engine fuelmetering device, the airplane is capa-ble of continued safe flight and land-ing.
(a) If there are propeller speed orpitch controls, they must be groupedand arranged to allow—
(1) Separate control of each pro-peller; and
(2) Simultaneous control of all pro-pellers.
(b) The controls must allow readysynchronization of all propellers onmultiengine airplanes.
§ 23.1153 Propeller feathering controls.
If there are propeller feathering con-trols installed, it must be possible tofeather each propeller separately. Eachcontrol must have a means to preventinadvertent operation.
[Doc. No. 27804, 61 FR 5138, Feb. 9, 1996]
§ 23.1155 Turbine engine reversethrust and propeller pitch settingsbelow the flight regime.
For turbine engine installations,each control for reverse thrust and forpropeller pitch settings below theflight regime must have means to pre-vent its inadvertent operation. Themeans must have a positive lock orstop at the flight idle position andmust require a separate and distinctoperation by the crew to displace thecontrol from the flight regime (forwardthrust regime for turbojet powered air-planes).
[Amdt. 23–7, 34 FR 13096, Aug. 13, 1969]
§ 23.1157 Carburetor air temperaturecontrols.
There must be a separate carburetorair temperature control for each en-gine.
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14 CFR Ch. I (1–1–01 Edition)§ 23.1163
§ 23.1163 Powerplant accessories.(a) Each engine mounted accessory
must—(1) Be approved for mounting on the
engine involved and use the provisionson the engines for mounting; or
(2) Have torque limiting means on allaccessory drives in order to prevent thetorque limits established for thosedrives from being exceeded; and
(3) In addition to paragraphs (a)(1) or(a)(2) of this section, be sealed to pre-vent contamination of the engine oilsystem and the accessory system.
(b) Electrical equipment subject toarcing or sparking must be installed tominimize the probability of contactwith any flammable fluids or vaporsthat might be present in a free state.
(c) Each generator rated at or morethan 6 kilowatts must be designed andinstalled to minimize the probabilityof a fire hazard in the event it malfunc-tions.
(d) If the continued rotation of anyaccessory remotely driven by the en-gine is hazardous when malfunctioningoccurs, a means to prevent rotationwithout interfering with the continuedoperation of the engine must be pro-vided.
(e) Each accessory driven by a gear-box that is not approved as part of thepowerplant driving the gearbox must—
(1) Have torque limiting means toprevent the torque limits establishedfor the affected drive from being ex-ceeded;
(2) Use the provisions on the gearboxfor mounting; and
(3) Be sealed to prevent contamina-tion of the gearbox oil system and theaccessory system.
§ 23.1165 Engine ignition systems.(a) Each battery ignition system
must be supplemented by a generatorthat is automatically available as analternate source of electrical energy toallow continued engine operation ifany battery becomes depleted.
(b) The capacity of batteries and gen-erators must be large enough to meetthe simultaneous demands of the en-
gine ignition system and the greatestdemands of any electrical system com-ponents that draw from the samesource.
(c) The design of the engine ignitionsystem must account for—
(1) The condition of an inoperativegenerator;
(2) The condition of a completely de-pleted battery with the generator run-ning at its normal operating speed; and
(3) The condition of a completely de-pleted battery with the generator oper-ating at idling speed, if there is onlyone battery.
(d) There must be means to warn ap-propriate crewmembers if malfunc-tioning of any part of the electricalsystem is causing the continuous dis-charge of any battery used for engineignition.
(e) Each turbine engine ignition sys-tem must be independent of any elec-trical circuit that is not used for as-sisting, controlling, or analyzing theoperation of that system.
(f) In addition, for commuter cat-egory airplanes, each turbopropeller ig-nition system must be an essentialelectrical load.
Components, lines, and fittings, ex-cept those subject to the provisions of§ 23.1351(e), located behind the engine-compartment firewall must be con-structed of such materials and locatedat such distances from the firewallthat they will not suffer damage suffi-cient to endanger the airplane if a por-tion of the engine side of the firewall issubjected to a flame temperature ofnot less than 2000 °F for 15 minutes.
[Amdt. 23–14, 38 FR 31816, Nov. 19, 1973]
§ 23.1183 Lines, fittings, and compo-nents.
(a) Except as provided in paragraph(b) of this section, each component,line, and fitting carrying flammablefluids, gas, or air in any area subject toengine fire conditions must be at leastfire resistant, except that flammablefluid tanks and supports which are partof and attached to the engine must befireproof or be enclosed by a fireproofshield unless damage by fire to anynon-fireproof part will not cause leak-age or spillage of flammable fluid.Components must be shielded or lo-cated so as to safeguard against the ig-nition of leaking flammable fluid.Flexible hose assemblies (hose and endfittings) must be shown to be suitablefor the particular application. An inte-gral oil sump of less than 25–quart ca-pacity on a reciprocating engine neednot be fireproof nor be enclosed by afireproof shield.
(b) Paragraph (a) of this section doesnot apply to—
(1) Lines, fittings, and componentswhich are already approved as part of atype certificated engine; and
(2) Vent and drain lines, and their fit-tings, whose failure will not result in,or add to, a fire hazard.
(a) For each multiengine airplane thefollowing apply:
(1) Each engine installation musthave means to shut off or otherwiseprevent hazardous quantities of fuel,oil, deicing fluid, and other flammableliquids from flowing into, within, orthrough any engine compartment, ex-cept in lines, fittings, and componentsforming an integral part of an engine.
(2) The closing of the fuel shutoffvalve for any engine may not make anyfuel unavailable to the remaining en-gines that would be available to thoseengines with that valve open.
(3) Operation of any shutoff meansmay not interfere with the later emer-gency operation of other equipmentsuch as propeller feathering devices.
(4) Each shutoff must be outside ofthe engine compartment unless anequal degree of safety is provided withthe shutoff inside the compartment.
(5) Not more than one quart of flam-mable fluid may escape into the enginecompartment after engine shutoff. Forthose installations where the flam-mable fluid that escapes after shut-down cannot be limited to one quart, itmust be demonstrated that this greateramount can be safely contained ordrained overboard.
(6) There must be means to guardagainst inadvertent operation of eachshutoff means, and to make it possiblefor the crew to reopen the shutoffmeans in flight after it has been closed.
(b) Turbine engine installations neednot have an engine oil system shutoffif—
(1) The oil tank is integral with, ormounted on, the engine; and
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14 CFR Ch. I (1–1–01 Edition)§ 23.1191
(2) All oil system components exter-nal to the engine are fireproof or lo-cated in areas not subject to enginefire conditions.
(c) Power operated valves must havemeans to indicate to the flight crewwhen the valve has reached the se-lected position and must be designed sothat the valve will not move from theselected position under vibration con-ditions likely to exist at the valve lo-cation.
§ 23.1191 Firewalls.(a) Each engine, auxiliary power
unit, fuel burning heater, and othercombustion equipment, must be iso-lated from the rest of the airplane byfirewalls, shrouds, or equivalentmeans.
(b) Each firewall or shroud must beconstructed so that no hazardous quan-tity of liquid, gas, or flame can passfrom the compartment created by thefirewall or shroud to other parts of theairplane.
(c) Each opening in the firewall orshroud must be sealed with close fit-ting, fireproof grommets, bushings, orfirewall fittings.
(d) [Reserved](e) Each firewall and shroud must be
fireproof and protected against corro-sion.
(f) Compliance with the criteria forfireproof materials or componentsmust be shown as follows:
(1) The flame to which the materialsor components are subjected must be2,000 ± 150°F.
(2) Sheet materials approximately 10inches square must be subjected to theflame from a suitable burner.
(3) The flame must be large enoughto maintain the required test tempera-ture over an area approximately fiveinches square.
(g) Firewall materials and fittingsmust resist flame penetration for atleast 15 minutes.
(h) The following materials may beused in firewalls or shrouds withoutbeing tested as required by this sec-tion:
(1) Stainless steel sheet, 0.015 inchthick.
(2) Mild steel sheet (coated with alu-minum or otherwise protected againstcorrosion) 0.018 inch thick.
(3) Terne plate, 0.018 inch thick.(4) Monel metal, 0.018 inch thick.(5) Steel or copper base alloy firewall
For aircooled radial engines, the en-gine power section and all portions ofthe exhaust sytem must be isolatedfrom the engine accessory compart-ment by a diaphragm that meets thefirewall requirements of § 23.1191.
[Amdt. 23–14, 38 FR 31823, Nov. 19, 1973]
§ 23.1193 Cowling and nacelle.(a) Each cowling must be constructed
and supported so that it can resist anyvibration, inertia, and air loads towhich it may be subjected in operation.
(b) There must be means for rapidand complete drainage of each part ofthe cowling in the normal ground andflight attitudes. Drain operation maybe shown by test, analysis, or both, toensure that under normal aerodynamicpressure distribution expected in serv-ice each drain will operate as designed.No drain may discharge where it willcause a fire hazard.
(c) Cowling must be at least fire re-sistant.
(d) Each part behind an opening inthe engine compartment cowling mustbe at least fire resistant for a distanceof at least 24 inches aft of the opening.
(e) Each part of the cowling subjectedto high temperatures due to its near-ness to exhaust sytem ports or exhaustgas impingement, must be fire proof.
(f) Each nacelle of a multiengine air-plane with supercharged engines mustbe designed and constructed so thatwith the landing gear retracted, a firein the engine compartment will notburn through a cowling or nacelle andenter a nacelle area other than the en-gine compartment.
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(g) In addition, for commuter cat-egory airplanes, the airplane must bedesigned so that no fire originating inany engine compartment can enter, ei-ther through openings or by burn-through, any other region where itwould create additional hazards.
§ 23.1195 Fire extinguishing systems.(a) For commuter category airplanes,
fire extinguishing systems must be in-stalled and compliance shown with thefollowing:
(1) Except for combustor, turbine,and tailpipe sections of turbine-engineinstallations that contain lines or com-ponents carrying flammable fluids orgases for which a fire originating inthese sections is shown to be control-lable, a fire extinguisher system mustserve each engine compartment;
(2) The fire extinguishing system, thequantity of the extinguishing agent,the rate of discharge, and the dischargedistribution must be adequate to extin-guish fires. An individual ‘‘one shot’’system may be used.
(3) The fire extinguishing system fora nacelle must be able to simulta-neously protect each compartment ofthe nacelle for which protection is pro-vided.
(b) If an auxiliary power unit is in-stalled in any airplane certificated tothis part, that auxiliary power unitcompartment must be served by a fireextinguishing system meeting the re-quirements of paragraph (a)(2) of thissection.
§ 23.1197 Fire extinguishing agents.For commuter category airplanes,
the following applies:(a) Fire extinguishing agents must—(1) Be capable of extinguishing
flames emanating from any burning offluids or other combustible materialsin the area protected by the fire extin-guishing system; and
(2) Have thermal stability over thetemperature range likely to be experi-
enced in the compartment in whichthey are stored.
(b) If any toxic extinguishing agent isused, provisions must be made to pre-vent harmful concentrations of fluid orfluid vapors (from leakage during nor-mal operation of the airplane or as aresult of discharging the fire extin-guisher on the ground or in flight) fromentering any personnel compartment,even though a defect may exist in theextinguishing system. This must beshown by test except for built-in car-bon dioxide fuselage compartment fireextinguishing systems for which—
(1) Five pounds or less of carbon diox-ide will be discharged, under estab-lished fire control procedures, into anyfuselage compartment; or
(2) Protective breathing equipment isavailable for each flight crewmemberon flight deck duty.
[Amdt. 23–34, 52 FR 1833, Jan. 15, 1987]
§ 23.1199 Extinguishing agent con-tainers.
For commuter category airplanes,the following applies:
(a) Each extinguishing agent con-tainer must have a pressure relief toprevent bursting of the container byexcessive internal pressures.
(b) The discharge end of each dis-charge line from a pressure relief con-nection must be located so that dis-charge of the fire extinguishing agentwould not damage the airplane. Theline must also be located or protectedto prevent clogging caused by ice orother foreign matter.
(c) A means must be provided foreach fire extinguishing agent containerto indicate that the container has dis-charged or that the charging pressureis below the established minimum nec-essary for proper functioning.
(d) The temperature of each con-tainer must be maintained, under in-tended operating conditions, to preventthe pressure in the container from—
(1) Falling below that necessary toprovide an adequate rate of discharge;or
(2) Rising high enough to cause pre-mature discharge.
(e) If a pyrotechnic capsule is used todischarge the extinguishing agent,each container must be installed sothat temperature conditions will not
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cause hazardous deterioration of thepyrotechnic capsule.
(a) There must be means that ensurethe prompt detection of a fire in—
(1) An engine compartment of—(i) Multiengine turbine powered air-
planes;(ii) Multiengine reciprocating engine
powered airplanes incorporatingturbochargers;
(iii) Airplanes with engine(s) locatedwhere they are not readily visible fromthe cockpit; and
(iv) All commuter category air-planes.
(2) The auxiliary power unit compart-ment of any airplane incorporating anauxiliary power unit.
(b) Each fire detector must be con-structed and installed to withstand thevibration, inertia, and other loads towhich it may be subjected in operation.
(c) No fire detector may be affectedby any oil, water, other fluids, orfumes that might be present.
(d) There must be means to allow thecrew to check, in flight, the func-tioning of each fire detector electriccircuit.
(e) Wiring and other components ofeach fire detector system in a des-ignated fire zone must be at least fireresistant.
§ 23.1301 Function and installation.Each item of installed equipment
must—(a) Be of a kind and design appro-
priate to its intended function.(b) Be labeled as to its identification,
function, or operating limitations, orany applicable combination of thesefactors;
(c) Be installed according to limita-tions specified for that equipment; and
(d) Function properly when installed.
[Amdt. 23–20, 42 FR 36968, July 18, 1977]
§ 23.1303 Flight and navigation instru-ments.
The following are the minimum re-quired flight and navigation instru-ments:
(a) An airspeed indicator.(b) An altimeter.(c) A direction indicator (non-
stabilized magnetic compass).(d) For reciprocating engine-powered
airplanes of more than 6,000 poundsmaximum weight and turbine enginepowered airplanes, a free air tempera-ture indicator or an air-temperatureindicator which provides indicationsthat are convertible to free-air.
(e) A speed warning device for—(1) Turbine engine powered airplanes;
and(2) Other airplanes for which VMO/
MMO and VD/MD are established under§§ 23.335(b)(4) and 23.1505(c) if VMO/MMOis greater than 0.8 VD/MD.
The speed warning device must giveeffective aural warning (differing dis-tinctively from aural warnings used forother purposes) to the pilots wheneverthe speed exceeds VMO plus 6 knots orMMO+0.01. The upper limit of the pro-duction tolerance for the warning de-vice may not exceed the prescribedwarning speed. The lower limit of thewarning device must be set to mini-mize nuisance warning;
(f) When an attitude display is in-stalled, the instrument design mustnot provide any means, accessible tothe flightcrew, of adjusting the relativepositions of the attitude reference sym-bol and the horizon line beyond thatnecessary for parallax correction.
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(g) In addition, for commuter cat-egory airplanes:
(1) If airspeed limitations vary withaltitude, the airspeed indicator musthave a maximum allowable airspeed in-dicator showing the variation of VMO
with altitude.(2) The altimeter must be a sensitive
type.(3) Having a passenger seating con-
figuration of 10 or more, excluding thepilot’s seats and that are approved forIFR operations, a third attitude instru-ment must be provided that:
(i) Is powered from a source inde-pendent of the electrical generatingsystem;
(ii) Continues reliable operation for aminimum of 30 minutes after total fail-ure of the electrical generating system;
(iii) Operates independently of anyother attitude indicating system;
(iv) Is operative without selectionafter total failure of the electrical gen-erating system;
(v) Is located on the instrumentpanel in a position acceptable to theAdministrator that will make it plain-ly visible to and usable by any pilot atthe pilot’s station; and
(vi) Is appropriately lighted duringall phases of operation.
The following are required power-plant instruments:
(a) For all airplanes. (1) A fuel quan-tity indicator for each fuel tank, in-stalled in accordance with § 23.1337(b).
(2) An oil pressure indicator for eachengine.
(3) An oil temperature indicator foreach engine.
(4) An oil quantity measuring devicefor each oil tank which meets the re-quirements of § 23.1337(d).
(5) A fire warning means for thoseairplanes required to comply with§ 23.1203.
(b) For reciprocating engine-poweredairplanes. In addition to the powerplantinstruments required by paragraph (a)of this section, the following power-plant instruments are required:
(1) An induction system air tempera-ture indicator for each engine equippedwith a preheater and having inductionair temperature limitations that canbe exceeded with preheat.
(2) A tachometer indicator for eachengine.
(3) A cylinder head temperature indi-cator for—
(i) Each air-cooled engine with cowlflaps;
(ii) [Reserved](iii) Each commuter category air-
plane.(4) For each pump-fed engine, a
means:(i) That continuously indicates, to
the pilot, the fuel pressure or fuel flow;or
(ii) That continuously monitors thefuel system and warns the pilot of anyfuel flow trend that could lead to en-gine failure.
(5) A manifold pressure indicator foreach altitude engine and for each en-gine with a controllable propeller.
(6) For each turbocharger installa-tion:
(i) If limitations are established foreither carburetor (or manifold) airinlet temperature or exhaust gas orturbocharger turbine inlet tempera-ture, indicators must be furnished foreach temperature for which the limita-tion is established unless it is shownthat the limitation will not be exceed-ed in all intended operations.
(ii) If its oil system is separate fromthe engine oil system, oil pressure andoil temperature indicators must be pro-vided.
(7) A coolant temperature indicatorfor each liquid-cooled engine.
(c) For turbine engine-powered air-planes. In addition to the powerplantinstruments required by paragraph (a)of this section, the following power-plant instruments are required:
(1) A gas temperature indicator foreach engine.
(2) A fuel flowmeter indicator foreach engine.
(3) A fuel low pressure warningmeans for each engine.
(4) A fuel low level warning means forany fuel tank that should not be de-pleted of fuel in normal operations.
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(5) A tachometer indicator (to indi-cate the speed of the rotors with estab-lished limiting speeds) for each engine.
(6) An oil low pressure warningmeans for each engine.
(7) An indicating means to indicatethe functioning of the powerplant iceprotection system for each engine.
(8) For each engine, an indicatingmeans for the fuel strainer or filter re-quired by § 23.997 to indicate the occur-rence of contamination of the straineror filter before it reaches the capacityestablished in accordance with§ 23.997(d).
(9) For each engine, a warning meansfor the oil strainer or filter required by§ 23.1019, if it has no bypass, to warn thepilot of the occurrence of contamina-tion of the strainer or filter screen be-fore it reaches the capacity establishedin accordance with § 23.1019(a)(5).
(10) An indicating means to indicatethe functioning of any heater used toprevent ice clogging of fuel systemcomponents.
