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Airworthiness Standards FAA FAR Part 25

1. Frequency electronic systems under 825 1317 The field Peak Average strength values for the HIRF environments and 15 soo 20 zm equipment HIRF test levels are expressed in root mean square units measured during the peak of 500 kHz 2 MHz 30 30 the modulation cycle 2 MHz 30 MHz 100 100 a HIRF environment is specified in the fol SU MB TOO Mz 1 1 lowing table 100 MHz 200 MHz 30 10 Table Environment I 200 MHz 400 MHz 10 10 Field strength volts meter 400 MHz 1 GHz 700 40 Frequency Peak Average 1 GHz 2 GHz 1 300 160 10 kHz 2 MHz 50 50 2 GHz 4 GHz 3 000 120 2 MHz 30 MHz 100 100 4 GHz 6 GHz 3 000 160 30 MHz 100 MHz 50 50 6 GHz 8 GHz 400 170 100 MHz 400 MHz 100 100 8 GHz 12 GHZ 1 230 230 400 MHz 700 MHz 700 50 12 GHz 16 GHz 730 180 700 MHz 1 GHz 700 100 18 GHz 40 GHz 600 150 1 GHz 2 GHz 2 000 200 In this table the higher field strength applies at X Gis Gis 3 000 200 the frequency band edges 6 GHz 8 GHz 1 000 200 8 GHz 12 GHz 3 000 300 12 GHz 18 GHz 2 000 200 18 GHz 40 GHz 600 200 In this table the higher field strength applies at the frequency band edges 228 ASA Part 25 Airworthiness Standards Transport Category c Equipment HIRF Test Level 1 1 From 10 kilohertz kHz to 400 megahertz MHz use conducted susceptibility tests with continuous wave CW and 1 kHz square wave modulation with 90 percent depth or greater Th
2. CUSHION SEAT CUSHION MOUNTING 172 Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 Figure 3 19 5 SPACES tid aqu gg i VAN Nl nl gg gn gg 1 1 OV 1 EE 111 B Ba 13 a rh Nay yg er INNEN NN 11 11s qot tlg vA gt 1115 e prea tay VV arb b 417g gg 27 4 SPACES 7 54 SPACES Qul 1274 11g X DELE Ha E03 uf Y ges P DRAFT TUBE EXTENSION BOLTS FOR FAA HOSE TEST BURNER pam d X TO DRAFT TUBE D ct 1 SECTION OF CONNECTING i T 1214 FLANGE 4 CONNECTING FLANGE BOLT HOLES MATERIAL 0 050 STAINLESS STEEL NOTE ONE HALF 14 OF TUBE EXTENSION SHOWN SECOND HALF MATES AT SPOTWELD OVERLAPS jc 33 ASA 173 Appendix to Part 25 Figure 4 152 x 305 x 19 mm 6 x 12 x 3 MARINITE BLOCK 19 mm 34 25 12 305 3 mm 174 Federal Aviation Regulations 25 mm 1 DIAMETER HOLE FOR 254 3 mm CALORIMETER MOUNTING k 10 1g gt SAK 8 76 3 mmf 12 305 3 mm K 21 5 16 546 3 mm VIEW SIDE
3. 37 Design and construction accessibility 25 611 aeroelastic stability 25 629 bearing factors 25 623 bird strike damage 25 631 casting factors 25 621 control surfaces hinges 25 657 nicae installation 25 655 proof of strength 25 657 control systems cables 25 689 details 25 685 flaps or slats 25 701 gust locks 25 679 3 joints 625 693 desse teet 63 lift drag devices 25 697 25 699 limit load tests 25 687 operation tests 25 683 stability augmentation 25 672 stops 25 675 takeoff warning 25 703 trim systems 25 677 control systems S25 671 25 703 emergency provisions 25 801 25 819 fabrication methods for 25 605 fasteners 625 607 cieesanes ees fire protection cargo or baggage compartments 825 855 25 858 compartment interiors 25 853 components 25 867 extinguishers 25 857 flammable fluids 25 863 flight controls 25 865 heaters 25 859 lavatories 25 854 fire protection 25 851 25 gt fitting factors 25 625 floats and hulls 25 751 25 755 general standards 25 601 25 631 58 61 landing gear specifications 25 721 25 737 material strength 25 613 materials for 25 603 personnel and cargo accommodations G20 A 1 25 193 5 recie pie pressurization 25 841 25 843
4. rough air 25 1517 Airworthiness standards flight compliance proof of 925 21 13 controllability maneuverability 25 143 25 149 cree 23 27 ground and water handling 25 231 25 239 eene 31 32 load distribution limits 25 23 13 stability 25 171 25 181 stalls 25 201 25 207 trim 25 161 high speed characteristics 25 253 out of trim characteristics 25 255 transport category airplanes applicability special requirements 25 1 25 2 eee vibration and buffeting 25 257 Automatic Takeoff Thrust Control System ATTCS Appendix l ss 216 Auxiliary power unit controls 625 1142 sanae ees 109 B Batteries 525 1953 iine eaae 123 c Carburetor air preheater design 25 1101 iE pte 107 Carburetor air temperature controls 25 1157 110 Center of gravity limits 25 27 14 Climb general 925 117 Climb one engine inoperative 25 121 224 Cockpit voice recorders 25 1457 133 Continued airworthiness instructions 25 1529 Appendix 137 214 Continuous gust design criteria Appendix G coit teet cic Li EE Ert 210 Controllability and maneuverability general 925 143 ai eie 23 ASA Part 25 Index Cooling general 25 1041 test procedures 25 tests 25 1043 Cowling 25 1193 Design airspeeds 25 335
5. function and installation 25 1307 instruments flight and navigation 25 1303 instruments powerplant 25 1305 miscellaneous 25 1307 systems installation 25 1309 ETOPS design requirements Appendix K ETOPS type design 25 3 2 102 125 126 126 124 127 ETOPS Appendix 221 EWIS electrical wiring interconnection system Appendix H sse 215 EWIS electrical wiring interconnection systems 25 1701 1733 142 145 Exhaust driven turbo supercharger 25 1127 109 Exhaust heat exchanger 25 1125 52 108 Exhaust piping 25 1123 e 108 Exhaust system general 25 1 r 108 Extended operations ETOPS Appendix K 221 Extinguishing agent containers 25 1199 114 238 Fire detector system 25 1203 extinguishers 25 851 extinguishing agents 25 1197 extinguishing system materials 25 1207 extinguishing systems 25 1195 113 protection compliance 25 1207 114 zones drainage of 25 1187 ventilation of 25 1187 Fire walls 25 1191 Flammability Exposure Evaluation Time FEET 25 981 Appendices M and 101 229 231 Flammable fluid carrying components 25 1183 fluids 25 1185 Flight controllability maneuverability 25 143 25 149
6. 126 Federal Aviation Regulations Where a Area A includes all directions in the adja cent dihedral angle that pass through the light source and intersect the common boundary plane at more than 10 degrees but less than 20 de grees and b Area B includes all directions in the adja cent dihedral angle that pass through the light source and intersect the common boundary plane at more than 20 degrees 25 1397 Color specifications Each position light color must have the applica ble International Commission on Illumination chromaticity coordinates as follows a Aviation red 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 x is not greater than y 0 170 and y is not less than 0 390 0 170 x c Aviation white is not less than 0 300 and not greater than 0 540 y is not less than x 0 040 or yo 0 010 whichever is the smaller and y is not greater than x 0 020 nor 0 636 0 400 x Where is the y coordinate of the Planck ian radiator for the value of x considered Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 27 36 FR 12972 July 10 1971 825 1399 Riding light a Each riding anchor light required for a sea plane or amphibian must be installed so that it 1 Show white light for at l
7. 0 883 V 3 0 306 E Ji 0 851 100 100 100 100 E 0 692 100 0 614 us 0 252 15 0 658 3 us 0263 125 4 100 100 100 1 70 percent of the dry runway braking coeffi cient of friction used to determine the dry runway accelerate stop distance or 2 The wet runway braking coefficient defined in paragraph c of this section except that a spe cific anti skid system efficiency if determined is appropriate for a grooved or porous friction course wet runway and the maximum tire to ground wet runway braking coefficient of friction is defined as ASA Part 25 Airworthiness Standards Transport Category 825 111 Tire Pressure psi Maximum Braking Coefficient tire to ground _ vy AY vy 50 TE 0 1470 100 1 050 100 2 673 100 2 683 100 0 403 100 0 859 _ vy MY yy 100 MAX 0 1106 100 0 813 100 2 130 100 2 200 100 0 317 100 0 807 MY v 200 0 0498 Too 0 398 100 1 140 100 1 285 100 0 140 100 0 701 wy XY NY vy v 300 Pg 0 0314 100 0 247 100 0 703 100 0 779 100 0 00954 100 0 614 Where of the allowable brake wear range remaining on Tire Pressure maximum airplane operating tire pressure psi 9 maximum tire to ground braking coefficient V airplane true ground speed knots and Linear interpolation may be used for tire pres sures other than those listed
8. 100 Ibs Rudder 300 Ibs 130 Ibs The critical parts of the aileron control system must be designed for a single tangential force with 43 825 399 a limit value equal to 1 25 times the couple force determined from these criteria D wheel diameter inches 3The unsymmetrical forces must be applied at one of the normal handgrip points on the periph ery of the control wheel Docket 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55466 Dec 20 1976 Amdt 25 72 55 FR 29776 July 20 1990 825 399 Dual control system a Each dual control system must be designed for the pilots operating in opposition using individ ual pilot forces not less than 1 0 75 times those obtained under 825 395 or 2 The minimum forces specified 25 397 c b The control system must be designed for pilot forces applied in the same direction using in dividual pilot forces not less than 0 75 times those obtained under 825 395 25 405 Secondary control system Secondary controls such as wheel brake spoiler and tab controls must be designed for the maximum forces that a pilot is likely to apply to those controls The following values may be used Pilot Control Force Limits Secondary Controls Control Limit pilot forces Miscellaneous Crank wheel or lever 3 50 Ibs but not less than 50 Ibs nor more than 150 Ibs R radius Applicable to any angle
9. Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 91 62 FR 40705 July 29 1997 ASA 825 493 825 485 Side load conditions In addition to 25 479 d 2 the following condi tions must be considered a For the side load condition the airplane is assumed to be in the level attitude with only the main wheels contacting the ground in accor dance with figure 5 of Appendix A b Side loads of 0 8 of the vertical reaction on one side acting inward and 0 6 of the vertical re action on the other side acting outward must be combined with one half of the maximum vertical ground reactions obtained in the level landing conditions These loads are assumed to be ap plied at the ground contact point and to be re sisted by the inertia of the airplane The drag loads may be assumed to be zero Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 91 62 FR 40705 July 29 1997 825 487 Rebound landing condition a The landing gear and its supporting struc ture must be investigated for the loads occurring during rebound of the airplane from the landing surface b With the landing gear fully extended and not in contact with the ground a load factor of 20 0 must act on the unsprung weights of the landing gear This load factor must act in the direction of motion of the unsprung weights as they reach their limiting positions in extending with relation to the sprung parts of the landing gea
10. e Except as provided in paragraph f 1 of this section means other than wheel brakes may be used to determine the accelerate stop dis tance if that means 1 Is safe and reliable 2 Is used so that consistent results can be ex pected under normal operating conditions and 3 Is such that exceptional skill is not required to control the airplane f The effects of available reverse thrust 1 Shall not be included as an additional means of deceleration when determining the ac celerate stop distance on a dry runway and 2 May be included as an additional means of deceleration using recommended reverse thrust procedures when determining the accelerate stop distance on a wet runway provided the require ments of paragraph e of this section are met g The landing gear must remain extended throughout the accelerate stop distance h If the accelerate stop distance includes a stopway with surface characteristics substantially different from those on the runway the takeoff data must include operational correction factors for the accelerate stop distance The correction factors must account for the particular surface characteristics of the stopway and the variations in these characteristics with seasonal weather conditions such as temperature rain snow and ice within the established operational limits i A flight test demonstration of the maximum brake kinetic energy accelerate stop distance must be condu
11. procedures 3 The maneuvering speed Va and a state ment that full application of rudder and aileron controls as well as maneuvers that involve angles of attack near the stall should be confined to speeds below this value 4 The flap extended speed Vfe and the perti nent flap positions and engine powers 5 The landing gear operating speed or speeds and a statement explaining the speeds as defined in 25 1515 a 6 The landing gear extended speed Vig if greater than Vi o and a statement that this is the maximum speed at which the airplane can be safely flown with the landing gear extended b Powerplant limitations The following infor mation must be furnished 1 Limitations required by 825 1521 and 825 1522 2 Explanation of the limitations when appro priate 3 Information necessary for marking the instru ments required by 25 1549 through 25 1553 c Weight and loading distribution The weight and center of gravity limitations established under 825 1519 must be furnished in the Airplane Flight Manual All of the following information including the weight distribution limitations established un der 825 1519 must be presented either in the Air plane Flight Manual or in a separate weight and balance control and loading document that is in ASA 825 1585 corporated by reference in the Airplane Flight Manual 1 The condition of the airplane and the items included in the empty weight as de
12. 1 An airspeed indicator If airspeed limitations vary with altitude the indicator must have a maxi mum allowable airspeed indicator showing the variation of Vyo with altitude 2 An altimeter sensitive 3 A rate of climb indicator vertical speed 4 A gyroscopic rate of turn indicator com bined with an integral slip skid indicator turn and bank indicator except that only a slip skid indica tor is required on large airplanes with a third atti tude instrument system usable through flight atti tudes of 360 of pitch and roll and installed in ac cordance with 121 305 k of this title 5 A bank and pitch indicator gyroscopically stabilized 6 A direction indicator gyroscopically stabi lized magnetic or nonmagnetic c The following flight and navigation instru ments are required as prescribed in this para graph 1 A speed warning device is required for tur bine engine powered airplanes and for airplanes with greater than 0 8 Vpe Mpr or 0 8 Vp Mp The speed warning device must give ef fective aural warning differing distinctively from ASA 825 1305 aural warnings used for other purposes to the pilots whenever the speed exceeds Vyo plus 6 knots or Myo 0 01 The upper limit of the pro duction tolerance for the warning device may not exceed the prescribed warning speed 2 A machmeter is required at each pilot sta tion for airplanes with compressibility limitations not othe
13. b Reciprocating engine installations Operat ing limitations relating to the following must be es tablished for reciprocating engine installations 1 Horsepower or torque r p m manifold pressure and time at critical pressure altitude and Sea level pressure altitude for i Maximum continuous power relating to un supercharged operation or to operation in each supercharger mode as applicable and 1 Takeoff power relating to unsupercharged operation or to operation in each supercharger mode as applicable 2 Fuel grade or specification 3 Cylinder head and oil temperatures 4 Any other parameter for which a limitation has been established as part of the engine type certificate except that a limitation need not be es tablished for a parameter that cannot be exceeded during normal operation due to the design of the installation or to another established limitation c Turbine engine installations Operating limi tations relating to the following must be estab lished for turbine engine installations 1 Horsepower torque or thrust r p m gas temperature and time for i Maximum continuous power or thrust relat ing to augmented or unaugmented operation as applicable ii Takeoff power or thrust relating to aug mented or unaugmented operation as applicable 2 Fuel designation or specification 3 Any other parameter for which a limitation has been established as part of the engine type cer
14. 3 Each exchanger must have cooling provi sions wherever it is subject to contact with ex haust gases and 4 No exhaust heat exchanger or muff may have any stagnant areas or liquid traps that would increase the probability of ignition of flammable fluids or vapors that might be present in case of the failure or malfunction of components carrying flammable fluids b If an exhaust heat exchanger is used for heating ventilating air 1 There must be a secondary heat exchanger between the primary exhaust gas heat exchanger and the ventilating air system or 2 Other means must be used to preclude the harmful contamination of the ventilating air Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 825 1127 Exhaust driven turbo superchargers a Each exhaust driven turbo supercharger must be approved or shown to be suitable for the particular application It must be installed and supported to ensure safe operation between nor mal inspections and overhauls In addition there must be provisions for expansion and flexibility between exhaust conduits and the turbine b There must be provisions for lubricating the turbine and for cooling turbine parts where tem peratures are critical c If the normal turbo supercharger control system malfunctions the turbine speed may not exceed its maximum allowable value Except for the waste gate operating components the com ponen
15. 825 203 d The airplane is considered stalled when the behavior of the airplane gives the pilot a clear and distinctive indication of an acceptable nature that the airplane is stalled Acceptable indications of a stall occurring either individually or in combina tion are 1 A nose down pitch that cannot be readily ar rested 2 Buffeting of a magnitude and severity that is a strong and effective deterrent to further speed reduction or 3 The pitch control reaches the aft stop and no further increase in pitch attitude occurs when the control is held full aft for a short time before recovery is initiated Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55466 Dec 20 1976 Amdt 25 42 43 FR 2322 Jan 16 1978 Amdt 25 84 60 FR 30750 June 9 1995 Amdt 25 108 67 FR 70827 Nov 26 2002 825 203 Stall characteristics a It must be possible to produce and to cor rect roll and yaw by unreversed use of the aileron and rudder controls up to the time the airplane is stalled No abnormal nose up pitching may occur The longitudinal control force must be positive up to and throughout the stall In addition it must be possible to promptly prevent stalling and to re cover from a stall by normal use of the controls b For level wing stalls the roll occurring be tween the stall and the completion of the recovery may not exceed approximately 20 degrees c For turning flight s
16. The selected design airspeeds are equivalent airspeeds EAS Estimated values of Vso and Vg must be conservative a Design cruising speed Vc For Vc the fol lowing apply 1 The minimum value of Vc must be suffi ciently greater than Vg to provide for inadvertent speed increases likely to occur as a result of se vere atmospheric turbulence 2 Except as provided in 25 335 d 2 Vc may not be less than Vg 1 32 Uggr with as specified in 25 341 a 5 i However Vc need not exceed the maximum speed in level flight at maximum continuous power for the corre sponding altitude tion and the use of pilot controlled drag devices ASA 3 At altitudes where Vp is limited by Mach number Vc may be limited to a selected Mach number b Design dive speed Vp Vp must be selected so that Vc Mc is not greater than 0 8 Vp Mp so that the minimum speed margin between Vc Mc and is the greater of the following values 1 From an initial condition of stabilized flight at Vc Mc the airplane is upset flown for 20 seconds along a flight path 7 5 below the initial path and then pulled up at a load factor of 1 5 g 0 5 g ac celeration increment The speed increase occur ring in this maneuver may be calculated if reliable or conservative aerodynamic data is used Power as specified in 25 175 b 1 iv is assumed until the pullup is initiated at which time power reduc may be assumed 37 825 337
17. c Cooling tests for each stage of flight must be continued until 1 The component and engine fluid tempera tures stabilize 2 The stage of flight is completed or 3 An operating limitation is reached d For reciprocating engine powered air planes it may be assumed for cooling test pur poses that the takeoff stage of flight is complete when the airplane reaches an altitude of 1 500 feet above the takeoff surface or reaches a point in the takeoff where the transition from the takeoff to the en route configuration is completed and a speed is reached at which compliance with 25 121 c is shown whichever point is at a higher altitude The airplane must be in the follow ing configuration 1 Landing gear retracted 2 Wing flaps in the most favorable position 3 Cowl flaps or other means of controlling the engine cooling supply in the position that pro vides adequate cooling in the hot day condition 4 Critical engine inoperative and its propeller stopped 5 Remaining engines at the maximum contin uous power available for the altitude e For hull seaplanes and amphibians cooling must be shown during taxiing downwind for 10 minutes at five knots above step speed Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 57 49 FR 6848 Feb 23 1984 INDUCTION SYSTEM 825 1091 Air induction a The air induction system for each engine and auxiliary power unit must supply 1
18. c Subject to appreciable variability because of uncertainties in manufacturing processes or in spection methods Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5674 April 8 1970 825 621 Casting factors a General The factors tests and inspections specified in paragraphs b through d of this sec tion must be applied in addition to those neces sary to establish foundry quality control The in spections must meet approved specifications Paragraphs c and d of this section apply to any structural castings except castings that are pres sure tested as parts of hydraulic or other fluid sys tems and do not support structural loads b Bearing stresses and surfaces The casting factors specified in paragraphs c and d of this section 1 Need not exceed 1 25 with respect to bear ing stresses regardless of the method of inspec tion used and 2 Need not be used with respect to the bear ing surfaces of a part whose bearing factor is larger than the applicable casting factor c Critical castings For each casting whose failure would preclude continued safe flight and landing of the airplane or result in serious injury to occupants the following apply ASA 825 623 1 Each critical casting must i Have a casting factor of not less than 1 25 and ii Receive 100 percent inspection by visual radiographic and magnetic particle or penetrant inspection methods or
19. foot at the particular weight under consideration 88u K 53 u _ 2w p density of air slugs ft3 mean geometric chord of the wing feet 9 acceleration due to gravity ft sec2 a 7 slope of the airplane normal force coefficient curve Cya per radian 2 At altitudes where Vc is limited by Mach number 38 Federal Aviation Regulations i Vg may be chosen to provide an optimum margin between low and high speed buffet bound aries and ii Vg need not be greater than Vc e Design flap speeds For the following apply 1 The design flap speed for each flap position established in accordance with 25 697 a must be sufficiently greater than the operating speed recommended for the corresponding stage of flight including balked landings to allow for prob able variations in control of airspeed and for tran sition from one flap position to another 2 If an automatic flap positioning or load limit ing device is used the speeds and corresponding flap positions programmed or allowed by the de vice may be used 3 may not be less than i 1 6 Vs4 with the flaps in takeoff position at maximum takeoff weight ii 1 8 with the flaps in approach position at maximum landing weight and iii 1 8 Vso with the flaps in landing position at maximum landing weight f Design drag device speeds Vpp The se lected design speed for each drag device must b
20. 0 804 is 50 700350 100 Uto MAX 200 90331 300 Higgy 700401 Where Tire Pressure maximum airplane operating tire pressure psi Ax maximum tire to ground braking coefficient V airplane true ground speed knots and Linear interpolation may be used for tire pres sures other than those listed 2 The maximum tire to ground wet runway braking coefficient of friction must be adjusted to take into account the efficiency of the anti skid System on a wet runway Anti skid system opera tion must be demonstrated by flight testing on a smooth wet runway and its efficiency must be de termined Unless a specific anti skid system effi ciency is determined from a quantitative analysis of the flight testing on a smooth wet runway the maximum tire to ground wet runway braking coef ficient of friction determined in paragraph c 1 of this section must be multiplied by the efficiency value associated with the type of anti skid system installed on the airplane Type of anti skid system Efficiency value On Off 0 30 Quasi Modulating 0 50 Full Modulating 0 80 d At the option of the applicant a higher wet runway braking coefficient of friction may be used for runway surfaces that have been grooved or treated with a porous friction course material For grooved and porous friction course runways the wet runway braking coefficient of friction is de fined as either 18
21. 3 The part of the ventilating air passage that surrounds the combustion chamber However no fire extinguishment is required in cabin ventilating air passages b Ventilating air ducts Each ventilating air duct passing through any fire zone must be fire proof In addition 1 Unless isolation is provided by fireproof valves or by equally effective means the ventilat ing air duct downstream of each heater must be fireproof for a distance great enough to ensure that any fire originating in the heater can be con tained in the duct and 2 Each part of any ventilating duct passing through any region having a flammable fluid sys tem must be constructed or isolated from that sys tem so that the malfunctioning of any component of that system cannot introduce flammable fluids or vapors into the ventilating airstream c Combustion air ducts Each combustion air duct must be fireproof for a distance great enough to prevent damage from backfiring or reverse flame propagation In addition 1 No combustion air duct may have a com mon opening with the ventilating airstream unless flames from backfires or reverse burning cannot enter the ventilating airstream under any operat ing condition including reverse flow or malfunc tioning of the heater or its associated compo nents and 2 No combustion air duct may restrict the prompt relief of any backfire that if so restricted could cause heater failure d Heater controls
22. ASA 825 1001 for a 15 minute flight comprised of a takeoff go around and landing at the airport of departure with the airplane configuration speed power and thrust the same as that used in meeting the appli cable takeoff approach and landing climb perfor mance requirements of this part b If a fuel jettisoning system is required it must be capable of jettisoning enough fuel within 15 minutes starting with the weight given in para graph a of this section to enable the airplane to meet the climb requirements of 25 119 and 25 121 d assuming that the fuel is jettisoned un der the conditions except weight found least fa vorable during the flight tests prescribed in para graph c of this section c Fuel jettisoning must be demonstrated be ginning at maximum takeoff weight with flaps and landing gear up and in 1 A power off glide at 1 3 Vsr1 2 A climb at the one engine inoperative best rate of climb speed with the critical engine inop erative and the remaining engines at maximum continuous power and 3 Level flight at 1 3 if the results of the tests in the conditions specified in paragraphs c 1 and 2 of this section show that this condition could be critical d During the flight tests prescribed in para graph c of this section it must be shown that 1 The fuel jettisoning system and its operation are free from fire hazard 2 The fuel discharges clear of any part of th
23. Appendix F to Part 25 FIGURE 5 3 Views Top View 19 in 483 mm 13 1 8 in 333 mm 1in 25 Side View Three Dimensional View 4 Pilot Burner The pilot burner used to ignite the specimen must be a Bernzomatic commer cial propane venturi torch with an axially symmet ric burner tip and a propane supply tube with an orifice diameter of 0 006 inches 0 15 mm The length of the burner tube must be 2 7 8 inches 71 mm The propane flow must be adjusted via gas pressure through an in line regulator to produce a blue inner cone length of 3 4 inch 19 mm A 3 4 inch 19 mm guide such as a thin strip of metal may be soldered to the top of the burner to aid in setting the flame height The overall flame length must be approximately 5 inches long 127 mm Provide a way to move the burner out of the igni ASA tion position so that the flame is horizontal and at least 2 inches 50 mm above the specimen plane See figure 6 195 Appendix to Part 25 Federal Aviation Regulations FIGURE 6 Propane Pilot Burner SJ 5in 127 mm 2 7 8 in 71 mm 5 Thermocouples Install a 24 American Wire Gauge AWG Type K Chromel Alumel thermo couple in the test chamber for temperature moni toring Insert it into the chamber through a small hole drilled through the back of the chamber Place the thermo
24. Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29780 July 20 1990 25 793 Floor surfaces The floor surface of all areas which are likely to become wet in service must have slip resistant properties Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 51 45 FR 7755 Feb 4 1980 825 795 Security considerations a Protection of flightcrew compartment If a flightdeck door is required by operating rules 1 The bulkhead door and any other accessi ble boundary separating the flightcrew compart ment from occupied areas must be designed to resist forcible intrusion by unauthorized persons and be capable of withstanding impacts of 300 joules 221 3 foot pounds 2 The bulkhead door and any other accessi ble boundary separating the flightcrew compart ment from occupied areas must be designed to resist a constant 250 pound 1 113 Newtons ten sile load on accessible handholds including the doorknob or handle 3 The bulkhead door and any other bound ary separating the flightcrew compartment from any occupied areas must be designed to resist penetration by small arms fire and fragmentation devices to a level equivalent to level Illa of the Na tional Institute of Justice NIJ Standard 0101 04 b Airplanes with a maximum certificated pas senger seating capacity of more than 60 persons or a maximum certificated takeoff gross weight of over 100 000 pounds 45 359
25. T ni 31 04 0 5 in 1 787 12 Second stage air distribution plate with 120 no 28 holes 6 00 0 25 in 152 6 mm 9 95 0 01 in 16 00 0 25 in 406 6 Air distribution plate 8 no 4 holes 4 00 0 25 in 102 6 mm 0 049 0 002 in 1 24 0 05 mm thick Thermopile Air flow Figure 1A Rate of Heat Release Apparatus ASA Appendix to Part 25 Federal Aviation Regulations 0 39 0 02 in 10 0 0 5 mm e Air manifold with Baffle 48 no 26 holes NOTE Seal and secure with bolts or clamps Section A A Air manifold with 48 no 26 holes 0 25 0 01 Specimen top lt 6 0 3 Window frame X 80mm 3 i Specimen face 13 25 0 25 in 337 0 6 mm 0 25 om 6 4mm Hinge typ 7 nominal outside diameter 26 in 711 mm 11 in 279mm inimum minimunr gt 1 50 0 03 in 10 75 0 02 38 1 19 1 1 5 38 gi nominal inside diameter 0 750 0 002 in 19 1 0 1 mm orifice E 24 gauge stainless steel thermopile connection End view Air distribution plate with 8 no 4 holes Figure 1B Rate of Heat Release Apparatus 186 ASA Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 Top Unless denoted otherwise all dimensions are in millimeters
26. c The average gradient of the stable slope of the stick force versus speed curve may not be less than 1 pound for each 6 knots d Within the free return speed range specified in paragraph b of this section it is permissible for the airplane without control forces to stabilize on speeds above or below the desired trim speeds if exceptional attention on the part of the pilot is not required to return to and maintain the desired trim speed and altitude Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 7 30 FR 13117 Oct 15 1965 825 175 Demonstration of static longitudinal stability Static longitudinal stability must be shown as follows a Climb The stick force curve must have a stable slope at speeds between 85 and 115 per cent of the speed at which the airplane 1 Is trimmed with i Wing flaps retracted ii Landing gear retracted iii Maximum takeoff weight and iv 75 percent of maximum continuous power for reciprocating engines or the maximum power or thrust selected by the applicant as an operating limitation for use during climb for turbine engines and 2 Is trimmed at the speed for best rate of climb except that the speed need not be less than 1 3 b Cruise Static longitudinal stability must be shown in the cruise condition as follows 1 With the landing gear retracted at high speed the stick force curve must have a stable slope at all speeds within a ran
27. iii At speed Vp Uc is equal to 1 the values ob tained under paragraph b 3 i of this appendix iv At speeds between Vg and Vc and between Vc and Vp Uc is equal to a value obtained by lin ear interpolation 4 When a stability augmentation system is in cluded in the analysis the effect of system nonlin earities on loads at the limit load level must be re alistically or conservatively accounted for c Mission analysis Limit loads must be deter mined in accordance with the following 1 The expected utilization of the airplane must be represented by one or more night profiles in which the load distribution and the variation with time of speed altitude gross weight and center of gravity position are defined These profiles must be divided into mission segments or blocks for analysis and average or effective values of the pertinent parameters defined for each segment 2 For each of the mission segments defined under paragraph c 1 of this appendix values of A and No must be determined by analysis A is defined as the ratio of root mean square incre mental load to root mean square gust velocity and No is the radius of gyration of the load power spectral density function about zero frequency 211 Appendix G to Part 25 The power spectral density of the atmospheric turbulence must be given by the equation set forth in paragraph b 2 of this appendix where t selected time interval y 7 net value of
28. ity are 1 0 vertically and 0 5 laterally The side ground reaction of each wheel must be 0 5 of the vertical reaction 825 497 Tail wheel yawing a A vertical ground reaction equal to the static load on the tail wheel in combination with a side component of equal magnitude is assumed b If there is a swivel the tail wheel is as sumed to be swiveled 90 to the airplane longitu dinal axis with the resultant load passing through the axle c If there is a lock steering device or shimmy damper the tail wheel is also assumed to be in the trailing position with the side load acting at the ground contact point 825 499 Nose wheel yaw and steering a A vertical load factor of 1 0 at the airplane center of gravity and a side component at the nose wheel ground contact equal to 0 8 of the ver tical ground reaction at that point are assumed b With the airplane assumed to be in static equilibrium with the loads resulting from the use of brakes on one side of the main landing gear the nose gear its attaching structure and the fu selage structure forward of the center of gravity must be designed for the following loads 1 A vertical load factor at the center of gravity of 1 0 ASA Part 25 Airworthiness Standards Transport Category 2 A forward acting load at the airplane center of gravity of 0 8 times the vertical load on one main gear 3 Side and vertical loads at the ground con tact point on the nose ge
29. nation of the kind of operation authorized requires consideration of the operating rules under which the airplane will be operated Unless an applicant desires approval for a more limited kind of opera tion It is assumed that each airplane certificated under this Part will operate under IFR conditions Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 3 30 FR 6067 April 29 1965 163 Appendix E to Part 25 APPENDIX E TO PART 25 1 LIMITED WEIGHT CREDIT FOR AIRPLANES EQUIPPED WiTH STANDBY POWER a Each applicant for an increase in the maxi mum certificated takeoff and landing weights of an airplane equipped with a type certificated standby power rocket engine may obtain an in crease as specified in paragraph b if 1 The installation of the rocket engine has been approved and it has been established by flight test that the rocket engine and its controls can be operated safely and reliably at the in crease in maximum weight and 2 The Airplane Flight Manual or the placard markings or manuals required in place thereof set forth in addition to any other operating limita tions the Administrator may require the increased weight approved under this regulation and a pro hibition against the operation of the airplane at the approved increased weight when i The installed standby power rocket engines have been stored or installed in excess of the time limit established by the manufacturer of th
30. sary switches to actuate the indicator or other means to inform the pilot that the gear is secured in the extended or retracted position This means must be designed as follows 1 If switches are used they must be located and coupled to the landing gear mechanical sys tems in a manner that prevents an erroneous indi cation of down and locked if the landing gear is not in a fully extended position or of and locked if the landing gear is not in the fully re tracted position The switches may be located where they are operated by the actual landing gear locking latch or device 2 The flightcrew must be given an aural warn ing that functions continuously or is periodically repeated if a landing is attempted when the land ing gear is not locked down 3 The warning must be given in sufficient time to allow the landing gear to be locked down or a go around to be made 4 There must be a manual shut off means readily available to the flightcrew for the warning required by paragraph 2 of this section such that it could be operated instinctively inadvert ently or by habitual reflexive action 5 The system used to generate the aural warning must be designed to eliminate false or in appropriate alerts 6 Failures of systems used to inhibit the landing gear aural warning that would prevent the warning System from operating must be improbable f Protection of equipment in wheel wells Equipment that i
31. 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 86 61 FR 5220 Feb 9 1996 FLiGHT MANEUVER AND GUST CONDITIONS 25 331 Symmetric maneuvering conditions a Procedure For the analysis of the maneu vering flight conditions specified in paragraphs b and c of this section the following provisions apply 1 Where sudden displacement of a control is specified the assumed rate of control surface dis placement may not be less than the rate that could be applied by the pilot through the control system 2 In determining elevator angles and chord wise load distribution in the maneuvering condi tions of paragraphs b and c of this section the effect of corresponding pitching velocities must be taken into account The in trim and out of trim flight conditions specified in 25 255 must be con sidered b Maneuvering balanced conditions Assum ing the airplane to be in equilibrium with zero pitch ing acceleration the maneuvering conditions A through on the maneuvering envelope in 25 333 b must be investigated c Pitch maneuver conditions The conditions specified in paragraphs c 1 and 2 of this sec 36 Federal Aviation Regulations tion must be investigated The movement of the pitch control surfaces may be adjusted to take into account limitations imposed by the maximum pilot effort specified by 25 397 b control system stops and any indirect effect imposed by limita tions i
32. 2 The minimum speed margin must be enough to provide for atmospheric variations such as horizontal gusts and penetration of jet streams and cold fronts and for instrument errors and airframe production variations These factors may be considered on a probability basis The margin at altitude where Mg is limited by com pressibility effects must not be less than 0 07M unless a lower margin is determined using a ratio nal analysis that includes the effects of any auto matic systems In any case the margin may not be reduced to less than 0 05M c Design maneuvering speed Va For Va the following apply 1 VA may not be less than V n where i n is the limit positive maneuvering load factor at Vc and ii is the stalling speed with flaps retracted 2 Va and Vs must be evaluated at the design weight and altitude under consideration 3 Va need not be more than Vc or the speed at which the positive Cymax curve intersects the positive maneuver load factor line whichever is less d Design speed for maximum gust intensity Vg 1 Vg may not be less than 1 2 d Vs 498w where the 1 g stalling speed based on with the flaps retracted at the particular weight under consideration Vc 7 design cruise speed knots equivalent airspeed Uret the reference gust velocity feet per second equivalent airspeed 25 341 a 5 i w average wing loading pounds per square
33. 38 41 FR 55468 Dec 20 1976 25 1516 Other speed limitations Any other limitation associated with speed must be established Docket No FAA 2000 8511 66 FR 34024 June 26 2001 25 1517 Rough air speed A rough air speed for use as the recom mended turbulence penetration airspeed 25 1585 a 8 must be established which 1 Is not greater than the design airspeed for maximum gust intensity selected for Vg and 2 Is not less than the minimum value of Vg specified in 25 335 d and 3 Is sufficiently less than Vio to ensure that likely speed variation during rough air encounters will not cause the overspeed warning to operate too frequently In the absence of a rational investi gation substantiating the use of other values VRA must be less than Vo 35 knots TAS Docket No 27902 61 FR 5222 Feb 9 1996 825 1519 Weight center of gravity and weight distribution The airplane weight center of gravity and weight distribution limitations determined under 825 23 through 25 27 must be established as operating limitations 825 1521 Powerplant limitations a General The powerplant limitations pre Scribed in this section must be established so that they do not exceed the corresponding limits for which the engines or propellers are type certifi cated and do not exceed the values on which 137 825 1522 compliance with any other requirement of this part is based
34. 4 The assisting means provided for each es cape route leading from a Type C exit must be au tomatically erected within 10 seconds from the time the opening means of the exit is actuated and that provided for the escape route leading from any other exit type must be automatically erected within 10 seconds after actuation of the erection system e If an integral stair is installed in a passenger entry door that is qualified as a passenger emer gency exit the stair must be designed so that un der the following conditions the effectiveness of passenger emergency egress will not be im paired 1 The door integral stair and operating mechanism have been subjected to the inertia forces specified in 25 561 b 3 acting sepa rately relative to the surrounding structure 2 The airplane is in the normal ground attitude and in each of the attitudes corresponding to col lapse of one or more legs of the landing gear Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29782 July 20 1990 Amdt 25 88 61 FR 57958 Nov 8 1996 Amdt 25 88 62 FR 1817 Jan 13 1997 Amdt 25 114 69 FR 24501 May 3 2004 525 811 Emergency exit marking a Each passenger emergency exit its means of access and its means of opening must be con spicuously marked b The identity and location of each passenger emergency exit must be recognizable from a dis tance equal to the width of the cabin c Means mu
35. 60 FR 6623 Feb 2 1995 Amdt 25 83 60 FR 11194 March 1 1995 Amdt 25 94 63 FR 8848 Feb 23 1998 Docket No FAA 2000 7909 Amdt 25 111 68 FR 45059 July 31 2003 Amdt 25 128 74 FR 25645 May 29 2009 210 ASA Part 25 Airworthiness Standards Transport Category APPENDIX G TO PART 25 CONTINUOUS GUST DESIGN CRITERIA The continuous gust design criteria in this ap pendix must be used in establishing the dynamic response of the airplane to vertical and lateral continuous turbulence unless a more rational cri teria is used The following gust load require ments apply to mission analysis and design enve lope analysis a The limit gust loads utilizing the continuous turbulence concept must be determined in accor dance with the provisions of either paragraph b or paragraphs c and d of this appendix b Design envelope analysis The limit loads must be determined in accordance with the fol lowing 1 All critical altitudes weights and weight dis tributions as specified in 825 321 b and all criti cal speeds within the ranges indicated in para graph b 3 of this appendix must be considered 2 Values of A8 ratio of root mean square in cremental load root mean square gust velocity must be determined by dynamic analysis The power spectral density of the atmospheric turbu lence must be as given by the equation 1 85 1 339 Lo 1 1 33910 6 Q o L n where power spectr
36. A significant transient may lead to a signifi cant reduction in safety margins an increase in flightcrew workload discomfort to the flightcrew or physical distress to the passengers or cabin crew possibly including non fatal injuries Signifi cant transients do not require in order to remain within or recover to the normal flight envelope any of the following i Exceptional strength ii Forces applied by the pilot which are greater than those specified in 25 143 c iii Accelerations or attitudes in the airplane that might result in further hazard to secured or non secured occupants Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50598 Oct 30 1978 Amdt 25 119 71 FR 18191 April 11 2006 piloting skill alertness or ASA 825 1333 825 1331 Instruments using a power supply a For each instrument required by 25 1303 b that uses a power supply the following apply 1 Each instrument must have a visual means integral with the instrument to indicate when power adequate to sustain proper instrument per formance is not being supplied The power must be measured at or near the point where it enters the instruments For electric instruments the power is considered to be adequate when the voltage is within approved limits 2 Each instrument must in the event of the failure of one power source be supplied by an other power source This may be accomplished automat
37. At speeds between Vrgc Mrc and the direction of the primary longitudinal control force may not reverse c Except as provided in paragraphs d and e of this section compliance with the provisions of paragraph a of this section must be demon strated in flight over the acceleration range 1 1 g to 2 5 g 2 0 g to 2 0 g and extrapolating by an accept able method to 1 g and 2 5 g d If the procedure set forth in paragraph c 2 of this section is used to demonstrate compliance and marginal conditions exist during flight test with regard to reversal of primary longitudinal control force flight tests must be accomplished from the normal acceleration at which a marginal condition is found to exist to the applicable limit specified in paragraph b 1 of this section e During flight tests required by paragraph a of this section the limit maneuvering load factors prescribed in 25 333 b and 25 337 and the maneuvering load factors associated with proba 33 825 255 ble inadvertent excursions beyond the boundaries of the buffet onset envelopes determined under 25 251 e need not be exceeded In addition the entry speeds for flight test demonstrations at normal acceleration values less than 1 g must be limited to the extent necessary to accomplish a recovery without exceeding VpE Mpr f In the out of trim condition specified in para graph a of this section it must be possible fro
38. Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 64 53 FR 17646 May 17 1988 55 825 563 825 563 Structural ditching provisions Structural strength considerations of ditching provisions must be in accordance with 825 801 e FATIGUE EVALUATION 825 571 Damage tolerance and fatigue evaluation of structure a General An evaluation of the strength de tail design and fabrication must show that cata strophic failure due to fatigue corrosion manu facturing defects or accidental damage will be avoided throughout the operational life of the air plane This evaluation must be conducted in ac cordance with the provisions of paragraphs b and e of this section except as specified in paragraph c of this section for each part of the structure that could contribute to a catastrophic failure such as wing empennage control sur faces and their systems the fuselage engine mounting landing gear and their related primary attachments For turbojet powered airplanes those parts that could contribute to a catastrophic failure must also be evaluated under paragraph d of this section In addition the following apply 1 Each evaluation required by this section must include i The typical loading spectra temperatures and humidities expected in service ii The identification of principal structural ele ments and detail design points the failure of which could cause catastrophic failu
39. Each EWIS associated with independent airplane power sources or power sources con nected in combination must be designed and in stalled to ensure adequate physical separation and electrical isolation so that a fault in any one airplane power source EWIS will not adversely af fect any other independent power sources In ad dition 1 Airplane independent electrical power Sources must not share a common ground termi nating location 2 Airplane system static grounds must not share a common ground terminating location with any of the airplane s independent electrical power Sources e Except to the extent necessary to provide electrical connection to the fuel systems compo nents the EWIS must be designed and installed with adequate physical separation from fuel lines and other fuel system components so that 1 An EWIS component failure will not create a hazardous condition 2 Any fuel leakage onto EWIS components will not create a hazardous condition f Except to the extent necessary to provide electrical connection to the hydraulic systems components EWIS must be designed and in stalled with adequate physical separation from hydraulic lines and other hydraulic system compo nents so that 1 An EWIS component failure will not create a hazardous condition 2 Any hydraulic fluid leakage onto EWIS com ponents will not create a hazardous condition g Except to the extent necessary to provide electrical connecti
40. Flight data recorders 25 1459 Flight guidance system 25 1329 Flight instruments 25 1303 Flight load factors 25 321 Flight maneuvering envelope 25 333 Flight manual airplane operating limitations 25 1583 operating procedures 25 1585 Fluid draining of 25 1455 Fuel jettisoning system controls 25 1161 jettisoning system 25 1001 Fuel pumps 25 997 Fuel system airworthiness standards 25 951 25 981 97 102 analysis tests 25 952 components 25 991 25 1001 i fuel flow 25 955 25 957 hot weather operation 25 967 lightning protection 25 954 pressure fueling 25 979 tanks 25 963 25 977 Fuel tank explosion prevention 25 981 Fuel Tank Flammability Assessment Method User s Manual 25 5 25 981 Appendix 12 102 232 Fuel tank flammability exposure Appendix N Fuel tank system flammability Appendix M G Gust and turbulence loads 25 341 39 H High energy rotors in equipment 25 1467 135 High lift devices 25 345 40 High intensity radiated fields HIRF protection from 825 1317 sssssanus 118 HIRF environments and equipment testlevels Appendix L Hulls 25 755 Hydraulic systems 25 ASA Part 25 Index Send additional term
41. No discharge of the extinguisher can cause structural damage 2 The capacity of each required built in fire extinguishing system must be adequate for any fire likely to occur in the compartment where used considering the volume of the compartment and the ventilation rate Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 74 56 FR 15456 April 16 1991 25 853 Compartment interiors For each compartment occupied by the crew or passengers the following apply a Materials including finishes or decorative surfaces applied to the materials must meet the applicable test criteria prescribed in Part of Ap pendix F of this part or other approved equivalent methods regardless of the passenger capacity of the airplane b Reserved c In addition to meeting the requirements of paragraph a of this section seat cushions ex cept those on flight crewmember seats must meet the test requirements of part of Appendix F of this part or other equivalent methods re gardless of the passenger capacity of the air plane d Except as provided in paragraph e of this section the following interior components of air planes with passenger capacities of 20 or more must also meet the test requirements of parts IV and V of Appendix F of this part or other ap proved equivalent method in addition to the flam ASA 825 854 mability requirements prescribed in paragraph a of this section 1 Interio
42. Seat except as the Administrator may require Ex cept as required by subparagraph g of this para graph no employee of the applicant may be Seated next to an emergency exit j Seat belts and shoulder harnesses as re quired must be fastened k Before the start of the demonstration ap proximately one half of the total average amount of carry on baggage blankets pillows and other similar articles must be distributed at several loca tions in aisles and emergency exit access ways to create minor obstructions I No prior indication may be given to any crewmember or passenger of the particular exits to be used in the demonstration m The applicant may not practice rehearse or describe the demonstration for the participants nor may any participant have taken part in this type of demonstration within the preceding 6 months n Prior to entering the demonstration aircraft the passengers may also be advised to follow di rections of crewmembers but may not be in structed on the procedures to be followed in the demonstration except with respect to safety pro cedures in place for the demonstration or which have to do with the demonstration site Prior to the start of the demonstration the pre takeoff passenger briefing required by 121 571 may be given Flight attendants may assign demonstra tion subjects to assist persons from the bottom of a slide consistent with their approved training program The airplane m
43. a A placard meeting the requirements of this section must be installed on or near the mag netic direction indicator b The placard must show the calibration of the instrument in level flight with the engines op erating c The placard must state whether the calibra tion was made with radio receivers on or off d Each calibration reading must be in terms of magnetic heading in not more than 45 degree increments 825 1549 Powerplant and auxiliary power unit instruments For each required powerplant and auxiliary power unit instrument as appropriate to the type of instrument a Each maximum and if applicable minimum safe operating limit must be marked with a red ra dial or a red line b Each normal operating range must be marked with a green arc or green line not extend ing beyond the maximum and minimum safe limits c Each takeoff and precautionary range must be marked with a yellow arc or a yellow line and d Each engine auxiliary power unit or pro peller speed range that is restricted because of excessive vibration stresses must be marked with red arcs or red lines Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15044 March 17 1977 825 1551 Oil quantity indication Each oil quantity indicating means must be marked to indicate the quantity of oil readily and accurately Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29786 Ju
44. as defined in 49 CFR part 572 subpart F Torso contact during rebound is acceptable and need not be measured 4 Pelvis If the pelvis of an ATD at any seat place impacts seat and or adjacent structure dur ing testing pelvic lateral acceleration injury crite ria must be substantiated by dynamic test or by rational analysis based on previous test s of a similar seat installation Pelvic lateral acceleration may not exceed 130g Pelvic acceleration data must be processed as defined in FMVSS part 571 214 section 56 13 5 49 CFR 571 214 5 Body to Wall Furnishing Contact If the seat is installed aft of a structure such as an interior wall or furnishing that may contact the pelvis up per arm chest or head of an occupant seated next to the structure the structure or a conserva tive representation of the structure and its stiff ness must be included in the tests It is recom mended but not required that the contact surface ofthe actual structure be covered with at least two inches of energy absorbing protective padding foam or equivalent such as Ensolite 6 Shoulder Strap Loads Where upper torso straps shoulder straps are used for sofa occu pants the tension loads in individual straps may not exceed 1 750 pounds If dual straps are used for restraining the upper torso the total strap ten sion loads may not exceed 2 000 pounds 7 Occupant Retention All side facing seats require end closures or other means to pr
45. complish and would not contribute materially to the objective sought and the Administrator finds that the experience with the DC 3 or L 18 air planes justifies it he is authorized to accept such measures of compliance as he finds will effec tively accomplish the basic objective 4 Establishment of new maximum certificated weights An applicant for approval of new maxi mum certificated weights shall apply for an amendment of the airworthiness certificate of the airplane and shall show that the weights sought have been established and the appropriate man ual material obtained as provided in this section ASA Part 25 Airworthiness Standards Transport Category Note Transport category performance require ments result in the establishment of maximum certificated weights for various altitudes Weights 25 200 to 26 900 for the DC 3 and 18 500 to 19 500 for the L 18 New maximum certificated weights of more than 25 200 but not more than 26 900 pounds for DC 3 and more than 18 500 but not more than 19 500 pounds for L 18 airplanes may be established in accordance with the transport category performance requirements of either Part 4a or Part 4b if the airplane at the new maximum weights can meet the structural re quirements of the elected part b Weights of more than 26 900 for the DC 3 and 19 500 for the L 18 New maximum certifi cated weights of more than 26 900 pounds for DC 3 and 19 500 pounds for L 18 airplanes s
46. face 3 At each point along the takeoff path start ing at the point at which the airplane reaches 400 feet above the takeoff surface the available gradi ent of climb may not be less than i 1 2 percent for two engine airplanes ii 1 5 percent for three engine airplanes and iii 1 7 percent for four engine airplanes 19 825 111 4 The airplane configuration may not be changed except for gear retraction and automatic propeller feathering and no change in power or thrust that requires action by the pilot may be made until the airplane is 400 feet above the take off surface and 5 If 25 105 a 2 requires the takeoff path to be determined for flight in icing conditions the air borne part of the takeoff must be based on the air plane drag i With the takeoff ice accretion defined in ap pendix C from a height of 35 feet above the take off surface up to the point where the airplane is 400 feet above the takeoff surface and ii With the final takeoff ice accretion defined in appendix C from the point where the airplane is 400 feet above the takeoff surface to the end of the takeoff path 4 Except for gear retraction and propeller feathering the airplane configuration may not be changed and no change in power or thrust that requires action by the pilot may be made until the airplane is 400 feet above the takeoff surface d The takeoff path must be determined by a continuous demonstrated takeoff
47. lowed in the analysis The analysis must be done in accordance with the methods and procedures set forth in the Fuel Tank Flammability Assess ment Method Users Manual dated May 2008 document number DOT FAA AR 05 8 incorpo rated by reference see 825 5 The parameters specified in sections N25 3 b and c of this ap pendix must be used in the fuel tank flammability exposure Monte Carlo analysis b The following parameters are defined in the Monte Carlo analysis and provided in paragraph N25 4 of this appendix 1 Cruise Ambient Temperature as defined in this appendix 2 Ground Ambient Temperature as defined in this appendix 3 Fuel Flash Point as defined in this appen dix 4 Flight Length Distribution as defined in Ta ble 2 of this appendix 5 Airplane Climb and Descent Profiles as de fined in the Fuel Tank Flammability Assessment Method Users Manual dated May 2008 docu ment number DOT FAA AR 05 8 incorporated by reference in 825 5 c Parameters that are specific to the particu lar airplane model under evaluation that must be provided as inputs to the Monte Carlo analysis are 1 Airplane cruise altitude 2 Fuel tank quantities If fuel quantity affects fuel tank flammability inputs to the Monte Carlo analysis must be provided that represent the ac tual fuel quantity within the fuel tank or compart ment of the fuel tank throughout each of the flights being evaluated Input values for this
48. on both sides of the horizontal 30 total about the most critical axis for 25 hours If motion about more than one axis is likely to be critical the tank must be rocked about each critical axis for 12 hours c Except where satisfactory operating experi ence with a similar tank in a similar installation is shown nonmetallic tanks must withstand the test 99 825 967 specified in paragraph b 5 of this section with fuel at a temperature of 110 F During this test a representative specimen of the tank must be in stalled in a supporting structure simulating the in stallation in the airplane d For pressurized fuel tanks it must be shown by analysis or tests that the fuel tanks can with stand the maximum pressure likely to occur on the ground or in flight Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6913 May 5 1967 Amdt 25 40 42 FR 15043 March 17 1977 825 967 Fuel tank installations a Each fuel tank must be supported so that tank loads resulting from the weight of the fuel in the tanks are not concentrated on unsupported tank surfaces In addition 1 There must be pads if necessary to pre vent chafing between the tank and its supports 2 Padding must be nonabsorbent or treated to prevent the absorption of fluids 3 If a flexible tank liner is used it must be sup ported so that it is not required to withstand fluid loads and 4 Each interior surf
49. their operating mechanisms and their supporting structures must be designed for critical loads occurring in the conditions pre scribed in 825 345 accounting for the loads oc curring during transition from one flap position and airspeed to another 825 459 Special devices The loading for special devices using aerody namic surfaces such as slots slats and spoilers must be determined from test data Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29776 July 20 1990 GROUND LOADS 825 471 General a Loads and equilibrium For limit ground loads 1 Limit ground loads obtained under this sub part are considered to be external forces applied to the airplane structure and 2 In each specified ground load condition the external loads must be placed in equilibrium with the linear and angular inertia loads in a rational or conservative manner b Critical centers of gravity The critical cen ters of gravity within the range for which certifica tion is requested must be selected so that the maximum design loads are obtained in each land ing gear element Fore and aft vertical and lat eral airplane centers of gravity must be consid ered Lateral displacements of the c g from the airplane centerline which would result in main gear loads not greater than 103 percent of the critical design load for symmetrical loading condi tions may be selected without considering the ef fects of t
50. to prevent de velopment of ignition sources within the fuel tank System pursuant to paragraph a of this section to prevent increasing the flammability exposure of the tanks above that permitted under paragraph b of this section and to prevent degradation of the performance and reliability of any means pro vided according to paragraphs a or c of this section These CDCCL inspections and proce dures must be included in the Airworthiness Limi tations section of the instructions for continued airworthiness required by 25 1529 Visible means of identifying critical features of the design must be placed in areas of the airplane where foreseeable maintenance actions repairs or al terations may compromise the critical design con figuration control limitations e g color coding of wire to identify separation limitation These visi ble means must also be identified as CDCCL Docket No FAA 1999 6411 66 FR 23130 May 2001 as amended by Amdt 25 125 73 FR 42494 July 21 2008 102 Federal Aviation Regulations FUEL SYSTEM COMPONENTS 25 991 Fuel pumps a Main pumps Each fuel pump required for proper engine operation or required to meet the fuel system requirements of this subpart other than those in paragraph b of this section is a main pump For each main pump provision must be made to allow the bypass of each positive dis placement fuel pump other than a fuel injection pump a pump that supplies th
51. wing flaps i retracted and ii extended the most unfavorable center of gravity position approved for landing with the maximum landing weight and with the most unfavorable center of gravity posi tion approved for landing regardless of weight and 3 Level flight at any speed from 1 3 to with the landing gear and flaps re tracted and from 1 3 Vsg4 to Vi g with the landing gear extended d Longitudinal directional and lateral trim The airplane must maintain longitudinal direc tional and lateral trim and for the lateral trim the angle of bank may not exceed five degrees at 1 3 Vsri during climbing flight with 1 The critical engine inoperative 2 The remaining engines at maximum contin uous power and 3 The landing gear and flaps retracted e Airplanes with four or more engines Each airplane with four or more engines must maintain trim in rectilinear flight 1 At the climb speed configuration and power required by 25 123 a for the purpose of establishing the rate of climb 2 With the most unfavorable center of gravity position and 3 At the weight at which the two engine inop erative climb is equal to at least 0 013 at an altitude of 5 000 feet Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5671 April 8 1970 Amdt 25 38 41 FR 55466 Dec 20 1976 Amdt 25 108 67 FR 70827 Nov 26 2002 STABILITY 825 171 Genera
52. 1 and 2 of this section may be used if substantiated by data on actual engine oil consumption 25 1013 Oil tanks a Installation Each oil tank installation must meet the requirements of 25 967 b Expansion space Oil tank expansion space must be provided as follows 1 Each oil tank used with a reciprocating en gine must have an expansion space of not less than the greater of 10 percent of the tank capacity or 0 5 gallon and each oil tank used with a tur bine engine must have an expansion space of not less than 10 percent of the tank capacity 2 Each reserve oil tank not directly connected to any engine may have an expansion space of not less than two percent of the tank capacity 104 Federal Aviation Regulations 3 It must be impossible to fill the expansion space inadvertently with the airplane in the nor mal ground attitude c Filler connection Each recessed oil tank filler connection that can retain any appreciable quantity of oil must have a drain that discharges clear of each part of the airplane In addition each oil tank filler cap must provide an oil tight seal d Vent Oil tanks must be vented as follows 1 Each oil tank must be vented from the top part of the expansion space so that venting is ef fective under any normal flight condition 2 Oil tank vents must be arranged so that condensed water vapor that might freeze and ob struct the line cannot accumulate at any point e Outl
53. 125 45 40 35 100 30 25 0 50 100 150 TIME MINUTES 200 250 TIME TEMPERATURE RELATIONSHIP Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 41 42 FR 36970 July 18 1977 Amdt 25 87 61 FR 28695 June 5 1996 Amdt 25 89 61 FR 63956 Dec 2 1996 825 832 Cabin ozone concentration a The airplane cabin ozone concentration during flight must be shown not to exceed 1 0 25 parts per million by volume sea level equivalent at any time above flight level 320 and 2 0 1 parts per million by volume sea level equivalent time weighted average during any 3 hour interval above flight level 270 b For the purpose of this section sea level equivalent refers to conditions of 25 and 760 millimeters of mercury pressure c Compliance with this section must be shown by analysis or tests based on airplane op erational procedures and performance limitations that demonstrate that either 1 The airplane cannot be operated at an alti tude which would result in cabin ozone concentra tions exceeding the limits prescribed by para graph a of this section or 2 The airplane ventilation system including any ozone control equipment will maintain cabin ozone concentrations at or below the limits pre Scribed by paragraph a of this section Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 50 45 FR 3883 Jan 1 1980 Amdt 25 56 47
54. 13 2000 or as subsequently amended or ii Sections 33 77 and 33 78 of this chapter in effect on April 30 1998 or as subsequently amended before December 13 2000 or iii Comply with 33 77 of this chapter in effect on October 31 1974 or as subsequently amended prior to April 30 1998 unless that en gine s foreign object ingestion service history has resulted in an unsafe condition or 94 Federal Aviation Regulations iv Be shown to have a foreign object ingestion Service history in similar installation locations which has not resulted in any unsafe condition b Engine isolation The powerplants must be arranged and isolated from each other to allow operation in at least one configuration so that the failure or malfunction of any engine or of any sys tem that can affect the engine will not 1 Prevent the continued safe operation of the remaining engines or 2 Require immediate action by any ber for continued safe operation c Control of engine rotation There must be means for stopping the rotation of any engine in dividually in flight except that for turbine engine installations the means for stopping the rotation of any engine need be provided only where con tinued rotation could jeopardize the safety of the airplane Each component of the stopping system on the engine side of the firewall that might be ex posed to fire must be at least fire resistant If hy draulic propeller f
55. 138 139 Minimum flight crew criteria for determining Appendix D 162 Mixture controls 25 1147 N Nacelle areas behind firewalls 25 1782 iu 111 Nacelle skin 25 1193 113 Navigation instruments 9525 1303 115 ASA Part 25 Index Oil filter 25 1019 fittings 25 1017 lines 25 1017 radiators 25 1023 strainer 25 1019 system drains 25 1021 System general 25 1011 tank tests 25 7075 tanks 25 1073 valves 25 1025 Operating limitations airspeed 25 1503 25 1517 136 auxiliary power unit 25 1522 137 for airworthiness transport category airplanes 25 1501 1533 136 137 maneuvering flight load factors S25 1531 137 minimum flight crew 25 1523 Appendix D powerplant 25 1521 weight center of gravity and weight distribution 25 1519 136 Oxygen distributing system standards 25 1445 distributing units standards 925 1447 m equipment protection from rupture 25 1453 133 generators chemical 25 1450 means for determining use 25 1449 supplemental flow of 25 1443 Oxygen equipment supply 25 1441 P Performance general Position lights color specifications 25 1397 dihedral angle required 25 1387 distribution intensity 25 1389 installation
56. 154 ASA Part 25 Airworthiness Standards Transport Category APPENDIX C TO PART 25 PART ATMOSPHERIC ICING CONDITIONS a Continuous maximum icing The maximum continuous intensity of atmospheric icing condi tions continuous maximum icing is defined by the variables of the cloud liquid water content the mean effective diameter of the cloud droplets the ambient air temperature and the interrelationship of these three variables as shown in figure 1 of this appendix The limiting icing envelope in terms of altitude and temperature is given in figure 2 of this appendix The inter relationship of cloud liq uid water content with drop diameter and altitude is determined from figures 1 and 2 The cloud liq uid water content for continuous maximum icing conditions of a horizontal extent other than 17 4 nautical miles is determined by the value of liquid water content of figure 1 multiplied by the appro priate factor from figure 3 of this appendix b Intermittent maximum icing The intermit tent maximum intensity of atmospheric icing con ditions intermittent maximum icing is defined by the variables of the cloud liquid water content the mean effective diameter of the cloud droplets the ambient air temperature and the interrelationship of these three variables as shown in figure 4 of this appendix The limiting icing envelope in terms of altitude and temperature is given in figure 5 of this appendix The inter relationship
57. 160 microlamberts or ii Be conspicuously located and well illumi nated by the emergency lighting even in condi tions of occupant crowding at the exit 3 Reserved 4 Each Type A Type B Type C Type I or Type Il passenger emergency exit with a locking mech anism released by rotary motion of the handle must be marked i With a red arrow with a shaft at least three fourths of an inch wide and a head twice the width of the shaft extending along at least 70 degrees of arc at a radius approximately equal to three fourths of the handle length ii So that the centerline of the exit handle is within 1 inch of the projected point of the arrow when the handle has reached full travel and has released the locking mechanism and iii With the word open in red letters 1 inch high placed horizontally near the head of the ar row f Each emergency exit that is required to be openable from the outside and its means of opening must be marked on the outside of the airplane In addition the following apply 1 The outside marking for each passenger emergency exit in the side of the fuselage must include a 2 inch colored band outlining the exit 2 Each outside marking including the band must have color contrast to be readily distinguish able from the surrounding fuselage surface The contrast must be such that if the reflectance of the darker color is 15 percent or less the reflectance of the lighter color must be at
58. 1964 as amended by Amdt 25 23 35 FR 5674 April 8 1970 57 825 601 Subpart D Design and Construction GENERAL 25 601 General The airplane may not have design features or details that experience has shown to be hazard ous or unreliable The suitability of each question able design detail and part must be established by tests 25 603 Materials The suitability and durability of materials used for parts the failure of which could adversely af fect safety must a Be established on the basis of experience or tests b Conform to approved specifications such as industry or military specifications or Technical Standard Orders that ensure their having the strength and other properties assumed in the de sign data and c Take into account the effects of environmen tal conditions such as temperature and humidity expected in service Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55466 Dec 20 1976 Amdt 25 46 43 FR 50595 Oct 30 1978 825 605 Fabrication methods a The methods of fabrication used must pro duce a consistently sound structure If a fabrica tion process such as gluing spot welding or heat treating requires close control to reach this ob jective the process must be performed under an approved process specification b Each new aircraft fabrication method must be substantiated by a test program Docket No 5066 29 FR 18291 Dec
59. 2 and 55 10 relative humidity for a minimum of 24 hours prior to testing e Apparatus Calibration 1 With the sliding platform out of the chamber install the calorimeter holding frame Push the platform back into the chamber and insert the cal orimeter into the first hole zero position See figure 7 Close the bottom door located below the sliding platform The distance from the centerline of the calorimeter to the radiant panel surface at this point must be 7 1 2 inches 1 8 191 mm 3 Prior to igniting the radiant panel ensure that the calorimeter face is clean and that there is water running through the calorimeter 2 Ignite the panel Adjust the fuel air mixture to achieve 1 5 BTUs ft second 5 1 7 Watts cm 5 at the zero position If using an electric panel set the power controller to achieve the proper heat flux Allow the unit to reach steady state this may take up to 1 hour The pilot burner must be off and in the down position during this time 197 Appendix to Part 25 3 After steady state conditions have been reached move the calorimeter 2 inches 51 mm from the zero position first hole to position 1 and record the heat flux Move the calorimeter to position 2 and record the heat flux Allow enough time at each position for the calorimeter to stabi lize Table 1 depicts typical calibration values at the three positions Table 1 Calibration Table Position BTUs ft2
60. 2007 825 1355 Distribution system a The distribution system includes the distri bution busses their associated feeders and each control and protective device b Reserved c If two independent sources of electrical power for particular equipment or systems are re quired by this chapter in the event of the failure of one power source for such equipment or system another power source including its separate feeder must be automatically provided or be manually selectable to maintain equipment or System operation Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5679 April 8 1970 Amdt 25 38 41 FR 55468 Dec 20 1976 825 1357 Circuit protective devices a Automatic protective devices must be used to minimize distress to the electrical system and hazard to the airplane in the event of wiring faults or serious malfunction of the system or connected equipment b The protective and control devices in the generating system must be designed to de ener gize and disconnect faulty power sources and power transmission equipment from their associ ated busses with sufficient rapidity to provide pro tection from hazardous over voltage and other malfunctioning c Each resettable circuit protective device must be designed so that when an overload or circuit fault exists it will open the circuit irrespec tive of the position of the operating control d If the ability to reset a
61. 24 1964 as amended by Amdt 25 46 43 FR 50595 Oct 30 1978 825 607 Fasteners a Each removable bolt screw nut pin or other removable fastener must incorporate two separate locking devices if 1 Its loss could preclude continued flight and landing within the design limitations of the air plane using normal pilot skill and strength or 2 Its loss could result in reduction in pitch yaw or roll control capability or response below that required by Subpart B of this chapter b The fasteners specified in paragraph a of this section and their locking devices may not be adversely affected by the environmental condi tions associated with the particular installation c No self locking nut may be used on any bolt subject to rotation in operation unless a nonfriction 58 Federal Aviation Regulations locking device is used in addition to the self locking device Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5674 April 8 1970 825 609 Protection of structure Each part of the structure must a Be suitably protected against deterioration or loss of strength in service due to any cause including 1 Weathering 2 Corrosion and 3 Abrasion and b Have provisions for ventilation and drainage where necessary for protection 825 611 Accessibility provisions a Means must be provided to allow inspection including inspection of principal structural ele
62. 24 hours Each specimen must re main in the conditioning environment until it is subjected to the flame 2 Specimen configuration Except for small parts and electrical wire and cable insulation ma terials must be tested either as section cut from a fabricated part as installed in the airplane or as a specimen simulating a cut section such as a specimen cut from a flat sheet of the material or a model of the fabricated part The specimen may be cut from any location in a fabricated part how ever fabricated units such as sandwich panels may not be separated for test Except as noted below the specimen thickness must be no thicker than the minimum thickness to be qualified for use in the airplane Test specimens of thick foam parts such as seat cushions must be inch in thickness Test specimens of materials that must meet the requirements of paragraph a 1 v of part of this appendix must be no more than 1 inch in thickness Electrical wire and cable speci mens must be the same size as used in the air plane In the case of fabrics both the warp and fill direction of the weave must be tested to deter mine the most critical flammability condition Specimens must be mounted in a metal frame so that the two long edges and the upper edge are held securely during the vertical test prescribed in subparagraph 4 of this paragraph and the two long edges and the edge away from the flame are held securely during the horizontal tes
63. 25 Airworthiness Standards Transport Category structural failures described in paragraphs d 4 and d 5 of this section need not be considered in showing compliance with this section if i The structural element could not fail due to discrete source damage resulting from the condi tions described in 25 571 e and ii A damage tolerance investigation in accor dance with 25 571 b shows that the maximum extent of damage assumed for the purpose of re sidual strength evaluation does not involve com plete failure of the structural element 9 Any damage failure or malfunction consid ered under 25 631 25 671 25 672 and 25 1309 10 Any other combination of failures malfunc tions or adverse conditions not shown to be ex tremely improbable e Flight flutter testing Full scale flight flutter tests at speeds up to must be conducted for new type designs and for modifications to a type design unless the modifications have been shown to have an insignificant effect on the aeroelastic stability These tests must demon strate that the airplane has a proper margin of damping at all speeds up to VpF Mpr and that there is no large and rapid reduction in damping as Vpr Mpr is approached If a failure malfunc tion or adverse condition is simulated during flight test in showing compliance with paragraph d of this section the maximum speed investi gated need not exceed Vrc Mrc if it is shown by co
64. 25 1385 minimum maximum intensities required tables 25 1391 25 1393 25 1395 Power source capacity 25 1310 Powerplant accessories 25 1163 110 airworthiness standards transport category airplanes 25 901 25 945 augmentation systems 25 945 automatic takeoff thrust control system ATTCS 25 904 controls general 25 1141 engines 25 903 inlet engine exhaust installation 25 901 negative acceleration 25 943 propeller clearance 25 925 propeller deicing 25 929 propeller vibration and fatigue 25 907 propeller drag limiting systems 25 937 Propellers 25 905 25 101 caneinenesnes 14 thrust reversers systems 825 933 25 934 96 turbine engine operating characteristics 25 930 eU EORR 96 239 Pressurization pneumatic systems 25 1438 130 Pressurized compartment loads 25 365 41 Propeller feathering controls 25 17153 110 feathering system 25 1027 s pitch controls 25 1149 ees pitch settings below flight regime 25 1155 110 reinforcement 25 875 speed controls 25 1149 speed pitch limits 25 33 Protective breathing equipment 825 1439 130 Public address system 25 1423 129 R Reverse thrust below the
65. 25 1531 25 1533 25 1535 25 1541 25 1543 25 1545 25 1547 25 1549 25 1551 25 1553 25 1555 25 1557 25 1561 25 1563 25 1581 25 1583 25 1585 25 1587 Weight center of gravity and weight distribution Powerplant limitations Auxiliary power unit limitations Minimum flight crew Kinds of operation Ambient air temperature and operating altitude Instructions for Continued Airworthiness Maneuvering flight load factors Additional operating limitations ETOPS approval MARKINGS AND PLACARDS General Instrument markings general Airspeed limitation information Magnetic direction indicator Powerplant and auxiliary power unit instruments Oil quantity indication Fuel quantity indicator Control markings Miscellaneous markings and placards Safety equipment Airspeed placard AIRPLANE FLIGHT MANUAL General Operating limitations Operating procedures Performance information Subpart H Electrical Wiring Interconnection Systems EWIS 25 1701 25 1703 25 1705 25 1707 25 1709 25 1711 25 1713 25 1715 25 1717 25 1719 25 1721 25 1723 25 1725 25 1727 25 1729 25 1731 25 1733 Definition Function and installation EWIS Systems and functions EWIS System separation EWIS System safety EWIS Component identification EWIS Fire protection EWIS Electrical bonding and protection against static electricity EWIS Circuit protective devices EWIS Ac
66. 4400 4600 0 0 0 0 0 0 0 0 1 0 2 0 2 3 25 2 5 24 4600 4800 0 0 0 0 0 0 0 0 0 6 1 5 1 8 2 0 2 0 2 0 4800 5000 0 0 0 0 0 0 0 0 02 1 0 14 1 5 1 6 1 5 5000 5200 0 0 0 0 0 0 0 0 0 0 0 8 1 1 1 3 1 3 1 3 5200 5400 0 0 0 0 0 0 0 0 0 0 0 8 12 15 1 6 1 6 5400 5600 0 0 0 0 0 0 0 0 0 0 0 9 ET 2 1 22 23 5600 5800 0 0 0 0 0 0 0 0 0 0 0 6 1 6 2 2 24 2 5 5800 6000 0 0 0 0 0 0 0 0 0 0 02 1 8 24 2 8 2 9 6000 6200 0 0 0 0 0 0 0 0 0 0 0 0 1 7 2 6 3 1 3 3 6200 6400 0 0 0 0 0 0 0 0 0 0 0 0 14 24 2 9 3 1 6400 6600 0 0 0 0 0 0 0 0 0 0 0 0 0 9 1 8 22 2 5 6600 6800 0 0 0 0 0 0 0 0 0 0 0 0 0 5 1 2 1 6 1 9 6800 7000 0 0 0 0 0 0 0 0 0 0 0 0 0 2 0 8 1 1 1 3 7000 7200 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 4 0 7 0 8 7200 7400 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 3 0 5 0 7 7400 7600 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 2 0 5 0 6 7600 7800 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 1 0 5 0 7 7800 8000 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 1 0 6 0 8 8000 8200 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 5 0 8 8200 8400 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 5 1 0 8400 8600 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 6 1 3 8600 8800 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 4 1 1 8800 9000 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 2 0 8 9000 9200 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 5 9200 9400 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 2 9400 9600 0 0 0 0 0 0 0 0 0 0 00 0 0 0 0 0 0 0 1 9600 9800 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 1 9800 10000 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 1 ASA 235 825 1733 c Overnight Temperature Drop For air planes on which FRM i
67. 825 1709 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5679 April 8 1970 Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 41 42 FR 36970 July 18 1977 Amdt 25 123 72 FR 63405 Nov 8 2007 825 1310 Power source capacity distribution a Each installation whose functioning is re quired for type certification or under operating rules and that requires a power supply is an es sential load on the power supply The power sources and the system must be able to supply the following power loads in probable operating combinations and for probable durations 1 Loads connected to the system with the System functioning normally 2 Essential loads after failure of any one prime mover power converter or energy storage device 3 Essential loads after failure of i Any one engine on two engine airplanes and ii Any two engines on airplanes with three or more engines ASA 825 1316 4 Essential loads for which an alternate Source of power is required after any failure or malfunction in any one power supply system dis tribution system or other utilization system b In determining compliance with paragraphs 2 and 3 of this section the power loads may be assumed to be reduced under a monitoring procedure consistent with safety in the kinds of operation authorized Loads not required in con trolled flight need not be considered for the two engine inope
68. AINON 5 9NIOI OIH3IHdSOINIV 0 sano19 WHOSIINWND WAWIXVW LNSLLINYSLNI ASA 160 Part 25 Airworthiness Standards Transport Category Appendix C to Part 25 INTERMITTENT MAXIMUM CUMULIFORM CLOUDS ATMOSPHERIC ICING CONDITIONS AMBIENT TEMPERATURE VS PRESSURE ALTITUDE SOURCE OF DATA NACA TN NO 2569 32 26 20 AMBIENT TEMPERATURE F T 16 Z T V 7 PA RM G 28 DASHED LINES INDICATE Z 7 POSSIBLE OF 7171 7 17171 LIMITS TAA IE oe 34 y iB 4 8 12 16 20 24 28 30 PRESSURE ALTITUDE 1000 FEET CII FIGURE 5 ASA 161 Federal Aviation Regulations Appendix C to Part 25 9 d8fn9l4 SATIN TVOLLOVN LN3 LX3 IVLNOZIHOH GL Ol 80 9050 VO 0 0909 Ov OF 0c IVLNOZIHOH HLIM YOLOVA LN31NOO HZlVM JO NOLIVIHVA SNOILIQNOO SGNO1D WHOSIINNND WOWIXVIN LNALLINYALNI 8226 ON NL VOVN geq Jo e
69. Amdt 25 38 41 FR 55468 Dec 20 1976 Amdt 25 72 55 FR 29786 July 20 1990 Amdt 25 92 63 FR 8321 Feb 18 1998 25 1535 ETOPS approval Except as provided in 25 3 each applicant seeking ETOPS type design approval must com ply with the provisions of Appendix K of this part Docket No FAA 2002 6717 72 FR 1873 Jan 16 2007 MARKINGS AND PLACARDS 25 1541 General a The airplane must contain 1 The specified markings and placards and 2 Any additional information instrument markings and placards required for the safe oper ation if there are unusual design operating or handling characteristics b Each marking and placard prescribed in paragraph a of this section 1 Must be displayed in a conspicuous place and 2 May not be easily erased disfigured or ob scured ASA 825 1551 825 1543 Instrument markings general For each instrument a When markings are on the cover glass of the instrument there must be means to maintain the correct alignment of the glass cover with the face of the dial and b Each instrument marking must be clearly visible to the appropriate crewmember Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29786 July 20 1990 825 1545 Airspeed limitation information The airspeed limitations required by 825 1583 a must be easily read and understood by the flight crew 825 1547 Magnetic direction indicator
70. BURNER mm 1 DIAMETER p 4 6 102 3 mm SSS SSS STEEL ANGLE 4 x 1 x Ve 25 x 25 x 3 mm WATER COOLED CALORIMETER RACK FITS INSIDE SEAT FRAME TOP VIEW CALORIMETER BRACKET ASA Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 Figure 5 10 16 254 3 mm 3 16 76 3 mm 1 25 mm 7 THERMOCOUPLES SIDE VIEW 21 54 T 805 3 mm gt STEEL ANGLE Tx 1 x fe BURNER CONE 25 x 25 x 3mm 4 10243 ii 124 305 3 mm RACKT FITS INSIDE SEAT FRAME TOP VIEW THERMOCOUPLE RAKE BRACKET ASA 175 Appendix to Part 25 PART TEST METHOD TO DETERMINE FLAME PENETRATION RESISTANCE OF CARGO COMPARTMENT LINERS a Criteria for Acceptance 1 At least three specimens of cargo compart ment sidewall or ceiling liner panels must be tested 2 Each specimen tested must simulate the cargo compartment sidewall or ceiling liner panel including any design features such as joints lamp assemblies etc the failure of which would affect the capability of the liner to safely contain a fire 3 There must be no flame penetration of any specimen within 5 minutes after application of the flame source and the peak temperature mea sured at 4 inches above the upper surface of the horizontal test sample must not exceed 400 F b Summary of Method This method provides a lab
71. Early ETOPS in accordance with 833 201 of this chap ter 2 The applicant must design the propulsion System to preclude failures or malfunctions that could result in an IFSD The applicant must show compliance with this requirement by analysis test in service experience on other airplanes or other means acceptable to the FAA If analysis is used the applicant must show that the propulsion System design will minimize failures and malfunc tions with the objective of achieving the following IFSD rates i An IFSD rate of 0 02 or less per 1 000 world fleet engine hours for type design approval up to and including 180 minutes ii An IFSD rate of 0 01 or less per 1 000 world fleet engine hours for type design approval beyond 180 minutes c Maintenance and operational procedures The applicant must validate all maintenance and operational procedures for ETOPS significant Systems The applicant must identify track and resolve any problems found during the validation in accordance with the problem tracking and reso lution system specified in section K25 2 2 h of this appendix d Propulsion system validation test 1 The installed engine configuration for which approval is being sought must comply with 33 201 c of this chapter The test engine must be configured with a complete airplane nacelle package including engine mounted equipment except for any configuration differences neces ASA Part 25 Airworthiness Stand
72. FAA 2005 22997 73 FR 42495 July 21 2008 unless otherwise noted N25 1 GENERAL a This appendix specifies the requirements for conducting fuel tank fleet average flammability exposure analyses required to meet 825 981 b and Appendix M of this part For fuel tanks in stalled in aluminum wings a qualitative assess ment is sufficient if it substantiates that the tank is a conventional unheated wing tank b This appendix defines parameters affecting fuel tank flammability that must be used in per forming the analysis These include parameters that affect all airplanes within the fleet such as a statistical distribution of ambient temperature fuel flash point flight lengths and airplane descent rate Demonstration of compliance also requires application of factors specific to the airplane model being evaluated Factors that need to be in cluded are maximum range cruise mach number typical altitude where the airplane begins initial cruise phase of flight fuel temperature during both ground and flight times and the performance of a flammability reduction means FRM if in stalled c The following definitions input variables and data tables must be used in the program to determine fleet average flammability exposure for a specific airplane model N25 2 DEFINITIONS a Bulk Average Fuel Temperature means the average fuel temperature within the fuel tank or different sections of the tank if the tank is sub divide
73. FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6913 May 5 1967 Amdt 25 36 39 FR 35461 Oct 1 1974 Amdt 25 57 49 FR 6849 Feb 23 1984 Amdt 25 101 65 FR 79710 Dec 19 2000 825 1185 Flammable fluids a Except for the integral oil sumps specified in 825 1183 no tank or reservoir that is a part of a system containing flammable fluids or gases may be in a designated fire zone unless the fluid contained the design of the system the materials used in the tank the shut off means and all con nections lines and control provide a degree of safety equal to that which would exist if the tank or reservoir were outside such a zone b There must be at least one half inch of clear airspace between each tank or reservoir and each firewall or shroud isolating a designated fire zone c Absorbent materials close to flammable fluid system components that might leak must be covered or treated to prevent the absorption of hazardous quantities of fluids Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 19 33 FR 15410 Oct 17 1968 Amdt 25 94 63 FR 8848 Feb 23 1998 825 1187 Drainage and ventilation of fire zones a There must be complete drainage of each part of each designated fire zone to minimize the hazards resulting from failure or malfunctioning of any component containing flammable fluids The drainage means must be 1 Effective under conditions expected to pre
74. FR 58489 Dec 30 1982 Amdt 25 94 63 FR 8848 Feb 23 1998 ASA 25 833 Combustion heating systems Combustion heaters must be approved Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29783 July 20 1990 PRESSURIZATION 825 841 Pressurized cabins a Pressurized cabins and compartments to be occupied must be equipped to provide a cabin pressure altitude of not more than 8 000 feet at the maximum operating altitude of the airplane under normal operating conditions 1 If certification for operation above 25 000 feet is requested the airplane must be designed that occupants will not be exposed to cabin pressure altitudes in excess of 15 000 feet after any probable failure condition in the pressuriza tion system 2 The airplane must be designed so that oc cupants will not be exposed to a cabin pressure altitude that exceeds the following after decom pression from any failure condition not shown to be extremely improbable i Twenty five thousand 25 000 feet for more than 2 minutes or ii Forty thousand 40 000 feet for any duration 3 Fuselage structure engine and system fail ures are to be considered in evaluating the cabin decompression 87 825 843 b Pressurized cabins must have at least the following valves controls and indicators for con trolling cabin pressure 1 Two pressure relief valves to automatically limit the positive pressure differential
75. FR 6067 April 29 1965 825 1525 Kinds of operation The kinds of operation to which the airplane is limited are established by the category in which it is eligible for certification and by the installed equipment 25 1527 Ambient air temperature and operating altitude The extremes of the ambient air temperature and operating altitude for which operation is al lowed as limited by flight structural powerplant functional or equipment characteristics must be established Docket No FAA 2000 8511 66 FR 34024 June 26 2001 825 1529 Instructions for Continued Airworthiness The applicant must prepare Instructions for Continued Airworthiness in accordance with Ap pendix H to this part that are acceptable to the Administrator The instructions may be incomplete at type certification if a program exists to ensure their completion prior to delivery of the first air plane or issuance of a standard certificate of air worthiness whichever occurs later Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 54 45 FR 60173 Sept 11 1980 825 1531 Maneuvering flight load factors Load factor limitations not exceeding the posi tive limit load factors determined from the maneu vering diagram in 25 333 b must be estab lished 825 1533 Additional operating limitations a Additional operating limitations must be es tablished as follows 1 The maximum takeoff weights must be es tablishe
76. Figure 2A Globar Radiant Panel ASA 187 Appendix to Part 25 Top Federal Aviation Regulations Reflector adjust slope top and bottom for uniform heat flux on sample 14 20 Machine Screw 75 19 Unless denoted otherwise all dimensions are in millimeters Figure 2B Globar Radiant Panel 188 ASA Part 25 Airworthiness Standards Transport Category 24 2a Steel Washer Vo 13 Nut ASA 50 mm OD 13mm ID 2 Radius 5 Flange SPRING ANGLE IRON RETAINER FRAME Washer 50 mm OD NN Spot Weld Frame NN M 165 13 mm ID 10 32 gt Drill amp bm Screw Tap for Weld Uu Nut 10 32 Machine Screw to Washer gt MOUNTING BRACKET Figure 3 Stainless Steel Wire EU AMPLE HOLDER Unless denoted otherwise all dimensions are in millimeters Appendix F to Part 25 189 Appendix to Part 25 Federal Aviation Regulations No 32 Drill Hole 7 9502 9 5 Tubing Leak Free Seal on 6 35 Pilot Tubing 2 Unless denoted otherwise all dimensions are in millimeters Figure 4 190 ASA Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 10 MM Figure 5 Thermocouple Position ASA 191 Appendix to Part 25 PART V TEST METHOD TO DETERMINE THE SMOKE EMISSION CHARACTERISTICS OF CABIN
77. Frow is i 0 3 W7 W7 less than 30 000 pounds ii 6 Wr 450 000 7 for Wr between 30 000 and 100 000 pounds and iii 0 15 Wy for WT over 100 000 pounds b For towing points not on the landing gear but near the plane of symmetry of the airplane the drag and side tow load components specified for the auxiliary gear apply For towing points lo cated outboard of the main gear the drag and side tow load components specified for the main gear apply Where the specified angle of swivel cannot be reached the maximum obtainable an gle must be used c The towing loads specified in paragraph d of this section must be reacted as follows 1 The side component of the towing load at the main gear must be reacted by a side force at the static ground line of the wheel to which the load is applied 2 The towing loads at the auxiliary gear and the drag components of the towing loads at the main gear must be reacted as follows i A reaction with a maximum value equal to the vertical reaction must be applied at the axle of the wheel to which the load is applied Enough airplane inertia to achieve equilibrium must be ap plied ii The loads must be reacted by airplane inertia 49 825 511 d The prescribed towing loads are as follows Federal Aviation Regulations Load Tow point Position Magnitude No Direction Main gear 0 75 FTow per main gear 1 Forward parallel to drag axis unit 2 Forwa
78. Fuel flow Flow between interconnected tanks Unusable fuel supply Fuel system hot weather operation Fuel tanks general Fuel tank tests Fuel tank installations Fuel tank expansion space Fuel tank sump Fuel tank filler connection Fuel tank vents and carburetor vapor vents Fuel tank outlet Pressure fueling system Fuel tank explosion prevention FUEL SYSTEM COMPONENTS Fuel pumps Fuel system lines and fittings Fuel system components Fuel valves Part 25 25 997 Fuel strainer or filter 25 999 Fuel system drains 25 1001 Fuel jettisoning system OiL SYSTEM 25 1011 General 25 1013 Oil tanks 25 1015 Oil tank tests 25 1017 Oil lines and fittings 25 1019 Oil strainer or filter 25 1021 Oil system drains 25 1023 Oil radiators 25 1025 Oil valves 25 1027 Propeller feathering system COOLING 25 1041 General 25 1043 Cooling tests 25 1045 Cooling test procedures INDUCTION SYSTEM 25 1091 Air induction 25 1093 Induction system icing protection 25 1101 Carburetor air preheater design 25 1103 Induction system ducts and air duct systems 25 1105 Induction system screens 25 1107 Inter coolers and after coolers EXHAUST SYSTEM 25 1121 General 25 1123 Exhaust piping 25 1125 Exhaust heat exchangers 25 1127 Exhaust driven turbo superchargers POWERPLANT CONTROLS AND ACCESSORIES 25 1141 Powerplant controls general 25 1142 Auxiliary power unit controls 25 1
79. MATERIALS a Summary of Method The specimens must be constructed conditioned and tested in the flaming mode in accordance with American Soci ety of Testing and Materials ASTM Standard Test Method ASTM F814 83 b Acceptance Criteria The specific optical smoke density Ds which is obtained by averag ing the reading obtained after 4 minutes with each of the three specimens shall not exceed 200 PART VI TEST METHOD TO DETERMINE THE FLAMMABILITY AND FLAME PROPAGATION CHARACTERISTICS OF THERMAL ACOUSTIC INSULATION MATERIALS Use this test method to evaluate the flammabil ity and flame propagation characteristics of ther mal acoustic insulation when exposed to both a radiant heat source and a flame a Definitions Flame propagation means the furthest dis tance of the propagation of visible flame towards the far end of the test specimen measured from the midpoint of the ignition source flame Measure this distance after initially applying the ignition Source and before all flame on the test specimen is extinguished The measurement is not a deter mination of burn length made after the test Radiant heat source means an electric or air propane panel Thermal acoustic insulation means a mate rial or system of materials used to provide thermal and or acoustic protection Examples include fi berglass or other batting material encapsulated by a film covering and foams Zero point means the point of applicatio
