14 CFR Part 29 — Airworthiness Standards: Transport Category Rotorcraft

Federal Aviation Administration, DOT Pt. 29 29.1203 Fire detector systems. M ISCELLANEOUS E QUIPMENT 29.1431 Electronic equipment.

Subpart F—Equipment 29.1433 Vacuum systems.

29.1435 Hydraulic systems.

G ENERAL 29.1439 Protective breathing equipment.

29.1301 Function and installation.

29.1457 Cockpit voice recorders.

29.1303 Flight and navigation instruments.

29.1459 Flight data recorders.

29.1305 Powerplant instruments.

29.1461 Equipment containing high energy 29.1307 Miscellaneous equipment.

rotors.

29.1309 Equipment, systems, and installa- tions.

Subpart G—Operating Limitations and 29.1316 Electrical and electronic system Information lightning protection.

29.1317 High-intensity Radiated Fields 29.1501 General.

(HIRF) Protection.

O PERATING L IMITATIONS I NSTRUMENTS : I NSTALLATION 29.1503 Airspeed limitations: general.

29.1321 Arrangement and visibility.

29.1505 Never-exceed speed.

29.1322 Warning, caution, and advisory 29.1509 Rotor speed.

lights.

29.1517 Limiting height-velocity envelope.

29.1323 Airspeed indicating system.

29.1519 Weight and center of gravity.

29.1325 Static pressure and pressure altim- 29.1521 Powerplant limitations.

eter systems.

29.1522 Auxiliary power unit limitations.

29.1327 Magnetic direction indicator.

29.1523 Minimum flight crew.

29.1329 Automatic pilot and flight guidance 29.1525 Kinds of operations.

system.

29.1527 Maximum operating altitude.

29.1331 Instruments using a power supply.

29.1529 Instructions for Continued Air- 29.1333 Instrument systems.

worthiness.

29.1337 Powerplant instruments.

M ARKINGS AND P LACARDS E LECTRICAL S YSTEMS AND E QUIPMENT 29.1541 General.

29.1351 General.

29.1543 Instrument markings: general.

29.1353 Energy storage systems.

29.1545 Airspeed indicator.

29.1355 Distribution system.

29.1547 Magnetic direction indicator.

29.1357 Circuit protective devices.

29.1549 Powerplant instruments.

29.1359 Electrical system fire and smoke 29.1551 Oil quantity indicator.

protection.

29.1553 Fuel quantity indicator.

29.1363 Electrical system tests.

29.1555 Control markings.

29.1557 Miscellaneous markings and plac- L IGHTS ards.

29.1381 Instrument lights.

29.1559 Limitations placard.

29.1383 Landing lights.

29.1561 Safety equipment.

29.1385 Position light system installation.

29.1565 Tail rotor.

29.1387 Position light system dihedral an- R OTORCRAFT F LIGHT M ANUAL gles.

29.1389 Position light distribution and in- 29.1581 General.

tensities.

29.1583 Operating limitations.

29.1391 Minimum intensities in the hori- 29.1585 Operating procedures.

zontal plane of forward and rear position 29.1587 Performance information.

lights.

29.1589 Loading information.

29.1393 Minimum intensities in any vertical A PPENDIX A TO P ART 29—I NSTRUCTIONS FOR plane of forward and rear position lights.

C ONTINUED A IRWORTHINESS 29.1395 Maximum intensities in overlapping APPENDIX B TO P ART 29—A IRWORTHINESS C RI - beams of forward and rear position TERIA FOR H ELICOPTER I NSTRUMENT lights.

F LIGHT 29.1397 Color specifications.

A PPENDIX C TO P ART 29—I CING C ERTIFICATION 29.1399 Riding light.

A PPENDIX D TO P ART 29—C RITERIA FOR D EM - 29.1401 Anticollision light system.

ONSTRATION OF E MERGENCY E VACUATION SAFETY EQUIPMENT P ROCEDURES U NDER § 29.803 A PPENDIX E TO P ART 29—HIRF E NVIRON - 29.1411 General.

MENTS AND E QUIPMENT HIRF T EST L EV - 29.1413 Safety belts: passenger warning de- ELS vice.

29.1415 Ditching equipment. A UTHORITY : 49 U.S.C. 106(f), 106(g), 40113, 29.1419 Ice protection. 44701–44702, 44704.

14 CFR Ch. I (1–1–25 Edition) § 29.1 S OURCE : Docket No. 5084, 29 FR 16150, Dec.

der harness that meets the require- 3, 1964, unless otherwise noted.

ments of paragraphs (a), (b), and (c) of this section.

Subpart A—General (a) Each occupant’s seat must have a combined safety belt and shoulder har- § 29.1 Applicability.

ness with a single-point release. Each pilot’s combined safety belt and shoul- (a) This part prescribes airworthiness der harness must allow each pilot, standards for the issue of type certifi- when seated with safety belt and shoul- cates, and changes to those certifi- der harness fastened, to perform all cates, for transport category rotor- functions necessary for flight oper- craft.

ations. There must be a means to se- (b) Transport category rotorcraft cure belts and harnesses, when not in must be certificated in accordance use, to prevent interference with the with either the Category A or Category operation of the rotorcraft and with B requirements of this part. A multien- rapid egress in an emergency.

gine rotorcraft may be type certifi- (b) Each occupant must be protected cated as both Category A and Category from serious head injury by a safety B with appropriate and different oper- belt plus a shoulder harness that will ating limitations for each category.

prevent the head from contacting any (c) Rotorcraft with a maximum injurious object.

weight greater than 20,000 pounds and (c) The safety belt and shoulder har- 10 or more passenger seats must be ness must meet the static and dynamic type certificated as Category A rotor- strength requirements, if applicable, craft.

specified by the rotorcraft type certifi- (d) Rotorcraft with a maximum cation basis.

weight greater than 20,000 pounds and (d) For purposes of this section, the nine or less passenger seats may be date of manufacture is either— type certificated as Category B rotor- (1) The date the inspection accept- craft provided the Category A require- ance records, or equivalent, reflect ments of Subparts C, D, E, and F of that the rotorcraft is complete and this part are met.

meets the FAA-Approved Type Design (e) Rotorcraft with a maximum Data; or weight of 20,000 pounds or less but with (2) The date that the foreign civil air- 10 or more passenger seats may be type worthiness authority certifies the certificated as Category B rotorcraft rotorcraft is complete and issues an provided the Category A requirements original standard airworthiness certifi- of §§ 29.67(a)(2), 29.87, 29.1517, and sub- cate, or equivalent, in that country.

parts C, D, E, and F of this part are met.

[Doc. No. 26078, 56 FR 41052, Aug. 16, 1991] (f) Rotorcraft with a maximum weight of 20,000 pounds or less and nine Subpart B—Flight or less passenger seats may be type certificated as Category B rotorcraft.

G ENERAL (g) Each person who applies under § 29.21 Proof of compliance.

Part 21 for a certificate or change de- scribed in paragraphs (a) through (f) of Each requirement of this subpart this section must show compliance must be met at each appropriate com- with the applicable requirements of bination of weight and center of grav- this part.

ity within the range of loading condi- tions for which certification is re- [Amdt. 29–21, 48 FR 4391, Jan. 31, 1983, as quested. This must be shown— amended by Amdt. 29–39, 61 FR 21898, May 10, 1996; 61 FR 33963, July 1, 1996] (a) By tests upon a rotorcraft of the type for which certification is re- § 29.2 Special retroactive require- quested, or by calculations based on, ments.

and equal in accuracy to, the results of For each rotorcraft manufactured testing; and after September 16, 1992, each applicant (b) By systematic investigation of must show that each occupant’s seat is each required combination of weight equipped with a safety belt and shoul- and center of gravity, if compliance Federal Aviation Administration, DOT § 29.29 cannot be reasonably inferred from lished for any rotorcraft-load combina- combinations investigated. tion if— (1) The rotorcraft-load combination [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as does not include human external cargo, amended by Amdt. 29–24, 49 FR 44435, Nov. 6, (2) Structural component approval 1984] for external load operations under ei- ther § 29.865 or under equivalent oper- § 29.25 Weight limits.

ational standards is obtained, (a) Maximum weight. The maximum (3) The portion of the total weight weight (the highest weight at which that is greater than the maximum compliance with each applicable re- weight established under paragraph (a) quirement of this part is shown) or, at of this section is made up only of the the option of the applicant, the highest weight of all or part of the jettisonable weight for each altitude and for each external load, practicably separable operating condi- (4) Structural components of the tion, such as takeoff, enroute oper- rotorcraft are shown to comply with ation, and landing, must be established the applicable structural requirements so that it is not more than— of this part under the increased loads (1) The highest weight selected by and stresses caused by the weight in- the applicant; crease over that established under (2) The design maximum weight (the paragraph (a) of this section, and highest weight at which compliance (5) Operation of the rotorcraft at a with each applicable structural loading total weight greater than the max- condition of this part is shown); or imum certificated weight established (3) The highest weight at which com- under paragraph (a) of this section is pliance with each applicable flight re- limited by appropriate operating limi- quirement of this part is shown. tations under § 29.865 (a) and (d) of this (4) For Category B rotorcraft with 9 part.

or less passenger seats, the maximum [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as weight, altitude, and temperature at amended by Amdt. 29–12, 41 FR 55471, Dec. 20, which the rotorcraft can safely operate 1976; Amdt. 29–43, 64 FR 43020, Aug. 6, 1999; near the ground with the maximum Amdt. 29–51, 73 FR 11001, Feb. 29, 2008] wind velocity determined under § 29.27 Center of gravity limits.

§ 29.143(c) and may include other dem- onstrated wind velocities and azi- The extreme forward and aft centers muths. The operating envelopes must of gravity and, where critical, the ex- be stated in the Limitations section of treme lateral centers of gravity must the Rotorcraft Flight Manual.

be established for each weight estab- (b) Minimum weight. The minimum lished under § 29.25. Such an extreme weight (the lowest weight at which may not lie beyond— compliance with each applicable re- (a) The extremes selected by the ap- quirement of this part is shown) must plicant; (b) The extremes within which the be established so that it is not less structure is proven; or than— (c) The extremes within which com- (1) The lowest weight selected by the pliance with the applicable flight re- applicant; quirements is shown.

(2) The design minimum weight (the lowest weight at which compliance [Amdt. 29–3, 33 FR 965, Jan. 26, 1968] with each structural loading condition of this part is shown); or § 29.29 Empty weight and cor- responding center of gravity.

(3) The lowest weight at which com- pliance with each applicable flight re- (a) The empty weight and cor- quirement of this part is shown.

responding center of gravity must be (c) Total weight with jettisonable exter- determined by weighing the rotorcraft nal load. A total weight for the rotor- without the crew and payload, but craft with a jettisonable external load with— attached that is greater than the max- (1) Fixed ballast; imum weight established under para- (2) Unusable fuel; and graph (a) of this section may be estab- (3) Full operating fluids, including— 14 CFR Ch. I (1–1–25 Edition) § 29.31 (i) Oil; (3) Adequate means to warn the pilot (ii) Hydraulic fluid; and of unsafe main rotor speeds.

(iii) Other fluids required for normal (c) Normal main rotor low pitch limit operation of rotorcraft systems, except (power off). It must be shown, with water intended for injection in the en- power off, that— gines.

(1) The normal main rotor low pitch (b) The condition of the rotorcraft at limit provides sufficient rotor speed, in the time of determining empty weight any autorotative condition, under the must be one that is well defined and most critical combinations of weight can be easily repeated, particularly and airspeed; and with respect to the weights of fuel, oil, (2) It is possible to prevent over- coolant, and installed equipment.

speeding of the rotor without excep- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- tional piloting skill.

eral Aviation Act of 1958 (49 U.S.C. 1354(a), (d) Emergency high pitch. If the main 1421, 1423, 1424, and 1425); and sec. 6(c) of the rotor high pitch stop is set to meet Dept. of Transportation Act (49 U.S.C.

1655(c))) paragraph (b)(1) of this section, and if that stop cannot be exceeded inadvert- [Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as ently, additional pitch may be made amended by Amdt. 29–15, 43 FR 2326, Jan. 16, 1978] available for emergency use.

(e) Main rotor low speed warning for § 29.31 Removable ballast.

helicopters. For each single engine heli- Removable ballast may be used in copter, and each multiengine heli- showing compliance with the flight re- copter that does not have an approved quirements of this subpart.

device that automatically increases power on the operating engines when § 29.33 Main rotor speed and pitch lim- one engine fails, there must be a main its.

rotor low speed warning which meets (a) Main rotor speed limits. A range of the following requirements: main rotor speeds must be established (1) The warning must be furnished to that— the pilot in all flight conditions, in- (1) With power on, provides adequate cluding power-on and power-off flight, margin to accommodate the variations when the speed of a main rotor ap- in rotor speed occurring in any appro- proaches a value that can jeopardize priate maneuver, and is consistent safe flight.

with the kind of governor or synchro- (2) The warning may be furnished ei- nizer used; and ther through the inherent aerodynamic (2) With power off, allows each appro- qualities of the helicopter or by a de- priate autorotative maneuver to be performed throughout the ranges of vice.

airspeed and weight for which certifi- (3) The warning must be clear and cation is requested.

distinct under all conditions, and must (b) Normal main rotor high pitch limit be clearly distinguishable from all (power on). For rotorcraft, except heli- other warnings. A visual device that copters required to have a main rotor requires the attention of the crew low speed warning under paragraph (e) within the cockpit is not acceptable by of this section, it must be shown, with itself.

power on and without exceeding ap- (4) If a warning device is used, the de- proved engine maximum limitations, vice must automatically deactivate that main rotor speeds substantially and reset when the low-speed condition less than the minimum approved main is corrected. If the device has an audi- rotor speed will not occur under any ble warning, it must also be equipped sustained flight condition. This must with a means for the pilot to manually be met by— (1) Appropriate setting of the main rotor high pitch stop; (2) Inherent rotorcraft characteris- tics that make unsafe low main rotor speeds unlikely; or Federal Aviation Administration, DOT § 29.51 silence the audible warning before the achieve the applicable rotorcraft per- low-speed condition is corrected. formance prescribed in this subpart.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

1421, 1423, 1424, and 1425); and sec. 6(c) of the of Transportation Act (49 U.S.C. 1655(c))) Dept. of Transportation Act (49 U.S.C.

1655(c))) [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–15, 43 FR 2326, Jan. 16, [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as 1978; Amdt. 29–24, 49 FR 44436, Nov. 6, 1984] amended by Amdt. 29–3, 33 FR 965, Jan. 26, 1968; Amdt. 29–15, 43 FR 2326, Jan. 16, 1978] § 29.49 Performance at minimum oper- ating speed.

P ERFORMANCE (a) For each Category A helicopter, § 29.45 General. the hovering performance must be de- termined over the ranges of weight, al- (a) The performance prescribed in titude, and temperature for which this subpart must be determined— takeoff data are scheduled— (1) With normal piloting skill and; (1) With not more than takeoff (2) Without exceptionally favorable power; conditions.

(2) With the landing gear extended; (b) Compliance with the performance and requirements of this subpart must be (3) At a height consistent with the shown— procedure used in establishing the (1) For still air at sea level with a takeoff, climbout, and rejected takeoff standard atmosphere and; paths.

(2) For the approved range of atmos- (b) For each Category B helicopter, pheric variables. the hovering performance must be de- termined over the ranges of weight, al- (c) The available power must cor- titude, and temperature for which cer- respond to engine power, not exceeding tification is requested, with— the approved power, less— (1) Takeoff power; (1) Installation losses; and (2) The landing gear extended; and (2) The power absorbed by the acces- (3) The helicopter in ground effect at sories and services at the values for a height consistent with normal take- which certification is requested and ap- off procedures.

proved.

(c) For each helicopter, the out-of- (d) For reciprocating engine-powered ground effect hovering performance rotorcraft, the performance, as affected must be determined over the ranges of by engine power, must be based on a weight, altitude, and temperature for relative humidity of 80 percent in a which certification is requested with standard atmosphere.

takeoff power.

(e) For turbine engine-powered rotor- (d) For rotorcraft other than heli- craft, the performance, as affected by copters, the steady rate of climb at the engine power, must be based on a rel- minimum operating speed must be de- ative humidity of— termined over the ranges of weight, al- titude, and temperature for which cer- (1) 80 percent, at and below standard tification is requested with— temperature; and (1) Takeoff power; and (2) 34 percent, at and above standard (2) The landing gear extended.

temperature plus 50 ° F.

[Doc. No. 24802, 61 FR 21898, May 10, 1996; 61 Between these two temperatures, the FR 33963, July 1, 1996] relative humidity must vary linearly.

(f) For turbine-engine-power rotor- § 29.51 Takeoff data: general.

craft, a means must be provided to per- (a) The takeoff data required by mit the pilot to determine prior to §§ 29.53, 29.55, 29.59, 29.60, 29.61, 29.62, takeoff that each engine is capable of 29.63, and 29.67 must be determined— developing the power necessary to (1) At each weight, altitude, and tem- perature selected by the applicant; and 14 CFR Ch. I (1–1–25 Edition) § 29.53 (2) With the operating engines within (1) The takeoff path must remain approved operating limitations. clear of the height-velocity envelope established in accordance with § 29.87; (b) Takeoff data must— (2) The rotorcraft must be flown to (1) Be determined on a smooth, dry, the engine failure point; at which hard surface; and point, the critical engine must be made (2) Be corrected to assume a level inoperative and remain inoperative for takeoff surface.

the rest of the takeoff; (c) No takeoff made to determine the (3) After the critical engine is made data required by this section may re- inoperative, the rotorcraft must con- quire exceptional piloting skill or tinue to the takeoff decision point, and alertness, or exceptionally favorable then attain V ; TOSS conditions.

(4) Only primary controls may be used while attaining V and while [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as TOSS amended by Amdt. 29–39, 61 FR 21899, May 10, establishing a positive rate of climb.

1996] Secondary controls that are located on the primary controls may be used after § 29.53 Takeoff: Category A.

a positive rate of climb and V are TOSS established but in no case less than 3 The takeoff performance must be de- seconds after the critical engine is termined and scheduled so that, if one made inoperative; and engine fails at any time after the start (5) After attaining V and a posi- TOSS of takeoff, the rotorcraft can— tive rate of a climb, the landing gear (a) Return to, and stop safely on, the may be retracted.

takeoff area; or (b) During the takeoff path deter- (b) Continue the takeoff and mination made in accordance with climbout, and attain a configuration paragraph (a) of this section and after and airspeed allowing compliance with attaining V and a positive rate of TOSS § 29.67(a)(2).

climb, the climb must be continued at [Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 a speed as close as practicable to, but FR 33963, July 1, 1996] not less than, V until the rotorcraft TOSS is 200 feet above the takeoff surface.

§ 29.55 Takeoff decision point (TDP): During this interval, the climb per- Category A.

formance must meet or exceed that re- (a) The TDP is the first point from quired by § 29.67(a)(1).

(c) During the continued takeoff, the which a continued takeoff capability is rotorcraft shall not descend below 15 assured under § 29.59 and is the last feet above the takeoff surface when the point in the takeoff path from which a takeoff decision point is above 15 feet.

rejected takeoff is assured within the (d) From 200 feet above the takeoff distance determined under § 29.62.

surface, the rotorcraft takeoff path (b) The TDP must be established in must be level or positive until a height relation to the takeoff path using no 1,000 feet above the takeoff surface is more than two parameters; e.g., air- attained with not less than the rate of speed and height, to designate the climb required by § 29.67(a)(2). Any sec- TDP.

ondary or auxiliary control may be (c) Determination of the TDP must used after attaining 200 feet above the include the pilot recognition time in- takeoff surface.

terval following failure of the critical (e) Takeoff distance will be deter- engine.

mined in accordance with § 29.61.

[Doc. No. 24802, 61 FR 21899, May 10, 1996] [Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 FR 33963, July 1, 1996, as amended by Amdt.

§ 29.59 Takeoff path: Category A.

29–44, 64 FR 45337, Aug. 19, 1999] (a) The takeoff path extends from the § 29.60 Elevated heliport takeoff path: point of commencement of the takeoff Category A.

procedure to a point at which the rotorcraft is 1,000 feet above the take- (a) The elevated heliport takeoff path off surface and compliance with extends from the point of commence- § 29.67(a)(2) is shown. In addition— ment of the takeoff procedure to a Federal Aviation Administration, DOT § 29.64 point in the takeoff path at which the (a) The takeoff path requirements of rotorcraft is 1,000 feet above the take- §§ 29.59 and 29.60 being used up to the off surface and compliance with TDP where the critical engine failure § 29.67(a)(2) is shown. In addition— is recognized and the rotorcraft is land- ed and brought to a complete stop on (1) The requirements of § 29.59(a) the takeoff surface; must be met; (2) While attaining V and a posi- (b) The remaining engines operating TOSS tive rate of climb, the rotorcraft may within approved limits; descend below the level of the takeoff (c) The landing gear remaining ex- surface if, in so doing and when clear- tended throughout the entire rejected ing the elevated heliport edge, every takeoff; and part of the rotorcraft clears all obsta- (d) The use of only the primary con- cles by at least 15 feet; trols until the rotorcraft is on the (3) The vertical magnitude of any de- ground. Secondary controls located on scent below the takeoff surface must be the primary control may not be used determined; and until the rotorcraft is on the ground.

Means other than wheel brakes may be (4) After attaining V and a posi- TOSS used to stop the rotorcraft if the means tive rate of climb, the landing gear are safe and reliable and consistent re- may be retracted.

sults can be expected under normal op- (b) The scheduled takeoff weight erating conditions.

must be such that the climb require- ments of § 29.67 (a)(1) and (a)(2) will be [Doc. No. 24802, 61 FR 21899, May 10, 1996, as met.

amended by Amdt. 29–44, 64 FR 45337, Aug. 19, (c) Takeoff distance will be deter- 1999] mined in accordance with § 29.61.

§ 29.63 Takeoff: Category B.

[Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 The horizontal distance required to FR 33963, July 1, 1996] take off and climb over a 50-foot obsta- § 29.61 Takeoff distance: Category A. cle must be established with the most unfavorable center of gravity. The (a) The normal takeoff distance is takeoff may be begun in any manner the horizontal distance along the take- if— off path from the start of the takeoff to (a) The takeoff surface is defined; the point at which the rotorcraft at- (b) Adequate safeguards are main- tains and remains at least 35 feet above tained to ensure proper center of grav- the takeoff surface, attains and main- ity and control positions; and tains a speed of at least V , and es- TOSS (c) A landing can be made safely at tablishes a positive rate of climb, as- any point along the flight path if an suming the critical engine failure oc- engine fails.

curs at the engine failure point prior to the takeoff decision point.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (b) For elevated heliports, the take- amended by Amdt. 29–12, 41 FR 55471, Dec. 20, off distance is the horizontal distance 1976] along the takeoff path from the start § 29.64 Climb: General.

of the takeoff to the point at which the rotorcraft attains and maintains a Compliance with the requirements of speed of at least V and establishes a TOSS §§ 29.65 and 29.67 must be shown at each positive rate of climb, assuming the weight, altitude, and temperature critical engine failure occurs at the en- within the operational limits estab- gine failure point prior to the takeoff lished for the rotorcraft and with the decision point.

most unfavorable center of gravity for each configuration. Cowl flaps, or other [Doc. No. 24802, 61 FR 21899, May 10, 1996] means of controlling the engine-cool- ing air supply, will be in the position § 29.62 Rejected takeoff: Category A.

that provides adequate cooling at the The rejected takeoff distance and temperatures and altitudes for which procedures for each condition where certification is requested.

takeoff is approved will be established with— [Doc. No. 24802, 61 FR 21900, May 10, 1996] 14 CFR Ch. I (1–1–25 Edition) § 29.65 which certification for use of 30-minute § 29.65 Climb: All engines operating.