(d) For turbojet/turbofan engine-pow-ered airplanes. In addition to the power-plant instruments required by para-graphs (a) and (c) of this section, thefollowing powerplant instruments arerequired:
(1) For each engine, an indicator toindicate thrust or to indicate a param-eter that can be related to thrust, in-cluding a free air temperature indi-cator if needed for this purpose.
(2) For each engine, a position indi-cating means to indicate to the flightcrew when the thrust reverser, if in-stalled, is in the reverse thrust posi-tion.
(e) For turbopropeller-powered air-planes. In addition to the powerplantinstruments required by paragraphs (a)and (c) of this section, the followingpowerplant instruments are required:
(1) A torque indicator for each en-gine.
(2) A position indicating means to in-dicate to the flight crew when the pro-peller blade angle is below the flightlow pitch position, for each propeller,unless it can be shown that such occur-rence is highly improbable.
The equipment necessary for an air-plane to operate at the maximum oper-ating altitude and in the kinds of oper-ation and meteorological conditionsfor which certification is requested andis approved in accordance with § 23.1559must be included in the type design.
(a) Each item of equipment, each sys-tem, and each installation:
(1) When performing its intendedfunction, may not adversely affect theresponse, operation, or accuracy ofany—
(i) Equipment essential to safe oper-ation; or
(ii) Other equipment unless there is ameans to inform the pilot of the effect.
(2) In a single-engine airplane, mustbe designed to minimize hazards to theairplane in the event of a probable mal-function or failure.
(3) In a multiengine airplane, must bedesigned to prevent hazards to the air-plane in the event of a probable mal-function or failure.
(4) In a commuter category airplane,must be designed to safeguard againsthazards to the airplane in the event oftheir malfunction or failure.
(b) The design of each item of equip-ment, each system, and each installa-tion must be examined separately andin relationship to other airplane sys-tems and installations to determine ifthe airplane is dependent upon its func-tion for continued safe flight and land-ing and, for airplanes not limited toVFR conditions, if failure of a systemwould significantly reduce the capa-bility of the airplane or the ability ofthe crew to cope with adverse oper-ating conditions. Each item of equip-ment, each system, and each installa-tion identified by this examination asone upon which the airplane is depend-ent for proper functioning to ensurecontinued safe flight and landing, orwhose failure would significantly re-duce the capability of the airplane or
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Federal Aviation Administration, DOT § 23.1309
the ability of the crew to cope with ad-verse operating conditions, must be de-signed to comply with the following ad-ditional requirements:
(1) It must perform its intended func-tion under any foreseeable operatingcondition.
(2) When systems and associatedcomponents are considered separatelyand in relation to other systems—
(i) The occurrence of any failure con-dition that would prevent the contin-ued safe flight and landing of the air-plane must be extremely improbable;and
(ii) The occurrence of any other fail-ure condition that would significantlyreduce the capability of the airplane orthe ability of the crew to cope with ad-verse operating conditions must be im-probable.
(3) Warning information must be pro-vided to alert the crew to unsafe sys-tem operating conditions and to enablethem to take appropriate correctiveaction. Systems, controls, and associ-ated monitoring and warning meansmust be designed to minimize crew er-rors that could create additional haz-ards.
(4) Compliance with the requirementsof paragraph (b)(2) of this section maybe shown by analysis and, where nec-essary, by appropriate ground, flight,or simulator tests. The analysis mustconsider—
(i) Possible modes of failure, includ-ing malfunctions and damage from ex-ternal sources;
(ii) The probability of multiple fail-ures, and the probability of undetectedfaults.;
(iii) The resulting effects on the air-plane and occupants, considering thestage of flight and operating condi-tions; and
(iv) The crew warning cues, correc-tive action required, and the crew’s ca-pability of determining faults.
(c) Each item of equipment, each sys-tem, and each installation whose func-tioning is required by this chapter andthat requires a power supply is an ‘‘es-sential load’’ on the power supply. Thepower sources and the system must beable to supply the following powerloads in probable operating combina-tions and for probable durations:
(1) Loads connected to the power dis-tribution system with the system func-tioning normally.
(2) Essential loads after failure of—(i) Any one engine on two-engine air-
planes; or(ii) Any two engines on an airplane
with three or more engines; or(iii) Any power converter or energy
storage device.(3) Essential loads for which an alter-
nate source of power is required, as ap-plicable, by the operating rules of thischapter, after any failure or malfunc-tion in any one power supply system,distribution system, or other utiliza-tion system.
(d) In determining compliance withparagraph (c)(2) of this section, thepower loads may be assumed to be re-duced under a monitoring procedureconsistent with safety in the kinds ofoperations authorized. Loads not re-quired in controlled flight need not beconsidered for the two-engine-inoper-ative condition on airplanes with threeor more engines.
(e) In showing compliance with thissection with regard to the electricalpower system and to equipment designand installation, critical environ-mental and atmospheric conditions, in-cluding radio frequency energy and theeffects (both direct and indirect) oflightning strikes, must be considered.For electrical generation, distribution,and utilization equipment required byor used in complying with this chapter,the ability to provide continuous, safeservice under forseeable environmentalconditions may be shown by environ-mental tests, design analysis, or ref-erence to previous comparable serviceexperience on other airplanes.
(f) As used in this section, ‘‘system’’refers to all pneumatic systems, fluidsystems, electrical systems, mechan-ical systems, and powerplant systemsincluded in the airplane design, exceptfor the following:
(1) Powerplant systems provided aspart of the certificated engine.
(2) The flight structure (such a wing,empennage, control surfaces and theirsystems, the fuselage, engine mount-ing, and landing gear and their related
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primary attachments) whose require-ments are specific in subparts C and Dof this part.
(a) Electronic display indicators, in-cluding those with features that makeisolation and independence betweenpowerplant instrument systems im-practical, must:
(1) Meet the arrangement and visi-bility requirements of § 23.1321.
(2) Be easily legible under all lightingconditions encountered in the cockpit,including direct sunlight, consideringthe expected electronic display bright-ness level at the end of an electronicdisplay indictor’s useful life. Specificlimitations on display system usefullife must be contained in the Instruc-tions for Continued Airworthiness re-quired by § 23.1529.
(3) Not inhibit the primary display ofattitude, airspeed, altitude, or power-plant parameters needed by any pilotto set power within established limita-tions, in any normal mode of oper-ation.
(4) Not inhibit the primary display ofengine parameters needed by any pilotto properly set or monitor powerplantlimitations during the engine startingmode of operation.
(5) Have an independent magnetic di-rection indicator and either an inde-pendent secondary mechanical altim-eter, airspeed indicator, and attitudeinstrument or individual electronicdisplay indicators for the altitude, air-speed, and attitude that are inde-pendent from the airplane’s primaryelectrical power system. These sec-ondary instruments may be installed inpanel positions that are displaced fromthe primary positions specified by§ 23.1321(d), but must be located wherethey meet the pilot’s visibility require-ments of § 23.1321(a).
(6) Incorporate sensory cues for thepilot that are equivalent to those inthe instrument being replaced by theelectronic display indicators.
(7) Incorporate visual displays of in-strument markings, required by§§ 23.1541 through 23.1553, or visual dis-plays that alert the pilot to abnormaloperational values or approaches to es-tablished limitation values, for eachparameter required to be displayed bythis part.
(b) The electronic display indicators,including their systems and installa-tions, and considering other airplanesystems, must be designed so that onedisplay of information essential forcontinued safe flight and landing willremain available to the crew, withoutneed for immediate action by any pilotfor continued safe operation, after anysingle failure or probable combinationof failures.
(c) As used in this section, ‘‘instru-ment’’ includes devices that are phys-ically contained in one unit, and de-vices that are composed of two or morephysically separate units or compo-nents connected together (such as a re-mote indicating gyroscopic directionindicator that includes a magneticsensing element, a gyroscopic unit, anamplifier, and an indicator connectedtogether). As used in this section, ‘‘pri-mary’’ display refers to the display of aparameter that is located in the instru-ment panel such that the pilot looks atit first when wanting to view that pa-rameter.
[Doc. No. 27806, 61 FR 5168, Feb. 9, 1996]
§ 23.1321 Arrangement and visibility.
(a) Each flight, navigation, and pow-erplant instrument for use by any re-quired pilot during takeoff, initialclimb, final approach, and landingmust be located so that any pilot seat-ed at the controls can monitor the air-plane’s flight path and these instru-ments with minimum head and eyemovement. The powerplant instru-ments for these flight conditions arethose needed to set power within pow-erplant limitations.
(b) For each multiengine airplane,identical powerplant instruments mustbe located so as to prevent confusion asto which engine each instrument re-lates.
(c) Instrument panel vibration maynot damage, or impair the accuracy of,any instrument.
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(d) For each airplane, the flight in-struments required by § 23.1303, and, asapplicable, by the operating rules ofthis chapter, must be grouped on theinstrument panel and centered as near-ly as practicable about the verticalplane of each required pilot’s forwardvision. In addition:
(1) The instrument that most effec-tively indicates the attitude must beon the panel in the top center position;
(2) The instrument that most effec-tively indicates airspeed must be adja-cent to and directly to the left of theinstrument in the top center position;
(3) The instrument that most effec-tively indicates altitude must be adja-cent to and directly to the right of theinstrument in the top center position;
(4) The instrument that most effec-tively indicates direction of flight,other than the magnetic direction indi-cator required by § 23.1303(c), must beadjacent to and directly below the in-strument in the top center position;and
(5) Electronic display indicators maybe used for compliance with paragraphs(d)(1) through (d)(4) of this sectionwhen such displays comply with re-quirements in § 23.1311.
(e) If a visual indicator is provided toindicate malfunction of an instrument,it must be effective under all probablecockpit lighting conditions.
(a) Each airspeed indicatinginstrucment must be calibrated to indi-cate true airspeed (at sea level with astandard atmosphere) with a minimumpracticable instrument calibrationerror when the corresponding pitot andstatic pressures are applied.
(b) Each airspeed system must becalibrated in flight to determine thesystem error. The system error, includ-ing position error, but excluding theairspeed indicator instrument calibra-tion error, may not exceed three per-cent of the calibrated airspeed or fiveknots, whichever is greater, through-out the following speed ranges:
(1) 1.3 VS1 to VMO/MMO or VNE, which-ever is appropriate with flaps re-tracted.
(2) 1.3 VS1 to VFE with flaps extended.(c) The design and installation of
each airspeed indicating system mustprovide positive drainage of moisturefrom the pitot static plumbing.
(d) If certification for instrumentflight rules or flight in icing conditionsis requested, each airspeed systemmust have a heated pitot tube or anequivalent means of preventing mal-function due to icing.
(e) In addition, for commuter cat-egory airplanes, the airspeed indi-cating system must be calibrated to de-termine the system error during theaccelerate-takeoff ground run. Theground run calibration must be ob-tained between 0.8 of the minimumvalue of V1, and 1.2 times the maximumvalue of V1 considering the approvedranges of altitude and weight. Theground run calibration must be deter-mined assuming an engine failure atthe minimum value of V1.
(f) For commuter category airplanes,where duplicate airspeed indicators arerequired, their respective pitot tubes
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must be far enough apart to avoid dam-age to both tubes in a collision with abird.
§ 23.1325 Static pressure system.(a) Each instrument provided with
static pressure case connections mustbe so vented that the influence of air-plane speed, the opening and closing ofwindows, airflow variations, moisture,or other foreign matter will least af-fect the accuracy of the instrumentsexcept as noted in paragraph (b)(3) ofthis section.
(b) If a static pressure system is nec-essary for the functioning of instru-ments, systems, or devices, it mustcomply with the provisions of para-graphs (b) (1) through (3) of this sec-tion.
(1) The design and installation of astatic pressure system must be suchthat—
(i) Positive drainage of moisture isprovided;
(ii) Chafing of the tubing, and exces-sive distortion or restriction at bendsin the tubing, is avoided; and
(iii) The materials used are durable,suitable for the purpose intended, andprotected against corrosion.
(2) A proof test must be conducted todemonstrate the integrity of the staticpressure system in the following man-ner:
(i) Unpressurized airplanes. Evacuatethe static pressure system to a pres-sure differential of approximately 1inch of mercury or to a reading on thealtimeter, 1,000 feet above the aircraftelevation at the time of the test. With-out additional pumping for a period of1 minute, the loss of indicated altitudemust not exceed 100 feet on the altim-eter.
(ii) Pressurized airplanes. Evacuatethe static pressure system until a pres-sure differential equivalent to the max-imum cabin pressure differential forwhich the airplane is type certificatedis achieved. Without additional pump-ing for a period of 1 minute, the loss ofindicated altitude must not exceed 2percent of the equivalent altitude of
the maximum cabin differential pres-sure or 100 feet, whichever is greater.
(3) If a static pressure system is pro-vided for any instrument, device, orsystem required by the operating rulesof this chapter, each static pressureport must be designed or located insuch a manner that the correlation be-tween air pressure in the static pres-sure system and true ambient atmos-pheric static pressure is not alteredwhen the airplane encounters icingconditions. An antiicing means or analternate source of static pressure maybe used in showing compliance withthis requirement. If the reading of thealtimeter, when on the alternate staticpressure system differs from the read-ing of the altimeter when on the pri-mary static system by more than 50feet, a correction card must be pro-vided for the alternate static system.
(c) Except as provided in paragraph(d) of this section, if the static pressuresystem incorporates both a primaryand an alternate static pressure source,the means for selecting one or theother source must be designed sothat—
(1) When either source is selected, theother is blocked off; and
(2) Both sources cannot be blockedoff simultaneously.
(d) For unpressurized airplanes, para-graph (c)(1) of this section does notapply if it can be demonstrated thatthe static pressure system calibration,when either static pressure source isselected, is not changed by the otherstatic pressure source being open orblocked.
(e) Each static pressure system mustbe calibrated in flight to determine thesystem error. The system error, in in-dicated pressure altitude, at sea-level,with a standard atmosphere, excludinginstrument calibration error, may notexceed ±30 feet per 100 knot speed forthe appropriate configuration in thespeed range between 1.3 VS0 with flapsextended, and 1.8 VS1 with flaps re-tracted. However, the error need not beless than 30 feet.
(f) [Reserved](g) For airplanes prohibited from
flight in instrument meteorological oricing conditions, in accordance with
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Federal Aviation Administration, DOT § 23.1329
§ 23.1559(b) of this part, paragraph (b)(3)of this section does not apply.
If a flight instrument pitot heatingsystem is installed to meet the require-ments specified in § 23.1323(d), an indi-cation system must be provided to in-dicate to the flight crew when thatpitot heating system is not operating.The indication system must complywith the following requirements:
(a) The indication provided must in-corporate an amber light that is inclear view of a flightcrew member.
(b) The indication provided must bedesigned to alert the flight crew if ei-ther of the following conditions exist:
(1) The pitot heating system isswitched ‘‘off.’’
(2) The pitot heating system isswitched ‘‘on’’ and any pitot tube heat-ing element is inoperative.
[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]
§ 23.1327 Magnetic direction indicator.
(a) Except as provided in paragraph(b) of this section—
(1) Each magnetic direction indicatormust be installed so that its accuracyis not excessively affected by the air-plane’s vibration or magnetic fields;and
(2) The compensated installation maynot have a deviation in level flight,greater than ten degrees on any head-ing.
(b) A magnetic nonstabilized direc-tion indicator may deviate more thanten degrees due to the operation ofelectrically powered systems such aselectrically heated windshields if ei-ther a magnetic stabilized direction in-dicator, which does not have a devi-ation in level flight greater than tendegrees on any heading, or a gyroscopicdirection indicator, is installed. Devi-ations of a magnetic nonstabilized di-rection indicator of more than 10 de-
grees must be placarded in accordancewith § 23.1547(e).
[Amdt. 23–20, 42 FR 36969, July 18, 1977]
§ 23.1329 Automatic pilot system.If an automatic pilot system is in-
stalled, it must meet the following:(a) Each system must be designed so
that the automatic pilot can—(1) Be quickly and positively dis-
engaged by the pilots to prevent itfrom interfering with their control ofthe airplane; or
(2) Be sufficiently overpowered byone pilot to let him control the air-plane.
(b) If the provisions of paragraph(a)(1) of this section are applied, thequick release (emergency) controlmust be located on the control wheel(both control wheels if the airplane canbe operated from either pilot seat) onthe side opposite the throttles, or onthe stick control, (both stick controls,if the airplane can be operated from ei-ther pilot seat) such that it can be op-erated without moving the hand fromits normal position on the control.
(c) Unless there is automatic syn-chronization, each system must have ameans to readily indicate to the pilotthe alignment of the actuating devicein relation to the control system it op-erates.
(d) Each manually operated controlfor the system operation must be read-ily accessible to the pilot. Each controlmust operate in the same plane andsense of motion as specified in § 23.779for cockpit controls. The direction ofmotion must be plainly indicated on ornear each control.
(e) Each system must be designed andadjusted so that, within the range ofadjustment available to the pilot, itcannot produce hazardous loads on theairplane or create hazardous deviationsin the flight path, under any flight con-dition appropriate to its use, eitherduring normal operation or in theevent of a malfunction, assuming thatcorrective action begins within a rea-sonable period of time.
(f) Each system must be designed sothat a single malfunction will notproduce a hardover signal in more thanone control axis. If the automatic pilotintegrates signals from auxiliary con-trols or furnishes signals for operation
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of other equipment, positive interlocksand sequencing of engagement to pre-vent improper operation are required.
(g) There must be protection againstadverse interaction of integrated com-ponents, resulting from a malfunction.
(h) If the automatic pilot system canbe coupled to airborne navigationequipment, means must be provided toindicate to the flight crew the currentmode of operation. Selector switch po-sition is not acceptable as a means ofindication.
For each instrument that uses apower source, the following apply:
(a) Each instrument must have an in-tegral visual power annunciator or sep-arate power indicator to indicate whenpower is not adequate to sustain properinstrument performance. If a separateindicator is used, it must be located sothat the pilot using the instrumentscan monitor the indicator with min-imum head and eye movement. Thepower must be sensed at or near thepoint where it enters the instrument.For electric and vacuum/pressure in-struments, the power is considered tobe adequate when the voltage or thevacuum/pressure, respectively, is with-in approved limits.
(b) The installation and power supplysystems must be designed so that—
(1) The failure of one instrument willnot interfere with the proper supply ofenergy to the remaining instrument;and
(2) The failure of the energy supplyfrom one source will not interfere withthe proper supply of energy from anyother source.
(c) There must be at least two inde-pendent sources of power (not drivenby the same engine on multiengine air-planes), and a manual or an automaticmeans to select each power source.
[Doc. No. 26344, 58 FR 18976, Apr. 9, 1993]
§ 23.1335 Flight director systems.If a flight director system is in-
stalled, means must be provided to in-
dicate to the flight crew its currentmode of operation. Selector switch po-sition is not acceptable as a means ofindication.
[Amdt. 23–20, 42 FR 36969, July 18, 1977]
§ 23.1337 Powerplant instruments in-stallation.