80. Maneuvering load factors lower than those specified in this section may be used if the air plane has design features that make it impossible to exceed these values in flight Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 525 341 Gust and turbulence loads a Discrete Gust Design Criteria The airplane is assumed to be subjected to symmetrical verti cal and lateral gusts in level flight Limit gust loads must be determined in accordance with the provi sions 1 Loads on each part of the structure must be determined by dynamic analysis The analysis must take into account unsteady aerodynamic characteristics and all significant structural de grees of freedom including rigid body motions 2 The shape of the gust must be 8 U 2 Cos H forO lt s lt 2H where s distance penetrated into the gust feet Ugs the design gust velocity in equivalent airspeed specified in paragraph a 4 of this section and H the gust gradient which is the distance feet parallel to the airplane s flight path for the gust to reach its peak velocity 3 A sufficient number of gust gradient dis tances in the range 30 feet to 350 feet must be in vestigated to find the critical response for each load quantity 4 The design gust velocity must be 1 6 Us B 0 sso where Uret the reference gust velocity in equivalent airspeed defined in pa
81. NYIN 0 GE 0 Gz 0 We o c Lon JZZ do 24 7 20 A 0nz m roH4 QN 2 INQIWIXVIN SSV TO 5 Q NL VOVN 90 42 JO 3osnos 90 H A jeonnewN 2 Jo eouejsip juejxe c Y 008 9 jug xo 2 1 000 zz TS eBues epninje A 6 0 YALANVIG 3ALLO3HJ3 NYAN SA LNALNOS ANON SGNO1D WHOSILVYLS INRIWIXVIN SNONNILNOO Part 25 Airworthiness Standards Transport Category 157 ASA Appendix C to Part 25 Federal Aviation Regulations CONTINUOUS MAXIMUM STRATIFORM CLOUDS ATMOSPHERIC ICING CONDITIONS AMBIENT TEMPERATURE VS PRESSURE ALTITUDE SOURCE OF DATA NACA TN NO 2569 32 26 20 AMBIENT TEMPERATURE F 22 0 PRESSURE ALTITUDE 1000 FT FIGURE 2 158 ASA Appendix C to Part 25 Part 2
82. ON ALL EDGES BETWEEN ANGLES AS SHOWN IN VIEW A A TEST STAND FRAME HORIZONTAL SPEC VERTICAL SPEC SUPPORT ANGLE TOP VIEW VIEW A A Typical lt 24 gt 46 m 16 BURNER CONE 48 BURNER ASSEMBLY BURNER SHIELD SUPPORT BRACE 1 x3 x UN STEEL U CHANNEL Y FRONT VIEW SIDE VIEW TEST STAND IS CONSTRUCTED WITH 1 x 1 x 16 STEEL ANGLES ALL JOINTS WELDED SUPPORT ANGLES ARE 1 dx er CUT TO FIT FIGURE 1 Test apparatus for horizontal and vertical mounting 178 ASA Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 CL 1 DIAMETER HOLE FOR CALORIMETER MOUNTING 12 gt x MARINITE BLOCK 1 6 12 3 4 LL Yl 25 gt TOP VIEW NOTE BRACKET IS CLAMPED TO TEST STAND WITH CALORIMETER WATER COOLED ch CENTERED OVER BURNER CONE CALORIMETER STEEL ANGLE x x BURNER SIDE VIEW FIGURE 2 Calorimeter bracket ASA 179 Appendix to Part 25 Federal Aviation Regulations SEVEN THERMOCOUPLES lt 24 gt 12 Yo e e NOTE BRACKET IS CLAMPED TO TEST TOP VIEW STAND WITH THERMOCOUPLES OFF CENTER OF BURNER CONE BY ONE INCH 1 STEEL ANGLE 2 m 8 SIDE VIEW FIGURE 3 Thermocouple rake bracket 180 ASA Part 25 Airworthiness Standards Transport Category PART
83. Regulations cant must comply with 825 1535 except that it need not comply with the following provisions of Appendix K K25 1 4 of this part i K25 1 4 a fuel system pressure and flow requirements ii K25 1 4 a 3 low fuel alerting and iii K25 1 4 c engine oil tank design 2 For ETOPS type design approval of an air plane beyond 180 minutes an applicant must comply with 825 1535 c Airplanes with more than two engines An applicant for ETOPS type design approval must comply with 825 1535 for an airplane manu factured on or after February 17 2015 except that for an airplane configured for a three person flight crew the applicant need not comply with Ap pendix K K25 1 4 a 3 of this part low fuel alert ing Docket No FAA 2002 6717 72 FR 1873 Jan 16 2007 825 5 Incorporations by reference a The materials listed in this section are incor porated by reference in the corresponding sec tions noted These incorporations by reference were approved by the Director of the Federal Reg ister in accordance with 5 U S C 552 a and 1 CFR part 51 These materials are incorporated as they exist on the date of the approval and notice of any change in these materials will be published in the Federal Register The materials are avail able for purchase at the corresponding addresses noted below and all are available for inspection at the National Archives and Records Administration NARA and at FAA Trans
84. The air required by that engine and auxiliary power unit under each operating condition for which certification is requested and 2 The air for proper fuel metering and mixture distribution with the induction system valves in any position ASA Part 25 Airworthiness Standards Transport Category b Each reciprocating engine must have an al ternate air source that prevents the entry of rain ice or any other foreign matter c Air intakes may not open within the cowling unless 1 That part of the cowling is isolated from the engine accessory section by means of a fireproof diaphragm or 2 For reciprocating engines there are means to prevent the emergence of backfire flames d For turbine engine powered airplanes and airplanes incorporating auxiliary power units 1 There must be means to prevent hazardous quantities of fuel leakage or overflow from drains vents or other components of flammable fluid Systems from entering the engine or auxiliary power unit intake system and 2 The airplane must be designed to prevent water or slush on the runway taxiway or other air port operating surfaces from being directed into the engine or auxiliary power unit air inlet ducts in hazardous quantities and the air inlet ducts must be located or protected so as to minimize the in gestion of foreign matter during takeoff landing and taxiing e If the engine induction system contains parts or components that could
85. a The compartment must meet one of the class requirements of 825 857 b Class B through Class E cargo or baggage compartments as defined in 825 857 must have a liner and the liner must be separate from but may be attached to the airplane structure c Ceiling and sidewall liner panels of Class C compartments must meet the test requirements of part of Appendix F of this part or other ap proved equivalent methods d All other materials used in the construction of the cargo or baggage compartment must meet the applicable test criteria prescribed in part of Appendix F of this part or other approved equiva lent methods e No compartment may contain any controls lines equipment or accessories whose damage or failure would affect safe operation unless those items are protected so that 1 They cannot be damaged by the movement of cargo in the compartment and 2 Their breakage or failure will not create a fire hazard f There must be means to prevent cargo or baggage from interfering with the functioning of the fire protective features of the compartment g Sources of heat within the compartment must be shielded and insulated to prevent igniting the cargo or baggage h Flight tests must be conducted to show compliance with the provisions of 825 857 con cerning 1 Compartment accessibility 2 The entries of hazardous quantities of smoke or extinguishing agent into compartments occupied by the cre
86. a complete stop or to a speed of approximately 3 knots for water landings from a point 50 feet above the landing surface must be determined for standard temperatures at each weight altitude and wind within the operational limits established by the applicant for the air plane 1 In non icing conditions and 2 In icing conditions with the landing ice ac cretion defined in appendix C if Vggr for icing con ditions exceeds Vref for non icing conditions by more than 5 knots CAS at the maximum landing weight b In determining the distance in paragraph a of this section 1 The airplane must be in the landing configu ration 2 A stabilized approach with a calibrated air speed of not less than Vref must be maintained down to the 50 foot height ASA Part 25 Airworthiness Standards Transport Category i In non icing conditions may not be less than A 1 23 Vsno B Vme established under 25 149 f and C A speed that provides the maneuvering ca pability specified in 25 143 h ii In icing conditions may not be less than A The speed determined b 2 i of this section B 1 23 Vsno with the landing ice accretion de fined in appendix C if that speed exceeds Vref for non icing conditions by more than 5 knots CAS and C A speed that provides the maneuvering ca pability specified in 825 143 h with the landing ice accretion defined in appendix C 3 Changes in config
87. a haz ardous condition Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15043 Mar 17 1977 25 953 Fuel system independence Each fuel system must meet the requirements of 25 903 b by a Allowing the supply of fuel to each engine through a system independent of each part of the system supplying fuel to any other engine or b Any other acceptable method 97 825 954 825 954 Fuel system lightning protection The fuel system must be designed and ar ranged to prevent the ignition of fuel vapor within the system by a Direct lightning strikes to areas having a high probability of stroke attachment b Swept lightning strokes to areas where swept strokes are highly probable and c Corona and streamering at fuel vent outlets Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 14 32 FR 11629 Aug 11 1967 825 955 Fuel flow a Each fuel system must provide at least 100 percent of the fuel flow required under each in tended operating condition and maneuver Com pliance must be shown as follows 1 Fuel must be delivered to each engine at a pressure within the limits specified in the engine type certificate 2 The quantity of fuel in the tank may not ex ceed the amount established as the unusable fuel supply for that tank under the requirements of 625 959 plus that necessary to show compliance with this section 3 Each main pump must be us
88. airplane must have a means to minimize the danger to the pilots from flying windshield fragments due to bird impact This must be shown for each transparent pane in the cockpit that 1 Appears in the front view of the airplane 2 Is inclined 15 degrees or more to the longi tudinal axis of the airplane and 3 Has any part of the pane located where its fragmentation will constitute a hazard to the pilots ASA 825 777 d The design of windshields and windows in pressurized airplanes must be based on factors peculiar to high altitude operation including the effects of continuous and cyclic pressurization loadings the inherent characteristics of the mate rial used and the effects of temperatures and temperature differentials The windshield and win dow panels must be capable of withstanding the maximum cabin pressure differential loads com bined with critical aerodynamic pressure and tem perature effects after any single failure in the in stallation or associated systems It may be as sumed that after a single failure that is obvious to the flight crew established under 825 1523 the cabin pressure differential is reduced from the maximum in accordance with appropriate operat ing limitations to allow continued safe flight of the airplane with a cabin pressure altitude of not more than 15 000 feet e The windshield panels in front of the pilots must be arranged so that assuming the loss of vi sion through any
89. ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum cylinder barrel temperature recorded during the cooling test Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2323 Jan 16 1978 825 1045 Cooling test procedures a Compliance with 825 1041 must be shown for the takeoff climb en route and landing stages of flight that correspond to the applicable perfor mance requirements The cooling tests must be conducted with the airplane in the configuration and operating under the conditions that are criti 106 Federal Aviation Regulations cal relative to cooling during each stage of flight For the cooling tests a temperature is stabilized when its rate of change is less than two degrees F per minute b Temperatures must be stabilized under the conditions from which entry is made into each stage of flight being investigated unless the entry condition normally is not one during which com ponent and the engine fluid temperatures would stabilize in which case operation through the full entry condition must be conducted before entry into the stage of flight being investigated in order to allow temperatures to reach their natural levels at the time of entry The takeoff cooling test must be preceded by a period during which the power plant component and engine fluid temperatures are stabilized with the engines at ground idle
90. and b In icing conditions with the landing ice ac cretion defined in appendix C and with a climb speed of Vrer determined in accordance with 25 125 b 2 ii Docket No FAA 2005 22840 72 FR 44666 Aug 8 2007 25 121 Climb One engine inoperative a Takeoff landing gear extended the criti cal takeoff configuration existing along the flight path between the points at which the airplane ASA 825 121 reaches Vi or and at which the landing gear is fully retracted and in the configuration used in 825 111 but without ground effect the steady gra dient of climb must be positive for two engine air planes and not less than 0 3 percent for three en gine airplanes or 0 5 percent for four engine air planes at Vi or and with 1 The critical engine inoperative and the re maining engines at the power or thrust available when retraction of the landing gear is begun in ac cordance with 825 111 unless there is a more crit ical power operating condition existing later along the flight path but before the point at which the landing gear is fully retracted and 2 The weight equal to the weight existing when retraction of the landing gear is begun de termined under 825 111 b Takeoff landing gear retracted In the take off configuration existing at the point of the flight path at which the landing gear is fully retracted and in the configuration used in 825 111 but with out ground effect 1 The ste
91. and ii Must provide illumination of not less than 0 03 foot candle measured normal to the direc tion of incident light at the ground and of the erected assist means where an evacuee would normally make first contact with the ground with the airplane in each of the attitudes correspond ing to the collapse of one or more legs of the land ing gear i The energy supply to each emergency light ing unit must provide the required level of illumi nation for at least 10 minutes at the critical ambi ent conditions after emergency landing j If storage batteries are used as the energy supply for the emergency lighting system they may be recharged from the airplane s main elec tric power system Provided That the charging circuit is designed to preclude inadvertent battery discharge into charging circuit faults Components of the emergency lighting sys tem including batteries wiring relays lamps and switches must be capable of normal operation af ter having been subjected to the inertia forces listed in 25 561 b I The emergency lighting system must be de signed so that after any single transverse vertical separation of the fuselage during crash landing 1 Not more than 25 percent of all electrically illuminated emergency lights required by this sec tion are rendered inoperative in addition to the lights that are directly damaged by the separation 2 Each electrically illuminated exit sign re quired unde
92. and c of this section e In the absence of a more rational analysis the nose gear vertical reaction prescribed in para graph d of this section must be calculated ac cording to the following formula po om IRAD N A B Where VN 7 Nose gear vertical reaction Wr Design takeoff weight A Horizontal distance between the c g of the airplane and the nose wheel B Horizontal distance between the c g of the airplane and the line joining the centers of the main wheels 48 Federal Aviation Regulations E Vertical height of the c g of the airplane above the ground in the 1 0 g static condition Coefficient of friction of 0 80 f Dynamic response factor 2 0 is to be used unless a lower factor is substantiated In the absence of other information the dynamic response factor f may be defined by the equation _ di xi Where amp isthe effective critical damping ratio of the rigid body pitching mode about the main landing gear effective ground contact point Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 Amdt 25 97 63 FR 29072 May 27 1998 825 495 Turning In the static position in accordance with figure 7 of Appendix A the airplane is assumed to exe cute a steady turn by nose gear steering or by application of sufficient differential power so that the limit load factors applied at the center of grav
93. and 25 863 2 Equipment that is located in designated fire zones and is used during emergency procedures must be at least fire resistant 3 EWIS components must meet the require ments of 825 1713 b Each vacuum air system line and fitting on the discharge side of the pump that might contain flammable vapors or fluids must meet the require ments of 825 1183 if the line or fitting is in a des ignated fire zone Other vacuum air systems com ponents in designated fire zones must be at least fire resistant c Oxygen equipment and lines must 1 Not be located in any designated fire zone 2 Be protected from heat that may be gener ated in or escape from any designated fire zone and 3 Be installed so that escaping oxygen cannot cause ignition of grease fluid or vapor accumula tions that are present in normal operation or as a result of failure or malfunction of any system Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29784 July 20 1990 Docket Nos FAA 2001 9634 FAA 2001 9633 FAA 2001 9638 FAA 2001 9637 Amdt 25 113 69 FR 12529 March 16 2004 ASA 825 899 MISCELLANEOUS 825 871 Leveling means There must be means for determining when the airplane is in a level position on the ground Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 825 875 Reinforcement near propellers a Each part of the airplane near
94. and arranged to allow 1 Separate control of each propeller and 2 Simultaneous control of all propellers c The controls must allow synchronization of all propellers d The propeller speed and pitch controls must be to the right of and at least one inch below the pilot s throttle controls 825 1153 Propeller feathering controls a There must be a separate propeller feather ing control for each propeller The control must have means to prevent its inadvertent operation b If feathering is accomplished by movement of the propeller pitch or speed control lever there must be means to prevent the inadvertent move ment of this lever to the feathering position during normal operation Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6913 May 5 1967 110 Federal Aviation Regulations 825 1155 Reverse thrust and propeller pitch settings below the flight regime Each control for reverse thrust and for propeller pitch settings below the flight regime must have means to prevent its inadvertent operation The means must have a positive lock or stop at the flight idle position and must require a separate and distinct operation by the crew to displace the control from the flight regime forward thrust re gime for turbojet powered airplanes Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6913 May 5 1967 25 1157 Carburetor air temperature control
95. and main float landing conditions a Symmetrical step bow and stern landing For symmetrical step bow and stern landings the limit water reaction load factors are those computed under 825 527 In addition 1 For symmetrical step landings the resultant water load must be applied at the keel through the center of gravity and must be directed per pendicularly to the keel line 2 For symmetrical bow landings the resultant water load must be applied at the keel one fifth of the longitudinal distance from the bow to the step and must be directed perpendicularly to the keel line and 3 For symmetrical stern landings the result ant water load must be applied at the keel at a point 85 percent of the longitudinal distance from nw 52 Federal Aviation Regulations the step to the stern post and must be directed perpendicularly to the keel line b Unsymmetrical landing for hull and single float seaplanes Unsymmetrical step bow and stern landing conditions must be investigated In addition 1 The loading for each condition consists of an upward component and a side component equal respectively to 0 75 and 0 25 tan times the resultant load in the corresponding symmetri cal landing condition and 2 The point of application and direction of the upward component of the load is the same as that in the symmetrical condition and the point of ap plication of the side component is at the same longitudin
96. and worst case ETOPS Signifi cant System failures and malfunctions that could occur in service The flight test must validate the airplane s flying qualities and performance with the demonstrated failures and malfunctions 224 Federal Aviation Regulations K25 2 2 Early ETOPS method An applicant for ETOPS type design approval using the Early ETOPS method must comply with the following requirements a Assessment of relevant experience with air planes previously certificated under part 25 The applicant must identify specific corrective actions taken on the candidate airplane to prevent rele vant design manufacturing operational and maintenance problems experienced on airplanes previously certificated under part 25 manufac tured by the applicant Specific corrective actions are not required if the nature of a problem is such that the problem would not significantly impact the safety or reliability of the airplane system in volved A relevant problem is a problem with an ETOPS group 1 significant system that has or could result in an IFSD or diversion The applicant must include in this assessment relevant prob lems of supplier provided ETOPS group 1 signifi cant systems and similar or identical equipment used on airplanes built by other manufacturers to the extent such information is reasonably avail able b Propulsion system design 1 The engine used in the applicant s airplane design must be approved as eligible for
97. as defined in part 1 of this chapter f For pressurized fuel tanks a means with fail safe features must be provided to prevent the buildup of an excessive pressure difference be tween the inside and the outside of the tank Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15043 March 17 1977 Amdt 25 69 54 FR 40354 Sept 29 1989 25 965 Fuel tank tests a It must be shown by tests that the fuel tanks as mounted in the airplane can withstand without failure or leakage the more critical of the pressures resulting from the conditions specified ASA 25 965 in paragraphs a 1 and 2 of this section In ad dition it must be shown by either analysis or tests that tank surfaces subjected to more critical pres sures resulting from the condition of paragraphs a 3 and 4 of this section are able to withstand the following pressures 1 An internal pressure of 3 5 psi 2 125 percent of the maximum air pressure developed in the tank from ram effect 3 Fluid pressures developed during maximum limit accelerations and deflections of the air plane with a full tank 4 Fluid pressures developed during the most adverse combination of airplane roll and fuel load b Each metallic tank with large unsupported or unstiffened flat surfaces whose failure or defor mation could cause fuel leakage must be able to withstand the following test or its equivalent with out lea
98. at any air speed up to 1 6 Vs with the flaps in the ap proach position at design landing weight and iii Any load factor up to those specified in 25 345 a for the flaps extended condition 2 Unless there are other means to decelerate the airplane in flight at this speed the landing gear the retracting mechanism and the airplane structure including wheel well doors must be de ASA 825 729 signed to withstand the flight loads occurring with the landing gear in the extended position at any speed up to 0 67 Vc 3 Landing gear doors their operating mecha nism and their supporting structures must be de signed for the yawing maneuvers prescribed for the airplane in addition to the conditions of air speed and load factor prescribed in paragraphs a 1 and 2 of this section b Landing gear lock There must be positive means to keep the landing gear extended in flight and on the ground c Emergency operation There must be an emergency means for extending the landing gear the event of 1 Any reasonably probable failure in the nor mal retraction system or 2 The failure of any single source of hydraulic electric or equivalent energy supply d Operation test The proper functioning of the retracting mechanism must be shown by oper ation tests e Position indicator and warning device If a retractable landing gear is used there must be a landing gear position indicator as well as neces
99. b The operation of the standby power unit and its control must have proven to be safe and reli able Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 6 30 FR 8468 July 2 1965 165 Appendix to Part 25 APPENDIX F TO PART 25 PART TEST CRITERIA AND PROCEDURES FOR SHOWING COMPLIANCE WITH 25 853 OR 525 855 a Material test criteria 1 Interior compartments occupied by crew or passengers i Interior ceiling panels interior wall panels partitions galley structure large cabinet walls structural flooring and materials used in the con struction of stowage compartments other than un derseat stowage compartments and compart ments for stowing small items such as magazines and maps must be self extinguishing when tested vertically in accordance with the applicable por tions of part of this appendix The average burn length may not exceed 6 inches and the average flame time after removal of the flame source may not exceed 15 seconds Drippings from the test specimen may not continue to flame for more than an average of 3 seconds after falling ii Floor covering textiles including draperies and upholstery seat cushions padding decora tive and nondecorative coated fabrics leather trays and galley furnishings electrical conduit air ducting joint and edge covering liners of Class B and E cargo or baggage compartments floor panels of Class B C D or E cargo or baggage c
100. bine powered airplanes 4 Maximum certificated takeoff weights The maximum certificated takeoff weights must be de termined at all altitudes and at ambient tempera tures if applicable at which performance credit is to be applied and may not exceed the weights es tablished in compliance with paragraphs a and b of this section a The conditions of paragraphs 2 b through d must be met at the maximum certifi cated takeoff weight b Without the use of standby power the air plane must meet all of the en route requirements of the applicable airworthiness regulations under which the airplane was originally certificated In addition turbine powered airplanes without the use of standby power must meet the final takeoff climb requirements prescribed in the applicable airworthiness regulations 5 Maximum certificated landing weights a The maximum certificated landing weights one engine inoperative approach and all engine operating landing climb must be determined at all altitudes and at ambient temperatures if applica ble at which performance credit is to be applied and must not exceed that established in compli ance with paragraph b of this section b The flight path with the engines operating at the power or thrust or both appropriate to the airplane configuration and with standby power in use must lie above the flight path without standby power in use at the maximum weight at which all of the applica
101. by the applicant but may not be less than determined under 25 149 e 2 V4 in terms of calibrated airspeed is se lected by the applicant however V4 may not be less than Ver plus the speed gained with critical engine inoperative during the time interval be tween the instant at which the critical engine is failed and the instant at which the pilot recog nizes and reacts to the engine failure as indicated by the pilot s initiation of the first action e g ap plying brakes reducing thrust deploying speed brakes to stop the airplane during accelerate stop tests b in terms of calibrated airspeed may not be less than 1 1 13 for 16 Federal Aviation Regulations i Two engine and three engine turbopropeller and reciprocating engine powered airplanes and ii Turbojet powered airplanes without provi sions for obtaining a significant reduction in the one engine inoperative power on stall speed 2 1 08 for i Turbopropeller and reciprocating engine powered airplanes with more than three engines and ii Turbojet powered airplanes with provisions for obtaining a significant reduction in the one en gine inoperative power on stall speed and 3 1 10 times Vmc established under 825 149 c V2 in terms of calibrated airspeed must be selected by the applicant to provide at least the gradient of climb required by 825 121 b but may not be less than 1
102. defined in terms of the design oper ating pressure DOP as follows 129 825 1435 Proof Ultimate Element xDOP xDOP 1 Tubes and fittings 1 5 3 0 2 Pressure vessels containing gas High pressure e g accumulators 3 0 4 0 Low pressure e g reservoirs 1 5 3 0 3 Hoses 2 0 4 0 4 All other elements 1 5 2 0 2 Withstand without deformation that would prevent it from performing its intended function the design operating pressure in combination with limit structural loads that may be imposed 3 Withstand without rupture the design oper ating pressure multiplied by a factor of 1 5 in com bination with ultimate structural load that can rea sonably occur simultaneously 4 Withstand the fatigue effects of all cyclic pressures including transients and associated externally induced loads taking into account the consequences of element failure and 5 Perform as intended under all environmen tal conditions for which the airplane is certificated b System design Each hydraulic system must 1 Have means located at a flightcrew station to indicate appropriate system parameters if i It performs a function necessary for contin ued safe flight and landing or ii In the event of hydraulic system malfunc tion corrective action by the crew to ensure con tinued safe flight and landing is necessary 2 Have means to ensure that system pres sures including transient pressure
103. does not have a combination cargo and passenger configuration at least one floor level exit must be located in each side near each end of the cabin 4 For an airplane that is required to have more than one passenger emergency exit for each side of the fuselage no passenger emer gency exit shall be more than 60 feet from any ad jacent passenger emergency exit on the same side of the same deck of the fuselage as mea sured parallel to the airplane s longitudinal axis between the nearest exit edges g Type and number required The maximum number of passenger seats permitted depends on the type and number of exits installed in each side of the fuselage Except as further restricted in paragraphs g 1 through g 9 of this section the maximum number of passenger seats permit ted for each exit of a specific type installed in each side of the fuselage is as follows TypeA 110 TypeB 75 TypeC 55 Type 45 Type ll 40 Type IIl 35 TypelV 9 1 For a passenger seating configuration of 1 to 9 seats there must be at least one Type IV or larger overwing exit in each side of the fuselage or if overwing exits are not provided at least one exit in each side that meets the minimum dimen sions of a Type III exit 2 For a passenger seating configuration of more than 9 seats each exit must be a Type 11 or larger exit 3 For a passenger seating configuration of 10 to 19 seats there must be at least one Type or larger exit in e
104. down distance as used in this section means the actual distance between the bottom of the required opening and a usable foot hold extending out from the fuselage that is large enough to be effective without searching by sight or feel c Over sized exits Openings larger than those specified in this section whether or not of rectangular shape may be used if the specified rectangular opening can be inscribed within the opening and the base of the inscribed rectangular opening meets the specified step up and step down heights d Asymmetry Exits of an exit pair need not be diametrically opposite each other nor of the same Size however the number of passenger seats permitted under paragraph g of this section is based on the smaller of the two exits e Uniformity Exits must be distributed as uni formly as practical taking into account passenger seat distribution f Location 1 Each required passenger emergency exit must be accessible to the passengers and lo 77 825 807 cated where it will afford the most effective means of passenger evacuation 2 If only one floor level exit per side is pre Scribed and the airplane does not have a tail cone or ventral emergency exit the floor level ex its must be in the rearward part of the passenger compartment unless another location affords a more effective means of passenger evacuation 3 If more than one floor level exit per side is prescribed and the airplane
105. each vapor vent return line must lead back to the fuel tank used for takeoff and landing 825 977 Fuel tank outlet a There must be a fuel strainer for the fuel tank outlet or for the booster pump This strainer must 1 For reciprocating engine powered air planes have 8 to 16 meshes per inch and 2 For turbine engine powered airplanes pre vent the passage of any object that could restrict fuel flow or damage any fuel system component b Reserved c The clear area of each fuel tank outlet strainer must be at least five times the area of the outlet line d The diameter of each strainer must be at least that of the fuel tank outlet e Each finger strainer must be accessible for inspection and cleaning Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6913 May 5 1967 Amdt 25 36 39 FR 35460 Oct 1 1974 825 979 Pressure fueling system For pressure fueling systems the following apply a Each pressure fueling system fuel manifold connection must have means to prevent the es cape of hazardous quantities of fuel from the sys tem if the fuel entry valve fails b An automatic shutoff means must be pro vided to prevent the quantity of fuel in each tank from exceeding the maximum quantity approved for that tank This means must 1 Allow checking for proper shutoff operation before each fueling of the tank and 2 Provide indication at each fueling station of fa
106. equivalent sitting in the normal upright position 1 A change in downward vertical velocity Av of not less than 35 feet per second with the air plane s longitudinal axis canted downward 30 de grees with respect to the horizontal plane and with the wings level Peak floor deceleration must occur in not more than 0 08 seconds after impact and must reach a minimum of 14g 2 A change in forward longitudinal velocity Av of not less than 44 feet per second with the airplane s longitudinal axis horizontal and yawed 10 degrees either right or left whichever would cause the greatest likelihood of the upper torso restraint system where installed moving off the ASA 25 562 occupant s shoulder and with the wings level Peak floor deceleration must occur in not more than 0 09 seconds after impact and must reach a minimum of 16g Where floor rails or floor fittings are used to attach the seating devices to the test fixture the rails or fittings must be misaligned with respect to the adjacent set of rails or fittings by at least 10 degrees vertically i e out of Parallel with one rolled 10 degrees c The following performance measures must not be exceeded during the dynamic tests con ducted in accordance with paragraph b of this section 1 Where upper torso straps are used for crew members tension loads in individual straps must not exceed 1 750 pounds If dual straps are used for restraining the upper torso th
107. figure 7 according to the criteria of paragraph c 3 iv of this part of this appendix 2 Ensure that the vertical plane of the burner cone is at a distance of 4 0 125 inch 102 3 mm from the outer surface of the horizontal stringers of the test specimen frame and that the burner and test frame are both situated at 30 angle with respect to vertical 3 When ready to begin the test direct the burner away from the test position to the warm up position so that the flame will not impinge on the specimens prematurely Turn on and light the burner and allow it to stabilize for 2 minutes 201 Appendix F to Part 25 4 To begin the test rotate the burner into the test position and simultaneously start the timing device 5 Expose the test specimens to the burner flame for 4 minutes and then turn off the burner Immediately rotate the burner out of the test posi tion 6 Determine where applicable the burn through time or the point at which the heat flux exceeds 2 0 Btu ft sec 2 27 W cm g Report 1 Identify and describe the specimen being tested 2 Report the number of insulation blanket specimens tested 3 Report the burnthrough time if any and the maximum heat flux on the back face of the in sulation blanket test specimen and the time at which the maximum occurred 202 Federal Aviation Regulations h Requirements 1 Each of the two insulation blanket test spec imens must not al
108. flight regime 25 1155 110 Rolling conditions 25 349 40 S Safety equipment ditching 25 1415 ee 128 general requirements 25 1411 ice protection 25 1479 megaphones 25 1421 Shutoff means 25 1789 Side load on engine 25 363 Signs 25 791 Speed control devices 25 373 Stall speed 25 103 Stall warning 25 207 iie 30 Static electricity protection against 25 899 Supercharger controls 25 1159 m Symmetric maneuvering conditions 25 331 36 240 T Takeoff distance and run 25 113 20 flightipatli 26257775 esses 21 pathr 625 111 cete speeds 25 107 Takeoff 25 105 Test criteria showing compliance Appendix F Tiri 625 1615 V Vacuum systems 25 1433 sss 129 Ww Weight limits 625 25 52 53 dota nies 13 ASA Part 25 Index
109. float structure including frames and bulkheads stringers and bottom plating must be designed under this sec tion b Local pressures For the design of the bot tom plating and stringers and their attachments to the supporting structure the following pressure distributions must be applied 1 For an unflared bottom the pressure at the chine is 0 75 times the pressure at the keel and the pressures between the keel and chine vary lin early in accordance with figure 3 of Appendix B The pressure at the keel psi is computed as fol lows 2 P CX where pressure p s i at the keel 0 00213 K hull station weighing factor in accordance with figure 2 of Appendix B seaplane stalling speed Knots at the de sign water takeoff weight with flaps extended in the appropriate takeoff position and angle of dead rise at keel in accordance with figure 1 of Appendix B 2 For a flared bottom the pressure at the be ginning of the flare is the same as that for an un flared bottom and the pressure between the chine and the beginning of the flare varies lin early in accordance with figure 3 of Appendix B The pressure distribution is the same as that pre scribed in paragraph b 1 of this section for an unflared bottom except that the pressure at the chine is computed as follows 2 251 P C x ch 3 tanB where Poh pressure p
110. floor 10 or less 112 15 11 through 19 12 20 20 or more 15 20 narrower width not less than 9 inches may be approved when substantiated by tests found necessary by the Administrator Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 15 32 FR 13265 Sept 20 1967 Amdt 25 38 41 FR 55466 Dec 20 1976 25 817 Maximum number of seats abreast On airplanes having only one passenger aisle no more than three seats abreast may be placed on each side of the aisle in any one row Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 15 32 FR 13265 Sept 20 1967 825 819 Lower deck service compartments including galleys For airplanes with a service compartment lo cated below the main deck which may be occu pied during taxi or flight but not during takeoff or landing the following apply a There must be at least two emergency evacuation routes one at each end of each lower deck service compartment or two having suffi cient separation within each compartment which could be used by each occupant of the lower deck service compartment to rapidly evacuate to the main deck under normal and emergency lighting conditions The routes must provide for the evacu ation of incapacitated persons with assistance The use of the evacuation routes may not be de pendent on any powered device The routes must be designed to minimize the possibility of block age which might result f
111. for its intended use Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 Amdt 25 40 42 FR 15042 March 17 1977 Amdt 25 57 49 FR 6848 Feb 23 1984 Amdt 25 72 55 FR 29784 July 20 1990 Amdt 25 73 55 FR 32861 Aug ASA Part 25 Airworthiness Standards Transport Category 10 1990 Amdt 25 94 63 FR 8848 Feb 23 1998 Amdt 25 95 63 FR 14798 March 26 1998 Amdt 25 100 65 FR 55854 Sept 14 2000 825 904 Automatic takeoff thrust control system ATTCS Each applicant seeking approval for installation of an engine power control system that automati cally resets the power or thrust on the operating engine s when any engine fails during the takeoff must comply with the requirements of Appendix I of this part Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 62 52 FR 43156 Nov 9 1987 825 905 Propellers a Each propeller must have a type certificate b Engine power and propeller shaft rotational speed may not exceed the limits for which the pro peller is certificated c The propeller blade pitch control system must meet the requirements of 8835 21 35 23 35 42 and 35 43 of this chapter d Design precautions must be taken to mini mize the hazards to the airplane in the event a propeller blade fails or is released by a hub failure The hazards which must be considered include damage to structure and vital systems due t
112. for the proposed config uration b Extinguishers must be evenly distributed throughout the cabin These extinguishers are in addition to those required by paragraph 14 of this SFAR unless it can be shown that the cooktop was installed in the immediate vicinity of the origi nal exits 16 Security The requirements of 525 795 not applicable to airplanes approved in accor dance with this SFAR Docket No FAA 2007 28250 SFAR No 109 74 FR 21541 May 8 2009 11 825 1 Subpart A General 825 1 Applicability a This part prescribes airworthiness standards for the issue of type certificates and changes to those certificates for transport category airplanes b Each person who applies under Part 21 for such a certificate or change must show compli ance with the applicable requirements in this part 825 2 Special retroactive requirements The following special retroactive requirements are applicable to an airplane for which the regula tions referenced in the type certificate predate the sections specified below a Irrespective of the date of application each applicant for a supplemental type certificate or an amendment to a type certificate involving an in crease in passenger seating capacity to a total greater than that for which the airplane has been type certificated must show that the airplane con cerned meets the requirements of 1 Sections 25 721 d 25 783 g 25 785 c 25 803 c 2
113. free entry through the passageway e No door may be installed between any pas senger seat that is occupiable for takeoff and landing and any passenger emergency exit such that the door crosses any egress path including aisles crossaisles and passageways f If it is necessary to pass through a doorway separating any crewmember seat except those seats on the flightdeck occupiable for takeoff and landing from any emergency exit the door must have a means to latch it in the open position The latching means must be able to withstand the loads imposed upon it when the door is subjected to the ultimate inertia forces relative to the sur rounding structure listed in 25 561 b Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 1 30 FR 3204 March 9 1965 Amdt 25 15 32 FR 13265 Sept 20 1967 Amdt 25 32 37 FR 3971 Feb 24 1972 Amdt 25 46 43 FR 50597 Oct 30 1978 Amdt 25 72 55 FR 29783 July 20 1990 Amdt 25 76 57 FR 19244 May 4 1992 Amdt 25 76 57 FR 29120 June 30 1992 Amdt 25 88 61 FR 57958 Nov 8 1996 Amdt 25 116 69 FR 62788 Oct 27 2004 Amdt 25 128 74 FR 25645 May 29 2009 ASA 825 819 825 815 Width of aisle The passenger aisle width at any point be tween seats must equal or exceed the values in the following table Minimum passenger aisle Passenger seating Width inches capacity Less than 25 25 and more in from floor from
114. gear in any normal position 2 The airplane trimmed for straight flight at a speed of 1 3 and 3 The power or thrust necessary to maintain level flight at 1 3 g Stall warning must also be provided in each abnormal configuration of the high lift devices that is likely to be used in flight following system fail ures including all configurations covered by Air plane Flight Manual procedures h For flight in icing conditions before the ice protection system has been activated and is per forming its intended function with the ice accre tion defined in appendix C part of this part the stall warning margin in straight and turning flight must be sufficient to allow the pilot to pre vent stalling without encountering any adverse flight characteristics when 1 The speed is reduced at rates not exceed ing one knot per second 2 The pilot performs the recovery maneuver in the same way as for flight in non icing condi tions and 3 The recovery maneuver is started no earlier than i One second after the onset of stall warning if stall warning is provided by the same means as for flight in non icing conditions or ii Three seconds after the onset of stall warn ing if stall warning is provided by a different means than for flight in non icing conditions i In showing compliance with paragraph h of this section if stall warning is provided by a differ ent means in icing conditions than for non
115. general Provision must be made to prevent the hazardous accumulation of water or ice on or in any heater control compo nent control system tubing or safety control e Heater safety controls For each combus tion heater there must be the following safety con trol means 1 Means independent of the components pro vided for the normal continuous control of air tem 91 825 863 perature airflow and fuel flow must be provided for each heater to automatically shut off the igni tion and fuel supply to that heater at a point re mote from that heater when any of the following 5 i The heat exchanger temperature exceeds safe limits ii The ventilating air temperature exceeds safe limits iii The combustion airflow becomes inade quate for safe operation iv The ventilating airflow becomes inadequate for safe operation 2 The means of complying with paragraph e 1 of this section for any individual heater must i Be independent of components serving any other heater whose heat output is essential for safe operation and 1 Keep the heater off until restarted by the crew 3 There must be means to warn the crew when any heater whose heat output is essential for safe operation has been shut off by the auto matic means prescribed in paragraph e 1 of this section f Air intakes Each combustion and ventilating air intake must be located so that no flammable fluids or vapors can enter th
116. ice accretion defined in appendix C part of this part that 1 The airplane is controllable in a pull up ma neuver up to 1 5 g load factor and 24 2 There is no pitch control force reversal dur ing a pushover maneuver down to 0 5 g load fac tor Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2321 Jan 16 1978 Amdt 25 84 60 FR 30749 June 9 1995 Amdt 25 108 67 FR 70826 Nov 26 2002 Amdt 25 121 72 FR 44667 Aug 8 2007 Amdt 25 129 74 FR 38339 Aug 3 2009 25 145 Longitudinal control a It must be possible at any point between the trim speed prescribed in 25 103 b 6 and stall identification as defined in 25 201 d to pitch the nose downward so that the acceleration to this selected trim speed is prompt with 1 The airplane trimmed at the trim speed pre scribed in 25 103 b 6 2 The landing gear extended 3 The wing flaps i retracted and ii ex tended and 4 Power i off and ii at maximum continuous power on the engines b With the landing gear extended no change in trim control or exertion of more than 50 pounds control force representative of the maximum short term force that can be applied readily by one hand may be required for the following ma neuvers ASA Part 25 Airworthiness Standards Transport Category 1 With power off flaps retracted and the air plane trimmed at 1 3 extend the flaps as ra
117. icing conditions compliance with 25 203 must be shown using the accretion defined in appendix C part of this part Compliance with this re quirement must be shown using the demonstra tion prescribed by 825 201 except that the decel eration rates of 825 201 c 2 need not be dem onstrated Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 7 30 FR 13118 Oct 15 1965 Amdt 25 42 43 FR 2322 Jan 16 1978 Amdt 25 108 67 FR 70827 Nov 26 2002 Amdt 25 121 72 FR 44668 Aug 8 2007 Amdt 25 129 74 FR 38339 Aug 3 2009 ASA 825 233 GROUND AND WATER HANDLING CHARACTERISTICS 825 231 Longitudinal stability and control a Landplanes may have no uncontrollable tendency to nose over in any reasonably ex pected operating condition or when rebound oc curs during landing or takeoff In addition 1 Wheel brakes must operate smoothly and may not cause any undue tendency to nose over and 2 If a tail wheel landing gear is used it must be possible during the takeoff ground run on con crete to maintain any attitude up to thrust line level at 75 percent of Vspq b For seaplanes and amphibians the most adverse water conditions safe for takeoff taxiing and landing must be established Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 108 67 FR 70828 Nov 26 2002 525 233 Directional stability and control a There may be no uncontrollable ground
118. in reaching the ground d Except as provided in paragraph a of this Appendix only the airplane s emergency lighting System may provide illumination e All emergency equipment required for the planned operation of the airplane must be in stalled f Each internal door or curtain must be in the takeoff configuration g Each crewmember must be seated in the normally assigned seat for takeoff and must re main in the seat until receiving the signal for com mencement of the demonstration Each crew member must be a person having knowledge of the operation of exits and emergency equipment and if compliance with 121 291 is also being demonstrated each flight attendant must be a member of a regularly scheduled line crew h A representative passenger load of persons in normal health must be used as follows 1 At least 40 percent of the passenger load must be female 2 At least 35 percent of the passenger load must be over 50 years of age 3 At least 15 percent of the passenger load must be female and over 50 years of age 4 Three life size dolls not included as part of the total passenger load must be carried by pas sengers to simulate live infants 2 years old or younger 220 Federal Aviation Regulations 5 Crewmembers mechanics and training personnel who maintain or operate the airplane in the normal course of their duties may not be used as passengers i No passenger may be assigned a specific
119. in the fuel tanks 2 Demonstrating that no temperature at each place inside each fuel tank where fuel ignition is possible will exceed the temperature determined under paragraph a 1 of this section This must be verified under all probable operating failure and malfunction conditions of each component whose operation failure or malfunction could in crease the temperature inside the tank 3 Demonstrating that an ignition source could not result from each single failure from each sin gle failure in combination with each latent failure condition not shown to be extremely remote and from all combinations of failures not shown to be extremely improbable The effects of manufactur ing variability aging wear corrosion and likely damage must be considered b Except as provided in paragraphs b 2 and c of this section no fuel tank Fleet Average Flammability Exposure on an airplane may ex ceed three percent of the Flammability Exposure Evaluation Time FEET as defined in Appendix N of this part or that of a fuel tank within the wing of the airplane model being evaluated whichever is greater If the wing is not a conventional unheated aluminum wing the analysis must be based on an assumed Equivalent Conventional Unheated Alu minum Wing Tank 101 825 991 1 Fleet Average Flammability Exposure is de termined in accordance with Appendix N of this part The assessment must be done in accor dance with the methods
120. inch above the top of the burner The flame must be applied for 15 seconds and then removed A min imum of 10 inches of specimen must be used for timing purposes approximately 11 7 inches must burn before the burning front reaches the timing zone and the average burn rate must be re corded 6 Forty five degree test A minimum of three specimens must be tested and the results aver aged The specimens must be supported at an angle of 45 to a horizontal surface The exposed surface when installed in the aircraft must be face down for the test The specimens must be ex posed to a Bunsen or Tirrill burner with a nominal inch I D tube adjusted to give a flame of 1 inches in height The minimum flame temperature measured by a calibrated thermocouple pyrome ter in the center of the flame must be 1550 F Suit able precautions must be taken to avoid drafts The flame must be applied for 30 seconds with one third contacting the material at the center of the specimen and then removed Flame time 167 Appendix to Part 25 glow time and whether the flame penetrates passes through the specimen must be recorded 7 Sixty degree test A minimum of three spec imens of each wire specification make and size must be tested The specimen of wire or cable in cluding insulation must be placed at an angle of 60 with the horizontal in the cabinet specified in subparagraph 3 of this paragraph with the cabi net door open during t
121. indicate the direction of the control movement relative to the airplane motion In addi tion there must be clearly visible means to indi cate the position of the trim device with respect to the range of adjustment c Trim control systems must be designed to prevent creeping in flight Trim tab controls must be irreversible unless the tab is appropriately bal anced and shown to be free from flutter d If an irreversible tab control system is used the part from the tab to the attachment of the irre versible unit to the airplane structure must consist of a rigid connection Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5675 April 8 1970 825 679 Control system gust locks a There must be a device to prevent damage to the control surfaces including tabs and to the control system from gusts striking the airplane while it is on the ground or water If the device when engaged prevents normal operation of the control surfaces by the pilot it must 1 Automatically disengage when the pilot op erates the primary flight controls in a normal man ner or ASA Part 25 Airworthiness Standards Transport Category 2 Limit the operation of the airplane so that the pilot receives unmistakable warning at the start of takeoff b The device must have means to preclude the possibility of it becoming inadvertently en gaged in flight 25 681 Limit load static tests a Compliance
122. instructions provided under 8833 5 and 35 3 of this chapter and ii The applicable provisions of this subpart 2 The components of the installation must be constructed arranged and installed so as to en sure their continued safe operation between nor mal inspections or overhauls 3 The installation must be accessible for nec essary inspections and maintenance and 4 The major components of the installation must be electrically bonded to the other parts of the airplane c For each powerplant and auxiliary power unit installation it must be established that no sin gle failure or malfunction or probable combination of failures will jeopardize the safe operation of the airplane except that the failure of structural ele ments need not be considered if the probability of such failure is extremely remote d Each auxiliary power unit installation must meet the applicable provisions of this subpart Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 Amdt 25 40 42 FR 15042 March 17 1977 Amdt 25 46 43 FR 50597 Oct 30 1978 Amdt 25 126 73 FR 63345 Oct 24 2008 25 903 Engines a Engine type certificate 1 Each engine must have a type certificate and must meet the applicable requirements of part 34 of this chapter 2 Each turbine engine must comply with one of the following i Sections 33 76 33 77 and 33 78 of this chapter in effect on December
123. interfere with the proper conduct of the test the fuel tank surfaces fuel lines and other fuel sys tem parts subject to cold air must be insulated to simulate insofar as practicable flight in hot weather Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6912 May 5 1967 Amdt 25 57 49 FR 6848 Feb 23 1984 25 963 Fuel tanks general a Each fuel tank must be able to withstand without failure the vibration inertia fluid and structural loads that it may be subjected to in op eration b Flexible fuel tank liners must be approved or must be shown to be suitable for the particular application c Integral fuel tanks must have facilities for in terior inspection and repair d Fuel tanks within the fuselage contour must be able to resist rupture and to retain fuel under the inertia forces prescribed for the emergency landing conditions in 25 561 In addition these tanks must be in a protected position so that ex posure of the tanks to scraping action with the ground is unlikely e Fuel tank access covers must comply with the following criteria in order to avoid loss of haz ardous quantities of fuel 1 All covers located in an area where experi ence or analysis indicates a strike is likely must be shown by analysis or tests to minimize pene tration and deformation by tire fragments low en ergy engine debris or other likely debris 2 All covers must be fire resistant
124. is attached to the heat flux calibration rig during calibration figure 4 Monitor the insulating block for deterioration and replace it when necessary Adjust the mounting as neces sary to ensure that the calorimeter face is parallel to the exit plane of the test burner cone SEE FIGURES 4 AND 5 AT THE END OF PART VII OF THIS APPENDIX iv Thermocouples Provide seven 1 8 inch 3 2 mm ceramic packed metal sheathed type K Chromel alumel grounded junction thermocou ples with a nominal 24 American Wire Gauge AWG size conductor for calibration Attach the thermocouples to a steel angle bracket to form a thermocouple rake for placement in the calibra tion rig during burner calibration figure 5 v Air velocity meter Use a vane type air ve locity meter to calibrate the velocity of air entering the burner An Omega Engineering Model HH30A is satisfactory Use a suitable adapter to attach the measuring device to the inlet side of the burner to prevent air from entering the burner other than through the measuring device which would produce erroneously low readings Use a flexible duct measuring 4 inches wide 102 mm by 20 feet long 6 1 meters to supply fresh air to the burner intake to prevent damage to the air ve locity meter from ingested soot An optional airbox permanently mounted to the burner intake area 199 Appendix to Part 25 can effectively house the air velocity meter and provide a mounting port for the f
125. least 45 percent Reflectance is the ratio of the luminous flux re flected by a body to the luminous flux it receives When the reflectance of the darker color is greater than 15 percent at least a 30 percent dif ference between its reflectance and the reflec tance of the lighter color must be provided 81 825 812 3 In the case of exists other than those in the side of the fuselage such as ventral or tail cone exists the external means of opening including instructions if applicable must be conspicuously marked in red or bright chrome yellow if the back ground color is such that red is inconspicuous When the opening means is located on only one side of the fuselage a conspicuous marking to that effect must be provided on the other side g Each sign required by paragraph d of this section may use the word exit in its legend in place of the term emergency exit Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 15 32 FR 13264 Sept 20 1967 Amdt 25 32 37 FR 3970 Feb 24 1972 Amdt 25 46 43 FR 50597 Oct 30 1978 43 FR 52495 Nov 13 1978 Amdt 25 79 58 FR 45229 Aug 26 1993 Amdt 25 88 61 FR 57958 Nov 8 1996 25 812 Emergency lighting a An emergency lighting system independent of the main lighting system must be installed However the sources of general cabin illumina tion may be common to both the emergency and the main lighting systems if the power su
126. level of safety in the takeoff approach and land ing regimes of flight equivalent to that prescribed in the regulations under which the airplane was originally certificated without standby power For the purposes of this Appendix standby power is power or thrust or both obtained from rocket en gines for a relatively short period and actuated only in cases of emergency The following provi sions apply 1 Takeoff general The takeoff data Scribed in paragraphs 2 and 3 of this Appendix must be determined at all weights and altitudes and at ambient temperatures if applicable at which performance credit is to be applied 2 Takeoff path a The one engine inoperative takeoff path with standby power in use must be determined in accordance with the performance requirements of the applicable airworthiness regulations b The one engine inoperative takeoff path excluding that part where the airplane is on or just above the takeoff surface determined in ac cordance with paragraph a of this section must lie above the one engine inoperative takeoff path without standby power at the maximum takeoff weight at which all of the applicable airworthiness requirements are met For the purpose of this comparison the flight path is considered to ex tend to at least a height of 400 feet above the takeoff surface c The takeoff path with all engines operating but without the use of standby power must reflect a conser
127. lubricating the engine during operation d Provision must be made to prevent sludge or other foreign matter from affecting the safe op eration of the propeller feathering system Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 COOLING 825 1041 General The powerplant and auxiliary power unit cooling provisions must be able to maintain the tempera tures of powerplant components engine fluids and auxiliary power unit components and fluids within the temperature limits established for these components and fluids under ground water and flight operating conditions and after normal en gine or auxiliary power unit shutdown or both Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 825 1043 Cooling tests a General Compliance with 825 1041 must be shown by tests under critical ground water 105 825 1045 and flight operating conditions For these tests the following apply 1 If the tests are conducted under conditions deviating from the maximum ambient atmo spheric temperature the recorded powerplant temperatures must be corrected under para graphs c and d of this section 2 No corrected temperatures determined un der paragraph a 1 of this section may exceed established limits 3 For reciprocating engines the fuel used during the cooling tests must be the minimum grade approved for the e
128. manufacturer as the source of this information if the applicant shows that the item has an exceptionally high degree of com plexity requiring specialized maintenance tech niques test equipment or expertise The recom mended overhaul periods and necessary cross references to the Airworthiness Limitations sec tion of the manual must also be included In addi tion the applicant must include an inspection pro gram that includes the frequency and extent of the inspections necessary to provide for the contin ued airworthiness of the airplane 2 Troubleshooting information describing probable malfunctions how to recognize those malfunctions and the remedial action for those malfunctions 3 Information describing the order and method of removing and replacing products and parts with any necessary precautions to be taken 4 Other general procedural instructions in cluding procedures for system testing during ground running symmetry checks weighing and determining the center of gravity lifting and shor ing and storage limitations c Diagrams of structural access plates and in formation needed to gain access for inspections when access plates are not provided d Details for the application of special inspec tion techniques including radiographic and ultra sonic testing where such processes are specified e Information needed to apply protective treatments to the structure after inspection f data relative to
129. markings or placards c The airborne equipment required for ex tended operations and flightcrew operating proce dures for this equipment d The system time capability for the following 1 The most limiting fire suppression system for Class C cargo or baggage compartments 2 The most limiting ETOPS significant system other than fire suppression systems for Class C cargo or baggage compartments e This statement The type design reliability and performance of this airplane engine combi nation has been evaluated under 14 CFR 25 1535 and found suitable for identify maximum ap proved diversion time extended operations ETOPS when the configuration maintenance and procedures standard contained in identify the CMP document are met The actual maxi mum approved diversion time for this airplane may be less based on its most limiting system time capability This finding does not constitute operational approval to conduct ETOPS K25 2 TWO ENGINE AIRPLANES An applicant for ETOPS type design approval of a two engine airplane must use one of the methods described in section K25 2 1 K25 2 2 or K25 2 3 of this appendix K25 2 1 Service experience method An applicant for ETOPS type design approval using the service experience method must com ply with sections K25 2 1 a and K25 2 1 b of this appendix before conducting the assessments specified in sections K25 2 1 c and K25 2 1 d of this appendix and the flight te
130. must not be less than those re sulting from application of the minimum forces prescribed in 25 397 c Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 72 55 FR 29776 July 20 1990 25 397 Control system loads a General The maximum and minimum pilot forces specified in paragraph c of this section are assumed to act at the appropriate control grips or pads in a manner simulating flight condi tions and to be reacted at the attachment of the control system to the control surface horn b Pilot effort effects In the control surface flight loading condition the air loads on movable surfaces and the corresponding deflections need not exceed those that would result in flight from the application of any pilot force within the ranges specified in paragraph c of this section Two thirds of the maximum values specified for the ai leron and elevator may be used if control surface hinge moments are based on reliable data In ap plying this criterion the effects of servo mecha nisms tabs and automatic pilot systems must be considered c Limit pilot forces and torques The limit pilot forces and torques are as follows Maximum Minimum Control forces or forces or torques torques Aileron Stick 100 Ibs 40 Ibs Wheel 80 D in Ibs 40 D in Ibs Elevator Stick 250 Ibs 100 Ibs Wheel symmetrical 300 Ibs 100 Ibs Wheel unsymmetrical
131. nishing the reacting inertia forces a Maneuvering The following conditions speeds and aileron deflections except as the de flections may be limited by pilot effort must be considered in combination with an airplane load factor of zero and of two thirds of the positive ma neuvering factor used in design In determining the required aileron deflections the torsional flex ibility of the wing must be considered in accor dance with 825 301 b 1 Conditions corresponding to steady rolling velocities must be investigated In addition condi tions corresponding to maximum angular acceler ation must be investigated for airplanes with en gines or other weight concentrations outboard of the fuselage For the angular acceleration condi tions zero rolling velocity may be assumed in the absence of a rational time history investigation of the maneuver 2 At VA a sudden deflection of the aileron to the stop is assumed 3 At Vc the aileron deflection must be that re quired to produce a rate of roll not less than that obtained in paragraph a 2 of this section 4 At Vp the aileron deflection must be that re quired to produce a rate of roll not less than one third of that in paragraph a 2 of this section b Unsymmetrical gusts The airplane is as sumed to be subjected to unsymmetrical vertical gusts in level flight The resulting limit loads must be determined from either the wing maximum air load derived directly f
132. not allow depressurization of the cabin to an unsafe level This safety assessment must include the physiological effects on the oc cupants 4 The open door during flight would not create aerodynamic interference that could preclude safe flight and landing 5 The airplane would meet the structural de sign requirements with the door open This as sessment must include the aeroelastic stability re quirements of 825 629 as well as the strength re quirements of subpart C of this part 6 The unlatching or opening of the door must not preclude safe flight and landing as a result of interaction with other systems or structures Docket No FAA 2003 14193 69 FR 24501 May 3 2004 825 785 Seats berths safety belts and harnesses a A seat or berth for a nonambulant person must be provided for each occupant who has reached his or her second birthday b Each seat berth safety belt harness and adjacent part of the airplane at each station des ignated as occupiable during takeoff and landing must be designed so that a person making proper use of these facilities will not suffer serious injury in an emergency landing as a result of the inertia forces specified in 25 561 and 25 562 c Each seat or berth must be approved d Each occupant of a seat that makes more than an 18 degree angle with the vertical plane containing the airplane centerline must be pro tected from head injury by a safety belt and an en er
133. not be less than that determined under 25 335 b or found necessary during the flight tests conducted under 825 253 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5680 April 8 1970 825 1507 Maneuvering speed The maneuvering speed must be established so that it does not exceed the design maneuver ing speed V4 determined under 25 335 c 825 1511 Flap extended speed The established flap extended speed must be established so that it does not exceed the de sign flap speed Vr chosen under 25 335 e and 25 345 for the corresponding flap positions and engine powers ASA 825 1521 825 1513 Minimum control speed The minimum control speed Vyc determined under 25 149 must be established as an operat ing limitation 25 1515 Landing gear speeds a The established landing gear operating speed or speeds o may not exceed the speed at which it is safe both to extend and to retract the landing gear as determined under 25 729 or by flight characteristics If the extension speed is not the same as the retraction speed the two speeds must be designated as Vi o gxr and re spectively b The established landing gear extended speed may not exceed the speed at which it is safe to fly with the landing gear secured in the fully extended position and that determined un der 825 729 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25
134. of any powerplant function necessary for safety b The ATTCS must be designed to 1 Apply thrust or power on the operating en gine s following any one engine failure during takeoff to achieve the maximum approved takeoff thrust or power without exceeding engine operat ing limits 2 Permit manual decrease or increase in thrust or power up to the maximum takeoff thrust or power approved for the airplane under existing conditions through the use of the power lever For airplanes equipped with limiters that automatically prevent engine operating limits from being ex ceeded under existing ambient conditions other means may be used to increase the thrust or power in the event of an ATTCS failure provided ASA Appendix to Part 25 the means is located on or forward of the power levers is easily identified and operated under all operating conditions by a single action of either pilot with the hand that is normally used to actuate the power levers and meets the requirements of 25 777 b and 3 Provide a means to verify to the flightcrew before takeoff that the ATTCS is in a condition to operate and 4 Provide a means for the flightcrew to deacti vate the automatic function This means must be designed to prevent inadvertent deactivation 125 6 POWERPLANT INSTRUMENTS In addition to the requirements of 25 1305 a A means must be provided to indicate when the ATTCS is in the armed or ready conditio
135. of cloud liq uid water content with drop diameter and altitude is determined from figures 4 and 5 The cloud liq uid water content for intermittent maximum icing conditions of a horizontal extent other than 2 6 nautical miles is determined by the value of cloud liquid water content of figure 4 multiplied by the appropriate factor in figure 6 of this appendix c Takeoff maximum icing The maximum in tensity of atmospheric icing conditions for takeoff takeoff maximum icing is defined by the cloud liquid water content of 0 35 g m3 the mean effec tive diameter of the cloud droplets of 20 microns and the ambient air temperature at ground level of minus 9 degrees Celsius 9 C The takeoff max imum icing conditions extend from ground level to a height of 1 500 feet above the level of the take off surface PART l AIRFRAME ICE ACCRETIONS FOR SHOWING COMPLIANCE WITH SUBPART B a Ice accretions General The most critical ice accretion in terms of airplane performance and handling qualities for each flight phase must be used to show compliance with the applicable airplane performance and handling requirements in icing conditions of subpart B of this part Appli cants must demonstrate that the full range of at mospheric icing conditions specified in part of this appendix have been considered including ASA Appendix C to Part 25 the mean effective drop diameter liquid water content and temperature appropriate to th
136. of ice on the airplane as a result of the drainage Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5680 April 8 1970 825 1457 Cockpit voice recorders a Each cockpit voice recorder required by the operating rules of this chapter must be approved and must be installed so that it will record the fol lowing 1 Voice communications transmitted from or received in the airplane by radio 2 Voice communications of flight crevmem bers on the flight deck 3 Voice communications of flight crewmem bers on the flight deck using the airplane s inter phone system 4 Voice or audio signals identifying navigation or approach aids introduced into a headset or speaker 5 Voice communications of flight crewmem bers using the passenger loudspeaker system if there is such a system and if the fourth channel is available in accordance with the requirements of paragraph c 4 ii of this section 6 If datalink communication equipment is in stalled all datalink communications using an ap proved data message set Datalink messages must be recorded as the output signal from the communications unit that translates the signal into usable data b The recording requirements of paragraph 2 of this section must be met by installing a cockpit mounted area microphone located in the best position for recording voice communications originating at the first and second pilot stations and voice com
137. of the cabin except that this provi sion does not apply to combination cargo passen ger configurations In addition a There must be a passageway leading from the nearest main aisle to each Type A Type B Type C Type 1 or Type Il emergency exit and be tween individual passenger areas Each passage way leading to a Type A or Type B exit must be un obstructed and at least 36 inches wide Passage ways between individual passenger areas and those leading to Type 1 Type or Type C emer gency exits must be unobstructed and at least 20 inches wide Unless there are two or more main aisles each Type A or B exit must be located so that there is passenger flow along the main aisle to that exit from both the forward and aft direc tions If two or more main aisles are provided there must be unobstructed cross aisles at least 20 inches wide between main aisles There must be 1 A cross aisle which leads directly to each passageway between the nearest main aisle and a Type A or B exit and 2 A cross aisle which leads to the immediate vicinity of each passageway between the nearest main aisle and a Type Type or Type III exit ex cept that when two Type III exits are located within three passenger rows of each other a single cross aisle may be used if it leads to the vicinity between the passageways from the nearest main aisle to each exit b Adequate space to allow crewmember s to assist in the evacuation of passen
138. of the door could result in a hazard must have a locking means to prevent the latches from becoming dis engaged The locking means must ensure suffi cient latching to prevent opening of the door even with a single failure of the latching mechanism 5 It must not be possible to position the lock in the locked position if the latch and the latching mechanism are not in the latched position 6 It must not be possible to unlatch the latches with the locks in the locked position Locks must be designed to withstand the limit loads re sulting from i The maximum operator effort when the latches are operated manually ii The powered latch actuators if installed and iii The relative motion between the latch and the structural counterpart 7 Each door for which unlatching would not result in a hazard is not required to have a locking mechanism meeting the requirements of para graphs d 3 through d 6 of this section e Warning caution and advisory indica tions Doors must be provided with the following indications 1 There must be a positive means to indicate at each door operator s station that all required operations to close latch and lock the door s have been completed 2 There must be a positive means clearly vis ible from each operator station for any door that could be a hazard if unlatched to indicate if the door is not fully closed latched and locked ASA Part 25 Airworthiness Standard
139. of the fuel that must be used for the flammability exposure analy sis is a function of the flash point of the fuel se lected by the Monte Carlo for a given flight The flammability envelope for the fuel is defined by the upper flammability limit UFL and lower flamma bility limit LFL as follows A LFL at sea level flash point temperature of the fuel at sea level minus 10 F LFL decreases from sea level value with increasing altitude at a rate of 1 F per 808 feet B UFL at sea level flash point temperature of the fuel at sea level plus 63 5 F UFL decreases from the sea level value with increasing altitude at a rate of 1 F per 512 feet 4 For each flight analyzed a separate random number must be generated for each of the three parameters ground ambient temperature cruise ambient temperature and fuel flash point using the Gaussian distribution defined in Table 1 of this appendix TABLE 1 GAUSSIAN DISTRIBUTION FOR GROUND AMBIENT TEMPERATURE CRUISE AMBIENT TEMPERATURE AND FUEL FLASH POINT Temperature in deg F Parameter Ground ambient Cruise ambient Fuel flash point FP temperature temperature P Mean Temp 59 95 70 120 1 std dev 20 14 8 8 Pos 1 std dev 17 28 8 8 b The Flight Length Distribution defined in Table 2 must be used in the Monte Carlo analysis 234 ASA Part 25 Airworthiness Standards Transport Category TABLE 2 FLIGHT LENGTH DISTRIBUTION Appendix N
140. operating limitations and operating procedures established by the applicant and provided in the Airplane Flight Manual 2 No changes in the load distribution limits of 25 23 the weight limits of 25 25 except where limited by performance requirements of this sub part and the center of gravity limits of 25 27 ASA 825 25 from those for non icing conditions are allowed for flight in icing conditions or with ice accretion Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5671 April 8 1970 Amdt 25 42 43 FR 2320 Jan 16 1978 Amdt 25 72 55 FR 29774 July 20 1990 Amdt 25 121 72 FR 44665 Aug 8 2007 25 23 Load distribution limits a Ranges of weights and centers of gravity within which the airplane may be safely operated must be established If a weight and center of gravity combination is allowable only within cer tain load distribution limits such as spanwise that could be inadvertently exceeded these limits and the corresponding weight and center of grav ity combinations must be established b The load distribution limits may not ex ceed 1 The selected limits 2 The limits at which the structure is proven or 3 The limits at which compliance with each applicable flight requirement of this subpart is shown 25 25 Weight limits a Maximum weights Maximum weights corre sponding to the airplane operating conditions such as ramp ground or wate
141. or as near to be ing in trim as practical g When maneuvering at a constant airspeed or Mach number up to the stick forces and the gradient of the stick force versus maneu vering load factor must lie within satisfactory lim Federal Aviation Regulations its The stick forces must not be so great as to make excessive demands on the pilot s strength when maneuvering the airplane and must not be low that the airplane can easily be over stressed inadvertently Changes of gradient that occur with changes of load factor must not cause undue difficulty in maintaining control of the air plane and local gradients must not be so low as to result in a danger of overcontrolling h The maneuvering capabilities in a constant speed coordinated turn at forward center of grav ity as specified in the following table must be free of stall warning or other characteristics that might interfere with normal maneuvering Configuration Speed Maneuvering bank angle Thrust power setting in a coordinated turn Takeoff V2 30 Asymmetric WAT Limited 1 Takeoff 2V3 XX 40 All engines operating climb 3 En route VETO 40 Asymmetric WAT Limited Landing VREF 40 Symmetric for 3 flight path angle 1A combination of weight altitude and temperature WAT such that the thrust or power setting produces the mini mum climb gradient specified in 25 121 for the flight condition 2 Airspeed approved fo
142. or de ice system is not functioning normally d For turbine engine powered airplanes the ice protection provisions of this section are con sidered to be applicable primarily to the airframe For the powerplant installation certain additional provisions of subpart E of this part may be found applicable e One of the following methods of icing detec tion and activation of the airframe ice protection system must be provided 1 A primary ice detection system that auto matically activates or alerts the flightcrew to acti vate the airframe ice protection system 2 A definition of visual cues for recognition of the first sign of ice accretion on a specified sur face combined with an advisory ice detection sys tem that alerts the flightcrew to activate the air frame ice protection system or 3 Identification of conditions conducive to air frame icing as defined by an appropriate static or total air temperature and visible moisture for use by the flightcrew to activate the airframe ice pro tection system f Unless the applicant shows that the airframe ice protection system need not be operated dur ing specific phases of flight the requirements of ASA Part 25 Airworthiness Standards Transport Category paragraph e of this section are applicable to all phases of flight 9 After the initial activation of the airframe ice protection system 1 The ice protection system must be designed to operate continuousl
143. over filling 825 959 Unusable fuel supply The unusable fuel quantity for each fuel tank and its fuel system components must be estab lished at not less than the quantity at which the first evidence of engine malfunction occurs under the most adverse fuel feed condition for all in tended operations and flight maneuvers involving fuel feeding from that tank Fuel system compo nent failures need not be considered Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5677 April 8 1970 Amdt 25 40 42 FR 15043 March 17 1977 825 961 Fuel system hot weather operation a The fuel system must perform satisfactorily in hot weather operation This must be shown by showing that the fuel system from the tank outlets to each engine is pressurized under all intended operations so as to prevent vapor formation or must be shown by climbing from the altitude of the airport elected by the applicant to the maximum altitude established as an operating limitation un der 825 1527 If a climb test is elected there may be no evidence of vapor lock or other malfunction ing during the climb test conducted under the fol lowing conditions 1 For reciprocating engine powered air planes the engines must operate at maximum continuous power except that takeoff power must be used for the altitudes from 1 000 feet below the critical altitude through the critical altitude The time interval during which takeoff
144. partial pressures 1 At cabin pressure altitudes above 10 000 feet up to and including 18 500 feet a mean tra cheal oxygen partial pressure of 100 mm Hg when breathing 15 liters per minute BTPS and with a tidal volume of 700 cc with a constant time interval between respirations 2 At cabin pressure altitudes above 18 500 feet up to and including 40 000 feet a mean tra cheal oxygen partial pressure of 83 8 mm Hg when breathing 30 liters per minute BTPS and with a tidal volume of 1 100 cc with a constant time interval between respirations d If first aid oxygen equipment is installed the minimum mass flow of oxygen to each user may not be less than four liters per minute STPD However there may be a means to decrease this flow to not less than two liters per minute STPD at any cabin altitude The quantity of oxygen re quired is based upon an average flow rate of three liters per minute per person for whom first aid ox ygen is required e If portable oxygen equipment is installed for use by crewmembers the minimum mass flow of supplemental oxygen is the same as specified in paragraph a or b of this section whichever is applicable 825 1445 Equipment standards for the oxygen distributing system a When oxygen is supplied to both crew and passengers the distribution system must be de signed for either 1 A source of supply for the flight crew on duty and a separate source for the passengers and
145. percent at and above standard tempera tures plus 50 F Between these two temperatures the relative humidity must vary linearly 2 For reciprocating engine powered air planes a relative humidity of 80 percent in a stan dard atmosphere Engine power corrections for vapor pressure must be made in accordance with the following table Specific Density ratio nm pressure e rw y plo i In Hg per Ib dry air 0 0023769 0 0 403 0 00849 0 99508 1 000 354 00773 96672 2 000 311 00703 93895 3 000 272 00638 91178 4 000 238 00578 88514 5 000 207 00523 85910 6 000 1805 00472 83361 7 000 1566 00425 80870 8 000 1356 00382 78434 9 000 1172 00343 76053 10 000 1010 00307 73722 15 000 0463 001710 62868 20 000 01978 000896 53263 25 000 00778 000436 44806 ASA Part 25 Airworthiness Standards Transport Category c The performance must correspond to the propulsive thrust available under the particular ambient atmospheric conditions the particular flight condition and the relative humidity specified in paragraph b of this section The available pro pulsive thrust must correspond to engine power or thrust not exceeding the approved power or thrust less 1 Installation losses and 2 The power or equivalent thrust absorbed by the accessories and services appropriate to the particular ambient atmospheric conditions and th
146. procedures peculiar to the particu lar type or model encountered in connection with routine operations 2 Non normal procedures for malfunction cases and failure conditions involving the use of special systems or the alternative use of regular Systems and 3 Emergency procedures for foreseeable but unusual situations in which immediate and pre cise action by the crew may be expected to sub stantially reduce the risk of catastrophe b Information or procedures not directly re lated to airworthiness or not under the control of the crew must not be included nor must any pro cedure that is accepted as basic airmanship c Information identifying each operating con dition in which the fuel system independence pre Scribed in 825 953 is necessary for safety must be furnished together with instructions for placing 141 825 1587 the fuel system in a configuration used to show compliance with that section d The buffet onset envelopes determined un der 825 251 must be furnished The buffet onset envelopes presented may reflect the center of gravity at which the airplane is normally loaded during cruise if corrections for the effect of differ ent center of gravity locations are furnished e Information must be furnished that indicates that when the fuel quantity indicator reads zero in level flight any fuel remaining in the fuel tank cannot be used safely in flight f Information on the total quantity of usable
147. protection of structure 25 609 static electricity protection against 25 899 93 thermal acoustic insulation materials 25 856 90 ventilation heating 25 831 25 833 Design fuel and oil loads 25 343 Designated fire zones 825 1181 105 106 105 113 237 Early ETOPS method Appendix K 223 225 Electrical appliances motors and transformers 25 1365 Electrical supply for emergency services 25 1362 124 Electrical systems and equipment circuit protection 25 1357 distribution system 25 1355 general airworthiness 25 13571 124 229 123 122 installation 25 1353 x 123 lights anticollision 25 1401 s126 instrument 25 1381 124 landing 25 1383 position 25 1385 25 1397 riding 25 1399 is tests 25 1363 wing icing detection lights 25 1 Electrical Wiring Interconnection System EWIS Appendix H iie 215 Electrical Wiring Interconnection Systems EWIS 25 1701 1733 Electronic equipment 25 1437 Empty weight 25 29 En route flight paths 25 Engine accessory section diaphragm 25 1192 controls 925 1143 ignition systems 25 1165 torque 25 361 Equipment airworthiness standards 25 1301 25 1461
148. reaction load factor that is the water reaction divided by seaplane weight 2 C4 empirical seaplane operations factor equal to 0 012 except that this factor may not be less than that necessary to obtain the minimum value of step load factor of 2 33 3 Vso seaplane stalling speed in knots with flaps extended in the appropriate landing position and with no slipstream effect 4 B angle of dead rise at the longitudinal station at which the load factor is being deter mined in accordance with figure 1 of Appendix B 5 W seaplane design landing weight in pounds 6 empirical hull station weighing factor in accordance with figure 2 of Appendix B 7 rx ratio of distance measured parallel to hull reference axis from the center of gravity of the seaplane to the hull longitudinal station at which the load factor is being computed to the ra dius of gyration in pitch of the seaplane the hull reference axis being a straight line in the plane of symmetry tangential to the keel at the main step c For a twin float seaplane because of the effect of flexibility of the attachment of the floats to the seaplane the factor K4 may be reduced at the bow and stern to 0 8 of the value shown in fig ure 2 of Appendix B This reduction applies only to the design of the carrythrough and seaplane structure Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 825 529 Hull
149. reduce the capability of the airplane or the ability of the flightcrew to respond to an adverse operating condition must be designed and in stalled so the system is not adversely affected when the equipment providing these functions is exposed to equipment HIRF test level 1 or 2 as described in appendix L to this part c Each electrical and electronic system that performs a function whose failure would reduce the capability of the airplane or the ability of the flightcrew to respond to an adverse operating condition must be designed and installed so the System is not adversely affected when the equip ment providing the function is exposed to equip ment HIRF test level 3 as described in appendix L to this part d Before December 1 2012 an electrical or electronic system that performs a function whose failure would prevent the continued safe flight and landing of an airplane may be designed and in stalled without meeting the provisions of para graph a provided 1 The system has previously been shown to comply with special conditions for HIRF pre scribed under 821 16 issued before December 1 2007 2 The HIRF immunity characteristics of the System have not changed since compliance with the special conditions was demonstrated and 3 The data used to demonstrate compliance with the special conditions is provided Docket No FAA 2006 23657 72 FR 44025 Aug 6 2007 118 Federal Aviation Regulations INSTRUM
150. remain on if 1 A single malfunction including a wire bun dle or junction box fire cannot result in loss of both the part turned off and the part turned on and 2 The parts turned on are electrically and me chanically isolated from the parts turned off Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 41 42 FR 36970 July 18 1977 Amdt 25 72 55 FR 29785 July 20 1990 25 1353 Electrical equipment and installations a Electrical equipment and controls must be installed so that operation of any one unit or sys tem of units will not adversely affect the simulta neous operation of any other electrical unit or sys tem essential to safe operation Any electrical in terference likely to be present in the airplane must not result in hazardous effects on the airplane or its systems b Storage batteries must be designed and in stalled as follows 1 Safe cell temperatures and pressures must be maintained during any probable charging or discharging condition No uncontrolled increase in cell temperature may result when the battery is recharged after previous complete discharge i At maximum regulated voltage or power ii During a flight of maximum duration and iii Under the most adverse cooling condition likely to occur in service 2 Compliance with paragraph b 1 of this section must be shown by test unless experience with similar batteries and installations has shown that m
151. requirement for bundles 1 The identification must be placed along the wire cable or wire bundle at appropriate intervals and in areas of the airplane where it is readily vis ible to maintenance repair or alteration person nel 2 If an EWIS component cannot be marked physically then other means of identification must be provided c The identifying markings required by para graphs a and b of this section must remain leg ible throughout the expected service life of the EWIS component d The means used for identifying each EWIS component as required by this section must not have an adverse effect on the performance of that component throughout its expected service life e Identification for EWIS modifications to the type design must be consistent with the identifica tion scheme of the original type design 825 1713 Fire protection EWIS a All EWIS components must meet the appli cable fire and smoke protection requirements of 25 831 c of this part b EWIS components that are located in des ignated fire zones and are used during emer gency procedures must be fire resistant c Insulation on electrical wire and electrical cable and materials used to provide additional protection for the wire and cable installed in any area of the airplane must be self extinguishing when tested in accordance with the applicable portions of Appendix F part of 14 CFR part 25 825 1715 Electrical bonding and prot
152. schedule imposed by center of gravity structural or other considerations of an airworthiness nature and to the ability of each engine to operate at all times from a single tank or source which is automati cally replenished if fuel is also stored in other tanks 4 The degree and duration of concentrated mental and physical effort involved in normal op eration and in diagnosing and coping with mal functions and emergencies 5 The extent of required monitoring of the fuel hydraulic pressurization electrical elec tronic deicing and other systems while en route 6 The actions requiring a crewmember to be unavailable at his assigned duty station includ ing observation of systems emergency operation of any control and emergencies in any compart ment 7 The degree of automation provided in the aircraft systems to afford after failures or mal functions automatic crossover or isolation of diffi culties to minimize the need for flight crew action to guard against loss of hydraulic or electric power to flight controls or to other essential systems 8 The communications and navigation work load ASA Appendix D to Part 25 9 The possibility of increased workload asso ciated with any emergency that may lead to other emergencies 10 Incapacitation of a flight crewmember whenever the applicable operating rule requires a minimum flight crew of at least two pilots c Kind of operation authorized The determi
153. should be designed to withstand 1 33 times the specified loads if these items are sub ject to severe wear and tear through frequent re moval e g quick change interior items d Seats and items of mass and their support ing structure must not deform under any loads up to those specified in paragraph b 3 of this sec tion in any manner that would impede subsequent rapid evacuation of occupants Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 Amdt 25 64 53 FR 17646 May 17 1988 Amdt 25 91 62 FR 40706 July 29 1997 25 562 Emergency landing dynamic conditions a The seat and restraint system in the air plane must be designed as prescribed in this sec tion to protect each occupant during an emer gency landing condition when 1 Proper use is made of seats safety belts and shoulder harnesses provided for in the de sign and 2 The occupant is exposed to loads resulting from the conditions prescribed in this section b Each seat type design approved for crew or passenger occupancy during takeoff and landing must successfully complete dynamic tests or be demonstrated by rational analysis based on dy namic tests of a similar type seat in accordance with each of the following emergency landing con ditions The tests must be conducted with an oc cupant simulated by a 170 pound anthropomor phic test dummy as defined by 49 CFR Part 572 Subpart B or its
154. size and so located as to per mit rapid evacuation by the crew One exit shall be provided on each side of the airplane or alterna tively a top hatch shall be provided Each exit must encompass an unobstructed rectangular opening of at least 19 by 20 inches unless satisfactory exit utility can be demonstrated by a typical crewmem ber Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29781 July 20 1990 Amdt 25 88 61 FR 57957 Nov 8 1996 Amdt 25 88 62 FR 1817 Jan 13 1997 Amdt 25 94 63 FR 8848 Feb 23 1998 Amdt 25 94 63 FR 12862 March 16 1998 Amdt 25 114 69 FR 24501 May 3 2004 825 809 Emergency exit arrangement a Each emergency exit including each flight crew emergency exit must be a moveable door or hatch in the external walls of the fuselage allow ing an unobstructed opening to the outside In ad dition each emergency exit must have means to permit viewing of the conditions outside the exit when the exit is closed The viewing means may be on or adjacent to the exit provided no obstruc tions exist between the exit and the viewing means Means must also be provided to permit viewing of the likely areas of evacuee ground con tact The likely areas of evacuee ground contact must be viewable during all lighting conditions with the landing gear extended as well as in all conditions of landing gear collapse b Each emergency exit must be openable from the ins
155. speed at maximum takeoff weight exceeds that in non icing conditions by more than the greater of 3 knots CAS or 3 percent of VSR or ii The degradation of the gradient of climb de termined in accordance with 825 121 b is greater than one half of the applicable actual to net take off flight path gradient reduction defined in 825 115 b b No takeoff made to determine the data re quired by this section may require exceptional pi loting skill or alertness c The takeoff data must be based on 1 In the case of land planes and amphibians i Smooth dry and wet hard surfaced run ways and ii At the option of the applicant grooved or po rous friction course wet hard surfaced runways d The takeoff data must include within the es tablished operational limits of the airplane the fol lowing operational correction factors 1 Not more than 50 percent of nominal wind components along the takeoff path opposite to the direction of takeoff and not less than 150 per cent of nominal wind components along the take off path in the direction of takeoff 2 Effective runway gradients Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 92 63 FR 8318 Feb 18 1998 Amdt 25 121 72 FR 44665 Aug 8 2007 825 107 Takeoff speeds a V4 must be established in relation to Ver as follows 1 Vgr is the calibrated airspeed at which the critical engine is assumed to fail must be se lected
156. system considering the probable pilot corrective action on the flight controls 1 At speeds between Vmc and Vp the loads resulting from power failure because of fuel flow interruption are considered to be limit loads 2 At speeds between Vyc and Vc the loads resulting from the disconnection of the engine compressor from the turbine or from loss of the turbine blades are considered to be ultimate loads 3 The time history of the thrust decay and drag build up occurring as a result of the pre scribed engine failures must be substantiated by test or other data applicable to the particular en gine propeller combination 4 The timing and magnitude of the probable pilot corrective action must be conservatively esti mated considering the characteristics of the par ticular engine propeller airplane combination b Pilot corrective action may be assumed to be initiated at the time maximum yawing velocity is reached but not earlier than two seconds after the engine failure The magnitude of the correc tive action may be based on the control forces specified in 25 397 b except that lower forces may be assumed where it is shown by analysis or test that these forces can control the yaw and roll resulting from the prescribed engine failure conditions ASA Part 25 Airworthiness Standards Transport Category 825 371 Gyroscopic loads The structure supporting any engine or auxil iary power unit must be designed for the loa