OEI power is requested; (a) The steady rate of climb must be (ii) The landing gear retracted; and determined— (iii) The speed selected by the appli- (1) With maximum continuous power; cant.

(2) With the landing gear retracted; (3) The steady rate of climb (or de- and scent) in feet per minute, at each alti- (3) At V for standard sea level condi- y tude and temperature at which the tions and at speeds selected by the ap- rotorcraft is expected to operate and at plicant for other conditions.

any weight within the range of weights (b) For each Category B rotorcraft for which certification is requested, except helicopters, the rate of climb must be determined with— determined under paragraph (a) of this (i) The critical engine inoperative section must provide a steady climb and the remaining engines at max- gradient of at least 1:6 under standard imum continuous power including con- sea level conditions.

tinuous OEI power, if approved, and at 30-minute OEI power for rotorcraft for (Secs. 313(a), 601, 603, 604, and 605 of the Fed- which certification for the use of 30- eral Aviation Act of 1958 (49 U.S.C. 1354(a), minute OEI power is requested; 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

of Transportation Act (49 U.S.C. 1655(c))) (ii) The landing gear retracted; and (iii) The speed selected by the appli- [Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as cant.

amended by Amdt. 29–15, 43 FR 2326, Jan. 16, (b) For multiengine Category B 1978; Amdt. 29–39, 61 FR 21900, May 10, 1996; 61 rotorcraft meeting the Category A en- FR 33963, July 1, 1996] gine isolation requirements, the steady § 29.67 Climb: One engine inoperative rate of climb (or descent) must be de- (OEI).

termined at the speed for best rate of climb (or minimum rate of descent) at (a) For Category A rotorcraft, in the each altitude, temperature, and weight critical takeoff configuration existing at which the rotorcraft is expected to along the takeoff path, the following operate, with the critical engine inop- apply: erative and the remaining engines at (1) The steady rate of climb without maximum continuous power including ground effect, 200 feet above the take- continuous OEI power, if approved, and off surface, must be at least 100 feet per at 30-minute OEI power for rotorcraft minute for each weight, altitude, and for which certification for the use of 30- temperature for which takeoff data are minute OEI power is requested.

to be scheduled with— (i) The critical engine inoperative [Doc. No. 24802, 61 FR 21900, May 10, 1996; 61 and the remaining engines within ap- FR 33963, July 1, 1996, as amended by Amdt.

29–44, 64 FR 45337, Aug. 19, 1999; 64 FR 47563, proved operating limitations, except Aug. 31, 1999] that for rotorcraft for which the use of 30-second/2-minute OEI power is re- § 29.71 Helicopter angle of glide: Cat- quested, only the 2-minute OEI power egory B.

may be used in showing compliance For each category B helicopter, ex- with this paragraph; cept multiengine helicopters meeting (ii) The landing gear extended; and the requirements of § 29.67(b) and the (iii) The takeoff safety speed selected powerplant installation requirements by the applicant.

of category A, the steady angle of glide (2) The steady rate of climb without must be determined in autorotation— ground effect, 1000 feet above the take- (a) At the forward speed for min- off surface, must be at least 150 feet per imum rate of descent as selected by the minute, for each weight, altitude, and applicant; temperature for which takeoff data are (b) At the forward speed for best glide to be scheduled with— angle; (i) The critical engine inoperative (c) At maximum weight; and and the remaining engines at max- (d) At the rotor speed or speeds se- imum continuous power including con- lected by the applicant.

tinuous OEI power, if approved, or at 30-minute OEI power for rotorcraft for [Amdt. 29–12, 41 FR 55471, Dec. 20, 1976] Federal Aviation Administration, DOT § 29.85 ical areas of the height-velocity enve- § 29.75 Landing: General.

lope determined in accordance with (a) For each rotorcraft— § 29.87.

(1) The corrected landing data must (b) It must be possible to make a safe be determined for a smooth, dry, hard, landing on a prepared landing surface and level surface; after complete power failure occurring (2) The approach and landing must during normal cruise.

not require exceptional piloting skill or exceptionally favorable conditions; [Doc. No. 24802, 61 FR 21900, May 10, 1996] and (3) The landing must be made with- § 29.81 Landing distance: Category A.

out excessive vertical acceleration or The horizontal distance required to tendency to bounce, nose over, ground land and come to a complete stop (or to loop, porpoise, or water loop.

a speed of approximately 3 knots for (b) The landing data required by water landings) from a point 50 ft §§ 29.77, 29.79, 29.81, 29.83, and 29.85 must above the landing surface must be de- be determined— termined from the approach and land- (1) At each weight, altitude, and tem- ing paths established in accordance perature for which landing data are ap- with § 29.79.

proved; [Doc. No. 24802, 64 FR 45338, Aug. 19, 1999] (2) With each operating engine within approved operating limitations; and § 29.83 Landing: Category B.

(3) With the most unfavorable center of gravity.

(a) For each Category B rotorcraft, the horizontal distance required to [Doc. No. 24802, 61 FR 21900, May 10, 1996] land and come to a complete stop (or to a speed of approximately 3 knots for § 29.77 Landing Decision Point (LDP): Category A. water landings) from a point 50 feet above the landing surface must be de- (a) The LDP is the last point in the termined with— approach and landing path from which (1) Speeds appropriate to the type of a balked landing can be accomplished rotorcraft and chosen by the applicant in accordance with § 29.85.

to avoid the critical areas of the (b) Determination of the LDP must height-velocity envelope established include the pilot recognition time in- under § 29.87; and terval following failure of the critical (2) The approach and landing made engine.

with power on and within approved [Doc. No. 24802, 64 FR 45338, Aug. 19, 1999] limits.

(b) Each multiengined Category B § 29.79 Landing: Category A.

rotorcraft that meets the powerplant (a) For Category A rotorcraft— installation requirements for Category (1) The landing performance must be A must meet the requirements of— determined and scheduled so that if the (1) Sections 29.79 and 29.81; or critical engine fails at any point in the (2) Paragraph (a) of this section.

approach path, the rotorcraft can ei- (c) It must be possible to make a safe ther land and stop safely or climb out landing on a prepared landing surface if and attain a rotorcraft configuration complete power failure occurs during and speed allowing compliance with normal cruise.

the climb requirement of § 29.67(a)(2); [Doc. No. 24802, 61 FR 21900, May 10, 1996; 61 (2) The approach and landing paths FR 33963, July 1, 1996] must be established with the critical engine inoperative so that the transi- § 29.85 Balked landing: Category A.

tion between each stage can be made smoothly and safely; For Category A rotorcraft, the (3) The approach and landing speeds balked landing path with the critical must be selected by the applicant and engine inoperative must be established must be appropriate to the type of so that— rotorcraft; and (a) The transition from each stage of (4) The approach and landing path the maneuver to the next stage can be must be established to avoid the crit- made smoothly and safely; 14 CFR Ch. I (1–1–25 Edition) § 29.87 (b) From the LDP on the approach (3) For power-on operations, under path selected by the applicant, a safe any condition of speed, power, and climbout can be made at speeds allow- rotor r.p.m. for which certification is ing compliance with the climb require- requested; and ments of § 29.67(a)(1) and (2); and (4) For power-off operations, under (c) The rotorcraft does not descend any condition of speed, and rotor r.p.m.

below 15 feet above the landing surface. for which certification is requested For elevated heliport operations, de- that is attainable with the controls scent may be below the level of the rigged in accordance with the approved landing surface provided the deck edge rigging instructions and tolerances; clearance of § 29.60 is maintained and (b) Be able to maintain any required the descent (loss of height) below the flight condition and make a smooth landing surface is determined. transition from any flight condition to any other flight condition without ex- [Doc. No. 24802, 64 FR 45338, Aug. 19, 1999] ceptional piloting skill, alertness, or strength, and without danger of ex- § 29.87 Height-velocity envelope.

ceeding the limit load factor under any (a) If there is any combination of operating condition probable for the height and forward velocity (including type, including— hover) under which a safe landing can- (1) Sudden failure of one engine, for not be made after failure of the critical multiengine rotorcraft meeting Trans- engine and with the remaining engines port Category A engine isolation re- (where applicable) operating within ap- quirements; proved limits, a height-velocity enve- (2) Sudden, complete power failure, lope must be established for— for other rotorcraft; and (1) All combinations of pressure alti- (3) Sudden, complete control system tude and ambient temperature for failures specified in § 29.695 of this part; which takeoff and landing are ap- and proved; and (c) Have any additional characteris- (2) Weight from the maximum weight tics required for night or instrument (at sea level) to the highest weight ap- operation, if certification for those proved for takeoff and landing at each kinds of operation is requested. Re- altitude. For helicopters, this weight quirements for helicopter instrument need not exceed the highest weight al- flight are contained in appendix B of lowing hovering out-of-ground effect at this part.

each altitude.

[Doc. No. 5084, 29 FR 16150, Dec. 8, 1964, as (b) For single-engine or multiengine amended by Amdt. 29–3, 33 FR 905, Jan. 26, rotorcraft that do not meet the Cat- 1968; Amdt. 29–12, 41 FR 55471, Dec. 20, 1976; egory A engine isolation requirements, Amdt. 29–21, 48 FR 4391, Jan. 31, 1983; Amdt.

the height-velocity envelope for com- 29–24, 49 FR 44436, Nov. 6, 1984] plete power failure must be estab- lished.

§ 29.143 Controllability and maneuver- ability.

[Doc. No. 24802, 61 FR 21901, May 10, 1996; 61 FR 33963, July 1, 1996] (a) The rotorcraft must be safely con- trollable and maneuverable— F LIGHT C HARACTERISTICS (1) During steady flight; and (2) During any maneuver appropriate § 29.141 General.

to the type, including— The rotorcraft must— (i) Takeoff; (a) Except as specifically required in (ii) Climb; the applicable section, meet the flight (iii) Level flight; characteristics requirements of this (iv) Turning flight; subpart— (v) Autorotation; and (1) At the approved operating alti- (vi) Landing (power on and power tudes and temperatures; off).

(2) Under any critical loading condi- (b) The margin of cyclic control must tion within the range of weights and allow satisfactory roll and pitch con- centers of gravity for which certifi- trol at V with— NE cation is requested; and (1) Critical weight; Federal Aviation Administration, DOT § 29.173 (2) Critical center of gravity; (1) The helicopter must be safely slowed to V (power-off), without ex- (3) Critical rotor r.p.m.; and NE ceptional pilot skill after the last oper- (4) Power off (except for helicopters ating engine is made inoperative at demonstrating compliance with para- power-on V .

graph (f) of this section) and power on. NE (2) At a speed of 1.1 V (power-off), NE (c) Wind velocities from zero to at the margin of cyclic control must least 17 knots, from all azimuths, must allow satisfactory roll and pitch con- be established in which the rotorcraft trol with power off.

can be operated without loss of control on or near the ground in any maneuver (Secs. 313(a), 601, 603, 604, and 605 of the Fed- appropriate to the type (such as cross- eral Aviation Act of 1958 (49 U.S.C. 1354(a), wind takeoffs, sideward flight, and 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C.

rearward flight), with— 1655(c))) (1) Critical weight; (2) Critical center of gravity; [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–3, 33 FR 965, Jan. 26, (3) Critical rotor r.p.m.; and 1968; Amdt. 29–15, 43 FR 2326, Jan. 16, 1978; (4) Altitude, from standard sea level Amdt. 29–24, 49 FR 44436, Nov. 6, 1984; Amdt.

conditions to the maximum takeoff 29–51, 73 FR 11001, Feb. 29, 2008] and landing altitude capability of the rotorcraft.

§ 29.151 Flight controls.

(d) Wind velocities from zero to at (a) Longitudinal, lateral, directional, least 17 knots, from all azimuths, must and collective controls may not exhibit be established in which the rotorcraft excessive breakout force, friction, or can be operated without loss of control preload.

out-of-ground effect, with— (b) Control system forces and free (1) Weight selected by the applicant; play may not inhibit a smooth, direct (2) Critical center of gravity; rotorcraft response to control system (3) Rotor r.p.m. selected by the appli- input.

cant; and [Amdt. 29–24, 49 FR 44436, Nov. 6, 1984] (4) Altitude, from standard sea level conditions to the maximum takeoff § 29.161 Trim control.

and landing altitude capability of the The trim control— rotorcraft.

(a) Must trim any steady longitu- (e) The rotorcraft, after (1) failure of dinal, lateral, and collective control one engine, in the case of multiengine forces to zero in level flight at any ap- rotorcraft that meet Transport Cat- propriate speed; and egory A engine isolation requirements, (b) May not introduce any undesir- or (2) complete power failure in the able discontinuities in control force case of other rotorcraft, must be con- gradients.

trollable over the range of speeds and altitudes for which certification is re- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as quested when such power failure occurs amended by Amdt. 29–24, 49 FR 44436, Nov. 6, 1984] with maximum continuous power and critical weight. No corrective action § 29.171 Stability: general.

time delay for any condition following power failure may be less than— The rotorcraft must be able to be (i) For the cruise condition, one sec- flown, without undue pilot fatigue or ond, or normal pilot reaction time strain, in any normal maneuver for a (whichever is greater); and period of time as long as that expected in normal operation. At least three (ii) For any other condition, normal landings and takeoffs must be made pilot reaction time.

during this demonstration.

(f) For helicopters for which a V NE (power-off) is established under § 29.173 Static longitudinal stability.

§ 29.1505(c), compliance must be dem- onstrated with the following require- (a) The longitudinal control must be ments with critical weight, critical designed so that a rearward movement center of gravity, and critical rotor of the control is necessary to obtain an r.p.m.: airspeed less than the trim speed, and a 14 CFR Ch. I (1–1–25 Edition) § 29.175 forward movement of the control is (d) Autorotation. Static longitudinal necessary to obtain an airspeed more stability must be shown in autorota- than the trim speed. tion at— (b) Throughout the full range of alti- (1) Airspeeds from the minimum rate tude for which certification is re- of descent airspeed ¥ 10 kt to the min- quested, with the throttle and collec- imum rate of descent airspeed + 10 kt, tive pitch held constant during the ma- with— neuvers specified in § 29.175(a) through (i) Critical weight; (d), the slope of the control position (ii) Critical center of gravity; versus airspeed curve must be positive. (iii) The landing gear extended; and However, in limited flight conditions (iv) The rotorcraft trimmed at the or modes of operation determined by minimum rate of descent airspeed.

the Administrator to be acceptable, the (2) Airspeeds from the best angle-of- slope of the control position versus air- glide airspeed ¥ 10kt to the best angle- speed curve may be neutral or negative of-glide airspeed + 10kt, with— if the rotorcraft possesses flight char- (i) Critical weight; acteristics that allow the pilot to (ii) Critical center of gravity; maintain airspeed within ± 5 knots of (iii) The landing gear retracted; and the desired trim airspeed without ex- (iv) The rotorcraft trimmed at the ceptional piloting skill or alertness. best angle-of-glide airspeed.

[Amdt. 29–24, 49 FR 44436, Nov. 6, 1984, as [Amdt. 29–51, 73 FR 11001, Feb. 29, 2008] amended by Amdt. 29–51, 73 FR 11001, Feb. 29, 2008] § 29.177 Static directional stability.

(a) The directional controls must op- § 29.175 Demonstration of static longi- erate in such a manner that the sense tudinal stability.

and direction of motion of the rotor- (a) Climb. Static longitudinal sta- craft following control displacement bility must be shown in the climb con- are in the direction of the pedal motion dition at speeds from Vy ¥ 10 kt to Vy with throttle and collective controls + 10 kt with— held constant at the trim conditions (1) Critical weight; specified in § 29.175(a), (b), (c), and (d).

(2) Critical center of gravity; Sideslip angles must increase with (3) Maximum continuous power; steadily increasing directional control (4) The landing gear retracted; and deflection for sideslip angles up to the (5) The rotorcraft trimmed at Vy.

lesser of— (b) Cruise. Static longitudinal sta- (1) ± 25 degrees from trim at a speed of bility must be shown in the cruise con- 15 knots less than the speed for min- dition at speeds from 0.8 V ¥ 10 kt to NE imum rate of descent varying linearly 0.8 V + 10 kt or, if V is less than 0.8 NE H to ± 10 degrees from trim at V ; NE V , from VH ¥ 10 kt to V + 10 kt, NE H (2) The steady-state sideslip angles with— established by § 29.351; (1) Critical weight; (3) A sideslip angle selected by the (2) Critical center of gravity; applicant, which corresponds to a (3) Power for level flight at 0.8 V or NE sideforce of at least 0.1g; or V , whichever is less; H (4) The sideslip angle attained by (4) The landing gear retracted; and maximum directional control input.

(5) The rotorcraft trimmed at 0.8 V NE (b) Sufficient cues must accompany or V , whichever is less.

H the sideslip to alert the pilot when ap- (c) V . Static longitudinal stability NE proaching sideslip limits.

must be shown at speeds from V ¥ 20 NE (c) During the maneuver specified in kt to V with— NE paragraph (a) of this section, the side- (1) Critical weight; slip angle versus directional control (2) Critical center of gravity; position curve may have a negative (3) Power required for level flight at slope within a small range of angles V ¥ 10 kt or maximum continuous NE around trim, provided the desired head- power, whichever is less; ing can be maintained without excep- (4) The landing gear retracted; and tional piloting skill or alertness.

(5) The rotorcraft trimmed at V ¥ NE 10 kt. [Amdt. 29–51, 73 FR 11001, Feb. 29, 2008] Federal Aviation Administration, DOT § 29.307 Unless otherwise provided, prescribed § 29.181 Dynamic stability: Category A rotorcraft. loads are limit loads.

(b) Unless otherwise provided, the Any short-period oscillation occur- specified air, ground, and water loads ring at any speed from V to V must Y NE must be placed in equilibrium with in- be positively damped with the primary ertia forces, considering each item of flight controls free and in a fixed posi- mass in the rotorcraft. These loads tion.

must be distributed to closely approxi- [Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] mate or conservatively represent ac- tual conditions.

G ROUND AND W ATER H ANDLING (c) If deflections under load would C HARACTERISTICS significantly change the distribution of external or internal loads, this redis- § 29.231 General.

tribution must be taken into account.

The rotorcraft must have satisfac- tory ground and water handling char- § 29.303 Factor of safety.

acteristics, including freedom from un- Unless otherwise provided, a factor of controllable tendencies in any condi- safety of 1.5 must be used. This factor tion expected in operation.

applies to external and inertia loads unless its application to the resulting § 29.235 Taxiing condition.

internal stresses is more conservative.

The rotorcraft must be designed to withstand the loads that would occur § 29.305 Strength and deformation.

when the rotorcraft is taxied over the (a) The structure must be able to roughest ground that may reasonably support limit loads without detri- be expected in normal operation.

mental or permanent deformation. At any load up to limit loads, the defor- § 29.239 Spray characteristics.

mation may not interfere with safe op- If certification for water operation is eration.

requested, no spray characteristics (b) The structure must be able to during taxiing, takeoff, or landing may support ultimate loads without failure.

obscure the vision of the pilot or dam- This must be shown by— age the rotors, propellers, or other (1) Applying ultimate loads to the parts of the rotorcraft.

structure in a static test for at least three seconds; or § 29.241 Ground resonance.

(2) Dynamic tests simulating actual The rotorcraft may have no dan- load application.

gerous tendency to oscillate on the § 29.307 Proof of structure.

ground with the rotor turning.

(a) Compliance with the strength and MISCELLANEOUS F LIGHT R EQUIREMENTS deformation requirements of this sub- part must be shown for each critical § 29.251 Vibration.

loading condition accounting for the Each part of the rotorcraft must be environment to which the structure free from excessive vibration under will be exposed in operation. Struc- each appropriate speed and power con- tural analysis (static or fatigue) may dition.

be used only if the structure conforms to those structures for which experi- Subpart C—Strength Requirements ence has shown this method to be reli- able. In other cases, substantiating G ENERAL load tests must be made.

(b) Proof of compliance with the § 29.301 Loads.

strength requirements of this subpart must include— (a) Strength requirements are speci- fied in terms of limit loads (the max- (1) Dynamic and endurance tests of imum loads to be expected in service) rotors, rotor drives, and rotor controls; and ultimate loads (limit loads multi- (2) Limit load tests of the control plied by prescribed factors of safety). system, including control surfaces; 14 CFR Ch. I (1–1–25 Edition) § 29.309 (3) Operation tests of the control sys- § 29.337 Limit maneuvering load fac- tem; tor.

(4) Flight stress measurement tests; The rotorcraft must be designed for— (5) Landing gear drop tests; and (a) A limit maneuvering load factor (6) Any additional tests required for ranging from a positive limit of 3.5 to new or unusual design features.

a negative limit of ¥ 1.0; or (b) Any positive limit maneuvering (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, load factor not less than 2.0 and any 1425) negative limit maneuvering load factor [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as of not less than ¥ 0.5 for which— amended by Amdt. 29–4, 33 FR 14106, Sept. 18, (1) The probability of being exceeded 1968; Amdt. 27–26, 55 FR 8001, Mar. 6, 1990] is shown by analysis and flight tests to be extremely remote; and § 29.309 Design limitations.

(2) The selected values are appro- The following values and limitations priate to each weight condition be- must be established to show compli- tween the design maximum and design ance with the structural requirements minimum weights.

of this subpart: [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (a) The design maximum and design amended by Amdt. 27–26, 55 FR 8002, Mar. 6, minimum weights.

1990] (b) The main rotor r.p.m. ranges, power on and power off. § 29.339 Resultant limit maneuvering loads.

(c) The maximum forward speeds for each main rotor r.p.m. within the The loads resulting from the applica- ranges determined under paragraph (b) tion of limit maneuvering load factors of this section.

are assumed to act at the center of (d) The maximum rearward and side- each rotor hub and at each auxiliary ward flight speeds. lifting surface, and to act in directions and with distributions of load among (e) The center of gravity limits cor- the rotors and auxiliary lifting sur- responding to the limitations deter- faces, so as to represent each critical mined under paragraphs (b), (c), and (d) maneuvering condition, including of this section.

power-on and power-off flight with the (f) The rotational speed ratios be- maximum design rotor tip speed ratio.

tween each powerplant and each con- The rotor tip speed ratio is the ratio of nected rotating component.

the rotorcraft flight velocity compo- (g) The positive and negative limit nent in the plane of the rotor disc to maneuvering load factors.

the rotational tip speed of the rotor blades, and is expressed as follows: F LIGHT L OADS V cos a § 29.321 General.

μ = (a) The flight load factor must be as- R Ω sumed to act normal to the longitu- where— dinal axis of the rotorcraft, and to be V = The airspeed along the flight path equal in magnitude and opposite in di- (f.p.s.); a = The angle between the projection, in the rection to the rotorcraft inertia load plane of symmetry, of the axis of no factor at the center of gravity.

feathering and a line perpendicular to (b) Compliance with the flight load the flight path (radians, positive when requirements of this subpart must be axis is pointing aft); shown— W = The angular velocity of rotor (radians (1) At each weight from the design per second); and minimum weight to the design max- R = The rotor radius (ft.).

imum weight; and § 29.341 Gust loads.

(2) With any practical distribution of disposable load within the operating Each rotorcraft must be designed to limitations in the Rotorcraft Flight withstand, at each critical airspeed in- Manual. cluding hovering, the loads resulting Federal Aviation Administration, DOT § 29.395 from vertical and horizontal gusts of 30 (4) The torque imposed by sudden en- feet per second. gine stoppage due to malfunction or structural failure (such as compressor § 29.351 Yawing conditions.

jamming).