(a) Instruments and instrument lines.(1) Each powerplant and auxiliarypower unit instrument line must meetthe requirements of § 23.993.
(2) Each line carrying flammablefluids under pressure must—
(i) Have restricting orifices or othersafety devices at the source of pressureto prevent the escape of excessive fluidif the line fails; and
(ii) Be installed and located so thatthe escape of fluids would not create ahazard.
(3) Each powerplant and auxiliarypower unit instrument that utilizesflammable fluids must be installed andlocated so that the escape of fluidwould not create a hazard.
(b) Fuel quantity indication. Theremust be a means to indicate to theflightcrew members the quantity of us-able fuel in each tank during flight. Anindicator calibrated in appropriateunits and clearly marked to indicatethose units must be used. In addition:
(1) Each fuel quantity indicator mustbe calibrated to read ‘‘zero’’ duringlevel flight when the quantity of fuelremaining in the tank is equal to theunusable fuel supply determined under§ 23.959(a);
(2) Each exposed sight gauge used asa fuel quantity indicator must be pro-tected against damage;
(3) Each sight gauge that forms atrap in which water can collect andfreeze must have means to allow drain-age on the ground;
(4) There must be a means to indicatethe amount of usable fuel in each tankwhen the airplane is on the ground(such as by a stick gauge);
(5) Tanks with interconnected outletsand airspaces may be considered as onetank and need not have separate indi-cators; and
(6) No fuel quantity indicator is re-quired for an auxiliary tank that isused only to transfer fuel to othertanks if the relative size of the tank,
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Federal Aviation Administration, DOT § 23.1351
the rate of fuel transfer, and operatinginstructions are adequate to—
(i) Guard against overflow; and(ii) Give the flight crewmembers
prompt warning if transfer is not pro-ceeding as planned.
(c) Fuel flowmeter system. If a fuelflowmeter system is installed, eachmetering component must have ameans to by-pass the fuel supply ifmalfunctioning of that component se-verely restricts fuel flow.
(d) Oil quantity indicator. There mustbe a means to indicate the quantity ofoil in each tank—
(1) On the ground (such as by a stickgauge); and
(2) In flight, to the flight crew mem-bers, if there is an oil transfer systemor a reserve oil supply system.
§ 23.1351 General.(a) Electrical system capacity. Each
electrical system must be adequate forthe intended use. In addition—
(1) Electric power sources, theirtransmission cables, and their associ-ated control and protective devices,must be able to furnish the requiredpower at the proper voltage to eachload circuit essential for safe oper-ation; and
(2) Compliance with paragraph (a)(1)of this section must be shown as fol-lows—
(i) For normal, utility, and acrobaticcategory airplanes, by an electricalload analysis or by electrical measure-ments that account for the electricalloads applied to the electrical systemin probable combinations and for prob-able durations; and
(ii) For commuter category air-planes, by an electrical load analysisthat accounts for the electrical loadsapplied to the electrical system inprobable combinations and for probabledurations.
(b) Function. For each electrical sys-tem, the following apply:
(1) Each system, when installed,must be—
(i) Free from hazards in itself, in itsmethod of operation, and in its effectson other parts of the airplane;
(ii) Protected from fuel, oil, water,other detrimental substances, and me-chanical damage; and
(iii) So designed that the risk of elec-trical shock to crew, passengers, andground personnel is reduced to a min-imum.
(2) Electric power sources must func-tion properly when connected in com-bination or independently.
(3) No failure or malfunction of anyelectric power source may impair theability of any remaining source to sup-ply load circuits essential for safe oper-ation.
(4) In addition, for commuter cat-egory airplanes, the following apply:
(i) Each system must be designed sothat essential load circuits can be sup-plied in the event of reasonably prob-able faults or open circuits includingfaults in heavy current carrying cables;
(ii) A means must be accessible inflight to the flight crewmembers forthe individual and collective dis-connection of the electrical powersources from the system;
(iii) The system must be designed sothat voltage and frequency, if applica-ble, at the terminals of all essentialload equipment can be maintainedwithin the limits for which the equip-ment is designed during any probableoperating conditions;
(iv) If two independent sources ofelectrical power for particular equip-ment or systems are required, theirelectrical energy supply must be en-sured by means such as duplicate elec-trical equipment, throwover switching,or multichannel or loop circuits sepa-rately routed; and
(v) For the purpose of complyingwith paragraph (b)(5) of this section,the distribution system includes thedistribution busses, their associatedfeeders, and each control and protec-tive device.
(c) Generating system. There must beat least one generator/alternator if theelectrical system supplies power toload circuits essential for safe oper-ation. In addition—
(1) Each generator/alternator must beable to deliver its continuous rated
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power, or such power as is limited byits regulation system.
(2) Generator/alternator voltage con-trol equipment must be able to depend-ably regulate the generator/alternatoroutput within rated limits.
(3) Automatic means must be pro-vided to prevent damage to any gener-ator/alternator and adverse effects onthe airplane electrical system due toreverse current. A means must also beprovided to disconnect each generator/alternator from the battery and othergenerators/alternators.
(4) There must be a means to give im-mediate warning to the flight crew of afailure of any generator/alternator.
(5) Each generator/alternator musthave an overvoltage control designedand installed to prevent damage to theelectrical system, or to equipment sup-plied by the electrical system thatcould result if that generator/alter-nator were to develop an overvoltagecondition.
(d) Instruments. A means must existto indicate to appropriate flight crew-members the electric power systemquantities essential for safe operation.
(1) For normal, utility, and acrobaticcategory airplanes with direct currentsystems, an ammeter that can beswitched into each generator feedermay be used and, if only one generatorexists, the ammeter may be in the bat-tery feeder.
(2) For commuter category airplanes,the essential electric power systemquantities include the voltage and cur-rent supplied by each generator.
(e) Fire resistance. Electrical equip-ment must be so designed and installedthat in the event of a fire in the enginecompartment, during which the surfaceof the firewall adjacent to the fire isheated to 2,000° F for 5 minutes or to alesser temperature substantiated bythe applicant, the equipment essentialto continued safe operation and locatedbehind the firewall will function satis-factorily and will not create an addi-tional fire hazard.
(f) External power. If provisions aremade for connecting external power tothe airplane, and that external powercan be electrically connected to equip-ment other than that used for enginestarting, means must be provided toensure that no external power supply
having a reverse polarity, or a reversephase sequence, can supply power tothe airplane’s electrical system. Theexternal power connection must be lo-cated so that its use will not result ina hazard to the airplane or ground per-sonnel.
(g) It must be shown by analysis,tests, or both, that the airplane can beoperated safely in VFR conditions, fora period of not less than five minutes,with the normal electrical power (elec-trical power sources excluding the bat-tery and any other standby electricalsources) inoperative, with critical typefuel (from the standpoint of flameoutand restart capability), and with theairplane initially at the maximum cer-tificated altitude. Parts of the elec-trical system may remain on if—
(1) A single malfunction, including awire bundle or junction box fire, can-not result in loss of the part turned offand the part turned on; and
(2) The parts turned on are elec-trically and mechanically isolatedfrom the parts turned off.
(a) Each storage battery must be de-signed and installed as prescribed inthis section.
(b) Safe cell temperatures and pres-sures must be maintained during anyprobable charging and discharging con-dition. No uncontrolled increase in celltemperature may result when the bat-tery is recharged (after previous com-plete discharge)—
(1) At maximum regulated voltage orpower;
(2) During a flight of maximum dura-tion; and
(3) Under the most adverse coolingcondition likely to occur in service.
(c) Compliance with paragraph (b) ofthis section must be shown by tests un-less experience with similar batteriesand installations has shown that main-taining safe cell temperatures andpressures presents no problem.
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Federal Aviation Administration, DOT § 23.1359
(d) No explosive or toxic gases emit-ted by any battery in normal oper-ation, or as the result of any probablemalfunction in the charging system orbattery installation, may accumulatein hazardous quantities within the air-plane.
(e) No corrosive fluids or gases thatmay escape from the battery may dam-age surrounding structures or adjacentessential equipment.
(f) Each nickel cadmium battery in-stallation capable of being used tostart an engine or auxiliary power unitmust have provisions to prevent anyhazardous effect on structure or essen-tial systems that may be caused by themaximum amount of heat the batterycan generate during a short circuit ofthe battery or of its individual cells.
(g) Nickel cadmium battery installa-tions capable of being used to start anengine or auxiliary power unit musthave—
(1) A system to control the chargingrate of the battery automatically so asto prevent battery overheating;
(2) A battery temperature sensingand over-temperature warning systemwith a means for disconnecting thebattery from its charging source in theevent of an over-temperature condi-tion; or
(3) A battery failure sensing andwarning system with a means for dis-connecting the battery from its charg-ing source in the event of battery fail-ure.
(h) In the event of a complete loss ofthe primary electrical power gener-ating system, the battery must be ca-pable of providing at least 30 minutesof electrical power to those loads thatare essential to continued safe flightand landing. The 30 minute time periodincludes the time needed for the pilotsto recognize the loss of generatedpower and take appropriate load shed-ding action.
§ 23.1357 Circuit protective devices.(a) Protective devices, such as fuses
or circuit breakers, must be installedin all electrical circuits other than—
(1) Main circuits of starter motorsused during starting only; and
(2) Circuits in which no hazard is pre-sented by their omission.
(b) A protective device for a circuitessential to flight safety may not beused to protect any other circuit.
(c) Each resettable circuit protectivedevice (‘‘trip free’’ device in which thetripping mechanism cannot be over-ridden by the operating control) mustbe designed so that—
(1) A manual operation is required torestore service after tripping; and
(2) If an overload or circuit fault ex-ists, the device will open the circuit re-gardless of the position of the oper-ating control.
(d) If the ability to reset a circuitbreaker or replace a fuse is essential tosafety in flight, that circuit breaker orfuse must be so located and identifiedthat it can be readily reset or replacedin flight.
(e) For fuses identified as replaceablein flight—
(1) There must be one spare of eachrating or 50 percent spare fuses of eachrating, whichever is greater; and
(2) The spare fuse(s) must be readilyaccessible to any required pilot.
(a) Each component of the electricalsystem must meet the applicable fireprotection requirements of §§ 23.863 and23.1182.
(b) Electrical cables, terminals, andequipment in designated fire zones thatare used during emergency proceduresmust be fire-resistant.
(c) Insulation on electrical wire andelectrical cable must be self-extin-guishing when tested at an angle of 60degrees in accordance with the applica-ble portions of appendix F of this part,or other approved equivalent methods.The average burn length must not ex-ceed 3 inches (76 mm) and the averageflame time after removal of the flamesource must not exceed 30 seconds.Drippings from the test specimen must
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not continue to flame for more than anaverage of 3 seconds after falling.
[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]
§ 23.1361 Master switch arrangement.
(a) There must be a master switch ar-rangement to allow ready disconnec-tion of each electric power source frompower distribution systems, except asprovided in paragraph (b) of this sec-tion. The point of disconnection mustbe adjacent to the sources controlledby the switch arrangement. If separateswitches are incorporated into themaster switch arrangement, a meansmust be provided for the switch ar-rangement to be operated by one handwith a single movement.
(b) Load circuits may be connected sothat they remain energized when themaster switch is open, if the circuitsare isolated, or physically shielded, toprevent their igniting flammable fluidsor vapors that might be liberated bythe leakage or rupture of any flam-mable fluid system; and
(1) The circuits are required for con-tinued operation of the engine; or
(2) The circuits are protected by cir-cuit protective devices with a rating offive amperes or less adjacent to theelectric power source.
(3) In addition, two or more circuitsinstalled in accordance with the re-quirements of paragraph (b)(2) of thissection must not be used to supply aload of more than five amperes.
(c) The master switch or its controlsmust be so installed that the switch iseasily discernible and accessible to acrewmember.
(a) Each electric connecting cablemust be of adequate capacity.
(b) Any equipment that is associatedwith any electrical cable installationand that would overheat in the event ofcircuit overload or fault must be flameresistant. That equipment and the elec-trical cables must not emit dangerousquantities of toxic fumes.
(c) Main power cables (including gen-erator cables) in the fuselage must bedesigned to allow a reasonable degreeof deformation and stretching withoutfailure and must—
(1) Be separated from flammable fluidlines; or
(2) Be shrouded by means of elec-trically insulated flexible conduit, orequivalent, which is in addition to thenormal cable insulation.
(d) Means of identification must beprovided for electrical cables, termi-nals, and connectors.
(e) Electrical cables must be in-stalled such that the risk of mechan-ical damage and/or damage cased byfluids vapors, or sources of heat, isminimized.
(f) Where a cable cannot be protectedby a circuit protection device or otheroverload protection, it must not causea fire hazard under fault conditions.
§ 23.1367 Switches.Each switch must be—(a) Able to carry its rated current;(b) Constructed with enough distance
or insulating material between currentcarrying parts and the housing so thatvibration in flight will not cause short-ing;
(c) Accessible to appropriate flightcrewmembers; and
(d) Labeled as to operation and thecircuit controlled.
LIGHTS
§ 23.1381 Instrument lights.The instrument lights must—(a) Make each instrument and con-
trol easily readable and discernible;(b) Be installed so that their direct
rays, and rays reflected from the wind-shield or other surface, are shieldedfrom the pilot’s eyes; and
(c) Have enough distance or insu-lating material between current car-rying parts and the housing so that vi-bration in flight will not cause short-ing.A cabin dome light is not an instru-ment light.
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Federal Aviation Administration, DOT § 23.1389
§ 23.1383 Taxi and landing lights.Each taxi and landing light must be
designed and installed so that:(a) No dangerous glare is visible to
the pilots.(b) The pilot is not seriously affected
by halation.(c) It provides enough light for night
operations.(d) It does not cause a fire hazard in
any configuration.
[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]
§ 23.1385 Position light system installa-tion.
(a) General. Each part of each posi-tion light system must meet the appli-cable requirements of this section andeach system as a whole must meet therequirements of §§ 23.1387 through23.1397.
(b) Left and right position lights. Leftand right position lights must consistof a red and a green light spaced lat-erally as far apart as practicable andinstalled on the airplane such that,with the airplane in the normal flyingposition, the red light is on the leftside and the green light is on the rightside.
(c) Rear position light. The rear posi-tion light must be a white light mount-ed as far aft as practicable on the tailor on each wing tip.
(d) Light covers and color filters. Eachlight cover or color filter must be atleast flame resistant and may notchange color or shape or lose any ap-preciable light transmission duringnormal use.
(a) Except as provided in paragraph(e) of this section, each position lightmust, as installed, show unbroken lightwithin the dihedral angles described inthis section.
(b) Dihedral angle L (left) is formedby two intersecting vertical planes, thefirst parallel to the longitudinal axis ofthe airplane, and the other at 110 de-grees to the left of the first, as viewedwhen looking forward along the longi-tudinal axis.
(c) Dihedral angle R (right) is formedby two intersecting vertical planes, thefirst parallel to the longitudinal axis ofthe airplane, and the other at 110 de-grees to the right of the first, as viewedwhen looking forward along the longi-tudinal axis.
(d) Dihedral angle A (aft) is formedby two intersecting vertical planesmaking angles of 70 degrees to theright and to the left, respectively, to avertical plane passing through the lon-gitudinal axis, as viewed when lookingaft along the longitudinal axis.
(e) If the rear position light, whenmounted as far aft as practicable in ac-cordance with § 23.1385(c), cannot showunbroken light within dihedral angle A(as defined in paragraph (d) of this sec-tion), a solid angle or angles of ob-structed visibility totaling not morethan 0.04 steradians is allowable withinthat dihedral angle, if such solid angleis within a cone whose apex is at therear position light and whose elementsmake an angle of 30° with a verticalline passing through the rear positionlight.
§ 23.1389 Position light distributionand intensities.
(a) General. The intensities prescribedin this section must be provided by newequipment with each light cover andcolor filter in place. Intensities mustbe determined with the light source op-erating at a steady value equal to theaverage luminous output of the sourceat the normal operating voltage of theairplane. The light distribution and in-tensity of each position light mustmeet the requirements of paragraph (b)of this section.
(b) Position lights. The light distribu-tion and intensities of position lightsmust be expressed in terms of min-imum intensities in the horizontalplane, minimum intensities in anyvertical plane, and maximum inten-sities in overlapping beams, within di-hedral angles L, R, and A, and mustmeet the following requirements:
(1) Intensities in the horizontal plane.Each intensity in the horizontal plane(the plane containing the longitudinal
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axis of the airplane and perpendicularto the plane of symmetry of the air-plane) must equal or exceed the valuesin § 23.1391.
(2) Intensities in any vertical plane.Each intensity in any vertical plane(the plane perpendicular to the hori-zontal plane) must equal or exceed theappropriate value in § 23.1393, where I isthe minimum intensity prescribed in§ 23.1391 for the corresponding angles inthe horizontal plane.
(3) Intensities in overlaps between adja-cent signals. No intensity in any over-lap between adjacent signals may ex-ceed the values in § 23.1395, except thathigher intensities in overlaps may beused with main beam intensities sub-stantially greater than the minimaspecified in §§ 23.1391 and 23.1393, if theoverlap intensities in relation to themain beam intensities do not adverselyaffect signal clarity. When the peak in-tensity of the left and right positionlights is more than 100 candles, themaximum overlap intensities betweenthem may exceed the values in § 23.1395if the overlap intensity in Area A isnot more than 10 percent of peak posi-tion light intensity and the overlap in-tensity in Area B is not more than 2.5percent of peak position light inten-sity.
(c) Rear position light installation. Asingle rear position light may be in-stalled in a position displaced laterallyfrom the plane of symmetry of an air-plane if—
(1) The axis of the maximum cone ofillumination is parallel to the flightpath in level flight; and
(2) There is no obstruction aft of thelight and between planes 70 degrees tothe right and left of the axis of max-imum illumination.
§ 23.1393 Minimum intensities in anyvertical plane of position lights.
Each position light intensity mustequal or exceed the applicable values inthe following table:
Angle above or below the horizontal plane Intensity, l
0° ......................................................................... 1.000° to 5° ................................................................ 0.905° to 10° .............................................................. 0.8010° to 15° ............................................................ 0.7015° to 20° ............................................................ 0.5020° to 30° ............................................................ 0.3030° to 40° ............................................................ 0.1040° to 90° ............................................................ 0.05
§ 23.1395 Maximum intensities in over-lapping beams of position lights.