157. system must meet the following requirements 1 Each lift control switch outside the lift ex cept emergency stop buttons must be designed to prevent the activation of the life if the lift door or the hatch required by paragraph 9 3 of this sec tion or both are open 2 An emergency stop button that when acti vated will immediately stop the lift must be in stalled within the lift and at each entrance to the lift 3 There must be a hatch capable of being used for evacuating persons from the lift that is openable from inside and outside the lift without tools with the lift in any position Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 53 45 FR 41593 June 19 1980 45 FR 43154 June 26 1980 Amdt 25 110 68 FR 36883 June 19 2003 825 820 Lavatory doors All lavatory doors must be designed to pre clude anyone from becoming trapped inside the lavatory If a locking mechanism is installed it must be capable of being unlocked from the out side without the aid of special tools Docket No FAA 2003 14193 69 FR 24501 May 3 2004 VENTILATION AND HEATING 825 831 Ventilation a Under normal operating conditions and in the event of any probable failure conditions of any system which would adversely affect the ventilat ing air the ventilation system must be designed to provide a sufficient amount of uncontaminated air to enable the crewmembers to perform their du ties without undue
158. than 20 inches wide by 60 inches high with corner radii not greater than seven inches in the pressure shell and incorpo rating an approved assist means in accordance with 25 810 a 25 additional passenger seats iii For a tail cone exit incorporating an opening in the pressure shell which is at least equivalent to Type IIl emergency exit with respect to dimen sions step up and step down distance and with the top of the opening not less than 56 inches from the passenger compartment floor 15 addi tional passenger seats h Other exits The following exits also must meet the applicable emergency exit requirements of 8825 809 through 25 812 and must be readily accessible 1 Each emergency exit in the passenger com partment in excess of the minimum number of re quired emergency exits 2 Any other floor level door or exit that is ac cessible from the passenger compartment and is as large or larger than Type exit but less than 46 inches wide 3 Any other ventral or tail cone passenger exit i Ditching emergency exits for passengers Whether or not ditching certification is requested ditching emergency exits must be provided in ac cordance with the following requirements unless the emergency exits required by paragraph g of this section already meet them 1 For airplanes that have a passenger seating configuration of nine or fewer seats excluding pilot seats one exit above the waterline in each sid
159. that require a stick force of more than 50 pounds with i Wing flaps center of gravity position and weight as specified in paragraph b 1 of this section ii Power required for level flight at a speed equal to Vyo 1 3 Vsg1 2 and iii The airplane trimmed for level flight with the power required in paragraph b 2 ii of this sec tion 3 With the landing gear extended the stick force curve must have a stable slope at all speeds within a range which is the greater of 15 percent of the trim speed plus the resulting free return speed range or 50 knots plus the resulting free return speed range above and below the trim speed except that the speed range need not in clude speeds less than 1 3 nor speeds greater than V g nor speeds that require a stick force of more than 50 pounds with i Wing flap center of gravity position and weight as specified in paragraph b 1 of this section ii 75 percent of maximum continuous power for reciprocating engines or for turbine engines the maximum cruising power selected by the ap plicant as an operating limitation except that the ASA Part 25 Airworthiness Standards Transport Category power need not exceed that required for level flight at Vie and iii The aircraft trimmed for level flight with the power required in paragraph b 3 ii of this sec tion c Approach The stick force curve must have a stable slope at speeds between Vsw and 1
160. the engine accessory section 4 Any auxiliary power unit compartment 5 Any fuel burning heater and other combus tion equipment installation described in 825 859 6 The compressor and accessory sections of turbine engines and 7 Combustor turbine and tailpipe sections of turbine engine installations that contain lines or components carrying flammable fluids or gases b Each designated fire zone must meet the requirements of 25 867 and 25 1185 through 25 1203 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6913 May 5 1967 Amdt 25 23 35 FR 5677 April 8 1970 Amdt 25 72 55 FR 29785 July 20 1990 825 1182 Nacelle areas behind firewalls and engine pod attaching structures containing flammable fluid lines a Each nacelle area immediately behind the firewall and each portion of any engine pod at taching structure containing flammable fluid lines must meet each requirement of 25 1103 b 25 1165 d and e 25 1183 25 1185 c 25 1187 25 1189 and 25 1195 through 25 1203 including those concerning designated fire zones However engine pod attaching structures need not contain fire detection or extinguishing means b For each area covered by paragraph a of this section that contains a retractable landing gear compliance with that paragraph need only be shown with the landing gear retracted Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25
161. the load or stress Yone value of the load or stress in one g level flight average number of exceedances of the indicated value of the load or stress in unit time symbol denoting summation over all mission segments No parameters determined by dynamic analysis as defined in paragraph c 2 of this appendix P4 P2 by 62 parameters defining the probability distributions of root mean square gust velocity to be read from Figures 1 and 2 of this appendix The limit gust loads must be read form the fre quency of exceedance curves at a frequency of exceedance of 2 x 10 exceedances per hour Both positive and negative load directions must be considered in determining the limit loads 212 Federal Aviation Regulations 3 For each of the load and stress quantities selected the frequency of exceedance must be determined as a function of load level by means of the equation 0 7 ivy ies a 1 4 If a stability augmentation system is utilized to reduce the gust loads consideration must be given to the fraction of flight time that the system may be inoperative The flight profiles of para graph c 1 of this appendix must include flight with the system inoperative for this fraction of the flight time When a stability augmentation system is included in the analysis the effect of system nonlinearities on loads at the limit load level must be conservatively acc
162. through 9 25 803 d and e 25 807 a c and d 25 809 f and h 25 811 25 812 25 813 a b and c 25 815 25 817 25 853 a and b 25 855 25 993 f and 25 1359 c in effect on October 24 1967 and 2 Sections 25 803 b and 25 803 c 1 in ef fect on April 23 1969 b Irrespective of the date of application each applicant for a supplemental type certificate or an amendment to a type certificate for an airplane manufactured after October 16 1987 must show that the airplane meets the requirements of 825 807 7 in effect on July 24 1989 c Compliance with subsequent revisions to the sections specified in paragraph a or b of this section may be elected or may be required in accordance with 821 101 a of this chapter Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29773 July 20 1990 Amdt No 25 99 65 FR 36266 June 7 2000 Amdt 25 99 66 FR 56989 Nov 14 2001 825 3 Special provisions for ETOPS type design approvals a Applicability This section applies to an ap plicant for ETOPS type design approval of an air plane 1 That has an existing type certificate on Feb ruary 15 2007 or 2 For which an application for an original type certificate was submitted before February 15 2007 b Airplanes with two engines 1 For ETOPS type design approval of an air plane up to and including 180 minutes an appli 12 Federal Aviation
163. to Vc and in 825 345 2 The limit gust conditions specified in 825 341 at the specified speeds up to Vc and in 825 345 3 The limit rolling conditions specified in 825 349 and the limit unsymmetrical conditions specified in 25 367 and 25 427 a through at speeds up to Vc 4 The limit yaw maneuvering conditions specified in 25 351 a at the specified speeds up to Vc 5 For pressurized cabins the following condi tions i The normal operating differential pressure combined with the expected external aerody namic pressures applied simultaneously with the flight loading conditions specified in paragraphs b 1 through 4 of this section if they have a significant effect ii The maximum value of normal operating dif ferential pressure including the expected external aerodynamic pressures during 1 g level flight multiplied by a factor of 1 15 omitting other loads ASA Part 25 Airworthiness Standards Transport Category 6 For landing gear and directly affected air frame structure the limit ground loading condi tions specified in 25 473 25 491 and 25 493 If significant changes in structural stiffness or ge ometry or both follow from a structural failure or partial failure the effect on damage tolerance must be further investigated c Fatigue safe life evaluation Compliance with the damage tolerance requirements of para graph b of this section is not required if the appli
164. to a prede termined value at the maximum rate of flow deliv ered by the pressure source The combined ca pacity of the relief valves must be large enough so that the failure of any one valve would not cause an appreciable rise in the pressure differential The pressure differential is positive when the in ternal pressure is greater than the external 2 Two reverse pressure differential relief valves or their equivalents to automatically pre vent a negative pressure differential that would damage the structure One valve is enough how ever if it is of a design that reasonably precludes its malfunctioning 3 A means by which the pressure differential can be rapidly equalized 4 An automatic or manual regulator for con trolling the intake or exhaust airflow or both for maintaining the required internal pressures and airflow rates 5 Instruments at the pilot or flight engineer station to show the pressure differential the cabin pressure altitude and the rate of change of the cabin pressure altitude 6 Warning indication at the pilot or flight engi neer station to indicate when the safe or preset pressure differential and cabin pressure altitude limits are exceeded Appropriate warning mark ings on the cabin pressure differential indicator meet the warning requirement for pressure differ ential limits and an aural or visual signal in addi tion to cabin altitude indicating means meets the warning requirement for c
165. to part 25 need not be demonstrated if it can be shown by test or a com bination of test and analysis that the maximum time for evacuation of all occupants does not ex ceed 45 seconds under the conditions specified in Appendix J to part 25 13 Fire Detection For airplanes with a type certificated passenger capacity of 20 or more there must be means that meet the requirements of 25 858 a through d to signal the flightcrew in the event of a fire in any isolated room not oc cupiable for taxi takeoff and landing which can be closed off from the rest of the cabin by a door The indication must identify the compartment where the fire is located This does not apply to 10 Federal Aviation Regulations lavatories which continue to be governed by 825 854 14 Cooktops Each cooktop must be de signed and installed to minimize any potential threat to the airplane passengers and crew Compliance with this requirement must be found in accordance with the following criteria a Means such as conspicuous burner on in dicators physical barriers or handholds must be installed to minimize the potential for inadvertent personnel contact with hot surfaces of both the cooktop and cookware Conditions of turbulence must be considered b Sufficient design means must be included to restrain cookware while in place on the cook top as well as representative contents soup sauces etc from the effects of flight loads and
166. to the direction of the incident light where an evacuee using the established es cape route would normally make first contact with the ground 2 At each non overwing emergency exit not required by 25 810 a to have descent assist means the illumination must be not less than 0 03 foot candle measured normal to the direction of the incident light on the ground surface with the landing gear extended where an evacuee is likely to make first contact with the ground outside the cabin h The means required in 25 810 a 1 and d to assist the occupants in descending to the ground must be illuminated so that the erected assist means is visible from the airplane 1 If the assist means is illuminated by exterior emergency lighting it must provide illumination of not less than 0 03 foot candle measured normal to the direction of the incident light at the ground end of the erected assist means where an evac uee using the established escape route would normally make first contact with the ground with the airplane in each of the attitudes correspond ing to the collapse of one or more legs of the land ing gear 2 If the emergency lighting subsystem illumi nating the assist means serves no other assist ASA 825 812 means is independent of the airplane s main emergency lighting system and is automatically activated when the assist means is erected the lighting provisions i May not be adversely affected by stowage
167. turbulence Restraints must be provided to pre clude hazardous movement of cookware and con tents These restraints must accommodate any cookware that is identified for use with the cook top Restraints must be designed to be easily uti lized and effective in service The cookware re straint system should also be designed so that it will not be easily disabled thus rendering it unus able Placarding must be installed which prohibits the use of cookware that cannot be accommo dated by the restraint system c Placarding must be installed which prohibits the use of cooktops i e power on any burner during taxi takeoff and landing d Means must be provided to address the possibility of a fire occurring on or in the immedi ate vicinity of the cooktop Two acceptable means of complying with this requirement are as follows 1 Placarding must be installed that prohibits any burner from being powered when the cooktop is unattended Note This would prohibit a single person from cooking on the cooktop and intermit tently serving food to passengers while any burner is powered A fire detector must be in stalled in the vicinity of the cooktop which pro vides an audible warning in the passenger cabin and a fire extinguisher of appropriate size and ex tinguishing agent must be installed in the immedi ate vicinity of the cooktop Access to the extin guisher may not be blocked by a fire on or around the cooktop 2 An automati
168. valve failures automatic fuel 222 Federal Aviation Regulations management system failures and normal electri cal power generation failures 1 If the engine has been certified for limited operation with negative engine fuel pump inlet pressures the following requirements apply i Airplane demonstration testing must cover worst case cruise and diversion conditions involv ing A Fuel grade and temperature B Thrust or power variations C Turbulence and negative G D Fuel system components degraded within their approved maintenance limits ii Unusable fuel quantity in the suction feed configuration must be determined in accordance with 825 959 2 For two engine airplanes to be certificated for ETOPS beyond 180 minutes one fuel boost pump in each main tank and at least one cross feed valve or other means for transferring fuel must be powered by an independent electrical power source other than the three power sources required to comply with section K25 1 3 b of this appendix This requirement does not apply if the normal fuel boost pressure crossfeed valve actu ation or fuel transfer capability is not provided by electrical power 3 An alert must be displayed to the flightcrew when the quantity of fuel available to the engines falls below the level required to fly to the destina tion The alert must be given when there is enough fuel remaining to safely complete a diver sion This alert must accoun
169. vertical load factor is 1 2 at the design landing weight and 1 0 at the design ramp weight A drag reaction equal to the vertical reaction multiplied by a coefficient of friction of 0 8 must be com bined with the vertical reaction and applied at the ground contact point of each wheel with brakes The following two attitudes in accordance with figure 6 of Appendix A must be considered 1 The level attitude with the wheels contacting the ground and the loads distributed between the main and nose gear Zero pitching acceleration is assumed 2 The level attitude with only the main gear contacting the ground and with the pitching mo ment resisted by angular acceleration c A drag reaction lower than that prescribed in this section may be used if it is substantiated that an effective drag force of 0 8 times the verti cal reaction cannot be attained under any likely loading condition d An airplane equipped with a nose gear must be designed to withstand the loads arising from the dynamic pitching motion of the airplane due to sudden application of maximum braking force The airplane is considered to be at design takeoff weight with the nose and main gears in contact with the ground and with a steady state vertical load factor of 1 0 The steady state nose gear reaction must be combined with the maxi mum incremental nose gear vertical reaction caused by the sudden application of maximum braking force as described in paragraphs b
170. which certification for oper ation above 40 000 feet is requested must be ap proved 25 1443 Minimum mass flow of supplemental oxygen a If continuous flow equipment is installed for use by flight crewmembers the minimum mass flow of supplemental oxygen required for each crewmember may not be less than the flow re quired to maintain during inspiration a mean tra cheal oxygen partial pressure of 149 mm Hg when breathing 15 liters per minute BTPS and with a maximum tidal volume of 700 cc with a constant time interval between respirations b If demand equipment is installed for use by flight crewmembers the minimum mass flow of supplemental oxygen required for each crew member may not be less than the flow required to maintain during inspiration a mean tracheal oxy gen partial pressure of 122 mm Hg up to and in cluding a cabin pressure altitude of 35 000 feet and 95 percent oxygen between cabin pressure altitudes of 35 000 and 40 000 feet when breath 131 825 1445 ing 20 liters per minute BTPS In addition there must be means to allow the crew to use undiluted oxygen at their discretion c For passengers and cabin attendants the minimum mass flow of supplemental oxygen re quired for each person at various cabin pressure altitudes may not be less than the flow required to maintain during inspiration and while using the ox ygen equipment including masks provided the following mean tracheal oxygen
171. whichever is greater d Each pressure altimeter must be approved and must be calibrated to indicate pressure alti tude in a standard atmosphere with a minimum practicable calibration error when the correspond ing static pressures are applied e Each system must be designed and in stalled so that the error in indicated pressure alti 119 825 1326 tude at sea level with a standard atmosphere excluding instrument calibration error does not result in an error of more than 30 feet per 100 knots speed for the appropriate configuration in the speed range between 1 23 Vspo with flaps ex tended and 1 7 with flaps retracted How ever the error need not be less than 30 feet f If an altimeter system is fitted with a device that provides corrections to the altimeter indica tion the device must be designed and installed in such manner that it can be bypassed when it mal functions unless an alternate altimeter system is provided Each correction device must be fitted with a means for indicating the occurrence of rea sonably probable malfunctions including power failure to the flight crew The indicating means must be effective for any cockpit lighting condition likely to occur g Except as provided in paragraph h of this section if the static pressure system incorporates both a primary and an alternate static pressure Source the means for selecting one or the other Source must be designed so that 1 W
172. 066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 32 37 FR 3972 Feb 24 1972 Amdt 25 38 41 FR 55468 Dec 20 1976 Amdt 25 72 55 FR 29786 July 20 1990 25 1561 Safety equipment a Each safety equipment control to be oper ated by the crew in emergency such as controls for automatic liferaft releases must be plainly marked as to its method of operation b Each location such as a locker or compart ment that carries any fire extinguishing signal ing or other life saving equipment must be marked accordingly c Stowage provisions for required emergency equipment must be conspicuously marked to identify the contents and facilitate the easy re moval of the equipment d Each liferaft must have obviously marked operating instructions e Approved survival equipment must be marked for identification and method of operation Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50598 Oct 30 1978 825 1563 Airspeed placard A placard showing the maximum airspeeds for flap extension for the takeoff approach and land ing positions must be installed in clear view of each pilot AIRPLANE FLIGHT MANUAL 825 1581 General a Furnishing information An Airplane Flight Manual must be furnished with each airplane and it must contain the following 1 Information required by 8825 1583 through 25 1587 2 Other information that is necessary for safe operation because of
173. 1 STALLS Stall demonstration Stall characteristics Stall warning 25 201 25 203 25 207 GROUND AND WATER HANDLING CHARACTERISTICS Longitudinal stability and control Directional stability and control Taxiing condition Wind velocities Spray characteristics control and stability on water 25 231 25 233 25 235 25 237 25 239 MISCELLANEOUS FLIGHT REQUIREMENTS 25 251 Vibration and buffeting 25 253 High speed characteristics 25 255 Out of trim characteristics Subpart C Structure GENERAL 25 301 25 303 25 305 25 307 Loads Factor of safety Strength and deformation Proof of structure FLIGHT LOADS 25 321 General FLIGHT MANEUVER AND GUST CONDITIONS 25 331 Symmetric maneuvering conditions 25 333 Flight maneuvering envelope 25 335 Design airspeeds 25 337 Limit maneuvering load factors 25 341 Gust and turbulence loads 25 343 Design fuel and oil loads 25 345 High lift devices 25 349 Rolling conditions 25 351 Yaw maneuver conditions SUPPLEMENTARY CONDITIONS Engine torque Side load on engine and auxiliary power unit mounts Pressurized compartment loads Unsymmetrical loads due to engine failure Gyroscopic loads Speed control devices 25 361 25 363 25 365 25 367 25 371 25 373 Part 25 CONTROL SURFACE AND SYSTEM LOADS 25 391 Control surface loads general 25 393 Loads parallel to hinge line 25 395 Control system 25 397 Control system load
174. 11 32 FR 6913 May 5 1967 25 1183 Flammable fluid carrying components a Except as provided in paragraph b of this section each line fitting and other component carrying flammable fluid in any area subject to en gine fire conditions and each component which conveys or contains flammable fluid in a desig nated fire zone must be fire resistant except that flammable fluid tanks and supports in a desig nated fire zone must be fireproof or be enclosed by a fireproof shield unless damage by fire to any non fireproof part will not cause leakage or spill age of flammable fluid Components must be 111 825 1185 shielded or located to safeguard against the igni tion of leaking flammable fluid An integral oil sump of less than 25 quart capacity on a recipro cating engine need not be fireproof nor be en closed by a fireproof shield b Paragraph a of this section does not apply to 1 Lines fittings and components which are already approved as part of a type certificated en gine and 2 Vent and drain lines and their fittings whose failure will not result in or add to a fire hazard c All components including ducts within a designated fire zone must be fireproof if when ex posed to or damaged by fire they could 1 Result in fire spreading to other regions of the airplane or 2 Cause unintentional operation of or inability to operate essential services or equipment Docket No 5066 29
175. 143 Engine controls 25 1145 Ignition switches 25 1147 Mixture controls 25 1149 Propeller speed and pitch controls 25 1153 Propeller feathering controls 25 1155 Reverse thrust and propeller pitch settings below the flight regime 25 1157 Carburetor air temperature controls 25 1159 Supercharger controls 25 1161 Fuel jettisoning system controls 25 1163 Powerplant accessories 25 1165 Engine ignition systems 25 1167 Accessory gearboxes POWERPLANT FIRE PROTECTION 25 1181 Designated fire zones regions included 25 1182 Nacelle areas behind firewalls and engine pod attaching structures containing flammable fluid lines 25 1183 25 1185 25 1187 25 1189 25 1191 25 1192 25 1193 25 1195 25 1197 25 1199 25 1201 25 1203 25 1207 25 1301 25 1303 25 1305 25 1307 25 1309 25 1310 25 1316 25 1317 25 1321 25 1322 25 1323 25 1325 25 1326 25 1327 25 1329 25 1331 25 1333 25 1335 25 1337 Federal Aviation Regulations Flammable fluid carrying components Flammable fluids Drainage and ventilation of fire zones Shutoff means Firewalls Engine accessory section diaphragm Cowling and nacelle skin Fire extinguishing systems Fire extinguishing agents Extinguishing agent containers Fire extinguishing system materials Fire detector system Compliance Subpart F Equipment GENERAL Function and installation Flight and navigation instruments Powerplant instruments Miscellaneous
176. 15 must accommodate enough rafts for the maximum number of occupants for which certification for ditching is requested 2 Liferafts must be stowed near exits through which the rafts can be launched during an un planned ditching 3 Rafts automatically or remotely released outside the airplane must be attached to the air plane by means of the static line prescribed in 825 1415 127 825 1415 4 The stowage provisions for each portable liferaft must allow rapid detachment and removal of the raft for use at other than the intended exits e Long range signaling device The stowage provisions for the long range signaling device re quired by 25 1415 must be near an exit available during an unplanned ditching f Life preserver stowage provisions The stow age provisions for life preservers described in 25 1415 must accommodate one life preserver for each occupant for which certification for ditch ing is requested Each life preserver must be within easy reach of each seated occupant g Life line stowage provisions f certification for ditching under 825 801 is requested there must be provisions to store life lines These provi sions must 1 Allow one life line to be attached to each side of the fuselage and 2 Be arranged to allow the life lines to be used to enable the occupants to stay on the wing after ditching Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 32 37 FR 3972 Feb 24
177. 152 by 305 mm by inch 19 mm thick calcium silicate insulating board which is attached to a steel angle bracket for placement in the test stand during burner calibration as shown in Fig ure 4 1 Because crumbling of the insulating board with service can result in misalignment of the cal ASA Appendix F to Part 25 orimeter the calorimeter must be monitored and the mounting shimmed as necessary to ensure that the calorimeter face is flush with the exposed plane of the insulating board in a plane parallel to the exit of the test burner cone 4 Thermocouples The seven thermocouples to be used for testing must be 146 to Y inch metal sheathed ceramic packed type K grounded thermocouples with a nominal 22 to 30 American wire gage AWG size conductor The Seven thermocouples must be attached to a steel angle bracket to form a thermocouple rake for placement in the test stand during burner calibra tion as shown in Figure 5 5 Apparatus Arrangement The test burner must be mounted on a suitable stand to position the exit of the burner cone a distance of 4 inches 102 3 mm from one side of the speci men mounting stand The burner stand should have the capability of allowing the burner to be swung away from the specimen mounting stand during warmup periods 6 Data Recording A recording potentiometer or other suitable calibrated instrument with an ap propriate range must be used to measure and record the output
178. 1972 Amdt 25 46 43 FR 50598 Oct 30 1978 Amdt 25 53 45 FR 41593 June 19 1980 Amdt 25 70 54 FR 43925 Oct 27 1989 Amdt 25 79 58 FR 45229 Aug 26 1993 Amdt 25 116 69 FR 62789 Oct 27 2004 25 1415 Ditching equipment a Ditching equipment used in airplanes to be certificated for ditching under 25 801 and re quired by the operating rules of this chapter must meet the requirements of this section b Each liferaft and each life preserver must be approved In addition 1 Unless excess rafts of enough capacity are provided the buoyancy and seating capacity be yond the rated capacity of the rafts must accom modate all occupants of the airplane in the event of a loss of one raft of the largest rated capacity and 2 Each raft must have a trailing line and must have a static line designed to hold the raft near the airplane but to release it if the airplane be comes totally submerged c Approved survival equipment must be at tached to each liferaft d There must be an approved survival type emergency locator transmitter for use in one life raft e For airplanes not certificated for ditching un der 25 801 and not having approved life preserv ers there must be an approved flotation means for each occupant This means must be within easy reach of each seated occupant and must be readily removable from the airplane 128 Federal Aviation Regulations Docket No 5066 29 FR 18291
179. 2 Vp plus speed increment attained in ac cordance with 25 111 c 2 before reaching a height of 35 feet above the takeoff surface and 3 A speed that provides the maneuvering ca pability specified in 25 143 h d is the calibrated airspeed at and above which the airplane can safely lift off the ground and continue the takeoff speeds must be se lected by the applicant throughout the range of thrust to weight ratios to be certificated These speeds may be established from free air data if these data are verified by ground takeoff tests e Vp in terms of calibrated airspeed must be selected in accordance with the conditions of paragraphs e 1 through 4 of this section 1 Vp may not be less than i V4 ii 105 percent of Vic iii The speed determined in accordance with 25 111 c 2 that allows reaching V2 before reaching a height of 35 feet above the takeoff sur face or iv A speed that if the airplane is rotated at its maximum practicable rate will result in a Vi of not less than 110 percent of in the all en gines operating condition and not less than 105 percent of Vy determined at the thrust to weight ratio corresponding to the one engine inoperative condition 2 For any given set of conditions such as weight configuration and temperature a single value of Vg obtained in accordance with this paragraph must be used to show compliance with both the one eng
180. 25 503 Pivoting a The airplane is assumed to pivot about one side of the main gear with the brakes on that side locked The limit vertical load factor must be 1 0 and the coefficient of friction 0 8 b The airplane is assumed to be in static equi librium with the loads being applied at the ground contact points in accordance with figure 8 of Ap pendix A 825 507 Reversed braking a The airplane must be in a three point static ground attitude Horizontal reactions parallel to the ground and directed forward must be applied at the ground contact point of each wheel with brakes The limit loads must be equal to 0 55 times the vertical load at each wheel or to the load developed by 1 2 times the nominal maximum static brake torque whichever is less ASA 825 509 b For airplanes with nose wheels the pitching moment must be balanced by rotational inertia c For airplanes with tail wheels the resultant of the ground reactions must pass through the center of gravity of the airplane 825 509 Towing loads a The towing loads specified in paragraph d of this section must be considered separately These loads must be applied at the towing fittings and must act parallel to the ground In addition 1 A vertical load factor equal to 1 0 must be considered acting at the center of gravity 2 The shock struts and tires must be in their static positions and 3 With Wr as the design ramp weight the towing load
181. 25 965 and 2 The test fluid must be oil at 250 F instead of the fluid specified in 25 965 c Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 36 39 FR 35461 Oct 1 1974 25 1017 Oil lines and fittings a Each oil line must meet the requirements of 25 993 and each oil line and fitting in any desig nated fire zone must meet the requirements of 25 1183 ASA Part 25 Airworthiness Standards Transport Category b Breather lines must be arranged so that 1 Condensed water vapor that might freeze and obstruct the line cannot accumulate at any point 2 The breather discharge does not constitute a fire hazard if foaming occurs or causes emitted oil to strike the pilot s windshield and 3 The breather does not discharge into the engine air induction system 825 1019 Oil strainer or filter a Each turbine engine installation must incor porate an oil strainer or filter through which all of the engine oil flows and which meets the following requirements 1 Each oil strainer or filter that has a bypass must be constructed and installed so that oil will flow at the normal rate through the rest of the sys tem with the strainer or filter completely blocked 2 The oil strainer or filter must have the ca pacity with respect to operating limitations estab lished for the engine to ensure that engine oil system functioning is not impaired when the oil is contaminated to a deg
182. 291 Dec 24 1964 as amended by Amdt 25 57 49 FR 6848 Feb 23 1984 ASA 825 1043 825 1023 Oil radiators a Each oil radiator must be able to withstand without failure any vibration inertia and oil pres sure load to which it would be subjected in opera tion b Each oil radiator air duct must be located so that in case of fire flames coming from normal openings of the engine nacelle cannot impinge di rectly upon the radiator 825 1025 Oil valves a Each oil shutoff must meet the require ments of 825 1189 b The closing of oil shutoff means may not prevent propeller feathering c Each oil valve must have positive stops or suitable index provisions in the on and off posi tions and must be supported so that no loads re sulting from its operation or from accelerated flight conditions are transmitted to the lines at tached to the valve 825 1027 Propeller feathering system a If the propeller feathering system depends on engine oil there must be means to trap an amount of oil in the tank if the supply becomes depleted due to failure of any part of the lubricat ing system other than the tank itself b The amount of trapped oil must be enough to accomplish the feathering operation and must be available only to the feathering pump c The ability of the system to accomplish feathering with the trapped oil must be shown This may be done on the ground using an auxil iary source of oil for
183. 3VL NOISSG Lv MOL LHOISM ONIGNV1 N9IS3G 1V MZ 0 8V39 ISON 40 N9IS3Q Q NIVW JO N9IS3G SI ISON 55 Q 2 C TH3HM JHL 3ONV IV8 OL ANVSSSOSN SONOS VILYANI ASA 150 Part 25 Airworthiness Standards Transport Category Appendix A to Part 25 CENTER OF ROTATION Vy and are static ground reactions For tail wheel type the airplane is the three point attitude Pivoting is assumed to take place about one main landing gear unit FIGURE 8 Pivoting nose or tail wheel type ASA 151 Appendix B to Part 25 Federal Aviation Regulations APPENDIX B TO PART 25 Forebody __ _ Afterbody t Unflared Bottom Flared Bottom FIGURE 1 Pictorial definition of angles dimensions and directions on a seaplane 152 ASA Part 25 Airworthiness Standards Transport Category Appendix B to Part 25 Forebody Length L gt Afterbody Length La K Vertical Loads L k fio Forebody Length L gt lt Afterbody Length L gt Bottom Pressures FIGURE 2 Hull station weighing factor ASA 153 Appendix B to Part 25 Federal Aviation Regulations UNFLARED SYMMETRICAL FLARED UNSYMMETRICAL Local Pressure Distributed Pressure FIGURE 3 Transverse pressure distributions
184. 4 1964 as amended by Amdt 25 23 35 FR 5677 April 8 1970 Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 57 49 FR 6849 Feb 23 1984 825 1145 Ignition switches a Ignition switches must control each engine ignition circuit on each engine b There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control c Each group of ignition switches except igni tion switches for turbine engines for which contin uous ignition is not required and each master ig nition control must have a means to prevent its in advertent operation Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15044 March 17 1977 25 1147 Mixture controls a If there are mixture controls each engine must have a separate control The controls must be grouped and arranged to allow 1 Separate control of each engine and 2 Simultaneous control of all engines b Each intermediate position of the mixture controls that corresponds to a normal operating setting must be identifiable by feel and sight c The mixture controls must be accessible to both pilots However if there is a separate flight engineer station with a control panel the controls need be accessible only to the flight engineer 25 1149 Propeller speed and pitch controls a There must be a separate propeller speed and pitch control for each propeller b The controls must be grouped
185. 5 Airworthiness Standards Transport Category SATIN TVOLLOVN LN3 LX3 IVLNOZIHOH 002 00c 00L 09 OV 02 0 0169879 9 d8nol4 8222 ON VOVN ejeq Jo soinos eouejsiq JejuozuoH pno 2 SA 40j9e4 1u9juo2 JejeM pinbr1 SNOILIGNOD SONIDO OI3aHdSOlNLV sano19 WHOSILWYLS NNWIXVW SNONNILNOO pinbr E ssejuoisueuilq 159 ASA Federal Aviation Regulations Appendix C to Part 25 y 3dnoli4 SNOMOIN HSLSWVIG 1 08448 NYIN 0G gy oe Gz Sg 0 FF J07 LY 77 y Zeb Ay BSA LL ETT RS y x Z7 dd n B XE At 7 ame Ly 17 gt Kador mi SLINI 30 LN31X3 wA 2 37 550 3 SINI ALON 19 8 lt YX Sg INCIATXVIW LNALLINYSLNI Wel SSV TO 2 9981 NL VOVN E 30 304 05 2 oz 20 9 2 5 uajxe Z ul 000 22 000 eBueJ m 3ALLOd3233 NYAN SA LN3 LNOO
186. 6971 July 18 1977 825 1439 Protective breathing equipment a If there is a class A B or E cargo compart ment protective breathing equipment must be in stalled for the use of appropriate crewmembers In addition protective breathing equipment must be installed in each isolated separate compart ment in the airplane including upper and lower lobe galleys in which crewmember occupancy is permitted during flight for the maximum number of crewmembers expected to be in the area dur ing any operation b For protective breathing equipment re quired by paragraph a of this section or by any operating rule of this chapter the following apply 1 The equipment must be designed to protect the flight crew from smoke carbon dioxide and other harmful gases while on flight deck duty and while combating fires in cargo compartments 2 The equipment must include i Masks covering the eyes nose and mouth or ii Masks covering the nose and mouth plus accessory equipment to cover the eyes 3 The equipment while in use must allow the flight crew to use the radio equipment and to com municate with each other while at their assigned duty stations 4 The part of the equipment protecting the eyes may not cause any appreciable adverse ef fect on vision and must allow corrective glasses to be worn ASA 825 1443 5 The equipment must supply protective oxy gen of 15 minutes duration per crewmember at a pressur
187. 7 5 with 1 Wing flaps in the approach position 2 Landing gear retracted 3 Maximum landing weight and 4 The airplane trimmed at 1 3 Vsg4 with enough power to maintain level flight at this speed d Landing The stick force curve must have a stable slope and the stick force may not exceed 80 pounds at speeds between Vs and 1 7 with 1 Wing flaps in the landing position 2 Landing gear extended 3 Maximum landing weight 4 Power or thrust off on the engines and 5 The airplane trimmed at 1 3 Vsgo with power or thrust off Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 7 30 FR 13117 Oct 15 1965 Amdt 25 108 67 FR 70827 Nov 26 2002 825 177 Static lateral directional stability a Reserved b Reserved c In straight steady sideslips the aileron and rudder control movements and forces must be substantially proportional to the angle of sideslip in a stable sense and the factor of proportionality must lie between limits found necessary for safe operation throughout the range of sideslip angles appropriate to the operation of the airplane At greater angles up to the angle at which full rudder is used or a rudder force of 180 pounds is ob tained the rudder pedal forces may not reverse and increased rudder deflection must be needed for increased angles of sideslip Compliance with this paragraph must be demonstrated for all land ing ge
188. 7 for airplanes to be approved for op eration above 45 000 feet omitting other loads e Any structure component or part inside or outside a pressurized compartment the failure of which could interfere with continued safe flight and landing must be designed to withstand the effects of a sudden release of pressure through an open ing in any compartment at any operating altitude resulting from each of the following conditions 1 The penetration of the compartment by a portion of an engine following an engine disinte gration 2 Any opening in any pressurized compart ment up to the size in square feet however small compartments may be combined with an adjacent pressurized compartment and both con sidered as a single compartment for openings that cannot reasonably be expected to be con fined to the small compartment The size Ho must be computed by the following formula PAs where Maximum opening in square feet need not exceed 20 square feet As 024 6240 Ag Maximum cross sectional area of the pressurized shell normal to the longitudinal axis in square feet and 3 The maximum opening caused by airplane or equipment failures not shown to be extremely improbable f In complying with paragraph e of this sec tion the fail safe features of the design may be considered in determining the probability of failure or penetration and probable size of openings pro vided that possib
189. 90 Amdt 25 108 67 FR 70827 Nov 26 2002 STALLS 825 201 Stall demonstration a Stalls must be shown in straight flight and in 30 degree banked turns with 1 Power off and 2 The power necessary to maintain level flight at 1 5 where Vsg4 corresponds to the refer ence stall speed at maximum landing weight with flaps in the approach position and the landing gear retracted b In each condition required by paragraph a of this section it must be possible to meet the ap plicable requirements of 525 203 with 1 Flaps landing gear and deceleration de vices in any likely combination of positions ap proved for operation 2 Representative weights within the range for which certification is requested 3 The most adverse center of gravity for re covery and 4 The airplane trimmed for straight flight at the speed prescribed in 25 103 b 6 c The following procedures must be used to show compliance with 825 203 1 Starting at a speed sufficiently above the stalling speed to ensure that a steady rate of speed reduction can be established apply the longitudinal control so that the speed reduction does not exceed one knot per second until the air plane is stalled 2 In addition for turning flight stalls apply the longitudinal control to achieve airspeed decelera tion rates up to 3 knots per second 3 As soon as the airplane is stalled recover by normal recovery techniques 29
190. ATONV TIVIS LON 1d3OX3 SNLLOVINOO AYNLONYLS ANY V3O 3 19NV 133HM 3SON 1101 M uvop er 6 33914 133HM 9 Bulpue jone1 z 34 914 3OHOH VILYSNI dO L AHVSSH3OSN 30H04 VILY NI YVINONY TASHM IYL ASA 148 Part 25 Airworthiness Standards Transport Category Appendix A to Part 25 THE AIRPLANE INERTIA LOADS REQUIRED TO BALANCE THE EXTERNAL FORCES SINGLE WHEEL LOAD FROM 2 WHEEL LEVEL B LANDING CONDITION NOSE OR TAIL WHEEL TYPE FIGURE 4 0 80V 0 60V V 6 V M ONE HALF THE MAXIMUM VERTICAL GROUND REACTION CONTAINED AT EACH MAIN GEAR IN THE LEVEL LANDING CONDITIONS NOSE GEAR GROUND REACTION 0 NOSE OR TAIL WHEEL TYPE AIRPLANE IN LEVEL ALTITUDE FIGURE 5 Lateral drift landing ASA 149 Federal Aviation Regulations Appendix A to Part 25 punog 34914 TASHM JSON 3dAL 133HM maso 5 Mago Is 290 6 NMOHS SV SNOLLOV3H T33HM JHL A8 ALJANO 33V ALIAVHO JO H3LN3O AV SHOLOVH VILYANI 3NV IdHl V JHL pexe18g 9 AYNOIA T33HM JSON 3dAL T33HM 11 gals Hova A Az NA adis Yad AZ adis Yad LHOI3M 23430 3XVL NOIS3G LV MO LHOISM ONIGNV1 N9IS3G Lv MZ LHOISM 34O 3
191. Amdt 25 32 37 FR 3969 Feb 24 1972 Amdt 25 46 43 FR 50596 Oct 30 1978 25 791 Passenger information signs and placards a If smoking is to be prohibited there must be at least one placard so stating that is legible to each person seated in the cabin If smoking is to be allowed and if the crew compartment is sepa rated from the passenger compartment there must be at least one sign notifying when smoking is prohibited Signs which notify when smoking is prohibited must be operable by a member of the flightcrew and when illuminated must be legible under all probable conditions of cabin illumination to each person seated in the cabin b Signs that notify when seat belts should be fastened and that are installed to comply with the operating rules of this chapter must be operable by a member of the flightcrew and when illumi nated must be legible under all probable condi tions of cabin illumination to each person seated in the cabin c A placard must be located on or adjacent to the door of each receptacle used for the disposal of flammable waste materials to indicate that use of the receptacle for disposal of cigarettes etc is prohibited d Lavatories must have No Smoking or No Smoking in Lavatory placards conspicuously lo cated on or adjacent to each side of the entry door ASA 25 795 e Symbols that clearly express the intent of the sign or placard may be used in lieu of letters
192. CATIONS AND MAINTENANCE ACCESS a Reliability indications must be provided to identify failures of the FRM that would otherwise be latent and whose identification is necessary to ensure the fuel tank with an FRM meets the fleet average flammability exposure requirements listed in paragraph M25 1 of this appendix includ ing when the FRM is inoperative b Sufficient accessibility to FRM reliability in dications must be provided for maintenance per sonnel or the flightcrew c The access doors and panels to the fuel tanks with FRMs including any tanks that com municate with a tank via a vent system and to any other confined spaces or enclosed areas that could contain hazardous atmosphere under nor mal conditions or failure conditions must be per manently stenciled marked or placarded to warn maintenance personnel of the possible presence of a potentially hazardous atmosphere ASA Part 25 Airworthiness Standards Transport Category M25 4 AIRWORTHINESS LIMITATIONS AND PROCEDURES a If FRM is used to comply with paragraph M25 1 of this appendix Airworthiness Limitations must be identified for all maintenance or inspec tion tasks required to identify failures of compo nents within the FRM that are needed to meet paragraph M25 1 of this appendix b Maintenance procedures must be devel oped to identify any hazards to be considered during maintenance of the FRM These proce dures must be included in the instruct
193. Dec 24 1964 as amended by Amdt 25 29 36 FR 18722 Sept 21 1971 Amdt 25 50 45 FR 38348 June 9 1980 Amdt 25 72 55 FR 29785 July 20 1990 as Amdt 25 82 59 FR 32057 June 21 1994 825 1419 Ice protection If the applicant seeks certification for flight in ic ing conditions the airplane must be able to safely operate in the continuous maximum and intermit tent maximum icing conditions of appendix C To establish this a An analysis must be performed to establish that the ice protection for the various components of the airplane is adequate taking into account the various airplane operational configurations and b To verify the ice protection analysis to check for icing anomalies and to demonstrate that the ice protection system and its components are effective the airplane or its components must be flight tested in the various operational configu rations in measured natural atmospheric icing conditions and as found necessary by one or more of the following means 1 Laboratory dry air or simulated icing tests or a combination of both of the components or models of the components 2 Flight dry air tests of the ice protection sys tem as a whole or of its individual components 3 Flight tests of the airplane or its compo nents in measured simulated icing conditions c Caution information such as an amber cau tion light or equivalent must be provided to alert the flightcrew when the anti ice
194. ENTS INSTALLATION 825 1321 Arrangement and visibility a Each flight navigation and powerplant in strument for use by any pilot must be plainly visi ble to him from his station with the minimum prac ticable deviation from his normal position and line of vision when he is looking forward along the flight path b The flight instruments required by 825 1303 must be grouped on the instrument panel and centered as nearly as practicable about the verti cal plane of the pilot s forward vision In addition 1 The instrument that most effectively indi cates attitude must be on the panel in the top cen ter position 2 The instrument that most effectively indi cates airspeed must be adjacent to and directly to the left of the instrument in the top center position 3 The instrument that most effectively indi cates altitude must be adjacent to and directly to the right of the instrument in the top center posi tion and 4 The instrument that most effectively indi cates direction of flight must be adjacent to and directly below the instrument in the top center po sition c Required powerplant instruments must be closely grouped on the instrument panel In addition 1 The location of identical powerplant instru ments for the engines must prevent confusion as to which engine each instrument relates and 2 Powerplant instruments vital to the safe op eration of the airplane must be plainly visible to the appropr
195. EVALUATION Damage tolerance and fatigue evaluation of structure LIGHTNING PROTECTION Lightning protection Subpart D Design and Construction 25 601 25 603 25 605 25 607 25 609 25 611 25 613 25 619 25 621 25 623 25 625 25 629 25 631 25 651 25 655 25 657 25 671 25 672 25 675 25 677 25 679 25 681 25 683 25 685 25 689 25 693 25 697 25 699 25 701 25 703 25 721 25 723 25 725 25 727 25 729 25 731 25 733 25 735 25 737 GENERAL General Materials Fabrication methods Fasteners Protection of structure Accessibility provisions Material strength properties and material design values Special factors Casting factors Bearing factors Fitting factors Aeroelastic stability requirements Bird strike damage CONTROL SURFACES Proof of strength Installation Hinges CONTROL SYSTEMS General Stability augmentation and automatic and power operated systems Stops Trim systems Control system gust locks Limit load static tests Operation tests Control system details Cable systems Joints Lift and drag devices controls Lift and drag device indicator Flap and slat interconnection Takeoff warning system LANDING GEAR General Shock absorption tests Reserved Reserved Retracting mechanism Wheels Tires Brakes and braking systems Skis ASA Part 25 Airworthiness Standards Transport Category 25 751 25 753 25 755
196. FLOATS AND HULLS Main float buoyancy Main float design Hulls PERSONNEL AND CARGO ACCOMMODATIONS 25 771 25 772 25 773 25 775 25 777 25 779 25 781 25 783 25 785 25 787 25 789 25 791 25 793 25 795 25 801 25 803 25 807 25 809 25 810 25 811 25 812 25 813 25 815 25 817 25 819 25 820 25 831 25 832 25 833 25 841 25 843 25 851 25 853 25 854 25 855 25 856 25 857 25 858 ASA Pilot compartment Pilot compartment doors Pilot compartment view Windshields and windows Cockpit controls Motion and effect of cockpit controls Cockpit control knob shape Fuselage doors Seats berths safety belts and harnesses Stowage compartments Retention of items of mass in passenger and crew compartments and galleys Passenger information signs and placards Floor surfaces Security considerations EMERGENCY PROVISIONS Ditching Emergency evacuation Emergency exits Emergency exit arrangement Emergency egress assist means and escape routes Emergency exit marking Emergency lighting Emergency exit access Width of aisle Maximum number of seats abreast Lower deck surface compartments including galleys Lavatory doors VENTILATION AND HEATING Ventilation Cabin ozone concentration Combustion heating systems PRESSURIZATION Pressurized cabins Tests for pressurized cabins FIRE PROTECTION Fire extinguishers Compartment interiors Lavatory fire protecti
197. IV TEST METHOD TO DETERMINE THE HEAT RELEASE RATE FROM CABIN MATERIALS EXPOSED TO RADIANT HEAT a Summary of Method Three or more speci mens representing the completed aircraft compo nent are tested Each test specimen is injected into an environmental chamber through which a constant flow of air passes The specimen s expo sure is determined by a radiant heat source ad justed to produce on the specimen the desired total heat flux of 3 5 W cm The specimen is tested with the exposed surface vertical Combus tion is initiated by piloted ignition The combustion products leaving the chamber are monitored in or der to calculate the release rate of heat b Apparatus The Ohio State University OSU rate of heat release apparatus as de Scribed below is used This is a modified version of the rate of heat release apparatus standardized by the American Society of Testing and Materials ASTM ASTM E 906 1 This apparatus is shown in Figures 1A and 1B of this part IV All exterior surfaces of the appa ratus except the holding chamber must be insu lated with 1 inch 25 mm thick low density high temperature fiberglass board insulation A gas keted door through which the sample injection rod slides must be used to form an airtight clo sure on the specimen hold chamber 2 Thermopile The temperature difference be tween the air entering the environmental chamber and that leaving must be monitored by a thermo pile h
198. Immediately swing the burner away from the test position Terminate test 7 minutes af ter initiating cushion exposure to the flame by use of a gaseous extinguishing agent i e Halon or 7 Determine the weight of the remains of the seat cushion specimen set left on the mounting stand to the nearest 0 02 pound 9 grams exclud ing all droppings 170 Federal Aviation Regulations h Test Report With respect to all specimen Sets tested for a particular seat cushion for which testing of compliance is performed the following information must be recorded 1 An identification and description of the specimens being tested 2 The number of specimen sets tested 3 The initial weight and residual weight of each set the calculated percentage weight loss of each set and the calculated average percentage weight loss for the total number of sets tested 4 The burn length for each set tested ASA Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 Figure 1 STEEL ANGLE i uim 1x 1x fe 25 x 25 x 3 see STEEL FLAT STOCK uo 1 x fg 88 38 x 3 mm 5 Y 12 304 3 lt SIDE VIEW FRONT VIEW 18 450 3 22 x 16 5613 mm TOP VIEW NOTE ALL JOINTS WELDED FLAT STOCK BUTT WELDED ALL MEASUREMENTS INSIDE ASA 171 Appendix to Part 25 Federal Aviation Regulations
199. Kilograms must be designed to limit the effects of an explosive or in cendiary device as follows 1 Flightdeck smoke protection Means must be provided to limit entry of smoke fumes and noxious gases into the flightdeck 2 Passenger cabin smoke protection Means must be provided to prevent passenger incapaci tation in the cabin resulting from smoke fumes and noxious gases as represented by the initial combined volumetric concentrations of 0 5996 carbon monoxide and 1 2396 carbon dioxide 3 Cargo compartment fire suppression An extinguishing agent must be capable of suppress ing a fire All cargo compartment fire suppression Systems must be designed to withstand the fol lowing effects including support structure dis placements or adjacent materials displacing against the distribution system i Impact or damage from a 0 5 inch diameter aluminum sphere traveling at 430 feet per second 131 1 meters per second 75 825 801 ii A 15 pound per square inch 103 4 kPa pressure load if the projected surface area of the component is greater than 4 square feet Any sin gle dimension greater than 4 feet 1 22 meters may be assumed to be 4 feet 1 22 meters in length and iii A 6 inch 0 152 meters displacement ex cept where limited by the fuselage contour from a single point force applied anywhere along the dis tribution system where relative movement be tween the system and its attachment can occur iv Par
200. Part 25 Airworthiness Standards Transport Category PART 25 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY AIRPLANES SPECIAL FEDERAL AVIATION REGULATIONS SFAR No 13 SFAR No 109 Subpart A General Sec 25 1 25 2 25 3 Applicability Special retroactive requirements Special provisions for ETOPS type design approvals 25 5 by reference Subpart B Flight GENERAL 25 21 Proof of compliance 25 23 Load distribution limits 25 25 Weight limits 25 27 Center of gravity limits 25 29 Empty weight and corresponding center of gravity 25 31 Removable ballast 25 33 Propeller speed and pitch limits PERFORMANCE General Stall speed Takeoff Takeoff speeds Accelerate stop distance Takeoff path Takeoff distance and takeoff run Takeoff flight path Climb general Landing climb All engines operating Climb One engine inoperative En route flight paths Landing 25 101 25 103 25 105 25 107 25 109 25 111 25 113 25 115 25 117 25 119 25 121 25 123 25 125 CONTROLLABILITY AND MANEUVERABILITY 25 143 General 25 145 Longitudinal control 25 147 Directional and lateral control 25 149 Minimum control speed TRIM 25 161 Trim STABILITY General Static longitudinal stability 25 171 25 173 ASA Part 25 25 175 Demonstration of static longitudinal stability Static lateral directional stability Dynamic stability 25 177 25 18
201. Position the burner in front of the thermo couple rake After checking for proper alignment rotate the burner to the warm up position turn on the blower motor igniters and fuel flow and light the burner Allow it to warm up for a period of 2 minutes Move the burner into the calibration posi tion and allow 1 minute for thermocouple stabili zation then record the temperature of each of the 7 thermocouples once every second for a period of 30 seconds Turn off burner rotate out of posi tion and allow to cool Calculate the average tem perature of each thermocouple over this 30 sec ond period and record The average temperature of each of the 7 thermocouples should be 1900 F 100 1038 56 6 If either the heat flux the temperatures are not within the specified range adjust the burner intake air velocity and repeat the proce dures of paragraphs 4 and 5 above to obtain the proper values Ensure that the inlet air velocity is within the range of 2150 ft min 50 ft min 10 92 0 25 m s 7 Calibrate prior to each test until consistency has been demonstrated After consistency has been confirmed several tests may be conducted with calibration conducted before and after a se ries of tests f Test procedure 1 Secure the two insulation blanket test spec imens to the test frame The insulation blankets should be attached to the test rig center vertical former using four spring clamps positioned as shown in
202. Spaces 4 0 102 mm 12 50 318 mm Y 0 50 13 mm 4 Bolts 1 0 25 7 Connection Flange 11 0 279 mm gt 6 0 152 mm Y lt 4 25 108 mm Material 0 050 1 3 mm stainless steel ASA 205 Appendix to Part 25 FIGURE 4 Calorimeter Position Relative to Burner Cone Federal Aviation Regulations 6 in x 12 in x 3 4 in G 1 in 25 mm 152 x 305 x19 mm dia hole for Marinite block calorimeter mounting lin 25 mm 6 1 8 in 152 x 3 mm 3 1 8 in 76 3 mm 12 1 8 l 305 3 mm Side View 3 4 in Burner Cone 19 mm i 1in 25 dia ke 4 1 8 in 102 mm Calorimeter Water cooled Rack mounts to ae test frame 25 x 25 x 3 mm Top View 206 ASA Part 25 Airworthiness Standards Transport Category FiGURE 5 Thermocouple Rake Position Relative to Burner Cone Appendix F to Part 25 Seven thermocouple wires Side View l BurneriCone 4 1 8 102 3 mm lt 1 in 25 mm Rack mounts to Steel angle test frame 1x1x 1 8 25 x 25 x 3 mm Top View ASA 207 Appendix to Part 25 FIGURE 6 Position of Backface Calorimeters Relative to Test Specimen Frame Federal Aviation Regulations Calorimeter 102 mm Burner Cone Draft Tube 208 ASA Part 25 Airwort
203. V4 for takeoff from a dry runway ASA 825 109 b The accelerate stop distance on a wet run way is the greater of the following distances 1 The accelerate stop distance on a dry run way determined in accordance with paragraph a of this section or 2 The accelerate stop distance determined in accordance with paragraph a of this section ex cept that the runway is wet and the corresponding wet runway values of Ver and V4 are used In de termining the wet runway accelerate stop dis tance the stopping force from the wheel brakes may never exceed i The wheel brakes stopping force determined in meeting the requirements of 25 101 i and paragraph a of this section and ii The force resulting from the wet runway braking coefficient of friction determined in accor dance with paragraphs c or d of this section as applicable taking into account the distribution of the normal load between braked and unbraked wheels at the most adverse center of gravity posi tion approved for takeoff c The wet runway braking coefficient of fric tion for a smooth wet runway is defined as a curve of friction coefficient versus ground speed and must be computed as follows 17 825 109 1 The maximum tire to ground wet runway braking coefficient of friction is defined as Tire Pressure psi Federal Aviation Regulations Maximum Braking Coefficient tire to ground 0 0437 a 0 320 5 3 0 805 5
204. X h V Vo Fo flow of methane at baseline 1pm F4 higher preset flow of methane 1pm Vo thermopile voltage at baseline mv V4 thermopile voltage at higher flow mv Ambient temperature K P Ambient pressure mm Hg P Water vapor pressure mm Hg 2 Heat release rates may be calculated from the reading of the thermopile output voltage at any instant of time as unu 02323m ASA Appendix F to Part 25 4 The specimen must be placed in the hold chamber with the radiation doors closed The air tight outer door must be secured and the record ing devices must be started The specimen must be retained in the hold chamber for 60 seconds plus or minus 10 seconds before injection The thermopile zero value must be determined dur ing the last 20 seconds of the hold period The sample must not be injected before completion of the Zero value determination 5 When the specimen is to be injected the ra diation doors must be opened After the specimen is injected into the environmental chamber the ra diation doors must be closed behind the speci men 6 Reserved 7 Injection of the specimen and closure of the inner door marks time zero A record of the ther mopile output with at least one data point per sec ond must be made during the time the specimen is in the environmental chamber 8 The test duration time is five minutes The lower pilot burner and the upper pilot burn
205. a new engine type certificate Problems in changed Systems only f Acceptance criteria The type and frequency of failures and malfunctions on ETOPS significant Systems that occur during the airplane flight test program and the airplane demonstration flight test program specified in section K25 3 2 d of this ap pendix must be consistent with the type and fre quency of failures and malfunctions that would be expected to occur on currently certificated air planes approved for ETOPS K25 3 3 Combined service experience and Early ETOPS method An applicant for ETOPS type design approval using the Early ETOPS method must comply with the following requirements a A service experience requirement of less than 15 000 engine hours for the world fleet of the candidate airplane engine combination b The Early ETOPS requirements of section K25 3 2 of this appendix except for the airplane demonstration specified in section K25 3 2 d of this appendix and c The flight test requirement of section K25 3 1 c of this appendix 227 Appendix L to Part 25 APPENDIX L TO PART 25 HIRF ENVIRONMENTS AND EQUIPMENT HIRF TEST LEVELS This appendix specifies the HIRF environments and equipment HIRF test levels for electrical and Federal Aviation Regulations b HIRF environment II is specified in the fol lowing table Table 1 Environment 11 Field strength volts meter
206. abin pressure altitude limits if it warns the flight crew when the cabin pressure altitude exceeds 10 000 feet 7 A warning placard at the pilot or flight engi neer station if the structure is not designed for pressure differentials up to the maximum relief valve setting in combination with landing loads 8 The pressure sensors necessary to meet the requirements of paragraphs b 5 and b 6 of this section and 25 1447 c must be located and the sensing system designed so that in the event of loss of cabin pressure in any passenger or crew compartment including upper and lower lobe galleys the warning and automatic presen tation devices required by those provisions will be actuated without any delay that would signifi cantly increase the hazards resulting from de compression Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55466 Dec 20 1976 Amdt 25 87 61 FR 28696 June 5 1996 88 Federal Aviation Regulations 825 843 Tests for pressurized cabins a Strength test The complete pressurized cabin including doors windows and valves must be tested as a pressure vessel for the pressure differential specified in 25 365 d b Functional tests The following functional tests must be performed 1 Tests of the functioning and capacity of the positive and negative pressure differential valves and of the emergency release valve to stimulate the effects of closed regulat
207. able ASTM K2 fuel Number 2 grade kerosene or ASTM D2 ASA Appendix F to Part 25 fuel Number 2 grade fuel oil or Number 2 diesel fuel are acceptable if the nominal fuel flow rate temperature and heat flux measurements con form to the requirements of paragraph Vll e of this appendix vii Fuel pressure regulator Provide a fuel pressure regulator adjusted to deliver a nominal 6 0 gal hr 0 378 L min flow rate An operating fuel pressure of 100 Ib in 0 71 MPa for a nomi nally rated 6 0 gal hr 80 spray angle nozzle such as a PL type delivers 6 0 0 2 gal hr 0 378 0 0126 L min SEE FIGURE 3 AT THE END OF PART VII OF THIS APPENDIX 3 Calibration rig and equipment i Construct individual calibration rigs to incor porate a calorimeter and thermocouple rake for the measurement of heat flux and temperature Position the calibration rigs to allow movement of the burner from the test rig position to either the heat flux or temperature position with minimal dif ficulty ii Calorimeter The calorimeter must be a total heat flux foil type Gardon Gage of an appropriate range such as 0 20 Btu ft 2 sec 0 22 7 W cm accurate to 3 of the indicated reading The heat flux calibration method must be in accor dance with paragraph VI b 7 of this appendix iii Calorimeter mounting Mount the calorime ter in a 6 by 12 0 125 inch 152 by 305 3 mm by 0 75 0 125 inch 19 mm 3 mm thick in sulating block which
208. able to the pilots without additional crewmember action after any single failure or combination of failures that is not shown to be extremely improb able and c Additional instruments systems or equip ment may not be connected to the operating sys tems for the required instruments unless provi sions are made to ensure the continued normal functioning of the required instruments in the event of any malfunction of the additional instru 121 825 1337 ments systems or equipment which is not shown to be extremely improbable Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5679 April 8 1970 Amdt 25 41 42 FR 36970 July 18 1977 825 1337 Powerplant instruments a Instruments and instrument lines 1 Each powerplant and auxiliary power unit instrument line must meet the requirements of 25 993 and 25 1183 2 Each line carrying flammable fluids under pressure must i Have restricting orifices or other safety de vices at the source of pressure to prevent the es cape of excessive fluid if the line fails and ii Be installed and located so that the escape of fluids would not create a hazard 3 Each powerplant and auxiliary power unit instrument that utilizes flammable fluids must be installed and located so that the escape of fluid would not create a hazard b Fuel quantity indicator There must be means to indicate to the flight crewmembers the quantity in gall
209. able to withstand without failure or leakage an internal pressure 1 5 times the maximum operat ing pressure 3 If a vent is provided the venting must be ef fective under all normal flight conditions 4 Reserved c Augmentation system drains must be de signed and located in accordance with 825 1455 if 1 The augmentation system fluid is subject to freezing and 2 The fluid may be drained in flight or during ground operation d The augmentation liquid tank capacity avail able for the use of each engine must be large enough to allow operation of the airplane under the approved procedures for the use of liquid aug mented power The computation of liquid con sumption must be based on the maximum ap proved rate appropriate for the desired engine output and must include the effect of temperature on engine performance as well as any other fac tors that might vary the amount of liquid required e This section does not apply to fuel injection systems ASA 825 953 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15043 March 17 1977 Amdt 25 72 55 FR 29785 July 20 1990 FUEL SYSTEM 825 951 General a Each fuel system must be constructed and arranged to ensure a flow of fuel at a rate and pressure established for proper engine and auxil iary power unit functioning under each likely oper ating condition including any maneuver for which certification is requested and du
210. ace of the tank compart ment must be smooth and free of projections that could cause wear of the liner unless i Provisions are made for protection of the liner at these points or ii The construction of the liner itself provides that protection b Spaces adjacent to tank surfaces must be ventilated to avoid fume accumulation due to mi nor leakage If the tank is in a sealed compart ment ventilation may be limited to drain holes large enough to prevent excessive pressure re sulting from altitude changes c The location of each tank must meet the re quirements of 25 1185 a d No engine nacelle skin immediately behind a major air outlet from the engine compartment may act as the wall of an integral tank e Each fuel tank must be isolated from per sonnel compartments by a fumeproof and fuel proof enclosure 25 969 Fuel tank expansion space Each fuel tank must have an expansion space of not less than 2 percent of the tank capacity It must be impossible to fill the expansion space in advertently with the airplane in the normal ground attitude For pressure fueling systems compli ance with this section may be shown with the means provided to comply with 25 979 b Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6913 May 5 1967 100 Federal Aviation Regulations 25 971 Fuel tank sump a Each fuel tank must have a sump with an ef fective capacity in the norma
211. ach lavatory served h Each receptacle used for the disposal of flammable waste material must be fully enclosed constructed of at least fire resistant materials and must contain fires likely to occur in it under normal use The capability of the receptacle to contain those fires under all probable conditions of wear misalignment and ventilation expected in service must be demonstrated by test Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29783 July 20 1990 Amdt 25 83 60 FR 6623 Feb 2 1995 Amdt 25 116 69 FR 62788 Oct 27 2004 825 854 Lavatory fire protection For airplanes with a passenger capacity of 20 or more a Each lavatory must be equipped with a smoke detector system or equivalent that pro vides a warning light in the cockpit or provides a warning light or audible warning in the passenger 89 825 855 cabin that would be readily detected by a flight at tendant and b Each lavatory must be equipped with a built in fire extinguisher for each disposal recepta cle for towels paper or waste located within the lavatory The extinguisher must be designed to discharge automatically into each disposal recep tacle upon occurrence of a fire in that receptacle Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 74 56 FR 15456 April 16 1991 25 855 Cargo or baggage compartments For each cargo or baggage compartment the following apply
212. ach side of the fuselage 4 For a passenger seating configuration of 20 to 40 seats there must be at least two exits one of which must be a Type or larger exit in each side of the fuselage 5 For a passenger seating configuration of 41 to 110 seats there must be at least two exits one of which must be a Type or larger exit in each side of the fuselage 78 Federal Aviation Regulations 6 For a passenger seating configuration of more than 110 seats the emergency exits in each side of the fuselage must include at least two Type or larger exits 7 The combined maximum number of pas senger seats permitted for all Type III exits is 70 and the combined maximum number of passen ger seats permitted for two Type III exits in each side of the fuselage that are separated by fewer than three passenger seat rows is 65 8 If a Type A Type B or Type C exit is in stalled there must be at least two Type C or larger exits in each side of the fuselage 9 If a passenger ventral or tail cone exit is in stalled and that exit provides at least the same rate of egress as a Type III exit with the airplane in the most adverse exit opening condition that would result from the collapse of one or more legs of the landing gear an increase in the passenger seating configuration is permitted as follows i For a ventral exit 12 additional passenger seats ii For a tail cone exit incorporating a floor level opening of not less
213. ach wheel with brakes capable of producing this ground reaction This nose tire load may not exceed 1 5 times the load rating of the tire 3 The ground reaction of the tire correspond ing to the most critical combination of airplane weight up to maximum ramp weight and center of gravity position combined with forces of 1 0g downward and 0 20g forward acting at the center of gravity The reactions in this case must be dis tributed to the nose and main wheels by the prin ciples of statics with a drag reaction equal to 0 20 times the vertical load at each wheel with brakes capable of producing this ground reaction This nose tire load may not exceed 1 5 times the load rating of the tire c When a landing gear axle is fitted with more than one wheel and tire assembly such as dual or dual tandem each wheel must be fitted with a suitable tire of proper fit with a speed rating ap proved by the Administrator that is not exceeded under critical conditions and with a load rating approved by the Administrator that is not ex ceeded by 1 The loads on each main wheel tire corre sponding to the most critical combination of air plane weight up to maximum weight and center of gravity position when multiplied by a factor of 1 07 and 2 Loads specified in paragraphs 2 b 1 b 2 and b 3 of this section on each nose wheel tire d Each tire installed on a retractable landing gear system must at the maximum size of
214. ady gradient of climb may not be less than 2 4 percent for two engine airplanes 2 7 percent for three engine airplanes and 3 0 percent for four engine airplanes at V2 with i The critical engine inoperative the remaining engines at the takeoff power or thrust available at the time the landing gear is fully retracted deter mined under 825 111 unless there is a more crit ical power operating condition existing later along the flight path but before the point where the air plane reaches a height of 400 feet above the take off surface and ii The weight equal to the weight existing when the airplane s landing gear is fully retracted determined under 825 111 2 The requirements of paragraph b 1 of this section must be met i In non icing conditions and ii In icing conditions with the takeoff ice accre tion defined in appendix C if in the configuration of 25 121 b with the takeoff ice accretion A The stall speed at maximum takeoff weight exceeds that in non icing conditions by more than the greater of 3 knots CAS percent of Vsp or B The degradation of the gradient of climb de termined in accordance with 825 121 b is greater than one half of the applicable actual to net take off flight path gradient reduction defined in 825 115 b c Final takeoff In the en route configuration at the end of the takeoff path determined in accor dance with 825 111 1 The steady gradient of climb may not be
215. agraphs b 3 i through iii of this sec tion do not apply to components that are redun dant and separated in accordance with paragraph c 2 of this section or are installed remotely from the cargo compartment c An airplane with a maximum certificated passenger seating capacity of more than 60 per sons or a maximum certificated takeoff gross weight of over 100 000 pounds 45 359 Kilo grams must comply with the following 1 Least risk bomb location An airplane must be designed with a designated location where a bomb or other explosive device could be placed to best protect flight critical structures and systems from damage in the case of detonation 2 Survivability of systems i Except where impracticable redundant air plane systems necessary for continued safe flight and landing must be physically separated at a minimum by an amount equal to a sphere of di ameter D 2 JH o where Hg is defined under 25 365 e 2 of this part and D need not exceed 5 05 feet 1 54 meters The sphere is applied everywhere within the fuselage limited by the forward bulkhead and the aft bulkhead of the passenger cabin and cargo compartment beyond which only one half the sphere is applied ii Where compliance with paragraph c 2 i of this section is impracticable other design pre cautions must be taken to maximize the surviv ability of those systems 3 Interior design to facilitate searches Design features mus
216. aintaining safe cell temperatures and pres sures presents no problem 3 No explosive or toxic gases emitted by any battery in normal operation or as the result of any probable malfunction in the charging system or battery installation may accumulate in hazardous quantities within the airplane 4 No corrosive fluids or gases that may es cape from the battery may damage surrounding airplane structures or adjacent essential equip ment 5 Each nickel cadmium battery installation must have provisions to prevent any hazardous effect on structure or essential systems that may be caused by the maximum amount of heat the battery can generate during a short circuit of the battery or of individual cells ASA 25 1357 6 Nickel cadmium battery installations must have i A system to control the charging rate of the battery automatically so as to prevent battery overheating ii A battery temperature sensing and over temperature warning system with a means for dis connecting the battery from its charging source in the event of an over temperature condition or iii A battery failure sensing and warning sys tem with a means for disconnecting the battery from its charging source in the event of battery failure c Electrical bonding must provide an ade quate electrical return path under both normal and fault conditions on airplanes having grounded electrical systems Docket No FAA 2004 18379 72 FR 63405 Nov 8
217. airplane is certificated i Brake wear indicators Means must be pro vided for each brake assembly to indicate when the heat sink is worn to the permissible limit The means must be reliable and readily visible j Overtemperature burst prevention Means must be provided in each braked wheel to prevent a wheel failure a tire burst or both that may re sult from elevated brake temperatures Addition 67 825 737 ally all wheels must meet the requirements of 25 731 d k Compatibility Compatibility of the wheel and brake assemblies with the airplane and its systems must be substantiated Docket No 1999 6063 67 FR 20420 April 24 2002 as amended by Amdt 25 108 67 FR 70828 Nov 26 2002 25 737 Skis Each ski must be approved The maximum limit load rating of each ski must equal or exceed the maximum limit load determined under the appli cable ground load requirements of this part FLOATS AND HULLS 25 751 Main float buoyancy Each main float must have a A buoyancy of 80 percent in excess of that required to support the maximum weight of the seaplane or amphibian in fresh water and b Not less than five watertight compartments approximately equal in volume 25 753 Main float design Each main float must be approved and must meet the requirements of 25 521 25 755 Hulls a Each hull must have enough watertight compartments so that with any two adjacent compartme
218. airplane unit served by the air e Each auxiliary power unit induction system duct must be fireproof for a sufficient distance up stream of the auxiliary power unit compartment to prevent hot gas reverse flow from burning through auxiliary power unit ducts and entering any other compartment or area of the airplane in which a hazard would be created resulting from the entry of hot gases The materials used to form the re mainder of the induction system duct and plenum chamber of the auxiliary power unit must be capa ble of resisting the maximum heat conditions likely to occur f Each auxiliary power unit induction system duct must be constructed of materials that will not absorb or trap hazardous quantities of flammable fluids that could be ignited in the event of a surge or reverse flow condition Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50597 Oct 30 1978 825 1105 Induction system screens If induction system screens are used a Each screen must be upstream of the car buretor b No screen may be in any part of the induc tion system that is the only passage through which air can reach the engine unless it can be deiced by heated air c No screen may be deiced by alcohol alone and d It must be impossible for fuel to strike any Screen 825 1107 Inter coolers and after coolers Each inter cooler and after cooler must be able to withstand any vibration inertia and ai
219. airplane weight may be assumed unless the presence of systems or procedures significantly affects the lift c The method of analysis of airplane and landing gear loads must take into account at least the following elements 1 Landing gear dynamic characteristics 2 Spin up and springback 3 Rigid body response 4 Structural dynamic response of the air frame if significant d The landing gear dynamic characteristics must be validated by tests as defined in 825 723 a e The coefficient of friction between the tires and the ground may be established by consider ing the effects of skidding velocity and tire pres sure However this coefficient of friction need not be more than 0 8 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 Amdt 25 91 62 FR 40705 July 29 1997 Amdt 25 103 66 FR 27394 May 16 2001 825 477 Landing gear arrangement Sections 25 479 through 25 485 apply to air planes with conventional arrangements of main 46 Federal Aviation Regulations and nose gears or main and tail gears when nor mal operating techniques are used 825 479 Level landing conditions a In the level attitude the airplane is assumed to contact the ground at forward velocity compo nents ranging from Vj 4 to 1 25 Vj 2 parallel to the ground under the conditions prescribed 825 473 with 1 Vi4 equal to Vso TAS at the appropriate landing wei
220. al density ft sec rad 6 7 root mean square gust velocity ft sec Q reduced frequency radians per foot L 2 500 ft 3 The limit loads must be obtained by multi plying the A values determined by the dynamic analysis by the following values of the gust veloc ity Uo i At speed Vc Uo 85 fps true gust velocity in the interval O to 30 000 ft altitude and is linearly decreased to 30 fps true gust velocity at 80 000 ft altitude Where the Administrator finds that a de sign is comparable to a similar design with exten sive satisfactory service experience it will be ac ceptable to select Uo at Vc less than 85 fps but not less than 75 fps with linear decrease from that value at 20 000 feet to 30 fps at 80 000 feet The following factors will be taken into account when assessing comparability to a similar design 1 The transfer function of the new design should exhibit no unusual characteristics as com pared to the similar design which will significantly affect response to turbulence e g coalescence of modal response in the frequency regime which can result in a significant increase of loads 2 The typical mission of the new airplane is substantially equivalent to that of the similar de sign ASA Appendix G to Part 25 3 The similar design should demonstrate the adequacy of the Uo selected ii At speed Vg Uo is equal to 1 32 times the values obtained under paragraph b 3 i of this appendix
221. al station as the upward component but is directed inward perpendicularly to the plane of symmetry at a point midway between the keel and chine lines c Unsymmetrical landing twin float sea planes The unsymmetrical loading consists of an upward load at the step of each float of 0 75 and a side load of 0 25 tan at one float times the step landing load reached under 825 527 The side load is directed inboard perpendicularly to the plane of symmetry midway between the keel and chine lines of the float at the same longitudinal station as the upward load 825 531 Hull and main float takeoff condition For the wing and its attachment to the hull or main float a The aerodynamic wing lift is assumed to be zero and b A downward inertia load corresponding to a load factor computed from the following formula must be applied _ _ 2 av py where n inertia load factor empirical seaplane operations factor equal to 0 004 seaplane stalling speed knots at the design takeoff weight with the flaps extended in the appropriate takeoff position angle of dead rise at the main step degrees and W design water takeoff weight in pounds Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 ASA Part 25 Airworthiness Standards Transport Category 825 533 Hull and main float bottom pressures a General The hull and main
222. an 6 1966 Amdt 25 25 35 FR 13192 Aug 19 1970 Amdt 25 37 40 FR 2577 Jan 14 1975 Amdt 25 41 42 FR 36971 July 18 1977 Amdt 25 65 53 FR 26144 July 11 1988 Amdt 25 124 73 FR 12563 March 7 2008 Amdt 25 124 74 FR 32800 July 9 2009 135 825 1461 825 1461 Equipment containing high energy rotors a Equipment containing high energy rotors must meet paragraph b c or d of this section b High energy rotors contained in equipment must be able to withstand damage caused by malfunctions vibration abnormal speeds and abnormal temperatures In addition 1 Auxiliary rotor cases must be able to con tain damage caused by the failure of high energy rotor blades and 2 Equipment control devices systems and instrumentation must reasonably ensure that no operating limitations affecting the integrity of high energy rotors will be exceeded in service c It must be shown by test that equipment containing high energy rotors can contain any fail ure of a high energy rotor that occurs at the high est speed obtainable with the normal speed con trol devices inoperative d Equipment containing high energy rotors must be located where rotor failure will neither en danger the occupants nor adversely affect contin ued safe flight Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 41 42 FR 36971 July 18 1977 136 Federal Aviation Regulations ASA Part 25 Airwort
223. and 2 For airplanes that have a passenger seating configuration excluding pilots seats of 10 seats or more the spillage of enough fuel from any part of the fuel system to constitute a fire hazard b Each airplane that has a passenger seating configuration excluding pilots seats of 10 seats or more must be designed so that with the airplane under control it can be landed on a paved runway with any one or more landing gear legs not ex tended without sustaining a structural component failure that is likely to cause the spillage of enough fuel to constitute a fire hazard c Compliance with the provisions of this sec tion may be shown by analysis or tests or both Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 32 37 FR 3969 Feb 24 1972 ASA Part 25 Airworthiness Standards Transport Category 825 723 Shock absorption tests a The analytical representation of the landing gear dynamic characteristics that is used in deter mining the landing loads must be validated by en ergy absorption tests A range of tests must be conducted to ensure that the analytical represen tation is valid for the design conditions specified in 825 473 1 The configurations subjected to energy ab sorption tests at limit design conditions must in clude at least the design landing weight or the de sign takeoff weight whichever produces the greater value of landing impact energy 2 The test attitude of the landi
224. and adverse conditions which must be considered in showing compliance with this section are 1 Any critical fuel loading conditions not shown to be extremely improbable which may re sult from mismanagement of fuel 2 Any single failure in any flutter damper sys tem 3 For airplanes not approved for operation in icing conditions the maximum likely ice accumula tion expected as a result of an inadvertent encoun ter 4 Failure of any single element of the struc ture supporting any engine independently mounted propeller shaft large auxiliary power unit or large externally mounted aerodynamic body such as an external fuel tank 5 For airplanes with engines that have propel lers or large rotating devices capable of significant dynamic forces any single failure of the engine structure that would reduce the rigidity of the rota tional axis 6 The absence of aerodynamic or gyroscopic forces resulting from the most adverse combina tion of feathered propellers or other rotating de vices capable of significant dynamic forces In ad dition the effect of a single feathered propeller or rotating device must be coupled with the failures of paragraphs d 4 and d 5 of this section 7 Any single propeller or rotating device capa ble of significant dynamic forces rotating at the highest likely overspeed 8 Any damage or failure condition required or selected for investigation by 825 571 The single ASA Part
225. and modes of damage due to fatigue corrosion or accidental damage Repeated load and static analyses supported by test evidence and if available service experience must also be incorporated in the evaluation Special consider ation for widespread fatigue damage must be in cluded where the design is such that this type of damage could occur It must be demonstrated with sufficient full scale fatigue test evidence that widespread fatigue damage will not occur within the design service goal of the airplane The type certificate may be issued prior to completion of full scale fatigue testing provided the Administra tor has approved a plan for completing the re quired tests and the airworthiness limitations section of the instructions for continued airworthi ness required by 825 1529 of this part specifies that no airplane may be operated beyond a num ber of cycles equal to 1 the number of cycles ac cumulated on the fatigue test article until such testing is completed The extent of damage for re sidual strength evaluation at any time within the operational life of the airplane must be consistent with the initial detectability and subsequent growth under repeated loads The residual strength evaluation must show that the remaining structure is able to withstand loads considered as static ultimate loads corresponding to the fol lowing conditions 1 The limit symmetrical maneuvering condi tions specified in 825 337 at all speeds up
226. and procedures set forth the Fuel Tank Flammability Assessment Method Users Manual dated May 2008 docu ment number DOT FAA AR 05 8 incorporated by reference see 825 5 2 Any fuel tank other than a main fuel tank on an airplane must meet the flammability exposure criteria of Appendix M to this part if any portion of the tank is located within the fuselage contour 3 As used in this paragraph i Equivalent Conventional Unheated Alumi num Wing Tank is an integral tank in an unheated semi monocoque aluminum wing of a subsonic airplane that is equivalent in aerodynamic perfor mance structural capability fuel tank capacity and tank configuration to the designed wing ii Fleet Average Flammability Exposure is de fined in Appendix N to this part and means the percentage of time each fuel tank ullage is flam mable for a fleet of an airplane type operating over the range of flight lengths iii Main Fuel Tank means a fuel tank that feeds fuel directly into one or more engines and holds required fuel reserves continually through out each flight c Paragraph b of this section does not apply to a fuel tank if means are provided to mitigate the effects of an ignition of fuel vapors within that fuel tank such that no damage caused by an ignition will prevent continued safe flight and landing d Critical design configuration control limita tions CDCCL inspections or other procedures must be established as necessary
227. ane Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 4 30 FR 6113 April 30 1965 25 772 Pilot compartment doors For an airplane that has a lockable door in stalled between the pilot compartment and the passenger compartment a For airplanes with a maximum passenger seating configuration of more than 20 seats the emergency exit configuration must be designed so that neither crewmembers nor passengers re quire use of the flightdeck door in order to reach the emergency exits provided for them and b Means must be provided to enable flight crewmembers to directly enter the passenger compartment from the pilot compartment if the cockpit door becomes jammed c There must be an emergency means to en able a flight attendant to enter the pilot compart ment in the event that the flightcrew becomes in capacitated Docket No 24344 55 FR 29777 July 20 1990 as amended by Amdt 25 106 67 FR 2127 Jan 15 2002 25 773 Pilot compartment view a Nonprecipitation conditions For nonprecipi tation conditions the following apply 1 Each pilot compartment must be arranged to give the pilots a sufficiently extensive clear and undistorted view to enable them to safely perform any maneuvers within the operating limi tations of the airplane including taxiing takeoff approach and landing 2 Each pilot compartment must be free of glare and reflection that could interfere with the normal dut
228. anes subsequent to the effective date of this regulation 1 It is not intended to waive compliance with such airworthiness requirements as are included in the operating parts of the Civil Air Regulations for specific types of operation 2 General modifications Except as modified in sections 3 and 4 of this regulation an applicant for approval of modifications to a DC 3 or L 18 air plane which result in changes in design or in changes to approved limitations shall show that the modifications were accomplished in accor dance with the rules of either Part 4a or Part 4b in effect on September 1 1953 which are applica ble to the modification being made Provided That an applicant may elect to accomplish a mod ification in accordance with the rules of Part 4b in effect on the date of application for the modifica tion in lieu of Part 4a or Part 4b as in effect on Federal Aviation Regulations September 1 1953 And provided further That each specific modification must be accomplished in accordance with all of the provisions contained in the elected rules relating to the particular modi fication 3 Specific conditions for approval appli cant for any approval of the following specific changes shall comply with section 2 of this regu lation as modified by the applicable provisions of this section a Increase in take off power limitation 1 200 to 1 350 horsepower The engine take off power limitation for the airplane m
229. anes with four or more engines Airplanes with four or more engines must be able to make 20 banked turns with and against the inoperative engines from steady flight at a speed equal to 1 3 with maximum con tinuous power and with the airplane in the config uration prescribed by paragraph b of this sec tion e Lateral control all engines operating With the engines operating roll response must allow normal maneuvers such as recovery from upsets 25 825 149 produced by gusts and the initiation of evasive maneuvers There must be enough excess lat eral control in sideslips up to sideslip angles that might be required in normal operation to allow a limited amount of maneuvering and to correct for gusts Lateral control must be enough at any speed up to to provide a peak roll rate necessary for safety without excessive control forces or travel Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2321 Jan 16 1978 Amdt 25 72 55 FR 29774 July 20 1990 Amdt 25 108 Amdt 25 108 67 FR 70827 Nov 26 2002 25 149 Minimum control speed a In establishing the minimum control speeds required by this section the method used to simu late critical engine failure must represent the most critical mode of powerplant failure with respect to controllability expected in service b Vuc is the calibrated airspeed at which when the critical engine is suddenly made ino
230. any common es cape route from two Type Ill passenger emer gency exits must be at least 42 inches wide that from any other passenger emergency exit must be at least 24 inches wide and 2 The escape route surface must have a re flectance of at least 80 percent and must be de fined by markings with a surface to marking con trast ratio of at least 5 1 d Means must be provided to assist evacuees to reach the ground for all Type C exits located over the wing and if the place on the airplane structure at which the escape route required in paragraph c of this section terminates is more than 6 feet from the ground with the airplane on the ground and the landing gear extended for all other exit types 1 If the escape route is over the flap the height of the terminal edge must be measured ASA Part 25 Airworthiness Standards Transport Category with the flap in the takeoff or landing position whichever is higher from the ground 2 The assisting means must be usable and self supporting with one or more landing gear legs collapsed and under a 25 knot wind directed from the most critical angle 3 The assisting means provided for each es cape route leading from a Type A or B emergency exit must be capable of carrying simultaneously two parallel lines of evacuees and the assisting means leading from any other exit type must be capable of carrying as many parallel lines of evac uees as there are required escape routes
231. approved equivalent non destructive inspection methods 2 For each critical casting with a casting fac tor less than 1 50 three sample castings must be static tested and shown to meet i The strength requirements of 825 305 at an ultimate load corresponding to a casting factor of 1 25 and ii The deformation requirements of 825 305 at a load of 1 15 times the limit load 3 Examples of these castings are structural attachment fittings parts of flight control systems control surface hinges and balance weight attach ments seat berth safety belt and fuel and oil tank supports and attachments and cabin pres sure valves d Noncritical castings For each casting other than those specified in paragraph c of this sec tion the following apply 1 Except as provided in paragraphs d 2 and 3 of this section the casting factors and corre sponding inspections must meet the following ta ble Casting factor Inspection 2 0 or more Less than 2 0 but more than 1 5 100 percent visual 100 percent visual and magnetic particle or penetrant or equivalent nondestructive inspection methods 1 25 through 1 50 100 percent visual magnetic particle or penetrant and radiographic or approved equivalent nondestructive inspection methods 2 The percentage of castings inspected by nonvisual methods may be reduced below that specified in paragraph d 1 of this section when an approved quali
232. ar and flap positions and symmetrical power conditions at speeds from 1 13 to Vee Vi E or Vrc Mrc as appropriate d The rudder gradients must meet the re quirements of paragraph c at speeds between except that the dihedral effect aileron deflection opposite the correspond ing rudder input may be negative provided the di vergence is gradual easily recognized and easily controlled by the pilot Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29774 July 20 1990 55 FR 37607 Sept 12 1990 Amdt 25 108 67 FR 70827 Nov 26 2002 ASA 825 201 825 181 Dynamic stability a Any short period oscillation not including combined lateral directional oscillations occur ring between 1 13 Vsr and maximum allowable speed appropriate to the configuration of the air plane must be heavily damped with the primary controls 1 Free and 2 In a fixed position b Any combined lateral directional oscilla tions Dutch roll occurring between 1 13 Vsg and maximum allowable speed appropriate to the configuration of the airplane must be positively damped with controls free and must be controlla ble with normal use of the primary controls with out requiring exceptional pilot skill Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2322 Jan 16 1978 Amdt 25 72 55 FR 29775 July 20 1990 55 FR 37607 Sept 12 19
233. ar that are required for static equilibrium 4 A side load factor at the airplane center of gravity of zero c If the loads prescribed in paragraph b of this section result in a nose gear side load higher than 0 8 times the vertical nose gear load the de sign nose gear side load may be limited to 0 8 times the vertical load with unbalanced yawing moments assumed to be resisted by airplane in ertia forces d For other than the nose gear its attaching structure and the forward fuselage structure the loading conditions are those prescribed in para graph b of this section except that 1 A lower drag reaction may be used if an ef fective drag force of 0 8 times the vertical reaction cannot be reached under any likely loading condi tion and 2 The forward acting load at the center of gravity need not exceed the maximum drag reac tion on one main gear determined in accordance with 25 493 b e With the airplane at design ramp weight and the nose gear in any steerable position the com bined application of full normal steering torque and vertical force equal to 1 33 times the maximum static reaction on the nose gear must be consid ered in designing the nose gear its attaching structure and the forward fuselage structure Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 Amdt 25 46 43 FR 50595 Oct 30 1978 Amdt 25 91 62 FR 40705 July 29 1997 8