(b) For reciprocating engines, the (a) Each rotorcraft must be designed mean torque for maximum continuous for the loads resulting from the maneu- power multiplied by— vers specified in paragraphs (b) and (c) (1) 1.33, for engines with five or more of this section, with— cylinders; and (1) Unbalanced aerodynamic mo- (2) Two, three, and four, for engines ments about the center of gravity with four, three, and two cylinders, re- which the aircraft reacts to in a ration- spectively.

al or conservative manner considering the principal masses furnishing the re- [Amdt. 29–26, 53 FR 34215, Sept. 2, 1988] acting inertia forces; and (2) Maximum main rotor speed. C ONTROL S URFACE AND S YSTEM L OADS (b) To produce the load required in § 29.391 General.

paragraph (a) of this section, in unac- celerated flight with zero yaw, at for- Each auxiliary rotor, each fixed or ward speeds from zero up to 0.6 V — NE movable stabilizing or control surface, (1) Displace the cockpit directional and each system operating any flight control suddenly to the maximum de- control must meet the requirements of flection limited by the control stops or §§ 29.395 through 29.399, 29.411, and by the maximum pilot force specified 29.427.

in § 29.397(a); [Amdt. 29–26, 55 FR 8002, Mar. 6, 1990, as (2) Attain a resulting sideslip angle amended by Amdt. 29–41, 62 FR 46173, Aug. 29, or 90 ° , whichever is less; and 1997] (3) Return the directional control suddenly to neutral.

§ 29.395 Control system.

(c) To produce the load required in (a) The reaction to the loads pre- paragraph (a) of the section, in unac- scribed in § 29.397 must be provided by— celerated flight with zero yaw, at for- (1) The control stops only; ward speeds from 0.6 V up to V or NE NE (2) The control locks only; V , whichever is less— H (3) The irreversible mechanism only (1) Displace the cockpit directional (with the mechanism locked and with control suddenly to the maximum de- the control surface in the critical posi- flection limited by the control stops or tions for the effective parts of the sys- by the maximum pilot force specified tem within its limit of motion); in § 29.397(a); (4) The attachment of the control (2) Attain a resulting sideslip angle system to the rotor blade pitch control or 15 ° , whichever is less, at the lesser horn only (with the control in the crit- speed of V or V ; NE H ical positions for the affected parts of (3) Vary the sideslip angles of para- the system within the limits of its mo- graphs (b)(2) and (c)(2) of this section tion); and directly with speed; and (5) The attachment of the control (4) Return the directional control system to the control surface horn suddenly to neutral.

(with the control in the critical posi- tions for the affected parts of the sys- [Amdt. 29–26, 55 FR 8002, Mar. 6, 1990, as amended by Amdt. 29–41, 62 FR 46173, Aug. 29, tem within the limits of its motion).

1997] (b) Each primary control system, in- cluding its supporting structure, must § 29.361 Engine torque.

be designed as follows: (1) The system must withstand loads The limit engine torque may not be resulting from the limit pilot forces less than the following: prescribed in § 29.397; (a) For turbine engines, the highest (2) Notwithstanding paragraph (b)(3) of— (1) The mean torque for maximum of this section, when power-operated continuous power multiplied by 1.25; actuator controls or power boost con- (2) The torque required by § 29.923; trols are used, the system must also (3) The torque required by § 29.927; or withstand the loads resulting from the 14 CFR Ch. I (1–1–25 Edition) § 29.397 limit pilot forces prescribed in § 29.397 (a) In opposition; and in conjunction with the forces output (b) In the same direction.

of each normally energized power de- vice, including any single power boost § 29.411 Ground clearance: tail rotor guard.

or actuator system failure; (3) If the system design or the normal (a) It must be impossible for the tail operating loads are such that a part of rotor to contact the landing surface the system cannot react to the limit during a normal landing.

pilot forces prescribed in § 29.397, that (b) If a tail rotor guard is required to part of the system must be designed to show compliance with paragraph (a) of withstand the maximum loads that can this section— be obtained in normal operation. The (1) Suitable design loads must be es- minimum design loads must, in any tablished for the guard: and case, provide a rugged system for serv- ice use, including consideration of fa- (2) The guard and its supporting tigue, jamming, ground gusts, control structure must be designed to with- inertia, and friction loads. In the ab- stand those loads.

sence of a rational analysis, the design loads resulting from 0.60 of the speci- § 29.427 Unsymmetrical loads.

fied limit pilot forces are acceptable (a) Horizontal tail surfaces and their minimum design loads; and supporting structure must be designed (4) If operational loads may be ex- for unsymmetrical loads arising from ceeded through jamming, ground gusts, yawing and rotor wake effects in com- control inertia, or friction, the system bination with the prescribed flight con- must withstand the limit pilot forces ditions.

specified in § 29.397, without yielding.

(b) To meet the design criteria of [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as paragraph (a) of this section, in the ab- amended by Amdt. 29–26, 55 FR 8002, Mar. 6, sence of more rational data, both of the 1990] following must be met: § 29.397 Limit pilot forces and torques. (1) One hundred percent of the max- imum loading from the symmetrical (a) Except as provided in paragraph flight conditions acts on the surface on (b) of this section, the limit pilot one side of the plane of symmetry, and forces are as follows: (1) For foot controls, 130 pounds. no loading acts on the other side.

(2) For stick controls, 100 pounds fore (2) Fifty percent of the maximum and aft, and 67 pounds laterally.

loading from the symmetrical flight (b) For flap, tab, stabilizer, rotor conditions acts on the surface on each brake, and landing gear operating con- side of the plane of symmetry, in oppo- trols, the following apply (R = radius in site directions.

inches): (c) For empennage arrangements (1) Crank wheel, and lever controls, [1 where the horizontal tail surfaces are + R]/3 × 50 pounds, but not less than 50 supported by the vertical tail surfaces, pounds nor more than 100 pounds for the vertical tail surfaces and sup- hand operated controls or 130 pounds porting structure must be designed for for foot operated controls, applied at the combined vertical and horizontal any angle within 20 degrees of the surface loads resulting from each pre- plane of motion of the control.

scribed flight condition, considered (2) Twist controls, 80R inch-pounds.

separately. The flight conditions must [Amdt. 29–12, 41 FR 55471, Dec. 20, 1976, as be selected so that the maximum de- amended by Amdt. 29–47, 66 FR 23538, May 9, sign loads are obtained on each surface.

2001] In the absence of more rational data, the unsymmetrical horizontal tail sur- § 29.399 Dual control system.

face loading distributions described in Each dual primary flight control sys- this section must be assumed.

tem must be able to withstand the loads that result when pilot forces not [Amdt. 27–26, 55 FR 8002, Mar. 6, 1990, as less than 0.75 times those obtained amended by Amdt. 29–31, 55 FR 38966, Sept.

under § 29.395 are applied— 21, 1990] Federal Aviation Administration, DOT § 29.481 G ROUND LOADS two wheels aft, and one or more wheels forward, of the center of gravity.

§ 29.471 General.

§ 29.479 Level landing conditions.

(a) Loads and equilibrium. For limit ground loads— (a) Attitudes. Under each of the load- (1) The limit ground loads obtained ing conditions prescribed in paragraph in the landing conditions in this part (b) of this section, the rotorcraft is as- must be considered to be external loads sumed to be in each of the following that would occur in the rotorcraft level landing attitudes: structure if it were acting as a rigid (1) An attitude in which each wheel body; and contacts the ground simultaneously.

(2) In each specified landing condi- (2) An attitude in which the aft tion, the external loads must be placed wheels contact the ground with the for- in equilibrium with linear and angular ward wheels just clear of the ground.

inertia loads in a rational or conserv- (b) Loading conditions. The rotorcraft ative manner.

must be designed for the following (b) Critical centers of gravity. The crit- landing loading conditions: ical centers of gravity within the range (1) Vertical loads applied under for which certification is requested § 29.471.

must be selected so that the maximum (2) The loads resulting from a com- design loads are obtained in each land- bination of the loads applied under ing gear element.

paragraph (b)(1) of this section with drag loads at each wheel of not less § 29.473 Ground loading conditions than 25 percent of the vertical load at and assumptions.

that wheel.

(a) For specified landing conditions, (3) The vertical load at the instant of a design maximum weight must be peak drag load combined with a drag used that is not less than the max- component simulating the forces re- imum weight. A rotor lift may be as- quired to accelerate the wheel rolling sumed to act through the center of assembly up to the specified ground gravity throughout the landing impact.

speed, with— This lift may not exceed two-thirds of (i) The ground speed for determina- the design maximum weight.

tion of the spin-up loads being at least (b) Unless otherwise prescribed, for 75 percent of the optimum forward each specified landing condition, the flight speed for minimum rate of de- rotorcraft must be designed for a limit scent in autorotation; and load factor of not less than the limit (ii) The loading conditions of para- inertia load factor substantiated under graph (b) applied to the landing gear § 29.725.

and its attaching structure only.

(c) Triggering or actuating devices (4) If there are two wheels forward, a for additional or supplementary energy distribution of the loads applied to absorption may not fail under loads es- those wheels under paragraphs (b)(1) tablished in the tests prescribed in and (2) of this section in a ratio of §§ 29.725 and 29.727, but the factor of 40:60.

safety prescribed in § 29.303 need not be (c) Pitching moments. Pitching mo- used.

ments are assumed to be resisted by— [Amdt. 29–3, 33 FR 966, Jan. 26, 1968] (1) In the case of the attitude in para- graph (a)(1) of this section, the forward § 29.475 Tires and shock absorbers.

landing gear; and (2) In the case of the attitude in para- Unless otherwise prescribed, for each graph (a)(2) of this section, the angular specified landing condition, the tires inertia forces.

must be assumed to be in their static position and the shock absorbers to be § 29.481 Tail-down landing conditions.

in their most critical position.

(a) The rotorcraft is assumed to be in § 29.477 Landing gear arrangement.

the maximum nose-up attitude allow- Sections 29.235, 29.479 through 29.485, ing ground clearance by each part of and 29.493 apply to landing gear with the rotorcraft.

14 CFR Ch. I (1–1–25 Edition) § 29.483 (b) In this attitude, ground loads are (b) The structure must be designed to assumed to act perpendicular to the withstand, at the ground contact point ground. of each wheel with brakes, a drag load of at least the lesser of— § 29.483 One-wheel landing conditions.

(1) The vertical load multiplied by a For the one-wheel landing condition, coefficient of friction of 0.8; and the rotorcraft is assumed to be in the (2) The maximum value based on lim- level attitude and to contact the iting brake torque.

ground on one aft wheel. In this atti- tude— § 29.497 Ground loading conditions: (a) The vertical load must be the landing gear with tail wheels.

same as that obtained on that side (a) General. Rotorcraft with landing under § 29.479(b)(1); and gear with two wheels forward and one (b) The unbalanced external loads wheel aft of the center of gravity must must be reacted by rotorcraft inertia.

be designed for loading conditions as prescribed in this section.

§ 29.485 Lateral drift landing condi- tions. (b) Level landing attitude with only the forward wheels contacting the ground. In (a) The rotorcraft is assumed to be in this attitude— the level landing attitude, with— (1) The vertical loads must be applied (1) Side loads combined with one-half under §§ 29.471 through 29.475; of the maximum ground reactions ob- (2) The vertical load at each axle tained in the level landing conditions must be combined with a drag load at of § 29.479(b)(1); and (2) The loads obtained under para- that axle of not less than 25 percent of graph (a)(1) of this section applied— that vertical load; and (i) At the ground contact point; or (3) Unbalanced pitching moments are (ii) For full-swiveling gear, at the assumed to be resisted by angular iner- center of the axle.

tia forces.

(b) The rotorcraft must be designed (c) Level landing attitude with all to withstand, at ground contact— wheels contacting the ground simulta- (1) When only the aft wheels contact neously. In this attitude, the rotorcraft the ground, side loads of 0.8 times the must be designed for landing loading vertical reaction acting inward on one conditions as prescribed in paragraph side and 0.6 times the vertical reaction (b) of this section.

acting outward on the other side, all (d) Maximum nose-up attitude with combined with the vertical loads speci- only the rear wheel contacting the fied in paragraph (a) of this section; ground. The attitude for this condition and must be the maximum nose-up attitude (2) When the wheels contact the expected in normal operation, includ- ground simultaneously— ing autorotative landings. In this atti- (i) For the aft wheels, the side loads tude— specified in paragraph (b)(1) of this sec- (1) The appropriate ground loads tion; and specified in paragraph (b)(1) and (2) of (ii) For the forward wheels, a side this section must be determined and load of 0.8 times the vertical reaction applied, using a rational method to ac- combined with the vertical load speci- count for the moment arm between the fied in paragraph (a) of this section.

rear wheel ground reaction and the § 29.493 Braked roll conditions.

rotorcraft center of gravity; or (2) The probability of landing with Under braked roll conditions with initial contact on the rear wheel must the shock absorbers in their static po- be shown to be extremely remote.

sitions— (e) Level landing attitude with only one (a) The limit vertical load must be based on a load factor of at least— forward wheel contacting the ground. In (1) 1.33, for the attitude specified in this attitude, the rotorcraft must be § 29.479(a)(1); and designed for ground loads as specified (2) 1.0, for the attitude specified in in paragraph (b)(1) and (3) of this sec- § 29.479(a)(2); and tion.

Federal Aviation Administration, DOT § 29.501 (f) Side loads in the level landing atti- must be combined with an equal side tude. In the attitudes specified in para- load.

graphs (b) and (c) of this section, the (2) The load specified in paragraph following apply: (h)(1) of this section must be applied to the rear landing gear— (1) The side loads must be combined (i) Through the axle, if there is a at each wheel with one-half of the max- swivel (the rear wheel being assumed imum vertical ground reactions ob- to be swiveled 90 degrees to the longi- tained for that wheel under paragraphs tudinal axis of the rotorcraft); or (b) and (c) of this section. In this condi- (ii) At the ground contact point if tion, the side loads must be— there is a lock, steering device or shim- (i) For the forward wheels, 0.8 times my damper (the rear wheel being as- the vertical reaction (on one side) act- sumed to be in the trailing position).

ing inward, and 0.6 times the vertical (i) Taxiing condition. The rotorcraft reaction (on the other side) acting out- and its landing gear must be designed ward; and for the loads that would occur when (ii) For the rear wheel, 0.8 times the the rotorcraft is taxied over the rough- vertical reaction.

est ground that may reasonably be ex- (2) The loads specified in paragraph pected in normal operation.

(f)(1) of this section must be applied— (i) At the ground contact point with § 29.501 Ground loading conditions: the wheel in the trailing position (for landing gear with skids.

non-full swiveling landing gear or for (a) General. Rotorcraft with landing full swiveling landing gear with a lock, gear with skids must be designed for steering device, or shimmy damper to the loading conditions specified in this keep the wheel in the trailing posi- section. In showing compliance with tion); or this section, the following apply: (ii) At the center of the axle (for full (1) The design maximum weight, cen- swiveling landing gear without a lock, ter of gravity, and load factor must be steering device, or shimmy damper).

determined under §§ 29.471 through (g) Braked roll conditions in the level 29.475.

landing attitude. In the attitudes speci- (2) Structural yielding of elastic fied in paragraphs (b) and (c) of this spring members under limit loads is ac- section, and with the shock absorbers ceptable.

in their static positions, the rotorcraft (3) Design ultimate loads for elastic must be designed for braked roll loads spring members need not exceed those as follows: obtained in a drop test of the gear (1) The limit vertical load must be with— based on a limit vertical load factor of (i) A drop height of 1.5 times that not less than— specified in § 29.725; and (i) 1.0, for the attitude specified in (ii) An assumed rotor lift of not more paragraph (b) of this section; and than 1.5 times that used in the limit (ii) 1.33, for the attitude specified in drop tests prescribed in § 29.725.

paragraph (c) of this section.

(4) Compliance with paragraph (b) (2) For each wheel with brakes, a through (e) of this section must be drag load must be applied, at the shown with— ground contact point, of not less than (i) The gear in its most critically de- the lesser of— flected position for the landing condi- (i) 0.8 times the vertical load; and tion being considered; and (ii) The maximum based on limiting (ii) The ground reactions rationally brake torque.

distributed along the bottom of the (h) Rear wheel turning loads in the skid tube.

static ground attitude. In the static (b) Vertical reactions in the level land- ground attitude, and with the shock ing attitude. In the level attitude, and absorbers and tires in their static posi- with the rotorcraft contacting the tions, the rotorcraft must be designed ground along the bottom of both skids, for rear wheel turning loads as follows: the vertical reactions must be applied (1) A vertical ground reaction equal as prescribed in paragraph (a) of this to the static load on the rear wheel section.

14 CFR Ch. I (1–1–25 Edition) § 29.505 (c) Drag reactions in the level landing (i) Equal to 1.33 times the maximum attitude. In the level attitude, and with weight; the rotorcraft contacting the ground (ii) Distributed symmetrically among along the bottom of both skids, the fol- the skids; lowing apply: (iii) Concentrated at the forward end of the straight part of the skid tube; (1) The vertical reactions must be and combined with horizontal drag reac- (iv) Applied only to the forward end tions of 50 percent of the vertical reac- of the skid tube and its attachment to tion applied at the ground.

the rotorcraft.

(2) The resultant ground loads must (2) With the rotorcraft in the level equal the vertical load specified in landing attitude, a vertical ground re- paragraph (b) of this section.

action load equal to one-half of the (d) Sideloads in the level landing atti- vertical load determined under para- tude. In the level attitude, and with the graph (b) of this section. This load rotorcraft contacting the ground along must be— the bottom of both skids, the following (i) Applied only to the skid tube and apply: its attachment to the rotorcraft; and (1) The vertical ground reaction must (ii) Distributed equally over 33.3 per- be— cent of the length between the skid (i) Equal to the vertical loads ob- tube attachments and centrally located tained in the condition specified in midway between the skid tube attach- paragraph (b) of this section; and ments.

(ii) Divided equally among the skids.

[Amdt. 29–3, 33 FR 966, Jan. 26, 1968, as (2) The vertical ground reactions amended by Amdt. 27–26, 55 FR 8002, Mar. 6, must be combined with a horizontal 1990] sideload of 25 percent of their value.

(3) The total sideload must be applied § 29.505 Ski landing conditions.

equally between skids and along the If certification for ski operation is length of the skids.

requested, the rotorcraft, with skis, (4) The unbalanced moments are as- must be designed to withstand the fol- sumed to be resisted by angular iner- lowing loading conditions (where P is tia.

the maximum static weight on each ski (5) The skid gear must be inves- with the rotorcraft at design maximum tigated for— weight, and n is the limit load factor (i) Inward acting sideloads; and determined under § 29.473(b)): (ii) Outward acting sideloads.

(a) Up-load conditions in which— (e) One-skid landing loads in the level (1) A vertical load of Pn and a hori- attitude. In the level attitude, and with zontal load of Pn/4 are simultaneously the rotorcraft contacting the ground applied at the pedestal bearings; and along the bottom of one skid only, the (2) A vertical load of 1.33 P is applied following apply: at the pedestal bearings.

(1) The vertical load on the ground (b) A side load condition in which a contact side must be the same as that side load of 0.35 Pn is applied at the obtained on that side in the condition pedestal bearings in a horizontal plane specified in paragraph (b) of this sec- perpendicular to the centerline of the tion.

rotorcraft.

(2) The unbalanced moments are as- (c) A torque-load condition in which sumed to be resisted by angular iner- a torque load of 1.33 P (in foot-pounds) tia.

is applied to the ski about the vertical (f) Special conditions. In addition to axis through the centerline of the ped- the conditions specified in paragraphs estal bearings.

(b) and (c) of this section, the rotor- § 29.511 Ground load: unsymmetrical craft must be designed for the fol- loads on multiple-wheel units.

lowing ground reactions: (1) A ground reaction load acting up (a) In dual-wheel gear units, 60 per- and aft at an angle of 45 degrees to the cent of the total ground reaction for longitudinal axis of the rotorcraft. the gear unit must be applied to one This load must be— wheel and 40 percent to the other.

Federal Aviation Administration, DOT § 29.547 (b) To provide for the case of one de- less it can be shown that full immer- flated tire, 60 percent of the specified sion of the float is unlikely, in which load for the gear unit must be applied case the highest likely float buoyancy to either wheel except that the vertical load must be applied that considers ground reaction may not be less than loading of the float immersed to create the full static value. restoring moments compensating for (c) In determining the total load on a upsetting moments caused by side gear unit, the transverse shift in the wind, asymmetrical rotorcraft loading, load centroid, due to unsymmetrical water wave action, and rotorcraft iner- load distribution on the wheels, may be tia.

neglected.

[Amdt. 29–3, 33 FR 966, Jan. 26, 196, as amend- [Amdt. 29–3, 33 FR 966, Jan. 26, 1968] ed by Amdt. 27–26, 55 FR 8002, Mar. 6, 1990] W ATER L OADS § 29.521 Float landing conditions.

If certification for float operation § 29.519 Hull type rotorcraft: Water- (including float amphibian operation) based and amphibian.

is requested, the rotorcraft, with (a) General. For hull type rotorcraft, floats, must be designed to withstand the structure must be designed to with- the following loading conditions (where stand the water loading set forth in the limit load factor is determined paragraphs (b), (c), and (d) of this sec- under § 29.473(b) or assumed to be equal tion considering the most severe wave to that determined for wheel landing heights and profiles for which approval gear): is desired. The loads for the landing (a) Up-load conditions in which— conditions of paragraphs (b) and (c) of (1) A load is applied so that, with the this section must be developed and dis- rotorcraft in the static level attitude, tributed along and among the hull and the resultant water reaction passes auxiliary floats, if used, in a rational vertically through the center of grav- and conservative manner, assuming a ity; and rotor lift not exceeding two-thirds of (2) The vertical load prescribed in the rotorcraft weight to act through- paragraph (a)(1) of this section is ap- out the landing impact.

plied simultaneously with an aft com- (b) Vertical landing conditions. The ponent of 0.25 times the vertical com- rotorcraft must initially contact the ponent most critical wave surface at zero for- (b) A side load condition in which— ward speed in likely pitch and roll atti- (1) A vertical load of 0.75 times the tudes which result in critical design total vertical load specified in para- loadings. The vertical descent velocity graph (a)(1) of this section is divided may not be less than 6.5 feet per second equally among the floats; and relative to the mean water surface.

(2) For each float, the load share de- (c) Forward speed landing conditions.

termined under paragraph (b)(1) of this The rotorcraft must contact the most section, combined with a total side critical wave at forward velocities load of 0.25 times the total vertical from zero up to 30 knots in likely load specified in paragraph (b)(1) of pitch, roll, and yaw attitudes and with this section, is applied to that float a vertical descent velocity of not less only.

than 6.5 feet per second relative to the mean water surface. A maximum for- [Amdt. 29–3, 33 FR 967, Jan. 26, 1968] ward velocity of less than 30 knots may M AIN C OMPONENT R EQUIREMENTS be used in design if it can be dem- onstrated that the forward velocity se- § 29.547 Main and tail rotor structure.

lected would not be exceeded in a nor- mal one-engine-out landing. (a) A rotor is an assembly of rotating (d) Auxiliary float immersion condition. components, which includes the rotor In addition to the loads from the land- hub, blades, blade dampers, the pitch ing conditions, the auxiliary float, and control mechanisms, and all other its support and attaching structure in parts that rotate with the assembly.

the hull, must be designed for the load (b) Each rotor assembly must be de- developed by a fully immersed float un- signed as prescribed in this section and 14 CFR Ch. I (1–1–25 Edition) § 29.549 must function safely for the critical (3) The loads prescribed in § 29.547 flight load and operating conditions. A (d)(1) and (e)(1)(i).

design assessment must be performed, (b) Auxiliary rotor thrust, the torque including a detailed failure analysis to reaction of each rotor drive system, identify all failures that will prevent and the balancing air and inertia loads continued safe flight or safe landing, occurring under accelerated flight con- and must identify the means to mini- ditions, must be considered.

mize the likelihood of their occurrence.