No position light intensity may ex-ceed the applicable values in the fol-lowing equal or exceed the applicablevalues in § 23.1389(b)(3):
Overlaps
Maximum intensity
Area A(candles)
Area B(candles)
Green in dihedral angle L ............. 10 1Red in dihedral angle R ................ 10 1Green in dihedral angle A ............. 5 1Red in dihedral angle A ................ 5 1Rear white in dihedral angle L ...... 5 1Rear white in dihedral angle R ..... 5 1
Where—(a) Area A includes all directions in
the adjacent dihedral angle that passthrough the light source and intersectthe common boundary plane at morethan 10 degrees but less than 20 de-grees; and
(b) Area B includes all directions inthe adjacent dihedral angle that passthrough the light source and intersect
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§ 23.1397 Color specifications.Each position light color must have
the applicable International Commis-sion on Illumination chromaticity co-ordinates as follows:
(a) Aviation red—
‘‘y’’ is not greater than 0.335; and‘‘z’’ is not greater than 0.002.
(b) Aviation green—
‘‘x’’ is not greater than 0.440¥0.320 y;‘‘x’’ is not greater than y ¥0.170; and‘‘y’’ is not less than 0.390¥0.170 x.
(c) Aviation white—
‘‘x’’ is not less than 0.300 and not greaterthan 0.540;
‘‘y’’ is not less than ‘‘x ¥0.040’’ or ‘‘y0
¥0.010,’’ whichever is the smaller; and‘‘y’’ is not greater than ‘‘x+0.020’’ nor
‘‘0.636¥0.400 x ’’;Where ‘‘y0’’ is the ‘‘y’’ coordinate of the
Planckian radiator for the value of ‘‘x’’ con-sidered.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964,amended by Amdt. 23–11, 36 FR 12971, July 10,1971]
§ 23.1399 Riding light.(a) Each riding (anchor) light re-
quired for a seaplane or amphibian,must be installed so that it can—
(1) Show a white light for at leasttwo miles at night under clear atmos-pheric conditions; and
(2) Show the maximum unbrokenlight practicable when the airplane ismoored or drifting on the water.
(b) Externally hung lights may beused.
§ 23.1401 Anticollision light system.(a) General. The airplane must have
an anticollision light system that:(1) Consists of one or more approved
anticollision lights located so thattheir light will not impair the flightcrewmembers’ vision or detract fromthe conspicuity of the position lights;and
(2) Meets the requirements of para-graphs (b) through (f) of this section.
(b) Field of coverage. The system mustconsist of enough lights to illuminate
the vital areas around the airplane,considering the physical configurationand flight characteristics of the air-plane. The field of coverage must ex-tend in each direction within at least75 degrees above and 75 degrees belowthe horizontal plane of the airplane,except that there may be solid anglesof obstructed visibility totaling notmore than 0.5 steradians.
(c) Flashing characteristics. The ar-rangement of the system, that is, thenumber of light sources, beam width,speed of rotation, and other character-istics, must give an effective flash fre-quency of not less than 40, nor morethan 100, cycles per minute. The effec-tive flash frequency is the frequency atwhich the airplane’s complete anti-collision light system is observed froma distance, and applies to each sectorof light including any overlaps thatexist when the system consists of morethan one light source. In overlaps,flash frequencies may exceed 100, butnot 180, cycles per minute.
(d) Color. Each anticollision lightmust be either aviation red or aviationwhite and must meet the applicable re-quirements of § 23.1397.
(e) Light intensity. The minimumlight intensities in any vertical plane,measured with the red filter (if used)and expressed in terms of ‘‘effective’’intensities, must meet the require-ments of paragraph (f) of this section.The following relation must be as-sumed:
II t dt
t te
t
t
=( )
+ −( )∫
1
2
0 2 2 1.where:
Ie =effective intensity (candles).I(t) =instantaneous intensity as a function of
time.t2¥t1 =flash time interval (seconds).
Normally, the maximum value of effec-tive intensity is obtained when t2 and t1are chosen so that the effective inten-sity is equal to the instantaneous in-tensity at t2 and t1.
(f) Minimum effective intensities foranticollision lights. Each anticollisionlight effective intensity must equal orexceed the applicable values in the fol-lowing table.
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Angle above or below the horizontal planeEffective in-tensity (can-
dles)
0° to 5° .............................................................. 4005° to 10° ............................................................ 24010° to 20° .......................................................... 8020° to 30° .......................................................... 4030° to 75° .......................................................... 20
(a) Required safety equipment to beused by the flight crew in an emer-gency, such as automatic liferaft re-leases, must be readily accessible.
(b) Stowage provisions for requiredsafety equipment must be furnishedand must—
(1) Be arranged so that the equip-ment is directly accessible and its loca-tion is obvious; and
(2) Protect the safety equipmentfrom damage caused by being subjectedto the inertia loads resulting from theultimate static load factors specified in§ 23.561(b)(3) of this part.
[Amdt. 23–17, 41 FR 55465, Dec. 20, 1976, asamended by Amdt. 23–36, 53 FR 30815, Aug. 15,1988]
§ 23.1415 Ditching equipment.
(a) Emergency flotation and sig-naling equipment required by any oper-ating rule in this chapter must be in-stalled so that it is readily available tothe crew and passengers.
(b) Each raft and each life preservermust be approved.
(c) Each raft released automaticallyor by the pilot must be attached to theairplane by a line to keep it alongsidethe airplane. This line must be weakenough to break before submerging theempty raft to which it is attached.
(d) Each signaling device required byany operating rule in this chapter,must be accessible, function satisfac-torily, and must be free of any hazardin its operation.
§ 23.1416 Pneumatic de-icer boot sys-tem.
If certification with ice protectionprovisions is desired and a pneumaticde-icer boot system is installed—
(a) The system must meet the re-quirements specified in § 23.1419.
(b) The system and its componentsmust be designed to perform their in-tended function under any normal sys-tem operating temperature or pressure,and
(c) Means to indicate to the flightcrew that the pneumatic de-icer bootsystem is receiving adequate pressureand is functioning normally must beprovided.
[Amdt. 23–23, 43 FR 50593, Oct. 30, 1978]
§ 23.1419 Ice protection.If certification with ice protection
provisions is desired, compliance withthe requirements of this section andother applicable sections of this partmust be shown:
(a) An analysis must be performed toestablish, on the basis of the airplane’soperational needs, the adequacy of theice protection system for the variouscomponents of the airplane. In addi-tion, tests of the ice protection systemmust be conducted to demonstrate thatthe airplane is capable of operatingsafely in continuous maximum andintermittent maximum icing condi-tions, as described in appendix C ofpart 25 of this chapter. As used in thissection, ‘‘Capable of operating safely,’’means that airplane performance, con-trollability, maneuverability, and sta-bility must not be less than that re-quired in part 23, subpart B.
(b) Except as provided by paragraph(c) of this section, in addition to theanalysis and physical evaluation pre-scribed in paragraph (a) of this section,the effectiveness of the ice protectionsystem and its components must beshown by flight tests of the airplane orits components in measured natural at-mospheric icing conditions and by oneor more of the following tests, as foundnecessary to determine the adequacy ofthe ice protection system—
(1) Laboratory dry air or simulatedicing tests, or a combination of both, of
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Federal Aviation Administration, DOT § 23.1435
the components or models of the com-ponents.
(2) Flight dry air tests of the ice pro-tection system as a whole, or its indi-vidual components.
(3) Flight test of the airplane or itscomponents in measured simulatedicing conditions.
(c) If certification with ice protectionhas been accomplished on prior typecertificated airplanes whose designs in-clude components that arethermodynamically and aero-dynamically equivalent to those usedon a new airplane design, certificationof these equivalent components may beaccomplished by reference to pre-viously accomplished tests, required in§ 23.1419 (a) and (b), provided that theapplicant accounts for any differencesin installation of these components.
(d) A means must be identified orprovided for determining the formationof ice on the critical parts of the air-plane. Adequate lighting must be pro-vided for the use of this means duringnight operation. Also, when monitoringof the external surfaces of the airplaneby the flight crew is required for oper-ation of the ice protection equipment,external lighting must be provided thatis adequate to enable the monitoring tobe done at night. Any illuminationthat is used must be of a type that willnot cause glare or reflection thatwould handicap crewmembers in theperformance of their duties. The Air-plane Flight Manual or other approvedmanual material must describe themeans of determining ice formationand must contain information for thesafe operation of the airplane in icingconditions.
[Doc. No. 26344, 58 FR 18977, Apr. 9, 1993]
MISCELLANEOUS EQUIPMENT
§ 23.1431 Electronic equipment.
(a) In showing compliance with§ 23.1309(b)(1) and (2) with respect toradio and electronic equipment andtheir installations, critical environ-mental conditions must be considered.
(b) Radio and electronic equipment,controls, and wiring must be installedso that operation of any unit or systemof units will not adversely affect the si-multaneous operation of any other
radio or electronic unit, or system ofunits, required by this chapter.
(c) For those airplanes required tohave more than one flightcrew mem-ber, or whose operation will requiremore than one flightcrew member, thecockpit must be evaluated to deter-mine if the flightcrew members, whenseated at their duty station, can con-verse without difficulty under the ac-tual cockpit noise conditions when theairplane is being operated. If the air-plane design includes provision for theuse of communication headsets, theevaluation must also consider condi-tions where headsets are being used. Ifthe evaluation shows conditions underwhich it will be difficult to converse,an intercommunication system mustbe provided.
(d) If installed communication equip-ment includes transmitter ‘‘off-on’’switching, that switching means mustbe designed to return from the ‘‘trans-mit’’ to the ‘‘off’’ position when it isreleased and ensure that the trans-mitter will return to the off (non trans-mitting) state.
(e) If provisions for the use of com-munication headsets are provided, itmust be demonstrated that theflightcrew members will receive allaural warnings under the actual cock-pit noise conditions when the airplaneis being operated when any headset isbeing used.
(a) Design. Each hydraulic systemmust be designed as follows:
(1) Each hydraulic system and its ele-ments must withstand, without yield-ing, the structural loads expected inaddition to hydraulic loads.
(2) A means to indicate the pressurein each hydraulic system which sup-plies two or more primary functionsmust be provided to the flight crew.
(3) There must be means to ensurethat the pressure, including transient(surge) pressure, in any part of the sys-tem will not exceed the safe limitabove design operating pressure and toprevent excessive pressure resultingfrom fluid volumetric changes in all
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14 CFR Ch. I (1–1–01 Edition)§ 23.1437
lines which are likely to remain closedlong enough for such changes to occur.
(4) The minimum design burst pres-sure must be 2.5 times the operatingpressure.
(b) Tests. Each system must be sub-stantiated by proof pressure tests.When proof tested, no part of any sys-tem may fail, malfunction, or experi-ence a permanent set. The proof load ofeach system must be at least 1.5 timesthe maximum operating pressure ofthat system.
(c) Accumulators. A hydraulic accu-mulator or reservoir may be installedon the engine side of any firewall if—
(1) It is an integral part of an engineor propeller system, or
(2) The reservoir is nonpressurizedand the total capacity of all such non-pressurized reservoirs is one quart orless.
For multiengine airplanes, engine-driven accessories essential to safe op-eration must be distributed among twoor more engines so that the failure ofany one engine will not impair safe op-eration through the malfunctioning ofthese accessories.
§ 23.1438 Pressurization and pneu-matic systems.
(a) Pressurization system elementsmust be burst pressure tested to 2.0times, and proof pressure tested to 1.5times, the maximum normal operatingpressure.
(b) Pneumatic system elements mustbe burst pressure tested to 3.0 times,and proof pressure tested to 1.5 times,the maximum normal operating pres-sure.
(c) An analysis, or a combination ofanalysis and test, may be substitutedfor any test required by paragraph (a)or (b) of this section if the Adminis-trator finds it equivalent to the re-quired test.
[Amdt. 23–20, 42 FR 36969, July 18, 1977]
§ 23.1441 Oxygen equipment and sup-ply.
(a) If certification with supplementaloxygen equipment is requested, or theairplane is approved for operations ator above altitudes where oxygen is re-quired to be used by the operatingrules, oxygen equipment must be pro-vided that meets the requirements ofthis section and §§ 23.1443 through23.1449. Portable oxygen equipmentmay be used to meet the requirementsof this part if the portable equipmentis shown to comply with the applicablerequirements, is identified in the air-plane type design, and its stowage pro-visions are found to be in compliancewith the requirements of § 23.561.
(b) The oxygen system must be freefrom hazards in itself, in its method ofoperation, and its effect upon othercomponents.
(c) There must be a means to allowthe crew to readily determine, duringthe flight, the quantity of oxygenavailable in each source of supply.
(d) Each required flight crewmembermust be provided with—
(1) Demand oxygen equipment if theairplane is to be certificated for oper-ation above 25,000 feet.
(2) Pressure demand oxygen equip-ment if the airplane is to be certifi-cated for operation above 40,000 feet.
(e) There must be a means, readilyavailable to the crew in flight, to turnon and to shut off the oxygen supply atthe high pressure source. This shutoffrequirement does not apply to chem-ical oxygen generators.
§ 23.1443 Minimum mass flow of sup-plemental oxygen.
(a) If continuous flow oxygen equip-ment is installed, an applicant mustshow compliance with the require-ments of either paragraphs (a)(1) and(a)(2) or paragraph (a)(3) of this sec-tion:
(1) For each passenger, the minimummass flow of supplemental oxygen re-quired at various cabin pressure alti-tudes may not be less than the flow re-quired to maintain, during inspirationand while using the oxygen equipment
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Federal Aviation Administration, DOT § 23.1443
(including masks) provided, the fol-lowing mean tracheal oxygen partialpressures:
(i) At cabin pressure altitudes above10,000 feet up to and including 18,500feet, a mean tracheal oxygen partialpressure of 100 mm. Hg when breathing15 liters per minute, Body Tempera-ture, Pressure, Saturated (BTPS) andwith a tidal volume of 700 cc. with aconstant time interval between res-pirations.
(ii) At cabin pressure altitudes above18,500 feet up to and including 40,000feet, a mean tracheal oxygen partialpressure of 83.8 mm. Hg when breathing30 liters per minute, BTPS, and with atidal volume of 1,100 cc. with a con-
stant time interval between respira-tions.
(2) For each flight crewmember, theminimum mass flow may not be lessthan the flow required to maintain,during inspiration, a mean tracheal ox-ygen partial pressure of 149 mm. Hgwhen breathing 15 liters per minute,BTPS, and with a maximum tidal vol-ume of 700 cc. with a constant time in-terval between respirations.
(3) The minimum mass flow of sup-plemental oxygen supplied for eachuser must be at a rate not less thanthat shown in the following figure foreach altitude up to and including themaximum operating altitude of the air-plane.
(b) If demand equipment is installedfor use by flight crewmembers, theminimum mass flow of supplementaloxygen required for each flight crew-member may not be less than the flowrequired to maintain, during inspira-tion, a mean tracheal oxygen partialpressure of 122 mm. Hg up to and in-cluding a cabin pressure altitude of35,000 feet, and 95 percent oxygen be-tween cabin pressure altitudes of 35,000and 40,000 feet, when breathing 20 litersper minute BTPS. In addition, theremust be means to allow the crew to useundiluted oxygen at their discretion.
(c) If first-aid oxygen equipment isinstalled, the minimum mass flow ofoxygen to each user may not be lessthan 4 liters per minute, STPD. How-ever, there may be a means to decreasethis flow to not less than 2 liters perminute, STPD, at any cabin altitude.The quantity of oxygen required isbased upon an average flow rate of 3 li-ters per minute per person for whomfirst-aid oxygen is required.
(d) As used in this section:(1) BTPS means Body Temperature,
and Pressure, Saturated (which is, 37°C, and the ambient pressure to whichthe body is exposed, minus 47 mm. Hg,
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14 CFR Ch. I (1–1–01 Edition)§ 23.1445
which is the tracheal pressure dis-placed by water vapor pressure whenthe breathed air becomes saturatedwith water vapor at 37 °C).
(2) STPD means Standard, Tempera-ture, and Pressure, Dry (which is, 0 °Cat 760 mm. Hg with no water vapor).
[Doc. No. 26344, 58 FR 18978, Apr. 9, 1993]
§ 23.1445 Oxygen distribution system.
(a) Except for flexible lines from oxy-gen outlets to the dispensing units, orwhere shown to be otherwise suitableto the installation, nonmetallic tubingmust not be used for any oxygen linethat is normally pressurized duringflight.
(b) Nonmetallic oxygen distributionlines must not be routed where theymay be subjected to elevated tempera-tures, electrical arcing, and releasedflammable fluids that might resultfrom any probable failure.
[Doc. No. 26344, 58 FR 18978, Apr. 9, 1993]
§ 23.1447 Equipment standards for ox-ygen dispensing units.
If oxygen dispensing units are in-stalled, the following apply:
(a) There must be an individual dis-pensing unit for each occupant forwhom supplemental oxygen is to besupplied. Each dispensing unit must:
(1) Provide for effective utilization ofthe oxygen being delivered to the unit.
(2) Be capable of being readily placedinto position on the face of the user.
(3) Be equipped with a suitable meansto retain the unit in position on theface.
(4) If radio equipment is installed,the flightcrew oxygen dispensing unitsmust be designed to allow the use ofthat equipment and to allow commu-nication with any other required crewmember while at their assigned dutystation.
(b) If certification for operation up toand including 18,000 feet (MSL) is re-quested, each oxygen dispensing unitmust:
(1) Cover the nose and mouth of theuser; or
(2) Be a nasal cannula, in which caseone oxygen dispensing unit coveringboth the nose and mouth of the usermust be available. In addition, each
nasal cannula or its connecting tubingmust have permanently affixed—
(i) A visible warning against smokingwhile in use;
(ii) An illustration of the correctmethod of donning; and
(iii) A visible warning against usewith nasal obstructions or head coldswith resultant nasal congestion.
(c) If certification for operationabove 18,000 feet (MSL) is requested,each oxygen dispensing unit mustcover the nose and mouth of the user.
(d) For a pressurized airplane de-signed to operate at flight altitudesabove 25,000 feet (MSL), the dispensingunits must meet the following:
(1) The dispensing units for pas-sengers must be connected to an oxy-gen supply terminal and be imme-diately available to each occupantwherever seated.
(2) The dispensing units for crew-members must be automatically pre-sented to each crewmember before thecabin pressure altitude exceeds 15,000feet, or the units must be of the quick-donning type, connected to an oxygensupply terminal that is immediatelyavailable to crewmembers at their sta-tion.
(e) If certification for operationabove 30,000 feet is requested, the dis-pensing units for passengers must beautomatically presented to each occu-pant before the cabin pressure altitudeexceeds 15,000 feet.
(f) If an automatic dispensing unit(hose and mask, or other unit) systemis installed, the crew must be providedwith a manual means to make the dis-pensing units immediately available inthe event of failure of the automaticsystem.
There must be a means to allow thecrew to determine whether oxygen isbeing delivered to the dispensing equip-ment.
[Amdt. 23–9, 35 FR 6387, Apr. 21, 1970]
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Federal Aviation Administration, DOT § 23.1457
§ 23.1450 Chemical oxygen generators.(a) For the purpose of this section, a
chemical oxygen generator is definedas a device which produces oxygen bychemical reaction.