234. ards Transport Category sary to accommodate test stand interfaces with the engine nacelle package At the conclusion of the test the propulsion system must be i Visually inspected according to the appli cant s on wing inspection recommendations and limits and ii Completely disassembled and the propul sion system hardware inspected to determine whether it meets the service limits specified in the Instructions for Continued Airworthiness submit ted in compliance with 825 1529 2 The applicant must identify track and re solve each cause or potential cause of IFSD loss of thrust control or other power loss encountered during this inspection in accordance with the problem tracking and resolution system specified in section K25 2 2 h of this appendix e New technology testing Technology new to the applicant including substantially new manu facturing techniques must be tested to substanti ate its suitability for the airplane design f APU validation test If an APU is needed to comply with this appendix one APU of the type to be certified with the airplane must be tested for 3 000 equivalent airplane operational cycles Fol lowing completion of the test the APU must be disassembled and inspected The applicant must identify track and resolve each cause or potential cause of an inability to start or operate the APU in flight as intended in accordance with the problem tracking and resolution system specified in
235. at is accessible to crewmembers in flight 4 At least one hand fire extinguisher must be located in or readily accessible for use in each galley located above or below the passenger compartment 5 Each hand fire extinguisher must be ap proved ASA Part 25 Airworthiness Standards Transport Category 6 At least one of the required fire extinguish ers located in the passenger compartment of an airplane with a passenger capacity of at least 31 and not more than 60 and at least two of the fire extinguishers located in the passenger compart ment of an airplane with a passenger capacity of 61 or more must contain Halon 1211 bromochlo rodifluoromethane CBrC4F5 or equivalent as the extinguishing agent The type of extinguishing agent used in any other extinguisher required by this section must be appropriate for the kinds of fires likely to occur where used 7 The quantity of extinguishing agent used in each extinguisher required by this section must be appropriate for the kinds of fires likely to occur where used 8 Each extinguisher intended for use in a per sonnel compartment must be designed to mini mize the hazard of toxic gas concentration b Built in fire extinguishers If a built in fire ex tinguisher is provided 1 Each built in fire extinguishing system must be installed so that i No extinguishing agent likely to enter per sonnel compartments will be hazardous to the oc cupants and ii
236. at occur during the airplane flight test program and the airplane demonstration flight test program specified in section K25 2 2 g of this ap pendix must be consistent with the type and fre quency of failures and malfunctions that would be expected to occur on currently certificated air planes approved for ETOPS K25 2 3 Combined service experience and Early ETOPS method An applicant for ETOPS type design approval using the combined service experience and Early ETOPS method must comply with the following requirements A service experience requirement of not less than 15 000 engine hours for the world fleet of the candidate airplane engine combination b The Early ETOPS requirements of K25 2 2 except for the airplane demonstration specified in section K25 2 2 g of this appendix and c The flight test requirement of section K25 2 1 e of this appendix K25 3 AIRPLANES WITH MORE THAN TWO ENGINES An applicant for ETOPS type design approval of an airplane with more than two engines must use one of the methods described in section K25 3 1 K25 3 2 or K25 3 3 of this appendix K25 3 1 Service experience method An applicant for ETOPS type design approval using the service experience method must com ply with section K25 3 1 a of this appendix before conducting the airplane systems assessment specified in K25 3 1 b and the flight test speci fied in section K25 3 1 c of this appendix Service experience The wo
237. automatically available as an alternate source of electrical en ergy to allow continued engine operation if any battery becomes depleted ASA Part 25 Airworthiness Standards Transport Category b The capacity of batteries and generators must be large enough to meet the simultaneous demands of the engine ignition system and the greatest demands of any electrical system com ponents that draw electrical energy from the same source c The design of the engine ignition system must account for 1 The condition of an inoperative generator 2 The condition of a completely depleted bat tery with the generator running at its normal oper ating speed and 3 The condition of a completely depleted bat tery with the generator operating at idling speed if there is only one battery d Magneto ground wiring for separate igni tion circuits that lies on the engine side of the fire wall must be installed located or protected to minimize the probability of simultaneous failure of two or more wires as a result of mechanical dam age electrical faults or other cause e No ground wire for any engine may be routed through a fire zone of another engine un less each part of that wire within that zone is fire proof f Each ignition system must be independent of any electrical circuit not used for assisting controlling or analyzing the operation of that system g There must be means to warn appropriate flight crewme
238. aving five hot and five cold 24 gauge Chromel Alumel junctions The hot junctions must be spaced across the top of the exhaust stack 38 inches 10 mm below the top of the chimney The thermocouples must have a 050 010 inch 1 3 3 mm diameter ball type welded tip One thermocouple must be located in the geometric center with the other four located 1 18 inch 30 mm from the center along the diagonal toward each of the corners Figure 5 of this part IV The cold junctions must be located in the pan below the lower air distribution plate see paragraph b 4 of this part IV Thermopile hot junctions must be cleared of soot deposits as needed to maintain the calibrated sensitivity 3 Radiation Source radiant heat source in corporating four Type LL silicon carbide elements 20 inches 508 mm long by 63 inch 16 mm O D must be used as shown in Figures 2A and 2B of this part IV The heat source must have a nominal resistance of 1 4 ohms and be capable of generating a flux up to 100 kW m2 The silicone carbide elements must be mounted in the stain less steel panel box by inserting them through 63 inch 16 mm holes in 03 inch 1 mm thick ce ramic fiber or calcium silicate millboard Locations of the holes in the pads and stainless steel cover ASA Appendix F to Part 25 plates are shown in Figure 2B of this part IV The truncated diamond shaped mask of 042 002 inch 1 07 05 mm stainless steel must be added
239. ay be increased to more than 1 200 horsepower but not to more than 1 350 horsepower per engine if the increase in power does not adversely affect the flight charac teristics of the airplane b Increase in take off power limitation to more than 1 350 horsepower The engine take off power limitation for the airplane may be increased to more than 1 350 horsepower per engine if com pliance is shown with the flight characteristics and ground handling requirements of Part 4b c Installation of engines of not more than 1 830 cubic inches displacement and not having a certificated take off rating of more than 1 350 horsepower Engines of not more than 1 830 cu bic inches displacement and not having a certifi cated take off rating of more than 1 350 horse power which necessitate a major modification of redesign of the engine installation may be in stalled if the engine fire prevention and fire pro tection are equivalent to that on the prior engine installation d Installation of engines of more than 1 830 cubic inches displacement or having certificated take off rating of more than 1 350 horsepower Engines of more than 1 830 cubic inches dis placement or having certificated take off rating of more than 1 350 horsepower may be installed if compliance is shown with the engine installation requirements of Part 4b Provided That where lit eral compliance with the engine installation re quirements of Part 4b is extremely difficult to ac
240. ay force are an accept able means of compliance b Injury criteria for multiple occupancy side facing seats The following requirements are only applicable to airplanes that are subject to 825 562 1 Existing Criteria All injury protection criteria of 825 562 c 1 through c 6 apply to the occu pants of side facing seating The Head Injury Cri terion HIC assessments are only required for head contact with the seat and or adjacent struc tures 2 Body to Body Contact Contact between the head pelvis torso or shoulder area of one An thropomorphic Test Dummy ATD with the head pelvis torso or shoulder area of the ATD in the ad jacent seat is not allowed during the tests con ducted in accordance with 25 562 b 1 and b 2 Contact during rebound is allowed 3 Thoracic Trauma If the torso of an ATD at the forward most seat place impacts the seat and or adjacent structure during testing compli ance with the Thoracic Trauma Index TTI injury criterion must be substantiated by dynamic test or by rational analysis based on previous test s of a similar seat installation data must be ac quired with a Side Impact Dummy SID as de fined by 49 CFR part 572 subpart F or an equiv alent ATD or a more appropriate ATD and must be Federal Aviation Regulations processed as defined in Federal Motor Vehicle Safety Standards FMVSS part 571 214 section 56 13 5 49 CFR 571 214 The must be less than 85
241. be a straight length of 25 inch 6 3 mm O D 03 inch 0 8 mm wall stainless steel tubing that is 14 inches 360 mm long One end of the tubing must be closed and three No 40 drill holes must be drilled into the tubing 2 38 inch 60 mm apart for gas ports all radiating in the same direction The first hole must be 19 inch 5 mm from the closed end of the tubing The tube must be positioned 75 inch 19 mm above and 75 inch 19 mm behind the exposed upper edge of the specimen The middle hole must be in the vertical plane perpendicular to the exposed surface of the specimen which passes through its vertical centerline and must be pointed toward the radiation source The gas supplied to the burner must be methane and must be adjusted to pro duce flame lengths of 1 inch 25 mm iii Optional Fourteen Hole Upper Pilot Burner This burner may be used in lieu of the standard three hole burner described in paragraph b 8 ii of this part IV The pilot burner must be a straight length of 25 inch 6 3 mm O D 03 inch 0 8 mm wall stainless steel tubing that is 15 75 inches 400 mm long One end of the tubing must 182 Federal Aviation Regulations be closed and 14 No 59 drill holes must be drilled into the tubing 50 inch 13 mm apart for gas ports all radiating in the same direction The first hole must be 50 inch 13 mm from the closed end of the tubing The tube must be posi tioned above the specimen holder s
242. be damaged by foreign objects entering the air inlet it must be shown by tests or if appropriate by analysis that the induction system design can withstand the foreign object ingestion test conditions of 8833 76 33 77 and 33 78 a 1 of this chapter without failure of parts or components that could create a hazard Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 40 42 FR 15043 March 17 1977 Amdt 25 57 49 FR 6849 Feb 23 1984 Amdt 25 100 65 FR 55854 Sept 14 2000 25 1093 Induction system icing protection a Reciprocating engines Each reciprocating engine air induction system must have means to prevent and eliminate icing Unless this is done by other means it must be shown that in air free of visible moisture at a temperature of 30 F each airplane with altitude engines using 1 Conventional venturi carburetors have a preheater that can provide a heat rise of 120 F with the engine at 60 percent of maximum contin uous power or 2 Carburetors tending to reduce the probabil ity of ice formation has a preheater that can pro vide a heat rise of 100 F with the engine at 60 percent of maximum continuous power b Turbine engines 1 Each turbine engine must operate through out the flight power range of the engine including ASA 25 1103 idling without the accumulation of ice on the en gine inlet system components or airfram
243. ble airworthiness requirements are met In addition the flight paths must comply with subparagraphs i and ii of this paragraph i The flight paths must be established without changing the appropriate airplane configuration ii The flight paths must be carried out for a minimum height of 400 feet above the point where standby power is actuated 6 Airplane configuration speed and power and thrust general Any change in the airplane s configuration speed and power or thrust or both must be made in accordance with the procedures established by the applicant for the operation of the airplane in service and must comply with paragraphs a through c of this section In addi tion procedures must be established for the exe cution of balked landings and missed ap proaches The Administrator must find that the proce dure can be consistently executed in service by crews of average skill b The procedure may not involve methods or the use of devices which have not been proven to be safe and reliable ASA Appendix E to Part 25 c Allowances must be made for such time de lays in the execution of the procedures as may be reasonably expected to occur during service 7 Installation and operation standby power The standby power unit and its installation must comply with paragraphs a and b of this section a The standby power unit and its installation must not adversely affect the safety of the airplane
244. ble within the limitations specified in paragraph a of this section it is assumed that the airplane yaws to the overswing sideslip angle c With the airplane yawed to the static equilib rium sideslip angle it is assumed that the cockpit rudder control is held so as to achieve the maxi mum rudder deflection available within the limita tions specified in paragraph a of this section d With the airplane yawed to the static equilib rium sideslip angle of paragraph c of this sec tion it is assumed that the cockpit rudder control is suddenly returned to neutral Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 46 43 FR 50595 Oct 30 1978 Amdt 25 72 55 FR 29775 July 20 1990 55 FR 37607 Sept 12 1990 55 FR 41415 Oct 11 1990 Amdt 25 86 61 FR 5222 Feb 9 1996 Amdt 25 91 62 FR 40704 July 29 1997 SUPPLEMENTARY CONDITIONS 525 361 Engine torque a Each engine mount and its supporting structure must be designed for the effects of 1 A limit engine torque corresponding to take off power and propeller speed acting simulta neously with 75 percent of the limit loads from flight condition A of 25 333 b 2 A limit torque corresponding to the maxi mum continuous power and propeller speed act ASA 825 365 ing simultaneously with the limit loads from flight condition A of 25 333 b and 3 For turbopropeller installations in a
245. blower motor Note The Omega HH30 air velocity meter measures 2 625 inches in diameter To calculate the intake airflow multiply the cross sectional area 0 03758 ft2 by the air velocity 2150 ft min to obtain 80 80 ft min An air velocity meter other than the HH30 unit can be used provided the calculated airflow of 80 80 ft3 min 2 29 m3 min is equivalent 3 Rotate the burner from the test position to the warm up position Prior to lighting the burner ensure that the calorimeter face is clean of soot deposits and there is water running through the calorimeter Examine and clean the burner cone of any evidence of buildup of products of combus tion soot etc Soot buildup inside the burner ASA Appendix F to Part 25 cone may affect the flame characteristics and cause calibration difficulties Since the burner cone may distort with time dimensions should be checked periodically 4 While the burner is still rotated to the warm up position turn on the blower motor igniters and fuel flow and light the burner Allow it to warm up for a period of 2 minutes Move the burner into the calibration position and allow 1 minute for calorim eter stabilization then record the heat flux once every second for a period of 30 seconds Turn off burner rotate out of position and allow to cool Calculate the average heat flux over this 30 sec ond duration The average heat flux should be 16 0 0 8 Btu ft sec 18 2 0 9 W cm 5
246. burner without turning the fuel or igniters on Mea sure the air velocity using a hot wire anemometer in the center of the draft tube across the face of the opening Adjust the damper such that the air velocity is in the range of 1550 to 1800 ft min If tabs are being used at the exit of the draft tube they must be removed prior to this measurement Reinstall the draft tube extension cone 2 Place the calorimeter on the test stand as shown in Figure 2 at a distance of 8 inches 203 mm from the exit of the burner cone to simulate the position of the horizontal test specimen 3 Turn on the burner allow it to run for 2 min utes for warm up and adjust the damper to pro duce a calorimeter reading of 8 0 0 5 BTU per ft 2 sec 9 1 0 6 Watts cm 4 Replace the calorimeter with the thermo couple rake see Figure 3 5 Turn on the burner and ensure that each of the seven thermocouples reads 1700 F 100 F 927 C 38 C to ensure steady state conditions have been achieved If the temperature is out of this range repeat steps 2 through 5 until proper readings are obtained 6 Turn off the burner and remove the thermo couple rake 7 Repeat 1 to ensure that the burner is in the correct range g Test Procedure 1 Mount a thermocouple of the same type as that used for calibration at a distance of 4 inches 102 mm above the horizontal ceiling test spec imen The thermocouple should be centered over the burner co
247. by a crewmember during normal operations must not cause dangerous inadvertent movement by the crewmember or injury to the crewmember Docket No FAA 2004 18379 72 FR 63406 Nov 8 2007 825 1362 Electrical supplies for emergency conditions A suitable electrical supply must be provided to those services required for emergency proce dures after an emergency landing or ditching The circuits for these services must be designed pro tected and installed so that the risk of the ser vices being rendered ineffective under these emergency conditions is minimized Docket No FAA 2004 18379 72 FR 63406 Nov 8 2007 825 1363 Electrical system tests a When laboratory tests of the electrical sys tem are conducted 1 The tests must be performed on a mock up using the same generating equipment used in the airplane 2 The equipment must simulate the electrical characteristics of the distribution wiring and con nected loads to the extent necessary for valid test results and 124 Federal Aviation Regulations 3 Laboratory generator drives must simulate the actual prime movers on the airplane with re spect to their reaction to generator loading in cluding loading due to faults b For each flight condition that cannot be sim ulated adequately in the laboratory or by ground tests on the airplane flight tests must be made 825 1365 Electrical appliances motors and transformers a Domestic appliances
248. c 20 1976 Amdt 25 42 43 FR 2320 Jan 16 1978 Amdt 25 92 63 FR 8318 Feb 18 1998 Amdt 25 94 63 FR 8848 Feb 23 1998 Amdt 25 108 67 FR 70826 Nov 26 2002 Amdt 25 121 72 FR 44665 Aug 8 2007 825 109 Accelerate stop distance a The accelerate stop distance on a dry run way is the greater of the following distances 1 The sum of the distances necessary to i Accelerate the airplane from a standing start with all engines operating to Ver for takeoff from a dry runway ii Allow the airplane to accelerate from Vef to the highest speed reached during the rejected takeoff assuming the critical engine fails at Ver and the pilot takes the first action to reject the takeoff at the V4 for takeoff from a dry runway and iii Come to a full stop on a dry runway from the speed reached as prescribed in paragraph a 1 ii of this section plus iv A distance equivalent to 2 seconds at the V4 for takeoff from a dry runway 2 The sum of the distances necessary to i Accelerate the airplane from a standing start with all engines operating to the highest speed reached during the rejected takeoff assuming the pilot takes the first action to reject the takeoff at the V4 for takeoff from a dry runway and ii With all engines still operating come to a full stop on dry runway from the speed reached as prescribed in paragraph a 2 i of this section plus iii A distance equivalent to 2 seconds at the
249. c thermally activated fire sup pression system must be installed to extinguish a fire at the cooktop and immediately adjacent sur faces The agent used in the system must be an approved total flooding agent suitable for use in an occupied area The fire suppression system must have a manual override The automatic acti vation of the fire suppression system must also automatically shut off power to the cooktop e The surfaces of the galley surrounding the cooktop which would be exposed to a fire on the cooktop surface or in cookware on the cooktop ASA Part 25 Airworthiness Standards Transport Category must be constructed of materials that comply with the flammability requirements of part 1 of Appen dix F to part 25 This requirement is in addition to the flammability requirements typically required of the materials in these galley surfaces During the selection of these materials consideration must also be given to ensure that the flammability char acteristics of the materials will not be adversely affected by the use of cleaning agents and uten 515 used to remove cooking stains f The cooktop must be ventilated with a sys tem independent of the airplane cabin and cargo ventilation system Procedures and time intervals must be established to inspect and clean or re place the ventilation system to prevent a fire haz ard from the accumulation of flammable oils and be included in the instructions for continued air worthines
250. cally heated have a clear surface area on each side of the plate of at least 2 by 2 inches 51 by 51 mm and be 1 8 inch 1 16 inch thick 3 2 1 6 mm F Center the 2 transducers on opposite sides of the plates at equal distances from the plate G The distance of the calorimeter to the plate must be no less than 0 0625 inches 1 6 mm nor greater than 0 375 inches 9 5 mm H The range used in calibration must be at least 0 3 5 BTUs ft2 second 0 3 9 Watts cm and no greater than 0 5 7 BTUs ft second 0 6 4 Watts cm 1 The recording device used must record the 2 transducers simultaneously or at least within 1 10 of each other 8 Calorimeter fixture With the sliding platform pulled out of the chamber install the calorimeter holding frame and place a sheet of non combusti ble material in the bottom of the sliding platform adjacent to the holding frame This will prevent heat losses during calibration The frame must be 13 1 8 inches 333 mm deep front to back by 8 inches 203 mm wide and must rest on the top of the sliding platform It must be fabricated of 1 8 inch 3 2 mm flat stock steel and have an opening that accommodates a 1 2 inch 12 7 mm thick piece of refractory board which is level with the top of the sliding platform The board must have three 1 inch 25 4 mm diameter holes drilled through the board for calorimeter insertion The distance to the radiant panel surface from the cen terline of the firs
251. cant establishes that their application for particu lar structure is impractical This structure must be shown by analysis supported by test evidence to be able to withstand the repeated loads of vari able magnitude expected during its service life without detectable cracks Appropriate safe life scatter factors must be applied d Sonic fatigue strength It must be shown by analysis supported by test evidence or by the service history of airplanes of similar structural design and sonic excitation environment that 1 Sonic fatigue cracks are not probable in any part of the flight structure subject to sonic excita tion or 2 Catastrophic failure caused by sonic cracks is not probable assuming that the loads pre scribed in paragraph b of this section are ap plied to all areas affected by those cracks e Damage tolerance discrete source evalu ation The airplane must be capable of success fully completing a flight during which likely struc tural damage occurs as a result of 1 Impact with a 4 pound bird when the veloc ity of the airplane relative to the bird along the air plane s flight path is equal to Vc at sea level or 0 85Vg at 8 000 feet whichever is more critical 2 Uncontained fan blade impact 3 Uncontained engine failure or 4 Uncontained high energy rotating machin ery failure The damaged structure must be able to withstand the static loads considered as ultimate loads which are reas
252. cator for each engine 2 A cylinder head temperature indicator for each air cooled engine 3 A manifold pressure indicator for each en gine 4 A fuel pressure indicator to indicate the pressure at which the fuel is supplied for each engine 5 A fuel flowmeter or fuel mixture indicator for each engine without an automatic altitude mix ture control 6 A tachometer for each engine 7 A device that indicates to the flight crew during flight any change in the power output for each engine with i An automatic propeller feathering system whose operation is initiated by a power output measuring system or 115 825 1307 ii A total engine piston displacement of 2 000 cubic inches or more 8 A means to indicate to the pilot when the propeller is in reverse pitch for each reversing propeller c For turbine engine powered airplanes In addition to the powerplant instruments required by paragraph a of this section the following powerplant instruments are required 1 A gas temperature indicator for each en gine 2 A fuel flowmeter indicator for each engine 3 A tachometer to indicate the speed of the rotors with established limiting speeds for each engine 4 A means to indicate to the flight crew the operation of each engine starter that can be oper ated continuously but that is neither designed for continuous operation nor designed to prevent hazard if it failed 5 An indicato
253. ceeding the inert level The applicant must include any times when 233 Appendix N to Part 25 oxygen evolution from the fuel in the tank or com partment under evaluation would result in a flam mable fuel tank The oxygen evolution rate that must be used is defined in the Fuel Tank Flamma bility Assessment Method User s Manual dated May 2008 document number DOT FAA AR 05 8 incorporated by reference in 825 5 6 If an inerting system FRM is used the ef fects of any air that may enter the fuel tank follow ing the last flight of the day due to changes in am bient temperature as defined in Table 4 during a 12 hour overnight period e The applicant must submit to the FAA Over sight Office for approval the fuel tank flammability analysis including the airplane specific parame ters identified under paragraph N25 3 c of this appendix and any deviations from the parameters identified in paragraph N25 3 b of this appendix that affect flammability exposure substantiating data and any airworthiness limitations and other conditions assumed in the analysis N25 4 VARIABLES AND DATA TABLES The following data must be used when con ducting a flammability exposure analysis to deter mine the fleet average flammability exposure Variables used to calculate fleet flammability ex posure must include atmospheric ambient tem peratures flight length flammability exposure evaluation time fuel flash point thermal charac teristi
254. ceramic sheathed type K grounded thermocouples with a nominal 30 American wire gage AWG size con ductor The seven thermocouples must be at tached to a steel angle bracket to form a thermo couple rake for placement in the test stand during burner calibration as shown in Figure 3 of this part of this appendix 5 Apparatus Arrangement The test burner must be mounted on a suitable stand to position the exit of the burner cone a distance of 8 inches from the ceiling liner panel and 2 inches from the sidewall liner panel The burner stand should have the capability of allowing the burner to be swung away from the test specimen during warm up periods 6 Instrumentation recording potentiometer other suitable instrument with an appropriate range must be used to measure and record the outputs of the calorimeter and the thermocouples 7 Timing Device A stopwatch or other device must be used to measure the time of flame applica tion and the time of flame penetration if it occurs e Preparation of Apparatus Before calibra tion all equipment must be turned on and allowed to stabilize and the burner fuel flow must be ad justed as specified in paragraph d 2 f Calibration To ensure the proper thermal output of the burner the following test must be made 1 Remove the burner extension from the end of the draft tube Turn on the blower portion of the ASA Part 25 Airworthiness Standards Transport Category
255. cessibility provisions EWIS Protection of EWIS Flammable fluid fire protection EWIS Powerplants EWIS Flammable fluid shutoff means EWIS Instructions for Continued Airworthiness EWIS Powerplant and APU fire detector system EWIS Fire detector systems general EWIS Part 25 SFAR No 13 to Part 25 APPENDICES TO PART 25 Appendix A to Part 25 Appendix B to Part 25 Appendix C to Part 25 Appendix D to Part 25 Appendix E to Part 25 Appendix F to Part 25 Appendix G to Part 25 Continuous Gust Design Criteria Appendix H to Part 25 Instructions for Continued Airworthiness Appendix to Part 25 Installation of an Automatic Takeoff Thrust Control System ATTCS Appendix J to Part 25 Emergency Evacuation Appendix K to Part 25 Extended Operations ETOPS Appendix L to Part 25 HIRF Environments and Equipment HIRF Test Levels Appendix M to Part 25 Fuel Tank System Flammability Reduction Means Appendix N to Part 25 Fuel Tank Flammability Exposure and Reliability Analysis Authority 49 U S C 106 g 40113 44701 44702 and 44704 Source Docket No 5066 29 FR 18291 Dec 24 1964 unless otherwise noted SFAR No 13 TO PART 25 1 Applicability Contrary provisions of the Civil Air Regulations regarding certification notwith standing this regulation shall provide the basis for approval by the Administrator of modifications of individual Douglas DC 3 and Lockheed L 18 airpl
256. ching acceleration nose down is assumed to be reached concurrently with the positive maneuvering load factor points A to 02 825 333 b This negative pitching ac celeration must be equal to at least 26n n 1 5 Radians sec 2 where n is the positive load factor at the speed under consideration and V is the airplane equivalent speed in knots ASA Part 25 Airworthiness Standards Transport Category Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 46 43 FR 50594 Oct 30 1978 43 FR 52495 Nov 13 1978 43 FR 54082 Nov 20 1978 Amdt 25 72 55 FR 29775 July 20 1990 55 FR 37607 Sept 12 1990 Amdt 25 86 61 FR 5220 Feb 9 1996 Amdt 25 91 62 FR 40704 July 29 1997 FLAPS UP Cy MAX E RS e oe ee n4 O O o Q x 1 Cy MAX FLAPS UP 25 335 25 333 Flight maneuvering envelope a General The strength requirements must be met at each combination of airspeed and load factor on and within the boundaries of the repre sentative maneuvering envelope V n diagram of paragraph b of this section This envelope must also be used in determining the airplane struc tural operating limitations as specified 25 1501 b Maneuvering envelope Docket 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 86 61 FR 5220 Feb 9 1996 25 335 Design airspeeds
257. circuit breaker or re place a fuse is essential to safety in flight that cir cuit breaker or fuse must be located and identified So that it can be readily reset or replaced in flight Where fuses are used there must be spare fuses 123 825 1360 for use in flight equal to at least 5096 of the num ber of fuses of each rating required for complete circuit protection e Each circuit for essential loads must have individual circuit protection However individual protection for each circuit in an essential load sys tem such as each position light circuit in a sys tem is not required f For airplane systems for which the ability to remove or reset power during normal operations is necessary the system must be designed so that circuit breakers are not the primary means to remove or reset system power unless specifically designed for use as a switch g Automatic reset circuit breakers may be used as integral protectors for electrical equip ment such as thermal cut outs if there is circuit protection to protect the cable to the equipment Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 123 72 FR 63405 Nov 8 2007 825 1360 Precautions against injury a Shock The electrical system must be de signed to minimize risk of electric shock to crew passengers and servicing personnel and to maintenance personnel using normal precau tions b Burns The temperature of any part that may be handled
258. controls must be located on top of the pedes tal aft of the throttles centrally or to the right of the pedestal centerline and not less than 10 inches aft of the landing gear control 69 825 779 f The landing gear control must be located forward of the throttles and must be operable by each pilot when seated with seat belt and shoul der harness if provided fastened g Control knobs must be shaped in accor dance with 825 781 In addition the knobs must be of the same color and this color must contrast with the color of control knobs for other purposes and the surrounding cockpit h If a flight engineer is required as part of the minimum flight crew established under 825 1523 the airplane must have a flight engineer station lo cated and arranged so that the flight crewmembers can perform their functions efficiently and without interfering with each other Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50596 Oct 30 1978 825 779 Motion and effect of cockpit controls Cockpit controls must be designed so that they operate in accordance with the following move ment and actuation a Aerodynamic controls 1 Primary Controls Motion and effect Aileron Right clockwise for right wing down Elevator Rearward for nose up Rudder Right pedal forward for nose right 2 Secondary Controls Motion and effect Flaps orauxiliary Forward for fla
259. couple so that it extends 11 inches 279 mm out from the back of the cham ber wall 11 1 2 inches 292 mm from the right side of the chamber wall and is 2 inches 51 mm below the radiant panel The use of other thermo couples is optional 6 Calorimeter The calorimeter must be a one inch cylindrical water cooled total heat flux density foil type Gardon Gage that has a range of 0 to 5 BTU ft2 second 0 to 5 7 Watts cm 7 Calorimeter calibration specification and procedure i Calorimeter specification A Foil diameter must be 0 25 0 005 inches 6 35 0 13 mm B Foil thickness must be 0 0005 0 0001 inches 0 013 0 0025 mm C Foil material must be thermocouple grade Constantan D Temperature measurement must be a Cop per Constantan thermocouple E The copper center wire diameter must be 0 0005 inches 0 013 mm F The entire face of the calorimeter must be lightly coated with Black Velvet paint having an emissivity of 96 or greater ii Calorimeter calibration A The calibration method must be by compar ison to a like standardized transducer B The standardized transducer must meet the specifications given in paragraph VI b 6 of this appendix C Calibrate the standard transducer against a primary standard traceable to the National Insti tute of Standards and Technology NIST D The method of transfer must be a heated graphite plate 196 E The graphite plate must be electri
260. critical inoperative engine This must be shown at 1 3 for heading changes up to 15 degrees except that the heading change at which the rudder pedal force is 150 pounds need not be exceeded and with 1 The critical engine inoperative and its pro peller in the minimum drag position 2 The power required for level flight at 1 3 but not more than maximum continuous power 3 The most unfavorable center of gravity 4 Landing gear retracted 5 Flaps in the approach position and 6 Maximum landing weight b Directional control airplanes with four or more engines Airplanes with four or more en gines must meet the requirements of paragraph a of this section except that 1 The two critical engines must be inoperative with their propellers if applicable in the minimum drag position 2 Reserved 3 The flaps must be in the most favorable climb position c Lateral control general It must be possible to make 20 banked turns with and against the in operative engine from steady flight at a speed equal to 1 3 with 1 The critical engine inoperative and its pro peller if applicable in the minimum drag position 2 The remaining engines at maximum contin uous power 3 The most unfavorable center of gravity 4 Landing gear i retracted and ii extended 5 Flaps in the most favorable climb position and 6 Maximum takeoff weight d Lateral control airpl
261. cs of the fuel tank overnight temperature drop and oxygen evolution from the fuel into the ullage a Atmospheric Ambient Temperatures and Fuel Properties 1 In order to predict flammability exposure during a given flight the variation of ground ambi ent temperatures cruise ambient temperatures and a method to compute the transition from ground to cruise and back again must be used The variation of the ground and cruise ambient temperatures and the flash point of the fuel is de Federal Aviation Regulations fined by a Gaussian curve given by the 50 per cent value and a 1 standard deviation value 2 Ambient Temperature Under the program the ground and cruise ambient temperatures are linked by a set of assumptions on the atmo sphere The temperature varies with altitude fol lowing the International Standard Atmosphere ISA rate of change from the ground ambient temperature until the cruise temperature for the flight is reached Above this altitude the ambient temperature is fixed at the cruise ambient temper ature This results in a variation in the upper atmo spheric temperature For cold days an inversion is applied up to 10 000 feet and then the ISA rate of change is used 3 Fuel properties i For Jet A fuel the variation of flash point of the fuel is defined by a Gaussian curve given by the 50 percent value and a 1 standard deviation as shown in Table 1 of this appendix ii The flammability envelope
262. cted with not more than 10 percent ASA each of the airplane wheel brakes Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2321 Jan 16 1978 Amdt 25 92 63 FR 8318 Feb 18 1998 825 111 Takeoff path a The takeoff path extends from a standing start to a point in the takeoff at which the airplane is 1 500 feet above the takeoff surface or at which the transition from the takeoff to the en route con figuration is completed and Vpro is reached whichever point is higher In addition 1 The takeoff path must be based on the pro cedures prescribed in 25 101 f 2 The airplane must be accelerated on the ground to at which point the critical engine must be made inoperative and remain inoperative for the rest of the takeoff and 3 After reaching the airplane must be ac celerated to V2 b During the acceleration to speed V5 the nose gear may be raised off the ground at a speed not less than Vg However landing gear re traction may not be begun until the airplane is air borne c During the takeoff path determination in ac cordance with paragraphs a and b of this sec tion 1 The slope of the airborne part of the takeoff path must be positive at each point 2 The airplane must reach V2 before it is 35 feet above the takeoff surface and must continue at a speed as close as practical to but not less than V2 until it is 400 feet above the takeoff sur
263. currence ii A history of unscheduled engine removal rates since introduction into service using 6 and 12 month rolling averages with a summary of the major causes for the removals iii A list of all propulsion system events whether or not caused by maintenance or flight crew error including dispatch delays cancella tions aborted takeoffs turnbacks diversions and flights that continue to destination after the event 223 Appendix K to Part 25 iv The total number of engine hours and cy cles the number of hours for the engine with the highest number of hours the number of cycles for the engine with the highest number of cycles and the distribution of hours and cycles v The mean time between failures MTBF of propulsion system components that affect reliabil ity vi A history of the IFSD rates since introduc tion into service using a 12 month rolling average 2 The cause or potential cause of each item listed in K25 2 1 c 1 i must have a corrective action or actions that are shown to be effective in preventing future occurrences Each corrective action must be identified in the CMP document specified in section K25 1 6 A corrective action is not required i For an item where the manufacturer is un able to determine a cause or potential cause ii For an event where it is technically unfeasi ble to develop a corrective action iii If the world fleet IFSD rate A Is at or below 0 02 pe
264. d at sea level with a standard atmo sphere with a minimum practicable instrument calibration error when the corresponding pitot and static pressures are applied b Each system must be calibrated to deter mine the system error that is the relation be tween IAS and CAS in flight and during the ac celerated takeoff ground run The ground run cali bration must be determined 1 From 0 8 of the minimum value of V4 to the maximum value of V5 considering the approved ranges of altitude and weight and 2 With the flaps and power settings corre sponding to the values determined in the estab lishment of the takeoff path under 825 111 as suming that the critical engine fails at the mini mum value of V4 c The airspeed error of the installation ex cluding the airspeed indicator instrument calibra tion error may not exceed three percent or five knots whichever is greater throughout the speed range from 1 to 1 23 Van with flaps retracted and 2 1 23 Vsno to Vre with flaps in the landing position d From 1 23 to the speed at which stall warning begins the IAS must change perceptibly with CAS and in the same sense and at speeds below stall warning speed the IAS must not change in an incorrect sense e From Vyo to 2 3 Vpr the IAS must change perceptibly with CAS and in the same sense and at higher speeds up to the IAS must not change in an incorrect sense f The
265. d as the weights at which compliance is ASA Part 25 Airworthiness Standards Transport Category shown with the applicable provisions of this part including the takeoff climb provisions of 25 121 a through c for altitudes and ambient temperatures 2 The maximum landing weights must be es tablished as the weights at which compliance is shown with the applicable provisions of this part including the landing and approach climb provi sions of 25 119 and 25 121 d for altitudes and ambient temperatures 3 The minimum takeoff distances must be es tablished as the distances at which compliance is shown with the applicable provisions of this part including the provisions of 25 109 and 25 113 for weights altitudes temperatures wind compo nents runway surface conditions dry and wet and runway gradients for smooth hard surfaced runways Additionally at the option of the appli cant wet runway takeoff distances may be estab lished for runway surfaces that have been grooved or treated with a porous friction course and may be approved for use on runways where such surfaces have been designed constructed and maintained in a manner acceptable to the Ad ministrator b The extremes for variable factors such as altitude temperature wind and runway gradi ents are those at which compliance with the ap plicable provisions of this part is shown Docket No 5066 29 FR 18291 Dec 24 1964 as amended by
266. d by Amdt 25 38 41 FR 55466 Dec 20 1976 Amdt 25 92 63 FR 8318 Feb 18 1998 525 103 Stall speed a The reference stall speed is a cali brated airspeed defined by the applicant Vsg may not be less than a 1 g stall speed Vsp is ex pressed as Vcr MAX A Zw 2 ASA 825 105 where VcLyax Calibrated airspeed obtained when the load factor corrected lift coefficient za 45 is first a maximum during the maneuver prescribed in paragraph c of this section In addition when the maneuver is limited by a device that abruptly pushes the nose down at a selected angle of attack e g a stick pusher Vct Ax may not be less than the speed existing at the instant the device operates Nzw Load factor normal to the flight path at VcLmax W Airplane gross weight S Aerodynamic reference wing area and q Dynamic pressure b is determined with 1 Engines idling or if that resultant thrust causes an appreciable decrease in stall speed not more than zero thrust at the stall speed 2 Propeller pitch controls if applicable in the takeoff position 3 The airplane in other respects such as flaps landing gear and ice accretions in the con dition existing in the test or performance standard in which is being used 4 The weight used when Vsp is being used as a factor to determine compliance with a required performance standard 5 The center of grav
267. d by baffles or compartments b Flammability Exposure Evaluation Time FEET The time from the start of preparing the airplane for flight through the flight and landing until all payload is unloaded and all passengers and crew have disembarked In the Monte Carlo program the flight time is randomly selected from the Flight Length Distribution Table 2 the pre flight times are provided as a function of the flight time and the post flight time is a constant 30 min utes c Flammable With respect to a fluid or gas flammable means susceptible to igniting readily or to exploding 14 CFR Part 1 Definitions A non flammable ullage is one where the fuel air vapor is too lean or too rich to burn or is inert as defined below For the purposes of this appendix a fuel tank that is not inert is considered flammable when the bulk average fuel temperature within the tank is within the flammable range for the fuel type 232 Federal Aviation Regulations being used For any fuel tank that is subdivided into sections by baffles or compartments the tank is considered flammable when the bulk average fuel temperature within any section of the tank that is not inert is within the flammable range for the fuel type being used d Flash Point The flash point of a flammable fluid means the lowest temperature at which the application of a flame to a heated sample causes the vapor to ignite momentarily or flash Table 1 of this append
268. d flight operations 2 The complete system must be tested to de termine proper functional performance and rela tion to the other systems including simulation of relevant failure conditions and to support or vali date element design 3 The complete hydraulic system s must be functionally tested on the airplane in normal oper ation over the range of motion of all associated user systems The test must be conducted at the System relief pressure or 1 25 times the DOP if a System pressure relief device is not part of the sys tem design Clearances between hydraulic system elements and other systems or structural elements must remain adequate and there must be no detri mental effects Docket No 28617 66 FR 27402 May 16 2001 ASA Part 25 Airworthiness Standards Transport Category 825 1438 Pressurization and pneumatic systems a Pressurization system elements must be burst pressure tested to 2 0 times and proof pres sure tested to 1 5 times the maximum normal op erating pressure b Pneumatic system elements must be burst pressure tested to 3 0 times and proof pressure tested to 1 5 times the maximum normal operat ing pressure c An analysis or a combination of analysis and test may be substituted for any test required by paragraph a or b of this section if the Ad ministrator finds it equivalent to the required test Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 41 42 FR 3
269. data must be obtained from ground and flight test data or the approved FAA fuel management proce dures 3 Airplane cruise mach number 4 Airplane maximum range ASA Appendix N to Part 25 5 Fuel tank thermal characteristics If fuel temperature affects fuel tank flammability inputs to the Monte Carlo analysis must be provided that represent the actual bulk average fuel tempera ture within the fuel tank at each point in time throughout each of the flights being evaluated For fuel tanks that are subdivided by baffles or compartments bulk average fuel temperature in puts must be provided for each section of the tank Input values for these data must be obtained from ground and flight test data or a thermal model of the tank that has been validated by ground and flight test data 6 Maximum airplane operating temperature limit as defined by any limitations in the airplane flight manual 7 Airplane Utilization The applicant must pro vide data supporting the number of flights per day and the number of hours per flight for the specific airplane model under evaluation If there is no ex isting airplane fleet data to support the airplane being evaluated the applicant must provide sub stantiation that the number of flights per day and the number of hours per flight for that airplane model is consistent with the existing fleet data they propose to use d Fuel Tank FRM Model If FRM is used an FAA approved Monte Carlo p
270. ddition to the conditions specified in paragraphs a 1 and 2 of this section a limit engine torque corre sponding to takeoff power and propeller speed multiplied by a factor accounting for propeller con trol system malfunction including quick feather ing acting simultaneously with 1g level flight loads In the absence of a rational analysis a fac tor of 1 6 must be used b For turbine engine installations the engine mounts and supporting structure must be de signed to withstand each of the following 1 A limit engine torque load imposed by sud den engine stoppage due to malfunction or struc tural failure such as compressor jamming 2 A limit engine torque load imposed by the maximum acceleration of the engine c The limit engine torque to be considered un der paragraph a of this section must be obtained by multiplying mean torque for the specified power and speed by a factor of 1 1 25 for turbopropeller installations 2 1 33 for reciprocating engines with five or more cylinders or 3 Two three or four for engines with four three or two cylinders respectively Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 46 43 FR 50595 Oct 30 1978 Amdt 25 72 55 FR 29776 July 20 1990 825 363 Side load on engine and auxiliary power unit mounts a Each engine and auxiliary power unit mount and its supporting structure must be des
271. design operating or han dling characteristics 3 Any limitation procedure or other informa tion established as a condition of compliance with the applicable noise standards of part 36 of this chapter ASA Part 25 Airworthiness Standards Transport Category b Approved information Each part of the manual listed in 25 1583 through 25 1587 that is appropriate to the airplane must be furnished verified and approved and must be segregated identified and clearly distinguished from each un approved part of that manual c Reserved d Each Airplane Flight Manual must include a table of contents if the complexity of the manual indicates a need for it Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2323 Jan 16 1978 Amdt 25 72 55 FR 29786 July 20 1990 825 1583 Operating limitations a Airspeed limitations The following airspeed limitations and any other airspeed limitations nec essary for safe operation must be furnished 1 The maximum operating limit speed and a statement that this speed limit may not be deliberately exceeded in any regime of flight climb cruise or descent unless a higher speed is authorized for flight test or pilot training 2 If an airspeed limitation is based upon com pressibility effects a statement to this effect and information as to any symptoms the probable be havior of the airplane and the recommended re
272. designed as prescribed in this section to protect each occupant under those conditions b The structure must be designed to give each occupant every reasonable chance of es caping serious injury in a minor crash landing when 1 Proper use is made of seats belts and all other safety design provisions 2 The wheels are retracted where applica ble and 3 The occupant experiences the following ulti mate inertia forces acting separately relative to the surrounding structure i Upward 3 0g ii Forward 9 0g iii Sideward 3 0g on the airframe and 4 0g on the seats and their attachments iv Downward 6 0g v Rearward 1 5g c For equipment cargo in the passenger compartments and any other large masses the following apply 1 Except as provided in paragraph c 2 of this section these items must be positioned so that if they break loose they will be unlikely to i Cause direct injury to occupants ASA Part 25 Airworthiness Standards Transport Category ii Penetrate fuel tanks or lines or cause fire or explosion hazard by damage to adjacent sys tems or iii Nullify any of the escape facilities provided for use after an emergency landing 2 When such positioning is not practical e g fuselage mounted engines or auxiliary power units each such item of mass shall be restrained under all loads up to those specified in paragraph b 3 of this section The local attachments for these items
273. discomfort or fatigue and to provide reasonable passenger comfort For nor 86 Federal Aviation Regulations mal operating conditions the ventilation system must be designed to provide each occupant with an airflow containing at least 0 55 pounds of fresh air per minute b Crew and passenger compartment air must be free from harmful or hazardous concentrations of gases or vapors In meeting this requirement the following apply 1 Carbon monoxide concentrations in excess of 1 part in 20 000 parts of air are considered haz ardous For test purposes any acceptable carbon monoxide detection method may be used 2 Carbon dioxide concentration during flight must be shown not to exceed 0 5 percent by vol ume sea level equivalent in compartments nor mally occupied by passengers or crewmembers c There must be provisions made to ensure that the conditions prescribed in paragraph b of this section are met after reasonably probable fail ures or malfunctioning of the ventilating heating pressurization or other systems and equipment d If accumulation of hazardous quantities of smoke in the cockpit area is reasonably probable smoke evacuation must be readily accomplished starting with full pressurization and without de pressurizing beyond safe limits e Except as provided in paragraph f of this section means must be provided to enable the oc cupants of the following compartments and areas to control the temperatur
274. distance on a dry runway is the greater of 1 The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 35 feet above the takeoff surface determined under 825 111 for a dry run way or 2 115 percent of the horizontal distance along the takeoff path with all engines operating from the start of the takeoff to the point at which the air plane is 35 feet above the takeoff surface as de termined by a procedure consistent with 825 111 b Takeoff distance on a wet runway is the greater of 1 The takeoff distance on a dry runway deter mined in accordance with paragraph a of this section or 2 The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 15 feet above the takeoff surface achieved in a manner consistent with the achievement of V2 before reaching 35 feet above the takeoff surface determined under 825 111 for a wet runway c If the takeoff distance does not include a clearway the takeoff run is equal to the takeoff distance If the takeoff distance includes a clear 1 The takeoff run on a dry runway is the greater of i The horizontal distance along the takeoff path from the start of the takeoff to a point equi distant between the point at which is reached and the point at which the airplane is 35 feet above the takeoff surface as determined un der 825 111 for a
275. document then new or additional CMP requirements that the ap plicant has demonstrated would achieve this IFSD rate must be added to the CMP document 3 For type design approval beyond 180 min utes An IFSD rate of 0 01 or less per 1 000 fleet engine hours unless otherwise approved by the FAA If the airplane engine combination does not meet this rate by compliance with an existing 120 minute or 180 minute CMP document then new or additional CMP requirements that the applicant has demonstrated would achieve this IFSD rate must be added to the CMP document c Propulsion system assessment 1 The applicant must conduct a propulsion System assessment based on the following data collected from the world fleet of the airplane en gine combination i A list of all IFSDs unplanned ground engine shutdowns and occurrences both ground and in flight when an engine was not shut down but en gine control or the desired thrust or power level was not achieved including engine flameouts Planned IFSDs performed during flight training need not be included For each item the applicant must provide A Each airplane and engine make model and serial number B Engine configuration and major alteration history C Engine position D Circumstances leading up to the engine shutdown or occurrence E Phase of flight or ground operation F Weather and other environmental condi tions and G Cause of engine shutdown or oc
276. dry runway or ii 115 percent of the horizontal distance along the takeoff path with all engines operating from the start of the takeoff to a point equidistant be tween the point at which Vi or is reached and the point at which the airplane is 35 feet above the takeoff surface determined by a procedure con sistent with 825 111 2 The takeoff run on wet runway is the greater of i The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 15 feet above the takeoff surface achieved in a manner consistent with the achievement of V2 before reaching 35 feet above the takeoff surface as determined under 825 111 for a wet runway or ii 115 percent of the horizontal distance along the takeoff path with all engines operating from the start of the takeoff to a point equidistant be tween the point at which Vi or is reached and the point at which the airplane is 35 feet above the ASA Part 25 Airworthiness Standards Transport Category takeoff surface determined by a procedure con sistent with 825 111 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5671 April 8 1970 Amdt 25 92 63 FR 8320 Feb 18 1998 825 115 Takeoff flight path a The takeoff flight path shall be considered to begin 35 feet above the takeoff surface at the end of the takeoff distance determined in accordance with 25 113 a or b as appropriate for
277. ds in cluding the gyroscopic loads arising from the con ditions specified in 25 331 25 341 a 25 349 25 351 25 473 25 479 and 25 481 with the en gine or auxiliary power unit at maximum rpm ap propriate to the condition For the purposes of compliance with this section the pitch maneuver in 825 331 c 1 must be carried out until the pos itive limit maneuvering load factor point in 25 333 b is reached Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 86 61 FR 5222 Feb 9 1996 Amdt 25 91 62 FR 40704 July 29 1997 825 373 Speed control devices If speed control devices such as spoilers and drag flaps are installed for use in en route condi tions a The airplane must be designed for the sym metrical maneuvers prescribed in 825 333 and 825 337 the yawing maneuvers prescribed in 825 351 and the vertical and later gust conditions prescribed in 25 341 a at each setting and the maximum speed associated with that setting and b If the device has automatic operating or load limiting features the airplane must be de signed for the maneuver and gust conditions pre Scribed in paragraph a of this section at the speeds and corresponding device positions that the mechanism allows Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29776 July 20 1990 Amdt 25 86 61 FR 5222 Feb 9 1996 CONTROL SURFACE AND SYSTEM LOADS 525 391 Control surfac
278. e airplane 3 Fuel or fumes do not enter any parts of the airplane and 4 The jettisoning operation does not ad versely affect the controllability of the airplane e For reciprocating engine powered air planes means must be provided to prevent jetti soning the fuel in the tanks used for takeoff and landing below the level allowing 45 minutes flight at 75 percent maximum continuous power How ever if there is an auxiliary control independent of the main jettisoning control the system may be designed to jettison the remaining fuel by means of the auxiliary jettisoning control f For turbine engine powered airplanes means must be provided to prevent jettisoning the fuel in the tanks used for takeoff and landing be low the level allowing climb from sea level to 10 000 feet and thereafter allowing 45 minutes cruise at a speed for maximum range However if there is an auxiliary control independent of the main jettisoning control the system may be de signed to jettison the remaining fuel by means of the auxiliary jettisoning control g The fuel jettisoning valve must be designed to allow flight personnel to close the valve during any part of the jettisoning operation 103 825 1011 h Unless it is shown that using any means in cluding flaps slots and slats for changing the air flow across or around the wings does not ad versely affect fuel jettisoning there must be a placard adjacent to the jettisoning contr
279. e conducted susceptibility current must start at a minimum of 0 6 milliamperes mA at 10 kHz in creasing 20 decibels dB per frequency decade to a minimum of 30 mA at 500 kHz 2 From 500 kHz to 40 MHz the conducted susceptibility current must be at least 30 mA 3 From 40 MHz to 400 MHz use conducted susceptibility tests starting at a minimum of 30 mA at 40 MHz decreasing 20 dB per frequency decade to a minimum of 3 mA at 400 MHz 4 From 100 MHz to 400 MHz use radiated susceptibility tests at a minimum of 20 volts per meter V m peak with CW and 1 kHz square wave modulation with 90 percent depth or greater 5 From 400 MHz to 8 gigahertz GHz use ra diated susceptibility tests at a minimum of 150 V m peak with pulse modulation of 4 percent duty cycle with a 1 kHz pulse repetition frequency This signal must be switched on and off at a rate of 1 Hz with a duty cycle of 50 percent ASA Appendix L to Part 25 d Equipment HIRF Test Level 2 Equipment HIRF test level 2 is HIRF environment II in table of this appendix reduced by acceptable aircraft transfer function and attenuation curves Testing must cover the frequency band of 10 kHz to 8 GHz e Equipment HIRF Test Level 3 1 From 10 kHz to 400 MHz use conducted susceptibility tests starting at a minimum of 0 15 mA at 10 kHz increasing 20 dB per frequency de cade to a minimum of 7 5 mA at 500 kHz 2 From 500 kHz to 40 MHz use conducted susceptibili
280. e doors which includes all doors hatches open able windows access panels covers etc on the exterior of the fuselage that do not require the use of tools to open or close This also applies to each door or hatch through a pressure bulkhead in ASA cluding any bulkhead that is specifically designed to function as a secondary bulkhead under the prescribed failure conditions of part 25 These doors must meet the requirements of this section taking into account both pressurized and unpres surized flight and must be designed as follows 71 825 783 1 Each door must have means to safeguard against opening in flight as a result of mechanical failure or failure of any single structural element 2 Each door that could be a hazard if it un latches must be designed so that unlatching dur ing pressurized and unpressurized flight from the fully closed latched and locked condition is ex tremely improbable This must be shown by safety analysis 3 Each element of each door operating sys tem must be designed or where impracticable distinctively and permanently marked to minimize the probability of incorrect assembly and adjust ment that could result in a malfunction 4 All sources of power that could initiate un locking or unlatching of any door must be auto matically isolated from the latching and locking Systems prior to flight and it must not be possible to restore power to the door during flight 5 Each rem
281. e rocket engine usually stenciled on the engine casing or ii The rocket engine fuel has been expended or discharged b The currently approved maximum takeoff and landing weights at which an airplane is certif icated without a standby power rocket engine in stallation may be increased by an amount that does not exceed any of the following 1 An amount equal in pounds to 0 014 IN where is the maximum usable impulse in pounds seconds available from each standby power rocket engine and N is the number of rocket engines in stalled 2 An amount equal to 5 percent of the maxi mum certificated weight approved in accordance with the applicable airworthiness regulations with out standby power rocket engines installed 3 An amount equal to the weight of the rocket engine installation 4 An amount that together with the currently approved maximum weight would equal the max imum structural weight established for the air plane without standby rocket engines installed II PERFORMANCE CREDIT FOR TRANSPORT CATEGORY AIRPLANES EQUIPPED WITH STANDBY POWER The Administrator may grant performance credit for the use of standby power on transport category airplanes However the performance credit applies only to the maximum certificated takeoff and landing weights the takeoff distance and the takeoff paths and may not exceed that found by the Administrator to result in an overall 164 Federal Aviation Regulations
282. e sufficiently greater than the speed recommended for the operation of the device to allow for proba ble variations in speed control For drag devices intended for use in high speed descents Vpp may not be less than Vp When an automatic drag de vice positioning or load limiting means is used the speeds and corresponding drag device posi tions programmed or allowed by the automatic means must be used for design Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 86 61 FR 5220 Feb 9 1996 25 91 62 FR 40704 July 29 1997 825 337 Limit maneuvering load factors a Except where limited by maximum static lift coefficients the airplane is assumed to be sub jected to symmetrical maneuvers resulting in the limit maneuvering load factors prescribed in this section Pitching velocities appropriate to the cor responding pull up and steady turn maneuvers must be taken into account b The positive limit maneuvering load factor n for any speed up to may not be less than 2 1 24 000 W 10 000 except that n may not be less than 2 5 and need not be greater than 3 8 where W is the design maximum takeoff weight c The negative limit maneuvering load fac tor 1 May not be less than 1 0 at speeds up to Vc and ASA Part 25 Airworthiness Standards Transport Category 2 Must vary linearly with speed from the value at Vc to zero at Vp d
283. e a stopwatch or other device accurate to 1 to measure the time of application of the burner flame and burnthrough time 8 Test chamber Perform tests in a suitable chamber to reduce or eliminate the possibility of test fluctuation due to air movement The cham ber must have a minimum floor area of 10 by 10 feet 305 by 305 cm i Ventilation hood Provide the test chamber with an exhaust system capable of removing the products of combustion expelled during tests c Test Specimens 1 Specimen preparation Prepare a minimum of three specimen sets of the same construction and configuration for testing 2 Insulation blanket test specimen i For batt type materials such as fiberglass the constructed finished blanket specimen as semblies must be 32 inches wide by 36 inches long 81 3 by 91 4 cm exclusive of heat sealed film edges ii For rigid and other non conforming types of insulation materials the finished test specimens 200 Federal Aviation Regulations must fit into the test rig in such a manner as to replicate the actual in service installation 3 Construction Make each of the specimens tested using the principal components i e insu lation fire barrier material if used and moisture barrier film and assembly processes representa tive seams and closures i Fire barrier material If the insulation blanket is constructed with a fire barrier material place the fire barrier material in a man
284. e altitude of 8 000 feet with a respiratory minute volume of 30 liters per minute BTPD If a demand oxygen system is used a supply of 300 liters of free oxygen at 70 F and 760 mm Hg pressure is considered to be of 15 minute dura tion at the prescribed altitude and minute volume If a continuous flow protective breathing system is used including a mask with a standard re breather bag a flow rate of 60 liters per minute at 8 000 feet 45 liters per minute at sea level and a supply of 600 liters of free oxygen at 70 F and 760 mm Hg pressure is considered to be of 15 minute duration at the prescribed altitude and minute volume BTPD refers to body temperature conditions that is 37 at ambient pressure dry 6 The equipment must meet the requirements of paragraphs b and c of 825 1441 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55468 Dec 20 1976 825 1441 Oxygen equipment and supply a If certification with supplemental oxygen equipment is requested the equipment must meet the requirements of this section and 25 1443 through 25 1453 b The oxygen system must be free from haz ards in itself in its method of operation and in its effect upon other components c There must be a means to allow the crew to readily determine during flight the quantity of ox ygen available in each source of supply d The oxygen flow rate and the oxygen equip ment for airplanes for
285. e and quantity of ventilat ing air supplied to their compartment or area inde pendently of the temperature and quantity of air supplied to other compartments and areas 1 The flight crew compartment 2 Crewmember compartments and areas other than the flight crew compartment unless the crewmember compartment or area is ventilated by air interchange with other compartments or areas under all operating conditions f Means to enable the flight crew to control the temperature and quantity of ventilating air supplied to the flight crew compartment indepen dently of the temperature and quantity of ventilat ing air supplied to other compartments are not re quired if all of the following conditions are met 1 The total volume of the flight crew and pas senger compartments is 800 cubic feet or less 2 The air inlets and passages for air to flow between flight crew and passenger compart ments are arranged to provide compartment tem peratures within 5 degrees F of each other and adequate ventilation to occupants in both com partments 3 The temperature and ventilation controls are accessible to the flight crew ASA Part 25 Airworthiness Standards Transport Category g The exposure time at any given temperature must not exceed the values shown in the following graph after any improbable failure condition Humidity 2700 N m 27 mbar Vapor Pressure 825 841 Degrees F 150 65 60 55 TEMPERATURE 50
286. e com ponents that would adversely affect engine opera tion or cause a serious loss of power or thrust i Under the icing conditions specified in Ap pendix C and ii In falling and blowing snow within the limita tions established for the airplane for such opera tion 2 Each turbine engine must idle for 30 min utes on the ground with the air bleed available for engine icing protection at its critical condition without adverse effect in an atmosphere that is at a temperature between 15 and 30 F between 9 and 1 C and has a liquid water content not less than 0 3 grams per cubic meter in the form of drops having a mean effective diameter not less than 20 microns followed by momentary opera tion at takeoff power or thrust During the 30 min utes of idle operation the engine may be run up periodically to a moderate power or thrust setting in a manner acceptable to the Administrator c Supercharged reciprocating engines For each engine having a supercharger to pressurize the air before it enters the carburetor the heat rise in the air caused by that supercharging at any alti tude may be utilized in determining compliance with paragraph a of this section if the heat rise utilized is that which will be available automati cally for the applicable altitude and operating con dition because of supercharging Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 Amd
287. e current mode of operation including any armed modes transitions and reversions Selec tor switch position is not an acceptable means of indication The controls and indications must be grouped and presented in a logical and consistent manner The indications must be visible to each pilot under all expected lighting conditions j Following disengagement of the autopilot a warning visual and auditory must be provided to each pilot and be timely and distinct from all other cockpit warnings Following disengagement of the autothrust function a caution must be provided to each pilot I The autopilot may not create a potential haz ard when the flightcrew applies an override force to the flight controls m During autothrust operation it must be possible for the flightcrew to move the thrust le vers without requiring excessive force The auto thrust may not create a potential hazard when the flightcrew applies an override force to the thrust levers n For purposes of this section a transient is a disturbance in the control or flight path of the air plane that is not consistent with response to flight crew inputs or environmental conditions 1 A minor transient would not significantly re duce safety margins and would involve flightcrew actions that are well within their capabilities A mi nor transient may involve a slight increase in flightcrew workload or some physical discomfort to passengers or cabin crew 2
288. e ef fects of increased dynamic pressure Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5671 April 8 1970 Amdt 25 54 45 FR 60172 Sept 11 1980 Amdt 25 72 55 FR 29775 July 20 1990 Amdt 25 84 60 FR 30750 June 9 1995 Amdt 25 121 72 FR 44668 Aug 8 2007 825 255 Out of trim characteristics a From an initial condition with the airplane trimmed at cruise speeds up to Vyo Myo the air plane must have satisfactory maneuvering stabil ity and controllability with the degree of out of trim in both the airplane nose up and nose down di rections which results from the greater of 1 A three second movement of the longitudi nal trim system at its normal rate for the particular flight condition with no aerodynamic load or an equivalent degree of trim for airplanes that do not have a power operated trim system except as limited by stops in the trim system including those required by 25 655 b for adjustable stabi lizers or 2 The maximum mistrim that can be sus tained by the autopilot while maintaining level flight in the high speed cruising condition b In the out of trim condition specified in paragraph a of this section when the normal ac celeration is varied from 1 g to the positive and negative values specified in paragraph c of this section 1 The stick force vs g curve must have a pos itive slope at any speed up to and including Vrc Mrc and 2
289. e flight conditions for example configuration speed an gle of attack and altitude The ice accretions for each flight phase are defined as follows 1 Takeoff ice is the most critical ice accretion on unprotected surfaces and any ice accretion on the protected surfaces appropriate to normal ice protection system operation occurring between liftoff and 400 feet above the takeoff surface as suming accretion starts at liftoff in the takeoff maximum icing conditions of part paragraph of this appendix 2 Final takeoff ice is the most critical ice ac cretion on unprotected surfaces and any ice ac cretion on the protected surfaces appropriate to normal ice protection system operation between 400 feet and either 1 500 feet above the takeoff surface or the height at which the transition from the takeoff to the en route configuration is com pleted and Vero is reached whichever is higher Ice accretion is assumed to start at liftoff in the takeoff maximum icing conditions of part para graph c of this appendix 3 En route ice is the critical ice accretion on the unprotected surfaces and any ice accretion on the protected surfaces appropriate to normal ice protection system operation during the en route phase 4 Holding ice is the critical ice accretion on the unprotected surfaces and any ice accretion on the protected surfaces appropriate to normal ice protection system operation during the hold ing flight pha
290. e heater system un der any operating 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 25 1121 and 25 1123 In addition there must be provisions in the design of the heater exhaust system to safely expel the products of combustion to prevent the occur rence of 1 Fuel leakage from the exhaust to surround ing compartments 2 Exhaust gas impingement on surrounding equipment or structure 3 Ignition of flammable fluids by the exhaust if the exhaust is in a compartment containing flammable fluid lines and 4 Restriction by the exhaust of the prompt re lief of backfires that if so restricted could cause heater failure h Heater fuel systems Each heater fuel sys tem must meet each powerplant fuel system re quirement affecting safe heater operation Each heater fuel system component within the ventilat ing airstream must be protected by shrouds so that no leakage from those components can enter the ventilating airstream i Drains There must be means to safely drain fuel that might accumulate within the combustion chamber or the heat exchanger In addition 92 Federal Aviation Regulations 1 Each part of any drain that operates at high temperatures must be protected in the same manner as heater exhausts and 2 Each drain must be protected from hazard ous ice accumulati
291. e loads general The control surfaces must be designed for the limit loads resulting from the flight conditions in 25 331 25 341 a 25 349 and 25 351 and the ground gust conditions in 825 415 considering the requirements for a Loads parallel to hinge line in 825 393 b Pilot effort effects in 825 397 c Trim tab effects in 825 407 d Unsymmetrical loads in 825 427 and e Auxiliary aerodynamic surfaces in 825 445 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 86 61 FR 5222 Feb 9 1996 825 393 Loads parallel to hinge line a Control surfaces and supporting hinge brackets must be designed for inertia loads acting parallel to the hinge line b In the absence of more rational data the in ertia loads may be assumed to be equal to KW where ASA 25 397 1 K 24 for vertical surfaces 2 K 12 for horizontal surfaces and 3 W weight of the movable surfaces 25 395 Control system a Longitudinal lateral directional and drag control system and their supporting structures must be designed for loads corresponding to 125 percent of the computed hinge moments of the movable control surface in the conditions pre scribed in 25 391 b The system limit loads except the loads re sulting from ground gusts need not exceed the loads that can be produced by the pilot or pilots and by automatic or power devices operating the controls c The loads
292. e no 1 Spray characteristics that would impair the pilot s view cause damage or result in the taking in of an undue quantity of water 2 Dangerously uncontrollable porpoising bounding or swinging tendency or 3 Immersion of auxiliary floats or sponsons wing tips propeller blades or other parts not de signed to withstand the resulting water loads b Compliance with the requirements of para graph a of this section must be shown 32 Federal Aviation Regulations 1 In water conditions from smooth to the most adverse condition established in accor dance with 825 231 2 In wind and cross wind velocities water cur rents and associated waves and swells that may reasonably be expected in operation on water 3 At speeds that may reasonably be expected in operation on water 4 With sudden failure of the critical engine at any time while on water and 5 At each weight and center of gravity posi tion relevant to each operating condition within the range of loading conditions for which certifica tion is requested c In the water conditions of paragraph b of this section and in the corresponding wind condi tions the seaplane or amphibian must be able to drift for five minutes with engines inoperative aided if necessary by a sea anchor MISCELLANEOUS FLIGHT REQUIREMENTS 825 251 Vibration and buffeting a The airplane must be demonstrated in flight to be free from any vibration and buff
293. e of a fluid system must be identified and defined Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 Amdt 25 46 43 FR 50597 Oct 30 1978 25 865 Fire protection of flight controls engine mounts and other flight structure Essential flight controls engine mounts and other flight structures located in designated fire zones or in adjacent areas which would be sub jected to the effects of fire in the fire zone must be constructed of fireproof material or shielded so that they are capable of withstanding the effects of fire Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 ASA Part 25 Airworthiness Standards Transport Category 825 867 Fire protection other components a Surfaces to the rear of the nacelles within one nacelle diameter of the nacelle centerline must be at least fire resistant b Paragraph a of this section does not apply to tail surfaces to the rear of the nacelles that could not be readily affected by heat flames or sparks coming from a designated fire zone or en gine compartment of any nacelle Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 825 869 Fire protection systems a Electrical system components 1 Components of the electrical system must meet the applicable fire and smoke protection re quirements of 25 831 c
294. e of the airplane meeting at least the dimen sions of a Type IV exit ASA Part 25 Airworthiness Standards Transport Category 2 For airplanes that have a passenger seating configuration of 10 of more seats excluding pilot seats one exit above the waterline in a side of the airplane meeting at least the dimensions of a Type Ill exit for each unit or part of a unit of 35 passenger seats but no less than two such exits in the passenger cabin with one on each side of the airplane The passenger seat exit ratio may be increased through the use of larger exits or other means provided it is shown that the evacuation capability during ditching has been improved ac cordingly 3 If it is impractical to locate side exits above the waterline the side exits must be replaced by an equal number of readily accessible overhead hatches of not less than the dimensions of a Type Ill exit except that for airplanes with a passenger configuration of 35 or fewer seats excluding pilot seats the two required Type III side exits need be replaced by only one overhead hatch j Flightcrew emergency exits For airplanes in which the proximity of passenger emergency exits to the flightcrew area does not offer a convenient and readily accessible means of evacuation of the flightcrew and for all airplanes having a passenger seating capacity greater than 20 flightcrew exits shall be located in the flightcrew area Such exits shall be of sufficient
295. e particular flight condition d Unless otherwise prescribed the applicant must select the takeoff en route approach and landing configurations for the airplane e The airplane configurations may vary with weight altitude and temperature to the extent they are compatible with the operating proce dures required by paragraph f of this section f Unless otherwise prescribed in determining the accelerate stop distances takeoff flight paths takeoff distances and landing distances changes in the airplane s configuration speed power and thrust must be made in accordance with proce dures established by the applicant for operation in service g Procedures for the execution of balked landings and missed approaches associated with the conditions prescribed in 25 119 and 25 121 d must be established h The procedures established under para graphs f and g of this section must 1 Be able to be consistently executed in ser vice by crews of average skill 2 Use methods or devices that are safe and reliable and 3 Include allowance for any time delays in the execution of the procedures that may reasonably be expected in service i The accelerate stop and landing distances prescribed in 25 109 and 25 125 respectively must be determined with all the airplane wheel brake assemblies at the fully worn limit of their al lowable wear range Docket No 5066 29 FR 18291 Dec 24 1964 as amende
296. e proper flow and pressure for fuel injection when the injection is not accomplished in a carburetor approved as part of the engine b Emergency pumps There must be emer gency pumps or another main pump to feed each engine immediately after failure of any main pump other than a fuel injection pump approved as part of the engine 25 993 Fuel system lines and fittings a Each fuel line must be installed and sup ported to prevent excessive vibration and to with stand loads due to fuel pressure and accelerated flight conditions b Each fuel line connected to components of the airplane between which relative motion could exist must have provisions for flexibility c Each flexible connection in fuel lines that may be under pressure and subjected to axial loading must use flexible hose assemblies d Flexible hose must be approved or must be shown to be suitable for the particular application e No flexible hose that might be adversely af fected by exposure to high temperatures may be used where excessive temperatures will exist dur ing operation or after engine shut down f Each fuel line within the fuselage must be de signed and installed to allow a reasonable degree of deformation and stretching without leakage Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 15 32 FR 13266 Sept 20 1967 25 994 Fuel system components Fuel system components in an engine nacelle or in the fuselage m
297. e require ments of 25 101 through 25 125 b There must be a propeller speed limiting means at the governor It must limit the maximum possible governed engine speed to a value not exceeding the maximum allowable r p m c The means used to limit the low pitch posi tion of the propeller blades must be set so that the engine does not exceed 103 percent of the maxi mum allowable engine rpm or 99 percent of an approved maximum overspeed whichever is greater with 1 The propeller blades at the low pitch limit and governor inoperative 14 Federal Aviation Regulations 2 The airplane stationary under standard at mospheric conditions with no wind and 3 The engines operating at the takeoff mani fold pressure limit for reciprocating engine pow ered airplanes or the maximum takeoff torque limit for turbopropeller engine powered airplanes Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 57 49 FR 6848 Feb 23 1984 Amdt 25 72 55 FR 29774 July 20 1990 PERFORMANCE 25 101 General a Unless otherwise prescribed airplanes must meet the applicable performance require ments of this subpart for ambient atmospheric conditions and still air b The performance as affected by engine power or thrust must be based on the following relative humidities 1 For turbine engine powered airplanes a rel ative humidity of i 80 percent at and below standard tempera tures and ii 34
298. e total strap ten sion loads must not exceed 2 000 pounds 2 The maximum compressive load measured between the pelvis and the lumbar column of the anthropomorphic dummy must not exceed 1 500 pounds 3 The upper torso restraint straps where in stalled must remain on the occupant s shoulder during the impact 4 The lap safety belt must remain on the oc cupant s pelvis during the impact 5 Each occupant must be protected from seri ous head injury under the conditions prescribed in paragraph b of this section Where head contact with seats or other structure can occur protection must be provided so that the head impact does not exceed a Head Injury Criterion HIC of 1 000 units The level of HIC is defined by the equation 2 5 HIC festo etn Where t4 is the initial integration time to is the final integration time and a t is the total acceleration vs time curve for the head strike and where t is in seconds and a is in units of gravity g 6 Where leg injuries may result from contact with seats or other structure protection must be provided to prevent axially compressive loads ex ceeding 2 250 pounds in each femur 7 The seat must remain attached at all points of attachment although the structure may have yielded 8 Seats must not yield under the tests speci fied in paragraphs b 1 and b 2 of this section to the extent they would impede rapid evacuation of the airplane occupants
299. e windshield and window panels over an area which would provide the visibility specified in paragraph a of this sec tion under all internal and external ambient condi tions including precipitation conditions in which the airplane is intended to be operated d Fixed markers or other guides must be in stalled at each pilot station to enable the pilots to position themselves in their seats for an optimum combination of outside visibility and instrument scan If lighted markers or guides are used they must comply with the requirements specified in 825 1381 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 Amdt 25 46 43 FR 50595 Oct 30 1978 Amdt 25 72 55 FR 29778 July 20 1990 Amdt 25 108 67 FR 70828 Nov 26 2002 Amdt 25 121 72 FR 44669 Aug 8 2007 825 775 Windshields and windows a Internal panes must be made of nonsplin tering material b Windshield panes directly in front of the pilots in the normal conduct of their duties and the supporting structures for these panes must withstand without penetration the impact of a four pound bird when the velocity of the airplane relative to the bird along the airplane s flight path is equal to the value of Vc at sea level selected under 25 335 a c Unless it can be shown by analysis or tests that the probability of occurrence of a critical windshield fragmentation condition is of a low or der the
300. e with an open ing 6 inches 152 mm high and 11 inches 280 mm wide as shown in Figure of Part II of this appendix 176 Federal Aviation Regulations iv Have a burner fuel pressure regulator that is adjusted to deliver a nominal 2 0 GPH of 2 Grade kerosene or equivalent Burner models which have been used success fully in testing are the Lenox Model OB 32 Carlin Model 200 CRD and Park Model DPL The basic burner is described in FAA Powerplant Engineer ing Report No 3A Standard Fire Test Apparatus and Procedure for Flexible Hose Assemblies dated March 1978 however the test settings specified in this appendix differ in some instances from those specified in the report 3 Calorimeter i The calorimeter to be used in testing must be a total heat flux Foil Type Gardon Gage of an ap propriate range approximately O to 15 0 British thermal unit BTU per 42 sec 0 17 0 watts cm The calorimeter must be mounted in a 6 inch by 12 inch 152 by 305 mm by 34 inch 19 mm thick in sulating block which is attached to a steel angle bracket for placement in the test stand during burner calibration as shown in Figure 2 of this part of this appendix ii The insulating block must be monitored for deterioration and the mounting shimmed as nec essary to ensure that the calorimeter face is par allel to the exit plane of the test burner cone 4 Thermocouples The seven thermocouples to be used for testing must be 146 inch
301. ea of the aircraft must 1 Be of a kind and design appropriate to its in tended function 2 Be installed according to limitations speci fied for the EWIS components 3 Perform the function for which it was in tended without degrading the airworthiness of the airplane 4 Be designed and installed in a way that will minimize mechanical strain b Selection of wires must take into account known characteristics of the wire in relation to each installation and application to minimize the risk of wire damage including any arc tracking phenomena c The design and installation of the main power cables including generator cables in the fuselage must allow for a reasonable degree of deformation and stretching without failure d EWIS components located in areas of known moisture accumulation must be protected to minimize any hazardous effects due to mois ture 825 1705 Systems and functions EWIS a EWIS associated with any system required for type certification or by operating rules must be considered an integral part of that system and must be considered in showing compliance with the applicable requirements for that system b For systems to which the following rules ap ply the components of EWIS associated with those systems must be considered an integral part of that system or systems and must be con sidered in showing compliance with the applicable requirements for that system 1 825 773 b 2 Pilot compa
302. ead within range repeat steps in paragraphs 1 through 4 and adjust the burner air intake damper until the proper readings are obtained The ther mocouple rake and the calorimeter should be used frequently to maintain and record calibrated test parameters Until the specific apparatus has demonstrated consistency each test should be calibrated After consistency has been confirmed Several tests may be conducted with the pre test calibration before and a calibration check after the series g Test Procedure The flammability of each set of specimens must be tested as follows 1 Record the weight of each set of seat bot tom and seat back cushion specimens to be tested to the nearest 0 02 pound 9 grams 2 Mount the seat bottom and seat back cush ion test specimens on the test stand as shown in Figure 2 securing the seat back cushion speci men to the test stand at the top 3 Swing the burner into position and ensure that the distance from the exit of the burner cone to the side of the seat bottom cushion specimen is 4 1 inches 102 3 mm 4 Swing the burner away from the test posi tion Turn on the burner and allow it to run for 2 minutes to provide adequate warmup of the burner cone and flame stabilization 5 To begin the test swing the burner into the test position and simultaneously start the timing device 6 Expose the seat bottom cushion specimen to the burner flame for 2 minutes and then turn off the burner
303. east 2 nautical miles at night under clear atmospheric conditions and 2 Show the maximum unbroken light practica ble when the airplane is moored or drifting on the water b Externally hung lights may be used 825 1401 Anticollision light system a General The airplane must have an anticol lision light system that 1 Consists of one or more approved anticolli sion lights located so that their light will not impair the crew s vision or detract from the conspicuity of the position lights and ASA Part 25 Airworthiness Standards Transport Category 2 Meets the requirements of paragraphs b through f of this section b Field of coverage The system must consist of enough lights to illuminate the vital areas around the airplane considering the physical con figuration and flight characteristics of the airplane The field of coverage must extend in each direc tion within at least 75 degrees above and 75 de grees below the horizontal plane of the airplane except that a solid angle or angles of obstructed visibility totaling not more than 0 03 steradians is allowable within a solid angle equal to 0 15 stera dians centered about the longitudinal axis in the rearward direction c Flashing characteristics The arrangement of the system that is the number of light sources beam width speed of rotation and other charac teristics must give an effective flash frequency of not less than 40 nor more than 100 c
304. eathering systems are used for this purpose the feathering lines must be at least fire resistant under the operating conditions that may be expected to exist during feathering d Turbine engine installations For turbine en gine installations 1 Design precautions must be taken to mini mize the hazards to the airplane in the event of an engine rotor failure or of a fire originating within the engine which burns through the engine case 2 The powerplant systems associated with engine control devices systems and instrumen tation must be designed to give reasonable as surance that those engine operating limitations that adversely affect turbine rotor structural integ rity will not be exceeded in service e Restart capability 1 Means to restart any engine in flight must be provided 2 An altitude and airspeed envelope must be established for in flight engine restarting and each engine must have a restart capability within that envelope 3 For turbine engine powered airplanes if the minimum windmilling speed of the engines fol lowing the inflight shutdown of all engines is in sufficient to provide the necessary electrical power for engine ignition a power source inde pendent of the engine driven electrical power generating system must be provided to permit in flight engine ignition for restarting f Auxiliary Power Unit Each auxiliary power unit must be approved or meet the requirements of the category
305. ection against static electricity EWIS a EWIS components used for electrical bond ing and protection against static electricity must meet the requirements of 825 899 b On airplanes having grounded electrical Systems electrical bonding provided by EWIS components must provide an electrical return path capable of carrying both normal and fault currents without creating a shock hazard or dam ASA 825 1725 age to the EWIS components other airplane sys tem components or airplane structure 825 1717 Circuit protective devices EWIS Electrical wires and cables must be designed and installed so they are compatible with the cir cuit protection devices required by 825 1357 so that a fire or smoke hazard cannot be created un der temporary or continuous fault conditions 825 1719 Accessibility provisions EWIS Access must be provided to allow inspection and replacement of any EWIS component as nec essary for continued airworthiness 825 1721 Protection of EWIS a No cargo or baggage compartment may contain any EWIS whose damage or failure may affect safe operation unless the EWIS is pro tected so that 1 It cannot be damaged by movement of cargo or baggage in the compartment 2 Its breakage or failure will not create a fire hazard b EWIS must be designed and installed to minimize damage and risk of damage to EWIS by movement of people in the airplane during all phases of flight maintenance and se
306. ed 1 Description of the specimen 2 Radiant heat flux to the specimen ex pressed in W cm 3 Data giving release rates of heat in kW m as a function of time either graphically or tabu lated at intervals no greater than 10 seconds The calibration factor must be recorded 184 Federal Aviation Regulations 4 If melting sagging delaminating or other behavior that affects the exposed surface area or the mode of burning occurs these behaviors must be reported together with the time at which such behaviors were observed 5 The peak heat release and the 2 minute in tegrated heat release rate must be reported ASA Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 5 20 1 in 132 2 mm Thermocouple placement in chimney 1 2 in 30 mm from center on diagonal Thermocouple on 0 02 in level 10 0 5 mm 7 0 0 1 in 178 3 mm Baffle 1 00 0 03 in x 3 00 0 03 in 25 1mm x 76 1 mm 7 0 018 0 002 in Chimney 0 018 0 002 in 0 38 0 05 mm thick 0 38 0 05 mm thick Outer cone 0 031 0 002 in 0 79 0 05 mm thick lnrier Gone 12 20 0 25 in 0 018 0 002 in 310 6mm 0 38 0 05 mm thick 13 0 0 25 in Air manifold 330 6mm 2 with 48 no 26 holes Y LY 0 75 0 02 in 19 0 0 5 lt gt 1 5 in 38mm nominal ingi 5 inside dd 0 002 diameter
307. ed from the airplane structure except for attachments Such panels must be subjected to the 45 degree angle test The flame may not penetrate pass through the material during application of the flame or subsequent to its removal The average flame time after removal of the flame source may not ex ceed 15 seconds and the average glow time may not exceed 10 seconds iv Insulation blankets and covers used to pro tect cargo must be constructed of materials that meet the requirements of paragraph a 1 ii of part of this appendix Tiedown equipment in cluding containers bins and pallets used in each cargo and baggage compartment must be con structed of materials that meet the requirements of paragraph a 1 v of part of this appendix 3 Electrical system components Insulation on electrical wire or cable installed in any area of the fuselage must be self extinguishing when subjected to the 60 degree test specified in part I of this appendix The average burn length may not exceed 3 inches and the average flame time after removal of the flame source may not exceed 30 seconds Drippings from the test specimen may not continue to flame for more than an average of 3 seconds after falling ASA Part 25 Airworthiness Standards Transport Category b Test Procedures 1 Conditioning Specimens must be condi tioned to 70 5 F and at 50 percent 5 percent relative humidity until moisture equilibrium is reached or for
308. ed in accordance with part 35 of this chapter and the vibration data obtained from compliance with paragraph a of this section For the purpose of this paragraph the propeller includes the hub blades blade retention component and any other propeller component whose failure due to fatigue could be catastrophic to the airplane This evalua tion must include 1 The intended loading spectra including all reasonably foreseeable propeller vibration and cyclic load patterns identified emergency condi tions allowable overspeeds and overtorques and the effects of temperatures and humidity ex pected in service 2 The effects of airplane and propeller operat ing and airworthiness limitations Docket No FAA 2007 27310 73 FR 63345 Oct 24 2008 825 925 Propeller clearance Unless smaller clearances are substantiated propeller clearances with the airplane at maxi mum weight with the most adverse center of gravity and with the propeller in the most adverse pitch position may not be less than the following a Ground clearance There must be a clear ance of at least seven inches for each airplane with nose wheel landing gear or nine inches for each airplane with tail wheel landing gear be tween each propeller and the ground with the landing gear statically deflected and in the level takeoff or taxiing attitude whichever is most criti cal In addition there must be positive clearance between the propeller and the gr
309. ed that is nec essary for each operating condition and attitude for which compliance with this section is shown and the appropriate emergency pump must be substituted for each main pump so used 4 If there is a fuel flowmeter it must be blocked and the fuel must flow through the meter or its bypass b If an engine can be supplied with fuel from more than one tank the fuel system must 1 For each reciprocating engine supply the full fuel pressure to that engine in not more than 20 seconds after switching to any other fuel tank containing usable fuel when engine malfunction ing becomes apparent due to the depletion of the fuel supply in any tank from which the engine can be fed and 2 For each turbine engine in addition to hav ing appropriate manual switching capability be designed to prevent interruption of fuel flow to that engine without attention by the flight crew when any tank supplying fuel to that engine is depleted of usable fuel during normal operation and any other tank that normally supplies fuel to that en gine alone contains usable fuel Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6912 May 5 1967 825 957 Flow between interconnected tanks If fuel can be pumped from one tank to another in flight the fuel tank vents and the fuel transfer 98 Federal Aviation Regulations System must be designed so that no structural damage to the tanks can occur because of