(c) Each engine mount and adjacent (c) The rotor structure must be de- fuselage structure must be designed to signed to withstand the following loads withstand the loads occurring under prescribed in §§ 29.337 through 29.341 and accelerated flight and landing condi- 29.351: tions, including engine torque.

(1) Critical flight loads.

(d) [Reserved] (2) Limit loads occurring under nor- 1 (e) If approval for the use of 2 ⁄2 - mal conditions of autorotation.

minute OEI power is requested, each (d) The rotor structure must be de- engine mount and adjacent structure signed to withstand loads simulating— must be designed to withstand the (1) For the rotor blades, hubs, and loads resulting from a limit torque flapping hinges, the impact force of equal to 1.25 times the mean torque for each blade against its stop during 1 2 ⁄ 2 -minute OEI power combined with 1g ground operation; and flight loads.

(2) Any other critical condition ex- (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, pected in normal operation.

1425) (e) The rotor structure must be de- signed to withstand the limit torque at [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–4, 33 FR 14106, Sept. 18, any rotational speed, including zero.

1968; Amdt. 29–26, 53 FR 34215, Sept. 2, 1988] In addition: (1) The limit torque need not be § 29.551 Auxiliary lifting surfaces.

greater than the torque defined by a Each auxiliary lifting surface must torque limiting device (where pro- be designed to withstand— vided), and may not be less than the greater of— (a) The critical flight loads in §§ 29.337 (i) The maximum torque likely to be through 29.341, and 29.351; transmitted to the rotor structure, in (b) the applicable ground loads in either direction, by the rotor drive or §§ 29.235, 29.471 through 29.485, 29.493, by sudden application of the rotor 29.505, and 29.521; and brake; and (c) Any other critical condition ex- (ii) For the main rotor, the limit en- pected in normal operation.

gine torque specified in § 29.361.

(2) The limit torque must be equally E MERGENCY L ANDING C ONDITIONS and rationally distributed to the rotor § 29.561 General.

blades.

(a) The rotorcraft, although it may (Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, be damaged in emergency landing con- 1425) ditions on land or water, must be de- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as signed as prescribed in this section to amended by Amdt. 29–4, 33 FR 14106, Sept. 18, protect the occupants under those con- 1968; Amdt. 29–40, 61 FR 21907, May 10, 1996] ditions.

§ 29.549 Fuselage and rotor pylon (b) The structure must be designed to structures.

give each occupant every reasonable chance of escaping serious injury in a (a) Each fuselage and rotor pylon crash landing when— structure must be designed to with- (1) Proper use is made of seats, belts, stand— and other safety design provisions; (1) The critical loads prescribed in (2) The wheels are retracted (where §§ 29.337 through 29.341, and 29.351; applicable); and (2) The applicable ground loads pre- scribed in §§ 29.235, 29.471 through 29.485, (3) Each occupant and each item of 29.493, 29.497, 29.505, and 29.521; and mass inside the cabin that could injure Federal Aviation Administration, DOT § 29.562 an occupant is restrained when sub- and landing must successfully com- jected to the following ultimate iner- plete dynamic tests or be demonstrated tial load factors relative to the sur- by rational analysis based on dynamic rounding structure: tests of a similar type seat in accord- (i) Upward—4g.

ance with the following criteria. The (ii) Forward—16g.

tests must be conducted with an occu- (iii) Sideward—8g.

pant simulated by a 170-pound (iv) Downward—20g, after the in- anthropomorphic test dummy (ATD), tended displacement of the seat device.

as defined by 49 CFR 572, Subpart B, or (v) Rearward—1.5g.

its equivalent, sitting in the normal (c) The supporting structure must be upright position.

designed to restrain under any ulti- (1) A change in downward velocity of mate inertial load factor up to those not less than 30 feet per second when specified in this paragraph, any item of the seat or other seating device is ori- mass above and/or behind the crew and ented in its nominal position with re- passenger compartment that could in- spect to the rotorcraft’s reference sys- jure an occupant if it came loose in an tem, the rotorcraft’s longitudinal axis emergency landing. Items of mass to be is canted upward 60 ° with respect to considered include, but are not limited the impact velocity vector, and the to, rotors, transmission, and engines.

rotorcraft’s lateral axis is perpen- The items of mass must be restrained dicular to a vertical plane containing for the following ultimate inertial load the impact velocity vector and the factors: rotorcraft’s longitudinal axis. Peak (1) Upward—1.5g.

floor deceleration must occur in not (2) Forward—12g.

more than 0.031 seconds after impact (3) Sideward—6g.

and must reach a minimum of 30g’s.

(4) Downward—12g.

(2) A change in forward velocity of (5) Rearward—1.5g.

not less than 42 feet per second when (d) Any fuselage structure in the area the seat or other seating device is ori- of internal fuel tanks below the pas- ented in its nominal position with re- senger floor level must be designed to spect to the rotorcraft’s reference sys- resist the following ultimate inertial tem, the rotorcraft’s longitudinal axis factors and loads, and to protect the is yawed 10 ° either right or left of the fuel tanks from rupture, if rupture is impact velocity vector (whichever likely when those loads are applied to would cause the greatest load on the that area: shoulder harness), the rotorcraft’s lat- (1) Upward—1.5g.

eral axis is contained in a horizontal (2) Forward—4.0g.

plane containing the impact velocity (3) Sideward—2.0g.

vector, and the rotorcraft’s vertical (4) Downward—4.0g.

axis is perpendicular to a horizontal [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as plane containing the impact velocity amended by Amdt. 29–29, 54 FR 47319, Nov. 13, vector. Peak floor deceleration must 1989; Amdt. 29–38, 61 FR 10438, Mar. 13, 1996] occur in not more than 0.071 seconds after impact and must reach a min- § 29.562 Emergency landing dynamic conditions. imum of 18.4g’s.

(3) Where floor rails or floor or side- (a) The rotorcraft, although it may wall attachment devices are used to at- be damaged in a crash landing, must be tach the seating devices to the air- designed to reasonably protect each oc- frame structure for the conditions of cupant when— this section, the rails or devices must (1) The occupant properly uses the be misaligned with respect to each seats, safety belts, and shoulder har- other by at least 10 ° vertically (i.e., nesses provided in the design; and pitch out of parallel) and by at least a (2) The occupant is exposed to loads 10 ° lateral roll, with the directions op- equivalent to those resulting from the tional, to account for possible floor conditions prescribed in this section.

warp.

(b) Each seat type design or other seating device approved for crew or (c) Compliance with the following passenger occupancy during takeoff must be shown: 14 CFR Ch. I (1–1–25 Edition) § 29.563 (1) The seating device system must probable water conditions at forward remain intact although it may experi- velocities from zero up to 30 knots in ence separation intended as part of its likely pitch, roll, and yaw attitudes.

design. The rotorcraft limit vertical descent (2) The attachment between the seat- velocity may not be less than 5 feet per ing device and the airframe structure second relative to the mean water sur- must remain intact although the struc- face. Rotor lift may be used to act ture may have exceeded its limit load. through the center of gravity through- (3) The ATD’s shoulder harness strap out the landing impact. This lift may or straps must remain on or in the im- not exceed two-thirds of the design mediate vicinity of the ATD’s shoulder maximum weight. A maximum forward during the impact. velocity of less than 30 knots may be (4) The safety belt must remain on used in design if it can be dem- the ATD’s pelvis during the impact. onstrated that the forward velocity se- (5) The ATD’s head either does not lected would not be exceeded in a nor- contact any portion of the crew or pas- mal one-engine-out touchdown.

senger compartment or, if contact is (b) Auxiliary or emergency float condi- made, the head impact does not exceed tions —(1) Floats fixed or deployed before a head injury criteria (HIC) of 1,000 as initial water contact. In addition to the determined by this equation. landing loads in paragraph (a) of this section, each auxiliary or emergency 2.5 float, or its support and attaching ⎡ ⎤ t structure in the airframe or fuselage, HIC t t a(t)dt = −

( )

⎢ ⎥ 2 1

∫

t must be designed for the load devel- t t 1 −

( )

⎢ ⎥ 2 1 ⎣ ⎦ oped by a fully immersed float unless it Where: a(t) is the resultant acceleration at can be shown that full immersion is the center of gravity of the head form ex- unlikely. If full immersion is unlikely, pressed as a multiple of g (the accelera- the highest likely float buoyancy load ¥ t is the time tion of gravity) and t 2 1 must be applied. The highest likely duration, in seconds, of major head im- buoyancy load must include consider- pact, not to exceed 0.05 seconds.

ation of a partially immersed float cre- (6) Loads in individual shoulder har- ating restoring moments to com- ness straps must not exceed 1,750 pensate the upsetting moments caused pounds. If dual straps are used for re- by side wind, unsymmetrical rotorcraft taining the upper torso, the total har- loading, water wave action, rotorcraft ness strap loads must not exceed 2,000 inertia, and probable structural dam- pounds.

age and leakage considered under (7) The maximum compressive load § 29.801(d). Maximum roll and pitch an- measured between the pelvis and the gles determined from compliance with lumbar column of the ATD must not § 29.801(d) may be used, if significant, to exceed 1,500 pounds.

determine the extent of immersion of (d) An alternate approach that each float. If the floats are deployed in achieves an equivalent or greater level flight, appropriate air loads derived of occupant protection, as required by from the flight limitations with the this section, must be substantiated on floats deployed shall be used in sub- a rational basis.

stantiation of the floats and their at- tachment to the rotorcraft. For this [Amdt. 29–29, 54 FR 47320, Nov. 13, 1989, as purpose, the design airspeed for limit amended by Amdt. 29–41, 62 FR 46173, Aug. 29, 1997] load is the float deployed airspeed op- erating limit multiplied by 1.11.

§ 29.563 Structural ditching provi- (2) Floats deployed after initial water sions.

contact. Each float must be designed for If certification with ditching provi- full or partial immersion prescribed in sions is requested, structural strength paragraph (b)(1) of this section. In addi- for ditching must meet the require- tion, each float must be designed for ments of this section and § 29.801(e). combined vertical and drag loads using (a) Forward speed landing conditions. a relative limit speed of 20 knots be- The rotorcraft must initially contact tween the rotorcraft and the water.

the most critical wave for reasonably The vertical load may not be less than

Federal Aviation Administration, DOT § 29.571 the highest likely buoyancy load deter- (2) The loading spectra as severe as mined under paragraph (b)(1) of this those expected in operations based on section. loads or stresses determined under paragraph (e)(1) of this section, includ- [Amdt. 27–26, 55 FR 8003, Mar. 6, 1990] ing external load operations, if applica- ble, and other high frequency power- F ATIGUE E VALUATION cycle operations.

(3) Takeoff, landing, and taxi loads § 29.571 Fatigue Tolerance Evaluation when evaluating the landing gear and of Metallic Structure.

other affected PSEs.

(a) A fatigue tolerance evaluation of (4) For each PSE identified in para- each principal structural element graph (d) of this section, a threat as- (PSE) must be performed, and appro- sessment which includes a determina- priate inspections and retirement time tion of the probable locations, types, or approved equivalent means must be and sizes of damage, taking into ac- established to avoid catastrophic fail- count fatigue, environmental effects, ure during the operational life of the intrinsic and discrete flaws, or acci- rotorcraft. The fatigue tolerance eval- dental damage that may occur during uation must consider the effects of manufacture or operation.

both fatigue and the damage deter- (5) A determination of the fatigue mined under paragraph (e)(4) of this tolerance characteristics for the PSE section. Parts to be evaluated include with the damage identified in para- PSEs of the rotors, rotor drive systems graph (e)(4) of this section that sup- between the engines and rotor hubs, ports the inspection and retirement controls, fuselage, fixed and movable times, or other approved equivalent control surfaces, engine and trans- means.

mission mountings, landing gear, and (6) Analyses supported by test evi- their related primary attachments.

dence and, if available, service experi- (b) For the purposes of this section, ence.

the term— (f) A residual strength determination (1) Catastrophic failure means an is required that substantiates the max- event that could prevent continued imum damage size assumed in the fa- safe flight and landing.

tigue tolerance evaluation. In deter- (2) Principal structural element (PSE) mining inspection intervals based on means a structural element that con- damage growth, the residual strength tributes significantly to the carriage of evaluation must show that the remain- flight or ground loads, and the fatigue ing structure, after damage growth, is failure of that structural element could able to withstand design limit loads result in catastrophic failure of the air- without failure.

craft.

(g) The effect of damage on stiffness, (c) The methodology used to estab- dynamic behavior, loads, and func- lish compliance with this section must tional performance must be considered.

be submitted to and approved by the (h) Based on the requirements of this Administrator.

section, inspections and retirement (d) Considering all rotorcraft struc- times or approved equivalent means ture, structural elements, and assem- must be established to avoid cata- blies, each PSE must be identified.

strophic failure. The inspections and (e) Each fatigue tolerance evaluation retirement times or approved equiva- required by this section must include: lent means must be included in the (1) In-flight measurements to deter- Airworthiness Limitations Section of mine the fatigue loads or stresses for the Instructions for Continued Air- the PSEs identified in paragraph (d) of worthiness required by Section 29.1529 this section in all critical conditions and Section A29.4 of Appendix A of this throughout the range of design limita- part.

tions required by § 29.309 (including al- (i) If inspections for any of the dam- titude effects), except that maneu- age types identified in paragraph (e)(4) vering load factors need not exceed the of this section cannot be established maximum values expected in oper- within the limitations of geometry, ations. inspectability, or good design practice, 14 CFR Ch. I (1–1–25 Edition) § 29.573 then supplemental procedures, in con- uations of the strength of composite junction with the PSE retirement PSEs and other parts, detail design time, must be established to minimize points, and fabrication techniques.

the risk of occurrence of these types of Each applicant must account for the damage that could result in a cata- effects of material and process varia- strophic failure during the operational bility along with environmental condi- life of the rotorcraft. tions in the strength and fatigue eval- uations. Each applicant must evaluate [Doc. No. FAA–2009–0413, Amdt. 29–55, 76 FR parts that include PSEs of the air- 75442, Dec. 2, 2011] frame, main and tail rotor drive sys- tems, main and tail rotor blades and § 29.573 Damage Tolerance and Fa- tigue Evaluation of Composite hubs, rotor controls, fixed and movable Rotorcraft Structures.

control surfaces, engine and trans- mission mountings, landing gear, other (a) Each applicant must evaluate the parts, detail design points, and fabrica- composite rotorcraft structure under tion techniques deemed critical by the the damage tolerance standards of FAA. Each damage tolerance evalua- paragraph (d) of this section unless the tion must include: applicant establishes that a damage (i) The identification of all PSEs; tolerance evaluation is impractical (ii) In-flight and ground measure- within the limits of geometry, ments for determining the loads or inspectability, and good design prac- stresses for all PSEs for all critical tice. If an applicant establishes that it conditions throughout the range of is impractical within the limits of ge- limits in § 29.309 (including altitude ef- ometry, inspectability, and good design fects), except that maneuvering load practice, the applicant must do a fa- factors need not exceed the maximum tigue evaluation in accordance with values expected in service; paragraph (e) of this section.

(iii) The loading spectra as severe as (b) The methodology used to estab- those expected in service based on lish compliance with this section must loads or stresses determined under be submitted to and approved by the paragraph (d)(1)(ii) of this section, in- Administrator.

cluding external load operations, if ap- (c) Definitions: plicable, and other operations includ- (1) Catastrophic failure is an event ing high-torque events; that could prevent continued safe (iv) A threat assessment for all PSEs flight and landing.

that specifies the locations, types, and (2) Principal Structural Elements (PSEs) sizes of damage, considering fatigue, are structural elements that con- environmental effects, intrinsic and tribute significantly to the carrying of discrete flaws, and impact or other ac- flight or ground loads, the failure of cidental damage (including the discrete which could result in catastrophic fail- source of the accidental damage) that ure of the rotorcraft.

(3) Threat Assessment is an assessment may occur during manufacture or oper- that specifies the locations, types, and ation; and sizes of damage, considering fatigue, (v) An assessment of the residual environmental effects, intrinsic and strength and fatigue characteristics of discrete flaws, and impact or other ac- all PSEs that supports the replacement cidental damage (including the discrete times and inspection intervals estab- source of the accidental damage) that lished under paragraph (d)(2) of this may occur during manufacture or oper- section.

ation. (2) Each applicant must establish re- (d) Damage Tolerance Evaluation: placement times, inspections, or other (1) Each applicant must show that procedures for all PSEs to require the catastrophic failure due to static and repair or replacement of damaged parts fatigue loads, considering the intrinsic before a catastrophic failure. These re- or discrete manufacturing defects or placement times, inspections, or other accidental damage, is avoided through- procedures must be included in the Air- out the operational life or prescribed worthiness Limitations Section of the inspection intervals of the rotorcraft Instructions for Continued Airworthi- by performing damage tolerance eval- ness required by § 29.1529.

Federal Aviation Administration, DOT § 29.602 (i) Replacement times for PSEs must the limits of geometry, inspectability, be determined by tests, or by analysis or good design practice, the applicant supported by tests, and must show that must do a fatigue evaluation of the the structure is able to withstand the particular composite rotorcraft struc- repeated loads of variable magnitude ture and: expected in-service. In establishing (1) Identify all PSEs considered in these replacement times, the following the fatigue evaluation; items must be considered: (2) Identify the types of damage for (A) Damage identified in the threat all PSEs considered in the fatigue eval- assessment required by paragraph uation; (d)(1)(iv) of this section; (3) Establish supplemental proce- (B) Maximum acceptable manufac- dures to minimize the risk of cata- turing defects and in-service damage strophic failure associated with the ( i.e. , those that do not lower the resid- damages identified in paragraph (d) of ual strength below ultimate design this section; and loads and those that can be repaired to (4) Include these supplemental proce- restore ultimate strength); and dures in the Airworthiness Limitations (C) Ultimate load strength capability section of the Instructions for Contin- after applying repeated loads.

(ii) Inspection intervals for PSEs ued Airworthiness required by § 29.1529.

must be established to reveal any dam- [Doc. No. FAA–2009–0660, Amdt. 29–59, 76 FR age identified in the threat assessment 74664, Dec. 1, 2011] required by paragraph (d)(1)(iv) of this section that may occur from fatigue or Subpart D—Design and other in-service causes before such Construction damage has grown to the extent that the component cannot sustain the re- G ENERAL quired residual strength capability. In establishing these inspection intervals, § 29.601 Design.

the following items must be consid- ered: (a) The rotorcraft may have no de- (A) The growth rate, including no- sign features or details that experience growth, of the damage under the re- has shown to be hazardous or unreli- peated loads expected in-service deter- able.

mined by tests or analysis supported (b) The suitability of each question- by tests; able design detail and part must be es- (B) The required residual strength for tablished by tests.

the assumed damage established after considering the damage type, inspec- § 29.602 Critical parts.

tion interval, detectability of damage, (a) Critical part. A critical part is a and the techniques adopted for damage part, the failure of which could have a detection. The minimum required re- catastrophic effect upon the rotocraft, sidual strength is limit load; and and for which critical characterists (C) Whether the inspection will de- have been identified which must be tect the damage growth before the controlled to ensure the required level minimum residual strength is reached of integrity.

and restored to ultimate load capa- bility, or whether the component will (b) If the type design includes critical require replacement.

parts, a critical parts list shall be es- (3) Each applicant must consider the tablished. Procedures shall be estab- effects of damage on stiffness, dynamic lished to define the critical design behavior, loads, and functional per- characteristics, identify processes that formance on all PSEs when substan- affect those characteristics, and iden- tiating the maximum assumed damage tify the design change and process size and inspection interval.

change controls necessary for showing (e) Fatigue Evaluation: If an appli- compliance with the quality assurance cant establishes that the damage toler- requirements of part 21 of this chapter.

ance evaluation described in paragraph (d) of this section is impractical within [Doc. No. 29311, 64 FR 46232, Aug. 24, 1999] 14 CFR Ch. I (1–1–25 Edition) § 29.603 vice is used in addition to the self-lock- § 29.603 Materials.

ing device.

The suitability and durability of ma- [Amdt. 29–5, 33 FR 14533, Sept. 27, 1968] terials used for parts, the failure of which could adversely affect safety, § 29.609 Protection of structure.

must— Each part of the structure must— (a) Be established on the basis of ex- (a) Be suitably protected against de- perience or tests; terioration or loss of strength in serv- (b) Meet approved specifications that ice due to any cause, including— ensure their having the strength and (1) Weathering; other properties assumed in the design (2) Corrosion; and data; and (3) Abrasion; and (c) Take into account the effects of (b) Have provisions for ventilation environmental conditions, such as tem- and drainage where necessary to pre- perature and humidity, expected in vent the accumulation of corrosive, service.

flammable, or noxious fluids.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- § 29.610 Lightning and static elec- eral Aviation Act of 1958 (49 U.S.C. 1354(a), tricity protection.

1421, 1423, 1424), and sec. 6(c), Dept. of Trans- (a) The rotorcraft structure must be portation Act (49 U.S.C. 1655(c))) protected against catastrophic effects [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as from lightning.

amended by Amdt. 29–12, 41 FR 55471, Dec. 20, (b) For metallic components, compli- 1976; Amdt. 29–17, 43 FR 50599, Oct. 30, 1978] ance with paragraph (a) of this section may be shown by— § 29.605 Fabrication methods.

(1) Electrically bonding the compo- (a) The methods of fabrication used nents properly to the airframe; or must produce consistently sound struc- (2) Designing the components so that tures. If a fabrication process (such as a strike will not endanger the rotor- gluing, spot welding, or heat-treating) craft.

requires close control to reach this ob- (c) For nonmetallic components, jective, the process must be performed compliance with paragraph (a) of this according to an approved process speci- section may be shown by— fication.

(1) Designing the components to min- (b) Each new aircraft fabrication imize the effect of a strike; or method must be substantiated by a (2) Incorporating acceptable means of test program. diverting the resulting electrical cur- rent to not endanger the rotorcraft.

(Secs. 313(a), 601, 603, 604, Federal Aviation (d) The electric bonding and protec- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), tion against lightning and static elec- sec. 6(c), Dept. of Transportation Act (49 tricity must— U.S.C. 1655(c))) (1) Minimize the accumulation of [Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as electrostatic charge; amended by Amdt. 29–17, 43 FR 50599, Oct. 30, (2) Minimize the risk of electric 1978] shock to crew, passengers, and service and maintenance personnel using nor- § 29.607 Fasteners.

mal precautions; (a) Each removable bolt, screw, nut, (3) Provide and electrical return pin, or other fastener whose loss could path, under both normal and fault con- jeopardize the safe operation of the ditions, on rotorcraft having grounded rotorcraft must incorporate two sepa- electrical systems; and rate locking devices. The fastener and (4) Reduce to an acceptable level the its locking devices may not be ad- effects of static electricity on the func- versely affected by the environmental tioning of essential electrical and elec- conditions associated with the par- tronic equipment.

ticular installation.

[Amdt. 29–24, 49 FR 44437, Nov. 6, 1984; Amdt.

(b) No self-locking nut may be used 29–40, 61 FR 21907, May 10, 1996; 61 FR 33963, on any bolt subject to rotation in oper- July 1, 1996; Amdt. 29–53, 76 FR 33135, June 8, ation unless a nonfriction locking de- 2011] Federal Aviation Administration, DOT § 29.621 which a specimen of each individual § 29.611 Inspection provisions.

item is tested before use and it is de- There must be means to allow close termined that the actual strength examination of each part that re- properties of that particular item will quires— equal or exceed those used in design.