(b) Each chemical oxygen generatormust be designed and installed in ac-cordance with the following require-ments:
(1) Surface temperature developed bythe generator during operation maynot create a hazard to the airplane orto its occupants.
(2) Means must be provided to relieveany internal pressure that may be haz-ardous.
(c) In addition to meeting the re-quirements in paragraph (b) of this sec-tion, each portable chemical oxygengenerator that is capable of sustainedoperation by successive replacement ofa generator element must be placardedto show—
(1) The rate of oxygen flow, in litersper minute;
(2) The duration of oxygen flow, inminutes, for the replaceable generatorelement; and
(3) A warning that the replaceablegenerator element may be hot, unlessthe element construction is such thatthe surface temperature cannot exceed100° F.
[Amdt. 23–20, 42 FR 36969, July 18, 1977]
§ 23.1451 Fire protection for oxygenequipment.
Oxygen equipment and lines must:(a) Not be installed in any designed
fire zones.(b) Be protected from heat that may
be generated in, or escape from, anydesignated fire zone.
(c) Be installed so that escaping oxy-gen cannot come in contact with andcause ignition of grease, fluid, or vaporaccumulations that are present in nor-mal operation or that may result fromthe failure or malfunction of any othersystem.
[Doc. No. 27806, 61 FR 5170, Feb. 9, 1996]
§ 23.1453 Protection of oxygen equip-ment from rupture.
(a) Each element of the oxygen sys-tem must have sufficient strength towithstand the maximum pressure andtemperature, in combination with any
externally applied loads arising fromconsideration of limit structural loads,that may be acting on that part of thesystem.
(b) Oxygen pressure sources and thelines between the source and the shut-off means must be:
(1) Protected from unsafe tempera-tures; and
(2) Located where the probability andhazard of rupture in a crash landingare minimized.
[Doc. No. 27806, 61 FR 5170, Feb. 9, 1996]
§ 23.1457 Cockpit voice recorders.(a) Each cockpit voice recorder re-
quired by the operating rules of thischapter must be approved and must beinstalled so that it will record the fol-lowing:
(1) Voice communications trans-mitted from or received in the airplaneby radio.
(2) Voice communications of flightcrewmembers on the flight deck.
(3) Voice communications of flightcrewmembers on the flight deck, usingthe airplane’s interphone system.
(4) Voice or audio signals identifyingnavigation or approach aids introducedinto a headset or speaker.
(5) Voice communications of flightcrewmembers using the passenger loud-speaker system, if there is such a sys-tem and if the fourth channel is avail-able in accordance with the require-ments of paragraph (c)(4)(ii) of this sec-tion.
(b) The recording requirements ofparagraph (a)(2) of this section must bemet by installing a cockpit-mountedarea microphone, located in the bestposition for recording voice commu-nications originating at the first andsecond pilot stations and voice commu-nications of other crewmembers on theflight deck when directed to those sta-tions. The microphone must be so lo-cated and, if necessary, the pre-amplifiers and filters of the recordermust be so adjusted or supplemented,so that the intelligibility of the re-corded communications is as high aspracticable when recorded under flightcockpit noise conditions and playedback. Repeated aural or visual play-back of the record may be used in eval-uating intelligibility.
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14 CFR Ch. I (1–1–01 Edition)§ 23.1459
(c) Each cockpit voice recorder mustbe installed so that the part of thecommunication or audio signals speci-fied in paragraph (a) of this section ob-tained from each of the followingsources is recorded on a separate chan-nel:
(1) For the first channel, from eachboom, mask, or handheld microphone,headset, or speaker used at the firstpilot station.
(2) For the second channel from eachboom, mask, or handheld microphone,headset, or speaker used at the secondpilot station.
(3) For the third channel—from thecockpit-mounted area microphone.
(4) For the fourth channel from:(i) Each boom, mask, or handheld
microphone, headset, or speaker usedat the station for the third and fourthcrewmembers.
(ii) If the stations specified in para-graph (c)(4)(i) of this section are not re-quired or if the signal at such a stationis picked up by another channel, eachmicrophone on the flight deck that isused with the passenger loudspeakersystem, if its signals are not picked upby another channel.
(5) And that as far as is practicableall sounds received by the microphonelisted in paragraphs (c)(1), (2), and (4) ofthis section must be recorded withoutinterruption irrespective of the posi-tion of the interphone-transmitter keyswitch. The design shall ensure thatsidetone for the flight crew is producedonly when the interphone, public ad-dress system, or radio transmitters arein use.
(d) Each cockpit voice recorder mustbe installed so that:
(1) It receives its electric power fromthe bus that provides the maximum re-liability for operation of the cockpitvoice recorder without jeopardizingservice to essential or emergencyloads.
(2) There is an automatic means tosimultaneously stop the recorder andprevent each erasure feature from func-tioning, within 10 minutes after crashimpact; and
(3) There is an aural or visual meansfor preflight checking of the recorderfor proper operation.
(e) The record container must be lo-cated and mounted to minimize the
probability of rupture of the containeras a result of crash impact and con-sequent heat damage to the recordfrom fire. In meeting this requirement,the record container must be as far aftas practicable, but may not be whereaft mounted engines may crush thecontainer during impact. However, itneed not be outside of the pressurizedcompartment.
(f) If the cockpit voice recorder has abulk erasure device, the installationmust be designed to minimize the prob-ability of inadvertent operation and ac-tuation of the device during crash im-pact.
(g) Each recorder container must:(1) Be either bright orange or bright
yellow;(2) Have reflective tape affixed to its
external surface to facilitate its loca-tion under water; and
(3) Have an underwater locating de-vice, when required by the operatingrules of this chapter, on or adjacent tothe container which is secured in suchmanner that they are not likely to beseparated during crash impact.
[Amdt. 23–35, 53 FR 26142, July 11, 1988]
§ 23.1459 Flight recorders.
(a) Each flight recorder required bythe operating rules of this chaptermust be installed so that:
(1) It is supplied with airspeed, alti-tude, and directional data obtainedfrom sources that meet the accuracyrequirements of §§ 23.1323, 23.1325, and23.1327, as appropriate;
(2) The vertical acceleration sensor isrigidly attached, and located longitu-dinally either within the approved cen-ter of gravity limits of the airplane, orat a distance forward or aft of theselimits that does not exceed 25 percentof the airplane’s mean aerodynamicchord;
(3) It receives its electrical powerpower from the bus that provides themaximum reliability for operation ofthe flight recorder without jeopard-izing service to essential or emergencyloads;
(4) There is an aural or visual meansfor preflight checking of the recorderfor proper recording of data in the stor-age medium.
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Federal Aviation Administration, DOT § 23.1505
(5) Except for recorders powered sole-ly by the engine-driven electrical gen-erator system, there is an automaticmeans to simultaneously stop a re-corder that has a data erasure featureand prevent each erasure feature fromfunctioning, within 10 minutes aftercrash impact; and
(b) Each nonejectable record con-tainer must be located and mounted soas to minimize the probability of con-tainer rupture resulting from crash im-pact and subsequent damage to therecord from fire. In meeting this re-quirement the record container mustbe located as far aft as practicable, butneed not be aft of the pressurized com-partment, and may not be where aft-mounted engines may crush the con-tainer upon impact.
(c) A correlation must be establishedbetween the flight recorder readings ofairspeed, altitude, and heading and thecorresponding readings (taking into ac-count correction factors) of the first pi-lot’s instruments. The correlationmust cover the airspeed range overwhich the airplane is to be operated,the range of altitude to which the air-plane is limited, and 360 degrees ofheading. Correlation may be estab-lished on the ground as appropriate.
(d) Each recorder container must:(1) Be either bright orange or bright
yellow;(2) Have reflective tape affixed to its
external surface to facilitate its loca-tion under water; and
(3) Have an underwater locating de-vice, when required by the operatingrules of this chapter, on or adjacent tothe container which is secured in sucha manner that they are not likely to beseparated during crash impact.
(e) Any novel or unique design oroperational characteristics of the air-craft shall be evaluated to determine ifany dedicated parameters must be re-corded on flight recorders in additionto or in place of existing requirements.CITA≤[Amdt. 23–35, 53 FR 26143, July11, 1988]
§ 23.1461 Equipment containing highenergy rotors.
(a) Equipment, such as AuxiliaryPower Units (APU) and constant speeddrive units, containing high energy ro-
tors must meet paragraphs (b), (c), or(d) of this section.
(b) High energy rotors contained inequipment must be able to withstanddamage caused by malfunctions, vibra-tion, abnormal speeds, and abnormaltemperatures. In addition—
(1) Auxiliary rotor cases must be ableto contain damage caused by the fail-ure of high energy rotor blades; and
(2) Equipment control devices, sys-tems, and instrumentation must rea-sonably ensure that no operating limi-tations affecting the integrity of highenergy rotors will be exceeded in serv-ice.
(c) It must be shown by test thatequipment containing high energy ro-tors can contain any failure of a highenergy rotor that occurs at the highestspeed obtainable with the normal speedcontrol devices inoperative.
(d) Equipment containing high en-ergy rotors must be located whererotor failure will neither endanger theoccupants nor adversely affect contin-ued safe flight.
[Amdt. 23–20, 42 FR 36969, July 18, 1977, asamended by Amdt. 23–49, 61 FR 5170, Feb. 9,1996]
Subpart G—Operating Limitationsand Information
§ 23.1501 General.
(a) Each operating limitation speci-fied in §§ 23.1505 through 23.1527 andother limitations and information nec-essary for safe operation must be es-tablished.
(b) The operating limitations andother information necessary for safeoperation must be made available tothe crewmembers as prescribed in§§ 23.1541 through 23.1589.
[Amdt. 23–21, 43 FR 2319, Jan. 16, 1978]
§ 23.1505 Airspeed limitations.
(a) The never-exceed speed VNE mustbe established so that it is—
(1) Not less than 0.9 times the min-imum value of VD allowed under§ 23.335; and
(2) Not more than the lesser of—(i) 0.9 VD established under § 23.335; or(ii) 0.9 times the maximum speed
shown under § 23.251.
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(b) The maximum structural cruisingspeed VNO must be established so thatit is—
(1) Not less than the minimum valueof VC allowed under § 23.335; and
(2) Not more than the lesser of—(i) VC established under § 23.335; or(ii) 0.89 VNE established under para-
graph (a) of this section.(c) Paragraphs (a) and (b) of this sec-
tion do not apply to turbine airplanesor to airplanes for which a design div-ing speed VD/MD is established under§ 23.335(b)(4). For those airplanes, amaximum operating limit speed (VMO/MMO-airspeed or Mach number, which-ever is critical at a particular altitude)must be established as a speed thatmay not be deliberately exceeded inany regime of flight (climb, cruise, ordescent) unless a higher speed is au-thorized for flight test or pilot trainingoperations. VMO/MMO must be estab-lished so that it is not greater than thedesign cruising speed VC/MC and so thatit is sufficiently below VD/MD and themaximum speed shown under § 23.251 tomake it highly improbable that thelatter speeds will be inadvertently ex-ceeded in operations. The speed marginbetween VMO/MMO and VD/MD or themaximum speed shown under § 23.251may not be less than the speed marginestablished between VC/MC and VD/MD
under § 23.335(b), or the speed marginfound necessary in the flight test con-ducted under § 23.253.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–7, 34 FR 13096, Aug. 13,1969]
§ 23.1507 Operating maneuveringspeed.
The maximum operating maneu-vering speed, VO, must be establishedas an operating limitation. VO is a se-lected speed that is not greater thanVS√n established in § 23.335(c).
[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]
§ 23.1511 Flap extended speed.(a) The flap extended speed VFE must
be established so that it is—(1) Not less than the minimum value
of VF allowed in § 23.345(b); and(2) Not more than VF established
under § 23.345(a), (c), and (d).(i) VF established under § 23.345; or(ii) VF established under § 23.457.
(b) Additional combinations of flapsetting, airspeed, and engine powermay be established if the structure hasbeen proven for the corresponding de-sign conditions.
§ 23.1513 Minimum control speed.The minimum control speed VMC, de-
termined under § 23.149, must be estab-lished as an operating limitation.
§ 23.1519 Weight and center of gravity.The weight and center of gravity lim-
itations determined under § 23.23 mustbe established as operating limitations.
§ 23.1521 Powerplant limitations.(a) General. The powerplant limita-
tions prescribed in this section must beestablished so that they do not exceedthe corresponding limits for which theengines or propellers are type certifi-cated. In addition, other powerplantlimitations used in determining com-pliance with this part must be estab-lished.
(b) Takeoff operation. The powerplanttakeoff operation must be limited by—
(1) The maximum rotational speed(rpm);
(2) The maximum allowable manifoldpressure (for reciprocating engines);
(3) The maximum allowable gas tem-perature (for turbine engines);
(4) The time limit for the use of thepower or thrust corresponding to thelimitations established in paragraphs(b)(1) through (3) of this section; and
(5) The maximum allowable cylinderhead (as applicable), liquid coolant andoil temperatures.
(c) Continuous operation. The contin-uous operation must be limited by—
(1) The maximum rotational speed;(2) The maximum allowable manifold
pressure (for reciprocating engines);(3) The maximum allowable gas tem-
perature (for turbine engines); and(4) The maximum allowable cylinder
head, oil, and liquid coolant tempera-tures.
(d) Fuel grade or designation. The min-imum fuel grade (for reciprocating en-gines), or fuel designation (for turbineengines), must be established so that itis not less than that required for the
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Federal Aviation Administration, DOT § 23.1541
operation of the engines within thelimitations in paragraphs (b) and (c) ofthis section.
(e) Ambient temperature. For all air-planes except reciprocating engine-powered airplanes of 6,000 pounds orless maximum weight, ambient tem-perature limitations (including limita-tions for winterization installations ifapplicable) must be established as themaximum ambient atmospheric tem-perature at which compliance with thecooling provisions of §§ 23.1041 through23.1047 is shown.
If an auxiliary power unit is in-stalled, the limitations established forthe auxiliary power must be specifiedin the operating limitations for the air-plane.[Doc. No. 26269, 58 FR 42166, Aug.6, 1993]
§ 23.1523 Minimum flight crew.
The minimum flight crew must be es-tablished so that it is sufficient for safeoperation considering—
(a) The workload on individual crew-members and, in addition for com-muter category airplanes, each crew-member workload determination mustconsider the following:
(1) Flight path control,(2) Collision avoidance,(3) Navigation,(4) Communications,(5) Operation and monitoring of all
essential airplane systems,(6) Command decisions, and(7) The accessibility and ease of oper-
ation of necessary controls by the ap-propriate crewmember during all nor-mal and emergency operations when atthe crewmember flight station;
(b) The accessibility and ease of oper-ation of necessary controls by the ap-propriate crewmember; and
(c) The kinds of operation authorizedunder § 23.1525.
The maximum passenger seating con-figuration must be established.
[Amdt. 23–10, 36 FR 2864, Feb. 11, 1971]
§ 23.1525 Kinds of operation.The kinds of operation authorized
(e.g. VFR, IFR, day or night) and themeteorological conditions (e.g. icing)to which the operation of the airplaneis limited or from which it is prohib-ited, must be established appropriateto the installed equipment.
[Doc. No. 26269, 58 FR 42166, Aug. 6, 1993]
§ 23.1527 Maximum operating altitude.(a) The maximum altitude up to
which operation is allowed, as limitedby flight, structural, powerplant, func-tional or equipment characteristics,must be established.
(b) A maximum operating altitudelimitation of not more than 25,000 feetmust be established for pressurized air-planes unless compliance with§ 23.775(e) is shown.
[Doc. No. 26269, 58 FR 42166, Aug. 6, 1993]
§ 23.1529 Instructions for ContinuedAirworthiness.
The applicant must prepare Instruc-tions for Continued Airworthiness inaccordance with appendix G to thispart that are acceptable to the Admin-istrator. The instructions may be in-complete at type certification if a pro-gram exists to ensure their completionprior to delivery of the first airplane orissuance of a standard certificate ofairworthiness, whichever occurs later.
[Amdt. 23–26, 45 FR 60171, Sept. 11, 1980]
MARKINGS AND PLACARDS
§ 23.1541 General.(a) The airplane must contain—(1) The markings and placards speci-
fied in §§ 23.1545 through 23.1567; and(2) Any additional information, in-
strument markings, and placards re-quired for the safe operation if it hasunusual design, operating, or handlingcharacteristics.
(b) Each marking and placard pre-scribed in paragraph (a) of this sec-tion—
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(1) Must be displayed in a con-spicuous place; and
(2) May not be easily erased, dis-figured, or obscured.
(c) For airplanes which are to be cer-tificated in more than one category—
(1) The applicant must select one cat-egory upon which the placards andmarkings are to be based; and
(2) The placards and marking infor-mation for all categories in which theairplane is to be certificated must befurnished in the Airplane Flight Man-ual.
§ 23.1545 Airspeed indicator.(a) Each airspeed indicator must be
marked as specified in paragraph (b) ofthis section, with the marks located atthe corresponding indicated airspeeds.
(b) The following markings must bemade:
(1) For the never-exceed speed VNE, aradial red line.
(2) For the caution range, a yellowarc extending from the red line speci-fied in paragraph (b)(1) of this sectionto the upper limit of the green arcspecified in paragraph (b)(3) of this sec-tion.
(3) For the normal operating range, agreen arc with the lower limit at VS1
with maximum weight and with land-ing gear and wing flaps retracted, andthe upper limit at the maximum struc-tural cruising speed VNO establishedunder § 23.1505(b).
(4) For the flap operating range, awhite arc with the lower limit at VS0 at
the maximum weight, and the upperlimit at the flaps-extended speed VFE
established under § 23.1511.(5) For reciprocating multiengine-
powered airplanes of 6,000 pounds orless maximum weight, for the speed atwhich compliance has been shown with§ 23.69(b) relating to rate of climb atmaximum weight and at sea level, ablue radial line.
(6) For reciprocating multiengine-powered airplanes of 6,000 pounds orless maximum weight, for the max-imum value of minimum control speed,VMC, (one-engine-inoperative) deter-mined under § 23.149(b), a red radialline.
(c) If VNE or VNO vary with altitude,there must be means to indicate to thepilot the appropriate limitationsthroughout the operating altituderange.
(d) Paragraphs (b)(1) through (b)(3)and paragraph (c) of this section do notapply to aircraft for which a maximumoperating speed VMO/MMO is estab-lished under § 23.1505(c). For those air-craft there must either be a maximumallowable airspeed indication showingthe variation of VMO/MMO with altitudeor compressibility limitations (as ap-propriate), or a radial red line markingfor VMO/MMO must be made at lowestvalue of VMO/MMO established for anyaltitude up to the maximum operatingaltitude for the airplane.
(a) A placard meeting the require-ments of this section must be installedon or near the magnetic direction indi-cator.
(b) The placard must show the cali-bration of the instrument in levelflight with the engines operating.
(c) The placard must state whetherthe calibration was made with radio re-ceivers on or off.