310. ed with a sin gle wheel and tire assembly the wheel must be fitted with a suitable tire of proper fit with a speed rating approved by the Administrator that is not exceeded under critical conditions and with a load rating approved by the Administrator that is not exceeded under 1 The loads on the main wheel tire corre sponding to the most critical combination of air plane weight up to maximum weight and center of gravity position and 2 The loads corresponding to the ground re actions in paragraph b of this section on the nose wheel tire except as provided in paragraphs b 2 and b 3 of this section b The applicable ground reactions for nose wheel tires are as follows 1 The static ground reaction for the tire corre sponding to the most critical combination of air plane weight up to maximum ramp weight and center of gravity position with a force of 1 0g act ing downward at the center of gravity This load may not exceed the load rating of the tire 2 The ground reaction of the tire correspond ing to the most critical combination of airplane 66 Federal Aviation Regulations weight up to maximum landing weight and cen ter of gravity position combined with forces of 1 0g downward and 0 31g forward acting at the center of gravity The reactions in this case must be dis tributed to the nose and main wheels by the prin ciples of statics with a drag reaction equal to 0 31 times the vertical load at e
311. eed the limits set forth in 825 143 d subject ASA Part 25 Airworthiness Standards Transport Category to the conditions set forth in paragraphs e and f of 825 143 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 121 72 FR 44669 Aug 8 2007 825 943 Negative acceleration No hazardous malfunction of an engine an auxiliary power unit approved for use in flight or any component or system associated with the powerplant or auxiliary power unit may occur when the airplane is operated at the negative ac celerations within the flight envelopes prescribed in 825 333 This must be shown for the greatest duration expected for the acceleration Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15043 March 17 1977 825 945 Thrust or power augmentation system a General Each fluid injection system must provide a flow of fluid at the rate and pressure es tablished for proper engine functioning under each intended operating condition If the fluid can freeze fluid freezing may not damage the air plane or adversely affect airplane performance b Fluid tanks Each augmentation system fluid tank must meet the following requirements 1 Each tank must be able to withstand without failure the vibration inertia fluid and structural loads that it may be subject to in operation 2 The tanks as mounted in the airplane must be
312. eel an gles as shown in Figure 1 The length of the mounting stand legs is 12 1 inches 305 3 mm The mounting stand must be used for mounting the test specimen seat bottom and seat back as shown in Figure 2 The mounting stand should also include a suitable drip pan lined with aluminum foil dull side up 2 Test Burner The burner to be used in test ing must i Be a modified gun type ii Have an 80 degree spray angle nozzle nominally rated for 2 25 gallons hour at 100 psi iii Have 12 inch 305 mm burner cone in stalled at the end of the draft tube with an open ing 6 inches 152 mm high and 11 inches 280 mm wide as shown in Figure 3 and iv Have a burner fuel pressure regulator that is adjusted to deliver a nominal 2 0 gallon hour of 2 Grade kerosene or equivalent required for the test Burner models which have been used success fully in testing are the Lennox Model OB 32 Car lin Model 200 CRD and Park Model DPL 3400 FAA published reports pertinent to this type of burner are 1 Powerplant Engineering Report No 3A Standard Fire Test Apparatus and Proce dure for Flexible Hose Assemblies dated March 1978 and 2 Report No DOT FAA RD 76 213 Reevaluation of Burner Characteristics for Fire Resistance Tests dated January 1977 3 Calorimeter i The calorimeter to be used in testing must be a 0 15 0 BTU ft2 sec 0 17 0 W cm calorim eter accurate 3 mounted in a 6 inch by 12 inch
313. eet above the cabin floor and 2 Proceed to the exits using the marking sys tem necessary to accomplish the actions in 25 812 e 1 and e 2 c Transverse Separation of the Fuselage In the event of a transverse separation of the fuse lage compliance must be shown with 25 812 1 except as follows 1 For each airplane type originally type certif icated with a maximum passenger seating capac ity of 9 or less not more than 50 percent of all electrically illuminated emergency lights required by 25 812 may be rendered inoperative in addi tion to the lights that are directly damaged by the separation 2 For each airplane type originally type certif icated with a maximum passenger seating capac ity of 10 to 19 not more than 33 percent of all electrically illuminated emergency lights required SFAR No 109 to Part 25 by 825 812 may be rendered inoperative in addi tion to the lights that are directly damaged by the Separation 10 Interior doors In lieu of the requirements of 25 813 e interior doors may be installed be tween passenger seats and exits provided the following requirements are met a Each door between any passenger seat occupiable for taxi takeoff and landing and any emergency exit must have a means to signal to the flightcrew at the flightdeck that the door is in the open position for taxi takeoff and landing b Appropriate procedures limitations must be established to ensure that any
314. egory rective action Systems controls and associated monitoring and warning means must be designed to minimize crew errors which could create addi tional hazards d Compliance with the requirements of para graph b of this section must be shown by analy sis and where necessary by appropriate ground flight or simulator tests The analysis must con sider 1 Possible modes of failure including mal functions and damage from external sources 2 The probability of multiple failures and un detected failures 3 The resulting effects on the airplane and oc cupants considering the stage of flight and oper ating conditions and 4 The crew warning cues corrective action re quired and the capability of detecting faults e In showing compliance with paragraphs a and b of this section with regard to the electrical System and equipment design and installation critical environmental conditions must be consid ered For electrical generation distribution and utilization equipment required by or used in com plying with this chapter except equipment covered by Technical Standard Orders containing environ mental test procedures the ability to provide con tinuous safe service under foreseeable environ mental conditions may be shown by environmental tests design analysis or reference to previous comparable service experience on other aircraft f EWIS must be assessed in accordance with the requirements of
315. elage compartment or 2 There is protective breathing equipment for each flight crewmember on flight deck duty 113 825 1199 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 40 42 FR 15044 March 17 1977 825 1199 Extinguishing agent containers a Each extinguishing agent container must have a pressure relief to prevent bursting of the container by excessive internal pressures b The discharge end of each discharge line from a pressure relief connection must be located so that discharge of the fire extinguishing agent would not damage the airplane The line must also be located or protected to prevent clogging caused by ice or other foreign matter c There must be a means for each fire extin guishing agent container to indicate that the con tainer has discharged or that the charging pres sure is below the established minimum necessary for proper functioning d The temperature of each container must be maintained under intended operating conditions to prevent the pressure in the container from 1 Falling below that necessary to provide an adequate rate of discharge or 2 Rising high enough to cause premature dis charge e If a pyrotechnic capsule is used to dis charge the extinguishing agent each container must be installed so that temperature conditions will not cause hazardous deterioration of the pyro technic capsule Docket N
316. ellaneous markings and placards a Baggage and cargo compartments and bal last location Each baggage and cargo compart ment and each ballast location must have a plac ard stating any limitations on contents including weight that are necessary under the loading re quirements However underseat compartments designed for the storage of carry on articles weighing not more than 20 pounds need not have a loading limitation placard b Powerplant fluid filler openings The follow ing apply 1 Fuel filler openings must be marked at or near the filler cover with i The word fuel ii For reciprocating engine powered airplanes the minimum fuel grade iii For turbine engine powered airplanes the permissible fuel designations and iv For pressure fueling systems the maxi mum permissible fueling supply pressure and the maximum permissible defueling pressure 140 Federal Aviation Regulations 2 Oil filler openings must be marked at or near the filler cover with the word oil 3 Augmentation fluid filler openings must be marked at or near the filler cover to identify the re quired fluid c Emergency exit placards Each emergency exit placard must meet the requirements of 25 811 d Doors Each door that must be used in or der to reach any required emergency exit must have a suitable placard stating that the door is to be latched in the open position during takeoff and landing Docket No 5
317. entify track and resolve each cause or potential cause of an inability to start or operate the APU in flight as intended in accordance with the problem tracking and resolution system specified in sec tion K25 3 2 e of this appendix ASA Part 25 Airworthiness Standards Transport Category d Airplane demonstration For each airplane engine combination to be approved for ETOPS the applicant must flight test at least one airplane to demonstrate that the airplane and its compo nents and equipment are capable of functioning properly during ETOPS flights and diversions of the longest duration for which the applicant seeks approval This flight testing may be performed in conjunction with but may not substitute for the flight testing required by 21 35 b 2 1 The airplane demonstration flight test pro gram must include i Flights simulating actual ETOPS including flight at normal cruise altitude step climbs and if applicable APU operation ii Maximum duration flights with maximum du ration diversions iii Maximum duration engine inoperative di versions distributed among the engines installed on the airplanes used for the airplane demonstra tion flight test program At least two one engine inoperative diversions must be conducted at max imum continuous thrust or power using the same engine iv Flights under non normal conditions to vali date the flightcrew s ability to safely conduct an ETOPS diversion wit
318. equipment Equipment systems and installations Power source capacity and distribution System lightning protection High Intensity Radiated Fields HIRF Protection INSTRUMENTS INSTALLATION Arrangement and visibility Warning caution and advisory lights Airspeed indicating system Static pressure systems Pitot heat indication systems Magnetic direction indicator Flight guidance system Instruments using a power supply Instrument systems Flight director systems Powerplant instruments ELECTRICAL SYSTEMS AND EQUIPMENT 25 1351 25 1353 25 1355 25 1357 25 1360 25 1362 25 1363 25 1365 25 1381 25 1383 25 1385 25 1387 25 1389 General Electrical equipment and installations Distribution system Circuit protective devices Precautions against injury Electrical supplies for emergency conditions Electrical system tests Electrical appliances motors and transformers LIGHTS Instrument lights Landing lights Position light system installation Position light system dihedral angles Position light distribution and intensities ASA Part 25 Airworthiness Standards Transport Category 25 1391 25 1393 25 1395 25 1397 25 1399 25 1401 25 1403 25 1411 25 1415 25 1419 25 1421 25 1423 25 1431 25 1433 25 1435 25 1438 25 1439 25 1441 25 1443 25 1445 25 1447 25 1449 25 1450 25 1453 25 1455 25 1457 25 1459 25 1461 Minimum intensities in the horizo
319. er and iii To which the cockpit voice recorder and cockpit mounted area microphone are switched automatically in the event that all other power to 134 Federal Aviation Regulations the cockpit voice recorder is interrupted either by normal shutdown or by any other loss of power to the electrical power bus and 6 It is in a separate container from the flight data recorder when both are required If used to comply with only the cockpit voice recorder re quirements a combination unit may be installed e The recorder container must be located and mounted to minimize the probability of rupture of the container as a result of crash impact and con sequent heat damage to the recorder from fire 1 Except as provided in paragraph e 2 of this section the recorder container must be lo cated as far aft as practicable but need not be outside of the pressurized compartment and may not be located where aft mounted engines may crush the container during impact 2 If two separate combination digital flight data recorder and cockpit voice recorder units are installed instead of one cockpit voice recorder and one digital flight data recorder the combination unit that is installed to comply with the cockpit voice recorder requirements may be located near the cockpit f If the cockpit voice recorder has a bulk era sure device the installation must be designed to minimize the probability of inadvertent operation and actuation of
320. er must remain lighted for the entire duration of the test except that there may be intermittent flame extin guishment for periods that do not exceed 3 sec onds Furthermore if the optional three hole up per burner is used at least two flamelets must re main lighted for the entire duration of the test except that there may be intermittent flame extin guishment of all three flamelets for periods that do not exceed 3 seconds 9 A minimum of three specimens must be tested f Calculations 1 The calibration factor is calculated as fol lows kw 2241 01433 1000 HRR Heat release rate kw m2 baseline voltage mv Vm measured thermopile voltage mv calibration factor kw mv 3 The integral of the heat release rate is the total heat release as a function of time and is cal culated by multiplying the rate by the data sam pling frequency in minutes and summing the time from zero to two minutes g Criteria The total positive heat release over the first two minutes of exposure for each of the three or more samples tested must be averaged and the peak heat release rate for each of the samples must be averaged The average total heat release must not exceed 65 kilowatt minutes 183 Appendix to Part 25 per square meter and the average peak heat re lease rate must not exceed 65 kilowatts per square meter h Report The test report must include the fol lowing for each specimen test
321. er than electrical means and must have an initial bright ness of at least 400 microlamberts The colors 82 Federal Aviation Regulations may be reversed in the case of a sign that is self illuminated by other than electrical means 2 For airplanes that have a passenger seating configuration excluding pilot seats of nine seats or less that are required by 25 811 d 1 2 and 3 must have red letters at least 1 inch high on a white background at least 2 inches high These signs may be internally electrically illumi nated or self illuminated by other than electrical means with an initial brightness of at least 160 microlamberts The colors may be reversed in the case of a sign that is self illuminated by other than electrical means c General illumination in the passenger cabin must be provided so that when measured along the centerline of main passenger aisle s and cross aisle s between main aisles at seat arm rest height and at 40 inch intervals the average il lumination is not less than 0 05 foot candle and the illumination at each 40 inch interval is not less than 0 01 foot candle A main passenger aisle s is considered to extend along the fuselage from the most forward passenger emergency exit or cabin occupant seat whichever is farther forward to the most rearward passenger emergency exit or cabin occupant seat whichever is farther aft d The floor of the passageway leading to each floor level passenge
322. ers of gravity within the limits for which certifica tion is requested must be considered to reach maximum design loads for each part of the sea plane structure Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 825 525 Application of loads a Unless otherwise prescribed the seaplane as a whole is assumed to be subjected to the loads corresponding to the load factors specified 825 527 b In applying the loads resulting from the load factors prescribed in 825 527 the loads may be distributed over the hull or main float bottom in order to avoid excessive local shear loads and bending moments at the location of water load ap plication using pressures not less than those pre scribed in 825 533 b c For twin float seaplanes each float must be treated as an equivalent hull on a fictitious sea plane with a weight equal to one half the weight of the twin float seaplane d Except in the takeoff condition of 825 531 the aerodynamic lift on the seaplane during the impact is assumed to be 24 of the weight of the seaplane 825 527 Hull and main float load factors a Water reaction load factors ny must be computed in the following manner 1 For the step landing case Ci Vso 2 p ny 51 825 529 2 For the bow and stern landing cases Ci Vigo Ky M 2 3 rav p y 0 b The following values are used 1 nw water
323. established for dry runways and must be at least 20 knots or 0 2 Vspo whichever is greater except that it need not exceed 25 knots 2 The crosswind component for takeoff estab lished without ice accretions is valid in icing condi tions 3 The landing crosswind component must be established for i Non icing conditions and ii Icing conditions with the landing ice accre tion defined in appendix C b For seaplanes and amphibians the follow ing applies 1 A 90 degree cross component of wind ve locity up to which takeoff and landing is safe un der all water conditions that may reasonably be expected in normal operation must be estab lished and must be at least 20 knots or 0 2 whichever is greater except that it need not ex ceed 25 knots 2 A wind velocity for which taxiing is safe in any direction under all water conditions that may reasonably be expected in normal operation must be established and must be at least 20 knots or 0 2 Vsno whichever is greater except that it need not exceed 25 knots Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2322 Jan 16 1978 Amdt 25 108 67 FR 70828 Nov 26 2002 Amdt 25 121 72 FR 44668 Aug 8 2007 825 239 Spray characteristics control and stability on water a For seaplanes and amphibians during take off taxiing and landing and in the condi tions set forth in paragraph b of this section there may b
324. et There must be means to prevent en trance into the tank itself or into the tank outlet of any object that might obstruct the flow of oil through the system No oil tank outlet may be en closed by any screen or guard that would reduce the flow of oil below a safe value at any operating temperature There must be a shutoff valve at the outlet of each oil tank used with a turbine engine unless the external portion of the oil system in cluding the oil tank supports is fireproof f Flexible oil tank liners Each flexible oil tank liner must be approved or must be shown to be suitable for the particular application Docket No 5066 29 FR 18291 Dec 24 as amended by Amdt 25 19 33 FR 15410 Oct 17 1968 Amdt 25 23 35 FR 5677 April 8 1970 Amdt 25 36 39 FR 35460 Oct 1 1974 Amdt 25 57 49 FR 6848 Feb 23 1984 Amdt 25 72 55 FR 29785 July 20 1990 25 1015 Oil tank tests Each oil tank must be designed and installed so that a It can withstand without failure each vibra tion inertia and fluid load that it may be sub jected to in operation and b It meets the provisions of 25 965 except 1 The test pressure i For pressurized tanks used with a turbine engine may not be less than 5 p s i plus the max imum operating pressure of the tank instead of the pressure specified in 25 965 a and ii For all other tanks may not be less than 5 p s i instead of the pressure specified in
325. eting that would prevent continued safe flight in any likely operating condition b Each part of the airplane must be demon strated in flight to be free from excessive vibration under any appropriate speed and power condi tions up to VpF Mpr The maximum speeds shown must be used in establishing the operating limita tions of the airplane in accordance with 825 1505 c Except as provided in paragraph d of this section there may be no buffeting condition in normal flight including configuration changes during cruise severe enough to interfere with the control of the airplane to cause excessive fatigue to the crew or to cause structural damage Stall warning buffeting within these limits is allowable d There may be no perceptible buffeting con dition in the cruise configuration in straight flight at any speed up to except that stall warn ing buffeting is allowable e For an airplane with Mp greater than 6 or with a maximum operating altitude greater than 25 000 feet the positive maneuvering load factors at which the onset of perceptible buffeting occurs must be determined with the airplane in the cruise configuration for the ranges of airspeed or Mach number weight and altitude for which the air plane is to be certificated The envelopes of load factor speed altitude and weight must provide a sufficient range of speeds and load factors for normal operations Probable inadvertent excur sions beyond t
326. event the ATD s pelvis from translating beyond the end of the seat at any time during testing 8 Test Parameters i All seat positions need to be occupied by ATDs for the longitudinal tests ii A minimum of one longitudinal test con ducted in accordance with the conditions speci fied in 25 562 b 2 is required to assess the in jury criteria as follows Note that if a seat is in stalled aft of structure such as an interior wall or furnishing that does not have a homogeneous surface an additional test or tests may be re quired to demonstrate that the injury criteria are met for the area which an occupant could contact For example different yaw angles could result in different injury considerations and may require Separate tests to evaluate A For configurations without structure such as a wall or bulkhead installed directly forward of the forward seat place Hybrid II ATDs or equiva lent must be in all seat places B For configurations with structure such as a wall or bulkhead installed directly forward of the forward seat place a side impact dummy or equivalent ATD or more appropriate ATD must be ASA Part 25 Airworthiness Standards Transport Category in the forward seat place and a Hybrid Il ATD or equivalent must be in all other seat places C The test may be conducted with or without deformed floor D The test must be conducted with either no yaw or 10 degrees yaw for evaluating occupant in
327. event the insertion of the maxi mum approved takeoff thrust or power or must be shown to be an improbable event 2 Shall not result in a significant loss or reduc tion in thrust or power or must be shown to be an extremely improbable event b The concurrent existence of an ATTCS fail ure and an engine failure during the critical time in terval must be shown to be extremely improbable c All applicable performance requirements of Part 25 must be met with an engine failure occur ring at the most critical point during takeoff with the ATTCS system functioning 125 4 THRUST SETTING The initial takeoff thrust or power setting on each engine at the beginning of the takeoff roll may not be less than any of the following a Ninety 90 percent of the thrust or power set by the ATTCS the maximum takeoff thrust or power approved for the airplane under existing ambient conditions b That required to permit normal operation of all safety related systems and equipment depen dent upon engine thrust or power lever position or c That shown to be free of hazardous engine response characteristics when thrust or power is advanced from the initial takeoff thrust or power to the maximum approved takeoff thrust or power 125 5 POWERPLANT CONTROLS a In addition to the requirements of 25 1141 no single failure or malfunction or probable combi nation thereof of the ATTCS including associated systems may cause the failure
328. f 70 de grees to the right and to the left respectively to a vertical plane passing through the longitudinal ASA 825 1389 axis as viewed when looking aft along the longi tudinal axis e If the rear position light when mounted as far aft as practicable in accordance with 25 1385 c cannot show unbroken light within dihedral angle A as defined in paragraph d of this section a solid angle or angles of obstructed visibility totaling not more than 0 04 steradians is allowable within that dihedral angle if such solid angle is within a cone whose apex is at the rear position light and whose elements make an angle of 30 with a vertical line passing through the rear position light Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 30 36 FR 21278 Nov 5 1971 25 1389 Position light distribution and intensities a General The intensities prescribed in this section must be provided by new equipment with light covers and color filters in place Intensities must be determined with the light source operat ing at a steady value equal to the average lumi nous output of the source at the normal operating voltage of the airplane The light distribution and intensity of each position light must meet the re quirements of paragraph b of this section b Forward and rear position lights The light distribution and intensities of forward and rear po sition lights must be expressed in terms of mini mu
329. f guidance material includ ing any relevant limitations or information 5 An explanation of significant or unusual flight or ground handling characteristics of the air plane 6 Corrections to indicated values of airspeed altitude and outside air temperature 7 An explanation of operational landing run way length factors included in the presentation of the landing distance if appropriate Docket No 2000 8511 66 FR 34024 June 26 2001 Amdt 25 108 67 FR 70828 Nov 26 2002 ASA Part 25 Airworthiness Standards Transport Category Subpart H Electrical Wiring Interconnection Systems EWIS Source FAA 2004 18379 72 FR 63406 Nov 8 2007 unless otherwise noted 825 1701 Definition a As used in this chapter electrical wiring interconnection system EWIS means any wire wiring device or combination of these including termination devices installed in any area of the airplane for the purpose of transmitting electrical energy including data and signals between two or more intended termination points This in cludes 1 Wires and cables 2 Bus bars 3 The termination point on electrical devices including those on relays interrupters switches contactors terminal blocks and circuit breakers and other circuit protection devices 4 Connectors including feed through connec tors 5 Connector accessories 6 Electrical grounding and bonding devices and their assoc
330. f run at which when the critical engine is sud denly made inoperative it is possible to maintain 26 Federal Aviation Regulations control of the airplane using the rudder control alone without the use of nosewheel steering as limited by 150 pounds of force and the lateral control to the extent of keeping the wings level to enable the takeoff to be safely continued using normal piloting skill In the determination of Vc assuming that the path of the airplane accelerat ing with all engines operating is along the center line of the runway its path from the point at which the critical engine is made inoperative to the point at which recovery to a direction parallel to the centerline is completed may not deviate more than 30 feet laterally from the centerline at any point must be established with 1 The airplane in each takeoff configuration or at the option of the applicant in the most criti cal takeoff configuration 2 Maximum available takeoff power or thrust on the operating engines 3 The most unfavorable center of gravity 4 The airplane trimmed for takeoff and 5 The most unfavorable weight in the range of takeoff weights f Vcc the minimum control speed during ap proach and landing with all engines operating is the calibrated airspeed at which when the critical engine is suddenly made inoperative it is possi ble to maintain control of the airplane with that en gine still inoperative a
331. fined in accor dance with 25 29 2 Loading instructions necessary to ensure loading of the airplane within the weight and cen ter of gravity limits and to maintain the loading within these limits in flight 3 If certification for more than one center of gravity range is requested the appropriate limita tions with regard to weight and loading proce dures for each separate center of gravity range d Flight crew The number and functions of the minimum flight crew determined under 825 1523 must be furnished e Kinds of operation The kinds of operation approved under 825 1525 must be furnished f Ambient air temperatures and operating alti tudes The extremes of the ambient air tempera tures and operating altitudes established under 825 1527 must be furnished g Reserved h Additional operating limitations The operat ing limitations established under 825 1533 must be furnished i Maneuvering flight load factors The positive maneuvering limit load factors for which the struc ture is proven described in terms of accelera tions must be furnished Docket No 5066 29 FR 1891 Dec 24 1964 as amended by Amdt 25 38 41 FR 55468 Dec 20 1976 Amdt 25 42 43 FR 2323 Jan 16 1978 Amdt 25 46 43 FR 50598 Oct 30 1978 Amdt 25 72 55 FR 29787 July 20 1990 Amdt 25 105 66 FR 34024 June 26 2001 825 1585 Operating procedures a Operating procedures must be furnished for 1 Normal
332. flated tires except that it may not be less than 1g and 4 Pivoting need not be considered f Towing conditions For one and for two de flated tires the towing load Frow must be 60 percent and 50 percent respectively of the load prescribed 825 519 Jacking and tie down provisions a General The airplane must be designed to withstand the limit load conditions resulting from the static ground load conditions of paragraph b of this section and if applicable paragraph c of this section at the most critical combinations of airplane weight and center of gravity The maxi mum allowable load at each jack pad must be specified b Jacking The airplane must have provisions for jacking and must withstand the following limit loads when the airplane is supported on jacks 1 For jacking by the landing gear at the maxi mum ramp weight of the airplane the airplane structure must be designed for a vertical load of 1 33 times the vertical static reaction at each jack ing point acting singly and in combination with a horizontal load of 0 33 times the vertical static re action applied in any direction 2 For jacking by other airplane structure at maximum approved jacking weight i The airplane structure must be designed for a vertical load of 1 33 times the vertical reaction at each jacking point acting singly and in combi nation with a horizontal load of 0 33 times the ver tical static reaction applied in any direct
333. flated tires except that for multiple wheel gear units with more than one shock strut a rational distribution of the ground reactions be tween the deflated and inflated tires accounting for the differences in shock strut extensions re sulting from a deflated tire may be used d Landing conditions For one and for two de flated tires the applied load to each gear unit is assumed to be 60 percent and 50 percent respec tively of the limit load applied to each gear for each of the prescribed landing conditions How ever for the drift landing condition of 825 485 100 percent of the vertical load must be applied e Taxiing and ground handling conditions For one and for two deflated tires 1 The applied side or drag load factor or both factors at the center of gravity must be the most critical value up to 50 percent and 40 percent re spectively of the limit side or drag load factors or both factors corresponding to the most severe condition resulting from consideration of the pre scribed taxiing and ground handling conditions ASA Part 25 Airworthiness Standards Transport Category 2 For the braked roll conditions of 825 493 a and b 2 the drag loads on each inflated tire may not be less than those at each tire for the symmet rical load distribution with no deflated tires 3 The vertical load factor at the center of grav ity must be 60 percent and 50 percent respec tively of the factor with no de
334. flutter deformation and vibration re quirements must also be met with zero fuel Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 18 33 FR 12226 Aug 30 1968 Amdt 25 72 55 FR 37607 Sept 12 1990 Amdt 25 86 61 FR 5221 Feb 9 1996 825 345 High lift devices a If wing flaps are to be used during takeoff approach or landing at the design flap speeds established for these stages of flight under 25 335 e and with the wing flaps in the corre sponding positions the airplane is assumed to be subjected to symmetrical maneuvers and gusts The resulting limit loads must correspond to the conditions determined as follows 1 Maneuvering to a positive limit load factor of 2 0 and 2 Positive and negative gusts of 25 ft sec EAS acting normal to the flight path in level flight Gust loads resulting on each part of the structure must be determined by rational analysis The analysis must take into account the unsteady aerodynamic characteristics and rigid body motions of the air craft The shape of the gust must be as described in 825 341 a 2 except that Uas 25 ft sec EAS H 12 5 c and mean geometric chord of the wing feet b The airplane must be designed for the con ditions prescribed in paragraph a of this section except that the airplane load factor need not ex ceed 1 0 taking into account as separate condi tions the effects of 1 Propeller slipstream correspond
335. following must be met for each cargo or baggage compartment with those provisions a The detection system must provide a visual indication to the flight crew within one minute after the start of a fire b The system must be capable of detecting a fire at a temperature significantly below that at which the structural integrity of the airplane is substantially decreased c There must be means to allow the crew to check in flight the functioning of each fire detec tor circuit d The effectiveness of the detection system must be shown for all approved operating configu rations and conditions Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 54 45 FR 60173 Sept 11 1980 Amdt 25 93 63 FR 8048 Feb 17 1998 ASA 825 859 825 859 Combustion heater fire protection a Combustion heater fire zones The following combustion heater fire zones must be protected from fire in accordance with the applicable provi sions of 25 1181 through 25 1191 and 8825 1195 through 25 1203 1 The region surrounding the heater if this re gion contains any flammable fluid system compo nents excluding the heater fuel system that could i Be damaged by heater malfunctioning or ii Allow flammable fluids or vapors to reach the heater in case of leakage 2 The region surrounding the heater if the heater fuel system has fittings that if they leaked would allow fuel or vapors to enter this region
336. fuel for each fuel tank must be furnished Docket No FAA 2000 8511 66 FR 34024 June 26 2001 825 1587 Performance information a Each Airplane Flight Manual must contain information to permit conversion of the indicated temperature to free air temperature if other than a free air temperature indicator is used to comply with the requirements of 25 1303 a 1 b Each Airplane Flight Manual must contain the performance information computed under the applicable provisions of this part including 25 115 25 123 and 25 125 for the weights al titudes temperatures wind components and run way gradients as applicable within the opera tional limits of the airplane and must contain the following 142 Federal Aviation Regulations 1 In each case the conditions of power con figuration and speeds and the procedures for handling the airplane and any system having a significant effect on the performance information 2 Vsr determined in accordance with 25 103 3 The following performance information de termined by extrapolation and computed for the range of weights between the maximum landing weight and the maximum takeoff weight i Climb in the landing configuration ii Climb in the approach configuration iii Landing distance 4 Procedures established under 825 101 f and g that are related to the limitations and infor mation required by 825 1533 and by this para graph b in the form o
337. g of a back cushion specimen and a bottom cushion specimen If a cushion in cluding outer dress covering is demonstrated to meet the requirements of this appendix using the oil burner test the dress covering of that cushion may be replaced with a similar dress covering provided the burn length of the replacement cov ering as determined by the test specified in 25 853 c does not exceed the corresponding burn length of the dress covering used on the cushion subjected to the oil burner test 4 For at least two thirds of the total number of specimen sets tested the burn length from the burner must not reach the side of the cushion op posite the burner The burn length must not ex ceed 17 inches Burn length is the perpendicular distance from the inside edge of the seat frame closest to the burner to the farthest evidence of damage to the test specimen due to flame im pingement including areas of partial or complete consumption charring or embrittlement but not including areas sooted stained warped or dis colored or areas where material has shrunk or melted away from the heat source 5 The average percentage weight loss must not exceed 10 percent Also at least two thirds of the total number of specimen sets tested must not exceed 10 percent weight loss All droppings fall ing from the cushions and mounting stand are to be discarded before the after test weight is deter mined The percentage weight loss for a speci me
338. g takeoff and landing by a flight attendant required by the operating rules of this chapter must be 1 Near a required floor level emergency exit except that another location is acceptable if the emergency egress of passengers would be en hanced with that location A flight attendant seat must be located adjacent to each Type A or B emergency exit Other flight attendant seats must be evenly distributed among the required floor level emergency exits to the extent feasible 2 To the extent possible without compromis ing proximity to a required floor level emergency exit located to provide a direct view of the cabin area for which the flight attendant is responsible 74 Federal Aviation Regulations 3 Positioned so that the seat will not interfere with the use of a passageway or exit when the seat is not in use 4 Located to minimize the probability that oc cupants would suffer injury by being struck by items dislodged from service areas stowage compartments or service equipment 5 Either forward or rearward facing with an energy absorbing rest that is designed to support the arms shoulders head and spine 6 Equipped with a restraint system consisting of a combined safety belt and shoulder harness unit with single point release There must means to secure each restraint system when not in use to prevent interference with rapid egress in an emergency i Each safety belt must be equipped with a metal t
339. gated at airspeeds up to Vc Mc Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 54 45 FR 60172 Sept 11 1980 Amdt 25 77 57 FR 28949 June 29 1992 Amdt 25 86 61 FR 5220 Feb 9 1996 525 307 Proof of structure a Compliance with the strength and deforma tion requirements of this subpart must be shown for each critical loading condition Structural anal ysis may be used only if the structure conforms to that for which experience has shown this method to be reliable The Administrator may require ulti mate load tests in cases where limit load tests may be inadequate b c Reserved d When static or dynamic tests are used to show compliance with the requirements of 25 305 b for flight structures appropriate mate rial correction factors must be applied to the test results unless the structure or part thereof being tested has features such that a number of ele ments contribute to the total strength of the struc ture and the failure of one element results in the redistribution of the load through alternate load paths Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 54 45 FR 60172 Sept 11 1980 Amdt 25 72 55 FR 29775 July 20 1990 FLIGHT LOADS 25 321 General a Flight load factors represent the ratio of the aerodynamic force component acting normal to the as
340. ge which is the 28 Federal Aviation Regulations greater of 15 percent of the trim speed plus the resulting free return speed range or 50 knots plus the resulting free return speed range above and below the trim speed except that the speed range need not include speeds less than 1 3 nor speeds greater than Vec Mrc nor speeds that require a stick force of more than 50 pounds with i The wing flaps retracted ii The center of gravity in the most adverse position see 825 27 iii The most critical weight between the maxi mum takeoff and maximum landing weights iv 75 percent of maximum continuous power for reciprocating engines or for turbine engines the maximum cruising power selected by the ap plicant as an operating limitation see 825 1521 except that the power need not exceed that re quired at Vyo Myo and v The airplane trimmed for level flight with the power required in paragraph b 1 iv of this section 2 With the landing gear retracted at low speed the stick force curve must have a stable slope at all speeds within a range which is the greater of 15 percent of the trim speed plus the resulting free return speed range or 50 knots plus the resulting free return speed range above and below the trim speed except that the speed range need not include speeds less than 1 3 nor speeds greater than the minimum speed of the applicable speed range prescribed in paragraph b 1 nor speeds
341. gers must be provided as follows 1 Each assist space must be a rectangle on the floor of sufficient size to enable crewmem ber standing erect to effectively assist evacuees The assist space must not reduce the unob structed width of the passageway below that re quired for the exit 2 For each Type A or B exit assist space must be provided at each side of the exit regardless of 84 Federal Aviation Regulations whether an assist means is 25 810 3 For each Type exit installed in an airplane with seating for more than 80 passen gers an assist space must be provided at one side of the passageway regardless of whether an assist means is required by 25 810 a 4 For each Type C or Il exit an assist space must be provided at one side of the passageway if an assist means is required by 25 810 a 5 For any tailcone exit that qualifies for 25 ad ditional passenger seats under the provisions of 25 807 g 9 ii an assist space must be vided if an assist means is required by 825 810 a 6 There must be a handle or handles at each assist space located to enable the crewmember to steady himself or herself i While manually activating the assist means where applicable and ii While assisting passengers during an evac uation c The following must be provided for each Type 1 or Type IV exit 1 There must be access from the nearest to each exit In addition fo
342. ght Manual that there are extra seats installed but that the number of passengers on the airplane must not exceed 60 Additionally there must be a placard installed ad jacent to each door that can be used as a passen ger boarding door that states that the maximum passenger capacity is 60 The placard must be clearly legible to passengers entering the air plane b For airplanes outfitted with interior doors under paragraph 10 of this SFAR the airplane flight manual AFM must include an appropriate limitation that the airplane must be staffed with at least the following number of flight attendants who meet the requirements of 14 CFR 91 533 b 1 The number of flight attendants required by 91 533 a 1 and 2 of this chapter and 2 At least one flight attendant if the airplane model was originally certified for 75 passengers or more c The AFM must include appropriate limita tion s to require a preflight passenger briefing describing the appropriate functions to be per formed by the passengers and the relevant fea tures of the airplane to ensure the safety of the passengers and crew d The airplane may not be offered for com mon carriage or operated for hire The operating limitations section of the AFM must be revised to prohibit any operations involving the carriage of persons or property for compensation or hire The operators may receive remuneration to the extent consistent with parts 125 and 91 subpart F of th
343. ght and in standard sea level condi tions and 2 Vi2 equal to Vso TAS at the appropriate landing weight and altitudes in a hot day tempera ture of 41 degrees F above standard 3 The effects of increased contact speed must be investigated if approval of downwind landings exceeding 10 knots is requested b For the level landing attitude for airplanes with tail wheels the conditions specified in this section must be investigated with the airplane horizontal reference line horizontal in accordance with Figure 2 of Appendix A of this part c For the level landing attitude for airplanes with nose wheels shown in Figure 2 of Appendix A of this part the conditions specified in this sec tion must be investigated assuming the following attitudes 1 An attitude in which the main wheels are as sumed to contact the ground with the nose wheel just clear of the ground and 2 If reasonably attainable at the specified de scent and forward velocities an attitude in which the nose and main wheels are assumed to con tact the ground simultaneously d In addition to the loading conditions pre Scribed in paragraph a of this section but with maximum vertical ground reactions calculated from paragraph a the following apply 1 The landing gear and directly affected at taching structure must be designed for the maxi mum vertical ground reaction combined with an aft acting drag component of not less than 25 of this maximum ve
344. gned and installed in accordance with the fol lowing requirements 1 Surface temperature developed by the gen erator during operation may not create a hazard to the airplane or to its occupants 2 Means must be provided to relieve any in ternal pressure that may be hazardous c In addition to meeting the requirements in paragraph b of this section each portable chem ical oxygen generator that is capable of sustained operation by successive replacement of a genera tor element must be placarded to show 1 The rate of oxygen flow in liters per minute 2 The duration of oxygen flow in minutes for the replaceable generator element and 3 A warning that the replaceable generator el ement may be hot unless the element construc tion is such that the surface temperature cannot exceed 100 degrees F Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 41 42 FR 36971 July 18 1977 ASA 825 1457 825 1453 Protection of oxygen equipment from rupture Oxygen pressure tanks and lines between tanks and the shutoff means must be a Protected from unsafe temperatures and b Located where the probability and hazards of rupture in a crash landing are minimized 825 1455 Draining of fluids subject to freezing If fluids subject to freezing may be drained overboard in flight or during ground operation the drains must be designed and located to prevent the formation of hazardous quantities
345. graph d 3 of this Part along only the 25 inch 6 mm perimeter A V shaped spring is used to hold the assembly together A detachable 50 x 50 x 5 91 inch 12 x 12 x 150 mm drip pan and two 020 inch 5 mm stainless steel wires as shown in Figure 3 of this part IV must be used for testing materials prone to melting and dripping The positioning of the spring and frame may be changed to accommodate different specimen thicknesses by inserting the retaining rod in differ ent holes on the specimen holder ii Since the radiation shield described in ASTM E 906 is not used a guide pin must be added to the injection mechanism This fits into a slotted metal plate on the injection mechanism outside of the holding chamber It can be used to provide accurate positioning of the specimen face after injection The front surface of the specimen must be 3 9 inches 100 mm from the closed ra diation doors after injection iii The specimen holder clips onto the mounted bracket Figure 3 of this part IV The mounting bracket must be attached to the injec tion rod by three screws that pass through a wide area washer welded onto a 1 5 inch 13 mm nut The end of the injection rod must be threaded to 181 Appendix to Part 25 screw into the nut and a 020 inch 5 1 mm thick wide area washer must be held between two 1 2 inch 13 mm nuts that are adjusted to tightly cover the hole in the radiation doors through which the injecti
346. gure 2 ASA Part 25 Airworthiness Standards Transport Category FIGURE 2 Internal Chimney Appendix F to Part 25 FIGURE 3b Air Propane Radiant Panel 1 2 in 13 mm Kaowool M board 16 gauge 1 16 in 1 6 mm aluminum sheet metal 1 8 in 3 2 mm angle iron 15 9 16 in e 394 mm gt 13 in 330 mm 5 1 8 in 4 16 3 16 gt 1 411 mm 2 Radiant heat source Mount the radiant heat energy source in a cast iron frame or equivalent An electric panel must have six 3 inch wide emit ter strips The emitter strips must be perpendicu lar to the length of the panel The panel must have a radiation surface of 12 7 8 by 18 1 2 inches 327 by 470 mm The panel must be capable of operating at temperatures up to 1300 F 704 C An air propane panel must be made of a porous refractory material and have a radiation surface of 12 by 18 inches 305 by 457 mm The panel must be capable of operating at temperatures up to 1 500 F 816 C See figures and 3b FIGURE 3a Electric Panel 18 7 8 in 480 mm emitter strips 6 ASA i Electric radiant panel The radiant panel must be 3 phase and operate at 208 volts A sin gle phase 240 volt panel is also acceptable Use a solid state power controller and microproces sor based controller to set the electric panel oper ating parameters ii Gas radiant panel Use propane liq
347. gy absorbing rest that will support the arms shoulders head and spine or by a safety belt and shoulder harness that will prevent the head from contacting any injurious object Each occu pant of any other seat must be protected from head injury by a safety belt and as appropriate to the type location and angle of facing of each Seat by one or more of the following 1 A shoulder harness that will prevent the head from contacting any injurious object 2 The elimination of any injurious object within striking radius of the head 3 An energy absorbing rest that will support the arms shoulders head and spine e Each berth must be designed so that the forward part has a padded end board canvas dia phragm or equivalent means that can withstand the static load reaction of the occupant when sub jected to the forward inertia force specified in 825 561 Berths must be free from corners and 73 825 785 protuberances likely to cause injury to a person occupying the berth during emergency condi tions f Each seat or berth and its supporting struc ture and each safety belt or harness and its an chorage must be designed for an occupant weight of 170 pounds considering the maximum load factors inertia forces and reactions among the occupant seat safety belt and harness for each relevant flight and ground load condition includ ing the emergency landing conditions prescribed in 825 561 In addition 1 The
348. h temperature g Each exhaust shroud must be ventilated or insulated to avoid during normal operation a temperature high enough to ignite any flammable fluids or vapors external to the shroud Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15043 March 17 1977 825 1123 Exhaust piping For powerplant and auxiliary power unit instal lations the following apply a Exhaust piping must be heat and corrosion resistant and must have provisions to prevent fail ure due to expansion by operating temperatures b Piping must be supported to withstand any vibration and inertia loads to which it would be subjected in operation and c Piping connected to components between which relative motion could exist must have means for flexibility Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15044 March 17 1977 825 1125 Exhaust heat exchangers For reciprocating engine powered airplanes the following apply a Each exhaust heat exchanger must be con structed and installed to withstand each vibration ASA Part 25 Airworthiness Standards Transport Category inertia and other load to which it would be sub jected in operation In addition 1 Each exchanger must be suitable for contin ued operation at high temperatures and resistant to corrosion from exhaust gases 2 There must be means for the inspection of the critical parts of each exchanger
349. h certification is requested can be evacuated from the airplane to the ground under simulated emergency conditions within 90 sec onds Compliance with this requirement must be shown by actual demonstration using the test cri teria outlined in Appendix J of this part unless the Administrator finds that a combination of analysis and testing will provide data equivalent to that which would be obtained by actual demonstration d e Reserved Docket No 24344 55 FR 29781 July 20 1990 525 807 Emergency exits a Type For the purpose of this part the types of exits are defined as follows 1 Type This type is a floor level exit with a rectangular opening of not less than 24 inches wide by 48 inches high with corner radii not greater than eight inches 2 Type II This type is a rectangular opening of not less than 20 inches wide by 44 inches high with corner radii not greater than seven inches Type Il exits must be floor level exits unless lo cated over the wing in which case they must not have a step up inside the airplane of more than 10 inches nor a step down outside the airplane of more than 17 inches 3 Type III This type is a rectangular opening of not less than 20 inches wide by 36 inches high with corner radii not greater than seven inches ASA 825 807 and with a step up inside the airplane of not more than 20 inches If the exit is located over the wing the step down outside the airplane may
350. h worst case ETOPS signifi cant system failures or malfunctions that could oc cur in service v Diversions to airports that represent air ports of the types used for ETOPS diversions vi Repeated exposure to humid and inclement weather on the ground followed by a long duration flight at normal cruise altitude 2 The airplane demonstration flight test pro gram must validate the adequacy of the airplane s flying qualities and performance and the flight crew s ability to safely conduct an ETOPS diver sion under the conditions specified in section K25 3 2 d 1 of this appendix 3 During the airplane demonstration flight test program each test airplane must be operated and maintained using the applicants recom mended operating and maintenance procedures 4 At the completion of the airplane demon stration each ETOPS significant system must un dergo an on wing inspection or test in accordance with the tasks defined in the proposed Instruc tions for Continued Airworthiness to establish its condition for continued safe operation Each en gine must also undergo a gas path inspection These inspections must be conducted in a man ner to identify abnormal conditions that could re sult in an IFSD or diversion The applicant must identify track and resolve any abnormal condi tions in accordance with the problem tracking and resolution system specified in section K25 3 2 e of this appendix e Problem tracking and resolutio
351. hall be established in accordance with the structural performance flight characteristics and ground handling requirements of Part 4b Provided That where literal compliance with the structural re quirements of Part 4b is extremely difficult to ac complish and would not contribute materially to the objective sought and the Administrator finds that the experience with the DC 3 or L 18 air planes justifies it he is authorized to accept such measures of compliance as he finds will effec tively accomplish the basic objective c Airplane flight manual performance operat ing information An approved airplane flight man ual shall be provided for each DC 3 and L 18 air plane which has had new maximum certificated weights established under this section The air plane flight manual shall contain the applicable performance information prescribed in that part of the regulations under which the new certificated weights were established and such additional in formation as may be necessary to enable the ap plication of the take off en route and landing lim itations prescribed for transport category air planes in the operating parts of the Civil Air Regulations d Performance operating limitations Each air plane for which new maximum certificated weights are established in accordance with paragraphs a or b of this section shall be considered a trans port category airplane for the purpose of comply ing with the performance operati
352. haust sys tem parts or exhaust gas impingement must be fireproof e Each airplane must 1 Be designed and constructed so that no fire originating in any fire zone can enter either through openings or by burning through external skin any other zone or region where it would cre ate additional hazards 2 Meet paragraph e 1 of this section with the landing gear retracted if applicable and ASA 825 1197 3 Have fireproof skin in areas subject to flame if a fire starts in the engine power or accessory sections 825 1195 Fire extinguishing systems a Except for combustor turbine and tail pipe sections of turbine engine installations that con tain lines or components carrying flammable flu ids or gases for which it is shown that a fire origi nating in these sections can be controlled there must be a fire extinguisher system serving each designated fire zone b The fire extinguishing system the quantity of the extinguishing agent the rate of discharge and the discharge distribution must be adequate to extinguish fires It must be shown by either ac tual or simulated flights tests that under critical airflow conditions in flight the discharge of the ex tinguishing agent in each designated fire zone specified in paragraph a of this section will pro vide an agent concentration capable of extin guishing fires in that zone and of minimizing the probability of reignition An individual one shot Syste
353. he airplane will allow the occupants to leave the airplane and enter the liferafts required ASA Part 25 Airworthiness Standards Transport Category by 825 1415 If compliance with this provision is shown by buoyancy and trim computations ap propriate allowances must be made for probable structural damage and leakage If the airplane has fuel tanks with fuel jettisoning provisions that can reasonably be expected to withstand a ditching without leakage the jettisonable volume of fuel may be considered as buoyancy volume e Unless the effects of the collapse of exter nal doors and windows are accounted for in the investigation of the probable behavior of the air plane in a water landing as prescribed in para graphs c and d of this section the external doors and windows must be designed to with stand the probable maximum local pressures Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29781 July 20 1990 25 803 Emergency evacuation a Each crew and passenger area must have emergency means to allow rapid evacuation in crash landings with the landing gear extended as well as with the landing gear retracted consider ing the possibility of the airplane being on fire b Reserved c For airplanes having a seating capacity of more than 44 passengers it must be shown that the maximum seating capacity including the num ber of crewmembers required by the operating rules for whic
354. he boundaries of the buffet onset envelopes may not result in unsafe conditions ASA Part 25 Airworthiness Standards Transport Category Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5671 April 8 1970 Amdt 25 72 55 FR 29775 July 20 1990 Amdt 25 77 57 FR 28949 June 29 1992 825 253 High speed characteristics a Speed increase and recovery characteris tics The following speed increase and recovery characteristics must be met 1 Operating conditions and characteristics likely to cause inadvertent speed increases in cluding upsets in pitch and roll must be simulated with the airplane trimmed at any likely cruise speed up to Vyo Mmo These conditions and characteristics include gust upsets inadvertent control movements low stick force gradient in re lation to control friction passenger movement leveling off from climb and descent from Mach to airspeed limit altitudes 2 Allowing for pilot reaction time after effective inherent or artificial speed warning occurs it must be shown that the airplane can be recovered to a normal attitude and its speed reduced to without i Exceptional piloting strength or skill ii Exceeding Vp Mp VpE Mpr or the structural limitations and iii Buffeting that would impair the pilot s ability to read the instruments or control the airplane for recovery 3 With the airplane trimmed at any speed up to t
355. he test or must be placed within a chamber approximately 2 feet high by 1 foot by 1 foot open at the top and at one vertical side front and which allows sufficient flow of air for complete combustion but which is free from drafts The specimen must be parallel to and ap proximately 6 inches from the front of the cham ber The lower end of the specimen must be held rigidly clamped The upper end of the specimen must pass over a pulley or rod and must have an appropriate weight attached to it so that the spec imen is held tautly throughout the flammability test The test specimen span between lower clamp and upper pulley or rod must be 24 inches and must be marked 8 inches from the lower end to indicate the central point for flame application A flame from a Bunsen or Tirrill burner must be applied for 30 seconds at the test mark The burner must be mounted underneath the test mark on the specimen perpendicular to the spec imen and at an angle of 30 to the vertical plane of the specimen The burner must have a nominal bore of inch and be adjusted to provide a 3 inch high flame with an inner cone approximately one third of the flame height The minimum tem perature of the hottest portion of the flame as measured with a calibrated thermocouple pyrom eter may not be less than 1750 F The burner must be positioned so that the hottest portion of the flame is applied to the test mark on the wire Flame time burn length and flaming t
356. hen either source is selected the other is blocked off and 2 Both sources cannot be blocked off simulta neously h For unpressurized airplanes paragraph g 1 of this section does not apply if it can be demonstrated that the static pressure system cal ibration when either static pressure source is se lected is not changed by the other static pressure Source being open or blocked Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 5 30 FR 8261 June 29 1965 Amdt 25 12 32 FR 7587 May 24 1967 Amdt 25 41 42 FR 36970 July 18 1977 Amdt 25 108 67 FR 70828 Nov 26 2002 825 1326 Pitot heat indication systems If a flight instrument pitot heating system is in stalled an indication system must be provided to indicate to the flight crew when that pitot heating System is not operating The indication system must comply with the following requirements a The indication provided must incorporate an amber light that is in clear view of a flight crew member b The indication provided must be designed to alert the flight crew if either of the following con ditions exist 1 The pitot heating system is switched off 2 The pitot heating system is switched on and any pitot tube heating element is inoperative Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 43 43 FR 10339 March 13 1978 120 Federal Aviation Regulations 825 1327 Magnetic direction indica
357. her fluids or fumes that might be present d There must be means to allow the crew to check in flight the functioning of each fire or overheat detector electric circuit e Components of each fire or overheat detec tor system in a fire zone must be at least fire re sistant f No fire or overheat detector system compo nent for any fire zone may pass through another fire zone unless 1 It is protected against the possibility of false warnings resulting from fires in zones through which it passes or 2 Each zone involved is simultaneously pro tected by the same detector and extinguishing system g Each fire detector system must be con structed so that when it is in the configuration for installation it will not exceed the alarm activation time approved for the detectors using the re sponse time criteria specified in the appropriate Technical Standard Order for the detector h EWIS for each fire or overheat detector sys tem in a fire zone must meet the requirements of 825 1731 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5678 April 8 1970 Amdt 25 26 36 FR 5493 March 24 1971 Amdt 25 123 72 FR 63405 Nov 8 2007 25 1207 Compliance Unless otherwise specified compliance with the requirements of 8825 1181 through 25 1203 must be shown by a full scale fire test or by one or more of the following methods a Tests of similar powerplant configurations b Test
358. here must be no reversal of the re sponse to control input about any axis at any speed up to VpE Mpr Any tendency to pitch roll or yaw must be mild and readily controllable us ing normal piloting techniques When the airplane is trimmed at Vyo Myo the slope of the elevator control force versus speed curve need not be sta ble at speeds greater than Vec Mrc but there must be a push force at all speeds up to Vpe Mpr and there must be no sudden or excessive reduc tion of elevator control force as is reached b Maximum speed for stability characteristics VrEc Mgc Vgc Mrec is the maximum speed at which requirements of 8825 143 g 25 147 E 25 175 b 1 25 177 and 25 181 must be met with flaps and landing gear retracted Ex cept as noted in 25 253 c may not be less than a speed midway between and VprE Mpg except that for altitudes where Mach number is the limiting factor need not ex ceed the Mach number at which effective speed warning occurs c Maximum speed for stability characteristics in icing conditions The maximum speed for stabil ity characteristics with the ice accretions defined in appendix C at which the requirements of 8825 143 g 25 147 e 25 175 b 1 25 177 and 25 181 must be met is the lower of ASA 825 255 1 300 knots CAS 2 Vec or 3 A speed at which is demonstrated that airframe will be free of ice accretion due to th
359. hese lateral c g displacements on the loading of the main gear elements or on the air plane structure provided 1 The lateral displacement of the c g results from random passenger or cargo disposition 45 825 473 within the fuselage or from random unsymmetri cal fuel loading or fuel usage and 2 Appropriate loading instructions for random disposable loads are included under the provi sions of 25 1583 c 1 to ensure that the lateral displacement of the center of gravity is main tained within these limits c Landing gear dimension data Figure 1 of Appendix A contains the basic landing gear di mension data Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 825 473 Landing load conditions and assumptions a For the landing conditions specified in 825 479 to 825 485 the airplane is assumed to contact the ground 1 In the attitudes defined in 25 479 and 825 481 2 With a limit descent velocity of 10 fps at the design landing weight the maximum weight for landing conditions at the maximum descent veloc ity and 3 With a limit descent velocity of 6 fps at the design takeoff weight the maximum weight for landing conditions at a reduced descent velocity 4 The prescribed descent velocities may be modified if it is shown that the airplane has design features that make it impossible to develop these velocities b Airplane lift not exceeding
360. hiness Standards Transport Category Subpart G Operating Limitations and Information 825 1501 General a Each operating limitation specified in 25 1503 through 25 1533 and other limitations and information necessary for safe operation must be established b The operating limitations and other informa tion necessary for safe operation must be made available to the crewmembers as prescribed in 8825 1541 through 25 1587 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2323 Jan 16 1978 OPERATING LIMITATIONS 825 1503 Airspeed limitations general When airspeed limitations are a function of weight weight distribution altitude or Mach num ber limitations corresponding to each critical combination of these factors must be established 825 1505 Maximum operating limit speed The maximum operating limit speed airspeed or Mach Number whichever is critical at a particular altitude is a speed that may not be deliberately exceeded in any regime of flight climb cruise or descent unless a higher speed is authorized for flight test or pilot training opera tions must be established so that it is not greater than the design cruising speed Vc and so that it is sufficiently below Vp Mp or Vpe Mpr to make it highly improbable that the latter speeds will be inadvertently exceeded in operations The speed margin between Vmo Mmo and Vp Mp or may
361. hiness Standards Transport Category Appendix F to Part 25 FIGURE 7 Test Specimen Installation on Test Frame gt 4 Calorimeter 2 Calorimeter 1 Spring Clip Squeezes Insulation Sample __ Field Blanket 4 102 mm E Hat Shaped Stringer Steel Z Former Burner Cone ASA 209 Appendix to Part 25 Federal Aviation Regulations FIGURE 8 Burner Information and Calibration Settings Burner Type Nozzle Type Park Model DPL 3400 Monarch Manufacturing Co Inc 80 PL Hollow Cone gt Inner Stator H215 End Stator F124 A Thermocouples Thermo Electric Co Inc Type K Grounded 1 8 Burner Cone Ceramic Packed Metal Sheathed Air Velocity Meter Thermocouple Omega Engineering Inc Model HH30A A Burner Calibration Requirements Fuel Flowrate 6 0 0 2 gal hr Heat Flux Transducer Air Velocity 2150 50 ft min Vatell Corporation Temperature 1900 100 F Model 1000 Series Heat Flux 16 0 0 8 Btu ft sec Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 32 37 FR 3972 Feb 24 1972 37 FR 5284 March 14 1972 Amdt 25 55 47 FR 13315 March 29 1982 Amdt 25 59 49 FR 43193 Oct 26 1984 Amdt 25 60 51 FR 18243 May 16 1986 Amdt 25 61 51 FR 26214 July 21 1986 51 FR 28322 Aug 7 1986 Amdt 25 66 53 FR 32573 Aug 25 1988 53 FR 37542 37671 Sept 27 1988 Amdt 25 72 55 FR 29787 July 20 1990 Amdt 25 83
362. ht inten sity and the overlap intensity in Area B is not greater than 2 5 percent of peak position light in tensity 825 1391 Minimum intensities in the horizontal plane of forward and rear position lights Each position light intensity must equal or exceed the applicable values in the following table Angle from right or Dihedral angle left of longitudinal Intensity light included axis measured candles from dead ahead L and R forward 0 to 10 40 red and green 10 to 20 30 20 to 110 5 A rear white 110 to 180 20 25 1393 Minimum intensities in any vertical plane of forward and rear position lights Each position light intensity must equal or ex ceed the applicable values in the following table Intensity 0 1 00 0 to 5 0 90 5 to 10 0 80 10 to 15 0 70 15 to 20 0 50 20 to 30 0 30 30 to 40 0 10 40 to 90 0 05 25 1395 Maximum intensities in overlapping beams of forward and rear position lights No position light intensity may exceed the ap plicable values in the following table except as provided in 25 1389 b 3 Maximum Intensity Overlaps AreaA AreaB candles candles Green in dihedral angle L 10 1 Red in dihedral angle R 10 1 Green in dihedral angle A 5 1 Red in dihedral angle A 5 1 Rear white in dihedral angle L 5 1 Rear white in dihedral angle 5 1
363. iate crewmembers d Instrument panel vibration may not damage or impair the accuracy of any instrument e If a visual indicator is provided to indicate malfunction of an instrument it must be effective under all probable cockpit lighting conditions Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5679 April 8 1970 Amdt 25 41 42 FR 36970 July 18 1977 25 1322 Warning caution and advisory lights If warning caution or advisory lights are in stalled in the cockpit they must unless otherwise approved by the Administrator be a Red for warning lights lights indicating a hazard which may require immediate corrective action b Amber for caution lights lights indicating the possible need for future corrective action c Green for safe operation lights and d Any other color including white for lights not described in paragraphs a through c of this section provided the color differs sufficiently from ASA Part 25 Airworthiness Standards Transport Category the colors prescribed in paragraphs a through c of this section to avoid possible confusion Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 825 1323 Airspeed indicating system For each airspeed indicating system the fol lowing apply a Each airspeed indicating instrument must be approved and must be calibrated to indicate true airspee
364. iate position and with no slipstream effect W seaplane design landing weight in pounds angle of dead rise at a station 3 4 of the distance from the bow to the step but need not be less than 15 degrees and ry ratio of the lateral distance between the center of gravity and the plane of symmetry of the float to the radius of gyration in roll c Bow loading The resultant limit load must be applied in the plane of symmetry of the float at a point one fourth of the distance from the bow to the step and must be perpendicular to the tangent to the keel line at that point The magnitude of the resultant load is that specified in paragraph b of this section d Unsymmetrical step loading The resultant water load consists of a component equal to 0 75 times the load specified in paragraph a of this section and a side component equal to 3 25 tan B times the load specified in paragraph b of this section The side load must be applied perpendic ularly to the plane of symmetry of the float at a point midway between the keel and the chine e Unsymmetrical bow loading The resultant water load consists of a component equal to 0 75 times the load specified in paragraph b of this section and a side component equal to 0 25 tan B times the load specified in paragraph c of this section The side load must be applied perpendic ularly to the plane of symmetry at a point midway between the keel and the chine f Immersed f
365. iated connections 7 Electrical splices 8 Materials used to provide additional protec tion for wires including wire insulation wire sleev ing and conduits that have electrical termination for the purpose of bonding 9 Shields or braids 10 Clamps and other devices used to route and support the wire bundle 11 Cable tie devices 12 Labels or other means of identification 13 Pressure seals 14 EWIS components inside shelves panels racks junction boxes distribution panels and back planes of equipment racks including but not limited to circuit board back planes wire inte gration units and external wiring of equipment b Except for the equipment indicated in para graph a 14 of this section EWIS components inside the following equipment and the external connectors that are part of that equipment are excluded from the definition in paragraph a of this section 1 Electrical equipment or avionics that are qualified to environmental conditions and testing procedures when those conditions and proce dures are i Appropriate for the intended function and op erating environment and ii Acceptable to the FAA 2 Portable electrical devices that are not part of the type design of the airplane This includes ASA 825 1705 personal entertainment devices and laptop com puters 3 Fiber optics 825 1703 Function and installation EWIS a Each EWIS component installed in any ar
366. ical engine inoperative and its propeller if applicable in the minimum drag position 1 At the minimum V2 for takeoff 2 During an approach and go around and 3 During an approach and landing d The following table prescribes for conven tional wheel type controls the maximum control forces permitted during the testing required by paragraphs a and c of this section Force in pounds applied to the control wheel or rudder pedals Pitch Roll Yaw For short term application 75 50 for pitch and roll control two hands available for control For short term application 50 25 for pitch and roll control one hand available for control For short term application 150 for yaw control For long term application 10 5 20 e Approved operating procedures or conven tional operating practices must be followed when demonstrating compliance with the control force limitations for short term application that are pre scribed in paragraph d of this section The air 23 825 143 plane must be in trim or as near to being in trim as practical in the preceding steady flight condi tion For the takeoff condition the airplane must be trimmed according to the approved operating procedures f When demonstrating compliance with the control force limitations for long term application that are prescribed in paragraph d of this sec tion the airplane must be in trim
367. ically or by manual means 3 If an instrument presenting navigation data receives information from sources external to that instrument and loss of that information would ren der the presented data unreliable the instrument must incorporate a visual means to warn the crew when such loss of information occurs that the presented data should not be relied upon b As used in this section instrument in cludes devices that are physically contained in one unit and devices that are composed of two or more physically separate units or components connected together such as a remote indicating gyroscopic direction indicator that includes a magnetic sensing element gyroscopic unit an amplifier and an indicator connected together Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 41 42 FR 36970 July 18 1977 825 1333 Instrument systems For systems that operate the instruments re quired by 25 1303 b which are located at each pilot s station a Means must be provided to connect the re quired instruments at the first pilot s station to op erating systems which are independent of the op erating systems at other flight crew stations or other equipment b The equipment systems and installations must be designed so that one display of the infor mation essential to the safety of flight which is provided by the instruments including attitude di rection airspeed and altitude will remain avail
368. ide and the outside except that sliding ASA 825 809 window emergency exits in the flight crew area need not be openable from the outside if other ap proved exits are convenient and readily accessi ble to the flight crew area Each emergency exit must be capable of being opened when there is no fuselage deformation 1 With the airplane in the normal ground atti tude and in each of the attitudes corresponding to collapse of one or more legs of the landing gear and 2 Within 10 seconds measured from the time when the opening means is actuated to the time when the exit is fully opened 3 Even though persons may be crowded against the door on the inside of the airplane c The means of opening emergency exits must be simple and obvious may not require ex ceptional effort and must be arranged and marked so that it can be readily located and oper ated even in darkness Internal exit opening means involving sequence operations such as operation of two handles or latches or the release of safety catches may be used for flightcrew emergency exits if it can be reasonably estab lished that these means are simple and obvious to crewmembers trained in their use d If a single power boost or single power oper ated system is the primary system for operating more than one exit in an emergency each exit must be capable of meeting the requirements of paragraph b of this section in the event of failure of the primary syste
369. ies of the minimum flight crew estab lished under 25 1523 This must be shown in day and night flight tests under nonprecipitation conditions b Precipitation conditions For precipitation conditions the following apply 1 The airplane must have a means to main tain a clear portion of the windshield during pre cipitation conditions sufficient for both pilots to have a sufficiently extensive view along the flight path in normal flight attitudes of the airplane This means must be designed to function without con tinuous attention on the part of the crew in i Heavy rain at speeds up to 1 5 with lift and drag devices retracted and ASA Part 25 Airworthiness Standards Transport Category ii The icing conditions specified in 825 1419 if certification for flight in icing conditions is re quested 2 The first pilot must have i A window that is openable under the condi tions prescribed in paragraph b 1 of this section when the cabin is not pressurized provides the view specified in that paragraph and gives suffi cient protection from the elements against impair ment of the pilot s vision or ii An alternate means to maintain a clear view under the conditions specified in paragraph b 1 of this section considering the probable damage due to a severe hail encounter c nternal windshield and window fogging The airplane must have a means to prevent fog ging of the internal portions of th
370. igned for a limit load factor in lateral direction for the side load on the engine and auxiliary power unit mount at least equal to the maximum load factor obtained in the yawing conditions but not less than 1 1 33 or 2 One third of the limit load factor for flight condition A as prescribed in 25 333 b b The side load prescribed in paragraph a of this section may be assumed to be independent of other flight conditions Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 91 61 FR 40704 July 29 1997 825 365 Pressurized compartment loads For airplanes with one or more pressurized compartments the following apply a The airplane structure must be strong enough to withstand the flight loads combined 41 825 365 with pressure differential loads from zero up to the maximum relief valve setting b The external pressure distribution in flight and stress concentrations and fatigue effects must be accounted for c If landings may be made with the compart ment pressurized landing loads must be com bined with pressure differential loads from zero up to the maximum allowed during landing d The airplane structure must be designed to be able to withstand the pressure differential loads corresponding to the maximum relief valve setting multiplied by a factor of 1 33 for airplanes to be approved for operation to 45 000 feet or by a factor of 1 6
371. il lumination is available from another source and 2 Be installed so that i Their direct rays are shielded from the pilot s eyes and ii No objectionable reflections are visible to the pilot b Unless undimmed instrument lights are sat isfactory under each expected flight condition there must be a means to control the intensity of illumination Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29785 July 20 1990 ASA Part 25 Airworthiness Standards Transport Category 825 1383 Landing lights a Each landing light must be approved and must be installed so that 1 No objectionable glare is visible to the pilot 2 The pilot is not adversely affected by hala tion and 3 It provides enough light for night landing b Except when one switch is used for the lights of a multiple light installation at one location there must be a separate switch for each light c There must be a means to indicate to the pilots when the landing lights are extended 825 1385 Position light system installation a General Each part of each position light System must meet the applicable requirements of this section and each system as a whole must meet the requirements of 8825 1387 through 25 1397 b Forward position lights Forward position lights must consist of red and a green light spaced laterally as far apart as practicable and in stalled forward on the airplane so tha
372. ilure of the shutoff means to stop the fuel flow at the maximum quantity approved for that tank ASA 25 981 c A means must be provided to prevent dam age to the fuel system in the event of failure of the automatic shutoff means prescribed in paragraph b of this section d The airplane pressure fueling system not including fuel tanks and fuel tank vents must withstand an ultimate load that is 2 0 times the load arising from the maximum pressures includ ing surge that is likely to occur during fueling The maximum surge pressure must be established with any combination of tank valves being either intentionally or inadvertently closed e The airplane defueling system not includ ing fuel tanks and fuel tank vents must withstand an ultimate load that is 2 0 times the load arising from the maximum permissible defueling pres sure positive or negative at the airplane fueling connection Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6913 May 5 1967 Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 72 55 FR 29785 July 20 1990 25 981 Fuel tank explosion prevention a No ignition source may be present at each point in the fuel tank or fuel tank system where catastrophic failure could occur due to ignition of fuel or vapors This must be shown by 1 Determining the highest temperature allow ing a safe margin below the lowest expected au toignition temperature of the fuel
373. ime of drip pings if any must be recorded The burn length determined in accordance with paragraph 8 of this paragraph must be measured to the nearest tenth of an inch Breaking of the wire specimens is not considered a failure 8 Burn length Burn length is the distance from the original edge to the farthest evidence of damage to the test specimen due to flame im pingement including areas of partial or complete consumption charring or embrittlement but not including areas sooted stained warped or dis colored nor areas where material has shrunk or melted away from the heat source PART 11 FLAMMABILITY OF SEAT CUSHIONS a Criteria for Acceptance Each seat cushion must meet the following criteria 1 At least three sets of seat bottom and seat back cushion specimens must be tested 168 Federal Aviation Regulations 2 If the cushion is constructed with a fire blocking material the fire blocking material must completely enclose the cushion foam core mate rial 3 Each specimen tested must be fabricated using the principal components i e foam core flotation material fire blocking material if used and dress covering and assembly processes representative seams and closures intended for use in the production articles If a different mate rial combination is used for the back cushion than for the bottom cushion both material combina tions must be tested as complete specimen sets each set consistin
374. ine inoperative and the all engines operating takeoff provisions 3 It must be shown that the one engine inop erative takeoff distance using a rotation speed of 5 knots less than Vg established in accordance with paragraphs e 1 and 2 of this section does not exceed the corresponding one engine inoperative takeoff distance using the established ASA Part 25 Airworthiness Standards Transport Category Vg The takeoff distances must be determined in accordance with 25 113 a 1 4 Reasonably expected variations in service from the established takeoff procedures for the operation of the airplane such as over rotation of the airplane and out of trim conditions may not result in unsafe flight characteristics or in marked increases in the scheduled takeoff distances es tablished in accordance with 25 113 a f Vior is the calibrated airspeed at which the airplane first becomes airborne 9 in terms of calibrated airspeed must be selected by the applicant to provide at least the gradient of climb required by 25 121 c but may not be less than 1 1 18 and 2 A speed that provides the maneuvering ca pability specified in 25 143 h h In determining the takeoff speeds V4 Vn and for flight in icing conditions the values of Vucc Vuc and determined for non icing conditions may be used Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55466 De
375. ined in determining the landing distance as prescribed in 825 125 2 Fluid lost from a brake hydraulic system fol lowing a failure in or in the vicinity of the brakes is insufficient to cause or support a hazardous fire on the ground or in flight c Brake controls The brake controls must be designed and constructed so that 1 Excessive control force is not required for their operation 2 If an automatic braking system is installed means are provided to i Arm and disarm the system and ii Allow the pilot s to override the system by use of manual braking d Parking brake The airplane must have a parking brake control that when selected on will without further attention prevent the airplane from rolling on a dry and level paved runway when the most adverse combination of maximum thrust on one engine and up to maximum ground idle thrust on any or all other engine s is applied The control must be suitably located or be ade quately protected to prevent inadvertent opera tion There must be indication in the cockpit when the parking brake is not fully released e Antiskid system If an antiskid system is in stalled 1 It must operate satisfactorily over the range of expected runway conditions without external adjustment 2 It must at all times have priority over the automatic braking system if installed f Kinetic energy capacity 1 Design landing stop The design landing stop is an opera
376. information a Airplane maintenance manual or section 1 Introduction information that includes an ex planation of the airplane s features and data to the extent necessary for maintenance or preventive maintenance 2 A description of the airplane and its sys tems and installations including its engines pro pellers and appliances 3 Basic control and operation information de Scribing how the airplane components and systems are controlled and how they operate including any special procedures and limitations that apply 4 Servicing information that covers details re garding servicing points capacities of tanks res ervoirs types of fluids to be used pressures appli cable to the various systems location of access ASA Appendix H to Part 25 panels for inspection and servicing locations of lu brication points lubricants to be used equipment required for servicing tow instructions and limita tions mooring jacking and leveling information b Maintenance instructions 1 Scheduling information for each part of the airplane and its engines auxiliary power units propellers accessories instruments and equip ment that provides the recommended periods at which they should be cleaned inspected ad justed tested and lubricated and the degree of inspection the applicable wear tolerances and work recommended at these periods However the applicant may refer to an accessory instru ment or equipment
377. ing to maxi mum continuous power at the design flap speeds Vg and with takeoff power at not less than 1 4 times the stalling speed for the particular flap po sition and associated maximum weight and 2 A head on gust of 25 feet per second veloc ity EAS c If flaps or other high lift devices are to be used in en route conditions and with flaps in the appropriate position at speeds up to the flap de 40 Federal Aviation Regulations sign speed chosen for these conditions the air plane is assumed to be subjected to symmetrical maneuvers and gusts within the range deter mined by 1 Maneuvering to a positive limit load factor as prescribed in 25 337 b and 2 The discrete vertical 25 341 d The airplane must be designed for neuvering load factor of 1 5g at the maximum take off weight with the wing flaps and similar high lift devices in the landing configurations Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50595 Oct 30 1978 Amdt 25 72 55 FR 37607 Sept 17 1990 Amdt 25 86 61 FR 5221 Feb 9 1996 Amdt 25 91 62 FR 40704 July 29 1997 525 349 Rolling conditions The airplane must be designed for loads result ing from the rolling conditions specified in para graphs a and b of this section Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the principal masses fur
378. ion ii The jacking pads and local structure must be designed for a vertical load of 2 0 times the vertical static reaction at each jacking point act ing singly and in combination with a horizontal load of 0 33 times the vertical static reaction ap plied in any direction c Tie down If tie down points are provided the main tie down points and local structure must withstand the limit loads resulting from a 65 knot horizontal wind from any direction Docket No 26129 59 FR 22102 April 28 1994 WATER LOADS 825 521 General a Seaplanes must be designed for the water loads developed during takeoff and landing with the seaplane in any attitude likely to occur in nor mal operation and at the appropriate forward and ASA 825 527 sinking velocities under the most severe sea con ditions likely to be encountered b Unless a more rational analysis of the water loads is made or the standards in ANC 3 are used 8825 523 through 25 537 apply c The requirements of this section and 5625 523 through 25 537 apply also to amphibians 825 523 Design weights and center of gravity positions a Design weights The water load require ments must be met at each operating weight up to the design landing weight except that for the takeoff condition prescribed in 825 531 the de sign water takeoff weight the maximum weight for water taxi and takeoff run must be used b Center of gravity positions The critical cen t
379. ion between gated positions and from the last gated position to the fully retracted posi tion The requirements of paragraph c of this section also apply to retractions from each ap proval landing position to the control position s associated with the high lift device configura tion s used to establish the go around proce dure s from that landing position In addition the first gated control position from the maximum landing position must correspond with a configu ration of the high lift devices used to establish a go around procedure from a landing configura tion Each gated control position must require a separate and distinct motion of the control to pass through the gated position and must have fea tures to prevent inadvertent movement of the con trol through the gated position It must only be possible to make this separate and distinct motion once the control has reached the gated position ASA 825 147 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5671 April 8 1970 Amdt 25 72 55 FR 29774 July 20 1990 Amdt 25 84 60 FR 30749 June 9 1995 Amdt 25 98 64 FR 6164 Feb 8 1999 Amdt 25 108 67 FR 70827 Nov 26 2002 825 147 Directional and lateral control a Directional control general It must be pos sible with the wings level to yaw into the opera tive engine and to safely make a reasonably sud den change in heading of up to 15 degrees in the direction of the
380. ions all combinations of altitudes and speeds encompassed by the Vp Mp versus altitude envelope enlarged at all points by an increase of 15 percent in equivalent airspeed at both constant Mach number and con 60 Federal Aviation Regulations stant altitude In addition a proper margin of sta bility must exist at all speeds up to Vp Mp and there must be no large and rapid reduction in sta bility as Vp Mp is approached The enlarged velope may be limited to Mach 1 0 when Mp is less than 1 0 at all design altitudes and 2 For the conditions described in 825 629 d below for all approved altitudes any airspeed up to the greater airspeed defined by i The Vp Mp envelope determined by 25 335 b or ii An altitude airspeed envelope defined by a 15 percent increase in equivalent airspeed above Vc at constant altitude from sea level to the alti tude of the intersection of 1 15 Vc with the exten sion of the constant cruise Mach number line Mc then a linear variation in equivalent airspeed to 05 at the altitude of the lowest Vc Mgc inter section then at higher altitudes up to the maxi mum flight altitude the boundary defined by a 05 Mach increase in Mc at constant altitude c Balance weights If concentrated balance weights are used their effectiveness and strength including supporting structure must be substanti ated d Failures malfunctions and adverse condi tions The failures malfunctions
381. ions for con tinued airworthiness ICA M25 5 RELIABILITY REPORTING The effects of airplane component failures on FRM reliability must be assessed on an on going basis The applicant holder must do the following a Demonstrate effective means to ensure col lection of FRM reliability data The means must provide data affecting FRM reliability such as component failures ASA Appendix M to Part 25 b Unless alternative reporting procedures are approved by the FAA Oversight Office as defined in part 26 of this subchapter provide a report to the FAA every six months for the first five years af ter service introduction After that period contin ued reporting every six months may be replaced with other reliability tracking methods found ac ceptable to the FAA or eliminated if it is estab lished that the reliability of the FRM meets and will continue to meet the exposure requirements of paragraph M25 1 of this appendix c Develop service instructions or revise the applicable airplane manual according to a sched ule approved by the FAA Oversight Office as de fined in part 26 of this subchapter to correct any failures of the FRM that occur in service that could increase any fuel tank s Fleet Average Flammability Exposure to more than that required by paragraph M25 1 of this appendix 231 Appendix N to Part 25 APPENDIX N TO PART 25 FUEL TANK FLAMMABILITY EXPOSURE AND RELIABILITY ANALYSIS Source Docket No
382. is chapter e A placard stating that Operations involving the carriage of persons or property for compensa SFAR 109 to Part 25 tion or hire are prohibited must be located in the area of the Airworthiness Certificate holder at the entrance to the flightdeck f For passenger capacities of 45 to 60 pas sengers analysis must be submitted that demon strates that the airplane can be evacuated in less than 90 seconds under the conditions specified in 825 803 and Appendix J to part 25 g In order for any airplane certified under this SFAR to be placed in part 135 or part 121 opera tions the airplane must be brought back into full compliance with the applicable operational part EQUIPMENT AND DESIGN 3 General Unless otherwise noted compli ance is required with the applicable certification basis for the airplane Some provisions of this SFAR impose alternative requirements to certain airworthiness standards that do not apply to air planes certificated to earlier standards Those air planes with an earlier certification basis are not required to comply with those alternative require ments 4 Occupant Protection a Firm Handhold In lieu of the requirements of 25 785 j there must be means provided to enable persons to steady themselves in moder ately rough air while occupying aisles that are along the cabin sidewall or where practicable bordered by seats seat backs providing a 25 pound minimum breakaw
383. is extremely re mote 2 Compliance with this section may be shown by failure analysis or testing or both for propeller Systems that allow propeller blades to move from the flight low pitch position to a position that is substantially less than that at the normal flight low pitch position The analysis may include or be supported by the analysis made to show compli ance with the requirements of 35 21 of this chap ter for the propeller and associated installation components Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29784 July 20 1990 96 Federal Aviation Regulations 525 934 Turbojet engine thrust reverser system tests Thrust reversers installed on turbojet engines must meet the requirements of 833 97 of this chapter Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5677 April 8 1970 525 937 Turbopropeller drag limiting systems Turbopropeller power airplane propeller drag limiting systems must be designed so that no sin gle failure or malfunction of any of the systems during normal or emergency operation results in propeller drag in excess of that for which the air plane was designed under 825 367 Failure of structural elements of the drag limiting systems need not be considered if the probability of this kind of failure is extremely remote 525 939 Turbine engine operating characteristics a Turbine engine operating characteristics mu
384. is introduced into the ullage of a fuel tank so that the ullage becomes non flammable Monte Carlo Analysis The analytical method that is specified in this appendix as the compliance means for assessing the fleet aver age flammability exposure time for a fuel tank Oxygen evolution occurs when oxygen dissolved in the fuel is released into the ullage as the pressure and temperature in the fuel tank are reduced I Standard deviation is a statistical measure of the dispersion or variation in a distribution equal to the square root of the arithmetic mean of the squares of the deviations from the arithmetic means m Transport Effects For purposes of this appendix transport effects are the change in fuel ASA Part 25 Airworthiness Standards Transport Category vapor concentration in a fuel tank caused by low fuel conditions and fuel condensation and vapor ization n Ullage The volume within the fuel tank not occupied by liquid fuel N25 3 FUEL TANK FLAMMABILITY EXPOSURE ANALYSIS a A flammability exposure analysis must be conducted for the fuel tank under evaluation to determine fleet average flammability exposure for the airplane and fuel types under evaluation For fuel tanks that are subdivided by baffles or com partments an analysis must be performed either for each section of the tank or for the section of the tank having the highest flammability expo sure Consideration of transport effects is not al
385. ith failure effects for the longest diversion time for which it seeks approval K25 1 3 Airplane systems Operation in icing conditions 1 The airplane must be certificated for opera tion in icing conditions in accordance with 825 1419 2 The airplane must be able to safely conduct an ETOPS diversion with the most critical ice ac cretion resulting from i Icing conditions encountered at an altitude that the airplane would have to fly following an en gine failure or cabin decompression ii A 15 minute hold in the continuous maxi mum icing conditions specified in Appendix C of this part with a liquid water content factor of 1 0 iii Ice accumulated during approach and land ing in the icing conditions specified in Appendix C of this part b Electrical power supply The airplane must be equipped with at least three independent Sources of electrical power c Time limited systems The applicant must define the system time capability of each ETOPS significant system that is time limited K25 1 4 Propulsion systems a Fuel system design Fuel necessary to complete an ETOPS flight including a diversion for the longest time for which the applicant seeks approval must be available to the operating en gines at the pressure and fuel flow required by 625 955 under any airplane failure condition not shown to be extremely improbable Types of fail ures that must be considered include but are not limited to crossfeed