(a) Recurring inspection; (b) Adjustment for proper alignment (Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), and functioning; or sec. 6(c), Dept. of Transportation Act (49 (c) Lubrication.

U.S.C. 1655(c))) § 29.613 Material strength properties [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as and design values.

amended by Amdt. 29–17, 43 FR 50599, Oct. 30, 1978; Amdt. 29–30, 55 FR 8003, Mar. 6, 1990] (a) Material strength properties must be based on enough tests of material § 29.619 Special factors.

meeting specifications to establish de- (a) The special factors prescribed in sign values on a statistical basis.

§§ 29.621 through 29.625 apply to each (b) Design values must be chosen to part of the structure whose strength minimize the probability of structural is— failure due to material variability. Ex- (1) Uncertain; cept as provided in paragraphs (d) and (2) Likely to deteriorate in service (e) of this section, compliance with before normal replacement; or this paragraph must be shown by se- (3) Subject to appreciable variability lecting design values that assure mate- due to— rial strength with the following prob- (i) Uncertainties in manufacturing ability— processes; or (1) Where applied loads are eventu- (ii) Uncertainties in inspection meth- ally distributed through a single mem- ods.

ber within an assembly, the failure of (b) For each part of the rotorcraft to which would result in loss of structural which §§ 29.621 through 29.625 apply, the integrity of the component, 99 percent factor of safety prescribed in § 29.303 probability with 95 percent confidence; must be multiplied by a special factor and equal to— (2) For redundant structures, those in (1) The applicable special factors pre- which the failure of individual ele- scribed in §§ 29.621 through 29.625; or ments would result in applied loads (2) Any other factor great enough to being safely distributed to other load- ensure that the probability of the part carrying members, 90 percent prob- being understrength because of the un- ability with 95 percent confidence.

certainties specified in paragraph (a) of (c) The strength, detail design, and this section is extremely remote.

fabrication of the structure must mini- mize the probability of disastrous fa- § 29.621 Casting factors.

tigue failure, particularly at points of (a) General. The factors, tests, and in- stress concentration.

spections specified in paragraphs (b) (d) Design values may be those con- and (c) of this section must be applied tained in the following publications in addition to those necessary to estab- (available from the Naval Publications lish foundry quality control. The in- and Forms Center, 5801 Tabor Avenue, spections must meet approved speci- Philadelphia, PA 19120) or other values fications. Paragraphs (c) and (d) of this approved by the Administrator: section apply to structural castings ex- (1) MIL—HDBK–5, ‘‘Metallic Mate- cept castings that are pressure tested rials and Elements for Flight Vehicle as parts of hydraulic or other fluid sys- Structure’’.

tems and do not support structural (2) MIL—HDBK–17, ‘‘Plastics for loads.

Flight Vehicles’’.

(3) ANC–18, ‘‘Design of Wood Aircraft (b) Bearing stresses and surfaces. The Structures’’. casting factors specified in paragraphs (4) MIL—HDBK–23, ‘‘Composite Con- (c) and (d) of this section— struction for Flight Vehicles’’. (1) Need not exceed 1.25 with respect (e) Other design values may be used if to bearing stresses regardless of the a selection of the material is made in method of inspection used; and 14 CFR Ch. I (1–1–25 Edition) § 29.623 (2) Need not be used with respect to of these properties by test of coupons the bearing surfaces of a part whose cut from the castings on a sampling bearing factor is larger than the appli- basis— cable casting factor. (i) A casting factor of 1.0 may be (c) Critical castings. For each casting used; and whose failure would preclude continued (ii) The castings must be inspected as safe flight and landing of the rotorcraft provided in paragraph (d)(1) of this sec- or result in serious injury to any occu- tion for casting factors of ‘‘1.25 through pant, the following apply: 1.50’’ and tested under paragraph (c)(2) (1) Each critical casting must— of this section.

(i) Have a casting factor of not less [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as than 1.25; and amended by Amdt. 29–41, 62 FR 46173, Aug. 29, (ii) Receive 100 percent inspection by 1997] visual, radiographic, and magnetic par- ticle (for ferromagnetic materials) or § 29.623 Bearing factors.

penetrant (for nonferromagnetic mate- (a) Except as provided in paragraph rials) inspection methods or approved (b) of this section, each part that has equivalent inspection methods.

clearance (free fit), and that is subject (2) For each critical casting with a to pounding or vibration, must have a casting factor less than 1.50, three sam- bearing factor large enough to provide ple castings must be static tested and for the effects of normal relative mo- shown to meet— tion.

(i) The strength requirements of (b) No bearing factor need be used on § 29.305 at an ultimate load cor- a part for which any larger special fac- responding to a casting factor of 1.25; tor is prescribed.

and (ii) The deformation requirements of § 29.625 Fitting factors.

§ 29.305 at a load of 1.15 times the limit For each fitting (part or terminal load.

used to join one structural member to (d) Noncritical castings. For each cast- another) the following apply: ing other than those specified in para- (a) For each fitting whose strength is graph (c) of this section, the following not proven by limit and ultimate load apply: tests in which actual stress conditions (1) Except as provided in paragraphs are simulated in the fitting and sur- (d)(2) and (3) of this section, the casting rounding structures, a fitting factor of factors and corresponding inspections at least 1.15 must be applied to each must meet the following table: part of— Casting factor Inspection (1) The fitting; (2) The means of attachment; and 2.0 or greater ............... 100 percent visual.

(3) The bearing on the joined mem- Less than 2.0, greater 100 percent visual, and magnetic than 1.5. particle (ferromagnetic materials), bers.

penetrant (nonferromagnetic ma- (b) No fitting factor need be used— terials), or approved equivalent (1) For joints made under approved inspection methods.

practices and based on comprehensive 1.25 through 1.50 ........ 100 percent visual, and magnetic particle (ferromagnetic materials), test data (such as continuous joints in penetrant (nonferromagnetic ma- metal plating, welded joints, and scarf terials), and radiographic or ap- joints in wood); and proved equivalent inspection (2) With respect to any bearing sur- methods.

face for which a larger special factor is (2) The percentage of castings in- used.

spected by nonvisual methods may be (c) For each integral fitting, the part reduced below that specified in para- must be treated as a fitting up to the graph (d)(1) of this section when an ap- point at which the section properties proved quality control procedure is es- become typical of the member.

tablished. (d) Each seat, berth, litter, safety (3) For castings procured to a speci- belt, and harness attachment to the fication that guarantees the mechan- structure must be shown by analysis, ical properties of the material in the tests, or both, to be able to withstand casting and provides for demonstration the inertia forces prescribed in Federal Aviation Administration, DOT § 29.672 § 29.561(b)(3) multiplied by a fitting fac- (b) The structural integrity of the tor of 1.33. mass balance installation must be sub- stantiated.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–42, 63 FR 43285, Aug. 12, [Amdt. 29–3, 33 FR 967, Jan. 26, 1968] 1998] § 29.661 Rotor blade clearance.

§ 29.629 Flutter and divergence.

There must be enough clearance be- tween the rotor blades and other parts Each aerodynamic surface of the of the structure to prevent the blades rotorcraft must be free from flutter from striking any part of the structure and divergence under each appropriate during any operating condition.

speed and power condition.

[Amdt. 29–3, 33 FR 967, Jan. 26, 1968] [Doc. No. 28008, 61 FR 21907, May 10, 1996] § 29.663 Ground resonance prevention § 29.631 Bird strike.

means.

The rotorcraft must be designed to (a) The reliability of the means for ensure capability of continued safe preventing ground resonance must be flight and landing (for Category A) or shown either by analysis and tests, or safe landing (for Category B) after im- reliable service experience, or by show- pact with a 2.2-lb (1.0 kg) bird when the ing through analysis or tests that mal- velocity of the rotorcraft (relative to function or failure of a single means the bird along the flight path of the will not cause ground resonance.

rotorcraft) is equal to V or V NE H (b) The probable range of variations, (whichever is the lesser) at altitudes up during service, of the damping action to 8,000 feet. Compliance must be of the ground resonance prevention shown by tests or by analysis based on means must be established and must be tests carried out on sufficiently rep- investigated during the test required resentative structures of similar de- by § 29.241.

sign.

[Amdt. 27–26, 55 FR 8003, Mar. 6, 1990] [Doc. No. 28008, 61 FR 21907, May 10, 1996; 61 FR 33963, July 1, 1996] C ONTROL S YSTEMS § 29.671 General.

R OTORS (a) Each control and control system § 29.653 Pressure venting and drain- must operate with the ease, smooth- age of rotor blades.

ness, and positiveness appropriate to (a) For each rotor blade— its function.

(1) There must be means for venting (b) Each element of each flight con- the internal pressure of the blade; trol system must be designed, or dis- (2) Drainage holes must be provided tinctively and permanently marked, to for the blade; and minimize the probability of any incor- rect assembly that could result in the (3) The blade must be designed to pre- malfunction of the system.

vent water from becoming trapped in (c) A means must be provided to it.

allow full control movement of all pri- (b) Paragraphs (a)(1) and (2) of this mary flight controls prior to flight, or section does not apply to sealed rotor a means must be provided that will blades capable of withstanding the allow the pilot to determine that full maximum pressure differentials ex- control authority is available prior to pected in service.

flight.

[Amdt. 29–3, 33 FR 967, Jan. 26, 1968] [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–24, 49 FR 44437, Nov. 6, § 29.659 Mass balance.

1984] (a) The rotor and blades must be § 29.672 Stability augmentation, auto- mass balanced as necessary to— matic, and power-operated systems.

(1) Prevent excessive vibration; and (2) Prevent flutter at any speed up to If the functioning of stability aug- the maximum forward speed. mentation or other automatic or 14 CFR Ch. I (1–1–25 Edition) § 29.673 power-operated system is necessary to function, failure, or jam of any auxil- show compliance with the flight char- iary interconnected control.

acteristics requirements of this part, [Amdt. 27–26, 55 FR 8003, Mar. 6, 1990] the system must comply with § 29.671 of this part and the following: § 29.675 Stops.

(a) A warning which is clearly distin- (a) Each control system must have guishable to the pilot under expected stops that positively limit the range of flight conditions without requiring the motionof the pilot’s controls.

pilot’s attention must be provided for (b) Each stop must be located in the any failure in the stability augmenta- system so that the range of travel of tion system or in any other automatic its control is not appreciably affected or power-operated system which could by— result in an unsafe condition if the pilot is unaware of the failure. Warning (1) Wear; systems must not activate the control (2) Slackness; or systems.

(3) Takeup adjustments.

(b) The design of the stability aug- (c) Each stop must be able to with- mentation system or of any other auto- stand the loads corresponding to the matic or power-operated system must design conditions for the system.

allow initial counteraction of failures (d) For each main rotor blade— without requiring exceptional pilot (1) Stops that are appropriate to the skill or strength, by overriding the blade design must be provided to limit failure by moving the flight controls in travel of the blade about its hinge the normal sense, and by deactivating points; and the failed system.

(2) There must be means to keep the (c) It must be show that after any blade from hitting the droop stops dur- single failure of the stability aug- ing any operation other than starting mentation system or any other auto- and stopping the rotor.

matic or power-operated system— (Secs. 313(a), 601, 603, 604, Federal Aviation (1) The rotorcraft is safely control- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), lable when the failure or malfunction sec. 6(c), Dept. of Transportation Act (49 occurs at any speed or altitude within U.S.C. 1655(c))) the approved operating limitations; (2) The controllability and maneuver- [Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29–17, 43 FR 50599, Oct. 30, ability requirements of this part are 1978] met within a practical operational flight envelope (for example, speed, al- § 29.679 Control system locks.

titude, normal acceleration, and rotor- craft configurations) which is described If there is a device to lock the con- in the Rotorcraft Flight Manual; and trol system with the rotorcraft on the (3) The trim and stability character- ground or water, there must be means istics are not impaired below a level to— needed to allow continued safe flight (a) Automatically disengage the lock and landing.

when the pilot operates the controls in a normal manner, or limit the oper- [Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] ation of the rotorcraft so as to give un- mistakable warning to the pilot before § 29.673 Primary flight controls.

takeoff; and Primary flight controls are those (b) Prevent the lock from engaging in used by the pilot for immediate control flight.

of pitch, roll, yaw, and vertical motion of the rotorcraft.

§ 29.681 Limit load static tests.

[Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] (a) Compliance with the limit load requirements of this part must be § 29.674 Interconnected controls.

shown by tests in which— Each primary flight control system (1) The direction of the test loads must provide for safe flight and landing produces the most severe loading in the and operate independently after a mal- control system; and Federal Aviation Administration, DOT § 29.695 (2) Each fitting, pulley, and bracket (7) No fairlead may cause a change in used in attaching the system to the cable direction of more than three de- main structure is included; grees.

(b) Compliance must be shown (by (8) No clevis pin subject to load or analyses or individual load tests) with motion and retained only by cotter the special factor requirements for pins may be used in the control sys- control system joints subject to angu- tem.

lar motion. (9) Turnbuckles attached to parts having angular motion must be in- § 29.683 Operation tests.

stalled to prevent binding throughout It must be shown by operation tests the range of travel.

(10) There must be means for visual that, when the controls are operated inspection at each fairlead, pulley, ter- from the pilot compartment with the minal, and turnbuckle.

control system loaded to correspond (e) Control system joints subject to with loads specified for the system, the angular motion must incorporate the system is free from— following special factors with respect (a) Jamming; to the ultimate bearing strength of the (b) Excessive friction; and softest material used as a bearing: (c) Excessive deflection.

(1) 3.33 for push-pull systems other § 29.685 Control system details.

than ball and roller bearing systems.

(2) 2.0 for cable systems.

(a) Each detail of each control sys- (f) For control system joints, the tem must be designed to prevent jam- manufacturer’s static, non-Brinell rat- ming, chafing, and interference from ing of ball and roller bearings may not cargo, passengers, loose objects, or the be exceeded.

freezing of moisture.

(b) There must be means in the cock- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as pit to prevent the entry of foreign ob- amended by Amdt. 29–12, 41 FR 55471, Dec. 20, 1976] jects into places where they would jam the system.

§ 29.687 Spring devices.

(c) There must be means to prevent the slapping of cables or tubes against (a) Each control system spring device other parts. whose failure could cause flutter or (d) Cable systems must be designed other unsafe characteristics must be as follows: reliable.

(1) Cables, cable fittings, turn- (b) Compliance with paragraph (a) of buckles, splices, and pulleys must be of this section must be shown by tests an acceptable kind.

simulating service conditions.

(2) The design of cable systems must § 29.691 Autorotation control mecha- prevent any hazardous change in cable nism.

tension throughout the range of travel under any operating conditions and Each main rotor blade pitch control temperature variations.

mechanism must allow rapid entry into (3) No cable smaller than ⁄ 8 inch di- autorotation after power failure.

ameter may be used in any primary § 29.695 Power boost and power-oper- control system.

ated control system.

(4) Pulley kinds and sizes must cor- respond to the cables with which they (a) If a power boost or power-oper- are used. The pulley-cable combina- ated control system is used, an alter- tions and strength values specified in nate system must be immediately MIL-HDBK-5 must be used unless they available that allows continued safe are inapplicable. flight and landing in the event of— (5) Pulleys must have close fitting (1) Any single failure in the power guards to prevent the cables from being portion of the system; or displaced or fouled. (2) The failure of all engines.

(6) Pulleys must lie close enough to (b) Each alternate system may be a the plane passing through the cable to duplicate power portion or a manually prevent the cable from rubbing against operated mechanical system. The the pulley flange. power portion includes the power 14 CFR Ch. I (1–1–25 Edition) § 29.723 action that would exist at the nose source (such as hydrualic pumps), and wheel, assuming that the mass of the such items as valves, lines, and actu- rotorcraft acts at the center of gravity ators.

and exerts a force of 1.0 g downward and (c) The failure of mechanical parts 0.25 g forward.

(such as piston rods and links), and the W = W for tailwheel units (lbs.) equal to t jamming of power cylinders, must be whichever of the following is critical— considered unless they are extremely (1) The static weight on the tailwheel with the rotorcraft resting on all wheels; or improbable.

(2) The vertical component of the ground L ANDING G EAR reaction that would occur at the tailwheel assuming that the mass of the rotorcraft § 29.723 Shock absorption tests. acts at the center of gravity and exerts a force of 1 g downward with the rotorcraft in The landing inertia load factor and the maximum nose-up attitude considered in the reserve energy absorption capacity the nose-up landing conditions.

of the landing gear must be substan- h = specified free drop height (inches).

tiated by the tests prescribed in L = ratio of assumed rotor lift to the rotor- §§ 29.725 and 29.727, respectively. These craft weight.

tests must be conducted on the com- d = deflection under impact of the tire (at plete rotorcraft or on units consisting the proper inflation pressure) plus the vertical component of the axle travel of wheel, tire, and shock absorber in (inches) relative to the drop mass.

their proper relation.

n = limit inertia load factor.

n = the load factor developed, during impact, j § 29.725 Limit drop test.

on the mass used in the drop test (i.e., The limit drop test must be con- the acceleration dv/dt in g ’s recorded in ducted as follows: the drop test plus 1.0).

(a) The drop height must be at least [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as 8 inches.

amended by Amdt. 29–3, 33 FR 967, Jan. 26, (b) If considered, the rotor lift speci- 1968] fied in § 29.473(a) must be introduced into the drop test by appropriate en- § 29.727 Reserve energy absorption drop test.

ergy absorbing devices or by the use of an effective mass.

The reserve energy absorption drop (c) Each landing gear unit must be test must be conducted as follows: tested in the attitude simulating the (a) The drop height must be 1.5 times landing condition that is most critical that specified in § 29.725(a).

from the standpoint of the energy to be (b) Rotor lift, where considered in a absorbed by it.

manner similar to that prescribed in (d) When an effective mass is used in § 29.725(b), may not exceed 1.5 times the showing compliance with paragraph (b) lift allowed under that paragraph.

of this section, the following formulae (c) The landing gear must withstand may be used instead of more rational this test without collapsing. Collapse computations. of the landing gear occurs when a member of the nose, tail, or main gear h d + − ( ) 1 L will not support the rotorcraft in the W W = × ; and e h d + proper attitude or allows the rotorcraft structure, other than landing gear and W e n n L = + j external accessories, to impact the W landing surface.

where: W = the effective weight to be used in the [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as e drop test (lbs.). amended by Amdt. 27–26, 55 FR 8003, Mar. 6, W = W for main gear units (lbs.), equal to 1990] M the static reaction on the particular unit § 29.729 Retracting mechanism.

with the rotorcraft in the most critical attitude. A rational method may be used For rotorcraft with retractable land- in computing a main gear static reac- ing gear, the following apply: tion, taking into consideration the mo- (a) Loads. The landing gear, retract- ment arm between the main wheel reac- ing mechanism, wheel well doors, and tion and the rotorcraft center of gravity.

supporting structure must be designed W = W for nose gear units (lbs.), equal to N the vertical component of the static re- for— Federal Aviation Administration, DOT § 29.735 (1) The loads occurring in any ma- (1) Maximum weight; and neuvering condition with the gear re- (2) Critical center of gravity.

tracted; (c) The maximum limit load rating of (2) The combined friction, inertia, each wheel must equal or exceed the and air loads occurring during retrac- maximum radial limit load determined tion and extension at any airspeed up under the applicable ground load re- to the design maximum landing gear quirements of this part.

operating speed; and (3) The flight loads, including those § 29.733 Tires.

in yawed flight, occurring with the Each landing gear wheel must have a gear extended at any airspeed up to the tire— design maximum landing gear extended (a) That is a proper fit on the rim of speed.

the wheel; and (b) Landing gear lock. A positive means must be provided to keep the (b) Of a rating that is not exceeded gear extended. under— (c) Emergency operation. When other (1) The design maximum weight; than manual power is used to operate (2) A load on each main wheel tire the gear, emergency means must be equal to the static ground reaction cor- provided for extending the gear in the responding to the critical center of event of— gravity; and (1) Any reasonably probable failure in (3) A load on nose wheel tires (to be the normal retraction system; or compared with the dynamic rating es- (2) The failure of any single source of tablished for those tires) equal to the hydraulic, electric, or equivalent en- reaction obtained at the nose wheel, ergy.

assuming that the mass of the rotor- (d) Operation tests. The proper func- craft acts as the most critical center of tioning of the retracting mechanism gravity and exerts a force of 1.0 g down- must be shown by operation tests.

ward and 0.25 g forward, the reactions (e) Position indicator. There must be being distributed to the nose and main means to indicate to the pilot when the wheels according to the principles of gear is secured in the extreme posi- statics with the drag reaction at the tions.

ground applied only at wheels with (f) Control. The location and oper- brakes.

ation of the retraction control must meet the requirements of §§ 29.777 and (c) Each tire installed on a retract- 29.779. able landing gear system must, at the (g) Landing gear warning. An aural or maximum size of the tire type expected equally effective landing gear warning in service, have a clearance to sur- device must be provided that functions rounding structure and systems that is continuously when the rotorcraft is in adequate to prevent contact between a normal landing mode and the landing the tire and any part of the structure gear is not fully extended and locked.

or systems.

A manual shutoff capability must be [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as provided for the warning device and the amended by Amdt. 29–12, 41 FR 55471, Dec. 20, warning system must automatically 1976] reset when the rotorcraft is no longer in the landing mode.

§ 29.735 Brakes.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as For rotorcraft with wheel-type land- amended by Amdt. 29–24, 49 FR 44437, Nov. 6, ing gear, a braking device must be in- 1984] stalled that is— § 29.731 Wheels. (a) Controllable by the pilot; (b) Usable during power-off landings; (a) Each landing gear wheel must be and approved.

(c) Adequate to— (b) The maximum static load rating of each wheel may not be less than the (1) Counteract any normal unbal- corresponding static ground reaction anced torque when starting or stopping with— the rotor; and 14 CFR Ch. I (1–1–25 Edition) § 29.737 (2) Hold the rotorcraft parked on a § 29.755 Hull buoyancy.

10-degree slope on a dry, smooth pave- Water-based and amphibian rotorcraft.

ment.

The hull and auxiliary floats, if used, [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as must have enough watertight compart- amended by Amdt. 29–24, 49 FR 44437, Nov. 6, ments so that, with any single com- 1984] partment of the hull or auxiliary floats flooded, the buoyancy of the hull and § 29.737 Skis.

auxiliary floats, and wheel tires if (a) The maximum limit load rating of used, provides a margin of positive each ski must equal or exceed the max- water stability great enough to mini- imum limit load determined under the mize the probability of capsizing the applicable ground load requirements of rotorcraft for the worst combination of this part.

wave heights and surface winds for (b) There must be a stabilizing means which approval is desired.

to maintain the ski in an appropriate [Amdt. 29–3, 33 FR 967, Jan. 26, 1968, as position during flight. This means amended by Amdt. 27–26, 55 FR 8003, Mar. 6, must have enough strength to with- 1990] stand the maximum aerodynamic and inertia loads on the ski.

§ 29.757 Hull and auxiliary float strength.

F LOATS AND H ULLS The hull, and auxiliary floats if used, § 29.751 Main float buoyancy.

must withstand the water loads pre- scribed by § 29.519 with a rational and (a) For main floats, the buoyancy conservative distribution of local and necessary to support the maximum weight of the rotorcraft in fresh water distributed water pressures over the must be exceeded by— hull and float bottom.

(1) 50 percent, for single floats; and [Amdt. 29–3, 33 FR 967, Jan. 26, 1968] (2) 60 percent, for multiple floats.