(d) Each calibration reading must bein terms of magnetic headings in notmore than 30 degree increments.
(e) If a magnetic nonstabilized direc-tion indicator can have a deviation of
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Federal Aviation Administration, DOT § 23.1557
more than 10 degrees caused by the op-eration of electrical equipment, theplacard must state which electricalloads, or combination of loads, wouldcause a deviation of more than 10 de-grees when turned on.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30FR 258, Jan. 9, 1965, as amended by Amdt. 23–20, 42 FR 36969, July 18, 1977]
§ 23.1549 Powerplant and auxiliarypower unit instruments.
For each required powerplant andauxiliary power unit instrument, as ap-propriate to the type of instruments—
(a) Each maximum and, if applicable,minimum safe operating limit must bemarked with a red radial or a red line;
(b) Each normal operating rangemust be marked with a green arc orgreen line, not extending beyond themaximum and minimum safe limits;
(c) Each takeoff and precautionaryrange must be marked with a yellowarc or a yellow line; and
(d) Each engine, auxiliary powerunit, or propeller range that is re-stricted because of excessive vibrationstresses must be marked with red arcsor red lines.
[Amdt. 23–12, 41 FR 55466, Dec. 20, 1976, asamended by Amdt. 23–28, 47 FR 13315, Mar. 29,1982; Amdt. 23–45, 58 FR 42166, Aug. 6, 1993]
§ 23.1551 Oil quantity indicator.
Each oil quantity indicator must bemarked in sufficient increments to in-dicate readily and accurately the quan-tity of oil.
§ 23.1553 Fuel quantity indicator.
A red radial line must be marked oneach indicator at the calibrated zeroreading, as specified in § 23.1337(b)(1).
[Doc. No. 27807, 61 FR 5193, Feb. 9, 1996]
§ 23.1555 Control markings.
(a) Each cockpit control, other thanprimary flight controls and simplepush button type starter switches,must be plainly marked as to its func-tion and method of operation.
(b) Each secondary control must besuitably marked.
(c) For powerplant fuel controls—(1) Each fuel tank selector control
must be marked to indicate the posi-
tion corresponding to each tank and toeach existing cross feed position;
(2) If safe operation requires the useof any tanks in a specific sequence,that sequence must be marked on ornear the selector for those tanks;
(3) The conditions under which thefull amount of usable fuel in any re-stricted usage fuel tank can safely beused must be stated on a placard adja-cent to the selector valve for thattank; and
(4) Each valve control for any engineof a multiengine airplane must bemarked to indicate the position cor-responding to each engine controlled.
(d) Usable fuel capacity must bemarked as follows:
(1) For fuel systems having no selec-tor controls, the usable fuel capacity ofthe system must be indicated at thefuel quantity indicator.
(2) For fuel systems having selectorcontrols, the usable fuel capacityavailable at each selector control posi-tion must be indicated near the selec-tor control.
(e) For accessory, auxiliary, andemergency controls—
(1) If retractable landing gear is used,the indicator required by § 23.729 mustbe marked so that the pilot can, at anytime, ascertain that the wheels are se-cured in the extreme positions; and
(2) Each emergency control must bered and must be marked as to methodof operation. No control other than anemergency control, or a control thatserves an emergency function in addi-tion to its other functions, shall be thiscolor.
(a) Baggage and cargo compartments,and ballast location. Each baggage andcargo compartment, and each ballastlocation, must have a placard statingany limitations on contents, includingweight, that are necessary under theloading requirements.
(b) Seats. If the maximum allowableweight to be carried in a seat is lessthan 170 pounds, a placard stating the
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lesser weight must be permanently at-tached to the seat structure.
(c) Fuel, oil, and coolant filler open-ings. The following apply:
(1) Fuel filter openings must bemarked at or near the filler coverwith—
(i) For reciprocating engine-poweredairplanes—
(A) The word ‘‘Avgas’’; and(B) The minimum fuel grade.(ii) For turbine engine-powered air-
planes—(A) The words ‘‘Jet Fuel’’; and(B) The permissible fuel designations,
or references to the Airplane FlightManual (AFM) for permissible fuel des-ignations.
(iii) For pressure fueling systems, themaximum permissible fueling supplypressure and the maximum permissibledefueling pressure.
(2) Oil filler openings must bemarked at or near the filler cover withthe word ‘‘Oil’’ and the permissible oildesignations, or references to the Air-plane Flight Manual (AFM) for permis-sible oil designations.
(3) Coolant filler openings must bemarked at or near the filler cover withthe word ‘‘Coolant’’.
(d) Emergency exit placards. Eachplacard and operating control for eachemergency exit must be red. A placardmust be near each emergency exit con-trol and must clearly indicate the loca-tion of that exit and its method of op-eration.
(e) The system voltage of each directcurrent installation must be clearlymarked adjacent to its exernal powerconnection.
(a) There must be a placard in clearview of the pilot stating—
(1) That the airplane must be oper-ated in accordance with the AirplaneFlight Manual; and
(2) The certification category of theairplane to which the placards apply.
(b) For airplanes certificated in morethan one category, there must be aplacard in clear view of the pilot stat-
ing that other limitations are con-tained in the Airplane Flight Manual.
(c) There must be a placard in clearview of the pilot that specifies the kindof operations to which the operation ofthe airplane is limited or from which itis prohibited under § 23.1525.
[Doc. No. 27807, 61 FR 5193, Feb. 9, 1996]
§ 23.1561 Safety equipment.(a) Safety equipment must be plainly
marked as to method of operation.(b) Stowage provisions for required
safety equipment must be marked forthe benefit of occupants.
§ 23.1563 Airspeed placards.There must be an airspeed placard in
clear view of the pilot and as close aspracticable to the airspeed indicator.This placard must list—
(a) The operating maneuvering speed,VO; and
(b) The maximum landing gear oper-ating speed VLO.
(c) For reciprocating multiengine-powered airplanes of more than 6,000pounds maximum weight, and turbineengine-powered airplanes, the max-imum value of the minimum controlspeed, VMC (one-engine-inoperative) de-termined under § 23.149(b).
[Amdt. 23–7, 34 FR 13097, Aug. 13, 1969, asamended by Amdt. 23–45, 58 FR 42166, Aug. 6,1993; Amdt. 23–50, 61 FR 5193, Feb. 9, 1996]
§ 23.1567 Flight maneuver placard.(a) For normal category airplanes,
there must be a placard in front of andin clear view of the pilot stating: ‘‘Noacrobatic maneuvers, including spins,approved.’’
(b) For utility category airplanes,there must be—
(1) A placard in clear view of thepilot stating: ‘‘Acrobatic maneuversare limited to the following ———’’(list approved maneuvers and the rec-ommended entry speed for each); and
(2) For those airplanes that do notmeet the spin requirements for acro-batic category airplanes, an additionalplacard in clear view of the pilot stat-ing: ‘‘Spins Prohibited.’’
(c) For acrobatic category airplanes,there must be a placard in clear view ofthe pilot listing the approved acrobaticmaneuvers and the recommended entry
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Federal Aviation Administration, DOT § 23.1583
airspeed for each. If inverted flight ma-neuvers are not approved, the placardmust bear a notation to this effect.
(d) For acrobatic category airplanesand utility category airplanes approvedfor spinning, there must be a placard inclear view of the pilot—
(1) Listing the control actions for re-covery from spinning maneuvers; and
(2) Stating that recovery must be ini-tiated when spiral characteristics ap-pear, or after not more than six turnsor not more than any greater numberof turns for which the airplane hasbeen certificated.
AIRPLANE FLIGHT MANUAL ANDAPPROVED MANUAL MATERIAL
§ 23.1581 General.(a) Furnishing information. An Air-
plane Flight Manual must be furnishedwith each airplane, and it must containthe following:
(1) Information required by §§ 23.1583through 23.1589.
(2) Other information that is nec-essary for safe operation because of de-sign, operating, or handling character-istics.
(3) Further information necessary tocomply with the relevant operatingrules.
(b) Approved information. (1) Except asprovided in paragraph (b)(2) of this sec-tion, each part of the Airplane FlightManual containing information pre-scribed in §§ 23.1583 through 23.1589must be approved, segregated, identi-fied and clearly distinguished fromeach unapproved part of that AirplaneFlight Manual.
(2) The requirements of paragraph(b)(1) of this section do not apply to re-ciprocating engine-powered airplanesof 6,000 pounds or less maximumweight, if the following is met:
(i) Each part of the Airplane FlightManual containing information pre-scribed in § 23.1583 must be limited tosuch information, and must be ap-proved, identified, and clearly distin-guished from each other part of theAirplane Flight Manual.
(ii) The information prescribed in§§ 23.1585 through 23.1589 must be deter-mined in accordance with the applica-ble requirements of this part and pre-sented in its entirety in a manner ac-ceptable to the Administrator.
(3) Each page of the Airplane FlightManual containing information pre-scribed in this section must be of atype that is not easily erased, dis-figured, or misplaced, and is capable ofbeing inserted in a manual provided bythe applicant, or in a folder, or in anyother permanent binder.
(c) The units used in the AirplaneFlight Manual must be the same asthose marked on the appropriate in-struments and placards.
(d) All Airplane Flight Manual oper-ational airspeeds, unless otherwisespecified, must be presented as indi-cated airspeeds.
(e) Provision must be made for stow-ing the Airplane Flight Manual in asuitable fixed container which is read-ily accessible to the pilot.
(f) Revisions and amendments. EachAirplane Flight Manual (AFM) mustcontain a means for recording the in-corporation of revisions and amend-ments.
The Airplane Flight Manual mustcontain operating limitations deter-mined under this part 23, including thefollowing—
(a) Airspeed limitations. The followinginformation must be furnished:
(1) Information necessary for themarking of the airspeed limits on theindicator as required in § 23.1545, andthe significance of each of those limitsand of the color coding used on the in-dicator.
(2) The speeds VMC, VO, VLE, and VLO,if established, and their significance.
(3) In addition, for turbine poweredcommuter category airplanes—
(i) The maximum operating limitspeed, VMO/MMO and a statement thatthis speed must not be deliberately ex-ceeded in any regime of flight (climb,cruise or descent) unless a higher speed
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14 CFR Ch. I (1–1–01 Edition)§ 23.1583
is authorized for flight test or pilottraining;
(ii) If an airspeed limitation is basedupon compressibility effects, a state-ment to this effect and information asto any symptoms, the probable behav-ior of the airplane, and the rec-ommended recovery procedures; and
(iii) The airspeed limits must beshown in terms of VMO/MMO instead ofVNO and VNE.
(b) Powerplant limitations. The fol-lowing information must be furnished:
(1) Limitations required by § 23.1521.(2) Explanation of the limitations,
when appropriate.(3) Information necessary for mark-
ing the instruments required by§ 23.1549 through § 23.1553.
(c) Weight. The airplane flight man-ual must include—
(1) The maximum weight; and(2) The maximum landing weight, if
the design landing weight selected bythe applicant is less than the max-imum weight.
(3) For normal, utility, and acrobaticcategory reciprocating engine-poweredairplanes of more than 6,000 poundsmaximum weight and for turbine en-gine-powered airplanes in the normal,utility, and acrobatic category, per-formance operating limitations as fol-lows—
(i) The maximum takeoff weight foreach airport altitude and ambient tem-perature within the range selected bythe applicant at which the airplanecomplies with the climb requirementsof § 23.63(c)(1).
(ii) The maximum landing weight foreach airport altitude and ambient tem-perature within the range selected bythe applicant at which the airplanecomplies with the climb requirementsof § 23.63(c)(2).
(4) For commuter category airplanes,the maximum takeoff weight for eachairport altitude and ambient tempera-ture within the range selected by theapplicant at which—
(i) The airplane complies with theclimb requirements of § 23.63(d)(1); and
(ii) The accelerate-stop distance de-termined under § 23.55 is equal to theavailable runway length plus thelength of any stopway, if utilized; andeither:
(iii) The takeoff distance determinedunder § 23.59(a) is equal to the availablerunway length; or
(iv) At the option of the applicant,the takeoff distance determined under§ 23.59(a) is equal to the available run-way length plus the length of anyclearway and the takeoff run deter-mined under § 23.59(b) is equal to theavailable runway length.
(5) For commuter category airplanes,the maximum landing weight for eachairport altitude within the range se-lected by the applicant at which—
(i) The airplane complies with theclimb requirements of § 23.63(d)(2) forambient temperatures within the rangeselected by the applicant; and
(ii) The landing distance determinedunder § 23.75 for standard temperaturesis equal to the available runwaylength.
(6) The maximum zero wing fuelweight, where relevant, as establishedin accordance with § 23.343.
(d) Center of gravity. The establishedcenter of gravity limits.
(e) Maneuvers. The following author-ized maneuvers, appropriate airspeedlimitations, and unauthorized maneu-vers, as prescribed in this section.
(1) Normal category airplanes. No acro-batic maneuvers, including spins, areauthorized.
(2) Utility category airplanes. A list ofauthorized maneuvers demonstrated inthe type flight tests, together with rec-ommended entry speeds and any otherassociated limitations. No other ma-neuver is authorized.
(3) Acrobatic category airplanes. A listof approved flight maneuvers dem-onstrated in the type flight tests, to-gether with recommended entry speedsand any other associated limitations.
(4) Acrobatic category airplanes andutility category airplanes approved forspinning. Spin recovery procedure es-tablished to show compliance with§ 23.221(c).
(5) Commuter category airplanes. Ma-neuvers are limited to any maneuverincident to normal flying, stalls, (ex-cept whip stalls) and steep turns inwhich the angle of bank is not morethan 60 degrees.
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Federal Aviation Administration, DOT § 23.1585
(f) Maneuver load factor. The positivelimit load factors in g’s, and, in addi-tion, the negative limit load factor foracrobatic category airplanes.
(g) Minimum flight crew. The numberand functions of the minimum flightcrew determined under § 23.1523.
(h) Kinds of operation. A list of thekinds of operation to which the air-plane is limited or from which it is pro-hibited under § 23.1525, and also a list ofinstalled equipment that affects anyoperating limitation and identificationas to the equipment’s required oper-ational status for the kinds of oper-ation for which approval has beengiven.
(i) Maximum operating altitude. Themaximum altitude established under§ 23.1527.
(j) Maximum passenger seating configu-ration. The maximum passenger seatingconfiguration.
(k) Allowable lateral fuel loading. Themaximum allowable lateral fuel load-ing differential, if less than the max-imum possible.
(l) Baggage and cargo loading. The fol-lowing information for each baggageand cargo compartment or zone—
(1) The maximum allowable load; and(2) The maximum intensity of load-
ing.(m) Systems. Any limitations on the
use of airplane systems and equipment.(n) Ambient temperatures. Where ap-
propriate, maximum and minimum am-bient air temperatures for operation.
(o) Smoking. Any restrictions onsmoking in the airplane.
(p) Types of surface. A statement ofthe types of surface on which oper-ations may be conducted. (See § 23.45(g)and § 23.1587 (a)(4), (c)(2), and (d)(4)).
§ 23.1585 Operating procedures.(a) For all airplanes, information
concerning normal, abnormal (if appli-cable), and emergency procedures andother pertinent information necessaryfor safe operation and the achievement
of the scheduled performance must befurnished, including—
(1) An explanation of significant orunusual flight or ground handling char-acteristics;
(2) The maximum demonstrated val-ues of crosswind for takeoff and land-ing, and procedures and informationpertinent to operations in crosswinds;
(3) A recommended speed for flight inrough air. This speed must be chosen toprotect against the occurrence, as a re-sult of gusts, of structural damage tothe airplane and loss of control (for ex-ample, stalling);
(4) Procedures for restarting any tur-bine engine in flight, including the ef-fects of altitude; and
(5) Procedures, speeds, and configura-tion(s) for making a normal approachand landing, in accordance with §§ 23.73and 23.75, and a transition to thebalked landing condition.
(6) For seaplanes and amphibians,water handling procedures and thedemonstrated wave height.
(b) In addition to paragraph (a) ofthis section, for all single-engine air-planes, the procedures, speeds, and con-figuration(s) for a glide following en-gine failure, in accordance with § 23.71and the subsequent forced landing,must be furnished.
(c) In addition to paragraph (a) ofthis section, for all multiengine air-planes, the following information mustbe furnished:
(1) Procedures, speeds, and configura-tion(s) for making an approach andlanding with one engine inoperative;
(2) Procedures, speeds, and configura-tion(s) for making a balked landingwith one engine inoperative and theconditions under which a balked land-ing can be performed safely, or a warn-ing against attempting a balked land-ing;
(3) The VSSE determined in § 23.149;and
(4) Procedures for restarting any en-gine in flight including the effects ofaltitude.
(d) In addition to paragraphs (a) andeither (b) or (c) of this section, as ap-propriate, for all normal, utility, andacrobatic category airplanes, the fol-lowing information must be furnished:
(1) Procedures, speeds, and configura-tion(s) for making a normal takeoff, in
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accordance with § 23.51 (a) and (b), and§ 23.53 (a) and (b), and the subsequentclimb, in accordance with § 23.65 and§ 23.69(a).
(2) Procedures for abandoning a take-off due to engine failure or other cause.
(e) In addition to paragraphs (a), (c),and (d) of this section, for all normal,utility, and acrobatic category multi-engine airplanes, the information mustinclude the following:
(1) Procedures and speeds for con-tinuing a takeoff following engine fail-ure and the conditions under whichtakeoff can safely be continued, or awarning against attempting to con-tinue the takeoff.
(2) Procedures, speeds, and configura-tions for continuing a climb followingengine failure, after takeoff, in accord-ance with § 23.67, or enroute, in accord-ance with § 23.69(b).
(f) In addition to paragraphs (a) and(c) of this section, for commuter cat-egory airplanes, the information mustinclude the following:
(1) Procedures, speeds, and configura-tion(s) for making a normal takeoff.
(2) Procedures and speeds for car-rying out an accelerate-stop in accord-ance with § 23.55.
(3) Procedures and speeds for con-tinuing a takeoff following engine fail-ure in accordance with § 23.59(a)(1) andfor following the flight path deter-mined under § 23.57 and § 23.61(a).
(g) For multiengine airplanes, infor-mation identifying each operating con-dition in which the fuel system inde-pendence prescribed in § 23.953 is nec-essary for safety must be furnished, to-gether with instructions for placingthe fuel system in a configuration usedto show compliance with that section.
(h) For each airplane showing com-pliance with § 23.1353 (g)(2) or (g)(3), theoperating procedures for disconnectingthe battery from its charging sourcemust be furnished.
(i) Information on the total quantityof usable fuel for each fuel tank, andthe effect on the usable fuel quantity,as a result of a failure of any pump,must be furnished.
(j) Procedures for the safe operationof the airplane’s systems and equip-ment, both in normal use and in the
ance information must be providedover the altitude and temperatureranges required by § 23.45(b).