386. ity position that results in the highest value of reference stall speed and 6 The airplane trimmed for straight flight at a speed selected by the applicant but not less than 1 13 Vsr and not greater than 1 3 c Starting from the stabilized trim condition apply the longitudinal control to decelerate the air plane so that the speed reduction does not ex ceed one knot per second d In addition to the requirements of paragraph a of this section when a device that abruptly pushes the nose down at a selected angle of at tack e g a stick pusher is installed the refer ence stall speed Vaga may not be less than 2 knots or 2 percent whichever is greater above the speed at which the device operates FAA 2002 13902 67 FR 70825 Nov 26 2002 as amended by Amdt 25 121 72 FR 44665 Aug 8 2007 825 105 Takeoff a The takeoff speeds prescribed by 825 107 the accelerate stop distance prescribed by 25 109 the takeoff path prescribed by 25 111 the takeoff distance and takeoff run prescribed by 25 113 and the net takeoff flight path prescribed 15 825 107 by 825 115 must be determined in the selected configuration for takeoff at each weight altitude and ambient temperature within the operational limits selected by the applicant 1 In non icing conditions and 2 In icing conditions if in the configuration of 25 121 b with the takeoff ice accretion defined in appendix C i The stall
387. ive the remaining engines at the go around power or thrust setting ii The maximum landing weight iii A climb speed established in connection with normal landing procedures but not exceed ing 1 4 and iv Landing gear retracted 2 The requirements of paragraph d 1 of this section must be met i In non icing conditions and ii In icing conditions with the approach ice ac cretion defined in appendix C The climb speed selected for non icing conditions may be used if the climb speed for icing conditions computed in accordance with paragraph d 1 iii of this sec tion does not exceed that for non icing conditions by more than the greater of 3 knots CAS or 3 per cent Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amat 25 84 60 FR 30749 June 9 1995 Amdt 25 108 67 FR 70826 Nov 26 2002 Amdt 25 121 72 FR 44666 Aug 8 2007 25 123 En route flight paths a For the en route configuration the flight paths prescribed in paragraph b and c of this section must be determined at each weight alti tude and ambient temperature within the operat ing limits established for the airplane The varia tion of weight along the flight path accounting for 22 Federal Aviation Regulations the progressive consumption of fuel and oil by the operating engines may be included in the compu tation The flight paths must be determined at a speed not less than Veto with 1 The
388. ived as follows Surface K Position of controls a Aileron 0 75 Control column locked or lashed in mid position b do 11 0 50 Ailerons at full throw c Elevator 1140 75 c Elevator full down d do 1140 75 d Elevator full up e Rudder 0 75 e Rudder in neutral f do 0 75 f Rudder at full throw 1A positive value of K indicates a moment tend ing to depress the surface while a negative value of K indicates a moment tending to raise the sur face Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29776 July 20 1990 Amdt 25 91 62 FR 40705 July 29 1997 25 427 Unsymmetrical loads a In designing the airplane for lateral gust yaw maneuver and roll maneuver conditions ac count must be taken of unsymmetrical loads on the empennage arising from effects such as slip stream and aerodynamic interference with the wing vertical fin and other aerodynamic surfaces b The horizontal tail must be assumed to be subjected to unsymmetrical loading conditions determined as follows 1 100 percent of the maximum loading from the symmetrical maneuver conditions of 25 331 and the vertical gust conditions of 25 341 a act ing separately on the surface on one side of the plane of symmetry and 2 80 percent of these loadings acting on the other side c For empennage arrangements where the horizontal tail surfaces have dihedral angles greater than
389. ix provides the flash point for the standard fuel to be used in the analysis e Fleet average flammability exposure is the percentage of the flammability exposure eval uation time FEET each fuel tank ullage is flam mable for a fleet of an airplane type operating over the range of flight lengths in a world wide range of environmental conditions and fuel prop erties as defined in this appendix f Gaussian Distribution is another name for the normal distribution a symmetrical frequency distribution having a precise mathematical for mula relating the mean and standard deviation of the samples Gaussian distributions yield bell shaped frequency curves having a preponder ance of values around the mean with progres sively fewer observations as the curve extends outward g Hazardous atmosphere An atmosphere that may expose maintenance personnel pas sengers or flight crew to the risk of death inca pacitation impairment of ability to self rescue that is escape unaided from a confined space injury or acute illness h Inert For the purpose of this appendix the tank is considered inert when the bulk average oxygen concentration within each compartment of the tank is 12 percent or less from sea level up to 10 000 feet altitude then linearly increasing from 12 percent at 10 000 feet to 14 5 percent at 40 000 feet altitude and extrapolated linearly above that altitude i Inerting process where a noncombusti ble gas
390. jury Deviating from the no yaw condition may not result in the critical area of contact not being eval uated The upper torso restraint straps where in stalled must remain on the occupant s shoulder during the impact condition of 25 562 b 2 c For the vertical test conducted in accor dance with the conditions specified 25 562 b 1 Hybrid II ATDs or equivalent must be used in all seat positions 5 Direct View In lieu of the requirements of 25 785 h 2 to the extent practical without com promising proximity to a required floor level emer gency exit the majority of installed flight attendant seats must be located to face the cabin area for which the flight attendant is responsible 6 Passenger Information Signs Compliance with 25 791 is required except that for 25 791 a when smoking is to be prohibited no tification to the passengers may be provided by a single placard so stating to be conspicuously lo cated inside the passenger compartment easily visible to all persons entering the cabin in the im mediate vicinity of each passenger entry door 7 Distance Between Exits For an airplane that is required to comply with 25 807 f 4 in ef fect as of July 24 1989 which has more than one passenger emergency exit on each side of the fu selage no passenger emergency exit may be more than 60 feet from any adjacent passenger emergency exit on the same side of the same deck of the fuselage as measured
391. kage or excessive deformation of the tank walls 1 Each complete tank assembly and its sup ports must be vibration tested while mounted to simulate the actual installation 2 Except as specified in paragraph b 4 of this section the tank assembly must be vibrated for 25 hours at an amplitude of not less than 19 of an inch unless another amplitude is substanti ated while 2 filled with water or other suitable test fluid 3 The test frequency of vibration must be as follows i If no frequency of vibration resulting from any r p m within the normal operating range of engine speeds is critical the test frequency of vi bration must be 2 000 cycles per minute ii If only one frequency of vibration resulting from any r p m within the normal operating range of engine speeds is critical that frequency of vi bration must be the test frequency iii If more than one frequency of vibration re sulting from any r p m within the normal operat ing range of engine speeds is critical the most critical of these frequencies must be the test fre quency 4 Under paragraphs b 3 ii and iii of this section the time of test must be adjusted to ac complish the same number of vibration cycles that would be accomplished in 25 hours at the frequency specified in paragraph b 3 i of this section 5 During the test the tank assembly must be rocked at the rate of 16 to 20 complete cycles per minute through an angle of 15
392. l The airplane must be longitudinally direction ally and laterally stable in accordance with the provisions of 25 173 through 25 177 In addi tion suitable stability and control feel static sta bility is required in any condition normally en countered in service if flight tests show it is nec essary for safe operation Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 7 30 FR 13117 Oct 15 1965 27 825 173 825 173 Static longitudinal stability Under the conditions specified in 825 175 the characteristics of the elevator control forces in cluding friction must be as follows a A pull must be required to obtain and main tain speeds below the specified trim speed and a push must be required to obtain and maintain speeds above the specified trim speed This must be shown at any speed that can be obtained ex cept speeds higher than the landing gear or wing flap operating limit speeds or whichever is appropriate or lower than the minimum speed for steady unstalled flight b The airspeed must return to within 10 per cent of the original trim speed for the climb ap proach and landing conditions specified in 825 175 c and d and must return to within 7 5 percent of the original trim speed for the cruis ing condition specified in 25 175 b when the control force is slowly released from any speed within the range specified in paragraph a of this section
393. l Aviation Regulations FIGURE 4 Sliding Platform oO 13 3 16 in 335 mm 43 3 4 in 4 7 16 in 1 2 in 63 5 mm 1 4 in 6 35 mm thick refractory material 43 mm bolts ii Attach a 1 2 inch 13 mm piece of Kaowool board or other high temperature material measuring 41 1 2 by 8 1 4 inches 1054 by 210 mm to the back of the platform This board Serves as a heat retainer and protects the test specimen from excessive preheating The height of this board must not impede the sliding platform movement in and out of the test chamber If the platform has been fabricated such that the back side of the platform is high enough to prevent ex cess preheating of the specimen when the sliding platform is out a retainer board is not necessary iii Place the test specimen horizontally on the non combustible board s Place a steel retain ing securing frame fabricated of mild steel having a thickness of 1 8 inch 3 2 mm and overall di mensions of 23 by 13 1 8 inches 584 by 333 mm with a specimen opening of 19 by 10 3 4 inches 483 by 273 mm over the test specimen The front back and right portions of the top flange of the frame must rest on the top of the sliding plat form and the bottom flanges must pinch all 4 sides of the test specimen The right bottom flange must be flush with the sliding platform See figure 5 194 ASA Part 25 Airworthiness Standards Transport Category
394. l ground attitude of not less than the greater of 0 10 percent of the tank capacity or one sixteenth of a gallon unless operating limitations are established to ensure that the accumulation of water in service will not exceed the sump capacity b Each fuel tank must allow drainage of any hazardous quantity of water from any part of the tank to its sump with the airplane in the ground at titude c Each fuel tank sump must have an accessi ble drain that 1 Allows complete drainage of the sump on the ground 2 Discharges clear of each part of the air plane and 3 Has manual or automatic means for positive locking in the closed position 25 973 Fuel tank filler connection Each fuel tank filler connection must prevent the entrance of fuel into any part of the airplane other than the tank itself In addition a Reserved b Each recessed filler connection that can re tain any appreciable quantity of fuel must have a drain that discharges clear of each part of the air plane c Each filler cap must provide a fuel tight seal and d Each fuel filling point except pressure fuel ing connection points must have a provision for electrically bonding the airplane to ground fueling equipment Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15043 March 17 1977 Amdt 25 72 55 FR 29785 July 20 1990 25 975 Fuel tank vents and carburetor vapor vents a Fuel tank vent
395. l must be designed to retract the surfaces from the fully extended posi 63 825 699 tion during steady flight at maximum continuous engine power at any speed below 9 0 knots Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5675 April 8 1970 Amdt 25 46 43 FR 50595 Oct 30 1978 Amdt 25 57 49 FR 6848 Feb 23 1984 525 699 Lift and drag device indicator a There must be means to indicate to the pilots the position of each lift or drag device hav ing separate control in the cockpit to adjust its position In addition an indication of unsymmetri cal operation or other malfunction in the lift or drag device systems must be provided when such indication is necessary to enable the pilots to pre vent or counteract an unsafe flight or ground con dition considering the effects on flight character istics and performance b There must be means to indicate to the pilots the takeoff en route approach and landing lift device positions c If any extension of the lift and drag devices beyond the landing position is possible the con trols must be clearly marked to identify this range of extension Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5675 April 8 1970 825 701 Flap and slat interconnection a Unless the airplane has safe flight charac teristics with the flaps or slats retracted on one Side and extended on the other the m
396. le Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5674 April 8 1970 825 672 Stability augmentation and automatic and power operated systems If the functioning of stability augmentation or other automatic or power operated systems is necessary to show compliance with the flight characteristics requirements of this part such systems must comply with 825 671 and the fol lowing a A warning which is clearly distinguishable to the pilot under expected flight conditions without requiring his attention must be provided for any failure in the stability augmentation system or in any other automatic or power operated system which could result in an unsafe condition if the pilot were not aware of the failure Warning sys tems must not activate the control systems b The design of the stability augmentation System or of any other automatic or power oper ated system must permit initial counteraction of failures of the type specified in 525 671 without requiring exceptional pilot skill or strength by ei ther the deactivation of the system or a failed por tion thereof or by overriding the failure by move ment of the flight controls in the normal sense c It must be shown that after any single failure of the stability augmentation system or any other automatic or power operated system 1 The airplane is safely controllable when the failure or malfunction occurs at any speed or alti tude
397. le improper operation of closure devices and inadvertent door openings are also considered Furthermore the resulting differential pressure loads must be combined in a rational and conservative manner with 1 g level flight 42 Federal Aviation Regulations loads and any loads arising from emergency de pressurization conditions These loads may be considered as ultimate conditions however any deformations associated with these conditions must not interfere with continued safe flight and landing The pressure relief provided by intercom partment venting may also be considered g Bulkheads floors and partitions in pressur ized compartments for occupants must be de signed to withstand the conditions specified in paragraph e of this section In addition reason able design precautions must be taken to mini mize the probability of parts becoming detached and injuring occupants while in their seats Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 54 45 FR 60172 Sept 11 1980 Amdt 25 71 55 FR 13477 April 10 1990 Amdt 25 72 55 FR 29776 July 20 1990 Amdt 25 87 61 FR 28695 June 5 1996 825 367 Unsymmetrical loads due to engine failure a The airplane must be designed for the un symmetrical loads resulting from the failure of the critical engine Turbopropeller airplanes must be designed for the following conditions in combina tion with a single malfunction of the propeller drag limiting
398. le tail surfaces must be installed so that there is no interference between any surfaces when one is held in its extreme position and the others are operated through their full angular movement b If an adjustable stabilizer is used it must have stops that will limit its range of travel to the maximum for which the airplane is shown to meet the trim requirements of 825 161 825 657 Hinges a For control surface hinges including ball roller and self lubricated bearing hinges the ap proved rating of the bearing may not be ex ceeded For nonstandard bearing hinge configu rations the rating must be established on the ba sis of experience or tests and in the absence of a rational investigation a factor of safety of not less than 6 67 must be used with respect to the ulti mate bearing strength of the softest material used as a bearing b Hinges must have enough strength and ri gidity for loads parallel to the hinge line Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5674 April 8 1970 CONTROL SYSTEMS 25 671 General a Each control and control system must oper ate with the ease smoothness and positiveness appropriate to its function b Each element of each flight control system must be designed or distinctively and perma nently marked to minimize the probability of in correct assembly that could result in the malfunc tioning of the system c The airplane must be sho
399. less than 1 2 percent for two engine airplanes 1 5 percent for three engine airplanes and 1 7 percent for four engine airplanes at Veto with 21 825 123 i The critical engine inoperative and the re maining engines at the available maximum contin uous power or thrust and ii The weight equal to the weight existing at the end of the takeoff path determined under 825 111 2 The requirements of paragraph c 1 of this section must be met i In non icing conditions and ii In icing conditions with the final takeoff ice accretion defined in appendix C if in the configu ration of 825 121 b with the takeoff ice accretion A The stall speed at maximum takeoff weight exceeds that in non icing conditions by more than the greater of 3 knots CAS percent of Vs or B The degradation of the gradient of climb de termined in accordance with 825 121 b is greater than one half of the applicable actual to net take off flight path gradient reduction defined in 25 115 b d Approach a configuration corresponding to the normal all engines operating procedure in which for this configuration does not exceed 110 percent of the Vsp for the related all engines operating landing configuration 1 The steady gradient of climb may not be less than 2 1 percent for two engine airplanes 2 4 percent for three engine airplanes and 2 7 percent for four engine airplanes with i The critical engine inoperat
400. lexible intake duct 4 Test specimen mounting frame Make the mounting frame for the test specimens of 1 8 inch 3 2 mm thick steel as shown in figure 1 except for the center vertical former which should be 1 4 inch 6 4 mm thick to minimize warpage The specimen mounting frame stringers horizontal should be bolted to the test frame formers verti cal such that the expansion of the stringers will not cause the entire structure to warp Use the mounting frame for mounting the two insulation blanket test specimens as shown in figure 2 5 Backface calorimeters Mount two total heat flux Gardon type calorimeters behind the insula tion test specimens on the back side cold area of the test specimen mounting frame as shown in figure 6 Position the calorimeters along the same plane as the burner cone centerline at a distance of 4 inches 102 mm from the vertical centerline of the test frame SEE FIGURE 6 AT THE END OF PART VII OF THIS APPENDIX i The calorimeters must be a total heat flux foil type Gardon Gage of an appropriate range such as 0 5 Btu ft sec 0 5 7 W cm accurate to 3 of the indicated reading The heat flux cali bration method must comply with paragraph VI b 7 of this appendix 6 Instrumentation Provide a recording poten tiometer or other suitable calibrated instrument with an appropriate range to measure and record the outputs of the calorimeter and the thermocou ples 7 Timing device Provid
401. licant must prepare Instructions for Continued Airworthiness applicable to EWIS in accordance with Appendix H sections H25 4 and H25 5 to this part that are approved by the FAA 825 1731 Powerplant and APU fire detector system EWIS a EWIS that are part of each fire or overheat detector system in a fire zone must be fire resis tant b No EWIS component of any fire or overheat detector system for any fire zone may pass through another fire zone unless 1 It is protected against the possibility of false warnings resulting from fires in zones through which it passes or 2 Each zone involved is simultaneously pro tected by the same detector and extinguishing system c EWIS that are part of each fire or overheat detector system in a fire zone must meet the re quirements of 825 1203 146 Federal Aviation Regulations 825 1733 Fire detector systems general EWIS EWIS associated with any installed fire protec tion system including those required by 8825 854 and 25 858 must be considered an integral part of the system in showing compliance with the ap plicable requirements for that system ASA Appendix to Part 25 Part 25 Airworthiness Standards Transport Category APPENDIX TO PART 25 uoisueuuip 1 Buipue oiseg FYNOIS T33HM 3SON AdAL T33HM 147 ASA Federal Aviation Regulations Appendix A to Part 25 T33HM 3SON
402. loat condition The resultant load must be applied at the centroid of the cross sec tion of the float at a point one third of the distance from the bow to the step The limit load compo nents are as follows vertical pg 50 2 3 2 aft KV go 2 3 side CPV KV 5 54 Federal Aviation Regulations where p mass density of water slugs ft V volume of float ft 2 coefficient of drag force equal to 0 133 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 are incapable of submerging at a speed of 0 8 Vso in normal operations Vso seaplane stalling speed knots with landing flaps extended in the appropriate position and with no slipstream effect and g acceleration due to gravity ft sec 2 g Float bottom pressures The float bottom pressures must be established under 825 533 except that the value of in the formulae may be taken as 1 0 The angle of dead rise to be used in determining the float bottom pressures is set forth in paragraph b of this section Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 825 537 Seawing loads Seawing design loads must be based on appli cable test data EMERGENCY LANDING CONDITIONS 825 561 General a The airplane although it may be damaged in emergency landing conditions on land or water must be
403. looping tendency in 90 cross winds up to a wind velocity of 20 knots or 0 2 Vsgo whichever is greater except that the wind velocity need not ex ceed 25 knots at any speed at which the airplane may be expected to be operated on the ground This may be shown while establishing the 90 cross component of wind velocity required by 825 237 b Landplanes must be satisfactorily controlla ble without exceptional piloting skill or alertness in power off landings at normal landing speed without using brakes or engine power to maintain a straight path This may be shown during power off landings made in conjunction with other tests c The airplane must have adequate direc tional control during taxiing This may be shown during taxiing prior to takeoffs made in conjunc tion with other tests Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5671 April 8 1970 Amdt 25 42 43 FR 2322 Jan 16 1978 Amdt 25 94 63 FR 8848 Feb 23 1998 Amdt 25 108 67 FR 70828 Nov 26 2002 31 825 235 825 235 Taxiing condition The shock absorbing mechanism may not damage the structure of the airplane when the air plane is taxied on the roughest ground that may reasonably be expected in normal operation 525 237 Wind velocities a For land planes and amphibians the follow ing applies 1 A 90 degree cross component of wind ve locity demonstrated to be safe for takeoff and landing must be
404. low fire or flame penetration in less than 4 minutes 2 Each of the two insulation blanket test spec imens must not allow more than 2 0 Btu ft sec 2 27 W cm on the cold side of the insulation specimens at a point 12 inches 30 5 cm from the face of the test rig SEE PART VII FIGURES BEGINNING ON THE NEXT PAGE ASA Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 FIGURE 1 Burnthrough Test Apparatus Specimen Holder Detail A 10 mm Horizontal Hat shaped Stringers Bolted to Vertical Formers Detail B To 0 75 1 19 mm 25 mm 0 5 n 13 mm All Material 0 123 3 mm Thickness Except Center Vertical Former 0 250 6 mm Thick ASA 203 Appendix to Part 25 FIGURE 2 Burnthrough Test Apparatus Federal Aviation Regulations Insulation 4 Pa 102mm Burner Cone 4 sage HN a 77 gt Draft Tube Test Frame 204 ASA Part 25 Airworthiness Standards Transport Category FIGURE 3 Appendix F to Part 25 Burner Draft Tube Extension Cone Diagram 15 50 394 mm 5 0 127 mm ee 10 Degree bends on Broken Line 12 0 305 mm 27 64 11 mm Spaces 0 05 13 12 25 311 3 75 95 r 7 50 Note One half of tube 191 mm extension shown Second half mates at spotweld overlaps gt 19 32 15 mm
405. ly 20 1990 139 825 1553 825 1553 Fuel quantity indicator If the unusable fuel supply for any tank exceeds one gallon or five percent of the tank capacity whichever is greater a red arc must be marked on its indicator extending from the calibrated zero reading to the lowest reading obtainable in level flight 825 1555 Control markings a Each cockpit control other than primary flight controls and controls whose function is obvi ous must be plainly marked as to its function and method of operation b Each aerodynamic control must be marked under the requirements of 825 677 and 25 699 c For powerplant fuel controls 1 Each fuel tank selector control must be marked to indicate the position corresponding to each tank and to each existing cross feed position 2 If safe operation requires the use of any tanks in a specific sequence that sequence must be marked on or adjacent to the selector for those tanks and 3 Each valve control for each engine must be marked to indicate the position corresponding to each engine controlled d For accessory auxiliary and emergency controls 1 Each emergency control including each fuel jettisoning and fluid shutoff must be colored red and 2 Each visual indicator required by 25 729 e must be marked so that the pilot can determine at any time when the wheels are locked in either ex treme position if retractable landing gear is used 825 1557 Misc
406. m Manual operation of the exit after failure of the primary system is acceptable e Each emergency exit must be shown by tests or by a combination of analysis and tests to meet the requirements of paragraphs b and c of this section f Each door must be located where persons using them will not be endangered by the propel lers when appropriate operating procedures are used g There must be provisions to minimize the probability of jamming of the emergency exits re sulting from fuselage deformation in a minor crash landing h When required by the operating rules for any large passenger carrying turbojet powered airplane each ventral exit and tailcone exit must be 1 Designed and constructed so that it cannot be opened during flight and 2 Marked with a placard readable from a dis tance of 30 inches and installed at a conspicuous location near the means of opening the exit stat ing that the exit has been designed and con structed so that it cannot be opened during flight i Each emergency exit must have a means to retain the exit in the open position once the exit is opened in an emergency The means must not re quire separate action to engage when the exit is 79 825 810 opened and must require positive action to disen gage Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 15 32 FR 13264 Sept 20 1967 Amdt 25 32 37 FR 3970 Feb 24 1972 Amdt 25 34 37 FR 25355 No
407. m an overspeed condition at Vpp Mppr to produce at least 1 5 g for recovery by applying not more than 125 pounds of longitudinal control force using ei ther the primary longitudinal control alone or the primary longitudinal control and the longitudinal trim system If the longitudinal trim is used to as sist in producing the required load factor it must be shown at Vpre Mpr that the longitudinal trim can be actuated in the airplane nose up direction with the primary surface loaded to correspond to the least of the following airplane nose up control forces 34 Federal Aviation Regulations 1 The maximum control forces expected in Service as specified in 25 301 and 25 397 2 The control force required to produce 1 5 g 3 The control force corresponding to buffeting or other phenomena of such intensity that it is a strong deterrent to further application of primary longitudinal control force Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2322 Jan 16 1978 ASA Part 25 Airworthiness Standards Transport Category Subpart C Structure GENERAL 825 301 Loads a Strength requirements are specified in terms of limit loads the maximum loads to be ex pected in service and ultimate loads limit loads multiplied by prescribed factors of safety Unless otherwise provided prescribed loads are limit loads b Unless otherwise provided the specified air ground and water loads
408. m intensities in the horizontal plane minimum intensities in any vertical plane and maximum in tensities in overlapping beams within dihedral an gles L R and A and must meet the following re quirements 1 Intensities in the horizontal plane Each in tensity in the horizontal plane the plane contain ing the longitudinal axis of the airplane and per pendicular to the plane of symmetry of the air plane must equal or exceed the values in 25 1391 2 Intensities in any vertical plane Each inten sity in any vertical plane the plane perpendicular to the horizontal plane must equal or exceed the appropriate value in 25 1393 where is the min imum intensity prescribed in 25 1391 for the cor responding angles in the horizontal plane 3 Intensities in overlaps between adjacent signals No intensity in any overlap between cent signals may exceed the values given in 25 1395 except that higher intensities in over laps may be used with main beam intensities sub stantially greater than the minima specified in 25 1391 and 25 1393 if the overlap intensities in relation to the main beam intensities do not ad versely affect signal clarity When the peak inten sity of the forward position lights is more than 100 candles the maximum overlap intensities be tween them may exceed the values given in 125 825 1391 825 1395 if the overlap intensity in Area is not more than 10 percent of peak position lig
409. m may be used for auxiliary power units fuel burning heaters and other combustion equip ment For each other designated fire zone two discharges must be provided each of which pro duces adequate agent concentration c The fire extinguishing system for a nacelle must be able to simultaneously protect each zone of the nacelle for which protection is provided Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50598 Oct 30 1978 825 1197 Fire extinguishing agents a Fire extinguishing agents must 1 Be capable of extinguishing flames emanat ing from any burning of fluids or other combustible materials in the area protected by the fire extin guishing system and 2 Have thermal stability over the temperature range likely to be experienced in the compartment in which they are stored b If any toxic extinguishing agent is used pro visions must be made to prevent harmful concen trations of fluid or fluid vapors from leakage dur ing normal operation of the airplane or as a result of discharging the fire extinguisher on the ground or in flight from entering any personnel compart ment even though a defect may exist in the extin guishing system This must be shown by test ex cept for built in carbon dioxide fuselage compart ment fire extinguishing systems for which 1 Five pounds or less of carbon dioxide will be discharged under established fire control proce dures into any fus
410. mbers if the malfunctioning of any part of the electrical system is causing the contin uous discharge of any battery necessary for en gine ignition h Each engine ignition system of a turbine powered airplane must be considered an essen tial electrical load Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5677 April 8 1970 Amdt 25 72 55 FR 29785 July 20 1990 825 1167 Accessory gearboxes For airplanes equipped with an accessory gear box that is not certificated as part of an engine a The engine with gearbox and connecting transmissions and shafts attached must be sub jected to the tests specified in 33 49 or 33 87 of this chapter as applicable b The accessory gearbox must meet the re quirements of 8833 25 and 33 53 or 33 91 of this chapter as applicable and c Possible misalignments and torsional load ings of the gearbox transmission and shaft sys tem expected to result under normal operating conditions must be evaluated Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 ASA 825 1183 POWERPLANT FIRE PROTECTION 825 1181 Designated fire zones regions included a Designated fire zones are 1 The engine power section 2 The engine accessory section 3 Except for reciprocating engines any com plete powerplant compartment in which no isola tion is provided between the engine power section and
411. me aft cen ter of gravity limitations must be established for each practicably separable operating condition No such limit may lie beyond a The extremes selected by the applicant b The extremes within which the structure is proven or c The extremes within which compliance with each applicable flight requirement is shown 25 29 Empty weight and corresponding center of gravity a The empty weight and corresponding center of gravity must be determined by weighing the air plane with 1 Fixed ballast 2 Unusable fuel determined under 825 959 and 3 Full operating fluids including i Oil ii Hydraulic fluid and iii Other fluids required for normal operation of airplane systems except potable water lavatory precharge water and fluids intended for injection in the engine b The condition of the airplane at the time of determining empty weight must be one that is well defined and can be easily repeated Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2320 Jan 16 1978 Amdt 25 72 55 FR 29774 July 20 1990 25 31 Removable ballast Removable ballast may be used on showing compliance with the flight requirements of this subpart 25 33 Propeller speed and pitch limits a The propeller speed and pitch must be lim ited to values that will ensure 1 Safe operation under normal operating con ditions and 2 Compliance with the performanc
412. mental conditions must be considered b Radio and electronic equipment must be supplied with power under the requirements of 25 1355 c Radio and electronic equipment controls and wiring must be installed so that operation of any one unit or system of units will not adversely affect the simultaneous operation of any other ra dio or electronic unit or system of units required by this chapter d Electronic equipment must be designed and installed such that it does not cause essential loads to become inoperative as a result of electri cal power supply transients or transients from other causes Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Docket Nos FAA 2001 9634 FAA 2001 9633 FAA 2001 9638 FAA 2001 9637 Amdt 25 113 69 FR 12529 March 16 2004 25 1433 Vacuum systems There must be means in addition to the normal pressure relief to automatically relieve the pres sure in the discharge lines from the vacuum air pump when the delivery temperature of the air be comes unsafe Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29785 July 20 1990 825 1435 Hydraulic systems a Element design Each element of the hy draulic system must be designed to 1 Withstand the proof pressure without per manent deformation that would prevent it from performing its intended functions and the ultimate pressure without rupture The proof and ultimate pressures are
413. ments and control systems replacement of parts normally requiring replacement adjustment and lubrication as necessary for continued airworthi ness The inspection means for each item must be practicable for the inspection interval for the item Nondestructive inspection aids may be used to inspect structural elements where it is impracti cable to provide means for direct visual inspection if it is shown that the inspection is effective and the inspection procedures are specified in the maintenance manual required by 825 1529 b EWIS must meet the accessibility require ments of 825 1719 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5674 April 8 1970 Amdt 25 123 72 FR 63404 Nov 8 2007 25 613 Material strength properties and material design values a Material strength properties must be based on enough tests of material meeting approved specifications to establish design values on a sta tistical basis b Material design values must be chosen to minimize the probability of structural failures due to material variability Except as provided in para graphs e and f of this section compliance must be shown by selecting material design values which assure material strength with the following probability 1 Where applied loads are eventually distrib uted through a single member within an assem bly the failure of which would result in loss of structural integrity of the compo
414. most unfavorable center of gravity 2 The critical engines inoperative 3 The remaining engines at the available maximum continuous power or thrust and 4 The means for controlling the engine cooling air supply in the position that provides adequate cooling in the hot day condition b The one engine inoperative net flight path data must represent the actual climb performance diminished by a gradient of climb of 1 1 percent for two engine airplanes 1 4 percent for three en gine airplanes and 1 6 percent for four engine airplanes 1 In non icing conditions and 2 In icing conditions with the en route ice ac cretion defined in appendix C if i A speed of 1 18 Vsp with the en route ice ac cretion exceeds the en route speed selected for non icing conditions by more than the greater of 3 knots CAS 3 percent of Vsp or ii The degradation of the gradient of climb is greater than one half of the applicable actual to net flight path reduction defined in paragraph b of this section c For three or four engine airplanes the two engine inoperative net flight path data must represent the actual climb performance dimin ished by a gradient of climb of 0 3 percent for three engine airplanes and 0 5 percent for four engine airplanes Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 121 72 FR 44666 Aug 8 2007 825 125 Landing a The horizontal distance necessary to land and to come to
415. munications of other crewmembers on the flight deck when directed to those stations The microphone must be so located and if neces sary the preamplifiers and filters of the recorder must be so adjusted or supplemented that the in telligibility of the recorded communications is as high as practicable when recorded under flight cockpit noise conditions and played back Re 133 825 1457 peated aural or visual playback of the record may be used in evaluating intelligibility c Each cockpit voice recorder must be in stalled so that the part of the communication or audio signals specified in paragraph a of this section obtained from each of the following Sources is recorded on a separate channel 1 For the first channel from each boom mask or hand held microphone headset or speaker used at the first pilot station 2 For the second channel from each boom mask or hand held microphone headset or speaker used at the second pilot station 3 For the third channel from the cockpit mounted area microphone 4 For the fourth channel from i Each boom mask or hand held micro phone headset or speaker used at the station for the third and fourth crew members or ii If the stations specified in paragraph c 4 i of this section are not required or if the signal at such a station is picked up by another channel each microphone on the flight deck that is used with the passenger loudspeaker system if its
416. must be condi tioned as described in Part 1 of this appendix 3 Mounting Each test specimen must be wrapped tightly on all sides of the specimen ex cept for the one surface that is exposed with a sin gle layer of 001 inch 025 mm aluminum foil e Procedure 1 The power supply to the radiant panel must be set to produce a radiant flux of 3 5 05 W cm as measured at the point the center of the speci men surface will occupy when positioned for the test The radiant flux must be measured after the air flow through the equipment is adjusted to the desired rate 2 After the pilot flames are lighted their posi tion must be checked as described in paragraph b 8 of this part IV 3 Air flow through the apparatus must be con trolled by a circular plate orifice located in a 1 5 inch 38 1 mm 1 0 pipe with two pressure mea suring points located 1 5 inches 38 mm up stream and 75 inches 19 mm downstream of the orifice plate The pipe must be connected to a manometer set at a pressure differential of 7 87 inches 200 mm of Hg See Figure 1B of this part IV The total air flow to the equipment is ap proximately 04 m3 seconds The stop on the ver tical specimen holder rod must be adjusted so that the exposed surface of the specimen is posi tioned 3 9 inches 100 mm from the entrance when injected into the environmental chamber _ Ei 0108722 273 P mole CH4STP WATT min 760 mole Ta
417. must be designed for deflections consistent with the primary control surface loading conditions obtainable within the pilot maneuvering effort considering possible op position from the trim tabs 825 415 Ground gust conditions a The control system must be designed as fol lows for control surface loads due to ground gusts and taxiing downwind 1 The control system between the stops near est the surfaces and the cockpit controls must be designed for loads corresponding to the limit hinge moments H of paragraph a 2 of this sec tion These loads need not exceed i The loads corresponding to the maximum pilot loads in 825 397 c for each pilot alone or ii 0 75 times these maximum loads for each pilot when the pilot forces are applied in the same direction 2 The control system stops nearest the sur faces the control system locks and the parts of the systems if any between these stops and locks and the control surface horns must be de signed for limit hinge moments H in foot pounds obtained from the formula 0034KV2cS where V 65 wind speed in knots K limit hinge moment factor for ground gusts derived in paragraph b of this section c mean chord of the control surface aft of the hinge line ft S area of the control surface aft of the hinge line sq ft ASA Part 25 Airworthiness Standards Transport Category b The limit hinge moment factor K for ground gusts must be der
418. must be designed and installed so that in the event of failures of the elec trical supply or control system the requirements of 825 1309 b c and d will be satisfied Do mestic appliances are items such as cooktops ovens coffee makers water heaters refrigera tors and toilet flush systems that are placed on the airplane to provide service amenities to pas sengers b Galleys and cooking appliances must be in stalled in a way that minimizes risk of overheat or fire c Domestic appliances particularly those in galley areas must be installed or protected so as to prevent damage or contamination of other equipment or systems from fluids or vapors which may be present during normal operation or as a result of spillage if such damage or contamina tion could create a hazardous condition d Unless compliance with 25 1309 b is pro vided by the circuit protective device required by 25 1357 a electric motors and transformers in cluding those installed in domestic systems must have a suitable thermal protection device to pre vent overheating under normal operation and fail ure conditions if overheating could create a smoke or fire hazard Docket No FAA 2004 18379 72 FR 63406 Nov 8 2007 LIGHTS 825 1381 Instrument lights a The instrument lights must 1 Provide sufficient illumination to make each instrument switch and other device necessary for safe operation easily readable unless sufficient
419. must be placed in equilibrium with inertia forces considering each item of mass in the airplane These loads must be distributed to conservatively approximate closely represent actual conditions Methods used to determine load intensities and distribution must be validated by flight load measurement un less the methods used for determining those loading conditions are shown to be reliable c If deflections under load would significantly change the distribution of external or internal loads this redistribution must be taken into ac count Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 825 303 Factor of safety Unless otherwise specified a factor of safety of 1 5 must be applied to the prescribed limit load which are considered external loads on the struc ture When a loading condition is prescribed in terms of ultimate loads a factor of safety need not be applied unless otherwise specified Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 825 305 Strength and deformation a The structure must be able to support limit loads without detrimental permanent deformation At any load up to limit loads the deformation may not interfere with safe operation b The structure must be able to support ulti mate loads without failure for at least 3 seconds However when proof of strength is shown by dy namic tests simulati
420. n and b If the inherent flight characteristics of the airplane do not provide adequate warning that an engine has failed a warning system that is inde pendent of the ATTCS must be provided to give the pilot a clear warning of any engine failure dur ing takeoff Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 62 52 FR 43156 Nov 9 1987 219 Appendix J to Part 25 APPENDIX J TO PART 25 EMERGENCY DEMONSTRATION The following test criteria and procedures must be used for showing compliance with 825 803 a The emergency evacuation must be con ducted with exterior ambient light levels of no greater than 0 3 foot candles prior to the activa tion of the airplane emergency lighting system The source s of the initial exterior ambient light level may remain active or illuminated during the actual demonstration There must however be no increase in the exterior ambient light level except for that due to activation of the airplane emer gency lighting system b The airplane must be in normal attitude with landing gear extended c Unless the airplane is equipped with an off wing descent means stands or ramps may be used for descent from the wing to the ground Safety equipment such as mats or inverted life rafts may be placed on the floor or ground to pro tect participants No other equipment that is not part of the emergency evacuation equipment of the airplane may be used to aid the participants
421. n 5 seconds and 4 For propeller airplanes hazardous flight characteristics must not be exhibited due to any propeller position achieved when the engine fails or during any likely subsequent movements of the engine or propeller controls Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2321 Jan 16 1978 Amdt 25 72 55 FR 29774 July 20 1990 55 FR 37607 Sept 12 1990 Amdt 25 84 60 FR 30750 June 9 1995 Amdt 25 108 67 FR 70827 Nov 26 2002 TRIM 825 161 Trim a General Each airplane must meet the trim requirements of this section after being trimmed and without further pressure upon or movement of either the primary controls or their correspond ing trim controls by the pilot or the automatic pilot b Lateral and directional trim The airplane must maintain lateral and directional trim with the most adverse lateral displacement of the center of gravity within the relevant operating limitations during normally expected conditions of operation ASA 825 171 including operation at any speed from 1 3 to c Longitudinal trim The airplane must main tain longitudinal trim during 1 A climb with maximum continuous power at a speed not more than 1 3 with the landing gear retracted and the flaps i retracted and ii in the takeoff position 2 A glide with power off at a speed not more than 1 3 with the landing gear extended the
422. n of the pilot burner to the test specimen b Test apparatus 192 Federal Aviation Regulations FIGURE 1 Radiant Panel Test Chamber 55 in 1397 mm Glass Viewing Window 1 Radiant panel test chamber Conduct tests in a radiant panel test chamber see figure 1 above Place the test chamber under an exhaust hood to facilitate clearing the chamber of smoke after each test The radiant panel test chamber must be an enclosure 55 inches 1397 mm long by 19 5 495 mm deep by 28 710 mm to 30 inches maximum 762 mm above the test spec imen Insulate the sides ends and top with a fi brous ceramic insulation such as Kaowool MTM board On the front side provide a 52 by 12 inch 1321 by 305 mm draft free high temperature glass window for viewing the sample during test ing Place a door below the window to provide ac cess to the movable specimen platform holder The bottom of the test chamber must be a sliding steel platform that has provision for securing the test specimen holder in a fixed and level position The chamber must have an internal chimney with exterior dimensions of 5 1 inches 129 mm wide by 16 2 inches 411 mm deep by 13 inches 330 mm high at the opposite end of the chamber from the radiant energy source The interior dimen sions must be 4 5 inches 114 mm wide by 15 6 inches 395 mm deep The chimney must extend to the top of the chamber see fi
423. n other types of airplanes to the extent such in formation is reasonably available c Airplane flight test The applicant must con duct a flight test to validate the flightcrew s ability to safely conduct an ETOPS diversion with an in operative engine and worst case ETOPS signifi cant system failures and malfunctions that could occur in service The flight test must validate the airplane s flying qualities and performance with the demonstrated failures and malfunctions K25 3 2 Early ETOPS method An applicant for ETOPS type design approval using the Early ETOPS method must comply with the following requirements a Maintenance and operational procedures The applicant must validate all maintenance and operational procedures for ETOPS significant Systems The applicant must identify track and re solve any problems found during the validation in accordance with the problem tracking and resolu tion system specified in section K25 3 2 e of this appendix b New technology testing Technology new to the applicant including substantially new manu facturing techniques must be tested to substanti ate its suitability for the airplane design c APU validation test If an APU is needed to comply with this appendix one APU of the type to be certified with the airplane must be tested for 3 000 equivalent airplane operational cycles Fol lowing completion of the test the APU must be disassembled and inspected The applicant must id
424. n set is the weight of the specimen set before testing less the weight of the specimen set after testing expressed as the percentage of the weight before testing b Test Conditions Vertical air velocity should average 25 10 fpm at the top of the back seat cushion Horizontal air velocity should be be low 10 fpm just above the bottom seat cushion Air velocities should be measured with the ventila tion hood operating and the burner motor off c Test Specimens 1 For each test one set of cushion specimens representing a seat bottom and seat back cushion must be used 2 The seat bottom cushion specimen must be 18 inches 457 3 mm wide by 20 ASA Part 25 Airworthiness Standards Transport Category inches 508 3 mm deep by 4 inches 102 3 mm thick exclusive of fabric closures and seam overlap 3 The seat back cushion specimen must be 18 inches 432 3 mm wide by 25 inches 635 3 mm high by 2 inches 51 3 mm thick exclusive of fabric closures and seam overlap 4 The specimens must be conditioned at 70 5 21 2 55 10 relative humidity for at least 24 hours before testing d Test Apparatus The arrangement of the test apparatus is shown in Figures 1 through 5 and must include the components described in this section Minor details of the apparatus may vary depending on the model burner used 1 Specimen Mounting Stand The mounting stand for the test specimens consists of st
425. n system ASA Appendix K to Part 25 1 The applicant must establish and maintain a problem tracking and resolution system The sys tem must i Contain a process for prompt reporting to the responsible FAA aircraft certification office of each occurrence reportable under 21 4 a 6 en countered during the phases of airplane and en gine development used to assess Early ETOPS eligibility ii Contain a process for notifying the responsi ble FAA aircraft certification office of each pro posed corrective action that the applicant deter mines necessary for each problem identified from the occurrences reported under section K25 3 2 h 1 i of this appendix The timing of the notification must permit appropriate FAA review before taking the proposed corrective action 2 If the applicant is seeking ETOPS type de sign approval of a change to an airplane engine combination previously approved for ETOPS the problem tracking and resolution system need only address those problems specified in the following table provided the applicant obtains prior authori zation from the FAA If the change does not require a new airplane type certificate and Then the Problem Tracking and Resolution System must address i Requires a new engine type certificate problems applicable to the new engine installation and for the remainder of the airplane problems in changed systems only ii Does not require
426. n the output side of the control system for example stalling torque or maximum rate obtain able by a power control system 1 Maximum pitch control displacement at Va The airplane is assumed to be flying in steady level flight point A1 25 333 b and the cockpit pitch control is suddenly moved to obtain extreme nose up pitching acceleration In defining the tail load the response of the airplane must be taken into account Airplane loads that occur subse quent to the time when normal acceleration at the c g exceeds the positive limit maneuvering load factor at point in 25 333 b or the resulting tailplane normal load reaches its maximum whichever occurs first need not be considered 2 Specified control displacement checked maneuver based on a rational pitching control motion vs time profile must be established in which the design limit load factor specified in 825 337 will not be exceeded Unless lesser val ues cannot be exceeded the airplane response must result in pitching accelerations not less than the following i A positive pitching acceleration nose up is assumed to be reached concurrently with the air plane load factor of 1 0 Points to Dy 25 333 b The positive acceleration must be equal to at least 32n n 1 5 Radians sec 2 v where n is the positive load factor at the speed under consideration and V is the airplane equivalent speed in knots ii A negative pit
427. nce on each side of the fuselage 8 Emergency Exit Signs In lieu of the re quirements of 25 811 d 1 and 2 a single sign at each exit may be installed provided a The sign can be read from the aisle while di rectly facing the exit and b The sign can be read from the aisle adja cent to the passenger seat that is farthest from the exit and that does not have an intervening bulkhead divider or exit 9 Emergency Lighting a Exit Signs In lieu of the requirements of 25 812 b 1 for airplanes that have a passen ger seating configuration excluding pilot seats of 19 seats or less the emergency exit signs re quired by 825 811 d 1 2 and 3 must have red letters at least 1 inch high on a white back ground at least 2 inches high These signs may be internally electrically illuminated or self illumi nated by other than electrical means with an ini tial brightness of at least 160 microlamberts The color may be reversed in the case of a sign that is self illuminated by other than electrical means b Floor Proximity Escape Path Marking In lieu of the requirements of 25 812 e 1 for cabin seating compartments that do not have the main cabin aisle entering and exiting the compart ment the following are applicable 1 After a passenger leaves any passenger seat in the compartment he she must be able to exit the compartment to the main cabin aisle us ing only markings and visual features not more that 4 f
428. nd maintain straight flight with an angle of bank of not more than 5 degrees VmcL must be established with 1 The airplane in the most critical configuration or at the option of the applicant each configura tion for approach and landing with all engines op erating 2 The most unfavorable center of gravity 3 The airplane trimmed for approach with all engines operating 4 The most favorable weight or at the option of the applicant as a function of weight 5 For propeller airplanes the propeller of the inoperative engine in the position it achieves with out pilot action assuming the engine fails while at the power or thrust necessary to maintain a three degree approach path angle and 6 Go around power or thrust setting on the operating engine s g For airplanes with three or more engines VmcL 2 the minimum control speed during ap proach and landing with one critical engine inop erative is the calibrated airspeed at which when a second critical engine is suddenly made inoper ative it is possible to maintain control of the air plane with both engines still inoperative and maintain straight flight with an angle of bank of not more than 5 degrees Vyc 2 must be established with 1 The airplane in the most critical configura tion or at the option of the applicant each config ASA Part 25 Airworthiness Standards Transport Category uration for approach and landing with one critical e
429. nd or pressure demand pressure demand mask with a diluter demand ASA Part 25 Airworthiness Standards Transport Category pressure breathing regulator type or other ap proved oxygen equipment shown to provide the same degree of protection for airplanes to be oper ated above 25 000 feet ii The pressure demand pressure demand mask with a diluter demand pressure breathing regulator type with mask mounted regulator or other approved oxygen equipment shown to pro vide the same degree of protection for airplanes operated at altitudes where decompressions that are not extremely improbable may expose the flightcrew to cabin pressure altitudes in excess of 34 000 feet 4 Portable oxygen equipment must be imme diately available for each cabin attendant The portable oxygen equipment must have the oxygen dispensing unit connected to the portable oxygen supply Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 41 42 FR 36971 July 18 1977 Amdt 25 87 61 FR 28696 June 5 1996 Amdt 25 116 69 FR 62789 Oct 27 2004 25 1449 Means for determining use of oxygen There must be a means to allow the crew to de termine whether oxygen is being delivered to the dispensing equipment 825 1450 Chemical oxygen generators a For the purpose of this section a chemical oxygen generator is defined as a device which produces oxygen by chemical reaction b Each chemical oxygen generator must be desi
430. nd test a minimum of three test specimens If an oriented film cover material is used prepare and test both the warp and fill directions 2 Construction Test specimens must include all materials used in construction of the insulation including batting film scrim tape etc Cut a piece of core material such as foam or fiberglass and cut a piece of film cover material if used large enough to cover the core material Heat sealing is the preferred method of preparing fiber glass samples since they can be made without compressing the fiberglass box sample Cover materials that are not heat sealable may be sta pled sewn or taped as long as the cover material is over cut enough to be drawn down the sides without compressing the core material The fas tening means should be as continuous as possi ble along the length of the seams The specimen thickness must be of the same thickness as in stalled in the airplane ASA 3 Specimen Dimensions To facilitate proper placement of specimens in the sliding platform housing cut non rigid core materials such as fi berglass 12 1 2 inches 318mm wide by 23 inches 584mm long Cut rigid materials such as foam 11 1 2 1 4 inches 292 mm 6mm wide by 23 inches 584mm long in order to fit properly in the sliding platform housing and provide a flat exposed surface equal to the opening in the hous ing d Specimen conditioning Condition the test specimens at 70 5 21
431. nds after deployment is begun Assisting means installed at Type C exits must be automati cally erected within 10 seconds from the time the opening means of the exit is actuated iii It must be of such length after full deploy ment that the lower end is self supporting on the ground and provides safe evacuation of occu pants to the ground after collapse of one or more legs of the landing gear iv It must have the capability in 25 knot winds directed from the most critical angle to deploy and with the assistance of only one person to re main usable after full deployment to evacuate oc cupants safely to the ground v For each system installation mockup or air plane installed five consecutive deployment and inflation tests must be conducted per exit with out failure and at least three tests of each such five test series must be conducted using a single representative sample of the device The sample devices must be deployed and inflated by the sys tem s primary means after being subjected to the 80 Federal Aviation Regulations inertia forces specified in 25 561 b If any part of the system fails or does not function properly during the required tests the cause of the failure or malfunction must be corrected by positive means and after that the full series of five con secutive deployment and inflation tests must be conducted without failure 2 The assisting means for flightcrew emer gency exits may be a ro
432. ne 2 Mount the test specimen on the test stand shown in Figure 1 in either the horizontal or verti cal position Mount the insulating material in the other position ASA Appendix F to Part 25 3 Position the burner so that flames will not impinge on the specimen turn the burner on and allow it to run for 2 minutes Rotate the burner to apply the flame to the specimen and simulta neously start the timing device 4 Expose the test specimen to the flame for 5 minutes and then turn off the burner The test may be terminated earlier if flame penetration is ob served 5 When testing ceiling liner panels record the peak temperature measured 4 inches above the sample 6 Record the time at which flame penetration occurs if applicable h Test Report The test report must include the following 1 A complete description of the materials tested including type manufacturer thickness and other appropriate data 2 Observations of the behavior of the test specimens during flame exposure such as delam ination resin ignition smoke etc including the time of such occurrence 3 The time at which flame penetration occurs if applicable for each of the three specimens tested 4 Panel orientation ceiling or sidewall SEE FIGURES BEGINNING ON THE NEXT PAGE 177 Appendix to Part 25 Federal Aviation Regulations N HORIZONTAL AND VERTICAL SPECIMENS ARE HORIZONTAL SPEC CLAMPED IN PLACE
433. nent 99 percent probability with 95 percent confidence 2 For redundant structure in which the failure of individual elements would result in applied loads being safely distributed to other load carry ing members 90 percent probability with 95 per cent confidence ASA Part 25 Airworthiness Standards Transport Category c The effects of environmental conditions such as temperature and moisture on material design values used in an essential component or structure must be considered where these effects are significant within the airplane operating enve lope d Reserved e Greater material design values may be used if a premium selection of the material is made in which a specimen of each individual item is tested before use to determine that the actual strength properties of that particular item will equal or exceed those used in design f Other material design values may be used if approved by the Administrator Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50595 Oct 30 1978 Amdt 25 72 55 FR 29776 July 20 1990 Amdt 25 112 68 FR 46430 Aug 5 2003 825 619 Special factors The factor of safety prescribed in 825 303 must be multiplied by the highest pertinent special fac tor of safety prescribed in 25 621 through 25 625 for each part of the structure whose strength is a Uncertain b Likely to deteriorate in service before nor mal replacement or
434. ner reflective of the installed arrangement For example if the ma terial will be placed on the outboard side of the in sulation material inside the moisture film place it the same way in the test specimen ii Insulation material Blankets that utilize more than one variety of insulation composition density etc must have specimen sets con structed that reflect the insulation combination used If however several blanket types use simi lar insulation combinations it is not necessary to test each combination if it is possible to bracket the various combinations iii Moisture barrier film lf a production blanket construction utilizes more than one type of mois ture barrier film perform separate tests on each combination For example if a polyimide film is used in conjunction with an insulation in order to enhance the burnthrough capabilities also test the same insulation when used with a polyvinyl fluoride film iv Installation on test frame Attach the blan ket test specimens to the test frame using 12 steel spring type clamps as shown in figure 7 Use the clamps to hold the blankets in place in both of the outer vertical formers as well as the center verti cal former 4 clamps per former The clamp sur faces should measure 1 inch by 2 inches 25 by 51 mm Place the top and bottom clamps 6 inches 15 2 cm from the top and bottom of the test frame respectively Place the middle clamps 8 inches 20 3 cm from the to
435. ng the landing impact must be considered in deter mining the maximum design loads for the fore and aft wheel pairs 2 Any differentials in tire diameters resulting from a combination of manufacturing tolerances tire growth and tire wear A maximum tire diame ter differential equal to 2 3 of the most unfavorable combination of diameter variations that is ob tained when taking into account manufacturing tolerances tire growth and tire wear may be as sumed 3 Any unequal tire inflation pressure assum ing the maximum variation to be 5 percent of the nominal tire inflation pressure 4 A runway crown of zero and a runway crown having a convex upward shape that may be ap proximated by a slope of 11 5 percent with the hor 50 izontal Runway crown effects must be consid ered with the nose gear unit on either slope of the crown 5 The airplane attitude 6 Any structural deflections c Deflated tires The effect of deflated tires on the structure must be considered with respect to the loading conditions specified in paragraphs d through f of this section taking into account the physical arrangement of the gear components In addition 1 The deflation of any one tire for each multi ple wheel landing gear unit and the deflation of any two critical tires for each landing gear unit us ing four or more wheels per unit must be consid ered and 2 The ground reactions must be applied to the wheels with in
436. ng actual load conditions the 3 second limit does not apply Static tests con ducted to ultimate load must include the ultimate deflections and ultimate deformation induced by the loading When analytical methods are used to show compliance with the ultimate load strength requirements it must be shown that 1 The effects of deformation are not significant 2 The deformations involved are fully ac counted for in the analysis or 3 The methods and assumptions used are suf ficient to cover the effects of these deformations ASA 825 321 c Where structural flexibility is such that any rate of load application likely to occur in the oper ating conditions might produce transient stresses appreciably higher than those corresponding to static loads the effects of this rate of application must be considered d Reserved e The airplane must be designed to withstand any vibration and buffeting that might occur in any likely operating condition up to Vp Mp including stall and probable inadvertent excursions beyond the boundaries of the buffet onset envelope This must be shown by analysis flight tests or other tests found necessary by the Administrator f Unless shown to be extremely improbable the airplane must be designed to withstand any forced structural vibration resulting from any fail ure malfunction or adverse condition in the flight control system These must be considered limit loads and must be investi
437. ng condition and 4 System transients due to switching fault clearing or other causes do not make essential loads inoperative and do not cause a smoke or fire hazard 5 There are means accessible in flight to ap propriate crewmembers for the individual and col lective disconnection of the electrical power Sources from the system 6 There are means to indicate to appropriate crewmembers the generating system quantities essential for the safe operation of the system such as the voltage and current supplied by each generator c External power If provisions are made for connecting external power to the airplane and that external power can be electrically connected to equipment other than that used for engine starting means must be provided to ensure that no external power supply having a reverse polar ity or a reverse phase sequence can supply power to the airplane s electrical system d Operation without normal electrical power It must be shown by analysis tests or both that ASA Part 25 Airworthiness Standards Transport Category the airplane can be operated safely in VFR condi tions for a period of not less than five minutes with the normal electrical power electrical power Sources excluding the battery inoperative with critical type fuel from the standpoint of flameout and restart capability and with the airplane ini tially at the maximum certificated altitude Parts of the electrical system may
438. ng gear unit and the application of appropriate drag loads during the test must simulate the airplane landing condi tions in a manner consistent with the development of rational or conservative limit loads b The landing gear may not fail in a test dem onstrating its reserve energy absorption capacity simulating a descent velocity of 12 f p s at design landing weight assuming airplane lift not greater than airplane weight acting during the landing im pact c In lieu of the tests prescribed in this section changes in previously approved design weights and minor changes in design may be substanti ated by analyses based on previous tests con ducted on the same basic landing gear system that has similar energy absorption characteristics Docket No FAA 1999 5835 66 FR 27394 May 16 2001 825 725 Reserved 825 727 Reserved 825 729 Retracting mechanism a General For airplanes with retractable land ing gear the following apply 1 The landing gear retracting mechanism wheel well doors and supporting structure must be designed for i The loads occurring in the flight conditions when the gear is in the retracted position ii The combination of friction loads inertia loads brake torque loads air loads and gyro Scopic loads resulting from the wheels rotating at a peripheral speed equal to 1 3 Vs with the flaps in takeoff position at design takeoff weight oc curring during retraction and extension
439. ng limitations ap plicable to the operations in which it is utilized 5 Reference Unless otherwise provided all references in this regulation to Part 4a and Part 4b are those parts of the Civil Air Regulations in effect on September 1 1953 This regulation supersedes Special Civil Air Regulation SR 398 and shall remain effective un til superseded or rescinded by the Board 19 FR 5039 Aug 11 1954 Redesignated at 29 FR 19099 Dec 30 1964 ASA SFAR No 109 to Part 25 SFAR No 109 PART 25 1 Applicability Contrary provisions of 14 CFR parts 21 25 and 119 of this chapter notwith standing an applicant is entitled to an amended type certificate or supplemental type certificate in the transport category if the applicant complies with all applicable provisions of this SFAR OPERATIONS 2 General a The passenger capacity may not exceed 60 If more than 60 passenger seats are installed then 1 If the extra seats are not suitable for occu pancy during taxi takeoff and landing each extra seat must be clearly marked e g a placard on the top of an armrest or a placard sewn into the top of the back cushion that the seat is not to be occupied during taxi takeoff and landing 2 If the extra seats are suitable for occupancy during taxi takeoff and landing i e meet all the strength and passenger injury criteria in part 25 then a note must be included in the Limitations Section of the Airplane Fli
440. ngine inoperative 2 The most unfavorable center of gravity 3 The airplane trimmed for approach with one critical engine inoperative 4 The most unfavorable weight or at the op tion of the applicant as a function of weight 5 For propeller airplanes the propeller of the more critical inoperative engine in the position it achieves without pilot action assuming the en gine fails while at the power or thrust necessary to maintain a three degree approach path angle and the propeller of the other inoperative engine feathered 6 The power or thrust on the operating en gine s necessary to maintain an approach path angle of three degrees when one critical engine is inoperative and 7 The power or thrust on the operating en gine s rapidly changed immediately after the second critical engine is made inoperative from the power or thrust prescribed in paragraph g 6 of this section to i Minimum power or thrust and ii Go around power or thrust setting h In demonstration of Vict and 2 1 The rudder force may not exceed 150 pounds 2 The airplane may not exhibit hazardous flight characteristics or require exceptional pilot ing skill alertness or strength 3 Lateral control must be sufficient to roll the airplane from an initial condition of steady flight through an angle of 20 degrees in the direction necessary to initiate a turn away from the inoper ative engine s in not more tha
441. ngines and the mixture settings must be those normally used in the flight stages for which the cooling tests are conducted The test procedures must be as prescribed in 825 1045 b Maximum ambient atmospheric tempera ture maximum ambient atmospheric tempera ture corresponding to sea level conditions of at least 100 degrees F must be established The as sumed temperature lapse rate is 3 6 degrees F per thousand feet of altitude above sea level until a temperature of 69 7 degrees F is reached above which altitude the temperature is consid ered constant at 69 7 degrees F However for winterization installations the applicant may se lect a maximum ambient atmospheric tempera ture corresponding to sea level conditions of less than 100 degrees F c Correction factor except cylinder barrels Unless a more rational correction applies temper atures of engine fluids and powerplant compo nents except cylinder barrels for which tempera ture limits are established must be corrected by adding to them the difference between the maxi mum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum component or fluid temperature recorded during the cooling test d Correction factor for cylinder barrel temper atures Unless a more rational correction applies cylinder barrel temperatures must be corrected by adding to them 0 7 times the difference between the maximum
442. not ex ceed 27 inches 4 Type IV This type is a rectangular opening of less than 19 inches wide by 26 inches high with corner radii not greater than 6 3 inches lo cated over the wing with a step up inside the air plane of not more than 29 inches and a step down outside the airplane of not more than 36 inches 5 Ventral This type is an exit from the passen ger compartment through the pressure shell and the bottom fuselage skin The dimensions and physical configuration of this type of exit must al low at least the same rate of egress as a Type I exit with the airplane in the normal ground atti tude with landing gear extended 6 Tailcone This type is an aft exit from the passenger compartment through the pressure shell and through an openable cone of the fuse lage aft of the pressure shell The means of open ing the tailcone must be simple and obvious and must employ a single operation 7 Type A This type is a floor level exit with a rectangular opening of not less than 42 inches wide by 72 inches high with corner radii not greater than seven inches 8 Type B This type is a floor level exit with a rectangular opening of not less than 32 inches wide by 72 inches high with corner radii not greater than six inches 9 Type C This type is a floor level exit with a rectangular opening of not less than 30 inches wide by 48 inches high with corner radii not greater than 10 inches b Step down distance Step
443. ns disengagement of any automatic control function of a flight guidance system may not result in a transient any greater than a significant tran sient as defined in paragraph n 2 of this sec tion f The function and direction of motion of each command reference control such as heading se lect or vertical speed must be plainly indicated on or adjacent to each control if necessary to prevent inappropriate use or confusion g Under any condition of flight appropriate to its use the flight guidance system may not pro duce hazardous loads on the airplane nor create hazardous deviations in the flight path This ap plies to both fault free operation and in the event of a malfunction and assumes that the pilot be gins corrective action within a reasonable period of time h When the flight guidance system is in use a means must be provided to avoid excursions be yond an acceptable margin from the speed range of the normal flight envelope If the airplane expe riences an excursion outside this range a means ASA Part 25 Airworthiness Standards Transport Category must be provided to prevent the flight guidance System from providing guidance or control to an unsafe speed i The flight guidance system functions con trols indications and alerts must be designed to minimize flightcrew errors and confusion concern ing the behavior and operation of the flight guid ance system Means must be provided to indicate th
444. ntal plane of forward and rear position lights Minimum intensities in any vertical plane of forward and rear position lights Maximum intensities in overlapping beams of forward and rear position lights Color specifications Riding light Anticollision light system Wing icing detection lights SAFETY EQUIPMENT General Ditching equipment Ice protection Megaphones Public address system MISCELLANEOUS EQUIPMENT Electronic equipment Vacuum systems Hydraulic systems Pressurization and pneumatic systems Protective breathing equipment Oxygen equipment and supply Minimum mass flow of supplemental oxygen Equipment standards for the oxygen distributing system Equipment standards for oxygen dispensing units Means for determining use of oxygen Chemical oxygen generators Protection of oxygen equipment from rupture Draining of fluids subject to freezing Cockpit voice recorders Flight data recorders Equipment containing high energy rotors Subpart G Operating Limitations and Information 25 1501 25 1503 25 1505 25 1507 25 1511 25 1513 25 1515 25 1516 25 1517 ASA General OPERATING LIMITATIONS Airspeed limitations general Maximum operating limit speed Maneuvering speed Flap extended speed Minimum control speed Landing gear speeds Other speed limitations Rough air speed 25 1519 25 1521 25 1522 25 1523 25 1525 25 1527 25 1529
445. ntifica tion method and requirements for identifying any changes to EWIS under 825 1711 5 Electrical load data and instructions for up dating that data b The EWIS ICA developed in accordance with the requirements of H25 5 a 1 must be in the form of a document appropriate for the infor mation to be provided and they must be easily recognizable as EWIS ICA This document must either contain the required EWIS ICA or specifi cally reference other portions of the ICA that con tain this information Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 54 45 FR 60177 Sept 11 1980 Amdt 25 68 54 FR 34329 Aug 18 1989 Amdt 25 102 66 FR 23130 May 7 2001 Amdt 25 123 72 FR 63408 Nov 8 2007 ASA Part 25 Airworthiness Standards Transport Category APPENDIX TO PART 25 INSTALLATION OF AN AUTOMATIC TAKEOFF THRUST CONTROL SYSTEM ATTCS 125 1 GENERAL a This appendix specifies additional require ments for installation of an engine power control System that automatically resets thrust or power on operating engine s in the event of any one en gine failure during takeoff b With the ATTCS and associated systems functioning normally as designed all applicable requirements of Part 25 except as provided in this appendix must be met without requiring any action by the crew to increase thrust or power 125 2 DEFINITIONS a Automatic Takeoff Thrust Control System ATTCS An ATTCS i
446. nts flooded the buoyancy of the hull and auxiliary floats and wheel tires if used pro vides a margin of positive stability great enough to minimize the probability of capsizing in rough fresh water b Bulkheads with watertight doors may be used for communication between compartments PERSONNEL AND CARGO ACCOMMODATIONS 25 771 Pilot compartment a Each pilot compartment and its equipment must allow the minimum flight crew established under 25 1523 to perform their duties without unreasonable concentration or fatigue b The primary controls listed in 25 779 a excluding cables and control rods must be lo cated with respect to the propellers so that no member of the minimum flight crew established under 25 1523 or part of the controls lies in the region between the plane of rotation of any in board propeller and the surface generated by a line passing through the center of the propeller hub making an angle of five degrees forward or aft of the plane of rotation of the propeller c If provision is made for a second pilot the airplane must be controllable with equal safety from either pilot seat 68 Federal Aviation Regulations d The pilot compartment must be constructed so that when flying in rain or snow it will not leak in a manner that will distract the crew or harm the structure e Vibration and noise characteristics of cock pit equipment may not interfere with safe opera tion of the airpl
447. o 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5678 April 8 1970 Amdt 25 40 42 FR 15044 March 17 1977 825 1201 Fire extinguishing system materials a No material in any fire extinguishing system may react chemically with any extinguishing agent So as to create a hazard b Each system component in an engine com partment must be fireproof 825 1203 Fire detector system a There must be approved quick acting fire or overheat detectors in each designated fire zone and in the combustion turbine and tailpipe sec tions of turbine engine installations in numbers and locations ensuring prompt detection of fire in those zones b Each fire detector system must be con structed and installed so that 1 It will withstand the vibration inertia and other loads to which it may be subjected in oper ation 2 There is a means to warn the crew in the event that the sensor or associated wiring within a designated fire zone is severed at one point un 114 Federal Aviation Regulations less the system continues to function as a satis factory detection system after the severing and 3 There is a means to warn the crew in the event of a short circuit in the sensor or associated wiring within a designated fire zone unless the System continues to function as a satisfactory de tection system after the short circuit c No fire or overheat detector may be affected by any oil water ot
448. o im pact of a failed or released blade and the unbal ance created by such failure or release Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 54 45 FR 60173 Sept 11 1980 Amdt 25 57 49 FR 6848 Feb 23 1984 Amdt 25 72 55 FR 29784 July 20 1990 Amdt 25 126 73 FR 63345 Oct 24 2008 825 907 Propeller vibration and fatigue This section does not apply to fixed pitch wood propellers of conventional design a The applicant must determine the magni tude of the propeller vibration stresses or loads including any stress peaks and resonant condi tions throughout the operational envelope of the airplane by either 1 Measurement of stresses or loads through direct testing or analysis based on direct testing of the propeller on the airplane and engine installa tion for which approval is sought or 2 Comparison of the propeller to similar pro pellers installed on similar airplane installations for which these measurements have been made b The applicant must demonstrate by tests analysis based on tests or previous experience on similar designs that the propeller does not ex perience harmful effects of flutter throughout the operational envelope of the airplane c The applicant must perform an evaluation of the propeller to show that failure due to fatigue will be avoided throughout the operational life of the ASA 825 929 propeller using the fatigue and structural data ob tain
449. o metal latching device j If the seat backs do not provide a firm hand hold there must be a handgrip or rail along each aisle to enable persons to steady themselves while using the aisles in moderately rough air k Each projecting object that would injure per sons seated or moving about the airplane in nor mal flight must be padded I Each forward observer s seat required by the operating rules must be shown to be suitable for use in conducting the necessary enroute inspec tion Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29780 July 20 1990 Amdt 25 88 61 FR 57956 Nov 8 1996 825 787 Stowage compartments a Each compartment for the stowage of cargo baggage carry on articles and equipment such as life rafts and any other stowage com partment must be designed for its placarded max imum weight of contents and for the critical load distribution at the appropriate maximum load fac tors corresponding to the specified flight and ground load conditions and to the emergency landing conditions of 25 561 b except that the forces specified in the emergency landing condi tions need not be applied to compartments lo cated below or forward of all occupants in the air plane If the airplane has a passenger seating configuration excluding pilots seats of 10 seats or more each stowage compartment in the pas senger cabin except for underseat and overhead compartments for pa
450. o that the holes are placed above the specimen as shown in Figure 1B of this part IV The fuel supplied to the burner must be methane mixed with air in a ratio of approximately 50 50 by volume The total gas flow must be adjusted to produce flame lengths of 1 inch 25 mm When the gas air ratio and the flow rate are properly adjusted approximately 25 inch 6 mm of the flame length appears yellow in color c Calibration of Equipment 1 Heat Release Rate A calibration burner as shown in Figure 4 must be placed over the end of the lower pilot flame tubing using a gas tight con nection The flow of gas to the pilot flame must be at least 99 percent methane and must be accu rately metered Prior to usage the wet test meter must be properly leveled and filled with distilled water to the tip of the internal pointer while no gas is flowing Ambient temperature and pressure of the water are based on the internal wet test meter temperature A baseline flow rate of approxi mately 1 liter min must be set and increased to higher preset flows of 4 6 8 6 and 4 liters min Immediately prior to recording methane flow rates a flow rate of 8 liters min must be used for 2 minutes to precondition the chamber This is not recorded as part of calibration The rate must be determined by using a stopwatch to time a com plete revolution of the wet test meter for both the baseline and higher flow with the flow returned to baseline before changing
451. of any fuel shutoff valve for any engine may not make fuel unavailable to the re maining engines c Operation of any shutoff may not interfere with the later emergency operation of other equip ment such as the means for feathering the pro peller d Each flammable fluid shutoff means and control must be fireproof or must be located and protected so that any fire in a fire zone will not af fect its operation e No hazardous quantity of flammable fluid may drain into any designated fire zone after shutoff f There must be means to guard against inad vertent operation of the shutoff means and to make it possible for the crew to reopen the shutoff means in flight after it has been closed g Each tank to engine shutoff valve must be located so that the operation of the valve will not be affected by powerplant or engine mount struc tural failure h Each shutoff valve must have a means to relieve excessive pressure accumulation unless a means for pressure relief is otherwise provided in the system ASA Part 25 Airworthiness Standards Transport Category Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5677 April 8 1970 Amdt 25 57 49 FR 6849 Feb 23 1984 825 1191 Firewalls a Each engine auxiliary power unit fuel burning heater other combustion equipment in tended for operation in flight and the combustion turbine and tailpipe sections of turbine engines must be i
452. of the hull or main float center line in accordance with figure 3 of Appendix B These pressures are uniform and must be applied simultaneously over the entire hull or main float bottom The loads obtained must be carried into the sidewall structure of the hull proper but need not be transmitted in a fore and aft direction as shear and bending loads Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 25 535 Auxiliary float loads a General Auxiliary floats and their attach ments and supporting structures must be de signed for the conditions prescribed in this sec tion In the cases specified in paragraphs b through e of this section the prescribed water loads may be distributed over the float bottom to avoid excessive local loads using bottom pres sures not less than those prescribed in paragraph g of this section b Step loading The resultant water load must be applied in the plane of symmetry of the float at a point three fourths of the distance from the bow to the step and must be perpendicular to the keel The resultant limit load is computed as follows except that the value of L need not exceed three 53 825 537 times the weight of the displaced water when the float is completely submerged C 2 3 273 2 tan e Try where L limit load Ibs Cs 0 0053 Vso seaplane stalling speed knots with landing flaps extended in the appropr
453. of the takeoff roll if the airplane is in a configuration including any of the following that would not allow a safe takeoff 1 The wing flaps or leading edge devices are not within the approved range of takeoff positions 2 Wing spoilers except lateral control spoil ers meeting the requirements of 825 671 speed brakes or longitudinal trim devices are in a posi tion that would not allow a safe takeoff b The warning required by paragraph a of this section must continue until 1 The configuration is changed to allow a safe takeoff 2 Action is taken by the pilot to terminate the takeoff roll 3 The airplane is rotated for takeoff or 4 The warning is manually deactivated by the pilot c The means used to activate the system must function properly throughout the ranges of takeoff weights altitudes and temperatures for which certification is requested Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 42 43 FR 2323 Jan 16 1978 LANDING GEAR 825 721 General a The main landing gear system must be de signed so that if it fails due to overloads during takeoff and landing assuming the overloads to act in the upward and aft directions the failure mode is not likely to cause 1 For airplanes that have passenger seating configuration excluding pilots seats of nine seats or less the spillage of enough fuel from any fuel System in the fuselage to constitute a fire hazard
454. ol to warn flight crewmembers against jettisoning fuel while the means that change the airflow are being used i The fuel jettisoning system must be de signed so that any reasonably probable single malfunction in the system will not result in a haz ardous condition due to unsymmetrical jettisoning of or inability to jettison fuel Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 18 33 FR 12226 Aug 30 1968 Amdt 25 57 49 FR 6848 Feb 23 1984 Amdt 25 108 67 FR 70828 Nov 26 2002 OIL SYSTEM 25 1011 General a Each engine must have an independent oil system that can supply it with an appropriate quantity of oil at a temperature not above that safe for continuous operation b The usable oil capacity may not be less than the product of the endurance of the airplane under critical operating conditions and the ap proved maximum allowable oil consumption of the engine under the same conditions plus a suitable margin to ensure system circulation Instead of a rational analysis of airplane range for the purpose of computing oil requirements for reciprocating engine powered airplanes the following fuel oil ratios may be used 1 For airplanes without a reserve oil or oil transfer system a fuel oil ratio of 30 1 by volume 2 For airplanes with either a reserve oil or oil transfer system a fuel oil ratio of 40 1 by volume c Fuel oil ratios higher than those prescribed in paragraphs b
455. ol surfaces and if applicable propellers are free from frost snow or ice at the start of the takeoff 2 The ice accretion starts at liftoff 3 The critical ratio of thrust power to weight 4 Failure of the critical engine occurs at Ver and 5 Crew activation of the ice protection system is in accordance with a normal operating proce dure provided in the Airplane Flight Manual ex cept that after beginning the takeoff roll it must be assumed that the crew takes no action to activate the ice protection system until the airplane is at least 400 feet above the takeoff surface 156 Federal Aviation Regulations e The ice accretion before the ice protection System has been activated and is performing its intended function is the critical ice accretion formed on the unprotected and normally pro tected surfaces before activation and effective op eration of the ice protection system in continuous maximum atmospheric icing conditions This ice accretion only applies in showing compliance to 25 143 j and 25 207 h and 25 207 i Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 121 72 FR 44669 Aug 8 2007 Amdt 25 129 74 FR 38340 Aug 3 2009 ASA Appendix C to Part 25 y 3sdnol4d SNOMOIIN 3ALLO32J3
456. om crash impact and subsequent damage to the record from fire In meeting this requirement the record container must be located as far aft as practicable but need not be aft of the pressurized compartment and may not be where aft mounted engines may crush the container upon impact c A correlation must be established between the flight recorder readings of airspeed altitude and heading and the corresponding readings taking into account correction factors of the first pilot s instruments The correlation must cover the airspeed range over which the airplane is to be operated the range of altitude to which the air plane is limited and 360 degrees of heading Cor relation may be established on the ground as ap propriate 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 location under water and 3 Have an underwater locating device when required by the operating rules of this chapter on or adjacent to the container which is secured in such a manner that they are not likely to be sepa rated during crash impact e Any novel or unique design or operational characteristics of the aircraft shall be evaluated to determine if any dedicated parameters must be recorded on flight recorders in addition to or in place of existing requirements Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 8 31 FR 127 J
457. ompartments cargo covers and transparencies molded and thermoformed parts air ducting joints and trim strips decorative and chafing that are constructed of materials not covered in subparagraph iv below must be self extinguish ing when tested vertically in accordance with the applicable portions of part of this appendix or other approved equivalent means The average burn length may not exceed 8 inches and the av erage flame time after removal of the flame source may not exceed 15 seconds Drippings from the test specimen may not continue to flame for more than an average of 5 seconds after fall ing iii Motion picture film must be safety film meeting the Standard Specifications for Safety Photographic Film PHI 25 available from the American National Standards Institute 1430 Broadway New York NY 10018 If the film travels through ducts the ducts must meet the require ments of subparagraph ii of this paragraph iv Clear plastic windows and signs parts con structed in whole or in part of elastomeric materi als edge lighted instrument assemblies consist ing of two or more instruments in a common housing seat belts shoulder harnesses and cargo and baggage tiedown equipment including containers bins pallets etc used in passenger or crew compartments may not have an average burn rate greater than 2 5 inches per minute when 166 Federal Aviation Regulations tested horizontally in accordance with the a
458. on Cargo or baggage compartments Thermal Acoustic insulation materials Cargo compartment classification Cargo or baggage compartment smoke or fire detection systems 25 859 25 863 25 865 25 867 25 869 25 871 25 875 25 899 25 901 25 903 25 904 25 905 25 907 25 925 25 929 25 933 25 934 25 937 25 939 25 941 25 943 25 945 25 951 25 952 25 953 25 954 25 955 25 957 25 959 25 961 25 963 25 965 25 967 25 969 25 971 25 973 25 975 25 977 25 979 25 981 25 991 25 993 25 994 25 995 Part 25 Combustion heater fire protection Flammable fluid fire protection Fire protection of flight controls engine mounts and other flight structure Fire protection other components Fire protection systems MISCELLANEOUS Leveling means Reinforcement near propellers Electrical bonding and protection against static electricity Subpart E Powerplant GENERAL Installation Engines Automatic takeoff thrust control system ATTCS Propellers Propeller vibration and fatigue Propeller clearance Propeller deicing Reversing systems Turbojet engine thrust reverser system tests Turbopropeller drag limiting systems Turbine engine operating characteristics Inlet engine and exhaust compatibility Negative acceleration Thrust or power augmentation system FUEL SYSTEM General Fuel system analysis and test Fuel system independence Fuel system lightning protection
459. on of any engine and if the landing dis tance would be noticeably increased when a land ing is made with that engine inoperative the land ing distance must be determined with that engine inoperative unless the use of compensating means will result in a landing distance not more than that with each engine operating paragraph ASA 825 143 Docket No FAA 2005 22840 72 FR 44667 Aug 8 2007 CONTROLLABILITY AND MANEUVERABILITY 825 143 General a The airplane must be safely controllable and maneuverable during 1 Takeoff 2 Climb 3 Level flight 4 Descent and 5 Landing b It must be possible to make a smooth tran sition from one flight condition to any other flight condition without exceptional piloting skill alert ness or strength and without danger of exceed ing the airplane limit load factor under any proba ble operating conditions including 1 The sudden failure of the critical engine 2 For airplanes with three or more engines the sudden failure of the second critical engine when the airplane is in the en route approach or landing configuration and is trimmed with the crit ical engine inoperative and 3 Configuration changes including deploy ment or retraction of deceleration devices c The airplane must be shown to be safely controllable and maneuverable with the critical ice accretion appropriate to the phase of flight de fined in appendix C and with the crit
460. on rod or calibration calorimeter pass 7 Calorimeter A total flux type calorimeter must be mounted in the center of a inch Kao wool M board inserted in the sample holder to measure the total heat flux The calorimeter must have a view angle of 180 degrees and be cali brated for incident flux The calorimeter calibration must be acceptable to the Administrator 8 Pilot Flame Positions Pilot ignition of the specimen must be accomplished by simulta neously exposing the specimen to a lower pilot burner and an upper pilot burner as described in paragraph b 8 i and b 8 ii or D 8 iii of this part IV respectively Since intermittent pilot flame extinguishment for more than 3 seconds would in validate the test results a spark ignitor may be in stalled to ensure that the lower pilot burner re mains lighted i Lower Pilot Burner The pilot flame tubing must be 25 inch 6 3 mm O D 03 inch 0 8 mm wall stainless steel tubing A mixture of 120 cm min of methane and 850 cm min of air must be fed to the lower pilot flame burner The normal position of the end of the pilot burner tub ing is 40 inch 10 mm from and perpendicular to the exposed vertical surface of the specimen The centerline at the outlet of the burner tubing must intersect the vertical centerline of the sample at a point 20 inch 5 mm above the lower exposed edge of the specimen ii Standard Three Hole Upper Burner The pilot burner must
461. on to the oxygen systems com ponents EWIS must be designed and installed with adequate physical separation from oxygen lines and other oxygen system components so 144 Federal Aviation Regulations that an EWIS component failure will not create a hazardous condition h Except to the extent necessary to provide electrical connection to the water waste systems components EWIS must be designed and in stalled with adequate physical separation from water waste lines and other water waste system components so that 1 An EWIS component failure will not create a hazardous condition 2 Any water waste leakage onto EWIS com ponents will not create a hazardous condition i EWIS must be designed and installed with adequate physical separation between the EWIS and flight or other mechanical control systems ca bles and associated system components so that 1 Chafing jamming or other interference are prevented 2 An EWIS component failure will not create a hazardous condition 3 Failure of any flight or other mechanical control systems cables or systems components will not damage the EWIS and create a hazard ous condition j EWIS must be designed and installed with adequate physical separation between the EWIS components and heated equipment hot air ducts and lines so that 1 An EWIS component failure will not create a hazardous condition 2 Any hot air leakage or heat generated onto EWIS components will n