(b) Each main float must have P ERSONNEL AND C ARGO enough water-tight compartments so A CCOMMODATIONS that, with any single main float com- partment flooded, the mainfloats will § 29.771 Pilot compartment.

provide a margin of positive stability For each pilot compartment— great enough to minimize the prob- (a) The compartment and its equip- ability of capsizing.

ment must allow each pilot to perform [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as his duties without unreasonable con- amended by Amdt. 29–3, 33 FR 967, Jan. 26, centration or fatigue; 1968] (b) If there is provision for a second pilot, the rotorcraft must be control- § 29.753 Main float design.

lable with equal safety from either (a) Bag floats. Each bag float must be pilot position. Flight and powerplant designed to withstand— controls must be designed to prevent (1) The maximum pressure differen- confusion or inadvertent operation tial that might be developed at the when the rotorcraft is piloted from ei- maximum altitude for which certifi- ther position; cation with that float is requested; and (c) The vibration and noise charac- (2) The vertical loads prescribed in teristics of cockpit appurtenances may § 29.521(a), distributed along the length not interfere with safe operation; of the bag over three-quarters of its (d) Inflight leakage of rain or snow projected area.

that could distract the crew or harm (b) Rigid floats. Each rigid float must the structure must be prevented.

be able to withstand the vertical, hori- zontal, and side loads prescribed in [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as § 29.521. An appropriate load distribu- amended by Amdt. 29–3, 33 FR 967, Jan. 26, tion under critical conditions must be 1968; Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] used.

Federal Aviation Administration, DOT § 29.779 erence cue, that imagery and sym- § 29.773 Pilot compartment view.

bology must be aligned with, and (a) Nonprecipitation conditions. For scaled to, the external scene.

nonprecipitation conditions, the fol- (3) The vision system must provide a lowing apply: means to allow the pilot using the dis- (1) Each pilot compartment must be play to immediately deactivate and re- arranged to give the pilots a suffi- activate the vision system imagery, on ciently extensive, clear, and undis- torted view for safe operation. demand, without removing the pilot’s (2) Each pilot compartment must be hands from the primary flight and free of glare and reflection that could power controls, or their equivalent.

interfere with the pilot’s view. If cer- (4) When the vision system is not in tification for night operation is re- operation it must permit the pilot quested, this must be shown by ground compartment to satisfy the require- or night flight tests.

ments of paragraphs (a) and (b) of this (b) Precipitation conditions. For pre- section.

cipitation conditions, the following [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as apply: amended by Amdt. 29–3, 33 FR 967, Jan. 26, (1) Each pilot must have a suffi- 1968; Docket FAA–2013–0485, Amdt. 29–56, 81 ciently extensive view for safe oper- FR 90170, Dec. 13, 2016; Docket FAA–2016–9275, ation— Amdt. 29–57, 83 FR 9423, Mar. 6, 2018] (i) In heavy rain at forward speeds up to V ; and H § 29.775 Windshields and windows.

(ii) In the most severe icing condi- tion for which certification is re- Windshields and windows must be quested. made of material that will not break (2) The first pilot must have a win- into dangerous fragments.

dow that— [Amdt. 29–31, 55 FR 38966, Sept. 21, 1990] (i) Is openable under the conditions prescribed in paragraph (b)(1) of this § 29.777 Cockpit controls.

section; and Cockpit controls must be— (ii) Provides the view prescribed in that paragraph. (a) Located to provide convenient op- (c) Vision systems with transparent dis- eration and to prevent confusion and plays. A vision system with a trans- inadvertent operation; and parent display surface located in the (b) Located and arranged with re- pilot’s outside field of view, such as a spect to the pilot’s seats so that there head up-display, head mounted display, is full and unrestricted movement of or other equivalent display, must meet each control without interference from the following requirements in non- the cockpit structure or the pilot’s precipitation and precipitation condi- clothing when pilots from 5 ′ 2 ″ to 6 ′ 0 ″ in tions: height are seated.

(1) While the vision system display is in operation, it must compensate for § 29.779 Motion and effect of cockpit interference with the pilot’s outside controls.

field of view such that the combination Cockpit controls must be designed so of what is visible in the display and that they operate in accordance with what remains visible through and the following movements and actu- around it, allows the pilot compart- ation: ment to satisfy the requirements of (a) Flight controls, including the col- paragraphs (a) and (b) of this section.

lective pitch control, must operate (2) The pilot’s view of the external with a sense of motion which cor- scene may not be distorted by the responds to the effect on the rotor- transparent display surface or by the craft.

vision system imagery. When the vi- (b) Twist-grip engine power controls sion system displays imagery or any must be designed so that, for lefthand symbology that is referenced to the im- agery and outside scene topography, operation, the motion of the pilot’s including attitude symbology, flight hand is clockwise to increase power path vector, and flight path angle ref- when the hand is viewed from the edge 14 CFR Ch. I (1–1–25 Edition) § 29.783 containing the index finger. Other en- ing mechanism fails. If the door does gine power controls, excluding the col- not meet the requirements of para- lective control, must operate with a graph (c) of this section with this de- forward motion to increase power. vice in place, suitable operating proce- (c) Normal landing gear controls dures must be established to prevent must operate downward to extend the the use of the device during takeoff and landing gear.

landing.

(g) If an integral stair is installed in [Amdt. 29–24, 49 FR 44437, Nov. 6, 1984] a passenger entry door that is qualified as a passenger emergency exit, the § 29.783 Doors.

stair must be designed so that under (a) Each closed cabin must have at the following conditions the effective- least one adequate and easily acces- ness of passenger emergency egress will sible external door.

not be impaired: (b) Each external door must be lo- (1) The door, integral stair, and oper- cated, and appropriate operating proce- ating mechanism have been subjected dures must be established, to ensure to the inertial forces specified in para- that persons using the door will not be graph (d) of this section, acting sepa- endangered by the rotors, propellers, rately relative to the surrounding engine intakes, and exhausts when the structure.

operating procedures are used.

(2) The rotorcraft is in the normal (c) There must be means for locking ground attitude and in each of the atti- crew and external passenger doors and tudes corresponding to collapse of one for preventing their opening in flight or more legs, or primary members, as inadvertently or as a result of mechan- applicable, of the landing gear.

ical failure. It must be possible to open (h) Nonjettisonable doors used as external doors from inside and outside ditching emergency exits must have the cabin with the rotorcraft on the means to enable them to be secured in ground even though persons may be the open position and remain secure for crowded against the door on the inside emergency egress in sea state condi- of the rotorcraft. The means of opening tions prescribed for ditching.

must be simple and obvious and so ar- ranged and marked that it can be read- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ily located and operated.

amended by Amdt. 29–20, 45 FR 60178, Sept.

(d) There must be reasonable provi- 11, 1980; Amdt. 29–29, 54 FR 47320, Nov. 13, sions to prevent the jamming of any 1989; Amdt. 27–26, 55 FR 8003, Mar. 6, 1990; Amdt. 29–31, 55 FR 38966, Sept. 21, 1990] external doors in a minor crash as a re- sult of fuselage deformation under the § 29.785 Seats, berths, litters, safety following ultimate inertial forces ex- belts, and harnesses.

cept for cargo or service doors not suit- (a) Each seat, safety belt, harness, able for use as an exit in an emergency: and adjacent part of the rotorcraft at (1) Upward—1.5g.

each station designated for occupancy (2) Forward—4.0g.

during takeoff and landing must be free (3) Sideward—2.0g.

(4) Downward—4.0g. of potentially injurious objects, sharp (e) There must be means for direct edges, protuberances, and hard surfaces visual inspection of the locking mecha- and must be designed so that a person nism by crewmembers to determine making proper use of these facilities whether the external doors (including will not suffer serious injury in an passenger, crew, service, and cargo emergency landing as a result of the doors) are fully locked. There must be inertial factors specified in § 29.561(b) visual means to signal to appropriate and dynamic conditions specified in crewmembers when normally used ex- § 29.562.

ternal doors are closed and fully (b) Each occupant must be protected locked. from serious head injury by a safety (f) For outward opening external belt plus a shoulder harness that will doors usable for entrance or egress, prevent the head from contacting any there must be an auxiliary safety injurious object, except as provided for latching device to prevent the door in § 29.562(c)(5). A shoulder harness from opening when the primary latch- (upper torso restraint), in combination Federal Aviation Administration, DOT § 29.785 with the safety belt, constitutes a is capable of being used without the torso restraint system as described in shoulder harness, the inertial forces TSO-C114.

specified must be met by the safety (c) Each occupant’s seat must have a belt alone.

combined safety belt and shoulder har- (h) When a headrest is used, the head- ness with a single-point release. Each rest and its supporting structure must pilot’s combined safety belt and shoul- be designed to resist the inertia forces der harness must allow each pilot when specified in § 29.561, with a 1.33 fitting seated with safety belt and shoulder factor and a head weight of at least 13 harness fastened to perform all func- pounds.

tions necessary for flight operations.

(i) Each seating device system in- There must be a means to secure belt cludes the device such as the seat, the and harness when not in use to prevent cushions, the occupant restraint sys- interference with the operation of the tem and attachment devices.

rotorcraft and with rapid egress in an (j) Each seating device system may emergency.

use design features such as crushing or (d) If seat backs do not have a firm separation of certain parts of the seat handhold, there must be hand grips or in the design to reduce occupant loads rails along each aisle to let the occu- for the emergency landing dynamic pants steady themselves while using the aisle in moderately rough air. conditions of § 29.562; otherwise, the (e) Each projecting object that would system must remain intact and must injure persons seated or moving about not interfere with rapid evacuation of in the rotorcraft in normal flight must the rotorcraft.

be padded.

(k) For purposes of this section, a lit- (f) Each seat and its supporting ter is defined as a device designed to structure must be designed for an occu- carry a nonambulatory person, pri- pant weight of at least 170 pounds, con- marily in a recumbent position, into sidering the maximum load factors, in- and on the rotorcraft. Each berth or ertial forces, and reactions between the litter must be designed to withstand occupant, seat, and safety belt or har- the load reaction of an occupant ness corresponding with the applicable weight of at least 170 pounds when the flight and ground-load conditions, in- occupant is subjected to the forward cluding the emergency landing condi- inertial factors specified in § 29.561(b).

tions of § 29.561(b). In addition— A berth or litter installed within 15 ° or (1) Each pilot seat must be designed less of the longitudinal axis of the for the reactions resulting from the ap- rotorcraft must be provided with a pad- plication of the pilot forces prescribed ded end-board, cloth diaphragm, or in § 29.397; and equivalent means that can withstand (2) The inertial forces prescribed in the forward load reaction. A berth or § 29.561(b) must be multiplied by a fac- litter oriented greater than 15 ° with tor of 1.33 in determining the strength of the attachment of— the longitudinal axis of the rotorcraft (i) Each seat to the structure; and must be equipped with appropriate re- (ii) Each safety belt or harness to the straints, such as straps or safety belts, seat or structure.

to withstand the forward reaction. In (g) When the safety belt and shoulder addition— harness are combined, the rated (1) The berth or litter must have a re- strength of the safety belt and shoulder straint system and must not have cor- harness may not be less than that cor- ners or other protuberances likely to responding to the inertial forces speci- cause serious injury to a person occu- fied in § 29.561(b), considering the occu- pying it during emergency landing con- pant weight of at least 170 pounds, con- ditions; and sidering the dimensional characteris- (2) The berth or litter attachment tics of the restraint system installa- and the occupant restraint system at- tion, and using a distribution of at tachments to the structure must be de- least a 60-percent load to the safety signed to withstand the critical loads belt and at least a 40-percent load to resulting from flight and ground load the shoulder harness. If the safety belt 14 CFR Ch. I (1–1–25 Edition) § 29.787 conditions and from the conditions pre- to minimize the probability that in an scribed in § 29.561(b). The fitting factor emergency landing on water, the be- required by § 29.625(d) shall be applied. havior of the rotorcraft would cause immediate injury to the occupants or [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as would make it impossible for them to amended by Amdt. 29–24, 49 FR 44437, Nov. 6, escape.

1984; Amdt. 29–29, 54 FR 47320, Nov. 13, 1989; (c) The probable behavior of the Amdt. 29–42, 63 FR 43285, Aug. 12, 1998] rotorcraft in a water landing must be § 29.787 Cargo and baggage compart- investigated by model tests or by com- ments.

parison with rotorcraft of similar con- figuration for which the ditching char- (a) Each cargo and baggage compart- acteristics are known. Scoops, flaps, ment must be designed for its plac- projections, and any other factors like- arded maximum weight of contents and for the critical load distributions at ly to affect the hydrodynamic charac- teristics of the rotorcraft must be con- the appropriate maximum load factors sidered.

corresponding to the specified flight (d) It must be shown that, under rea- and ground load conditions, except the sonably probable water conditions, the emergency landing conditions of flotation time and trim of the rotor- § 29.561.

craft will allow the occupants to leave (b) There must be means to prevent the rotorcraft and enter the liferafts the contents of any compartment from required by § 29.1415. If compliance with becoming a hazard by shifting under this provision is shown by bouyancy the loads specified in paragraph (a) of and trim computations, appropriate al- this section.

lowances must be made for probable (c) Under the emergency landing con- structural damage and leakage. If the ditions of § 29.561, cargo and baggage rotorcraft has fuel tanks (with fuel jet- compartments must— tisoning provisions) that can reason- (1) Be positioned so that if the con- ably be expected to withstand a ditch- tents break loose they are unlikely to ing without leakage, the jettisonable cause injury to the occupants or re- volume of fuel may be considered as strict any of the escape facilities pro- bouyancy volume.

vided for use after an emergency land- (e) Unless the effects of the collapse ing; or of external doors and windows are ac- (2) Have sufficient strength to with- counted for in the investigation of the stand the conditions specified in probable behavior of the rotorcraft in a § 29.561, including the means of re- water landing (as prescribed in para- straint and their attachments required graphs (c) and (d) of this section), the by paragraph (b) of this section. Suffi- external doors and windows must be cient strength must be provided for the designed to withstand the probable maximum authorized weight of cargo maximum local pressures.

and baggage at the critical loading dis- tribution.

[Amdt. 29–12, 41 FR 55472, Dec. 20, 1976] (d) If cargo compartment lamps are installed, each lamp must be installed § 29.803 Emergency evacuation.

so as to prevent contact between lamp (a) Each crew and passenger area bulb and cargo.

must have means for rapid evacuation in a crash landing, with the landing [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55472, Dec. 20, gear (1) extended and (2) retracted, con- 1976; Amdt. 29–31, 55 FR 38966, Sept. 21, 1990] sidering the possibility of fire.

(b) Passenger entrance, crew, and § 29.801 Ditching.

service doors may be considered as (a) If certification with ditching pro- emergency exits if they meet the re- visions is requested, the rotorcraft quirements of this section and of must meet the requirements of this §§ 29.805 through 29.815.

section and §§ 29.807(d), 29.1411 and (c) [Reserved] 29.1415. (d) Except as provided in paragraph (b) Each practicable design measure, (e) of this section, the following cat- compatible with the general character- egories of rotorcraft must be tested in istics of the rotorcraft, must be taken accordance with the requirements of

Federal Aviation Administration, DOT § 29.807 appendix D of this part to demonstrate (1) Type I. This type must have a rec- that the maximum seating capacity, tangular opening of not less than 24 including the crewmembers required by inches wide by 48 inches high, with cor- the operating rules, can be evacuated ner radii not greater than one-third the from the rotorcraft to the ground with- width of the exit, in the passenger area in 90 seconds: in the side of the fuselage at floor level (1) Rotorcraft with a seating capacity and as far away as practicable from of more than 44 passengers. areas that might become potential fire hazards in a crash.

(2) Rotorcraft with all of the fol- (2) Type II. This type is the same as lowing: Type I, except that the opening must (i) Ten or more passengers per pas- be at least 20 inches wide by 44 inches senger exit as determined under high.

§ 29.807(b).

(3) Type III. This type is the same as (ii) No main aisle, as described in Type I, except that— § 29.815, for each row of passenger seats.

(i) The opening must be at least 20 (iii) Access to each passenger exit for inches wide by 36 inches high; and each passenger by virtue of design fea- (ii) The exits need not be at floor tures of seats, such as folding or break- level.

over seat backs or folding seats.

(4) Type IV. This type must have a (e) A combination of analysis and rectangular opening of not less than 19 tests may be used to show that the inches wide by 26 inches high, with cor- rotorcraft is capable of being evacu- ner radii not greater than one-third the ated within 90 seconds under the condi- width of the exit, in the side of the fu- tions specified in § 29.803(d) if the Ad- selage with a step-up inside the rotor- ministrator finds that the combination craft of not more than 29 inches.

of analysis and tests will provide data, with respect to the emergency evacu- Openings with dimensions larger than ation capability of the rotorcraft, those specified in this section may be equivalent to that which would be ob- used, regardless of shape, if the base of tained by actual demonstration. the opening has a flat surface of not less than the specified width.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (b) Passenger emergency exits; side-of- amended by Amdt. 29–3, 33 FR 967, Jan. 26, fuselage. Emergency exits must be ac- 1968; Amdt. 27–26, 55 FR 8004, Mar. 6, 1990] cessible to the passengers and, except as provided in paragraph (d) of this sec- § 29.805 Flight crew emergency exits.

tion, must be provided in accordance (a) For rotorcraft with passenger with the following table: emergency exits that are not conven- ient to the flight crew, there must be Emergency exits for each Passenger seating side of the fuselage flight crew emergency exits, on both capacity sides of the rotorcraft or as a top Type I Type II Type III Type IV hatch, in the flight crew area.

(b) Each flight crew emergency exit 11 through 19 .......... ............ ............ 1 or 2 located so as to allow rapid evacuation 60 through 79 .......... 1 ............ 1 or 2 of the flight crew. This must be shown by test.

(c) Passenger emergency exits; other (c) Each exit must not be obstructed than side-of-fuselage. In addition to the by water or flotation devices after a requirements of paragraph (b) of this ditching. This must be shown by test, section— demonstration, or analysis.

(1) There must be enough openings in the top, bottom, or ends of the fuselage [Amdt. 29–3, 33 FR 968, Jan. 26, 1968, as amended by Amdt. 27–26, 55 FR 8004, Mar. 6, to allow evacuation with the rotorcraft 1990] on its side; or (2) The probability of the rotorcraft § 29.807 Passenger emergency exits.

coming to rest on its side in a crash (a) Type. For the purpose of this part, landing must be extremely remote.

the types of passenger emergency exit (d) Ditching emergency exits for pas- are as follows: sengers. If certification with ditching 14 CFR Ch. I (1–1–25 Edition) § 29.809 provisions is requested, ditching emer- (b) Each emergency exit must be gency exits must be provided in accord- openable from the inside and from the ance with the following requirements outside.

and must be proven by test, demonstra- (c) The means of opening each emer- tion, or analysis unless the emergency gency exit must be simple and obvious exits required by paragraph (b) of this and may not require exceptional effort.

section already meet these require- (d) There must be means for locking ments.

each emergency exit and for preventing (1) For rotorcraft that have a pas- opening in flight inadvertently or as a senger seating configuration, excluding result of mechanical failure.

pilots seats, of nine seats or less, one (e) There must be means to minimize exit above the waterline in each side of the probability of the jamming of any the rotorcraft, meeting at least the di- emergency exit in a minor crash land- mensions of a Type IV exit.

ing as a result of fuselage deformation (2) For rotorcraft that have a pas- under the ultimate inertial forces in senger seating configuration, excluding § 29.783(d).

pilots seats, of 10 seats or more, one (f) Except as provided in paragraph exit above the waterline in a side of the (h) of this section, each land-based rotorcraft meeting at least the dimen- rotorcraft emergency exit must have sions of a Type III exit, for each unit an approved slide as stated in para- (or part of a unit) of 35 passenger seats, graph (g) of this section, or its equiva- but no less than two such exits in the lent, to assist occupants in descending passenger cabin, with one on each side to the ground from each floor level exit of the rotorcraft. However, where it and an approved rope, or its equivalent, has been shown through analysis, for all other exits, if the exit threshold ditching demonstrations, or any other is more that 6 feet above the ground— tests found necessary by the Adminis- (1) With the rotorcraft on the ground trator, that the evacuation capability and with the landing gear extended; of the rotorcraft during ditching is im- (2) With one or more legs or part of proved by the use of larger exits, or by the landing gear collapsed, broken, or other means, the passenger seat to exit not extended; and ratio may be increased.

(3) With the rotorcraft resting on its (3) Flotation devices, whether stowed side, if required by § 29.803(d).

or deployed, may not interfere with or (g) The slide for each passenger emer- obstruct the exits.

gency exit must be a self-supporting (e) Ramp exits. One Type I exit only, slide or equivalent, and must be de- or one Type II exit only, that is re- signed to meet the following require- quired in the side of the fuselage under ments: paragraph (b) of this section, may be (1) It must be automatically de- installed instead in the ramp of floor ployed, and deployment must begin ramp rotorcraft if— during the interval between the time (1) Its installation in the side of the the exit opening means is actuated fuselage is impractical; and from inside the rotorcraft and the time (2) Its installation in the ramp meets the exit is fully opened. However, each § 29.813.

passenger emergency exit which is also (f) Tests. The proper functioning of a passenger entrance door or a service each emergency exit must be shown by door must be provided with means to test.

prevent deployment of the slide when the exit is opened from either the in- [Amdt. 29–3, 33 FR 968, Jan. 26, 1968, as side or the outside under non- amended by Amdt. 29–12, 41 FR 55472, Dec. 20, 1976; Amdt. 27–26, 55 FR 8004, Mar. 6, 1990] emergency conditions for normal use.

(2) It must be automatically erected § 29.809 Emergency exit arrangement.

within 10 seconds after deployment is begun.

(a) Each emergency exit must consist of a movable door or hatch in the ex- (3) It must be of such length after full ternal walls of the fuselage and must deployment that the lower end is self- provide an unobstructed opening to the supporting on the ground and provides outside. safe evacuation of occupants to the Federal Aviation Administration, DOT § 29.811 ground after collapse of one or more overwater flights if the rotorcraft is legs or part of the landing gear. capsized and the cabin is submerged.

(4) It must have the capability, in 25- (b) The identity and location of each knot winds directed from the most passenger emergency exit must be rec- critical angle, to deploy and, with the ognizable from a distance equal to the assistance of only one person, to re- width of the cabin.

main usable after full deployment to (c) The location of each passenger evacuate occupants safely to the emergency exit must be indicated by a ground.

sign visible to occupants approaching (5) Each slide installation must be along the main passenger aisle. There qualified by five consecutive deploy- must be a locating sign— ment and inflation tests conducted (per (1) Next to or above the aisle near exit) without failure, and at least three each floor emergency exit, except that tests of each such five-test series must one sign may serve two exits if both ex- be conducted using a single representa- ists can be seen readily from that sign; tive sample of the device. The sample and devices must be deployed and inflated (2) On each bulkhead or divider that by the system’s primary means after prevents fore and aft vision along the being subjected to the inertia forces passenger cabin, to indicate emergency specified in § 29.561(b). If any part of the exits beyond and obscured by it, except system fails or does not function prop- that if this is not possible the sign may erly during the required tests, the be placed at another appropriate loca- cause of the failure or malfunction tion.

must be corrected by positive means (d) Each passenger emergency exit and after that, the full series of five marking and each locating sign must consecutive deployment and inflation have white letters 1 inch high on a red tests must be conducted without fail- background 2 inches high, be self or ure.

electrically illuminated, and have a (h) For rotorcraft having 30 or fewer minimum luminescence (brightness) of passenger seats and having an exit at least 160 microlamberts. The colors threshold more than 6 feet above the may be reversed if this will increase ground, a rope or other assist means the emergency illumination of the pas- may be used in place of the slide speci- senger compartment.

fied in paragraph (f) of this section, (e) The location of each passenger provided an evacuation demonstration emergency exit operating handle and is accomplished as prescribed in instructions for opening must be § 29.803(d) or (e).

shown— (i) If a rope, with its attachment, is (1) For each emergency exit, by a used for compliance with paragraph (f), marking on or near the exit that is (g), or (h) of this section, it must— readable from a distance of 30 inches; (1) Withstand a 400-pound static load; and and (2) Attach to the fuselage structure (2) For each Type I or Type II emer- at or above the top of the emergency gency exit with a locking mechanism exit opening, or at another approved released by rotary motion of the han- location if the stowed rope would re- dle, by— duce the pilot’s view in flight. (i) A red arrow, with a shaft at least three-fourths inch wide and a head [Amdt. 29–3, 33 FR 968, Jan. 26, 1968, as twice the width of the shaft, extending amended by Amdt. 29–29, 54 FR 47321, Nov. 13, along at least 70 degrees of arc at a ra- 1989; Amdt. 27–26, 55 FR 8004, Mar. 6, 1990] dius approximately equal to three- § 29.811 Emergency exit marking.

fourths of the handle length; and (ii) The word ‘‘open’’ in red letters 1 (a) Each passenger emergency exit, inch high, placed horizontally near the its means of access, and its means of head of the arrow.

opening must be conspicuously marked for the guidance of occupants using the (f) Each emergency exit, and its exits in daylight or in the dark. Such means of opening, must be marked on markings must be designed to remain the outside of the rotorcraft. In addi- visible for rotorcraft equipped for tion, the following apply: 14 CFR Ch. I (1–1–25 Edition) § 29.812 (1) There must be a 2-inch colored tact with the ground outside the cabin.

band outlining each passenger emer- The exterior emergency lighting may gency exit, except small rotorcraft be provided by either interior or exte- with a maximum weight of 12,500 rior sources with light intensity meas- pounds or less may have a 2-inch col- urements made with the emergency ored band outlining each exit release exits open.

lever or device of passenger emergency (c) Each light required by paragraph exits which are normally used doors.