(a) For all airplanes, the following in-formation must be furnished—
(1) The stalling speeds VSO and VS1
with the landing gear and wing flapsretracted, determined at maximumweight under § 23.49, and the effect onthese stalling speeds of angles of bankup to 60 degrees;
(2) The steady rate and gradient ofclimb with all engines operating, deter-mined under § 23.69(a);
(3) The landing distance, determinedunder § 23.75 for each airport altitudeand standard temperature, and thetype of surface for which it is valid;
(4) The effect on landing distances ofoperation on other than smooth hardsurfaces, when dry, determined under§ 23.45(g); and
(5) The effect on landing distances ofrunway slope and 50 percent of theheadwind component and 150 percent ofthe tailwind component.
(b) In addition to paragraph (a) ofthis section, for all normal, utility, andacrobatic category reciprocating en-gine-powered airplanes of 6,000 poundsor less maximum weight, the steadyangle of climb/descent, determinedunder § 23.77(a), must be furnished.
(c) In addition to paragraphs (a) and(b) of this section, if appropriate, fornormal, utility, and acrobatic categoryairplanes, the following informationmust be furnished—
(1) The takeoff distance, determinedunder § 23.53 and the type of surface forwhich it is valid.
(2) The effect on takeoff distance ofoperation on other than smooth hardsurfaces, when dry, determined under§ 23.45(g);
(3) The effect on takeoff distance ofrunway slope and 50 percent of theheadwind component and 150 percent ofthe tailwind component;
(4) For multiengine reciprocating en-gine-powered airplanes of more than6,000 pounds maximum weight and mul-tiengine turbine powered airplanes, the
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(5) For multiengine airplanes, theenroute rate and gradient of climb/de-scent with one engine inoperative, de-termined under § 23.69(b); and
(6) For single-engine airplanes, theglide performance determined under§ 23.71.
(d) In addition to paragraph (a) ofthis section, for commuter categoryairplanes, the following informationmust be furnished—
(1) The accelerate-stop distance de-termined under § 23.55;
(2) The takeoff distance determinedunder § 23.59(a);
(3) At the option of the applicant, thetakeoff run determined under § 23.59(b);
(4) The effect on accelerate-stop dis-tance, takeoff distance and, if deter-mined, takeoff run, of operation onother than smooth hard surfaces, whendry, determined under § 23.45(g);
(5) The effect on accelerate-stop dis-tance, takeoff distance, and if deter-mined, takeoff run, of runway slopeand 50 percent of the headwind compo-nent and 150 percent of the tailwindcomponent;
(6) The net takeoff flight path deter-mined under § 23.61(b);
(7) The enroute gradient of climb/de-scent with one engine inoperative, de-termined under § 23.69(b);
(8) The effect, on the net takeoffflight path and on the enroute gradientof climb/descent with one engine inop-erative, of 50 percent of the headwindcomponent and 150 percent of the tail-wind component;
(9) Overweight landing performanceinformation (determined by extrapo-lation and computed for the range ofweights between the maximum landingand maximum takeoff weights) as fol-lows—
(i) The maximum weight for each air-port altitude and ambient temperatureat which the airplane complies withthe climb requirements of § 23.63(d)(2);and
(ii) The landing distance determinedunder § 23.75 for each airport altitudeand standard temperature.
(10) The relationship between IASand CAS determined in accordancewith § 23.1323 (b) and (c).
(11) The altimeter system calibrationrequired by § 23.1325(e).
[Doc. No. 27807, 61 FR 5194, Feb. 9, 1996]
§ 23.1589 Loading information.The following loading information
must be furnished:(a) The weight and location of each
item of equipment that can be easilyremoved, relocated, or replaced andthat is installed when the airplane wasweighed under the requirement of§ 23.25.
(b) Appropriate loading instructionsfor each possible loading condition be-tween the maximum and minimumweights established under § 23.25, to fa-cilitate the center of gravity remain-ing within the limits established under§ 23.23.
APPENDIX A TO PART 23—SIMPLIFIEDDESIGN 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 foran airplane having a maximum weight of6,000 pounds or less and the following con-figuration:
(1) A single engine excluding turbine pow-erplants;
(2) A main wing located closer to the air-plane’s center of gravity than to the aft, fu-selage-mounted, empennage;
(3) A main wing that contains a quarter-chord sweep angle of not more than 15 de-grees fore or aft;
(4) A main wing that is equipped with trail-ing-edge controls (ailerons or flaps, or both);
(5) A main wing aspect ratio not greaterthan 7;
(6) A horizontal tail aspect ratio not great-er than 4;
(7) A horizontal tail volume coefficient notless than 0.34;
(8) A vertical tail aspect ratio not greaterthan 2;
(9) A vertical tail platform area not great-er than 10 percent of the wing platform area;and
(10) Symmetrical airfoils must be used inboth the horizontal and vertical tail designs.
(b) Appendix A criteria may not be used onany airplane configuration that contains anyof the following design features:
(1) Canard, tandem-wing, close-coupled, ortailless arrangements of the lifting surfaces;
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(2) Biplane or multiplane wing arrange-ments;
(3) T-tail, V-tail, or cruciform-tail (+) ar-rangements;
(4) Highly-swept wing platform (more than15-degrees of sweep at the quarter-chord),delta planforms, or slatted lifting surfaces;or
(5) Winglets or other wing tip devices, oroutboard fins.A23.3 Special symbols.n1 =Airplane Positive Maneuvering Limit
tor at VC.n4 =Airplane Negative Gust Limit Load Fac-
tor at VC.nflap =Airplane Positive Limit Load Factor
With Flaps Fully Extended at VF.
A23.5 Certification in more than one category.The criteria in this appendix may be used
for certification in the normal, utility, andacrobatic categories, or in any combinationof these categories. If certification in morethan one category is desired, the design cat-egory weights must be selected to make theterm ‘‘ n1W ’’ constant for all categories orgreater for one desired category than forothers. The wings and control surfaces (in-cluding wing flaps and tabs) need only be in-vestigated for the maximum value of ‘‘ n1W’’, or for the category corresponding to themaximum design weight, where ‘‘ n1W ’’ isconstant. If the acrobatic category is se-lected, a special unsymmetrical flight loadinvestigation in accordance with paragraphsA23.9(c)(2) and A23.11(c)(2) of this appendixmust be completed. The wing, wing carry-through, and the horizontal tail structuresmust be checked for this condition. Thebasic fuselage structure need only be inves-tigated for the highest load factor designcategory selected. The local supportingstructure for dead weight items need only bedesigned for the highest load factor imposedwhen the particular items are installed inthe airplane. The engine mount, however,must be designed for a higher side load fac-tor, if certification in the acrobatic category
is desired, than that required for certifi-cation in the normal and utility categories.When designing for landing loads, the land-ing gear and the airplane as a whole needonly be investigated for the category cor-responding to the maximum design weight.These simplifications apply to single-engineaircraft of conventional types for which ex-perience is available, and the Administratormay require additional investigations foraircraft with unusual design features.A23.7 Flight loads.
(a) Each flight load may be consideredindependent of altitude and, except for thelocal supporting structure for dead weightitems, only the maximum design weight con-ditions must be investigated.
(b) Table 1 and figures 3 and 4 of this ap-pendix must be used to determine values ofn1, n2, n3, and n4, corresponding to the max-imum design weights in the desired cat-egories.
(c) Figures 1 and 2 of this appendix must beused to determine values of n3 and n4 cor-responding to the minimum flying weights inthe desired categories, and, if these load fac-tors are greater than the load factors at thedesign weight, the supporting structure fordead weight items must be substantiated forthe resulting higher load factors.
(d) Each specified wing and tail loading isindependent of the center of gravity range.The applicant, however, must select a c.g.range, and the basic fuselage structure mustbe investigated for the most adverse deadweight loading conditions for the c.g. rangeselected.
(e) The following loads and loading condi-tions are the minimums for which strengthmust be provided in the structure:
(1) Airplane equilibrium. The aerodynamicwing loads may be considered to act normalto the relative wind, and to have a mag-nitude of 1.05 times the airplane normalloads (as determined from paragraphs A23.9(b) and (c) of this appendix) for the positiveflight conditions and a magnitude equal tothe airplane normal loads for the negativeconditions. Each chordwise and normal com-ponent of this wing load must be considered.
(2) Minimum design airspeeds. The minimumdesign airspeeds may be chosen by the appli-cant except that they may not be less thanthe minimum speeds found by using figure 3of this appendix. In addition, VCmin need notexceed values of 0.9 VH actually obtained atsea level for the lowest design weight cat-egory for which certification is desired. Incomputing these minimum design airspeeds,n1 may not be less than 3.8.
(3) Flight load factor. The limit flight loadfactors specified in Table 1 of this appendixrepresent the ratio of the aerodynamic forcecomponent (acting normal to the assumedlongitudinal axis of the airplane) to theweight of the airplane. A positive flight load
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factor is an aerodynamic force acting up-ward, with respect to the airplane.A23.9 Flight conditions.
(a) General. Each design condition in para-graphs (b) and (c) of this section must beused to assure sufficient strength for eachcondition of speed and load factor on orwithin the boundary of a V¥n diagram forthe airplane similar to the diagram in figure4 of this appendix. This diagram must also beused to determine the airplane structural op-erating limitations as specified in§§ 23.1501(c) through 23.1513 and § 23.1519.
(b) Symmetrical flight conditions. The air-plane must be designed for symmetricalflight conditions as follows:
(1) The airplane must be designed for atleast the four basic flight conditions, ‘‘A’’,‘‘D’’, ‘‘E’’, and ‘‘G’’ as noted on the flight en-velope of figure 4 of this appendix. In addi-tion, the following requirements apply:
(i) The design limit flight load factors cor-responding to conditions ‘‘D’’ and ‘‘E’’ of fig-ure 4 must be at least as great as those speci-fied in Table 1 and figure 4 of this appendix,and the design speed for these conditionsmust be at least equal to the value of VD
found from figure 3 of this appendix.(ii) For conditions ‘‘A’’ and ‘‘G’’ of figure 4,
the load factors must correspond to thosespecified in Table 1 of this appendix, and thedesign speeds must be computed using theseload factors with the maximum static liftcoefficient CNA determined by the applicant.However, in the absence of more precisecomputations, these latter conditions maybe based on a value of CNA =±1.35 and the de-sign speed for condition ‘‘A’’ may be lessthan VAmin.
(iii) Conditions ‘‘C’’ and ‘‘F’’ of figure 4need only be investigated when n3 W/S or n4
W/S are greater than n1 W/S or n2 W/S of thisappendix, respectively.
(2) If flaps or other high lift devices in-tended for use at the relatively low airspeedof approach, landing, and takeoff, are in-stalled, the airplane must be designed for thetwo flight conditions corresponding to thevalues of limit flap-down factors specified inTable 1 of this appendix with the flaps fullyextended at not less than the design flapspeed VFmin from figure 3 of this appendix.
(c) Unsymmetrical flight conditions. Each af-fected structure must be designed for unsym-metrical loadings as follows:
(1) The aft fuselage-to-wing attachmentmust be designed for the critical verticalsurface load determined in accordance withparagraph SA23.11(c)(1) and (2) of this appen-dix.
(2) The wing and wing carry-through struc-tures must be designed for 100 percent of con-dition ‘‘A’’ loading on one side of the planeof symmetry and 70 percent on the oppositeside for certification in the normal and util-ity categories, or 60 percent on the opposite
side for certification in the acrobatic cat-egory.
(3) The wing and wing carry-through struc-tures must be designed for the loads result-ing from a combination of 75 percent of thepositive maneuvering wing loading on bothsides of the plane of symmetry and the max-imum wing torsion resulting from ailerondisplacement. The effect of aileron displace-ment on wing torsion at VC or VA using thebasic airfoil moment coefficient modifiedover the aileron portion of the span, must becomputed as follows:
(ii) Cm=Cm ¥0.01δµ(down aileron side) wingbasic airfoil, where δµ is the up aileron de-flection and δ d is the down aileron deflec-tion.
(4) ∆ critical, which is the sum of δµ+δ dmust be computed as follows:
(i) Compute ∆α and ∆B from the formulas:
∆ ∆
∆ ∆
aA
C
p
bA
D
p
V
V
V
V
= ×
= ×
and
0 5.
Where ∆P = the maximum total deflection(sum of both aileron deflections) at VA withVA, VC, and VD described in subparagraph (2)of § 23.7(e) of this appendix.
(ii) Compute K from the formula:
KC V
C V
m b D
m a C
=−( )−( )
0 01
0 01
2
2
.
.
δ
δwhere δα is the down aileron deflection cor-responding to ∆α, and δ b is the down ailerondeflection corresponding to ∆ b as computedin step (i).
(iii) If K is less than 1.0, ∆α is ∆ critical andmust be used to determine δU and δd. In thiscase, VC is the critical speed which must beused in computing the wing torsion loadsover the aileron span.
(iv) If K is equal to or greater than 1.0, ∆Bis ∆ critical and must be used to determineδU and δD. In this case, Vd is the criticalspeed which must be used in computing thewing torsion loads over the aileron span.
(d) Supplementary conditions; rear lift truss;engine torque; side load on engine mount. Eachof the following supplementary conditionsmust be investigated:
(1) In designing the rear lift truss, the spe-cial condition specified in § 23.369 may be in-vestigated instead of condition ‘‘G’’ of figure4 of this appendix. If this is done, and if cer-tification in more than one category is de-sired, the value of W/S used in the formula
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appearing in § 23.369 must be that for the cat-egory corresponding to the maximum grossweight.
(2) Each engine mount and its supportingstructures must be designed for the max-imum limit torque corresponding to METOpower and propeller speed acting simulta-neously with the limit loads resulting fromthe maximum positive maneuvering flightload factor n1. The limit torque must be ob-tained by multiplying the mean torque by afactor of 1.33 for engines with five or morecylinders. For 4, 3, and 2 cylinder engines,the factor must be 2, 3, and 4, respectively.
(3) Each engine mount and its supportingstructure must be designed for the loads re-sulting from a lateral limit load factor of notless than 1.47 for the normal and utility cat-egories, or 2.0 for the acrobatic category.
A23.11 Control surface loads.
(a) General. Each control surface load mustbe determined using the criteria of para-graph (b) of this section and must lie withinthe simplified loadings of paragraph (c) ofthis section.
(b) Limit pilot forces. In each control surfaceloading condition described in paragraphs (c)through (e) of this section, the airloads onthe movable surfaces and the correspondingdeflections need not exceed those whichcould be obtained in flight by employing themaximum limit pilot forces specified in thetable in § 23.397(b). If the surface loads arelimited by these maximum limit pilot forces,the tabs must either be considered to be de-flected to their maximum travel in the direc-tion which would assist the pilot or the de-flection must correspond to the maximumdegree of ‘‘out of trim’’ expected at the speedfor the condition under consideration. Thetab load, however, need not exceed the valuespecified in Table 2 of this appendix.
(c) Surface loading conditions. Each surfaceloading condition must be investigated asfollows:
(1) Simplified limit surface loadings for thehorizontal tail, vertical tail, aileron, wingflaps, and trim tabs are specified in figures 5and 6 of this appendix.
(i) The distribution of load along the spanof the surface, irrespective of the chordwiseload distribution, must be assumed propor-tional to the total chord, except on horn bal-ance surfaces.
(ii) The load on the stabilizer and elevator,and the load on fin and rudder, must be dis-tributed chordwise as shown in figure 7 ofthis appendix.
(iii) In order to ensure adequate torsionalstrength and to account for maneuvers andgusts, the most severe loads must be consid-ered in association with every center of pres-sure position between the leading edge andthe half chord of the mean chord of the sur-face (stabilizer and elevator, or fin and rud-der).
(iv) To ensure adequate strength underhigh leading edge loads, the most severe sta-bilizer and fin loads must be further consid-ered as being increased by 50 percent overthe leading 10 percent of the chord with theloads aft of this appropriately decreased toretain the same total load.
(v) The most severe elevator and rudderloads should be further considered as beingdistributed parabolically from three timesthe mean loading of the surface (stabilizerand elevator, or fin and rudder) at the lead-ing edge of the elevator and rudder, respec-tively, to zero at the trailing edge accordingto the equation:
P x wc x
cf
( ) ( )( )
=−
32
2
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Where—P(x)=local pressure at the chordwise stations
x,c=chord length of the tail surface,cf=chord length of the elevator and rudder
respectively, andw̆≤=average surface loading as specified in
Figure A5.(vi) The chordwise loading distribution for
ailerons, wing flaps, and trim tabs are speci-fied in Table 2 of this appendix.
(2) If certification in the acrobatic cat-egory is desired, the horizontal tail must beinvestigated for an unsymmetrical load of100 percent w on one side of the airplane cen-terline and 50 percent on the other side ofthe airplane centerline.
(d) Outboard fins. Outboard fins must meetthe requirements of § 23.445.
(e) Special devices. Special devices mustmeet the requirements of § 23.459.A23.13 Control system loads.
(a) Primary flight controls and systems. Eachprimary flight control and system must bedesigned as follows:
(1) The flight control system and its sup-porting structure must be designed for loadscorresponding to 125 percent of the computedhinge moments of the movable control sur-face in the conditions prescribed in A23.11 ofthis appendix. In addition—
(i) The system limit loads need not exceedthose that could be produced by the pilot andautomatic devices operating the controls;and
(ii) The design must provide a rugged sys-tem for service use, including jamming,ground gusts, taxiing downwind, control in-ertia, and friction.
(2) Acceptable maximum and minimumlimit pilot forces for elevator, aileron, andrudder controls are shown in the table in§ 23.397(b). These pilots loads must be as-sumed to act at the appropriate control gripsor pads as they would under flight condi-tions, and to be reacted at the attachmentsof the control system to the control surfacehorn.
(b) Dual controls. If there are dual controls,the systems must be designed for pilots oper-ating in opposition, using individual pilotloads equal to 75 percent of those obtained inaccordance with paragraph (a) of this sec-tion, except that individual pilot loads maynot be less than the minimum limit pilotforces shown in the table in § 23.397(b).
(c) Ground gust conditions. Ground gust con-ditions must meet the requirements of§ 23.415.
(d) Secondary controls and systems. Sec-ondary controls and systems must meet therequirements of § 23.405.
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
at c. g.0 ................... 0 ...................... 0 ...................... 0 ......................... 0.
(4) ................... (1) ................... (1), (3), and (4) .. (3) and (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, withlinear 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 from25 percent deflection to 100 percent deflection unless otherwise shown and the load factor must be used with whatever shockabsorber 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-UP AND SPRING-BACK LOADS
D23.1 Wheel spin-up loads.(a) The following method for determining
wheel spin-up loads for landing conditions isbased on NACA T.N. 863. However, the dragcomponent used for design may not be lessthan the drag load prescribed in § 23.479(b).