462. on under any operating condi tion Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6912 May 5 1967 Amdt 25 23 35 FR 5676 April 8 1970 25 863 Flammable fluid fire protection a In each area where flammable fluids or va pors might escape by leakage of a fluid system there must be means to minimize the probability of ignition of the fluids and vapors and the result ant hazards if ignition does occur b Compliance with paragraph a of this sec tion must be shown by analysis or tests and the following factors must be considered 1 Possible sources and paths of fluid leakage and means of detecting leakage 2 Flammability characteristics of fluids in cluding effects of any combustible or absorbing materials 3 Possible ignition sources including electri cal faults overheating of equipment and malfunc tioning of protective devices 4 Means available for controlling or extin guishing a fire such as stopping flow of fluids shutting down equipment fireproof containment or use of extinguishing agents 5 Ability of airplane components that are criti cal to safety of flight to withstand fire and heat c If action by the flight crew is required to pre vent or counteract a fluid fire e g equipment shutdown or actuation of a fire extinguisher quick acting means must be provided to alert the crew d Each area where flammable fluids or vapors might escape by leakag
463. onably expected to occur on the flight Dynamic effects on these static loads need not be considered Corrective action to be taken by the pilot following the incident such as limiting ma neuvers avoiding turbulence and reducing speed must be considered If significant changes in struc tural stiffness or geometry or both follow from a structural failure or partial failure the effect on damage tolerance must be further investigated Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 45 43 FR 46242 Oct 5 1978 Amdt 25 54 45 FR 60173 Sept 11 1980 Amdt 25 72 55 FR 29776 July 20 1990 Amdt 25 86 61 FR 5222 Feb 9 1996 Amdt 25 96 63 FR 15714 March 31 1998 Amdt 25 92 63 FR 23338 April 28 1998 ASA 825 581 LIGHTNING PROTECTION 825 581 Lightning protection a The airplane must be protected against cat astrophic effects from lightning b For metallic components compliance with paragraph a of this section may be shown by 1 Bonding the components properly to the air frame or 2 Designing the components so that a strike will not endanger the airplane c For nonmetallic components compliance with paragraph a of this section may be shown 1 Designing the components to minimize the effect of a strike or 2 Incorporating acceptable means of diverting the resulting electrical current so as not to endan ger the airplane Docket No 5066 29 FR 18291 Dec 24
464. one panel one or more panels remain available for use by a pilot seated at a pilot station to permit continued safe flight and landing Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 Amdt 25 38 41 FR 55466 Dec 20 1976 825 777 Cockpit controls a Each cockpit control must be located to pro vide convenient operation and to prevent confu sion and inadvertent operation b The direction of movement of cockpit con trols must meet the requirements of 825 779 Wherever practicable the sense of motion involved in the operation of other controls must correspond to the sense of the effect of the operation upon the airplane or upon the part operated Controls of a variable nature using a rotary motion must move clockwise from the off position through an increas ing range to the full on position c The controls must be located and arranged with respect to the pilots seats so that there is full and unrestricted movement of each control without interference from the cockpit structure or the clothing of the minimum flight crew estab lished under 825 1523 when any member of this flight crew from 5 2 to 6 3 in height is seated with the seat belt and shoulder harness if pro vided fastened d Identical powerplant controls for each en gine must be located to prevent confusion as to the engines they control e Wing flap controls and other auxiliary lift de vice
465. ons or equivalent units of usable fuel in each tank during flight In addition 1 Each fuel quantity indicator must be cali brated to read zero during level flight when the quantity of fuel remaining in the tank is equal to the unusable fuel supply determined under 825 959 2 Tanks with interconnected outlets and air spaces may be treated as one tank and need not have separate indicators and 3 Each exposed sight gauge used as a fuel quantity indicator must be protected against damage c Fuel flowmeter system If a fuel flowmeter System is installed each metering component must have a means for bypassing the fuel supply if malfunction of that component severely restricts fuel flow d Oil quantity indicator There must be a stick gauge or equivalent means to indicate the quan tity of oil in each tank If an oil transfer or reserve oil supply system is installed there must be a means to indicate to the flight crew in flight the quantity of oil in each tank e Turbopropeller blade position indicator Re quired turbopropeller blade position indicators must begin indicating before the blade moves more than eight degrees below the flight low pitch stop The source of indication must directly sense the blade position f Fuel pressure indicator There must be means to measure fuel pressure in each system supplying reciprocating engines at a point down stream of any fuel pump except fuel injection pumps In addi
466. ontinued Airworthiness ICA applicable 10 EWIS as defined by 25 1701 that are approved by the FAA and include the following 1 Maintenance and inspection requirements for the EWIS developed with the use of an en hanced zonal analysis procedure that includes i Identification of each zone of the airplane ii Identification of each zone that contains EWIS iii Identification of each zone containing EWIS that also contains combustible materials iv Identification of each zone in which EWIS is in close proximity to both primary and back up hy draulic mechanical or electrical flight controls and lines v Identification of A Tasks and the intervals for performing those tasks that will reduce the likelihood of igni tion sources and accumulation of combustible material and B Procedures and the intervals for perform ing those procedures that will effectively clean the EWIS components of combustible material if there is not an effective task to reduce the likeli hood of combustible material accumulation 216 Federal Aviation Regulations vi Instructions for protections and caution in formation that will minimize contamination and ac cidental damage to EWIS as applicable during performance of maintenance alteration or re pairs 2 Acceptable EWIS maintenance practices in a standard format 3 Wire separation requirements as deter mined under 825 1707 4 Information explaining the EWIS ide
467. or by synthesis from segments If the takeoff path is determined by the segmental method 1 The segments must be clearly defined and must be related to the distinct changes in the con figuration power or thrust and speed 2 The weight of the airplane the configura tion and the power or thrust must be constant throughout each segment and must correspond to the most critical condition prevailing in the seg ment 3 The flight path must be based on the air plane s performance without ground effect and 4 The takeoff path data must be checked by continuous demonstrated takeoffs up to the point at which the airplane is out of ground effect and its speed is stabilized to ensure that the path is conservative relative to the continuous path The airplane is considered to be out of the ground effect when it reaches a height equal to its wing span e For airplanes equipped with standby power rocket engines the takeoff path may be deter mined in accordance with section Il of Appendix E Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 6 30 FR 8468 July 2 1965 Amdt 25 42 43 FR 2321 Jan 16 1978 Amdt 25 54 45 FR 60172 Sept 11 1980 Amdt 25 72 55 FR 29774 July 20 1990 Amdt 25 94 63 FR 8848 Feb 23 1998 Amdt 25 108 67 FR 70826 Nov 26 2002 Amdt 25 121 72 FR 44666 Aug 8 2007 20 Federal Aviation Regulations 825 113 Takeoff distance and takeoff run a Takeoff
468. or valves 2 Tests of the pressurization system to show proper functioning under each possible condition of pressure temperature and moisture up to the maximum altitude for which certification is re quested 3 Flight tests to show the performance of the pressure supply pressure and flow regulators in dicators and warning signals in steady and stepped climbs and descents at rates correspond ing to the maximum attainable within the operating limitations of the airplane up to the maximum alti tude for which certification is requested 4 Tests of each door and emergency exit to show that they operate properly after being sub jected to the flight tests prescribed in paragraph b 3 of this section FiRE PROTECTION 825 851 Fire extinguishers a Hand fire extinguishers 1 The following minimum number of hand fire extinguishers must be conveniently located and evenly distributed in passenger compartments Passenger capacity 7 through 30 No of extinguishers 31 through 60 61 through 200 201 through 300 301 through 400 401 through 500 NIJA Oo rm 501 through 600 601 through 700 8 2 At least one hand fire extinguisher must be conveniently located in the pilot compartment 3 At least one readily accessible hand fire ex tinguisher must be available for use in each Class A or Class B cargo or baggage compartment and in each Class E cargo or baggage compartment th
469. oratory test procedure for measuring the ca pability of cargo compartment lining materials to resist flame penetration with a 2 gallon per hour GPH 2 Grade kerosene or equivalent burner fire source Ceiling and sidewall liner panels may be tested individually provided a baffle is used to simulate the missing panel Any specimen that passes the test as a ceiling liner panel may be used as a sidewall liner panel c Test Specimens 1 The specimen to be tested must measure 16 inches 406 3 mm by 24 inches 6103 mm 2 The specimens must be conditioned at 70 F 5 F 21 2 and 55 5 humidity for at least 24 hours before testing d Test Apparatus The arrangement of the test apparatus which is shown in Figure of Part II and Figures 1 through 3 of this part of Appendix F must include the components described in this section Minor details of the apparatus may vary depending on the model of the burner used 1 Specimen Mounting Stand The mounting stand for the test specimens consists of steel an gles as shown in Figure 1 2 Test Burner The burner to be used in test ing must i Be a modified gun type ii Use a suitable nozzle and maintain fuel pressure to yield a 2 GPH fuel flow For example an 80 degree nozzle nominally rated at 2 25 GPH and operated at 85 pounds per square inch PSI gage to deliver 2 03 GPH iii Have a 12 inch 305 mm burner extension installed at the end of the draft tub
470. ot create a hazardous condition For systems for which redundancy is re quired by certification rules by operating rules or as a result of the assessment required by 825 1709 EWIS components associated with those systems must be designed and installed with adequate physical separation I Each EWIS must be designed and installed So there is adequate physical separation between it and other aircraft components and aircraft struc ture and so that the EWIS is protected from sharp edges and corners to minimize potential for abrasion chafing vibration damage and other types of mechanical damage 825 1709 System safety EWIS Each EWIS must be designed and installed so that a Each catastrophic failure condition 1 Is extremely improbable and 2 Does not result from a single failure b Each hazardous failure condition is ex tremely remote ASA Part 25 Airworthiness Standards Transport Category 825 1711 Component identification EWIS a EWIS components must be labeled or otherwise identified using a consistent method that facilitates identification of the EWIS compo nent its function and its design limitations if any b For systems for which redundancy is re quired by certification rules by operating rules or as a result of the assessment required by 825 1709 EWIS components associated with those systems must be specifically identified with component part number function and separation
471. other crewmembers or 2 A common source of supply with means to separately reserve the minimum supply required by the flight crew on duty b Portable walk around oxygen units of the continuous flow diluter demand and straight de mand kinds may be used to meet the crew or pas senger breathing requirements 132 Federal Aviation Regulations 825 1447 Equipment standards for oxygen dispensing units If oxygen dispensing units are installed the fol lowing apply a There must be an individual dispensing unit for each occupant for whom supplemental oxygen is to be supplied Units must be designed to cover the nose and mouth and must be equipped with a suitable means to retain the unit in position on the face Flight crew masks for supplemental oxygen must have provisions for the use of communica tion equipment b If certification for operation up to and includ ing 25 000 feet is requested an oxygen supply terminal and unit of oxygen dispensing equipment for the immediate use of oxygen by each crew member must be within easy reach of that crew member For any other occupants the supply ter minals and dispensing equipment must be lo cated to allow the use of oxygen as required by the operating rules in this chapter c f certification for operation above 25 000 feet is requested there must be oxygen dispensing equipment meeting the following requirements 1 There must be an oxygen dispensing unit connected to ox
472. otion of flaps or slats on opposite sides of the plane of symmetry must be synchronized by a mechanical interconnection or approved equivalent means b If wing flap or slat interconnection or equivalent means is used it must be designed to account for the applicable unsymmetrical loads including those resulting from flight with the en gines on one side of the plane of symmetry inop erative and the remaining engines at takeoff power c For airplanes with flaps or slats that are not subjected to slipstream conditions the structure must be designed for the loads imposed when the wing flaps or slats on one side are carrying the most severe load occurring in the prescribed sym metrical conditions and those on the other side are carrying not more than 80 percent of that load d The interconnection must be designed for the loads resulting when interconnected flap or slat surfaces on one side of the plane of symme try are jammed and immovable while the surfaces on the other side are free to move and the full power of the surface actuating system is applied Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29777 July 20 1990 64 Federal Aviation Regulations 825 703 Takeoff warning system A takeoff warning system must be installed and must meet the following requirements a The system must provide to the pilots an au ral warning that is automatically activated during the initial portion
473. ound when in the level takeoff attitude with the critical tire s com pletely deflated and the corresponding landing gear strut bottomed b Water clearance There must be a clearance of at least 18 inches between each propeller and the water unless compliance with 25 239 a can be shown with a lesser clearance c Structural clearance There must be 1 At least one inch radial clearance between the blade tips and the airplane structure plus any additional radial clearance necessary to prevent harmful vibration 2 At least one half inch longitudinal clearance between the propeller blades or cuffs and station ary parts of the airplane and 3 Positive clearance between other rotating parts of the propeller or spinner and stationary parts of the airplane Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29784 July 20 1990 525 929 Propeller deicing a For airplanes intended for use where icing may be expected there must be a means to pre 95 825 933 vent or remove hazardous ice accumulation on propellers or on accessories where ice accumula tion would jeopardize engine performance b If combustible fluid is used for propeller de icing 25 1181 through 25 1185 and 25 1189 apply 825 933 Reversing systems a For turbojet reversing systems 1 Each system intended for ground operation only must be designed so that during any reversal in flight the engine will p
474. ounos J0 2eJ JUS UOD Jaye pinbr1 ssejuorsueuulq 4 ASA 162 Part 25 Airworthiness Standards Transport Category APPENDIX D TO PART 25 Criteria for determining minimum flight crew The following are considered by the Agency in determin ing the minimum flight crew under 825 1523 a Basic workload functions The following ba sic workload functions are considered 1 Flight path control 2 Collision avoidance 3 Navigation 4 Communications 5 Operation and monitoring of aircraft en gines and systems 6 Command decisions b Workload factors The following workload factors are considered significant when analyzing and demonstrating workload for minimum flight crew determination 1 The accessibility ease and simplicity of op eration of all necessary flight power and equip ment controls including emergency fuel shutoff valves electrical controls electronic controls pres surization system controls and engine controls 2 The accessibility and conspicuity of all nec essary instruments and failure warning devices such as fire warning electrical system malfunc tion and other failure or caution indicators The extent to which such instruments or devices direct the proper corrective action is also considered 3 The number urgency and complexity of op erating procedures with particular consideration given to the specific fuel management
475. ounted for d Supplementary design envelope analysis In addition to the limit loads defined by paragraph c of this appendix limit loads must also be de termined in accordance with paragraph b of this appendix except that 1 In paragraph b 3 i of this appendix the value of Uc 85 fps true gust velocity is replaced by Uc 60 fps true gust velocity on the interval 0 to 30 000 ft altitude and is linearly decreased to 25 fps true gust velocity at 80 000 ft altitude and 2 In paragraph b of this appendix the refer ence to paragraphs b 3 i through b 3 iii of this appendix is to be understood as referring to the paragraph as modified by paragraph d 1 ASA Part 25 Airworthiness Standards Transport Category Appendix G to Part 25 10 01 P 001 0001 P AND P VALUES 00001 000001 FIGURE 1 80 70 60 o Ww t e 000 20 10 0 213 Appendix G to Part 25 Federal Aviation Regulations b AND b VALUES FIGURE 2 80 70 o e 10 0001 20 10 0 214 ASA Part 25 Airworthiness Standards Transport Category APPENDIX H TO PART 25 INSTRUCTIONS FOR CONTINUED AIRWORTHINESS H25 1 GENERAL a This appendix specifies requirements for preparation of Instructions for Contin
476. ovable bolt screw nut pin or other removable fastener must meet the locking requirements of 825 607 6 Certain doors as specified by 25 807 h must also meet the applicable requirements of 25 809 through 25 812 for emergency exits b Opening by persons There must be a means to safeguard each door against opening during flight due to inadvertent action by persons In addition design precautions must be taken to minimize the possibility for a person to open a door intentionally during flight If these precau tions include the use of auxiliary devices those devices and their controlling systems must be de signed so that 1 No single failure will prevent more than one exit from being opened and 2 Failures that would prevent opening of the exit after landing are improbable c Pressurization prevention means There must be a provision to prevent pressurization of the airplane to an unsafe level if any door subject to pressurization is not fully closed latched and locked 1 The provision must be designed to function after any single failure or after any combination of failures not shown to be extremely improbable 2 Doors that meet the conditions described in paragraph h of this section are not required to have a dedicated pressurization prevention means if from every possible position of the door it will remain open to the extent that it prevents pressurization or safely close and latch as pres suriza
477. oving between the fully open and fully closed position Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15044 March 17 1977 Amdt 25 72 55 FR 29785 July 20 1990 825 1142 Auxiliary power unit controls Means must be provided on the flight deck for starting stopping and emergency shutdown of each installed auxiliary power unit Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50598 Oct 30 1978 25 1143 Engine controls a There must be a separate power or thrust control for each engine b Power and thrust controls must be arranged to allow 1 Separate control of each engine and 2 Simultaneous control of all engines c Each power and thrust control must provide a positive and immediately responsive means of controlling its engine d For each fluid injection other than fuel sys tem and its controls not provided and approved as part of the engine the applicant must show that the flow of the injection fluid is adequately con trolled e If a power or thrust control incorporates a fuel shutoff feature the control must have a means to prevent the inadvertent movement of the control into the shutoff position The means must 1 Have a positive lock or stop at the idle posi tion and 2 Require a separate and distinct operation to place the control in the shutoff position 109 825 1145 Docket No 5066 29 FR 18291 Dec 2
478. p and bottom clamps SEE FIGURE 7 AT THE END OF PART VII OF THIS APPENDIX Note For blanket materials that cannot be in stalled in accordance with figure 7 above the blankets must be installed in manner approved by the FAA v Conditioning Condition the specimens at 70 5 21 2 and 55 10 relative hu midity for a minimum of 24 hours prior to testing d Preparation of apparatus 1 Level and center the frame assembly to en sure alignment of the calorimeter and or thermo couple rake with the burner cone 2 Turn on the ventilation hood for the test chamber Do not turn on the burner blower Mea sure the airflow of the test chamber using a vane anemometer or equivalent measuring device The vertical air velocity just behind the top of the upper ASA Part 25 Airworthiness Standards Transport Category insulation blanket test specimen must be 100 50 ft min 0 51 0 25 m s The horizontal air velocity at this point must be less than 50 ft min 0 25 m s 3 If a calibrated flow meter is not available measure the fuel flow rate using a graduated cyl inder of appropriate size Turn on the burner mo tor fuel pump after insuring that the igniter sys tem is turned off Collect the fuel via a plastic or rubber tube into the graduated cylinder for a 2 minute period Determine the flow rate in gallons per hour The fuel flow rate must be 6 0 0 2 gal lons per hour 0 378 0 0126 L min e Calibra
479. parallel to the airplane s longitudinal axis between the nearest exit edges unless the following conditions are met a Each passenger seat must be located within 30 feet from the nearest exit on each side of the fuselage as measured parallel to the air plane s longitudinal axis between the nearest exit edge and the front of the seat bottom cushion b The number of passenger seats located be tween two adjacent pairs of emergency exits commonly referred to as a passenger zone or between a pair of exits and a bulkhead or a com partment door commonly referred to as dead end zone may not exceed the following 1 For zones between two pairs of exits 50 percent of the combined rated capacity of the two pairs of emergency exits 2 For zones between one pair of exits and a bulkhead 40 percent of the rated capacity of the pair of emergency exits c The total number of passenger seats in the airplane may not exceed 33 percent of the maxi mum seating capacity for the airplane model us ASA SFAR No 109 to Part 25 ing the exit ratings listed in 25 807 g for the original certified exits or the maximum allowable after modification when exits are deactivated whichever is less d A distance of more than 60 feet between adjacent passenger emergency exits on the same side of the same deck of the fuselage as mea sured parallel to the airplane s longitudinal axis between the nearest exit edges is allowed only o
480. pe or any other means demonstrated to be suitable for the purpose If the assisting means is a rope or an approved device equivalent to a rope it must be i Attached to the fuselage structure at or above the top of the emergency exit opening or for a device at a pilot s emergency exit window at another approved location if the stowed device or its attachment would reduce the pilot s view in flight ii Able with its attachment to withstand a 400 pound static load b Assist means from the cabin to the wing are required for each Type A or Type B exit located above the wing and having a stepdown unless the exit without an assist means can be shown to have a rate of passenger egress at least equal to that of the same type of nonoverwing exit If an assist means is required it must be automatically deployed and automatically erected concurrent with the opening of the exit In the case of assist means installed at Type C exits it must be self supporting within 10 seconds from the time the opening means of the exits is actuated For all other exit types it must be self supporting 6 sec onds after deployment is begun c An escape route must be established from each overwing emergency exit and except for flap surfaces suitable as slides covered with a slip resistant surface Except where a means for channeling the flow of evacuees is provided 1 The escape route from each Type A or Type B passenger emergency exit or
481. peed prescribed in paragraphs c and d of this section Except for showing compliance with the stall warning margin prescribed in paragraph h 3 ii of this section stall warning for flight in icing conditions must be provided by the same means as stall warning for flight in non icing conditions c When the speed is reduced at rates not ex ceeding one knot per second stall warning must begin in each normal configuration at a speed Vsw exceeding the speed at which the stall is identified in accordance with 825 201 d by not less than five knots or five percent CAS which ever is greater Once initiated stall warning must continue until the angle of attack is reduced to ap proximately that at which stall warning began d In addition to the requirement of paragraph c of this section when the speed is reduced at rates not exceeding one knot per second in straight flight with engines idling and at the cen ter of gravity position specified in 25 103 b 5 Vsw in each normal configuration must exceed Vsg by not less than three knots or three percent CAS whichever is greater e In icing conditions the stall warning margin in straight and turning flight must be sufficient to allow the pilot to prevent stalling as defined in 25 201 d when the pilot starts a recovery ma neuver not less than three seconds after the on Set of stall warning When demonstrating compli ance with this paragraph the pilot must perform
482. pendix F to this part or other approved equiva lent test requirements This requirement does not apply to thermal acoustic insulation installations that the FAA finds would not contribute to fire pen etration resistance Docket No FAA 2000 7909 68 FR 45059 July 31 2003 825 857 Cargo compartment classification a Class A A Class A cargo or baggage com partment is one in which 1 The presence of a fire would be easily dis covered by a crewmember while at his station and 2 Each part of the compartment is easily ac cessible in flight b Class B A Class B cargo or baggage com partment is one in which 1 There is sufficient access in flight to enable a crewmember to effectively reach any part of the compartment with the contents of a hand fire ex tinguisher 2 When the access provisions are being used no hazardous quantity of smoke flames or extinguishing agent will enter any compartment occupied by the crew or passengers 3 There is a separate approved smoke detec tor or fire detector system to give warning at the pilot or flight engineer station c Class C A Class C cargo or baggage com partment is one not meeting the requirements for either a Class A or B compartment but in which ASA Part 25 Airworthiness Standards Transport Category 1 There is a separate approved smoke detec tor or fire detector system to give warning at the pilot or flight engineer station 2 There i
483. per ative it is possible to maintain control of the air plane with that engine still inoperative and main tain straight flight with an angle of bank of not more than 5 degrees Vc may not exceed 1 13 with 1 Maximum available takeoff power or thrust on the engines 2 The most unfavorable center of gravity 3 The airplane trimmed for takeoff 4 The maximum sea level takeoff weight or any lesser weight necessary to show Vme 5 The airplane in the most critical takeoff con figuration existing along the flight path after the airplane becomes airborne except with the land ing gear retracted 6 The airplane airborne and the ground effect negligible and 7 If applicable the propeller of the inoperative engine i Windmilling ii In the most probable position for the specific design of the propeller control or iii Feathered if the airplane has an automatic feathering device acceptable for showing compli ance with the climb requirements of 525 121 d The rudder forces required to maintain con trol at Vmc may not exceed 150 pounds nor may it be necessary to reduce power or thrust of the op erative engines During recovery the airplane may not assume any dangerous attitude or re quire exceptional piloting skill alertness or strength to prevent a heading change of more than 20 degrees e the minimum control speed on the ground is the calibrated airspeed during the take of
484. phone 202 307 2942 or iii National Archives and Records Administra tion NARA For information on the availability of this material at NARA go to http www archives gov federal register code of federal regulations ibr locations html or call 202 741 6030 3 You may obtain copies of NIJ Standard 0101 04 from the National Criminal Justice Refer ence Service P O Box 6000 Rockville MD 20849 6000 telephone 800 851 3420 Docket No FAA 2006 26722 73 FR 63879 Oct 28 2008 EMERGENCY PROVISIONS 825 801 Ditching a If certification with ditching provisions is re quested the airplane must meet the requirements of this section and 25 807 e 25 1411 and 25 1415 a b Each practicable design measure compati ble with the general characteristics of the air plane must be taken to minimize the probability that in an emergency landing on water the behav ior of the airplane would cause immediate injury to the occupants or would make it impossible for them to escape c The probable behavior of the airplane in a water landing must be investigated by model tests or by comparison with airplanes of similar config uration for which the ditching characteristics are known Scoops flaps projections and any other factor likely to affect the hydrodynamic character istics of the airplane must be considered d It must be shown that under reasonably probable water conditions the flotation time and trim of t
485. pidly as possible while maintaining the airspeed at approximately 30 percent above the reference stall speed existing at each instant throughout the maneuver 2 Repeat paragraph b 1 except initially ex tend the flaps and then retract them as rapidly as possible 3 Repeat paragraph b 2 except at the go around power or thrust setting 4 With power off flaps retracted and the air plane trimmed at 1 3 rapidly set go around power or thrust while maintaining the same air speed 5 Repeat paragraph b 4 except with flaps extended 6 With power off flaps extended and the air plane trimmed at 1 3 Vsg4 obtain and maintain airspeeds between Vsy and either 1 6 or Vee whichever is lower c It must be possible without exceptional pi loting skill to prevent loss of altitude when com plete retraction of the high lift devices from any position is begun during steady straight level flight at 1 08 for propeller powered air planes or 1 13 Vsg4 for turbojet powered air planes with 1 Simultaneous movement of the power or thrust controls to the go around power or thrust setting 2 The landing gear extended and 3 The critical combinations of landing weights and altitudes d If gated high lift device control positions are provided paragraph c of this section applies to retractions of the high lift devices from any posi tion from the maximum landing position to the first gated posit
486. plus or minus 10 degrees or are supported by the vertical tail surfaces the sur faces and the supporting structure must be de signed for gust velocities specified in 25 341 a acting in any orientation at right angles to the flight path d Unsymmetrical loading on the empennage arising from buffet conditions of 25 305 e must be taken into account Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 Amdt 25 86 61 FR 5222 Feb 9 1996 25 445 Auxiliary aerodynamic surfaces a When significant the aerodynamic influ ence between auxiliary aerodynamic surfaces such as outboard fins and winglets and their sup porting aerodynamic surfaces must be taken into ASA 825 471 account for all loading conditions including pitch roll and yaw maneuvers and gusts as specified in 825 341 a acting at any orientation at right an gles to the flight path b To provide for unsymmetrical loading when outboard fins extend above and below the hori zontal surface the critical vertical surface loading load per unit area determined under 825 391 must also be applied as follows 1 100 percent to the area of the vertical sur faces above or below the horizontal surface 2 80 percent to the area below or above the horizontal surface Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 86 61 FR 5222 Feb 9 1996 825 457 Wing flaps Wing flaps
487. ply with paragraphs a through f of this section but must have a means to prevent inadvertent opening during flight 2 Inward opening removable emergency exits that are not normally removed except for mainte nance purposes or emergency evacuation and flight deck openable windows need not comply with paragraphs c and f of this section 3 Maintenance doors that meet the conditions of paragraph h of this section and for which a placard is provided limiting use to maintenance access need not comply with paragraphs c and f of this section h Doors that are not a hazard For the pur poses of this section a door is considered not to be a hazard in the unlatched condition during flight provided it can be shown to meet all of the following conditions 1 Doors in pressurized compartments would remain in the fully closed position if not restrained by the latches when subject to a pressure greater than 1 2 psi Opening by persons either inadvert ently or intentionally need not be considered in making this determination 2 The door would remain inside the airplane or remain attached to the airplane if it opens ei ther in pressurized or unpressurized portions of the flight This determination must include the ASA 825 785 consideration of inadvertent and intentional open ing by persons during either pressurized or un pressurized portions of the flight 3 The disengagement of the latches during flight would
488. port Airplane Director ate Aircraft Certification Service 1601 Lind Ave nue SW Renton Washington 98057 3356 For in formation on the availability of this material at NARA call 202 741 6030 or go to http www archives gov federal register code of federal regulations ibr locations html b The following materials are available for purchase from the following address The Na tional Technical Information Services NTIS Springfield Virginia 22166 1 Fuel Tank Flammability Assessment Method Users Manual dated May 2008 docu ment number DOT FAA AR 05 8 IBR approved for 825 981 and Appendix N It can also be ob tained at the following Web site http www fire tc faa gov systems fueltank FTFAM stm 2 Reserved Docket No FAA 2005 22997 73 FR 42494 July 21 2008 ASA Part 25 Airworthiness Standards Transport Category Subpart B Flight GENERAL 25 21 Proof of compliance a Each requirement of this subpart must be met at each appropriate combination of weight and center of gravity within the range of loading conditions for which certification is requested This must shown 1 By tests upon an airplane of the type for which certification is requested or by calculations based on and equal in accuracy to the results of testing and 2 By systematic investigation of each proba ble combination of weight and center of gravity if compliance cannot be reasonably inferred from combination
489. posure to humid and inclement weather on the ground followed by a long duration flight at normal cruise altitude 2 The airplane demonstration flight test pro gram must validate the adequacy of the airplane s flying qualities and performance and the flight crew s ability to safely conduct an ETOPS diver sion under the conditions specified in section K25 2 2 g 1 of this appendix 3 During the airplane demonstration flight test program each test airplane must be operated and maintained using the applicants recom mended operating and maintenance procedures 4 At the completion of the airplane demon stration flight test program each ETOPS signifi cant system must undergo an on wing inspection or test in accordance with the tasks defined in the proposed Instructions for Continued Airworthi ness to establish its condition for continued safe operation Each engine must also undergo a gas path inspection These inspections must be con ducted in a manner to identify abnormal condi tions that could result in an IFSD or diversion The applicant must identify track and resolve any ab normal conditions in accordance with the problem tracking and resolution system specified in sec tion K25 2 2 h of this appendix h Problem tracking and resolution system 1 The applicant must establish and maintain a problem tracking and resolution system The sys tem must i Contain a process for prompt reporting to the responsible FAA ai
490. power is used may not be less than the takeoff time limitation 2 For turbine engine powered airplanes the engines must operate at takeoff power for the time interval selected for showing the takeoff flight path and at maximum continuous power for the rest of the climb 3 The weight of the airplane must be the weight with full fuel tanks minimum crew and the ballast necessary to maintain the center of gravity within allowable limits 4 The climb airspeed may not exceed i For reciprocating engine powered airplanes the maximum airspeed established for climbing from takeoff to the maximum operating altitude with the airplane in the following configuration A Landing gear retracted B Wing flaps in the most favorable position C Cowl flaps or other means of controlling the engine cooling supply in the position that pro vides adequate cooling in the hot day condition D Engine operating within the maximum con tinuous power limitations ASA Part 25 Airworthiness Standards Transport Category E Maximum takeoff weight and ii For turbine engine powered airplanes the maximum airspeed established for climbing from takeoff to the maximum operating altitude 5 The fuel temperature must be at least 110 F b The test prescribed in paragraph a of this section may be performed in flight or on the ground under closely simulated flight conditions If a flight test is performed in weather cold enough to
491. pplica ble portions of this appendix v Except for small parts such as knobs han dles rollers fasteners clips grommets rub strips pulleys and small electrical parts that would not contribute significantly to the propaga tion of a fire and for electrical wire and cable insu lation materials in items not specified in para graphs a 1 i ii iii or iv of part of this ap pendix may not have a burn rate greater than 4 0 inches per minute when tested horizontally in ac cordance with the applicable portions of this ap pendix 2 Cargo and baggage compartments not oc cupied by crew or passengers i Reserved ii A cargo or baggage compartment defined in 825 857 as Class B or E must have a liner con structed of materials that meet the requirements of paragraph a 1 ii of part of this appendix and separated from the airplane structure except for attachments In addition such liners must be subjected to the 45 degree angle test The flame may not penetrate pass through the material during application of the flame or subsequent to its removal The average flame time after removal of the flame source may not exceed 15 seconds and the average glow time may not exceed 10 seconds iii A cargo or baggage compartment defined in 825 857 as Class B C D or E must have floor panels constructed of materials which meet the requirements of paragraph a 1 ii of part of this appendix and which are separat
492. pply to the emergency lighting system is independent of the power supply to the main lighting system The emergency lighting system must include 1 Illuminated emergency exit marking and lo cating signs sources of general cabin illumina tion interior lighting in emergency exit areas and floor proximity escape path marking 2 Exterior emergency lighting b Emergency exit signs 1 For airplanes that have a passenger seating configuration excluding pilot seats of 10 seats or more must meet the following requirements i Each passenger emergency exit locator sign required by 25 811 d 1 and each passen ger emergency exit marking sign required by 25 811 d 2 must have red letters at least 11 inches high on an illuminated white background and must have an area of at least 21 square inches excluding the letters The lighted back ground to letter contrast must be at least 10 1 The letter height to stroke width ratio may not be more than 7 1 nor less than 6 1 These signs must be internally electrically illuminated with a background brightness of at least 25 foot lam berts and a high to low background contrast no greater than 3 1 ii Each passenger emergency exit sign re quired by 25 811 d 3 must have red letters at least 11 5 inches high on a white background hav ing an area of at least 21 square inches excluding the letters These signs must be internally electri cally illuminated or self illuminated by oth
493. prehensive test data such as continuous joints in metal plating welded joints and scarf joints in wood or 2 With respect to any bearing surface for which a larger special factor is used c For each integral fitting the part must be treated as a fitting up to the point at which the section properties become typical of the member d For each seat berth safety belt and har ness the fitting factor specified in 25 785 f 3 applies Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5674 April 8 1970 Amdt 25 72 55 FR 29776 July 20 1990 825 629 Aeroelastic stability requirements a General The aeroelastic stability evalua tions required under this section include flutter di vergence control reversal and any undue loss of stability and control as a result of structural defor mation The aeroelastic evaluation must include whirl modes associated with any propeller or ro tating device that contributes significant dynamic forces Compliance with this section must be shown by analyses wind tunnel tests ground vi bration tests flight tests or other means found necessary by the Administrator b Aeroelastic stability envelopes The air plane must be designed to be free from aeroelas tic instability for all configurations and design con ditions within the aeroelastic stability envelopes as follows 1 For normal conditions without failures mal functions or adverse condit
494. ps up rearward for lift devices flaps down Trim tabs or Rotate to produce similar rotation of equivalent the airplane about an axis parallel to the axis of the control b Powerplant and auxiliary controls 1 Powerplant Controls Motion and effect Forward to increase forward thrust and rearward to increase rearward thrust Power or thrust Propellers Forward to increase rpm Mixture Forward or upward for rich Carburetor air heat Forward or upward for cold Supercharger Forward or upward for low blower For turbosuperchargers forward upward or clockwise to increase pressure 70 Federal Aviation Regulations 2 Auxiliary Controls Motion and effect Down to extend Landing gear Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29778 July 20 1990 ASA Part 25 Airworthiness Standards Transport Category 825 781 Cockpit control knob shape Cockpit control knobs must conform to the gen eral shapes but not necessarily the exact sizes or specific proportions in the following figure 825 783 MIXTURE CONTROL KNOB C POWER OR THRUST KNOB R SUPERCHARGER CONTROL KNOB ES PROPELLER CONTROL KNOB Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29779 July 20 1990 825 783 Fuselage doors a General This section applies to fuselag
495. r 825 489 Ground handling conditions Unless otherwise prescribed the landing gear and airplane structure must be investigated for the conditions in 825 491 through 25 509 with the airplane at the design ramp weight the maxi mum weight for ground handling conditions No wing lift may be considered The shock absorbers and tires may be assumed to be in their static po sition Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 825 491 Taxi takeoff and landing roll Within the range of appropriate ground speeds and approved weights the airplane structure and landing gear are assumed to be subjected to loads not less than those obtained when the air craft is operating over the roughest ground that may reasonably be expected in normal operation Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 91 62 FR 40705 July 29 1997 825 493 Braked roll conditions a An airplane with a tail wheel is assumed to be in the level attitude with the load on the main wheels in accordance with figure 6 of Appendix 47 825 495 A The limit vertical load factor is 1 2 at the design landing weight and 1 0 at the design ramp weight A drag reaction equal to the vertical reaction mul tiplied by a coefficient of friction of 0 8 must be combined with the vertical ground reaction and applied at the ground contact point b For an airplane with a nose wheel the limit
496. r 25 811 d 2 remains operative ex clusive of those that are directly damaged by the separation and 3 At least one required exterior emergency light for each side of the airplane remains opera tive exclusive of those that are directly damaged by the separation Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 15 32 FR 13265 Sept 20 1967 Amdt 25 28 36 FR 16899 Aug 26 1971 Amdt 25 32 37 FR 3971 Feb 24 1972 Amdt 25 46 43 FR 50597 Oct 30 1978 Amdt 25 58 49 FR 43186 Oct 26 1984 Amdt 25 88 61 FR 57958 Nov 8 1996 Amdt 25 116 69 FR 62788 Oct 27 2004 Amdt 25 128 74 FR 25645 May 29 2009 83 825 813 525 813 Emergency exit access Each required emergency exit must be acces sible to the passengers and located where it will afford an effective means of evacuation Emer gency exit distribution must be as uniform as practical taking passenger distribution into ac count however the size and location of exits on both sides of the cabin need not be symmetrical If only one floor level exit per side is prescribed and the airplane does not have a tail cone or ven tral emergency exit the floor level exit must be in the rearward part of the passenger compartment unless another location affords a more effective means of passenger evacuation Where more than one floor level exit per side is prescribed at least one floor level exit per side must be located near each end
497. r pressure load to which it would be subjected in operation EXHAUST SYSTEM 825 1121 General For powerplant and auxiliary power unit instal lations the following apply 108 Federal Aviation Regulations a Each exhaust system must ensure safe dis posal of exhaust gases without fire hazard or car bon monoxide contamination in any personnel compartment For test purposes any acceptable carbon monoxide detection method may be used to show the absence of carbon monoxide b Each exhaust system part with a surface hot enough to ignite flammable fluids or vapors must be located or shielded so that leakage from any system carrying flammable fluids or vapors will not result in a fire caused by impingement of the fluids or vapors on any part of the exhaust System including shields for the exhaust system c Each component that hot exhaust gases could strike or that could be subjected to high temperatures from exhaust system parts must be fireproof All exhaust system components must be Separated by fireproof shields from adjacent parts of the airplane that are outside the engine and auxiliary power unit compartments d No exhaust gases may discharge so as to cause a fire hazard with respect to any flammable fluid vent or drain e No exhaust gases may discharge where they will cause a glare seriously affecting pilot vi sion at night f Each exhaust system component must be ventilated to prevent points of excessively hig
498. r 1 000 world fleet en gine hours for approval up to and including 180 minute ETOPS or B Is at or below 0 01 per 1 000 world fleet en gine hours for approval greater than 180 minute ETOPS d Airplane systems assessment The appli cant must conduct an airplane systems assess ment The applicant must show that the airplane systems comply with 25 1309 b using available in service reliability data for ETOPS significant Systems on the candidate airplane engine combi nation Each cause or potential cause of a rele vant design manufacturing operational and maintenance problem occurring in service must have a corrective action or actions that are shown to be effective in preventing future occurrences Each corrective action must be identified in the CMP document specified in section K25 1 6 of this appendix A corrective action is not required if the problem would not significantly impact the safety or reliability of the airplane system in volved A relevant problem is a problem with an ETOPS group 1 significant system that has or could result in an IFSD or diversion The appli cant must include in this assessment relevant problems with similar or identical equipment in stalled on other types of airplanes to the extent such information is reasonably available e Airplane flight test The applicant must con duct a flight test to validate the flightcrew s ability to safely conduct an ETOPS diversion with an in operative engine
499. r all engines operating initial climb 3 That thrust or power setting which in the event of failure of the critical engine and without any crew action to adjust the thrust or power of the remaining engines would result in the thrust or power specified for the takeoff condition at V2 or any lesser thrust or power setting that is used for all engines operating initial climb procedures i When demonstrating compliance with 25 143 in icing conditions 1 Controllability must be demonstrated with the ice accretion defined in appendix C that is most critical for the particular flight phase 2 It must be shown that a push force is re quired throughout a pushover maneuver down to a zero g load factor or the lowest load factor ob tainable if limited by elevator power or other de sign characteristic of the flight control system It must be possible to promptly recover from the maneuver without exceeding a pull control force of 50 pounds and 3 Any changes in force that the pilot must ap ply to the pitch control to maintain speed with in creasing sideslip angle must be steadily increas ing with no force reversals unless the change in control force is gradual and easily controllable by the pilot without using exceptional piloting skill alertness or strength j For flight in icing conditions before the ice protection system has been activated and is per forming its intended function it must be demon strated in flight with the
500. r ceiling and wall panels other than lighting lenses and windows 2 Partitions other than transparent panels needed to enhance cabin safety 3 Galley structure including exposed sur faces of stowed carts and standard containers and the cavity walls that are exposed when a full complement of such carts or containers is not car ried and 4 Large cabinets and cabin stowage compart ments other than underseat stowage compart ments for stowing small items such as magazines and maps e The interiors of compartments such as pilot compartments galleys lavatories crew rest quar ters cabinets and stowage compartments need not meet the standards of paragraph d of this section provided the interiors of such compart ments are isolated from the main passenger cabin by doors or equivalent means that would normally be closed during an emergency landing condition f Smoking is not allowed in lavatories If smoking is allowed in any area occupied by the crew or passengers an adequate number of self contained removable ashtrays must be provided in designated smoking sections for all seated oc cupants g Regardless of whether smoking is allowed in any other part of the airplane lavatories must have self contained removable ashtrays located conspicuously on or near the entry side of each lavatory door except that one ashtray may serve more than one lavatory door if the ashtray can be seen readily from the cabin side of e
501. r each Type III exit in an airplane that has a passenger seating configura tion of 60 or more i Except as provided in paragraph c 1 ii the access must be provided by an unobstructed passageway that is at least 10 inches in width for interior arrangements in which the adjacent seat rows on the exit side of the aisle contain no more than two seats or 20 inches in width for interior arrangements in which those rows contain three seats The width of the passageway must be mea sured with adjacent seats adjusted to their most adverse position The centerline of the required passageway width must not be displaced more than 5 inches horizontally from that of the exit ii In lieu of one 10 or 20 inch passageway there may be two passageways between seat rows only that must be at least 6 inches in width and lead to an unobstructed space adjacent to each exit Adjacent exits must not share a com mon passageway The width of the passageways must be measured with adjacent seats adjusted to their most adverse position The unobstructed space adjacent to the exit must extend vertically from the floor to the ceiling or bottom of sidewall stowage bins inboard from the exit for a distance not less than the width of the narrowest passen ger seat installed on the airplane and from the forward edge of the forward passageway to the aft edge of the aft passageway The exit opening must be totally within the fore and aft bounds of the unobst
502. r emergency exit between the main aisles and the exit openings must be provided with illumination that is not less than 0 02 foot candle measured along a line that is within 6 inches of and parallel to the floor and is centered on the passenger evacuation path e Floor proximity emergency escape path marking must provide emergency evacuation guidance for passengers when all sources of illu mination more than 4 feet above the cabin aisle floor are totally obscured In the dark of the night the floor proximity emergency escape path mark ing must enable each passenger to 1 After leaving the passenger seat visually identify the emergency escape path along the cabin aisle floor to the first exits or pair of exits for ward and aft of the seat and 2 Readily identify each exit from the emer gency escape path by reference only to markings and visual features not more than 4 feet above the cabin floor f Except for subsystems provided in accor dance with paragraph h of this section that serve no more than one assist means are independent of the airplane s main emergency lighting system and are automatically activated when the assist means is erected the emergency lighting system must be designed as follows 1 The lights must be operable manually from the flight crew station and from a point in the pas senger compartment that is readily accessible to a normal flight attendant seat ASA Part 25 Airworthiness Standard
503. r taxi takeoff en route and landing environmental conditions such as altitude and temperature and loading conditions such as zero fuel weight center of gravity position and weight distribution must be established so that they are not more than 1 The highest weight selected by the appli cant for the particular conditions or 2 The highest weight at which compliance with each applicable structural loading and flight requirement is shown except that for airplanes equipped with standby power rocket engines the maximum weight must not be more than the high est weight established in accordance with Appen dix E of this part or 3 The highest weight at which compliance is shown with the certification requirements of Part 36 of this chapter b Minimum weight The minimum weight the lowest weight at which compliance with each ap plicable requirement of this part is shown must be established so that it is not less than 1 The lowest weight selected by the applicant 2 The design minimum weight the lowest weight at which compliance with each structural loading condition of this part is shown or 3 The lowest weight at which compliance with each applicable flight requirement is shown Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5671 April 8 1970 Amdt 25 63 53 FR 16365 May 6 1988 13 825 27 25 27 Center of gravity limits The extreme forward and the extre
504. r to indicate the functioning of the powerplant ice protection system for each engine 6 An indicator for the fuel strainer or filter re quired by 825 997 to indicate the occurrence of contamination of the strainer or filter before it reaches the capacity established in accordance with 825 997 d 7 A warning means for the oil strainer or filter required by 825 1019 if it has no bypass to warn the pilot of the occurrence of contamination of the strainer or filter screen before it reaches the capac ity established in accordance with 25 1019 a 2 8 An indicator to indicate the proper function ing of any heater used to prevent ice clogging of fuel system components d For turbojet engine powered airplanes In addition to the powerplant instruments required by paragraphs a and c of this section the fol lowing powerplant instruments are required 1 An indicator to indicate thrust or a parame ter that is directly related to thrust to the pilot The indication must be based on the direct measure ment of thrust or of parameters that are directly related to thrust The indicator must indicate a change in thrust resulting from any engine mal function damage or deterioration 2 A position indicating means to indicate to the flight crew when the thrust reversing device is in the reverse thrust position for each engine us ing a thrust reversing device 3 An indicator to indicate rotor system unbal ance e Fo
505. r turbopropeller powered airplanes In addition to the powerplant instruments required by paragraphs a and c of this section the fol lowing powerplant instruments are required 1 A torque indicator for each engine 2 Position indicating means to indicate to the flight crew when the propeller blade angle is be low the flight low pitch position for each propeller 116 Federal Aviation Regulations f For airplanes equipped with fluid systems other than fuel for thrust or power augmenta tion an approved means must be provided to in dicate the proper functioning of that system to the flight crew Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5678 April 8 1970 Amdt 25 35 39 FR 1831 Jan 15 1974 Amdt 25 36 39 FR 35461 Oct 1 1974 Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 54 45 FR 60173 Sept 11 1980 Amdt 25 72 55 FR 29785 July 20 1990 825 1307 Miscellaneous equipment The following is required miscellaneous equip ment a Reserved b Two or more independent sources of electri cal energy c Electrical protective devices as prescribed in this part d Two systems for two way radio communica tions with controls for each accessible from each pilot station designed and installed so that failure of one system will not preclude operation of the other system The use of a common antenna sys tem is acceptable if adequate reliability is sho
506. ragraph a 5 of this section Fg the flight profile alleviation factor defined in paragraph a 6 of this section 5 The following reference gust velocities apply i At the airplane design speed Vc Positive and negative gusts with reference gust velocities of 56 0 ft sec EAS must be considered at sea level The reference gust velocity may be reduced ASA 825 343 linearly from 56 0 ft sec EAS at sea level to 44 0 ft sec EAS at 15000 feet The reference gust ve locity may be further reduced linearly from 44 0 ft sec EAS at 15000 feet to 26 0 ft sec EAS at 50000 feet ii At the airplane design speed Vp The refer ence gust velocity must be 0 5 times the value ob tained under 25 341 a 5 i 6 The flight profile alleviation factor Fg must be increased linearly from the sea level value to a value of 1 0 at the maximum operating altitude de fined in 25 1527 At sea level the flight profile al leviation factor is determined by the following equation F 05 F 8 gz gm Where mo E 1 87 250000 Foy 4 Maximum Landing Weight Maximum Take off Weight Maximum Zero Fuel Weight Maximum Take off Weight R 2 Maximum operating altitude defined in 825 1527 7 When a stability augmentation system is in cluded in the analysis the effect of any significant System nonlinearities should be accounted for when deriving limit loads from limit gust condi tion
507. rative condition on airplanes with three or more engines Docket No FAA 2004 18379 72 FR 63405 Nov 8 2007 825 1316 System lightning protection a For functions whose failure would contribute to or cause a condition that would prevent the continued safe flight and landing of the airplane each electrical and electronic system that per forms these functions must be designed and in stalled to ensure that the operation and opera tional capabilities of the systems to perform these functions are not adversely affected when the air plane is exposed to lightning b For functions whose failure would contrib ute to or cause a condition that would reduce the capability of the airplane or the ability of the flight crew to cope with adverse operating conditions each electrical and electronic system that per forms these functions must be designed and in stalled to ensure that these functions can be re covered in a timely manner after the airplane is exposed to lightning c Compliance with the lightning protection criteria prescribed in paragraphs a and b of this section must be shown for exposure to a se vere lightning environment The applicant must design for and verify that aircraft electrical elec tronic systems are protected against the effects of lightning by 1 Determining the lightning strike zones for the airplane 2 Establishing the external lightning environ ment for the zones 3 Establishing the in
508. rcraft certification office of each occurrence reportable under 21 4 a 6 en countered during the phases of airplane and en gine development used to assess Early ETOPS eligibility ii Contain a process for notifying the responsi ble FAA aircraft certification office of each pro posed corrective action that the applicant deter mines necessary for each problem identified from the occurrences reported under section K25 2 2 h 1 i of this appendix The timing of the notifi cation must permit appropriate FAA review before taking the proposed corrective action 2 If the applicant is seeking ETOPS type de sign approval of a change to an airplane engine combination previously approved for ETOPS the problem tracking and resolution system need only address those problems specified in the following table provided the applicant obtains prior authori zation from the FAA 225 Appendix to Part 25 If the change does not require a new airplane type certificate and Then the Problem Tracking and Resolution System must address i Requires a new engine type certificate All problems applicable to the new engine installation and for the remainder of the airplane problems in changed systems only ii Does not require a new engine type certificate Problems in changed systems only i Acceptance criteria The type and frequency of failures and malfunctions on ETOPS significant Systems th
509. rd at 30 to drag axis 3 Aft parallel to drag axis 4 Aft at 30 to drag axis Auxiliary gear Swiveled forward 1 0 Frow 5 Forward 6 Aft Swiveled aft do 7 Forward 8 Aft Swiveled 45 from forward 0 5 Frow 9 Forward in plane of wheel 10 Aft in plane of wheel Swiveled 45 from aft do 11 Forward in plane of wheel 12 Aft in plane of wheel Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 825 511 Ground load unsymmetrical loads on multiple wheel units a General Multiple wheel landing gear units are assumed to be subjected to the limit ground loads prescribed in this subpart under paragraphs b through f of this section In addition 1 A tandem strut gear arrangement is a multi ple wheel unit and 2 In determining the total load on a gear unit with respect to the provisions of paragraphs b through f of this section the transverse shift in the load centroid due to unsymmetrical load dis tribution on the wheels may be neglected b Distribution of limit loads to wheels tires in flated The distribution of the limit loads among the wheels of the landing gear must be estab lished for each landing taxiing and ground han dling condition taking into account the effects of the following factors 1 The number of wheels and their physical ar rangements For truck type landing gear units the effects of any seesaw motion of the truck duri
510. re must be no indication of airspeed that would cause undue difficulty to the pilot during the takeoff between the initiation of rotation and the achievement of a steady climbing condition g The effects of airspeed indicating system lag may not introduce significant takeoff indicated airspeed bias or significant errors in takeoff or ac celerate stop distances h Each system must be arranged so far as practicable to prevent malfunction or serious er ror due to the entry of moisture dirt or other sub stances i Each system must have a heated pitot tube or an equivalent means of preventing malfunction due to icing ASA 825 1325 j Where duplicate airspeed indicators are re quired their respective pitot tubes must be far enough apart to avoid damage to both tubes in a collision with a bird Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 57 49 FR 6849 Feb 23 1984 Amdt 25 108 67 FR 70828 Nov 26 2002 Amdt 25 109 67 FR 76656 Dec 12 2002 825 1325 Static pressure systems a Each instrument with static air case con nections must be vented to the outside atmo sphere through an appropriate piping system b Each static port must be designed and lo cated in such manner that the static pressure sys tem performance is least affected by airflow varia tion or by moisture or other foreign matter and that the correlation between air pressure in the static pressure system and
511. re of the air plane and An analysis supported by test evidence of the principal structural elements and detail de sign points identified in paragraph a 1 ii of this section 2 The service history of airplanes of similar structural design taking due account of differ ences in operating conditions and procedures may be used in the evaluations required by this section 3 Based on the evaluations required by this section inspections or other procedures must be established as necessary to prevent catastrophic failure and must be included in the Airworthiness Limitations Section of the Instructions for Contin ued Airworthiness required by 825 1529 Inspec tion thresholds for the following types of structure must be established based on crack growth anal yses and or tests assuming the structure con tains an initial flaw of the maximum probable size that could exist as a result of manufacturing or service induced damage i Single load path structure and ii Multiple load path fail safe structure and crack arrest fail safe structure where it cannot be demonstrated that load path failure partial fail ure or crack arrest will be detected and repaired during normal maintenance inspection or opera 56 Federal Aviation Regulations tion of an airplane prior to failure of the remaining structure b Damage tolerance evaluation The evalua tion must include a determination of the probable locations
512. ree with respect to particle size and density that is greater than that estab lished for the engine under Part 33 of this chapter 3 The oil strainer or filter unless it is installed at an oil tank outlet must incorporate an indicator that will indicate contamination before it reaches the capacity established in accordance with para graph a 2 of this section 4 The bypass of a strainer or filter must be constructed and installed so that the release of collected contaminants is minimized by appropri ate location of the bypass to ensure that collected contaminants are not in the bypass flow path 5 An oil strainer or filter that has no bypass except one that is installed at an oil tank outlet must have a means to connect it to the warning System required in 825 1305 c 7 b Each oil strainer or filter in a powerplant in stallation using reciprocating engines must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer or filter element completely blocked Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 36 39 FR 35461 Oct 1 1974 Amdt 25 57 49 FR 6848 Feb 23 1984 825 1021 Oil system drains A drain or drains must be provided to allow safe drainage of the oil system Each drain must a Be accessible and b Have manual or automatic means for posi tive locking in the closed position Docket No 5066 29 FR 18
513. ring which the en gine or auxiliary power unit is permitted to be in operation b Each fuel system must be arranged so that any air which is introduced into the system will not result in 1 Power interruption for more than 20 sec onds for reciprocating engines or 2 Flameout for turbine engines c Each fuel system for a turbine engine must be capable of sustained operation throughout its flow and pressure range with fuel initially satu rated with water at 80 F and having 0 75cc of free water per gallon added and cooled to the most critical condition for icing likely to be encountered in operation d Each fuel system for a turbine engine pow ered airplane must meet the applicable fuel vent ing requirements of part 34 of this chapter Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5677 April 8 1970 Amdt 25 36 39 FR 35460 Oct 1 1974 Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 73 55 FR 32861 Aug 10 1990 25 952 Fuel system analysis and test a Proper fuel system functioning under all probable operating conditions must be shown by analysis and those tests found necessary by the Administrator Tests if required must be made using the airplane fuel system or a test article that reproduces the operating characteristics of the portion of the fuel system to be tested b The likely failure of any heat exchanger us ing fuel as one of its fluids may not result in
514. rld fleet for the airplane engine combination must accumulate a minimum of 250 000 engine hours The FAA may reduce this number of hours if the applicant iden tifies compensating factors that are acceptable to the FAA The compensating factors may include experience on another airplane but experience on the candidate airplane must make up a signifi 226 Federal Aviation Regulations cant portion of the total required service experi ence b Airplane systems assessment The appli cant must conduct an airplane systems assess ment The applicant must show that the airplane systems comply with the 25 1309 b using avail able in service reliability data for ETOPS signifi cant systems on the candidate airplane engine combination Each cause or potential cause of a relevant design manufacturing operational or maintenance problem occurring in service must have a corrective action or actions that are shown to be effective in preventing future occurrences Each corrective action must be identified in the CMP document specified in section K25 1 6 of this appendix A corrective action is not required if the problem would not significantly impact the safety or reliability of the airplane system in volved A relevant problem is a problem with an ETOPS group 1 significant system that has or could result in an IFSD or diversion The applicant must include in this assessment relevant prob lems with similar or identical equipment installed o
515. roduce no more than flight idle thrust In addition it must be shown by analysis or test or both that i Each operable reverser can be restored to the forward thrust position and 1 The airplane is capable of continued safe flight and landing under any possible position of the thrust reverser 2 Each system intended for inflight use must be designed so that no unsafe condition will result during normal operation of the system or from any failure or reasonably likely combination of failures of the reversing system under any antic ipated condition of operation of the airplane in cluding ground operation Failure of structural ele ments need not be considered if the probability of this kind of failure is extremely remote 3 Each system must have means to prevent the engine from producing more than idle thrust when the reversing system malfunctions except that it may produce any greater forward thrust that is shown to allow directional control to be maintained with aerodynamic means alone under the most critical reversing condition expected in operation b For propeller reversing systems 1 Each system intended for ground operation only must be designed so that no single failure or reasonably likely combination of failures or mal function of the system will result in unwanted re verse thrust under any expected operating condi tion Failure of structural elements need not be considered if this kind of failure
516. rogram must be used to show compliance with the flammability re quirements of 825 981 and Appendix M of this part The program must determine the time peri ods during each flight phase when the fuel tank or compartment with the FRM would be flammable The following factors must be considered in es tablishing these time periods 1 Any time periods throughout the flammabil ity exposure evaluation time and under the full range of expected operating conditions when the FRM is operating properly but fails to maintain a non flammable fuel tank because of the effects of the fuel tank vent system or other causes 2 If dispatch with the system inoperative un der the Master Minimum Equipment List MMEL is requested the time period assumed in the reli ability analysis 60 flight hours must be used for a 10 day MMEL dispatch limit unless an alternative period has been approved by the Administrator 3 Frequency and duration of time periods of FRM inoperability substantiated by test or analy sis acceptable to the FAA caused by latent or known failures including airplane system shut downs and failures that could cause the FRM to shut down or become inoperative 4 Effects of failures of the FRM that could in crease the flammability exposure of the fuel tank 5 If an FRM is used that is affected by oxygen concentrations in the fuel tank the time periods when oxygen evolution from the fuel results in the fuel tank or compartment ex
517. rom 25 341 a or the wing maximum airload derived indirectly from the verti cal load factor calculated from 825 341 It must be assumed that 100 percent of the wing air load gust criteria in ASA Part 25 Airworthiness Standards Transport Category acts on one side of the airplane and 80 percent of the wing air load acts on the other side Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5672 April 8 1970 Amdt 25 86 61 FR 5222 Feb 9 1996 Amdt 25 94 63 FR 8848 Feb 23 1998 825 351 Yaw maneuver conditions The airplane must be designed for loads result ing from the yaw maneuver conditions specified in paragraphs a through d of this section at speeds from Vyc to Vp Unbalanced aerodynamic moments about the center of gravity must be re acted in a rational or conservative manner consid ering the airplane inertia forces In computing the tail loads the yawing velocity may be assumed to be zero a With the airplane in unaccelerated flight at zero yaw it is assumed that the cockpit rudder control is suddenly displaced to achieve the re sulting rudder deflection as limited by 1 The control system on control surface Stops or 2 A limit pilot force of 300 pounds from to Va and 200 pounds from Vc Mc to Vp Mp with a linear variation between Va and Vc Mc b With the cockpit rudder control deflected so as always to maintain the maximum rudder de flection availa
518. rom fire mechanical or structural failure or persons standing on top of or against the escape routes In the event the air plane s main power system or compartment main lighting system should fail emergency illumina tion for each lower deck service compartment must be automatically provided b There must be a means for two way voice communication between the flight deck and each lower deck service compartment which remains available following loss of normal electrical power generating system 85 825 820 c There must be an aural emergency alarm System audible during normal and emergency conditions to enable crewmembers on the flight deck and at each required floor level emergency exit to alert occupants of each lower deck service compartment of an emergency situation d There must be a means readily detectable by occupants of each lower deck service com partment that indicates when seat belts should be fastened e If a public address system is installed in the airplane speakers must be provided in each lower deck service compartment f For each occupant permitted in a lower deck service compartment there must be a forward or aft facing seat which meets the requirements of 25 785 d and must be able to withstand maxi mum flight loads when occupied g For each powered lift system installed be tween a lower deck service compartment and the main deck for the carriage of persons or equip ment or both the
519. rovided in the lines and connec tions and d Have the capacity with respect to operating limitations established for the engine to ensure that engine fuel system functioning is not im paired with the fuel contaminated to a degree with respect to particle size and density that is greater than that established for the engine in Part 33 of this chapter Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 36 39 FR 35460 Oct 1 1974 Amdt 25 57 49 FR 6848 Feb 23 1984 825 999 Fuel system drains a Drainage of the fuel system must be accom plished by the use of fuel strainer and fuel tank sump drains b Each drain required by paragraph a of this section must 1 Discharge clear of all parts of the airplane 2 Have manual or automatic means for posi tive locking in the closed position and 3 Have a drain valve i That is readily accessible and which can be easily opened and closed and ii That is either located or protected to prevent fuel spillage in the event of a landing with landing gear retracted Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55467 Dec 20 1976 825 1001 Fuel jettisoning system a A fuel jettisoning system must be installed on each airplane unless it is shown that the air plane meets the climb requirements of 25 119 and 25 121 d at maximum takeoff weight less the actual or computed weight of fuel necessary
520. rovided that the ac ceptance rate of the stand or ramp is no greater than the acceptance rate of the means available on the airplane for descent from the wing during an actual crash situation evacuees using stands or ramps allowed by paragraph c of this Appen dix are considered to be on the ground when they are on the stand or ramp Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29788 July 20 1990 Amdt 25 79 Aug 26 1993 Amdt 25 117 69 FR 67499 Nov 17 2004 221 Appendix K to Part 25 APPENDIX K TO PART 25 EXTENDED OPERATIONS ETOPS Source Docket No FAA 2002 6717 72 FR 1873 Jan 16 2007 unless otherwise noted This appendix specifies airworthiness require ments for the approval of an airplane engine com bination for extended operations ETOPS For two engine airplanes the applicant must comply with sections K25 1 and K25 2 of this appendix For airplanes with more than two engines the ap plicant must comply with sections K25 1 and K25 3 of this appendix K25 1 DESIGN REQUIREMENTS K25 1 1 Part 25 compliance The airplane engine combination must comply with the requirements of part 25 considering the maximum flight time and the longest diversion time for which the applicant seeks approval K25 1 2 Human factors An applicant must consider crew workload op erational implications and the crew s and passen gers physiological needs during continued opera tion w
521. rrelation of the flight test data with other test data or analyses that the airplane is free from any aeroelastic instability at all speeds within the altitude airspeed envelope described in para graph b 2 of this section Docket No 26007 57 FR 28949 June 29 1992 25 631 Bird strike damage The empennage structure must be designed to assure capability of continued safe flight and land ing of the airplane after impact with an 8 pound bird when the velocity of the airplane relative to the bird along the airplane s flight path is equal to Vc at sea level selected under 25 335 a Com pliance with this section by provision of redundant structure and protected location of control system elements or protective devices such as splitter plates or energy absorbing material is acceptable Where compliance is shown by analysis tests or both use of data on airplanes having similar structural design is acceptable Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5674 April 8 1970 ASA 825 671 CONTROL SURFACES 525 651 Proof of strength a Limit load tests of control surfaces are re quired These tests must include the horn or fitting to which the control system is attached b Compliance with the special factors require ments of 25 619 through 25 625 and 25 657 for control surface hinges must be shown by analysis or individual load tests 825 655 Installation a Movab
522. rtical ground reaction 2 The most severe combination of loads that are likely to arise during a lateral drift landing must be taken into account In absence of a more rational analysis of this condition the following must be investigated i A vertical load equal to 75 of the maximum ground reaction of 825 473 must be considered in combination with a drag and side load of 4096 and 25 respectively of that vertical load ii The shock absorber and tire deflections must be assumed to be 75 of the deflection cor responding to the maximum ground reaction of 25 473 a 2 This load case need not be consid ered in combination with flat tires ASA Part 25 Airworthiness Standards Transport Category 3 The combination of vertical and drag com ponents is considered to be acting at the wheel axle centerline Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5673 April 8 1970 Amdt 25 91 62 FR 40705 July 29 1997 Amdt 25 91 62 FR 45481 Aug 27 1997 825 481 Tail down landing conditions a In the tail down attitude the airplane is as sumed to contact the ground at forward velocity components ranging from V 4 to Vi parallel to the ground under the conditions prescribed in 825 473 with 1 Vi4 equal to Vso TAS at the appropriate landing weight and in standard sea level condi tions and 2 Vi2 equal to Vso TAS at the appropriate landing weight and altitudes in a hot day
523. rtment view 2 825 981 Fuel tank ignition prevention 3 825 1165 Engine ignition systems 4 825 1310 Power source capacity and distri bution 5 825 1316 System lightning protection 6 825 1331 a 2 Instruments using a power supply 7 825 1351 General 8 825 1355 Distribution system 9 825 1360 Precautions against injury 10 825 1362 Electrical supplies for emer gency conditions 11 825 1365 Electrical appliances motors and transformers 12 825 1431 c and d Electronic equipment 143 825 1707 825 1707 System separation EWIS a Each EWIS must be designed and installed with adequate physical separation from other EWIS and airplane systems so that an EWIS component failure will not create a hazardous condition Unless otherwise stated for the pur poses of this section adequate physical separa tion must be achieved by separation distance or by a barrier that provides protection equivalent to that separation distance b Each EWIS must be designed and installed so that any electrical interference likely to be present in the airplane will not result in hazardous effects upon the airplane or its systems c Wires and cables carrying heavy current and their associated EWIS components must be designed and installed to ensure adequate physi cal separation and electrical isolation so that damage to circuits associated with essential func tions will be minimized under fault conditions d
524. ructed space 2 In addition to the access i For airplanes that have a passenger seating configuration of 20 or more the projected opening required by ASA Part 25 Airworthiness Standards Transport Category of the exit provided must be obstructed and there must be no interference in opening the exit by seats berths or other protrusions including any seatback in the most adverse position for a dis tance from that exit not less than the width of the narrowest passenger seat installed on the air plane ii For airplanes that have a passenger seating configuration of 19 or fewer there may be minor obstructions in this region if there are compen sating factors to maintain the effectiveness of the exit 3 For each Type Ill exit regardless of the pas senger capacity of the airplane in which it is in stalled there must be placards that i Are readable by all persons seated adjacent to and facing a passageway to the exit ii Accurately state or illustrate that proper method of opening the exit including the use of handholds and iii If the exit is a removable hatch state the weight of the hatch and indicate an appropriate lo cation to place the hatch after removal d If it is necessary to pass through a pas sageway between passenger compartments to reach any required emergency exit from any seat in the passenger cabin the passageway must be unobstructed However curtains may be used if they allow
525. rvicing c EWIS must be designed and installed to minimize damage and risk of damage to EWIS by items carried onto the aircraft by passengers or cabin crew 825 1723 Flammable fluid fire protection EWIS EWIS components located in each area where flammable fluid or vapors might escape by leak age of a fluid system must be considered a poten tial ignition source and must meet the require ments of 825 863 825 1725 Powerplants EWIS a EWIS associated with any powerplant must be designed and installed so that the failure of an EWIS component will not prevent the continued safe operation of the remaining powerplants or re quire immediate action by any crewmember for continued safe operation in accordance with the requirements of 25 903 b b Design precautions must be taken to mini mize hazards to the airplane due to EWIS dam age in the event of a powerplant rotor failure or a fire originating within the powerplant that burns through the powerplant case in accordance with the requirements of 825 903 d 1 145 SFAR No 109 to Part 25 825 1727 Flammable fluid shutoff means EWIS EWIS associated with each flammable fluid shutoff means and control must be fireproof or must be located and protected so that any fire in a fire zone will not affect operation of the flammable fluid shutoff means in accordance with the re quirements of 825 1189 825 1729 Instructions for Continued Airworthiness EWIS The app
526. rwise indicated to the pilot by the air speed indicating system required under para graph b 1 of this section Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5678 April 8 1970 Amdt 25 24 35 FR 7108 May 6 1970 Amdt 25 38 41 FR 55467 Dec 20 1976 Amdt 25 90 62 FR 13253 March 19 1997 825 1305 Powerplant instruments The following are required powerplant instru ments a For all airplanes 1 A fuel pressure warning means for each en gine or a master warning means for all engines with provision for isolating the individual warning means from the master warning means 2 A fuel quantity indicator for each fuel tank 3 An oil quantity indicator for each oil tank 4 An oil pressure indicator for each indepen dent pressure oil system of each engine 5 An oil pressure warning means for each en gine or a master warning means for all engines with provision for isolating the individual warning means from the master warning means 6 An oil temperature indicator for each engine 7 Fire warning indicators 8 An augmentation liquid quantity indicator appropriate for the manner in which the liquid is to be used in operation for each tank b For reciprocating engine powered planes In addition to the powerplant instruments required by paragraph a of this section the fol lowing powerplant instruments are required 1 A carburetor air temperature indi
527. s Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55466 Dec 20 1976 825 689 Cable systems a Each cable cable fitting turnbuckle splice and pulley must be approved In addition 1 No cable smaller than 1 inch in diameter may be used in the aileron elevator or rudder Systems and 2 Each cable system must be designed so that there will be no hazardous change in cable ASA 825 697 tension throughout the range of travel under oper ating conditions and temperature variations b Each kind and size of pulley must corre spond to the cable with which it is used Pulleys and sprockets must have closely fitted guards to prevent the cables and chains from being dis placed or fouled Each pulley must lie in the plane passing through the cable so that the cable does not rub against the pulley flange c Fairleads must be installed so that they do not cause a change in cable direction of more than three degrees d Clevis pins subject to load or motion and re tained only by cotter pins may not be used in the control system e Turnbuckles must be attached to parts hav ing angular motion in a manner that will positively prevent binding throughout the range of travel f There must be provisions for visual inspection of fairleads pulleys terminals and turnbuckles 825 693 Joints Control system joints in push pull systems that are subject to angular motion except tho
528. s There must be a separate carburetor air tem perature control for each engine 825 1159 Supercharger controls Each supercharger control must be accessible to the pilots or if there is a separate flight engineer station with a control panel to the flight engineer 825 1161 Fuel jettisoning system controls Each fuel jettisoning system control must have guards to prevent inadvertent operation No con trol may be near any fire extinguisher control or other control used to combat fire 25 1163 Powerplant accessories a Each engine mounted accessory must 1 Be approved for mounting on the engine in volved 2 Use the provisions on the engine for mount ing and 3 Be sealed to prevent contamination of the engine oil system and the accessory system b Electrical equipment subject to arcing or sparking must be installed to minimize the proba bility of contact with any flammable fluids or va pors that might be present in a free state c If continued rotation of an engine driven cabin supercharger or of any remote accessory driven by the engine is hazardous if malfunction ing occurs there must be means to prevent rota tion without interfering with the continued opera tion of the engine Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 57 49 FR 6849 Feb 23 1984 25 1165 Engine ignition systems a Each battery ignition system must be sup plemented by a generator that is
529. s b Continuous Gust Design Criteria The dy namic response of the airplane to vertical and lat eral continuous turbulence must be taken into ac count The continuous gust design criteria of Ap pendix G of this part must be used to establish the dynamic response unless more rational criteria are shown Docket No 27902 61 FR 5221 Feb 9 1996 as amended by Amdt 25 86 61 FR 9533 March 8 1996 825 343 Design fuel and oil loads a The disposable load combinations must in clude each fuel and oil load in the range from zero fuel and oil to the selected maximum fuel and oil load A structural reserve fuel condition not ex ceeding 45 minutes of fuel under the operating conditions in 25 1001 e and f as applicable may be selected b If a structural reserve fuel condition is se lected it must be used as the minimum fuel 39 825 345 weight condition for showing compliance with the flight load requirements as prescribed in this sub part In addition 1 The structure must be designed for a condi tion of zero fuel and oil in the wing at limit loads corresponding to i A maneuvering load factor of 2 25 and ii The gust conditions of 25 341 a but as suming 85 of the design velocities prescribed in 825 341 a 4 2 Fatigue evaluation of the structure must ac count for any increase in operating stresses re sulting from the design condition of paragraph b 1 of this section and 3 The
530. s 25 399 Dual control system 25 405 Secondary control system 25 407 Trim tab effects 25 409 Tabs 25 415 Ground gust conditions 25 427 Unsymmetrical loads 25 445 Auxiliary aerodynamic surfaces 25 457 Wing flaps 25 459 Special devices GROUND LOADS 25 471 General 25 473 Landing load conditions and assumptions 25 477 Landing gear arrangement 25 479 Level landing conditions 25 481 Tail down landing conditions 25 483 One gear landing conditions 25 485 Side load conditions 25 487 Rebound landing condition 25 489 Ground handling conditions 25 491 Taxi takeoff and landing roll 25 493 Braked roll conditions 25 495 Turning 25 497 Tail wheel yawing 25 499 Nose wheel yaw and steering 25 503 Pivoting 25 507 Reversed braking 25 509 Towing loads 25 511 Ground load unsymmetrical loads on multiple wheel units 25 519 Jacking and tie down provisions WATER LOADS 25 521 General 25 523 Design weights and center of gravity positions 25 525 Application of loads 25 527 Hull and main float load factors 25 529 Hull and main float landing conditions 25 531 Hull and main float takeoff condition 25 533 Hull and main float bottom pressures 25 535 Auxiliary float loads 25 537 Seawing loads EMERGENCY LANDING CONDITIONS 25 561 General 25 562 Emergency landing dynamic conditions 25 563 Structural ditching provisions 25 571 25 581 Federal Aviation Regulations FATIGUE
531. s Burnthrough time means the time in sec onds for the burner flame to penetrate the test specimen and or the time required for the heat flux to reach 2 0 Btu ft2sec 2 27 W cm on the inboard side at a distance of 12 inches 30 5 cm from the front surface of the insulation blanket test ASA Part 25 Airworthiness Standards Transport Category frame whichever is sooner The burnthrough time is measured at the inboard side of each of the in sulation blanket specimens Insulation blanket specimen means one of two specimens positioned in either side of the test rig at an angle of 30 with respect to vertical Specimen set means two insulation blanket specimens Both specimens must represent the same production insulation blanket construction and materials proportioned to correspond to the specimen size b Apparatus 1 The arrangement of the test apparatus is shown in figures 1 and 2 and must include the ca pability of swinging the burner away from the test specimen during warm up SEE FIGURE 1 AT THE END OF PART VII OF THIS APPENDIX 2 Test burner The test burner must be a mod ified gun type such as the Park Model DPL 3400 Flame characteristics are highly dependent on actual burner setup Parameters such as fuel pressure nozzle depth stator position and intake airflow must be properly adjusted to achieve the correct flame output SEE FIGURE 2 AT THE END OF PART VII OF THIS APPENDIX i Nozzle A nozzle mu
532. s Each fuel tank must be vented from the top part of the expansion space so that venting is effective under any normal flight condition In addition 1 Each vent must be arranged to avoid stop page by dirt or ice formation 2 The vent arrangement must prevent siphon ing of fuel during normal operation 3 The venting capacity and vent pressure lev els must maintain acceptable differences of pres sure between the interior and exterior of the tank during i Normal flight operation ii Maximum rate of ascent and descent and iii Refueling and defueling where applicable 4 Airspaces of tanks with interconnected out lets must be interconnected ASA Part 25 Airworthiness Standards Transport Category 5 There may be no point in any vent line where moisture can accumulate with the airplane in the ground attitude or the level flight attitude unless drainage is provided and 6 No vent or drainage provision may end at any point i Where the discharge of fuel from the vent outlet would constitute a fire hazard or ii From which fumes could enter personnel compartments b Carburetor vapor vents Each carburetor with vapor elimination connections must have a vent line to lead vapors back to one of the fuel tanks In addition 1 Each vent system must have means to avoid stoppage by ice and 2 If there is more than one fuel tank and it is necessary to use the tanks in a definite se quence
533. s The ventilation system ducting must be protected by a flame arrestor Note The appli cant may find additional useful information in So ciety of Automotive Engineers Aerospace Rec ommended Practice 85 Rev E entitled Air Con ditioning Systems for Subsonic Airplanes dated August 1 1991 g Means must be provided to contain spilled foods or fluids in a manner that will prevent the creation of a slipping hazard to occupants and will not lead to the loss of structural strength due to airplane corrosion h Cooktop installations must provide ade quate space for the user to immediately escape a hazardous cooktop condition i A means to shut off power to the cooktop must be provided at the galley containing the cooktop and in the cockpit If additional switches are introduced in the cockpit revisions to smoke or fire emergency procedures of the AFM will be required j If the cooktop is required to have a lid to en close the cooktop there must be a means to auto matically shut off power to the cooktop when the lid is closed ASA 15 Hand Held Fire Extinguishers a For airplanes that were originally type certif icated with more than 60 passengers the number of hand held fire extinguishers must be the greater of 1 That provided in accordance with the re quirements of 825 851 2 A number equal to the number of originally type certificated exit pairs regardless of whether the exits are deactivated
534. s Transport Category 2 There must be a flight crew warning light which illuminates when power is on in the airplane and the emergency lighting control device is not armed 3 The cockpit control device must have an on off and armed position so that when armed in the cockpit or turned on at either the cockpit or flight attendant station the lights will ei ther light or remain lighted upon interruption ex cept an interruption caused by a transverse verti cal separation of the fuselage during crash land ing of the airplane s normal electric power There must be a means to safeguard against inadvert ent operation of the control device from the armed or on positions g Exterior emergency lighting must be pro vided as follows 1 At each overwing emergency exit the illumi nation must be i Not less than 0 03 foot candle measured normal to the direction of the incident light on a 2 square foot area where an evacuee is likely to make his first step outside the cabin ii Not less than 0 05 foot candle measured normal to the direction of the incident light for a minimum width of 42 inches for a Type A overwing emergency exit and two feet for all other overwing emergency exits along the 30 percent of the slip resistant portion of the escape route required in 25 810 c that is farthest from the exit and iii Not less than 0 03 foot candle on the ground surface with the landing gear extended measured normal
535. s Transport Category 3 There must be a visual means on the flight deck to signal the pilots if any door is not fully closed latched and locked The means must be designed such that any failure or combination of failures that would result in an erroneous closed latched and locked indication is improbable for i Each door that is subject to pressurization and for which the initial opening movement is not inward or ii Each door that could be a hazard if un latched 4 There must be an aural warning to the pilots prior to or during the initial portion of takeoff roll if any door is not fully closed latched and locked and its opening would prevent a safe takeoff and return to landing f Visual inspection provision Each door for which unlatching of the door could be a hazard must have a provision for direct visual inspection to determine without ambiguity if the door is fully closed latched and locked The provision must be permanent and discernible under operational lighting conditions or by means of a flashlight or equivalent light source g Certain maintenance doors removable emergency exits and access panels Some doors not normally opened except for mainte nance purposes or emergency evacuation and Some access panels need not comply with certain paragraphs of this section as follows 1 Access panels that are not subject to cabin pressurization and would not be a hazard if open during flight need not com
536. s which may remain powered by the same source when all other power sources are inoperative and 2 An additional time duration in its standby state appropriate or required for any other loads that are powered by the same source and that are essential to safety of flight or required during emergency conditions b Be capable of operation within 10 seconds by a flight attendant at those stations in the pas senger compartment from which they system is accessible c Be intelligible at all passenger seats lavato ries and flight attendant seats and work stations d Be designed so that no unused unstowed microphone will render the system inoperative e Be capable of functioning independently of any required crewmember interphone system ASA 825 1435 f Be accessible for immediate use from each of two flight crewmember stations in the pilot com partment g For each required floor level passenger emergency exit which has an adjacent flight at tendant seat have a microphone which is readily accessible to the seated flight attendant except that one microphone may serve more than one exit provided the proximity of the exits allows un assisted verbal communication between seated flight attendants Docket No 26003 58 FR 45229 Aug 26 1993 825 1431 Electronic equipment a In showing compliance with 25 1309 a and b with respect to radio and electronic equip ment and their installations critical environ