(a) or (b) of this section must be oper- (2) Each outside marking, including able manually from the cockpit station the band, must have color contrast to and from a point in the passenger com- be readily distinguishable from the sur- partment that is readily accessible.

rounding fuselage surface. The contrast The cockpit control device must have must be such that, if the reflectance of an ‘‘on,’’ ‘‘off,’’ and ‘‘armed’’ position the darker color is 15 percent or less, so that when turned on at the cockpit the reflectance of the lighter color or passenger compartment station or must be at least 45 percent. ‘‘Reflec- when armed at the cockpit station, the tance’’ is the ratio of the luminous flux emergency lights will either illuminate reflected by a body to the luminous or remain illuminated upon interrup- flux it receives. When the reflectance tion of the rotorcraft’s normal electric of the darker color is greater than 15 power.

percent, at least a 30 percent difference between its reflectance and the reflec- (d) Any means required to assist the tance of the lighter color must be pro- occupants in descending to the ground vided.

must be illuminated so that the erect- (g) Exits marked as such, though in ed assist means is visible from the excess of the required number of exits, rotorcraft.

must meet the requirements for emer- (1) The assist means must be pro- gency exits of the particular type.

vided with an illumination of not less Emergency exits need only be marked than 0.03 foot-candle (measured normal with the word ‘‘Exit.’’ to the direction of the incident light) [Amdt. 29–3, 33 FR 968, Jan. 26, 1968, as at the ground end of the erected assist amended by Amdt. 29–24, 49 FR 44438, Nov. 6, means where an evacuee using the es- 1984; Amdt. 27–26, 55 FR 8004, Mar. 6, 1990; tablished escape route would normally Amdt. 29–31, 55 FR 38967, Sept. 21, 1990] make first contact with the ground, with the rotorcraft in each of the atti- § 29.812 Emergency lighting.

tudes corresponding to the collapse of For transport Category A rotorcraft, one or more legs of the landing gear.

the following apply: (2) If the emergency lighting sub- (a) A source of light with its power system illuminating the assist means supply independent of the main light- is independent of the rotorcraft’s main ing system must be installed to— emergency lighting system, it— (1) Illuminate each passenger emer- gency exit marking and locating sign; (i) Must automatically be activated and when the assist means is erected; (2) Provide enough general lighting (ii) Must provide the illumination re- in the passenger cabin so that the aver- quired by paragraph (d)(1); and age illumination, when measured at 40- (iii) May not be adversely affected by inch intervals at seat armrest height stowage.

on the center line of the main pas- (e) The energy supply to each emer- senger aisle, is at least 0.05 foot-candle.

gency lighting unit must provide the (b) Exterior emergency lighting must required level of illumination for at be provided at each emergency exit.

least 10 minutes at the critical ambient The illumination may not be less than conditions after an emergency landing.

0.05 foot-candle (measured normal to (f) If storage batteries are used as the the direction of incident light) for min- energy supply for the emergency light- imum width on the ground surface, ing system, they may be recharged with landing gear extended, equal to from the rotorcraft’s main electrical the width of the emergency exit where power system provided the charging an evacuee is likely to make first con- Federal Aviation Administration, DOT § 29.851 circuit is designed to preclude inad- Minimum main passenger aisle width vertent battery discharge into charg- Passenger seating capacity Less than 25 Inches ing circuit faults.

25 inches and more from floor from floor [Amdt. 29–24, 49 FR 44438, Nov. 6, 1984] (inches) (inches) § 29.813 Emergency exit access. 20 or more ................................. 15 20 A narrower width not less than 9 inches may be approved (a) Each passageway between pas- when substantiated by tests found necessary by the senger compartments, and each pas- Administrator.

sageway leading to Type I and Type II emergency exits, must be— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55472, Dec. 20, (1) Unobstructed; and 1976] (2) At least 20 inches wide.

(b) For each emergency exit covered § 29.831 Ventilation.

by § 29.809(f), there must be enough (a) Each passenger and crew compart- space adjacent to that exit to allow a ment must be ventilated, and each crewmember to assist in the evacu- crew compartment must have enough ation of passengers without reducing fresh air (but not less than 10 cu. ft. per the unobstructed width of the passage- minute per crewmember) to let crew- way below that required for that exit.

members perform their duties without (c) There must be access from each undue discomfort or fatigue.

aisle to each Type III and Type IV exit, (b) Crew and passenger compartment and air must be free from harmful or haz- (1) For rotorcraft that have a pas- ardous concentrations of gases or va- senger seating configuration, excluding pors.

pilot seats, of 20 or more, the projected (c) The concentration of carbon mon- opening of the exit provided must not oxide may not exceed one part in 20,000 be obstructed by seats, berths, or other parts of air during forward flight. If the protrusions (including seatbacks in any concentration exceeds this value under position) for a distance from that exit other conditions, there must be suit- of not less than the width of the nar- able operating restrictions.

rowest passenger seat installed on the rotorcraft; (d) There must be means to ensure compliance with paragraphs (b) and (c) (2) For rotorcraft that have a pas- senger seating configuration, excluding of this section under any reasonably probable failure of any ventilating, pilot seats, of 19 or less, there may be heating, or other system or equipment.

minor obstructions in the region de- scribed in paragraph (c)(1) of this sec- § 29.833 Heaters.

tion, if there are compensating factors to maintain the effectiveness of the Each combustion heater must be ap- exit.

proved.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as F IRE P ROTECTION amended by Amdt. 29–12, 41 FR 55472, Dec. 20, 1976] § 29.851 Fire extinguishers.

§ 29.815 Main aisle width. (a) Hand fire extinguishers. For hand fire extinguishers the following apply: The main passenger aisle width be- (1) Each hand fire extinguisher must tween seats must equal or exceed the be approved.

values in the following table: (2) The kinds and quantities of each Minimum main passenger extinguishing agent used must be ap- aisle width propriate to the kinds of fires likely to Passenger seating capacity Less than 25 Inches occur where that agent is used.

25 inches and more from floor from floor (3) Each extinguisher for use in a per- (inches) (inches) sonnel compartment must be designed to minimize the hazard of toxic gas 10 or less ................................... 12 15 11 through 19 ............................ 12 20 concentrations.

14 CFR Ch. I (1–1–25 Edition) § 29.853 (b) Built-in fire extinguishers. If a tinguishing when tested vertically in built-in fire extinguishing system is re- accordance with the applicable portion quired— of appendix F of Part 25 of this chapter, (1) The capacity of each system, in or other approved equivalent methods.

relation to the volume of the compart- The average burn length may not ex- ment where used and the ventilation ceed 8 inches and the average flame rate, must be adequate for any fire time after removal of the flame source likely to occur in that compartment. may not exceed 15 seconds. Drippings (2) Each system must be installed so from the test specimen may not con- that— tinue to flame for more than an aver- (i) No extinguishing agent likely to age of 5 seconds after falling.

enter personnel compartments will be (3) Acrylic windows and signs, parts present in a quantity that is hazardous constructed in whole or in part of to the occupants; and elastometric materials, edge lighted (ii) No discharge of the extinguisher instrument assemblies consisting of can cause structural damage.

two or more instruments in a common housing, seat belts, shoulder harnesses, § 29.853 Compartment interiors.

and cargo and baggage tiedown equip- For each compartment to be used by ment, including containers, bins, pal- the crew or passengers— lets, etc., used in passenger or crew (a) The materials (including finishes compartments, may not have an aver- or decorative surfaces applied to the age burn rate greater than 2.5 inches materials) must meet the following per minute when tested horizontally in test criteria as applicable: accordance with the applicable por- (1) Interior ceiling panels, interior tions of appendix F of Part 25 of this wall panels, partitions, galley struc- chapter, or other approved equivalent ture, large cabinet walls, structural methods.

flooring, and materials used in the con- (4) Except for electrical wire and struction of stowage compartments cable insulation, and for small parts (other than underseat stowage com- (such as knobs, handles, rollers, fas- partments and compartments for stow- teners, clips, grommets, rub strips, pul- ing small items such as magazines and leys, and small electrical parts) that maps) must be self-extinguishing when the Administrator finds would not con- tested vertically in accordance with tribute significantly to the propaga- the applicable portions of appendix F tion of a fire, materials in items not of Part 25 of this chapter, or other ap- specified in paragraphs (a)(1), (a)(2), or proved equivalent methods. The aver- (a)(3) of this section may not have a age burn length may not exceed 6 burn rate greater than 4 inches per inches and the average flame time minute when tested horizontally in ac- after removal of the flame source may cordance with the applicable portions not exceed 15 seconds. Drippings from of appendix F of Part 25 of this chapter, the test specimen may not continue to or other approved equivalent methods.

flame for more than an average of 3 (b) In addition to meeting the re- seconds after falling. quirements of paragraph (a)(2), seat (2) Floor covering, textiles (including cushions, except those on flight crew- draperies and upholstery), seat cush- member seats, must meet the test re- ions, padding, decorative and non- quirements of Part II of appendix F of decorative coated fabrics, leather, Part 25 of this chapter, or equivalent.

trays and galley furnishings, electrical (c) If smoking is to be prohibited, conduit, thermal and acoustical insula- there must be a placard so stating, and tion and insulation covering, air duct- if smoking is to be allowed— ing, joint and edge covering, cargo (1) There must be an adequate num- compartment liners, insulation blan- ber of self-contained, removable ash- kets, cargo covers, and transparencies, trays; and molded and thermoformed parts, air (2) Where the crew compartment is ducting joints, and trim strips (decora- separated from the passenger compart- tive and chafing) that are constructed ment, there must be at least one illu- of materials not covered in paragraph minated sign (using either letters or (a)(3) of this section, must be self ex- symbols) notifying all passengers when Federal Aviation Administration, DOT § 29.859 smoking is prohibited. Signs which no- (iv) Notwithstanding § 29.1439(a), pro- tify when smoking is prohibited must— tective breathing equipment is not re- quired.

(i) When illuminated, be legible to (b) No compartment may contain any each passenger seated in the passenger controls, wiring, lines, equipment, or cabin under all probable lighting condi- accessories whose damage or failure tions; and would affect safe operation, unless (ii) Be so constructed that the crew those items are protected so that— can turn the illumination on and off.

(1) They cannot be damaged by the (d) Each receptacle for towels, paper, movement of cargo in the compart- or waste must be at least fire-resistant ment; and and must have means for containing (2) Their breakage or failure will not possible fires; create a fire hazard.

(e) There must be a hand fire extin- (c) The design and sealing of inacces- guisher for the flight crewmembers; sible compartments must be adequate and to contain compartment fires until a (f) At least the following number of landing and safe evacuation can be hand fire extinguishers must be con- made.

veniently located in passenger com- (d) Each cargo and baggage compart- partments: ment that is not sealed so as to contain cargo compartment fires completely Fire extin- Passenger capacity guishers without endangering the safety of a rotorcraft or its occupants must be de- signed, or must have a device, to en- sure detection of fires or smoke by a crewmember while at his station and to prevent the accumulation of harm- (Secs. 313(a), 601, 603, 604, Federal Aviation ful quantities of smoke, flame, extin- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), guishing agents, and other noxious sec. 6(c), Dept. of Transportation Act (49 gases in any crew or passenger com- U.S.C. 1655(c))) partment. This must be shown in [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as flight.

amended by Amdt. 29–3, 33 FR 969, Jan. 26, (e) For rotorcraft used for the car- 1968; Amdt. 29–17, 43 FR 50600, Oct. 30, 1978; riage of cargo only, the cabin area may Amdt. 29–18, 45 FR 7756, Feb. 4, 1980; Amdt.

be considered a cargo compartment 29–23, 49 FR 43200, Oct. 26, 1984] and, in addition to paragraphs (a) through (d) of this section, the fol- § 29.855 Cargo and baggage compart- lowing apply: ments.

(1) There must be means to shut off (a) Each cargo and baggage compart- the ventilating airflow to or within the ment must be construced of or lined compartment. Controls for this purpose with materials in accordance with the must be accessible to the flight crew in following: the crew compartment.

(1) For accessible and inaccessible (2) Required crew emergency exits compartments not occupied by pas- must be accessible under all cargo sengers or crew, the material must be loading conditions.

at least fire resistant.

(3) Sources of heat within each com- (2) Materials must meet the require- partment must be shielded and insu- ments in § 29.853(a)(1), (a)(2), and (a)(3) lated to prevent igniting the cargo.

for cargo or baggage compartments in [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as which— amended by Amdt. 29–3, 33 FR 969, Jan. 26, (i) The presence of a compartment 1968; Amdt. 29–24, 49 FR 44438, Nov. 6, 1984; fire would be easily discovered by a Amdt. 27–26, 55 FR 8004, Mar. 6, 1990] crewmember while at the crew- member’s station; § 29.859 Combustion heater fire pro- tection.

(ii) Each part of the compartment is easily accessible in flight; (a) Combustion heater fire zones. The (iii) The compartment has a volume following combustion heater fire zones of 200 cubic feet or less; and must be protected against fire under 14 CFR Ch. I (1–1–25 Edition) § 29.859 the applicable provisions of §§ 29.1181 or in any heater control component, through 29.1191, and 29.1195 through control system tubing, or safety con- 29.1203: trol.

(1) The region surrounding any heat- (e) Heater safety controls. For each er, if that region contains any flam- combustion heater, safety control mable fluid system components (in- means must be provided as follows: cluding the heater fuel system), that (1) Means independent of the compo- could— nents provided for the normal contin- (i) Be damaged by heater malfunc- uous control of air temperature, air- tioning; or flow, and fuel flow must be provided, (ii) Allow flammable fluids or vapors for each heater, to automatically shut to reach the heater in case of leakage. off the ignition and fuel supply of that (2) Each part of any ventilating air heater at a point remote from that passage that— heater when any of the following oc- (i) Surrounds the combustion cham- curs: ber; and (i) The heat exchanger temperature (ii) Would not contain (without dam- exceeds safe limits.

age to other rotorcraft components) (ii) The ventilating air temperature any fire that may occur within the pas- exceeds safe limits.

sage. (iii) The combustion airflow becomes (b) Ventilating air ducts. Each ven- inadequate for safe operation.

tilating air duct passing through any (iv) The ventilating airflow becomes fire zone must be fireproof. In addi- inadequate for safe operation.

tion— (2) The means of complying with (1) Unless isolation is provided by paragraph (e)(1) of this section for any fireproof valves or by equally effective individual heater must— means, the ventilating air duct down- (i) Be independent of components stream of each heater must be fireproof serving any other heater whose heat for a distance great enough to ensure output is essential for safe operation; that any fire originating in the heater and can be contained in the duct; and (ii) Keep the heater off until re- (2) Each part of any ventilating duct started by the crew.

passing through any region having a (3) There must be means to warn the flammable fluid system must be so crew when any heater whose heat out- constructed or isolated from that sys- put is essential for safe operation has tem that the malfunctioning of any been shut off by the automatic means component of that system cannot in- prescribed in paragraph (e)(1) of this troduce flammable fluids or vapors section.

into the ventilating airstream. (f) Air intakes. Each combustion and (c) Combustion air ducts. Each com- ventilating air intake must be where bustion air duct must be fireproof for a no flammable fluids or vapors can distance great enough to prevent dam- enter the heater system under any op- age from backfiring or reverse flame erating condition— propagation. In addition— (1) During normal operation; or (1) No combustion air duct may com- (2) As a result of the malfunction of municate with the ventilating air- any other component.

stream unless flames from backfires or (g) Heater exhaust. Each heater ex- reverse burning cannot enter the ven- haust system must meet the require- tilating airstream under any operating ments of §§ 29.1121 and 29.1123. In addi- condition, including reverse flow or tion— malfunction of the heater or its associ- (1) Each exhaust shroud must be ated components; and sealed so that no flammable fluids or (2) No combustion air duct may re- hazardous quantities of vapors can strict the prompt relief of any backfire reach the exhaust systems through that, if so restricted, could cause heat- joints; and er failure. (2) No exhaust system may restrict (d) Heater controls; general. There the prompt relief of any backfire that, must be means to prevent the haz- if so restricted, could cause heater fail- ardous accumulation of water or ice on ure.

Federal Aviation Administration, DOT § 29.865 (h) Heater fuel systems. Each heater (2) Flammability characteristics of fuel system must meet the powerplant fluids, including effects of any combus- fuel system requirements affecting safe tible or absorbing materials.

heater operation. Each heater fuel sys- (3) Possible ignition sources, includ- ing electrical faults, overheating of tem component in the ventilating air- stream must be protected by shrouds equipment, and malfunctioning of pro- so that no leakage from those compo- tective devices.

nents can enter the ventilating air- (4) Means available for controlling or stream. extinguishing a fire, such as stopping (i) Drains. There must be means for flow of fluids, shutting down equip- safe drainage of any fuel that might ac- ment, fireproof containment, or use of cumulate in the combustion chamber extinguishing agents.

(5) Ability of rotorcraft components or the heat exchanger. In addition— (1) Each part of any drain that oper- that are critical to safety of flight to ates at high temperatures must be pro- withstand fire and heat.

tected in the same manner as heater (c) If action by the flight crew is re- exhausts; and quired to prevent or counteract a fluid (2) Each drain must be protected fire (e.g. equipment shutdown or actu- against hazardous ice accumulation ation of a fire extinguisher), quick act- under any operating condition. ing means must be provided to alert the crew.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (d) Each area where flammable fluids amended by Amdt. 29–2, 32 FR 6914, May 5, or vapors might escape by leakage of a 1967] fluid system must be identified and de- § 29.861 Fire protection of structure, fined.

controls, and other parts.

(Secs. 313(a), 601, 603, 604, Federal Aviation Each part of the structure, controls, Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), and the rotor mechanism, and other sec. 6(c), Dept. of Transportation Act (49 parts essential to controlled landing U.S.C. 1655(c))) and (for category A) flight that would [Amdt. 29–17, 43 FR 50600, Oct. 30, 1978] be affected by powerplant fires must be isolated under § 29.1191, or must be— E XTERNAL L OADS (a) For category A rotorcraft, fire- § 29.865 External loads.

proof; and (b) For Category B rotorcraft, fire- (a) It must be shown by analysis, proof or protected so that they can per- test, or both, that the rotorcraft exter- form their essential functions for at nal load attaching means for rotor- least 5 minutes under any foreseeable craft-load combinations to be used for powerplant fire conditions.

nonhuman external cargo applications can withstand a limit static load equal [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as to 2.5, or some lower load factor ap- amended by Amdt. 27–26, 55 FR 8005, Mar. 6, proved under §§ 29.337 through 29.341, 1990] multiplied by the maximum external § 29.863 Flammable fluid fire protec- load for which authorization is re- tion.

quested. It must be shown by analysis, (a) In each area where flammable test, or both that the rotorcraft exter- fluids or vapors might escape by leak- nal load attaching means and cor- age of a fluid system, there must be responding personnel carrying device means to minimize the probability of system for rotorcraft-load combina- ignition of the fluids and vapors, and tions to be used for human external the resultant hazards if ignition does cargo applications can withstand a occur. limit static load equal to 3.5 or some (b) Compliance with paragraph (a) of lower load factor, not less than 2.5, ap- this section must be shown by analysis proved under §§ 29.337 through 29.341, or tests, and the following factors must multiplied by the maximum external be considered: load for which authorization is re- (1) Possible sources and paths of fluid quested. The load for any rotorcraft- leakage, and means of detecting leak- load combination class, for any exter- age. nal cargo type, must be applied in the 14 CFR Ch. I (1–1–25 Edition) § 29.865 vertical direction. For jettisonable ex- (B) The minimum level of protection ternal loads of any applicable external required for jettisonable rotorcraft- cargo type, the load must also be ap- load combinations used for human ex- plied in any direction making the max- ternal cargo is a radio frequency field imum angle with the vertical that can strength of 200 volts per meter.

(iii) Be protected against any failure be achieved in service but not less than that could be induced by a failure mode 30 ° . However, the 30 ° angle may be re- of any other electrical or mechanical duced to a lesser angle if— rotorcraft system.

(1) An operating limitation is estab- (c) For rotorcraft-load combinations lished limiting external load oper- to be used for human external cargo ations to such angles for which compli- applications, the rotorcraft must— ance with this paragraph has been (1) For jettisonable external loads, shown; or have a quick-release system that meets (2) It is shown that the lesser angle the requirements of paragraph (b) of can not be exceeded in service.

this section and that— (b) The external load attaching (i) Provides a dual actuation device means, for jettisonable rotorcraft-load for the primary quick release sub- combinations, must include a quick-re- system, and lease system to enable the pilot to re- (ii) Provides a separate dual actu- lease the external load quickly during ation device for the backup quick re- flight. The quick-release system must lease subsystem; consist of a primary quick release sub- (2) Have a reliable, approved per- system and a backup quick release sub- sonnel carrying device system that has system that are isolated from one an- the structural capability and personnel other. The quick release system, and safety features essential for external the means by which it is controlled, occupant safety; must comply with the following: (3) Have placards and markings at all (1) A control for the primary quick appropriate locations that clearly state release subsystem must be installed ei- the essential system operating instruc- ther on one of the pilot’s primary con- tions and, for the personnel carrying trols or in an equivalently accessible device system, ingress and egress in- location and must be designed and lo- structions; cated so that it may be operated by ei- (4) Have equipment to allow direct ther the pilot or a crewmember with- intercommunication among required out hazardously limiting the ability to crewmembers and external occupants; control the rotorcraft during an emer- (5) Have the appropriate limitations gency situation.

and procedures incorporated in the (2) A control for the backup quick re- flight manual for conducting human lease subsystem, readily accessible to external cargo operations; and either the pilot or another crew- (6) For human external cargo applica- member, must be provided.

tions requiring use of Category A (3) Both the primary and backup rotorcraft, have one-engine-inoperative quick release subsystems must— hover performance data and procedures (i) Be reliable, durable, and function in the flight manual for the weights, properly with all external loads up to altitudes, and temperatures for which and including the maximum external external load approval is requested.

limit load for which authorization is (d) The critically configured jettison- requested.

able external loads must be shown by a (ii) Be protected against electro- combination of analysis, ground tests, magnetic interference (EMI) from ex- and flight tests to be both transport- ternal and internal sources and against able and releasable throughout the ap- lightning to prevent inadvertent load proved operational envelope without release.

hazard to the rotorcraft during normal (A) The minimum level of protection flight conditions. In addition, these ex- required for jettisonable rotorcraft- ternal loads—must be shown to be re- load combinations used for nonhuman leasable without hazard to the rotor- external cargo is a radio frequency craft during emergency flight condi- field strength of 20 volts per meter. tions.