FHmax =1/re √ 2Iw(VH—Vc)nFVmax/tS
where—
FHmax=maximum rearward horizontal forceacting on the wheel (in pounds);
re=effective rolling radius of wheel under im-pact based on recommended operatingtire pressure (which may be assumed tobe equal to the rolling radius under astatic load of njWe) in feet;
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Iw=rotational mass moment of inertia ofrolling assembly (in slug feet);
VH=linear velocity of airplane parallel toground at instant of contact (assumed tobe 1.2 VS0,in feet per second);
Vc=peripheral speed of tire, if prerotation isused (in feet per second) (there must be apositive means of pre-rotation beforepre-rotation may be considered);
n=equals effective coefficient of friction (0.80may be used);
FVmax=maximum vertical force on wheel(pounds)= njWe, where We and nj are de-fined in § 23.725;
ts=time interval between ground contact andattainment of maximum vertical forceon wheel (seconds). (However, if thevalue of FVmax, from the above equationexceeds 0.8 FVmax, the latter value mustbe used for FHmax.)
(b) The equation assumes a linear vari-ation of load factor with time until the peakload is reached and under this assumption,the equation determines the drag force atthe time that the wheel peripheral velocityat radius re equals the airplane velocity.Most shock absorbers do not exactly follow alinear variation of load factor with time.Therefore, rational or conservative allow-ances must be made to compensate for thesevariations. On most landing gears, the timefor wheel spin-up will be less than the timerequired to develop maximum vertical loadfactor for the specified rate of descent andforward velocity. For exceptionally largewheels, a wheel peripheral velocity equal tothe ground speed may not have been attainedat the time of maximum vertical gear load.However, as stated above, the drag spin-upload need not exceed 0.8 of the maximumvertical loads.
(c) Dynamic spring-back of the landinggear and adjacent structure at the instantjust after the wheels come up to speed mayresult in dynamic forward acting loads ofconsiderable magnitude. This effect must bedetermined, in the level landing condition,by assuming that the wheel spin-up loadscalculated by the methods of this appendixare reversed. Dynamic spring-back is likelyto become critical for landing gear unitshaving wheels of large mass or high landingspeeds.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, asamended by Amdt. 23–45, 58 FR 42167, Aug. 6,1993]
APPENDIX E TO PART 23 [RESERVED]
APPENDIX F TO PART 23—TESTPROCEDURE
Acceptable test procedure for self-extin-guishing materials for showing compliancewith §§ 23.853, 23.855 and 23.1359.
(a) Conditioning. Specimens must be condi-tioned to 70 degrees F, plus or minus 5 de-grees, and at 50 percent plus or minus 5 per-cent relative humidity until moisture equi-librium is reached or for 24 hours. Only onespecimen at a time may be removed from theconditioning environment immediately be-fore subjecting it to the flame.
(b) Specimen configuration. Except as pro-vided for materials used in electrical wireand cable insulation and in small parts, ma-terials must be tested either as a section cutfrom a fabricated part as installed in the air-plane or as a specimen simulating a cut sec-tion, such as: a specimen cut from a flatsheet of the material or a model of the fab-ricated part. The specimen may be cut fromany location in a fabricated part; however,fabricated units, such as sandwich panels,may not be separated for a test. The speci-men thickness must be no thicker than theminimum thickness to be qualified for use inthe airplane, except that: (1) Thick foamparts, such as seat cushions, must be testedin 1⁄2 inch thickness; (2) when showing com-pliance with § 23.853(d)(3)(v) for materialsused in small parts that must be tested, thematerials must be tested in no more than 1⁄8inch thickness; (3) when showing compliancewith § 23.1359(c) for materials used in elec-trical wire and cable insulation, the wire andcable specimens must be the same size asused in the airplane. In the case of fabrics,both the warp and fill direction of the weavemust be tested to determine the most crit-ical flammability conditions. When per-forming the tests prescribed in paragraphs(d) and (e) of this appendix, the specimenmust be mounted in a metal frame so that (1)in the vertical tests of paragraph (d) of thisappendix, the two long edges and the upperedge are held securely; (2) in the horizontaltest of paragraph (e) of this appendix, thetwo long edges and the edge away from theflame are held securely; (3) the exposed areaof the specimen is at least 2 inches wide and12 inches long, unless the actual size used inthe airplane is smaller; and (4) the edge towhich the burner flame is applied must notconsist of the finished or protected edge ofthe specimen but must be representative ofthe actual cross section of the material orpart installed in the airplane. When per-forming the test prescribed in paragraph (f)of this appendix, the specimen must bemounted in metal frame so that all fouredges are held securely and the exposed areaof the specimen is at least 8 inches by 8inches.
(c) Apparatus. Except as provided in para-graph (g) of this appendix, tests must be con-ducted in a draft-free cabinet in accordancewith Federal Test Method Standard 191Method 5903 (revised Method 5902) which isavailable from the General Services Admin-istration, Business Service Center, Region 3,Seventh and D Streets SW., Washington,
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D.C. 20407, or with some other approvedequivalent method. Specimens which are toolarge for the cabinet must be tested in simi-lar draft-free conditions.
(d) Vertical test. A minimum of three speci-mens must be tested and the results aver-aged. For fabrics, the direction of weave cor-responding to the most critical flammabilityconditions must be parallel to the longest di-mension. Each specimen must be supportedvertically. The specimen must be exposed toa Bunsen or Tirrill burner with a nominal 3⁄8-inch I.D. tube adjusted to give a flame of 11⁄2inches in height. The minimum flame tem-perature measured by a calibrated thermo-couple pryometer in the center of the flamemust be 1550° F. The lower edge of the speci-men must be three-fourths inch above thetop edge of the burner. The flame must beapplied to the center line of the lower edge ofthe specimen. For materials covered by§§ 23.853(d)(3)(i) and 23.853(f), the flame mustbe applied for 60 seconds and then removed.For materials covered by § 23.853(d)(3)(ii), theflame must be applied for 12 seconds andthen removed. Flame time, burn length, andflaming time of drippings, if any, must be re-corded. The burn length determined in ac-cordance with paragraph (h) of this appendixmust be measured to the nearest one-tenthinch.
(e) Horizontal test. A minimum of threespecimens must be tested and the resultsaveraged. Each specimen must be supportedhorizontally. The exposed surface when in-stalled in the airplane must be face down forthe test. The specimen must be exposed to aBunsen burner or Tirrill burner with a nomi-nal 3⁄8-inch I.D. tube adjusted to give a flameof 11⁄2 inches in height. The minimum flametemperature measured by a calibrated ther-mocouple pyrometer in the center of theflame must be 1550° F. The specimen must bepositioned so that the edge being tested isthree-fourths of an inch above the top of, andon the center line of, the burner. The flamemust be applied for 15 seconds and then re-moved. A minimum of 10 inches of the speci-men must be used for timing purposes, ap-proximately 11⁄2 inches must burn before theburning front reaches the timing zone, andthe average burn rate must be recorded.
(f) Forty-five degree test. A minimum ofthree specimens must be tested and the re-sults averaged. The specimens must be sup-ported at an angle of 45 degrees to a hori-zontal surface. The exposed surface when in-stalled in the aircraft must be face down forthe test. The specimens must be exposed toa Bunsen or Tirrill burner with a nominal 3⁄8inch I.D. tube adjusted to give a flame of 11⁄2inches in height. The minimum flame tem-perature measured by a calibrated thermo-couple pyrometer in the center of the flamemust be 1550°F. Suitable precautions must betaken to avoid drafts. The flame must be ap-plied for 30 seconds with one-third con-
tacting the material at the center of thespecimen and then removed. Flame time,glow time, and whether the flame penetrates(passes through) the specimen must be re-corded.
(g) Sixty-degree test. A minimum of threespecimens of each wire specification (makeand size) must be tested. The specimen ofwire or cable (including insulation) must beplaced at an angle of 60 degrees with the hor-izontal in the cabinet specified in paragraph(c) of this appendix, with the cabinet dooropen during the test or placed within achamber approximately 2 feet high × 1 foot ×1 foot, open at the top and at one verticalside (front), that allows sufficient flow of airfor complete combustion but is free fromdrafts. The specimen must be parallel to andapproximately 6 inches from the front of thechamber. The lower end of the specimenmust be held rigidly clamped. The upper endof the specimen must pass over a pulley orrod and must have an appropriate weight at-tached to it so that the specimen is heldtautly throughout the flammability test.The test specimen span between lower clampand upper pulley or rod must be 24 inchesand must be marked 8 inches from the lowerend to indicate the central point for flameapplication. A flame from a Bunsen or Tirrillburner must be applied for 30 seconds at thetest mark. The burner must be mounted un-derneath the test mark on the specimen, per-pendicular to the specimen and at an angleof 30 degrees to the vertical plane of thespecimen. The burner must have a nominalbore of three-eighths inch, and must be ad-justed to provide a three-inch-high flamewith an inner cone approximately one-thirdof the flame height. The minimum tempera-ture of the hottest portion of the flame, asmeasured with a calibrated thermocouplepyrometer, may not be less than 1,750 °F. Theburner must be positioned so that the hot-test portion of the flame is applied to thetest mark on the wire. Flame time, burnlength, and flaming time drippings, if any,must be recorded. The burn length deter-mined in accordance with paragraph (h) ofthis appendix must be measured to the near-est one-tenth inch. Breaking of the wirespecimen is not considered a failure.
(h) Burn length. Burn length is the distancefrom the original edge to the farthest evi-dence of damage to the test specimen due toflame impingement, including areas of par-tial or complete consumption, charring, orembrittlement, but not including areas soot-ed, stained, warped, or discolored, nor areaswhere material has shrunk or melted awayfrom the heat source.
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14 CFR Ch. I (1–1–01 Edition)Pt. 23, App. G
APPENDIX G TO PART 23—INSTRUCTIONSFOR CONTINUED AIRWORTHINESS
G23.1 General. (a) This appendix specifiesrequirements for the preparation of Instruc-tions for Continued Airworthiness as re-quired by § 23.1529.
(b) The Instructions for Continued Air-worthiness for each airplane must includethe Instructions for Continued Airworthinessfor each engine and propeller (hereinafterdesignated ‘products’), for each appliance re-quired by this chapter, and any required in-formation relating to the interface of thoseappliances and products with the airplane. IfInstructions for Continued Airworthiness arenot supplied by the manufacturer of an ap-pliance or product installed in the airplane,the Instructions for Continued Airworthinessfor the airplane must include the informa-tion essential to the continued airworthinessof the airplane.
(c) The applicant must submit to the FAAa program to show how changes to the In-structions for Continued Airworthiness madeby the applicant or by the manufacturers ofproducts and appliances installed in the air-plane will be distributed.
G23.2 Format. (a) The Instructions forContinued Airworthiness must be in theform of a manual or manuals as appropriatefor the quantity of data to be provided.
(b) The format of the manual or manualsmust provide for a practical arrangement.
G23.3 Content. The contents of the manualor manuals must be prepared in the Englishlanguage. The Instructions for ContinuedAirworthiness must contain the followingmanuals or sections, as appropriate, and in-formation:
(a) Airplane maintenance manual or section.(1) Introduction information that includes anexplanation of the airplane’s features anddata to the extent necessary for mainte-nance or preventive maintenance.
(2) A description of the airplane and itssystems and installations including its en-gines, propellers, and appliances.
(3) Basic control and operation informationdescribing how the airplane components andsystems are controlled and how they oper-ate, including any special procedures andlimitations that apply.
(4) Servicing information that covers de-tails regarding servicing points, capacities oftanks, reservoirs, types of fluids to be used,pressures applicable to the various systems,location of access panels for inspection andservicing, locations of lubrication points, lu-bricants to be used, equipment required forservicing, tow instructions and limitations,mooring, jacking, and leveling information.
(b) Maintenance instructions. (1) Schedulinginformation for each part of the airplane andits engines, auxiliary power units, propellers,accessories, instruments, and equipment
that provides the recommended periods atwhich they should be cleaned, inspected, ad-justed, tested, and lubricated, and the degreeof inspection, the applicable wear tolerances,and work recommended at these periods.However, the applicant may refer to an ac-cessory, instrument, or equipment manufac-turer as the source of this information if theapplicant shows that the item has an excep-tionally high degree of complexity requiringspecialized maintenance techniques, testequipment, or expertise. The recommendedoverhaul periods and necessary cross ref-erence to the Airworthiness Limitations sec-tion of the manual must also be included. Inaddition, the applicant must include an in-spection program that includes the fre-quency and extent of the inspections nec-essary to provide for the continued air-worthiness of the airplane.
(2) Troubleshooting information describingprobable malfunctions, how to recognizethose malfunctions, and the remedial actionfor those malfunctions.
(3) Information describing the order andmethod of removing and replacing productsand parts with any necessary precautions tobe taken.
(4) Other general procedural instructionsincluding procedures for system testing dur-ing ground running, symmetry checks,weighing and determining the center of grav-ity, lifting and shoring, and storage limita-tions.
(c) Diagrams of structural access platesand information needed to gain access for in-spections when access plates are not pro-vided.
(d) Details for the application of special in-spection techniques including radiographicand ultrasonic testing where such processesare specified.
(e) Information needed to apply protectivetreatments to the structure after inspection.
(f) All data relative to structural fastenerssuch as identification, discard recommenda-tions, and torque values.
(g) A list of special tools needed.(h) In addition, for commuter category air-
planes, the following information must befurnished:
(1) Electrical loads applicable to the var-ious systems;
(2) Methods of balancing control surfaces;(3) Identification of primary and secondary
structures; and(4) Special repair methods applicable to
the airplane.G23.4 Airworthiness Limitations section. The
Instructions for Continued Airworthinessmust contain a section titled AirworthinessLimitations that is segregated and clearly
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Federal Aviation Administration, DOT Pt. 23, App. H
distinguishable from the rest of the docu-ment. This section must set forth each man-datory replacement time, structural inspec-tion interval, and related structural inspec-tion procedure required for type certifi-cation. If the Instructions for Continued Air-worthiness consist of multiple documents,the section required by this paragraph mustbe included in the principal manual. Thissection must contain a legible statement ina prominent location that reads: ‘‘The Air-worthiness Limitations section is FAA ap-proved and specifies maintenance requiredunder §§ 43.16 and 91.403 of the Federal Avia-tion Regulations unless an alternative pro-gram has been FAA approved.’’
APPENDIX H TO PART 23—INSTALLATIONOF AN AUTOMATIC POWER RESERVE(APR) SYSTEM
H23.1, General.(a) This appendix specifies requirements
for installation of an APR engine power con-trol system that automatically advancespower or thrust on the operating engine(s) inthe event any engine fails during takeoff.
(b) With the APR system and associatedsystems functioning normally, all applicable
requirements (except as provided in this ap-pendix) must be met without requiring anyaction by the crew to increase power orthrust.
H23.2, Definitions.(a) Automatic power reserve system means
the entire automatic system used only dur-ing takeoff, including all devices both me-chanical and electrical that sense enginefailure, transmit signals, actuate fuel con-trols or power levers on operating engines,including power sources, to achieve thescheduled power increase and furnish cockpitinformation on system operation.
(b) Selected takeoff power, notwithstandingthe definition of ‘‘Takeoff Power’’ in part 1of the Federal Aviation Regulations, meansthe power obtained from each initial powersetting approved for takeoff.
(c) Critical Time Interval, as illustrated infigure H1, means that period starting at V1
minus one second and ending at the intersec-tion of the engine and APR failure flightpath line with the minimum performance allengine flight path line. The engine and APRfailure flight path line intersects the one-en-gine-inoperative flight path line at 400 feetabove the takeoff surface. The engine andAPR failure flight path is based on the air-plane’s performance and must have a posi-tive gradient of at least 0.5 percent at 400feet above the takeoff surface.
H23.3, Reliability and performance require-ments.
(a) It must be shown that, during the crit-ical time interval, an APR failure that in-creases or does not affect power on either en-
gine will not create a hazard to the airplane,or it must be shown that such failures areimprobable.
(b) It must be shown that, during the crit-ical time interval, there are no failure modes
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14 CFR Ch. I (1–1–01 Edition)Pt. 23, App. H
of the APR system that would result in afailure that will decrease the power on eitherengine or it must be shown that such failuresare extremely improbable.
(c) It must be shown that, during the crit-ical time interval, there will be no failure ofthe APR system in combination with an en-gine failure or it must be shown that suchfailures are extremely improbable.
(d) All applicable performance require-ments must be met with an engine failureoccurring at the most critical point duringtakeoff with the APR system functioningnormally.
H23.4, Power setting.The selected takeoff power set on each en-
gine at the beginning of the takeoff roll maynot be less than—
(a) The power necessary to attain, at V1, 90percent of the maximum takeoff power ap-proved for the airplane for the existing con-ditions;
(b) That required to permit normal oper-ation of all safety-related systems and equip-ment that are dependent upon engine poweror power lever position; and
(c) That shown to be free of hazardous en-gine response characteristics when power isadvanced from the selected takeoff powerlevel to the maximum approved takeoffpower.
H23.5, Powerplant controls—general.(a) In addition to the requirements of
§ 23.1141, no single failure or malfunction (orprobable combination thereof) of the APR,including associated systems, may cause thefailure of any powerplant function necessaryfor safety.
(b) The APR must be designed to—(1) Provide a means to verify to the flight
crew before takeoff that the APR is in an op-erating condition to perform its intendedfunction;
(2) Automatically advance power on theoperating engines following an engine failureduring takeoff to achieve the maximum at-
tainable takeoff power without exceeding en-gine operating limits;
(3) Prevent deactivation of the APR bymanual adjustment of the power levers fol-lowing an engine failure;
(4) Provide a means for the flight crew todeactivate the automatic function. Thismeans must be designed to prevent inad-vertent deactivation; and
(5) Allow normal manual decrease or in-crease in power up to the maximum takeoffpower approved for the airplane under theexisting conditions through the use of powerlevers, as stated in § 23.1141(c), except as pro-vided under paragraph (c) of H23.5 of this ap-pendix.
(c) For airplanes equipped with limitersthat automatically prevent engine operatinglimits from being exceeded, other meansmay be used to increase the maximum levelof power controlled by the power levers inthe event of an APR failure. The means mustbe located on or forward of the power levers,must be easily identified and operated underall operating conditions by a single action ofany pilot with the hand that is normallyused to actuate the power levers, and mustmeet the requirements of § 23.777 (a), (b), and(c).
H23.6, Powerplant instruments.In addition to the requirements of § 23.1305:(a) A means must be provided to indicate
when the APR is in the armed or ready con-dition.
(b) If the inherent flight characteristics ofthe airplane do not provide warning that anengine has failed, a warning system inde-pendent of the APR must be provided to givethe pilot a clear warning of any engine fail-ure during takeoff.
(c) Following an engine failure at V1 orabove, there must be means for the crew toreadily and quickly verify that the APR hasoperated satisfactorily.
[Doc. 26344, 58 FR 18979, Apr. 9, 1993]
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Federal Aviation Administration, DOT Pt. 23, App. I
APPENDIX I TO PART 23—SEAPLANE LOADS
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