537. s an approved built in fire extinguish ing or suppression system controllable from the cockpit 3 There are means to exclude hazardous quan tities of smoke flames or extinguishing agent from any compartment occupied by the crew or passen gers 4 There are means to control ventilation and drafts within the compartment so that the extin guishing agent used can control any fire that may start within the compartment d Reserved e Class E A Class E cargo compartment is one on airplanes used only for the carriage of cargo and in which 1 Reserved 2 There is a separate approved smoke or fire detector system to give warning at the pilot or flight engineer station 3 There are means to shut off the ventilating airflow to or within the compartment and the controls for these means are accessible to the flight crew in the crew compartment 4 There are means to exclude hazardous quantities of smoke flames or noxious gases from the flight crew compartment and 5 The required crew emergency exits are ac cessible under any cargo loading condition Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 32 37 FR 3972 Feb 24 1972 Amdt 25 60 51 FR 18243 May 16 1986 Amdt 25 93 63 FR 8048 Feb 17 1998 525 858 Cargo or baggage compartment smoke or fire detection systems If certification with cargo or baggage compart ment smoke or fire detection provisions is re quested the
538. s and pres sures from fluid volumetric changes in elements that are likely to remain closed long enough for such changes to occur are within the design ca pabilities of each element such that they meet the requirements defined in 25 1435 a 1 through a 5 3 Have means to minimize the release of harmful or hazardous concentrations of hydraulic fluid or vapors into the crew and passenger com partments during flight 4 Meet the applicable requirements of 25 863 25 1183 25 1185 and 25 1189 if a flammable hydraulic fluid is used and 5 Be designed to use any suitable hydraulic fluid specified by the airplane manufacturer which must be identified by appropriate markings as re quired by 25 1541 c Tests Tests must be conducted on the hy draulic system s and or subsystem s and ele ments except that analysis may be used in place of or to supplement testing where the analysis is shown to be reliable and appropriate All internal 130 Federal Aviation Regulations and external influences must be taken into ac count to an extent necessary to evaluate their ef fects and to assure reliable system and element functioning and integration Failure or unaccept able deficiency of an element or system must be corrected and be sufficiently retested where nec essary 1 The system s subsystem s or element s must be subjected to performance fatigue and endurance tests representative of airplane ground an
539. s defined as the entire auto matic system used on takeoff including all de vices both mechanical and electrical that sense engine failure transmit signals actuate fuel con trols or power levers or increase engine power by other means on operating engines to achieve scheduled thrust or power increases and furnish cockpit information on system operation ASA Appendix to Part 25 b Critical Time Interval When conducting an ATTCS takeoff the critical time interval is be tween V4 minus 1 second and a point on the min imum performance all engine flight path where assuming a simultaneous occurrence of an en gine and ATTCS failure the resulting minimum flight path thereafter intersects the Part 25 re quired actual flight path at no less than 400 feet above the takeoff surface This time interval is shown in the following illustration 217 Appendix to Part 25 Height above runway surface ft 218 Engine and ATTC failure Federal Aviation Regulations Flight path with ATTCS and engine failure Critical time interval ASA Part 25 Airworthiness Standards Transport Category 125 3 PERFORMANCE AND SYSTEM RELIABILITY REQUIREMENTS The applicant must comply with the perfor mance and ATTCS reliability requirements as fol lows a An ATTCS failure or a combination of fail ures in the ATTCS during the critical time interval 1 Shall not pr
540. s essential to safe operation of the airplane and that is located in wheel wells must be protected from the damaging effects of 65 825 731 1 A bursting tire unless it is shown that a tire cannot burst from overheat and 2 A loose tire tread unless it is shown that a loose tire tread cannot cause damage Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5676 April 8 1970 Amdt 25 42 43 FR 2323 Jan 16 1978 Amdt 25 72 55 FR 29777 July 20 1990 Amdt 25 75 56 FR 63762 Dec 5 1991 825 731 Wheels a Each main and nose wheel must be ap proved b The maximum static load rating of each wheel may not be less than the corresponding static ground reaction with 1 Design maximum weight and 2 Critical center of gravity c The maximum limit load rating of each wheel must equal or exceed the maximum radial limit load determined under the applicable ground load requirements of this part d Overpressure burst prevention Means must be provided in each wheel to prevent wheel failure and tire burst that may result from exces sive pressurization of the wheel and tire assem bly e Braked wheels Each braked wheel must meet the applicable requirements of 525 735 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29777 July 20 1990 Amdt 25 107 67 FR 20420 April 24 2002 825 733 Tires a When a landing gear axle is fitt
541. s i at the chine 0 0016 K2 hull station weighing factor in accordance with figure 2 of Appendix B seaplane stalling speed at the design water takeoff weight with flaps extended in the appropriate takeoff position and angle of dead rise at appropriate station The area over which these pressures are applied must simulate pressures occurring during high lo ASA 825 535 calized impacts on the hull or float but need not extend over an area that would induce critical stresses in the frames or in the overall structure c Distributed pressures For the design of the frames keel and chine structure the following pressure distributions apply 1 Symmetrical pressures are computed as follows K5Vsg P CX 4 tan where P pressure p s i 0 078 C4 with C4 computed under 825 527 K hull station weighing factor determined in accordance with figure 2 of Appendix B Vso seaplane stalling speed Knots with landing flaps extended in the appropriate position and with no slipstream effect and Vso seaplane stalling speed with landing flaps extended in the appropriate position and with no slipstream effect and angle of dead rise at appropriate station 2 The unsymmetrical pressure distribution consists of the pressures prescribed in paragraph c 1 of this section on one side of the hull or main float centerline and one half of that pressure on the other side
542. s installed the overnight temperature drop for this appendix is defined us ing 1 A temperature at the beginning of the over night period that equals the landing temperature of the previous flight that is a random value based on a Gaussian distribution and 2 An overnight temperature drop that is a ran dom value based on a Gaussian distribution 3 For any flight that will end with an overnight ground period one flight per day out of an aver age number of flights per day depending on utili zation of the particular airplane model being eval uated the landing outside air temperature OAT is to be chosen as a random value from the follow ing Gaussian curve Federal Aviation Regulations TABLE 3 LANDING OUTSIDE AIR TEMPERATURE Landing outside air Parameter temperature F Mean Temperature 58 68 negative 1 std dev 20 55 positive 1 std dev 13 21 4 The outside ambient air temperature OAT overnight temperature drop is to be chosen as a random value from the following Gaussian curve TABLE 4 OUTSIDE AIR TEMPERATURE OAT DROP Parameter OAT drop temperature F Mean Temp 12 0 1 std dev 6 0 d Number of Simulated Flights Required in Analysis order for the Monte Carlo analysis to be valid for showing compliance with the fleet average and warm day flammability exposure require ments the applicant must run the analysis for a minimum number of flights to ensure that the flee
543. s investigated b Reserved c The controllability stability trim and stalling characteristics of the airplane must be shown for each altitude up to the maximum expected in op eration d Parameters critical for the test being con ducted such as weight loading center of gravity and inertia airspeed power and wind must be maintained within acceptable tolerances of the critical values during flight testing e If compliance with the flight characteristics requirements is dependent upon a stability aug mentation system or upon any other automatic or power operated system compliance must be shown with 25 671 and 25 672 f In meeting the requirements of 25 105 d 25 125 25 233 and 25 237 the wind velocity must be measured at a height of 10 meters above the surface or corrected for the difference be tween the height at which the wind velocity is mea sured and the 10 meter height g The requirements of this subpart associ ated with icing conditions apply only if the appli cant is seeking certification for flight in icing condi tions 1 Each requirement of this subpart except 25 121 a 25 123 c 25 143 b 1 and b 2 25 149 25 201 c 2 25 207 c and d 25 239 and 25 251 b through e must be met in icing conditions Compliance must be shown using the ice accretions defined in appendix C assuming normal operation of the airplane and its ice pro tection system in accordance with the
544. s must use the subset of those flights that begin with a sea level ground ambient temperature of 80 F standard day plus 21 F at mosphere or above from the flammability expo sure analysis done for overall performance 2 For the ground and takeoff climb phases of flight the average flammability exposure must be calculated by dividing the time during the specific flight phase the fuel tank is flammable by the total time of the specific flight phase 3 Compliance with this paragraph may be shown using only those flights for which the air plane is dispatched with the flammability reduc tion means operational 230 Federal Aviation Regulations M25 2 SHOWING COMPLIANCE a The applicant must provide data from analy sis ground testing and flight testing or any com bination of these that 1 Validate the parameters used in the analy sis required by paragraph M25 1 of this appendix 2 Substantiate that the FRM is effective at lim iting flammability exposure in all compartments of each tank for which the FRM is used to show compliance with paragraph M25 1 of this appen dix and 3 Describe the circumstances under which the FRM would not be operated during each phase of flight b The applicant must validate that the FRM meets the requirements of paragraph M25 1 of this appendix with any airplane or engine configu ration affecting the performance of the FRM for which approval is sought M25 3 RELIABILITY INDI
545. s of components c Service experience of aircraft with similar powerplant configurations d Analysis Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 46 43 FR 50598 Oct 30 1978 ASA Part 25 Airworthiness Standards Transport Category Subpart F Equipment GENERAL 825 1301 Function and installation a Each item of installed equipment must 1 Be of a kind and design appropriate to its in tended function 2 Be labeled as to its identification function or operating limitations or any applicable combi nation of these factors 3 Be installed according to limitations speci fied for that equipment and 4 Function properly when installed b EWIS must meet the requirements of Sub part H of this part Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 123 72 FR 63405 Nov 8 2007 825 1303 Flight and navigation instruments a The following flight and navigation instru ments must be installed so that the instrument is visible from each pilot station 1 A free air temperature indicator or an air temperature indicator which provides indications that are convertible to free air temperature 2 A clock displaying hours minutes and sec onds with a sweep second pointer or digital pre sentation 3 A direction indicator nonstabilized mag netic compass b The following flight and navigation instru ments must be installed at each pilot station
546. s of the calorimeter and the ther mocouples 7 Weight Scale Weighing Device A device must be used that with proper procedures may determine the before and after test weights of each set of seat cushion specimens within 0 02 pound 9 grams A continuous weighing system is preferred 8 Timing Device A stopwatch or other device calibrated to 1 second must be used to mea sure the time of application of the burner flame and self extinguishing time or test duration e Preparation of Apparatus Before calibra tion all equipment must be turned on and the burner fuel must be adjusted as specified in para graph d 2 f Calibration To ensure the proper thermal output of the burner the following test must be made 1 Place the calorimeter on the test stand as shown in Figure 4 at a distance of 4 inches 102 3 mm from the exit of the burner cone 2 Turn on the burner allow it to run for 2 min utes for warmup and adjust the burner air intake damper to produce a reading of 10 5 0 5 BTU ft2 11 9 0 6 w cm on the calorimeter to en sure steady state conditions have been achieved Turn off the burner 3 Replace the calorimeter with the thermo couple rake Figure 5 4 Turn on the burner and ensure that the ther mocouples are reading 1900 100 F 1038 38 C to ensure steady state conditions have been achieved 169 Appendix to Part 25 5 If the calorimeter and thermocouples do not r
547. s to asa asa2fly com 1 Icing conditions maximum intensities tables Appendix C 154 Ignition switches 25 1145 110 Induction system ducts 25 1103 107 icing protection 25 1093 2 107 screens 25 1105 108 1222 424 In flight shutdown IFSD rates Appendix K Injury precautions 25 1360 Installation of airspeed indication system 25 1323 flight guidance system 25 1329 instrument systems 25 1333 E instruments using a power supply 25 1331 121 119 magnetic direction indicator 25 1327 120 pitot heat indication system 25 13206 120 powerplant instruments 25 1337 122 static pressure system 25 1325 119 warning caution advisory lights 825 1322 118 Instrument markings 25 1543 138 Instruments installation arrangement and visibility 25 1321 118 108 Inter coolers 25 1107 L Landing 25 125 Landing climb 25 119 Lightning protection for systems Load factors limit maneuvering 25 337 38 Loads gyroscopic 25 371 unsymmetrical 25 367 25 427 Markings and placards airspeed information 25 1545 2 138 cockpit control 25 1555 139 magnetic direction indicator 25 1547 138 powerplant APU instruments 25 1549 138 safety equipment 25 1567 2 199 specifications 25 1541 25 1563
548. se 5 Approach ice is the critical ice accretion on the unprotected surfaces and any ice accretion on the protected surfaces appropriate to normal ice protection system operation following exit from the holding flight phase and transition to the most critical approach configuration 6 Landing ice is the critical ice accretion on the unprotected surfaces and any ice accretion on the protected surfaces appropriate to normal ice protection system operation following exit from the approach flight phase and transition to the fi nal landing configuration b In order to reduce the number of ice accre tions to be considered when demonstrating com pliance with the requirements of 825 21 g any of the ice accretions defined in paragraph a of this section may be used for any other flight phase if it is shown to be more critical than the specific ice accretion defined for that flight phase Configura tion differences and their effects on ice accretions must be taken into account c The ice accretion that has the most adverse effect on handling qualities may be used for air plane performance tests provided any difference in performance is conservatively taken into ac count 155 Appendix C to Part 25 d For both unprotected and protected parts the ice accretion for the takeoff phase may be de termined by calculation assuming the takeoff maximum icing conditions defined in appendix C and assuming that 1 Airfoils contr
549. se in ball and roller bearing systems must have a spe cial factor of safety of not less than 3 33 with re spect to the ultimate bearing strength of the soft est material used as a bearing This factor may be reduced to 2 0 for joints in cable control systems For ball or roller bearings the approved ratings may not be exceeded Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29777 July 20 1990 825 697 Lift and drag devices controls a Each lift device control must be designed so that the pilots can place the device in any takeoff en route approach or landing position estab lished under 825 101 d Lift and drag devices must maintain the selected positions except for movement produced by an automatic positioning or load limiting device without further attention by the pilots b Each lift and drag device control must be designed and located to make inadvertent opera tion improbable Lift and drag devices intended for ground operation only must have means to pre vent the inadvertent operation of their controls in flight if that operation could be hazardous c The rate of motion of the surfaces in re sponse to the operation of the control and the characteristics of the automatic positioning or load limiting device must give satisfactory flight and performance characteristics under steady or changing conditions of airspeed engine power and airplane attitude d The lift device contro
550. sec Watts cm Zero Position 1 5 14 Position 1 1 51 1 50 1 49 1 71 1 70 1 69 Position 2 1 43 1 44 1 62 1 63 4 Open the bottom door remove the eter and holder fixture Use caution as the fixture is very hot f Test Procedure 1 Ignite the pilot burner Ensure that it is at least 2 inches 51 mm above the top of the plat form The burner must not contact the specimen until the test begins 2 Place the test specimen in the sliding plat form holder Ensure that the test sample surface is level with the top of the platform At zero point the specimen surface must be 7 1 2 inches 1 8 inch 191 mm 3 below the radiant panel 3 Place the retaining securing frame over the test specimen It may be necessary due to com pression to adjust the sample up or down in or der to maintain the distance from the sample to the radiant panel 7 1 2 inches 1 8 inch 191 mm 3 at zero position With film fiberglass assem blies it is critical to make a slit in the film cover to purge any air inside This allows the operator to maintain the proper test specimen position level with the top of the platform and to allow ventila tion of gases during testing A longitudinal slit ap proximately 2 inches 51mm in length must be centered 3 inches 1 2 inch 76mm x 13mm from the left flange of the securing frame A utility knife is acceptable for slitting the film cover 4 Immediately push the
551. sec tion K25 2 2 h of this appendix g Airplane demonstration For each airplane engine combination to be approved for ETOPS the applicant must flight test at least one airplane to demonstrate that the airplane and its compo nents and equipment are capable of functioning properly during ETOPS flights and diversions of the longest duration for which the applicant seeks approval This flight testing may be performed in conjunction with but may not substitute for the flight testing required by 21 35 b 2 of this chap ter 1 The airplane demonstration flight test pro gram must include i Flights simulating actual ETOPS including flight at normal cruise altitude step climbs and if applicable APU operation ii Maximum duration flights with maximum du ration diversions iii Maximum duration engine inoperative di versions distributed among the engines installed on the airplanes used for the airplane demonstra tion flight test program At least two one engine inoperative diversions must be conducted at max imum continuous thrust or power using the same engine iv Flights under non normal conditions to demonstrate the flightcrew s ability to safely con duct an ETOPS diversion with worst case ETOPS ASA Appendix K to Part 25 significant system failures or malfunctions that could occur in service v Diversions to airports that represent air ports of the types used for ETOPS diversions vi Repeated ex
552. sig nals are not picked up by another channel 5 As far as is practicable all sounds received by the microphone listed in paragraphs c 1 2 and 4 of this section must be recorded without interruption irrespective of the position of the in terphone transmitter key switch The design shall ensure that sidetone for the flight crew is pro duced only when the interphone public address System or radio transmitters are in use d Each cockpit voice recorder must be in stalled so that 1 i It receives its electrical power from the bus that provides the maximum reliability for oper ation of the cockpit voice recorder without jeopar dizing service to essential or emergency loads ii It remains powered for as long as possible without jeopardizing emergency operation of the airplane 2 There is an automatic means to simulta neously stop the recorder and prevent each era sure feature from functioning within 10 minutes after crash impact 3 There is an aural or visual means for preflight checking of the recorder for proper operation 4 Any single electrical failure external to the re corder does not disable both the cockpit voice re corder and the flight data recorder 5 It has an independent power source i That provides 10 1 minutes of electrical power to operate both the cockpit voice recorder and cockpit mounted area microphone ii That is located as close as practicable to the cockpit voice record
553. sliding platform into the chamber and close the bottom door 5 Bring the pilot burner flame into contact with the center of the specimen at the zero point and simultaneously start the timer The pilot burner must be at a 27 angle with the sample and be ap proximately 1 2 inch 12 mm above the sample See figure 7 A stop as shown in figure 8 allows the operator to position the burner correctly each time 198 Federal Aviation Regulations FiGURE 8 Propane Burner Stop 6 Leave the burner in position for 15 seconds and then remove to a position at least 2 inches 51 mm above the specimen g Report 1 Identify and describe the test specimen 2 Report any shrinkage or melting of the test specimen 3 Report the flame propagation distance If this distance is less than 2 inches report this as a pass no measurement required 4 Report the after flame time h Requirements 1 There must be no flame propagation be yond 2 inches 51 mm to the left of the centerline of the pilot flame application 2 The flame time after removal of the pilot burner may not exceed 3 seconds on any speci men PART VII TEST METHOD TO DETERMINE THE BURNTHROUGH RESISTANCE OF THERMAL ACOUSTIC INSULATION MATERIALS Use the following test method to evaluate the burnthrough resistance characteristics of aircraft thermal acoustic insulation materials when ex posed to a high intensity open flame a Definition
554. solated from the rest of the airplane by firewalls shrouds or equivalent means b Each firewall and shroud must be 1 Fireproof 2 Constructed so that no hazardous quantity of air fluid or flame can pass from the compart ment to other parts of the airplane 3 Constructed so that each opening is sealed with close fitting fireproof grommets bushings or firewall fittings and 4 Protected against corrosion 825 1192 Engine accessory section diaphragm For reciprocating engines the engine power section and all portions of the exhaust system must be isolated from the engine accessory com partment by a diaphragm that complies with the firewall requirements of 825 1191 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5678 April 8 1970 25 1193 Cowling and nacelle skin a Each cowling must be constructed and sup ported so that it can resist any vibration inertia and air load to which it may be subjected in oper ation b Cowling must meet the drainage and venti lation requirements of 25 1187 c On airplanes with a diaphragm isolating the engine power section from the engine accessory section each part of the accessory section cowl ing subject to flame in case of fire in the engine power section of the powerplant must 1 Be fireproof and 2 Meet the requirements of 25 1191 d Each part of the cowling subject to high temperatures due to its nearness to ex
555. ssenger convenience must be completely enclosed b There must be a means to prevent the con tents in the compartments from becoming a haz ard by shifting under the loads specified in para graph a of this section For stowage compart ments in the passenger and crew cabin if the means used is a latched door the design must ASA Part 25 Airworthiness Standards Transport Category take into consideration the wear and deterioration expected in service c If cargo compartment lamps are installed each lamp must be installed so as to prevent con tact between lamp bulb and cargo Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 32 37 FR 3969 Feb 24 1972 Amdt 25 38 41 FR 55466 Dec 20 1976 Amdt 25 51 45 FR 7755 Feb 4 1980 825 789 Retention of items of mass in passenger and crew compartments and galleys a Means must be provided to prevent each item of mass that is part of the airplane type de sign in a passenger or crew compartment or gal ley from becoming a hazard by shifting under the appropriate maximum load factors corresponding to the specified flight and ground load conditions and to the emergency landing conditions of 825 561 b b Each interphone restraint system must be designed so that when subjected to the load fac tors specified in 25 561 b 3 the interphone will remain in its stowed position Docket No 5066 29 FR 18291 Dec 24 1964 as amended by
556. st be investigated in flight to determine that no adverse characteristics such as stall surge or flameout are present to a hazardous degree during normal and emergency operation within the range of operating limitations of the airplane and of the engine b Reserved c The turbine engine air inlet system may not as a result of air flow distortion during normal op eration cause vibration harmful to the engine Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 11 32 FR 6912 May 5 1967 Amdt 25 40 42 FR 15043 March 17 1977 525 941 Inlet engine and exhaust compatibility For airplanes using variable inlet or exhaust System geometry or both a The system comprised of the inlet engine including thrust augmentation systems if incor porated and exhaust must be shown to function properly under all operating conditions for which approval is sought including all engine rotating speeds and power settings and engine inlet and exhaust configurations b The dynamic effects of the operation of these including consideration of probable mal functions upon the aerodynamic control of the airplane may not result in any condition that would require exceptional skill alertness or strength on the part of the pilot to avoid exceeding an opera tional or structural limitation of the airplane and c In showing compliance with paragraph b of this section the pilot strength required may not exc
557. st be provided to assist the occu pants in locating the exits in conditions of dense smoke d The location of each passenger emergency exit must be indicated by a sign visible to occu pants approaching along the main passenger aisle or aisles There must be 1 A passenger emergency exit locator sign above the aisle or aisles near each passenger emergency exit or at another overhead location if it is more practical because of low headroom ex cept that one sign may serve more than one exit if each exit can be seen readily from the sign ASA 25 811 2 A passenger emergency exit marking sign next to each passenger emergency exit except that one sign may serve two such exits if they both can be seen readily from the sign and 3 A sign on each bulkhead or divider that pre vents fore and aft vision along the passenger cabin to indicate emergency exits beyond and ob scured by the bulkhead or divider except that if this is not possible the sign may be placed at an other appropriate location e The location of the operating handle and in structions for opening exits from the inside of the airplane must be shown in the following manner 1 Each passenger emergency exit must have on or near the exit a marking that is readable from a distance of 30 inches 2 Each Type A Type B Type C or Type pas senger emergency exit operating handle must i Be self illuminated with an initial brightness of at least
558. st maintain the fuel pressure to yield a nominal 6 0 gal hr 0 378 L min fuel flow A Monarch manufactured 80 PL hollow cone nozzle nominally rated at 6 0 gal hr at 100 Ib in2 0 71 MPa delivers a proper spray pattern 1 Fuel Rail The fuel rail must be adjusted to position the fuel nozzle at a depth of 0 3125 inch 8 mm from the end plane of the exit stator which must be mounted in the end of the draft tube iii Internal Stator The internal stator located in the middle of the draft tube must be positioned at a depth of 3 75 inches 95 mm from the tip of the fuel nozzle The stator must also be posi tioned such that the integral igniters are located at an angle midway between the 10 and 11 o clock position when viewed looking into the draft tube Minor deviations to the igniter angle are accept able if the temperature and heat flux requirements conform to the requirements of paragraph VII e of this appendix iv Blower Fan The cylindrical blower fan used to pump air through the burner must measure 5 25 inches 133 mm in diameter by 3 5 inches 89 mm in width v Burner cone Install a 12 0 125 inch 305 3 mm burner extension cone at the end of the draft tube The cone must have an opening 6 0 125 inch 152 3 mm high and 11 0 125 inch 280 3 mm wide see figure 3 vi Fuel Use JP 8 Jet A or their international equivalent at a flow rate of 6 0 0 2 gal hr 0 378 0 0126 L min If this fuel is unavail
559. st specified in sec tion K25 2 1 e of this appendix a Service experience The world fleet for the airplane engine combination must accumulate a minimum of 250 000 engine hours The FAA may reduce this number of hours if the applicant iden tifies compensating factors that are acceptable to the FAA The compensating factors may include experience on another airplane but experience on the candidate airplane must make up a signifi cant portion of the total service experience b In flight shutdown IFSD rates The demon strated 12 month rolling average IFSD rate for the world fleet of the airplane engine combination must be commensurate with the level of ETOPS approval being sought 1 For type design approval up to and includ ing 120 minutes An IFSD rate of 0 05 or less per ASA Appendix K to Part 25 1 000 world fleet engine hours unless otherwise approved by the FAA Unless the IFSD rate is 0 02 or less per 1 000 world fleet engine hours the applicant must provide a list of corrective ac tions in the CMP document specified in section K25 1 6 of this appendix that when taken would result in an IFSD rate of 0 02 or less per 1 000 fleet engine hours 2 For type design approval up to and includ ing 180 minutes An IFSD rate of 0 02 or less per 1 000 world fleet engine hours unless otherwise approved by the FAA If the airplane engine com bination does not meet this rate by compliance with an existing 120 minute CMP
560. structural analysis and testing of the seats berths and their supporting structures may be determined by assuming that the critical load in the forward sideward downward upward and rearward directions as determined from the pre Scribed flight ground and emergency landing conditions acts separately or using selected combinations of loads if the required strength in each specified direction is substantiated The for ward load factor need not be applied to safety belts for berths 2 Each pilot seat must be designed for the re actions resulting from the application of the pilot forces prescribed in 825 395 3 The inertia forces specified in 825 561 must be multiplied by a factor of 1 33 instead of the fit ting factor prescribed in 825 625 in determining the strength of the attachment of each seat to the structure and each belt or harness to the seat or structure g Each seat at a flight deck station must have a restraint system consisting of a combined safety belt and shoulder harness with a single point re lease that permits the flight deck occupant when seated with the restraint system fastened to per form all of the occupant s necessary flight deck functions There must be a means to secure each combined restraint system when not in use to pre vent interference with the operation of the air plane and with rapid egress in an emergency h Each seat located in the passenger com partment and designated for use durin
561. structural fasteners such as identification discard recommendations and torque values g A list of special tools needed 215 Appendix to Part 25 H25 4 AIRWORTHINESS LIMITATIONS SECTION a The Instructions for Continued Airworthi ness must contain a section titled Airworthiness Limitations that is segregated and clearly distin guishable from the rest of the document This sec tion must set forth 1 Each mandatory replacement time struc tural inspection interval and related structural in spection procedures approved under 25 571 2 Each mandatory replacement time inspec tion interval related inspection procedure and all critical design configuration control limitations ap proved under 25 981 for the fuel tank system 3 Any mandatory replacement time of EWIS components as defined in 25 1701 b If the Instructions for Continued Airworthi ness consist of multiple documents the section required by this paragraph must be included in the principal manual This section must contain a leg ible statement in a prominent location that reads The Airworthiness Limitations section is FAA ap proved and specifies maintenance required under 43 16 and 91 403 of the Federal Aviation Regu lations unless an alternative program has been FAA approved H25 5 ELECTRICAL WIRING INTERCONNECTION SYSTEM EWIS INSTRUCTIONS FOR CONTINUED AIRWORTHINESS a The applicant must prepare Instructions for C
562. such door is in the open configuration for takeoff and landing c Each door between any passenger seat and any exit must have dual means to retain it in the open position each of which is capable of react ing the inertia loads specified in 825 561 d Doors installed across a longitudinal aisle must translate laterally to open and close e g pocket doors e Each door between any passenger seat and any exit must be frangible in either direction f Each door between any passenger seat and any exit must be operable from either side and if a locking mechanism is installed it must be capa ble of being unlocked from either side without the use of special tools 11 Width of Aisle Compliance is required with 825 815 except that aisle width may be re duced to 0 inches between passenger seats dur ing in flight operations only provided that the ap plicant demonstrates that all areas of the cabin are easily accessible by a crew member in the event of an emergency e g in flight fire decom pression Additionally instructions must be pro vided at each passenger seat for restoring the aisle width required by 825 815 Procedures must be established and documented in the AFM to en sure that the required aisle widths are provided during taxi takeoff and landing 12 Materials for Compartment Interiors Compliance is required with the applicable provi sions of 825 853 except that compliance with Ap pendix F parts IV and V
563. sumed longitudinal axis of the airplane to the weight of the airplane A positive load factor is 35 825 331 one in which the aerodynamic force acts upward with respect to the airplane b Considering compressibility effects at each speed compliance with the flight load require ments of this subpart must be shown 1 At each critical altitude within the range of altitudes selected by the applicant 2 At each weight from the design minimum weight to the design maximum weight appropriate to each particular flight load condition and 3 For each required altitude and weight for any practicable distribution of disposable load within the operating limitations recorded in the Airplane Flight Manual c Enough points on and within the boundaries of the design envelope must be investigated to ensure that the maximum load for each part of the airplane structure is obtained d The significant forces acting on the airplane must be placed in equilibrium in a rational or con servative manner The linear inertia forces must be considered in equilibrium with the thrust and all aerodynamic loads while the angular pitch ing inertia forces must be considered in equilib rium with thrust and all aerodynamic moments in cluding moments due to loads on components such as tail surfaces and nacelles Critical thrust values in the range from zero to maximum contin uous thrust must be considered Docket No 5066 29 FR 18291 Dec 24
564. t 25 40 42 FR 15043 March 17 1977 Amdt 25 57 49 FR 6849 Feb 23 1984 Amdt 25 72 55 FR 29785 July 20 1990 825 1101 Carburetor air preheater design Each carburetor air preheater must be de signed and constructed to a Ensure ventilation of the preheater when the engine is operated in cold air b Allow inspection of the exhaust manifold parts that it surrounds and c Allow inspection of critical parts of the pre heater itself 825 1103 Induction system ducts and air duct systems a Each induction system duct upstream of the first stage of the engine supercharger and of the auxiliary power unit compressor must have a drain to prevent the hazardous accumulation of fuel and moisture in the ground attitude No drain may discharge where it might cause a fire hazard b Each induction system duct must be 107 825 1105 1 Strong enough to prevent induction system failures resulting from normal backfire conditions and 2 Fire resistant if it is in any fire zone for which a fire extinguishing system is required ex cept that ducts for auxiliary power units must be fireproof within the auxiliary power unit fire zone c Each duct connected to components be tween which relative motion could exist must have means for flexibility d For turbine engine and auxiliary power unit bleed air duct systems no hazard may result if a duct failure occurs at any point between the air duct source and the
565. t with the airplane in the normal flying position the red light is on the left side and the green light is on the right side Each light must be approved c Rear position light The rear position light must be a white light mounted as far aft as practi cable on the tail or on each wing tip and must be approved d Light covers and color filters Each light cover or color filter must be at least flame resistant and may not change color or shape or lose any ap preciable light transmission during normal use Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55468 Dec 20 1976 825 1387 Position light system dihedral angles a Except as provided in paragraph e of this section each forward and rear position light must as installed show unbroken light within the dihe dral angles described in this section b Dihedral angle L left is formed by two in tersecting vertical planes the first parallel to the longitudinal axis of the airplane and the other at 110 degrees to the left of the first as viewed when looking forward along the longitudinal axis c Dihedral angle R right is formed by two in tersecting vertical planes the first parallel to the longitudinal axis of the airplane and the other at 110 degrees to the right of the first as viewed when looking forward along the longitudinal axis d Dihedral angle A aft is formed by two inter secting vertical planes making angles o
566. t av erage and warm day flammability exposure for the fuel tank under evaluation meets the applicable flammability limits defined in Table 5 of this appendix TABLE 5 FLAMMABILITY EXPOSURE LIMIT Minimum number of flights in Monte Carlo analysis Maximum acceptable Monte Carlo average fuel tank flammability exposure percent to meet 3 percent requirements Maximum acceptable Monte Carlo average fuel tank flammability exposure percent to meet 7 percent Part 26 requirements 10 000 2 91 6 79 100 000 2 98 6 96 1 000 000 3 00 7 00 236 ASA Send additional terms to asa asa2fly com A Accelerate stop distance 25 109 17 Accessory gearboxes 25 1167 After cooler 25 1107 Air duct systems 25 1103 Air induction 25 1091 Aircraft structure control surface and system loads 9825 391 25 459 43 45 emergency landing conditions 25 561 25 563 fatigue evaluation 25 571 flight maneuver and gust conditions 25 331 25 351 5 ssi ae 36 41 general load factors limits 25 301 25 307 ground loads 25 471 25 519 lightning protection 25 587 water loads 25 521 25 537 Airplane flight manual 25 1581 25 1587 139 141 Airspeed limitations flap extended speed 25 1511 general 25 1503 landing gear 25 1515 gs maneuvering speed 25 1507 maximum operating 25 1505 minimum control speed 25 1513
567. t be incorporated that will deter con cealment or promote discovery of weapons ex plosives or other objects from a simple inspection in the following areas of the airplane cabin i Areas above the overhead bins must be de signed to prevent objects from being hidden from view in a simple search from the aisle Designs that prevent concealment of objects with volumes 20 cubic inches and greater satisfy this require ment ii Toilets must be designed to prevent the pas sage of solid objects greater than 2 0 inches in di ameter 76 Federal Aviation Regulations iii Life preservers or their storage locations must be designed so that tampering is evident d Exceptions Airplanes used solely to trans port cargo only need to meet the requirements of paragraphs b 1 b 3 and c 2 of this sec tion e Material Incorporated by Reference You must use National Institute of Justice NIJ Stan dard 0101 04 Ballistic Resistance of Personal Body Armor June 2001 Revision A to establish ballistic resistance as required by paragraph 3 of this section 1 The Director of the Federal Register ap proved the incorporation by reference of this doc ument under 5 U S C 552 a and 1 CFR part 51 2 You may review copies of NIJ Standard 0101 04 at the i FAA Transport Airplane Directorate 1601 Lind Avenue SW Renton Washington 98055 ii National Institute of Justice NIJ http www ojp usdoj gov nij tele
568. t for abnormal fuel management or transfer between tanks and pos sible loss of fuel This paragraph does not apply to airplanes with a required flight engineer b APU design If an APU is needed to comply with this appendix the applicant must demon strate that 1 The reliability of the APU is adequate to meet those requirements and 2 If it is necessary that the APU be able to start in flight it is able to start at any altitude up to the maximum operating altitude of the airplane or 45 000 feet whichever is lower and run for the re mainder of any flight c Engine oil tank design The engine oil tank filler cap must comply with 33 71 c 4 of this chapter K25 1 5 Engine condition monitoring Procedures for engine condition monitoring must be specified and validated in accordance with Part 33 Appendix A paragraph A33 3 c of this chapter K25 1 6 Configuration maintenance and pro cedures The applicant must list any configuration oper ating and maintenance requirements hardware life limits MMEL constraints and ETOPS ap proval in a CMP document ASA Part 25 Airworthiness Standards Transport Category K25 1 7 Airplane flight manual The airplane flight manual must contain the fol lowing information applicable to the ETOPS type design approval a Special limitations including any limitation associated with operation of the airplane up to the maximum diversion time being approved b Required
569. t hole zero position must be 7 1 2 1 8 inches 191 3 mm The distance be tween the centerline of the first hole to the center line of the second hole must be 2 inches 51 mm It must also be the same distance from the center line of the second hole to the centerline of the third hole See figure 7 A calorimeter holding frame that differs in construction is acceptable as long as the height from the centerline of the first hole to the radiant panel and the distance between holes is the same as described in this paragraph ASA Part 25 Airworthiness Standards Transport Category Appendix F to Part 25 FIGURE 7 Calorimeter Holding Frame 4 Back of Chamber Position O Position 1 Position 2 1 1 16 in 27 mm eme 13 1 4 in 337 mm Front of Chamber via connections K to propane supply _ 2 in 51 mm B Y 2in 51 mm 9 Instrumentation Provide a calibrated re cording device with an appropriate range or a computerized data acquisition system to measure and record the outputs of the calorimeter and the thermocouple The data acquisition system must be capable of recording the calorimeter output ev ery second during calibration 10 Timing device Provide a stopwatch or other device accurate to 1 second hour to mea sure the time of application of the pilot burner flame c Test specimens 1 Specimen preparation Prepare a
570. t prescribed in subparagraph 5 of this paragraph The ex posed area of the specimen must be at least 2 inches wide and 12 inches long unless the actual size used in the airplane is smaller The edge to which the burner flame is applied must not consist of the finished or protected edge of the specimen but must be representative of the actual cross section of the material or part as installed in the airplane The specimen must be mounted in a metal frame so that all four edges are held se curely and the exposed area of the specimen is at least 8 inches by 8 inches during the 45 test pre Scribed in subparagraph 6 of this paragraph 3 Apparatus Except as provided in subpara graph 7 of this paragraph tests must be con ducted in a draft free cabinet in accordance with Federal Test Method Standard 191 Model 5903 revised Method 5902 for the vertical test or Method 5906 for horizontal test available from the General Services Administration Business Service Center Region 3 Seventh amp D Streets SW Washington DC 20407 Specimens which are too large for the cabinet must be tested in sim ilar draft free conditions ASA Appendix F to Part 25 4 Vertical test minimum of three specimens must be tested and results averaged For fabrics the direction of weave corresponding to the most critical flammability conditions must be parallel to the longest dimension Each specimen must be supported vertically The specimen m
571. talls the action of the air plane after the stall may not be so violent or ex treme as to make it difficult with normal piloting Skill to effect a prompt recovery and to regain control of the airplane The maximum bank angle that occurs during the recovery may not exceed 1 Approximately 60 degrees in the original di rection of the turn or 30 degrees in the opposite direction for deceleration rates up to 1 knot per second and 2 Approximately 90 degrees in the original di rection of the turn or 60 degrees in the opposite direction for deceleration rates in excess of 1 knot per second Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 84 60 FR 30750 June 9 1995 825 207 Stall warning a Stall warning with sufficient margin to pre vent inadvertent stalling with the flaps and landing gear in any normal position must be clear and dis tinctive to the pilot in straight and turning flight b The warning must be furnished either through the inherent aerodynamic qualities of the airplane or by a device that will give clearly distin guishable indications under expected conditions 30 Federal Aviation Regulations of flight However a visual stall warning device that requires the attention of the crew within the Cockpit is not acceptable by itself If a warning de vice is used it must provide a warning in each of the airplane configurations prescribed in para graph a of this section at the s
572. tempera ture of 41 degrees F above standard 3 The combination of vertical and drag com ponents considered to be acting at the main wheel axle centerline b For the tail down landing condition for air planes with tail wheels the main and tail wheels are assumed to contact the ground simulta neously in accordance with figure 3 of Appendix A Ground reaction conditions on the tail wheel are assumed to act 1 Vertically and 2 Up and aft through the axle at 45 degrees to the ground line c For the tail down landing condition for air planes with nose wheels the airplane is assumed to be at an attitude corresponding to either the stalling angle or the maximum angle allowing clearance with the ground by each part of the air plane other than the main wheels in accordance with figure 3 of Appendix A whichever is less Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 91 62 FR 40705 July 29 1997 Amdt 25 94 63 FR 8848 Feb 23 1998 825 483 One gear landing conditions For the one gear landing conditions the air plane is assumed to be in the level attitude and to contact the ground on one main landing gear in accordance with Figure 4 of Appendix A of this part In this attitude a The ground reactions must be the same as those obtained on that side under 825 479 d 1 and b Each unbalanced external load must be re acted by airplane inertia in a rational or conserva tive manner
573. ternal environment 4 Identifying all the electrical and electronic Systems that are subject to the requirements of this section and their locations on or within the airplane 5 Establishing the susceptibility of the sys tems to the internal and external lightning envi ronment 6 Designing protection and 7 Verifying that the protection is adequate Docket No 25912 59 FR 22116 April 28 1994 117 825 1317 825 1317 High Intensity Radiated Fields HIRF Protection a Except as provided in paragraph d of this section each electrical and electronic system that performs a function whose failure would prevent the continued safe flight and landing of the air plane must be designed and installed so that 1 The function is not adversely affected dur ing and after the time the airplane is exposed to environment as described in appendix L to this part 2 The system automatically recovers normal operation of that function in a timely manner af ter the airplane is exposed to HIRF environment as described in appendix L to this part unless the System s recovery conflicts with other operational or functional requirements of the system and 3 The system is not adversely affected during and after the time the airplane is exposed to HIRF environment as described in appendix L to this part b Each electrical and electronic system that performs a function whose failure would signifi cantly
574. the device during crash impact 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 location under water and 3 Have an underwater locating device when required by the operating rules of this chapter on or adjacent to the container which is secured in such manner that they are not likely to be sepa rated during crash impact Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 2 30 FR 3932 March 26 1965 Amdt 25 16 32 FR 13914 Oct 6 1967 Amdt 25 41 42 FR 36971 July 18 1977 Amdt 25 65 53 FR 26143 July 11 1988 Amdt 25 124 73 FR 12563 March 7 2008 Amdt 25 124 74 FR 32800 July 9 2009 ASA Part 25 Airworthiness Standards Transport Category 825 1459 Flight data recorders a Each flight recorder required by the operat ing rules of this chapter must be installed so that 1 It is supplied with airspeed altitude and di rectional data obtained from sources that meet the accuracy requirements of 25 1323 25 1325 and 25 1327 as appropriate 2 The vertical acceleration sensor is rigidly at tached and located longitudinally either within the approved center of gravity limits of the airplane or at a distance forward or aft of these limits that does not exceed 25 percent of the airplane s mean aerodynamic chord 3 i It receives its electrical power from the b
575. the propeller tips must be strong and stiff enough to withstand the effects of the induced vibration and of ice thrown from the propeller b No window may be near the propeller tips unless it can withstand the most severe ice impact likely to occur 25 899 Electrical bonding and protection against static electricity a Electrical bonding and protection against static electricity must be designed to minimize ac cumulation of electrostatic charge that would cause 1 Human injury from electrical shock 2 Ignition of flammable vapors or 3 Interference with installed electrical elec tronic equipment b Compliance with paragraph a of this sec tion may be shown by 1 Bonding the components properly to the air frame or 2 Incorporating other acceptable means to dissipate the static charge so as not to endanger the airplane personnel or operation of the in stalled electrical electronic systems Docket No FAA 2004 18379 72 FR 63405 Nov 8 2007 93 825 901 Subpart E Powerplant GENERAL 525 901 Installation a For the purpose of this part the airplane powerplant installation includes each component that 1 Is necessary for propulsion 2 Affects the control of the major propulsive units or 3 Affects the safety of the major propulsive units between normal inspections or overhauls b For each powerplant 1 The installation must comply with i The installation
576. the recovery maneuver in the same way as for the airplane in non icing conditions Compliance with this requirement must be demonstrated in flight with the speed reduced at rates not exceeding one knot per second with 1 The more critical of the takeoff ice and final takeoff ice accretions defined in appendix C for each configuration used in the takeoff phase of flight 2 The en route ice accretion defined in appen dix C for the en route configuration 3 The holding ice accretion defined in appen dix C for the holding configuration s 4 The approach ice accretion defined in ap pendix C for the approach configuration s and 5 The landing ice accretion defined in appen dix C for the landing and go around configura tion s f The stall warning margin must be sufficient in both non icing and icing conditions to allow the pilot to prevent stalling when the pilot starts a re covery maneuver not less than one second after the onset of stall warning in slow down turns with at least 1 5 g load factor normal to the flight path ASA Part 25 Airworthiness Standards Transport Category and airspeed deceleration rates of at least 2 knots per second When demonstrating compliance with this paragraph for icing conditions the pilot must perform the recovery maneuver in the same way as for the airplane in non icing conditions Compliance with this requirement must be dem onstrated in flight with 1 The flaps and landing
577. the run way surface condition b The net takeoff flight path data must be de termined so that they represent the actual takeoff flight paths determined in accordance with 825 111 and with paragraph a of this section re duced at each point by a gradient of climb equal to 1 0 8 percent for two engine airplanes 2 0 9 percent for three engine airplanes and 3 1 0 percent for four engine airplanes c The prescribed reduction in climb gradient may be applied as an equivalent reduction in ac celeration along that part of the takeoff flight path at which the airplane is accelerated in level flight Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 92 63 FR 8320 Feb 18 1998 825 117 Climb general Compliance with the requirements of 25 119 and 25 121 must be shown at each weight alti tude and ambient temperature within the opera tional limits established for the airplane and with the most unfavorable center of gravity for each configuration 825 119 Landing climb All engines operating In the landing configuration the steady gradi ent of climb may not be less than 3 2 percent with the engines at the power or thrust that is available 8 seconds after initiation of movement of the power or thrust controls from the minimum flight idle to the go around power or thrust setting a In non icing conditions with a climb speed of Vggr determined accordance with 25 125 b 2 i
578. the tire type expected in service have a clearance to sur rounding structure and systems that is adequate to prevent unintended contact between the tire and any part of the structure or systems e For an airplane with a maximum certificated takeoff weight of more than 75 000 pounds tires mounted on braked wheels must be inflated with dry nitrogen or other gases shown to be inert so that the gas mixture in the tire does not contain oxygen in excess of 5 percent by volume unless it can be shown that the tire liner material will not produce a volatile gas when heated or that means are provided to prevent tire temperatures from reaching unsafe levels Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 48 44 FR 68752 Nov 29 1979 Amdt 25 72 55 FR 29777 July 20 1990 Amdt 25 78 58 FR 11781 Feb 26 1993 ASA Part 25 Airworthiness Standards Transport Category 825 735 Brakes and braking systems a Approval Each assembly consisting of a wheel s and brake s must be approved b Brake system capability The brake system associated systems and components must be de signed and constructed so that 1 If any electrical pneumatic hydraulic or mechanical connecting or transmitting element fails or if any single source of hydraulic or other brake operating energy supply is lost it is possi ble to bring the airplane to rest with a braked roll stopping distance of not more than two times that obta
579. tificate except that a limitation need not be es tablished for a parameter that cannot be exceeded during normal operation due to the design of the installation or to another established limitation d Ambient temperature An ambient tempera ture limitation including limitations for winterization installations if applicable must be established as the maximum ambient atmospheric temperature established in accordance with 25 1043 b Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29786 July 20 1990 825 1522 Auxiliary power unit limitations If an auxiliary power unit is installed in the air plane limitations established for the auxiliary power unit including categories of operation must be specified as operating limitations for the airplane Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29786 July 20 1990 138 Federal Aviation Regulations 825 1523 Minimum flight crew The minimum flight crew must be established so that it is sufficient for safe operation consider ing a The workload on individual crewmembers b The accessibility and ease of operation of necessary controls by the appropriate crewmem ber and c The kind of operation authorized under 825 1525 The criteria used in making the deter minations required by this section are set forth in Appendix D Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 3 30
580. tion 122 Federal Aviation Regulations 1 If necessary for the maintenance of proper fuel delivery pressure there must be a connection to transmit the carburetor air intake static pres sure to the proper pump relief valve connection and 2 If a connection is required under paragraph f 1 of this section the gauge balance lines must be independently connected to the carburetor in let pressure to avoid erroneous readings Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 40 42 FR 15044 March 17 1977 ELECTRICAL SYSTEMS AND EQUIPMENT 825 1351 General a Electrical system capacity The required generating capacity and number and kinds of power sources must 1 Be determined by an electrical load analy Sis and 2 Meet the requirements of 825 1309 b Generating system The generating system includes electrical power sources main power busses transmission cables and associated con trol regulation and protective devices It must be designed so that 1 Power sources function properly when inde pendent and when connected in combination 2 No failure or malfunction of any power Source can create a hazard or impair the ability of remaining sources to supply essential loads 3 The system voltage and frequency as ap plicable at the terminals of all essential load equipment can be maintained within the limits for which the equipment is designed during any probable operati
581. tion 1 Position the burner in front of the calorime ter so that it is centered and the vertical plane of the burner cone exit is 4 0 125 inches 102 3 mm from the calorimeter face Ensure that the horizontal centerline of the burner cone is offset 1 inch below the horizontal centerline of the calo rimeter figure 8 Without disturbing the calorime ter position rotate the burner in front of the ther mocouple rake such that the middle thermocou ple number 4 of 7 is centered on the burner cone SEE FIGURE 8 AT THE END OF PART VII OF THIS APPENDIX Ensure that the horizontal centerline of the burner cone is also offset 1 inch below the hori zontal centerline of the thermocouple tips Re check measurements by rotating the burner to each position to ensure proper alignment be tween the cone and the calorimeter and thermo couple rake Note The test burner mounting sys tem must incorporate detents that ensure proper centering of the burner cone with respect to both the calorimeter and the thermocouple rakes so that rapid positioning of the burner can be achieved during the calibration procedure 2 Position the air velocity meter in the adapter or airbox making certain that no gaps exist where air could leak around the air velocity measuring device Turn on the blower motor while ensuring that the fuel solenoid and igniters are off Adjust the air intake velocity to a level of 2150 ft min 10 92 m s then turn off the
582. tion takes place This must also be shown with any single failure and malfunction except that i With failures or malfunctions in the latching mechanism it need not latch after closing and ii With jamming as a result of mechanical fail ure or blocking debris the door need not close and latch if it can be shown that the pressurization 72 Federal Aviation Regulations loads on the jammed door or mechanism would not result in an unsafe condition d Latching and locking The latching and locking mechanisms must be designed as follows 1 There must be a provision to latch each door 2 The latches and their operating mechanism must be designed so that under all airplane flight and ground loading conditions with the door latched there is no force or torque tending to un latch the latches In addition the latching system must include a means to secure the latches in the latched position This means must be indepen dent of the locking system 3 Each door subject to pressurization and for which the initial opening movement is not inward must i Have an individual lock for each latch ii Have the lock located as close as practica ble to the latch and iii Be designed so that during pressurized flight no single failure in the locking system would prevent the locks from restraining the latches nec essary to secure the door 4 Each door for which the initial opening movement is inward and unlatching
583. tional landing stop at maximum landing weight The design landing stop brake ki netic energy absorption requirement of each wheel brake and tire assembly must be deter mined It must be substantiated by dynamometer testing that the wheel brake and tire assembly is capable of absorbing not less than this level of ki netic energy throughout the defined wear range of the brake The energy absorption rate derived from the airplane manufacturer s braking require ments must be achieved The mean deceleration must not be less than 10 fps ASA 825 735 2 Maximum kinetic energy accelerate stop The maximum kinetic energy accelerate stop is a rejected takeoff for the most critical combination of airplane takeoff weight and speed The acceler ate stop brake kinetic energy absorption require ment of each wheel brake and tire assembly must be determined It must be substantiated by dynamometer testing that the wheel brake and tire assembly is capable of absorbing not less than this level of kinetic energy throughout the de fined wear range of the brake The energy ab sorption rate derived from the airplane manufac turers braking requirements must be achieved The mean deceleration must not be less than 6 fps 3 Most severe landing stop The most severe landing stop is a stop at the most critical combina tion of airplane landing weight and speed The most severe landing stop brake kinetic energy ab sorption requirement of each
584. to 20 80 20 to 30 40 30 to 75 20 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 27 36 FR 12972 July 10 1971 Amdt 25 41 42 FR 36970 July 18 1977 825 1403 Wing icing detection lights Unless operations at night in known or forecast icing conditions are prohibited by an operating limitation a means must be provided for illuminat ing or otherwise determining the formation of ice on the parts of the wings that are critical from the standpoint of ice accumulation Any illumination that is used must be of a type that will not cause glare or reflection that would handicap crewmem bers in the performance of their duties Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55468 Dec 20 1976 SAFETY EQUIPMENT 825 1411 General a Accessibility Required safety equipment to be used by the crew in an emergency must be readily accessible b Stowage provisions Stowage provisions for required emergency equipment must be fur nished and must 1 Be arranged so that the equipment is di rectly accessible and its location is obvious and 2 Protect the safety equipment from inadvert ent damage c Emergency exit descent device The stow age provisions for the emergency exit descent de vices required by 25 810 a must be at each exit for which they are intended d Liferafts 1 The stowage provisions for the liferafts de scribed in 825 14
585. to Part 25 Flight length NM Airplane maximum range nautical miles NM From To 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000 Distribution of flight lengths percentage of total 0 200 11 7 7 5 62 5 5 4 7 4 0 34 3 0 2 6 2 3 200 400 27 3 19 9 17 0 15 2 13 2 114 97 8 5 7 5 6 7 400 600 46 3 40 0 35 7 32 6 28 5 24 9 21 2 18 7 16 4 14 8 600 800 10 3 11 6 11 0 10 2 9 1 8 0 6 9 6 1 54 4 8 800 1000 44 8 5 8 6 82 7 4 6 6 5 7 5 0 4 5 4 0 1000 1200 0 0 4 8 5 3 5 3 4 8 4 3 3 8 3 3 3 0 27 1200 1400 0 0 3 6 44 4 5 4 2 3 8 3 3 3 0 27 2 4 1400 1600 0 0 2 2 3 3 3 5 3 3 3 1 2 7 2 4 2 2 2 0 1600 1800 0 0 1 2 2 3 2 6 2 5 24 21 1 9 diz 1 6 1800 2000 0 0 0 7 22 2 6 2 6 2 5 2 2 2 0 1 8 1 7 2000 2200 0 0 0 0 1 6 2 1 2 2 2 1 1 9 1 7 1 6 1 4 2200 2400 0 0 0 0 1 1 1 6 1 7 4 7 1 6 14 1 3 12 2400 2600 0 0 0 0 0 7 1 2 1 4 1 4 1 3 12 1 1 1 0 2600 2800 0 0 0 0 0 4 0 9 1 0 1 1 1 0 0 9 0 9 0 8 2800 3000 0 0 0 0 0 2 0 6 0 7 0 8 0 7 0 7 0 6 0 6 3000 3200 0 0 0 0 0 0 0 6 0 8 0 8 0 8 0 8 0 7 0 7 3200 3400 0 0 0 0 0 0 0 7 ti 1 2 1 2 1 1 1 1 1 0 3400 3600 0 0 0 0 0 0 0 7 1 3 1 6 1 6 1 5 1 5 1 4 3600 3800 0 0 0 0 0 0 0 9 22 2 7 2 8 27 2 6 2 5 3800 4000 0 0 0 0 0 0 0 5 2 0 2 6 2 8 2 8 2 7 2 6 4000 4200 0 0 0 0 0 0 0 0 2 1 3 0 3 2 3 3 3 2 31 4200 4400 0 0 0 0 0 0 0 0 14 2 2 2 5 2 6 2 6 25
586. to provide uniform heat flux density over the area occupied by the vertical sample 4 Air Distribution System The air entering the environmental chamber must be distributed by a 25 inch 6 3 mm thick aluminum plate having eight No 4 drill holes located 2 inches 51 mm from sides on 4 inch 102 mm centers mounted at the base of the environmental chamber A sec ond plate of 18 gauge stainless steel having 120 evenly spaced No 28 drill holes must be mounted 6 inches 152 mm above the aluminum plate A well regulated air supply is required The air supply manifold at the base of the pyramidal section must have 48 evenly spaced No 26 drill holes located 38 inch 10 mm from the inner edge of the manifold resulting in an airflow split of approximately three to one within the apparatus 5 Exhaust Stack An exhaust stack 5 25 x 2 75 inches 133 x 70 mm in cross section and 10 inches 254 mm long fabricated from 28 gauge stainless steel must be mounted on the outlet of the pyramidal section A 1 0 x 3 0 inch 25 x 76 mm baffle plate of 0 18 002 inch 50 05 mm stainless steel must be centered inside the stack perpendicular to the air flow 3 inches 76 mm above the base of the stack 6 Specimen Holders i The specimen must be tested in a vertical orientation The specimen holder Figure 3 of this part IV must incorporate a frame that touches the specimen which is wrapped with aluminum foil as required by para
587. to the next higher flow The thermopile baseline voltage must be mea sured The gas flow to the burner must be in creased to the higher preset flow and allowed to burn for 2 0 minutes and the thermopile voltage must be measured The sequence must be re peated until all five values have been determined The average of the five values must be used as the calibration factor The procedure must be re peated if the percent relative standard deviation is greater than 5 percent Calculations are shown in paragraph f of this part IV 2 Flux Uniformity Uniformity of flux over the specimen must be checked periodically and after each heating element change to determine if it is within acceptable limits of plus or minus 5 percent 3 As noted in paragraph b 2 of this part IV thermopile hot junctions must be cleared of soot deposits as needed to maintain the calibrated sensitivity d Preparation of Test Specimens 1 The test specimens must be representative of the aircraft component in regard to materials and construction methods The standard size for ASA Part 25 Airworthiness Standards Transport Category the test specimens is 5 91 03 x 5 91 03 inches 149 1 x 149 1 mm The thickness of the specimen must be the same as that of the air craft component it represents up to a maximum thickness of 1 75 inches 45 mm Test specimens representing thicker components must be 1 75 inches 45 mm 2 Conditioning Specimens
588. tor a Each magnetic direction indicator must be installed so that its accuracy is not excessively af fected by the airplane s vibration or magnetic fields b The compensated installation may not have a deviation in level flight greater than 10 degrees on any heading 825 1329 Flight guidance system a Quick disengagement controls for the auto pilot and autothrust functions must be provided for each pilot The autopilot quick disengagement controls must be located on both control wheels or equivalent The autothrust quick disengage ment controls must be located on the thrust con trol levers Quick disengagement controls must be readily accessible to each pilot while operating the control wheel or equivalent and thrust con trol levers b The effects of a failure of the system to dis engage the autopilot or autothrust functions when manually commanded by the pilot must be as sessed in accordance with the requirements of 825 1309 c Engagement or switching of the flight guid ance system a mode or a sensor may not cause a transient response of the airplane s flight path any greater than a minor transient as defined in paragraph n 1 of this section d Under normal conditions the disengage ment of any automatic control function of a flight guidance system may not cause a transient re sponse of the airplane s flight path any greater than a minor transient e Under rare normal and non normal condi tio
589. true ambient atmo spheric static pressure is not changed when the airplane is exposed to the continuous and inter mittent maximum icing conditions defined in Ap pendix C of this part c The design and installation of the static pressure system must be such that 1 Positive drainage of moisture is provided chafing of the tubing and excessive distortion or restriction at bends in the tubing is avoided and the materials used are durable suitable for the purpose intended and protected against corro sion and 2 It is airtight except for the port into the atmo sphere A proof test must be conducted to dem onstrate the integrity of the static pressure system in the following manner i Unpressurized airplanes Evacuate the static pressure system to a pressure differential of approximately 1 inch of mercury or to a reading on the altimeter 1 000 feet above the airplane el evation at the time of the test Without additional pumping for a period of 1 minute the loss of indi cated altitude must not exceed 100 feet on the al timeter ii Pressurized airplanes Evacuate the static pressure system until a pressure differential equivalent to the maximum cabin pressure differ ential for which the airplane is type certificated is achieved Without additional pumping for a period of 1 minute the loss of indicated altitude must not exceed 2 percent of the equivalent altitude of the maximum cabin differential pressure or 100 feet
590. ts provided for meeting this requirement must be independent of the normal turbo super charger controls POWERPLANT CONTROLS AND ACCESSORIES 825 1141 Powerplant controls general Each powerplant control must be located ar ranged and designed under 25 777 through 25 781 and marked under 825 1555 In addition it must meet the following requirements a Each control must be located so that it can not be inadvertently operated by persons enter ing leaving or moving normally in the cockpit b Each flexible control must be approved or must be shown to be suitable for the particular ap plication ASA 825 1143 c Each control must have sufficient strength and rigidity to withstand operating loads without failure and without excessive deflection d Each control must be able to maintain any Set position without constant attention by flight crewmembers and without creep due to control loads or vibration e The portion of each powerplant control lo cated in a designated fire zone that is required to be operated in the event of fire must be at least fire resistant f Powerplant valve controls located in the cockpit must have 1 For manual valves positive stops or in the case of fuel valves suitable index provisions in the open and closed position and 2 For power assisted valves a means to indi cate to the flight crew when the valve i Is in the fully open or fully closed position or ii Is m
591. ty control procedure is estab lished 3 For castings procured to a specification that guarantees the mechanical properties of the ma terial in the casting and provides for demonstra tion of these properties by test of coupons cut from the castings on a sampling basis i A casting factor of 1 0 may be used and ii The castings must be inspected as provided in paragraph d 1 of this section for casting fac tors of 1 25 through 1 50 and tested under para graph c 2 of this section 825 623 Bearing factors a Except as provided in paragraph b of this section each part that has clearance free fit 59 825 625 and that is subject to pounding or vibration must have a bearing factor large enough to provide for the effects of normal relative motion b No bearing factor need be used for a part for which any larger special factor is prescribed 825 625 Fitting factors For each fitting a part or terminal used to join one structural member to another the following apply a For each fitting whose strength is not proven by limit and ultimate load tests in which ac tual stress conditions are simulated in the fitting and surrounding structures a fitting factor of at least 1 15 must be applied to each part of 1 The fitting 2 The means of attachment and 3 The bearing on the joined members b No fitting factor need be used 1 For joints made under approved practices and based on com
592. ty tests at a minimum of 7 5 mA 3 From 40 MHz to 400 MHz use conducted susceptibility tests starting at a minimum of 7 5 mA at 40 MHz decreasing 20 dB per frequency decade to a minimum of 0 75 mA at 400 MHz 4 From 100 MHz to 8 GHz use radiated sus ceptibility tests at a minimum of 5 V m Docket No FAA 2006 23657 72 FR 44026 Aug 6 2007 229 Appendix M to Part 25 APPENDIX M TO PART 25 FUEL TANK SYSTEM FLAMMABILITY REDUCTION MEANS Source Docket No FAA 2005 22997 73 FR 42494 July 21 2008 unless otherwise noted M25 1 FUEL TANK FLAMMABILITY EXPOSURE REQUIREMENTS a The Fleet Average Flammability Exposure of each fuel tank as determined in accordance with Appendix N of this part may not exceed 3 percent of the Flammability Exposure Evaluation Time FEET as defined in Appendix N of this part As a portion of this 3 percent if flammability reduction means FRM are used each of the fol lowing time periods may not exceed 1 8 percent of the FEET 1 When any FRM is operational but the fuel tank is not inert and the tank is flammable and 2 When any FRM is inoperative and the tank is flammable b The Fleet Average Flammability Exposure as defined in Appendix N of this part of each fuel tank may not exceed 3 percent of the portion of the FEET occurring during either ground or take off climb phases of flight during warm days The analysis must consider the following conditions 1 The analysi
593. ued Airwor thiness as required by 25 1529 25 1729 and applicable provisions of parts 21 and 26 of this chapter b The Instructions for Continued Airworthi ness for each airplane must include the Instruc tions for Continued Airworthiness for each engine and propeller hereinafter designated products for each appliance required by this chapter and any required information relating to the interface of those appliances and products with the air plane If Instructions for Continued Airworthiness are not supplied by the manufacturer of an appli ance or product installed in the airplane the In structions for Continued Airworthiness for the air plane must include the information essential to the continued airworthiness of the airplane c The applicant must submit to the FAA a pro gram to show how changes to the Instructions for Continued Airworthiness made by the applicant or by the manufacturers or products and appliances installed in the airplane will be distributed H25 2 FORMAT a The Instructions for Continued Airworthiness must be in the form of a manual or manuals as ap propriate for the quantity of data to be provided b The format of the manual or manuals must provide for a practical arrangement H25 3 CONTENT The contents of the manual or manuals must be prepared in the English language The Instruc tions for Continued Airworthiness must contain the following manuals or sections as appropriate and
594. uid pe troleum gas 2 1 UN 1075 for the radiant panel fuel The panel fuel system must consist of a ven turi type aspirator for mixing gas and air at ap proximately atmospheric pressure Provide suit able instrumentation for monitoring and control ling the flow of fuel and air to the panel Include an air flow gauge an air flow regulator and a gas pressure gauge iii Radiant panel placement Mount the panel in the chamber at 30 to the horizontal specimen plane and 7 1 2 inches above the zero point of the specimen 3 Specimen holding system i The sliding platform serves as the housing for test specimen placement Brackets may be at tached via wing nuts to the top lip of the platform in order to accommodate various thicknesses of test specimens Place the test specimens on a sheet of Kaowool M board or 1260 Standard Board manufactured by Thermal Ceramics and available in Europe or equivalent either resting on the bottom lip of the sliding platform or on the base of the brackets It may be necessary to use multiple sheets of material based on the thickness of the test specimen to meet the sample height requirement Typically these non combustible sheets of material are available in 1 4 inch 6 mm thicknesses See figure 4 A sliding platform that is deeper than the 2 inch 50 8mm platform shown in figure 4 is also acceptable as long as the sample height requirement is met 193 Appendix to Part 25 Federa
595. uration power or thrust and speed must be made in accordance with the established procedures for service operation 4 The landing must be made without exces sive vertical acceleration tendency to bounce nose over ground loop porpoise or water loop 5 The landings may not require exceptional piloting skill or alertness c For landplanes and amphibians the landing distance on land must be determined on a level smooth dry hard surfaced runway In addition 1 The pressures on the wheel braking sys tems may not exceed those specified by the brake manufacturer 2 The brakes may not be used so as to cause excessive wear of brakes or tires and 3 Means other than wheel brakes may be used if that means i Is safe and reliable ii Is used so that consistent results can be ex pected in service and iii Is such that exceptional skill is not required to control the airplane d For seaplanes and amphibians the landing distance on water must be determined on smooth water e For skiplanes the landing distance on snow must be determined on smooth dry snow f The landing distance data must include cor rection factors for not more than 50 percent of the nominal wind components along the landing path opposite to the direction of landing and not less than 150 percent of the nominal wind components along the landing path in the direction of landing g If any device is used that depends on the operati
596. us that provides the maximum reliability for oper ation of the flight data recorder without jeopardiz ing service to essential or emergency loads ii It remains powered for as long as possible without jeopardizing emergency operation of the airplane 4 There is an aural or visual means for pre flight checking of the recorder for proper recording of data in the storage medium 5 Except for recorders powered solely by the engine driven electrical generator system there is an automatic means to simultaneously stop a recorder that has a data erasure feature and pre vent each erasure feature from functioning within 10 minutes after crash impact 6 There is a means to record data from which the time of each radio transmission either to or from ATC can be determined 7 Any single electrical failure external to the recorder does not disable both the cockpit voice recorder and the flight data recorder and 8 It is in a separate container from the cockpit voice recorder when both are required If used to comply with only the flight data recorder require ments a combination unit may be installed If a combination unit is installed as a cockpit voice re corder to comply with 25 1457 e 2 a combina tion unit must be used to comply with this flight data recorder requirement ASA 825 1459 b Each nonejectable record container must be located and mounted so as to minimize the probability of container rupture resulting fr
597. ust be configured to prevent disclosure of the active emergency exits to dem onstration participants in the airplane until the start of the demonstration p Exits used in the demonstration must con sist of one exit from each exit pair The demon stration may be conducted with the escape slides if provided inflated and the exits open at the be ginning of the demonstration In this case all exits must be configured such that the active exits are not disclosed to the occupants If this method is used the exit preparation time for each exit uti lized must be accounted for and exits that are not to be used in the demonstration must not be indi cated before the demonstration has started The exits to be used must be representative of all of the emergency exits on the airplane and must be designated by the applicant subject to approval by the Administrator At least one floor level exit must be used ASA Part 25 Airworthiness Standards Transport Category q Except as provided in paragraph c of this section all evacuees must leave the airplane by a means provided as part of the airplane s equip ment r The applicant s approved procedures must be fully utilized except the flightcrew must take no active role in assisting others inside the cabin dur ing the demonstration ASA Appendix J to Part 25 s The evacuation time period is completed when the last occupant has evacuated the air plane and is on the ground P
598. ust be ex posed to a Bunsen or Tirrill burner with a nominal inch I D tube adjusted to give a flame of 1 inches in height The minimum flame temperature measured by a calibrated thermocouple pyrome ter in the center of the flame must 1550 F The lower edge of the specimen must be inch above the top edge of the burner The flame must be applied to the center line of the lower edge of the specimen For materials covered by para graph a 1 i of part of this appendix the flame must be applied for 60 seconds and then re moved For materials covered by paragraph 1 1 of part of this appendix the flame must be applied for 12 seconds and then removed Flame time burn length and flaming time of drip pings if any may be recorded The burn length determined in accordance with subparagraph 7 of this paragraph must be measured to the near est tenth of an inch 5 Horizontal test A minimum of three speci mens must be tested and the results averaged Each specimen must be supported horizontally The exposed surface when installed in the air craft must be face down for the test The speci men must be exposed to a Bunsen or Tirrill burner with a nominal 4 I D tube adjusted to give a flame of 1 inches in height The minimum flame temperature measured by a calibrated thermo couple pyrometer in the center of the flame must be 1550 F The specimen must be positioned so that the edge being tested is centered
599. ust be protected from damage which could result in spillage of enough fuel to constitute a fire hazard as a result of a wheels up landing on a paved runway Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 57 49 FR 6848 Feb 23 1984 25 995 Fuel valves In addition to the requirements of 25 1189 for shutoff means each fuel valve must a Reserved b Be supported so that no loads resulting from their operation or from accelerated flight ASA Part 25 Airworthiness Standards Transport Category conditions are transmitted to the lines attached to the valve Docket No 5066 29 FR 18291 Dec 24 1964 amended by Amdt 25 40 42 FR 15043 17 1977 825 997 Fuel strainer or filter There must be a fuel strainer or filter between the fuel tank outlet and the inlet of either the fuel metering device or an engine driven positive dis placement pump whichever is nearer the fuel tank outlet This fuel strainer or filter must a Be accessible for draining and cleaning and must incorporate a screen or element which is easily removable b Have a sediment trap and drain except that it need not have a drain if the strainer or filter is easily removable for drain purposes c Be mounted so that its weight is not sup ported by the connecting lines or by the inlet or outlet connections of the strainer or filter itself un less adequate strength margins under all loading conditions are p
600. v 30 1972 Amdt 25 46 43 FR 50597 Oct 30 1978 Amdt 25 47 44 FR 61325 Oct 25 1979 Amdt 25 72 55 FR 29782 July 20 1990 Amdt 25 114 69 FR 24501 May 3 2004 Amdt 25 116 69 FR 62788 Oct 27 2004 825 810 Emergency egress assist means and escape routes a Each non over wing Type A Type B or Type C exit and any other non over wing landplane emergency exit more than 6 feet from the ground with the airplane on the ground and the landing gear extended must have an approved means to assist the occupants in descending to the ground 1 The assisting means for each passenger emergency exit must be a self supporting slide or equivalent and in the case of a Type A or Type B exits it must be capable of carrying simulta neously two parallel lines of evacuees In addition the assisting means must be designed to meet the following requirements i It must be automatically deployed and de ployment must begin during the interval between the time the exit opening means is actuated from inside the airplane and the time the exit is fully opened However each passenger emergency exit which is also a passenger entrance door or a service door must be provided with means to pre vent deployment of the assisting means when it is opened from either the inside or the outside under nonemergency conditions for normal use ii Except for assisting means installed at Type C exits it must be automatically erected within 6 seco
601. vail when drainage is needed and 2 Arranged so that no discharged fluid will cause an additional fire hazard 112 Federal Aviation Regulations b Each designated fire zone must be venti lated to prevent the accumulation of flammable vapors c No ventilation opening may be where it would allow the entry of flammable fluids vapors or flame from other zones d Each ventilation means must be arranged so that no discharged vapors will cause an addi tional fire hazard e Unless the extinguishing agent capacity and rate of discharge are based on maximum air flow through a zone there must be means to al low the crew to shut off sources of forced ventila tion to any fire zone except the engine power sec tion of the nacelle and the combustion heater ven tilating air ducts 825 1189 Shutoff means a Each engine installation and each fire zone specified in 25 1181 a 4 and 5 must have a means to shut off or otherwise prevent hazardous quantities of fuel oil deicer and other flammable fluids from flowing into within or through any designated fire zone except that shutoff means are not required for 1 Lines fittings and components forming an integral part of an engine and 2 systems for turbine engine installations in which all components of the system in a desig nated fire zone including oil tanks are fireproof or located in areas not subject to engine fire con ditions b The closing
602. vatively greater overall level of perfor mance than the one engine inoperative takeoff path established in accordance with paragraph a of this section The margin must be estab lished by the Administrator to insure safe day to day operations but in no case may it be less than 15 percent The all engines operating takeoff path must be determined by a procedure consistent with that established in complying with paragraph a of this section d For reciprocating engine powered air planes the takeoff path to be scheduled in the Airplane Flight Manual must represent the one engine operative takeoff path determined in ac cordance with paragraph a of this section and modified to reflect the procedure see paragraph 6 established by the applicant for flap retraction and attainment of the en route speed The sched uled takeoff path must have a positive slope at all points of the airborne portion and at no point must itlie above the takeoff path specified in paragraph a of this section 3 Takeoff distance The takeoff distance must be the horizontal distance along the one engine inoperative take off path determined in accor dance with paragraph 2 a from the start of the takeoff to the point where the airplane attains a height of 50 feet above the takeoff surface for re ASA Part 25 Airworthiness Standards Transport Category ciprocating engine powered airplanes and a height of 35 feet above the takeoff surface for tur
603. w or passengers and 3 The dissipation of the extinguishing agent in Class C compartments i During the above tests it must be shown that no inadvertent operation of smoke or fire de tectors in any compartment would occur as a re sult of fire contained in any other compartment either during or after extinguishment unless the 90 Federal Aviation Regulations extinguishing system floods each such compart ment simultaneously j Cargo or baggage compartment electrical wiring interconnection system components must meet the requirements of 825 1721 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29784 July 20 1990 Amdt 25 93 63 FR 8048 Feb 17 1998 Amdt 25 116 69 FR 62789 Oct 27 2004 Amdt 25 123 72 FR 63405 Nov 8 2007 25 856 Thermal Acoustic insulation materials a Thermal acoustic insulation material in stalled in the fuselage must meet the flame prop agation test requirements of part VI of Appendix F to this part or other approved equivalent test re quirements This requirement does not apply to small parts as defined in part of Appendix F of this part b For airplanes with a passenger capacity of 20 or greater thermal acoustic insulation materi als including the means of fastening the materi als to the fuselage installed in the lower half of the airplane fuselage must meet the flame pene tration resistance test requirements of part VII of Ap
604. wheel brake and tire assembly must be determined It must be sub stantiated by dynamometer testing that at the de clared fully worn limit s of the brake heat sink the wheel brake and tire assembly is capable of ab Sorbing not less than this level of kinetic energy The most severe landing stop need not be consid ered for extremely improbable failure conditions or if the maximum kinetic energy accelerate stop energy is more severe g In the landing case the minimum speed rat ing of each main wheel brake assembly that is the initial speed used in the dynamometer tests may not be more than the V used in the determi nation of kinetic energy in accordance with para graph f of this section assuming that the test procedures for wheel brake assemblies involve a specified rate of deceleration and therefore for the same amount of kinetic energy the rate of en ergy absorption the power absorbing ability of the brake varies inversely with the initial speed h Stored energy systems An indication to the flightcrew of the usable stored energy must be provided if a stored energy system is used to show compliance with paragraph b 1 of this section The available stored energy must be suf ficient for 1 At least 6 full applications of the brakes when an antiskid system is not operating and 2 Bringing the airplane to a complete stop when an antiskid system is operating under all runway surface conditions for which the
605. with the limit load requirements of this Part must be shown by tests in which 1 The direction of the test loads produces the most severe loading in the control system and 2 Each fitting pulley and bracket used in at taching the system to the main structure is in cluded b Compliance must be shown by analyses or individual load tests with the special factor re quirements for control system joints subject to an gular motion 825 683 Operation tests It must be shown by operation tests that when portions of the control system subject to pilot ef fort loads are loaded to 80 percent of the limit load specified for the system and the powered portions of the control system are loaded to the maximum load expected in normal operation the system is free from a Jamming b Excessive friction and c Excessive deflection Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5675 April 8 1970 825 685 Control system details a Each detail of each control system must be designed and installed to prevent jamming chaf ing and interference from cargo passengers loose objects or the freezing of moisture b There must be means in the cockpit to pre vent the entry of foreign objects into places where they would jam the system c There must be means to prevent the slap ping of cables or tubes against other parts d Sections 25 689 and 25 693 apply to cable Systems and joint
606. within 20 of plane of control Twist 133 in Ibs Push pull To be chosen by applicant Limited to flap tab stabilizer spoiler and landing gear operation controls 825 407 Trim tab effects The effects of trim tabs on the control surface design conditions must be accounted for only where the surface loads are limited by maximum pilot effort In these cases the tabs are consid ered to be deflected in the direction that would as sist the pilot and the deflections are a For elevator trim tabs those required to trim the airplane at any point within the positive por tion of the pertinent flight envelope in 25 333 b except as limited by the stops and b For aileron and rudder trim tabs those re quired to trim the airplane in the critical unsym metrical power and loading conditions with ap propriate allowance for rigging tolerances 44 Federal Aviation Regulations 825 409 Tabs a Trim tabs Trim tabs must be designed to withstand loads arising from all likely combina tions of tab setting primary control position and airplane speed obtainable without exceeding the flight load conditions prescribed for the airplane as a whole when the effect of the tab is opposed by pilot effort forces up to those specified in 825 397 b b Balancing tabs Balancing tabs must be de signed for deflections consistent with the primary control surface loading conditions c Servo tabs Servo tabs
607. within the approved operating limitations that is critical for the type of failure being considered 2 The controllability and maneuverability re quirements of this part are met within a practical 62 Federal Aviation Regulations operational flight envelope for example speed altitude normal acceleration and airplane config urations which is described in the Airplane Flight Manual and 3 The trim stability and stall characteristics are not impaired below a level needed to permit continued safe flight and landing Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5675 April 8 1970 825 675 Stops a Each control system must have stops that positively limit the range of motion of each movable aerodynamic surface controlled by the system b Each stop must be located so that wear slackness or take up adjustments will not ad versely affect the control characteristics of the air plane because of a change in the range of surface travel c Each stop must be able to withstand any loads corresponding to the design conditions for the control system Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 38 41 FR 55466 Dec 20 1976 825 677 Trim systems a Trim controls must be designed to prevent inadvertent or abrupt operation and to operate in the plane and with the sense of motion of the air plane b There must be means adjacent to the trim control to
608. wn e Two systems for radio navigation with con trols for each accessible from each pilot station designed and installed so that failure of one sys tem will not preclude operation of the other sys tem The use of a common antenna system is ac ceptable if adequate reliability is shown Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 23 35 FR 5678 April 8 1970 Amdt 25 46 43 FR 50598 Oct 30 1978 Amdt 25 54 45 FR 60173 Sept 11 1980 Amdt 25 72 55 FR 29785 July 20 1990 25 1309 Equipment systems and installations a The equipment systems and installations whose functioning is required by this subchapter must be designed to ensure that they perform their intended functions under any foreseeable operating condition b The airplane systems and associated com ponents considered separately and in relation to other systems must be designed so that 1 The occurrence of any failure condition which would prevent the continued safe flight and landing of the airplane is extremely improbable and 2 The occurrence of any other failure condi tions which would reduce the capability of the air plane or the ability of the crew to cope with ad verse operating conditions is improbable c Warning information must be provided to alert the crew to unsafe system operating condi tions and to enable them to take appropriate cor ASA Part 25 Airworthiness Standards Transport Cat
609. wn by analysis tests or both to be capable of continued safe flight and landing after any of the following failures or jamming in the flight control system and sur faces including trim lift drag and feel systems within the normal flight envelope without requir ing exceptional piloting skill or strength Probable malfunctions must have only minor effects on control system operation and must be capable of being readily counteracted by the pilot 61 825 672 1 Any single failure excluding jamming for example disconnection or failure of mechanical elements or structural failure of hydraulic compo nents such as actuators control spool housing and valves 2 Any combination of failures not shown to be extremely improbable excluding jamming for ex ample dual electrical or hydraulic system failures or any single failure in combination with any prob able hydraulic or electrical failure 3 Any jam in a control position normally en countered during takeoff climb cruise normal turns descent and landing unless the jam is shown to be extremely improbable or can be alle viated A runaway of a flight control to an adverse position and jam must be accounted for if such runaway and subsequent jamming is not ex tremely improbable d The airplane must be designed so that it is controllable if all engines fail Compliance with this requirement may be shown by analysis where that method has been shown to be reliab
610. y 2 The airplane must be equipped with a sys tem that automatically cycles the ice protection System or 3 An ice detection system must be provided to alert the flightcrew each time the ice protection System must be cycled h Procedures for operation of the ice protec tion system including activation and deactivation must be established and documented in the Air plane Flight Manual Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 72 55 FR 29785 July 20 1990 Amdt 25 121 72 FR 44669 Aug 8 2007 Amdt 25 129 74 FR 38339 Aug 3 2009 825 1421 Megaphones If a megaphone is installed a restraining means must be provided that is capable of restraining the megaphone when it is subjected to the ultimate in ertia forces specified in 25 561 b 3 Docket No 5066 29 FR 18291 Dec 24 1964 as amended by Amdt 25 41 42 FR 36970 July 18 1977 MISCELLANEOUS EQUIPMENT 825 1423 Public address system A public address system required by this chap ter must a Be powerable when the aircraft is in flight or stopped on the ground after the shutdown or fail ure of all engines and auxiliary power units or the disconnection or failure of all power sources de pendent on their continued operation for 1 A time duration of at least 10 minutes in cluding an aggregate time duration of at least 5 minutes of announcements made by flight and cabin crewmembers considering all other load
611. ycles per minute The effective flash frequency is the fre quency at which the airplane s complete anticolli sion light system is observed from a distance and applies to each sector of light including any over laps that exist when the system consists of more than one light source In overlaps flash frequen cies may exceed 100 but not 180 cycles per minute d Color Each anticollision light must be either aviation red or aviation white and must meet the applicable requirements of 825 1397 e Light intensity The minimum light intensities in all vertical planes measured with the red filter if used and expressed in terms of effective in tensities must meet the requirements of para graph f of this section The following relation must be assumed t oat TPE Ee gt 0 2 5 1 where effective intensity candles I t instantaneous intensity as a function of time tg t4 flash time interval seconds Normally the maximum value of effective intensity is obtained when tz and t4 are chosen so that effective intensity is equal to the instantaneous in tensity at t and t4 ASA 825 1411 f Minimum effective intensities for anticollision lights Each anticollision light effective intensity must equal or exceed the applicable values in the following table Angle above or below the Effective Intensity horizontal plane candles 0 to 5 400 5 to 10 240 10
612. ygen supply terminals immediately available to each occupant wherever seated and at least two oxygen dispensing units connected to oxygen terminals in each lavatory The total num ber of dispensing units and outlets in the cabin must exceed the number of seats by at least 10 percent The extra units must be as uniformly dis tributed throughout the cabin as practicable If cer tification for operation above 30 000 feet is re quested the dispensing units providing the re quired oxygen flow must be automatically presented to the occupants before the cabin pres sure altitude exceeds 15 000 feet The crew must be provided with a manual means of making the dispensing units immediately available in the event of failure of the automatic system 2 Each flight crewmember on flight deck duty must be provided with a quick donning type oxy gen dispensing unit connected to an oxygen sup ply terminal This dispensing unit must be immedi ately available to the flight crewmember when seated at his station and installed so that it i Can be placed on the face from its ready po sition properly secured sealed and supplying oxygen upon demand with one hand within five seconds and without disturbing eyeglasses or causing delay in proceeding with emergency du ties and ii Allows while in place the performance of normal communication functions 3 The oxygen dispensing equipment for the flight crewmembers must be i The diluter dema

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