Federal Aviation Administration, DOT § 29.903 (e) A placard or marking must be in- (2) Each component of the installa- stalled next to the external-load at- tion must be constructed, arranged, taching means clearly stating any and installed to ensure its continued operational limitations and the max- safe operation between normal inspec- imum authorized external load as dem- tions or overhauls for the range of tem- onstrated under § 29.25 and this section.

perature and altitude for which ap- (f) The fatigue evaluation of § 29.571 proval is requested.

of this part does not apply to rotor- (3) Accessibility must be provided to craft-load combinations to be used for allow any inspection and maintenance nonhuman external cargo except for necessary for continued airworthiness; the failure of critical structural ele- and ments that would result in a hazard to (4) Electrical interconnections must the rotorcraft. For rotorcraft-load be provided to prevent differences of combinations to be used for human ex- potential between major components of ternal cargo, the fatigue evaluation of the installation and the rest of the § 29.571 of this part applies to the entire rotorcraft.

quick release and personnel carrying (5) Axial and radial expansion of tur- device structural systems and their at- bine engines may not affect the safety tachments.

of the installation.

[Amdt. 29–12, 41 FR 55472, Dec. 20, 1976, as (6) Design precautions must be taken amended by Amdt. 27–26, 55 FR 8005, Mar. 6, to minimize the possibility of incorrect 1990; Amdt. 29–43, 64 FR 43020, Aug. 6, 1999] assembly of components and equipment essential to safe operation of the rotor- M ISCELLANEOUS craft, except where operation with the incorrect assembly can be shown to be § 29.871 Leveling marks.

extremely improbable.

There must be reference marks for (c) For each powerplant and auxiliary leveling the rotorcraft on the ground.

power unit installation, it must be es- § 29.873 Ballast provisions. tablished that no single failure or mal- function or probable combination of Ballast provisions must be designed failures will jeopardize the safe oper- and constructed to prevent inadvertent ation of the rotorcraft except that the shifting of ballast in flight.

failure of structural elements need not be considered if the probability of any Subpart E—Powerplant such failure is extremely remote.

(d) Each auxiliary power unit instal- G ENERAL lation must meet the applicable provi- sions of this subpart.

§ 29.901 Installation.

(a) For the purpose of this part, the (Secs. 313(a), 601, 603, 604, Federal Aviation powerplant installation includes each Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 part of the rotorcraft (other than the U.S.C. 1655(c))) main and auxiliary rotor structures) that— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (1) Is necessary for propulsion; amended by Amdt. 29–3, 33 FR 969, Jan. 26, (2) Affects the control of the major 1968; Amdt. 29–13, 42 FR 15046, Mar. 17, 1977; Amdt. 29–17, 43 FR 50600, Oct. 30, 1978; Amdt.

propulsive units; or 29–26, 53 FR 34215, Sept. 2, 1988; Amdt. 29–36, (3) Affects the safety of the major 60 FR 55776, Nov. 2, 1995] propulsive units between normal in- spections or overhauls.

§ 29.903 Engines.

(b) For each powerplant installa- tion— (a) Engine type certification. Each en- (1) The installation must comply gine must have an approved type cer- with— tificate. Reciprocating engines for use in helicopters must be qualified in ac- (i) The installation instructions pro- vided under § 33.5 of this chapter; and cordance with § 33.49(d) of this chapter (ii) The applicable provisions of this or be otherwise approved for the in- subpart. tended usage.

14 CFR Ch. I (1–1–25 Edition) § 29.907 (b) Category A; engine isolation. For (3) Following the in-flight shutdown each category A rotorcraft, the power- of all engines, in-flight engine restart plants must be arranged and isolated capability must be provided.

from each other to allow operation, in (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 at least one configuration, so that the U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 failure or malfunction of any engine, or U.S.C. 1655(c)) the failure of any system that can af- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as fect any engine, will not— amended by Amdt. 29–12, 41 FR 55472, Dec. 20, (1) Prevent the continued safe oper- 1976; Amdt. 29–26, 53 FR 34215, Sept. 2, 1988; ation of the remaining engines; or Amdt. 29–31, 55 FR 38967, Sept. 21, 1990; 55 FR (2) Require immediate action, other 41309, Oct. 10, 1990; Amdt. 29–36, 60 FR 55776, than normal pilot action with primary Nov. 2, 1995] flight controls, by any crewmember to § 29.907 Engine vibration.

maintain safe operation.

(c) Category A; control of engine rota- (a) Each engine must be installed to tion. For each Category A rotorcraft, prevent the harmful vibration of any there must be a means for stopping the part of the engine or rotorcraft.

rotation of any engine individually in (b) The addition of the rotor and the flight, except that, for turbine engine rotor drive system to the engine may installations, the means for stopping not subject the principal rotating parts the engine need be provided only where of the engine to excessive vibration necessary for safety. In addition— stresses. This must be shown by a vi- (1) Each component of the engine bration investigation.

stopping system that is located on the § 29.908 Cooling fans.

engine side of the firewall, and that might be exposed to fire, must be at For cooling fans that are a part of a least fire resistant; or powerplant installation the following (2) Duplicate means must be avail- apply: able for stopping the engine and the (a) Category A. For cooling fans in- controls must be where all are not like- stalled in Category A rotorcraft, it ly to be damaged at the same time in must be shown that a fan blade failure case of fire.

will not prevent continued safe flight (d) Turbine engine installation. For either because of damage caused by the turbine engine installations— failed blade or loss of cooling air.

(1) Design precautions must be taken (b) Category B. For cooling fans in- to minimize the hazards to the rotor- stalled in category B rotorcraft, there craft in the event of an engine rotor must be means to protect the rotor- failure; and craft and allow a safe landing if a fan (2) The powerplant systems associ- blade fails. It must be shown that— ated with engine control devices, sys- (1) The fan blade would be contained tems, and instrumentation must be de- in the case of a failure; signed to give reasonable assurance (2) Each fan is located so that a fan that those engine operating limitations blade failure will not jeopardize safety; that adversely affect engine rotor or structural integrity will not be exceed- (3) Each fan blade can withstand an ed in service.

ultimate load of 1.5 times the cen- (e) Restart capability. (1) A means to trifugal force expected in service, lim- restart any engine in flight must be ited by either— provided.

(i) The highest rotational speeds (2) Except for the in-flight shutdown achievable under uncontrolled condi- of all engines, engine restart capability tions; or must be demonstrated throughout a (ii) An overspeed limiting device.

flight envelope for the rotorcraft.

(c) Fatigue evaluation. Unless a fa- tigue evaluation under § 29.571 is con- ducted, it must be shown that cooling Federal Aviation Administration, DOT § 29.923 fan blades are not operating at reso- trol of the rotorcraft when the device nant conditions within the operating is operating.

limits of the rotorcraft.

(5) If the rotors must be phased for intermeshing, each system must pro- (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 vide constant and positive phase rela- U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 tionship under any operating condi- U.S.C. 1655 (c)) tion.

[Amdt. 29–13, 42 FR 15046, Mar. 17, 1977, as (6) If a rotor dephasing device is in- amended by Amdt. 29–26, 53 FR 34215, Sept. 2, corporated, there must be means to 1988] keep the rotors locked in proper phase before operation.

R OTOR D RIVE S YSTEM [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as § 29.917 Design.

amended by Amdt. 29–12, 41 FR 55472, Dec. 20, 1976; Amdt. 29–40, 61 FR 21908, May 10, 1996] (a) General. The rotor drive system includes any part necessary to trans- § 29.921 Rotor brake.

mit power from the engines to the rotor hubs. This includes gear boxes, If there is a means to control the ro- shafting, universal joints, couplings, tation of the rotor drive system inde- rotor brake assemblies, clutches, sup- pendently of the engine, any limita- porting bearings for shafting, any at- tions on the use of that means must be tendant accessory pads or drives, and specified, and the control for that any cooling fans that are a part of, at- means must be guarded to prevent in- tached to, or mounted on the rotor advertent operation.

drive system.

§ 29.923 Rotor drive system and con- (b) Design assessment. A design assess- trol mechanism tests.

ment must be performed to ensure that the rotor drive system functions safely (a) Endurance tests, general. Each over the full range of conditions for rotor drive system and rotor control which certification is sought. The de- mechanism must be tested, as pre- sign assessment must include a de- scribed in paragraphs (b) through (n) tailed failure analysis to identify all and (p) of this section, for at least 200 failures that will prevent continued hours plus the time required to meet safe flight or safe landing and must the requirements of paragraphs (b)(2), identify the means to minimize the (b)(3), and (k) of this section. These likelihood of their occurrence. tests must be conducted as follows: (c) Arrangement. Rotor drive systems (1) Ten-hour test cycles must be used, must be arranged as follows: except that the test cycle must be ex- (1) Each rotor drive system of multi- tended to include the OEI test of para- engine rotorcraft must be arranged so graphs (b)(2) and (k), of this section if that each rotor necessary for operation OEI ratings are requested.

and control will continue to be driven (2) The tests must be conducted on by the remaining engines if any engine the rotorcraft.

fails. (3) The test torque and rotational speed must be— (2) For single-engine rotorcraft, each rotor drive system must be so arranged (i) Determined by the powerplant that each rotor necessary for control in limitations; and autorotation will continue to be driven (ii) Absorbed by the rotors to be ap- by the main rotors after disengage- proved for the rotorcraft.

ment of the engine from the main and (b) Endurance tests; takeoff run. The auxiliary rotors.

takeoff run must be conducted as fol- (3) Each rotor drive system must in- lows: corporate a unit for each engine to (1) Except as prescribed in para- automatically disengage that engine graphs (b)(2) and (b)(3) of this section, from the main and auxiliary rotors if the takeoff torque run must consist of that engine fails.

1 hour of alternate runs of 5 minutes at (4) If a torque limiting device is used takeoff torque and the maximum speed in the rotor drive system, it must be for use with takeoff torque, and 5 min- located so as to allow continued con- utes at as low an engine idle speed as 14 CFR Ch. I (1–1–25 Edition) § 29.923 practicable. The engine must be de- applicant). At least one run sequence clutched from the rotor drive system, must be conducted from a simulated and the rotor brake, if furnished and so ‘‘flight idle’’ condition. When con- intended, must be applied during the ducted on a bench test, the test se- first minute of the idle run. During the quence must be conducted following remaining 4 minutes of the idle run, stabilization at take-off power.

the clutch must be engaged so that the (ii) For the purpose of this para- engine drives the rotors at the min- graph, an affected power input includes imum practical r.p.m. The engine and all parts of the rotor drive system the rotor drive system must be acceler- which can be adversely affected by the ated at the maximum rate. When de- application of higher or asymmetric clutching the engine, it must be decel- torque and speed prescribed by the erated rapidly enough to allow the op- test.

eration of the overrunning clutch.

(iii) This test may be conducted on a (2) For helicopters for which the use representative bench test facility when of a 2 ⁄2 -minute OEI rating is requested, engine limitations either preclude re- the takeoff run must be conducted as peated use of this power or would re- prescribed in paragraph (b)(1) of this sult in premature engine removals dur- section, except for the third and sixth ing the test. The loads, the vibration runs for which the takeoff torque and frequency, and the methods of applica- the maximum speed for use with take- off torque are prescribed in that para- tion to the affected rotor drive system graph. For these runs, the following components must be representative of apply: rotorcraft conditions. Test components (i) Each run must consist of at least must be those used to show compliance one period of 2 ⁄ 2 minutes with takeoff with the remainder of this section.

torque and the maximum speed for use (c) Endurance tests; maximum contin- with takeoff torque on all engines.

uous run. Three hours of continuous op- (ii) Each run must consist of at least eration at maximum continuous torque one period, for each engine in sequence, and the maximum speed for use with during which that engine simulates a maximum continuous torque must be power failure and the remaining en- conducted as follows: gines are run at the 2 ⁄2 -minute OEI (1) The main rotor controls must be torque and the maximum speed for use operated at a minimum of 15 times 1 1 with 2 ⁄2 -minute OEI torque for 2 ⁄2 min- each hour through the main rotor pitch utes.

positions of maximum vertical thrust, (3) For multiengine, turbine-powered maximum forward thrust component, rotorcraft for which the use of 30-sec- maximum aft thrust component, max- ond/2-minute OEI power is requested, imum left thrust component, and max- the takeoff run must be conducted as imum right thrust component, except prescribed in paragraph (b)(1) of this that the control movements need not section except for the following: (i) Immediately following any one 5- produce loads or blade flapping motion minute power-on run required by para- exceeding the maximum loads of mo- graph (b)(1) of this section, simulate a tions encountered in flight.

failure for each power source in turn, (2) The directional controls must be and apply the maximum torque and the operated at a minimum of 15 times maximum speed for use with 30-second each hour through the control ex- OEI power to the remaining affected tremes of maximum right turning drive system power inputs for not less torque, neutral torque as required by than 30 seconds. Each application of 30- the power applied to the main rotor, second OEI power must be followed by and maximum left turning torque.

two applications of the maximum (3) Each maximum control position torque and the maximum speed for use must be held for at least 10 seconds, with the 2 minute OEI power for not and the rate of change of control posi- less than 2 minutes each; the second tion must be at least as rapid as that application must follow a period at sta- for normal operation.

bilized continuous or 30 minute OEI power (whichever is requested by the Federal Aviation Administration, DOT § 29.923 (d) Endurance tests; 90 percent of max- that the control positions need not imum continuous run. One hour of con- produce loads or blade flapping motion tinuous operation at 90 percent of max- exceeding the maximum loads or mo- imum continuous torque and the max- tions encountered in flight): imum speed for use with 90 percent of (1) For full vertical thrust, 20 per- maximum continuous torque must be cent.

conducted.

(2) For the forward thrust compo- (e) Endurance tests; 80 percent of max- nent, 50 percent.

imum continuous run. One hour of con- (3) For the right thrust component, tinuous operation at 80 percent of max- 10 percent.

imum continuous torque and the min- (4) For the left thrust component, 10 imum speed for use with 80 percent of percent.

maximum continuous torque must be (5) For the aft thrust component, 10 conducted.

percent.

(f) Endurance tests; 60 percent of max- (j) Endurance tests, clutch and brake imum continuous run. Two hours or, for engagements. A total of at least 400 helicopters for which the use of either clutch and brake engagements, includ- 30-minute OEI power or continuous OEI ing the engagements of paragraph (b) power is requested, 1 hour of contin- of this section, must be made during uous operation at 60 percent of max- the takeoff torque runs and, if nec- imum continuous torque and the min- essary, at each change of torque and imum speed for use with 60 percent of speed throughout the test. In each maximum continuous torque must be clutch engagement, the shaft on the conducted.

driven side of the clutch must be accel- (g) Endurance tests; engine malfunc- erated from rest. The clutch engage- tioning run. It must be determined ments must be accomplished at the whether malfunctioning of compo- speed and by the method prescribed by nents, such as the engine fuel or igni- the applicant. During deceleration tion systems, or whether unequal en- after each clutch engagement, the en- gine power can cause dynamic condi- gines must be stopped rapidly enough tions detrimental to the drive system.

to allow the engines to be automati- If so, a suitable number of hours of op- cally disengaged from the rotors and eration must be accomplished under rotor drives. If a rotor brake is in- those conditions, 1 hour of which must stalled for stopping the rotor, the be included in each cycle, and the re- clutch, during brake engagements, maining hours of which must be ac- must be disengaged above 40 percent of complished at the end of the 20 cycles.

maximum continuous rotor speed and If no detrimental condition results, an the rotors allowed to decelerate to 40 additional hour of operation in compli- percent of maximum continuous rotor ance with paragraph (b) of this section speed, at which time the rotor brake must be conducted in accordance with must be applied. If the clutch design the run schedule of paragraph (b)(1) of does not allow stopping the rotors with this section without consideration of the engine running, or if no clutch is paragraph (b)(2) of this section.

provided, the engine must be stopped (h) Endurance tests; overspeed run. One before each application of the rotor hour of continuous operation must be brake, and then immediately be started conducted at maximum continuous after the rotors stop.

torque and the maximum power-on overspeed expected in service, assum- (k) Endurance tests; OEI power run — ing that speed and torque limiting de- (1) 30-minute OEI power run. For rotor- vices, if any, function properly. craft for which the use of 30-minute (i) Endurance tests; rotor control posi- OEI power is requested, a run at 30- tions. When the rotor controls are not minute OEI torque and the maximum being cycled during the tie-down tests, speed for use with 30-minute OEI the rotor must be operated, using the torque must be conducted as follows: procedures prescribed in paragraph (c) For each engine, in sequence, that en- of this section, to produce each of the gine must be inoperative and the re- maximum thrust positions for the fol- maining engines must be run for a 30- lowing percentages of test time (except minute period.

14 CFR Ch. I (1–1–25 Edition) § 29.927 (2) Continuous OEI power run. For ducted with transmission and gearbox rotorcraft for which the use of contin- lubricant temperatures, at the location uous OEI power is requested, a run at prescribed for measurement, not lower continuous OEI torque and the max- than the maximum operating tempera- imum speed for use with continuous ture for which approval is requested; OEI torque must be conducted as fol- (2) For pressure lubricated systems, lows: For each engine, in sequence, at least three 10-hour cycles required that engine must be inoperative and by this section must be conducted with the remaining engines must be run for the lubricant pressure, at the location 1 hour. prescribed for measurement, not higher (3) The number of periods prescribed than the minimum operating pressure in paragraph (k)(1) or (k)(2) of this sec- for which approval is requested; and tion may not be less than the number (3) The test conditions of paragraphs of engines, nor may it be less than two. (p)(1) and (p)(2) of this section must be (l) [Reserved] applied simultaneously and must be ex- (m) Any components that are af- tended to include operation at any one- fected by maneuvering and gust loads engine-inoperative rating for which ap- must be investigated for the same proval is requested.

flight conditions as are the main ro- (Secs. 313(a), 601, 603, 604, Federal Aviation tors, and their service lives must be de- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), termined by fatigue tests or by other sec. 6(c), Dept. of Transportation Act (49 acceptable methods. In addition, a U.S.C. 1655(c))) level of safety equal to that of the [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as main rotors must be provided for— amended by Amdt. 29–1, 30 FR 8778, July 13, (1) Each component in the rotor drive 1965; Amdt. 29–17, 43 FR 50600, Oct. 30, 1978; system whose failure would cause an Amdt. 29–26, 53 FR 34215, Sept. 2, 1988; Amdt.

uncontrolled landing; 29–31, 55 FR 38967, Sept. 21, 1990; Amdt. 29–34, (2) Each component essential to the 59 FR 47768, Sept. 16, 1994; Amdt. 29–40, 61 FR 21908, May 10, 1996; Amdt. 29–42, 63 FR 43285, phasing of rotors on multirotor rotor- Aug. 12, 1998] craft, or that furnishes a driving link for the essential control of rotors in § 29.927 Additional tests.

autorotation; and (a) Any additional dynamic, endur- (3) Each component common to two ance, and operational tests, and vibra- or more engines on multiengine rotor- tory investigations necessary to deter- craft.

mine that the rotor drive mechanism is (n) Special tests. Each rotor drive sys- safe, must be performed.

tem designed to operate at two or more (b) If turbine engine torque output to gear ratios must be subjected to special the transmission can exceed the high- testing for durations necessary to sub- est engine or transmission torque stantiate the safety of the rotor drive limit, and that output is not directly system.

controlled by the pilot under normal (o) Each part tested as prescribed in operating conditions (such as where this section must be in a serviceable the primary engine power control is ac- condition at the end of the tests. No in- complished through the flight control), tervening disassembly which might af- the following test must be made: fect test results may be conducted.

(p) Endurance tests; operating lubri- (1) Under conditions associated with cants. To be approved for use in rotor all engines operating, make 200 appli- drive and control systems, lubricants cations, for 10 seconds each, of torque must meet the specifications of lubri- that is at least equal to the lesser of— cants used during the tests prescribed (i) The maximum torque used in by this section. Additional or alternate meeting § 29.923 plus 10 percent; or lubricants may be qualified by equiva- (ii) The maximum torque attainable lent testing or by comparative analysis under probable operating conditions, of lubricant specifications and rotor assuming that torque limiting devices, drive and control system characteris- if any, function properly.

tics. In addition— (2) For multiengine rotorcraft under (1) At least three 10-hour cycles re- conditions associated with each engine, quired by this section must be con- in turn, becoming inoperative, apply to Federal Aviation Administration, DOT § 29.935 the remaining transmission torque in- (3) Overspeed runs must be made with puts the maximum torque attainable the rotors in the flattest pitch for smooth operation.

under probable operating conditions, assuming that torque limiting devices, (e) The tests prescribed in paragraphs if any, function properly. Each trans- (b) and (d) of this section must be con- mission input must be tested at this ducted on the rotorcraft and the torque must be absorbed by the rotors to be maximum torque for at least fifteen installed, except that other ground or minutes.

flight test facilities with other appro- (c) Lubrication system failure. For lu- priate methods of torque absorption brication systems required for proper may be used if the conditions of sup- operation of rotor drive systems, the port and vibration closely simulate the following apply: conditions that would exist during a (1) Category A. Unless such failures test on the rotorcraft.

are extremely remote, it must be (f) Each test prescribed by this sec- shown by test that any failure which tion must be conducted without inter- results in loss of lubricant in any nor- vening disassembly and, except for the mal use lubrication system will not lubrication system failure test re- prevent continued safe operation, al- quired by paragraph (c) of this section, though not necessarily without dam- each part tested must be in a service- age, at a torque and rotational speed able condition at the conclusion of the prescribed by the applicant for contin- test.

ued flight, for at least 30 minutes after perception by the flightcrew of the lu- (Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423 1424), brication system failure or loss of lu- sec. 6(c), Dept. of Transportation Act (49 bricant.

U.S.C. 1655(c))) (2) Category B. The requirements of [Amdt. 29–3, 33 FR 969, Jan. 26, 1968, as Category A apply except that the rotor amended by Amdt. 29–17, 43 FR 50601, Oct. 30, drive system need only be capable of 1978; Amdt. 29–26, 53 FR 34216, Sept. 2, 1988] operating under autorotative condi- tions for at least 15 minutes.

§ 29.931 Shafting critical speed.

(d) Overspeed test. The rotor drive sys- (a) The critical speeds of any shafting tem must be subjected to 50 overspeed must be determined by demonstration runs, each 30 ± 3 seconds in duration, at except that analytical methods may be not less than either the higher of the used if reliable methods of analysis are rotational speed to be expected from an available for the particular design.

engine control device failure or 105 per- (b) If any critical speed lies within, cent of the maximum rotational speed, or close to, the operating ranges for including transients, to be expected in idling, power-on, and autorotative con- service. If speed and torque limiting ditions, the stresses occurring at that devices are installed, are independent speed must be within safe limits. This of the normal engine control, and are must be shown by tests.

shown to be reliable, their rotational (c) If analytical methods are used and speed limits need not be exceeded.

show that no critical speed lies within These runs must be conducted as fol- the permissible operating ranges, the lows: margins between the calculated crit- (1) Overspeed runs must be alternated ical speeds and the limits of the allow- with stabilizing runs of from 1 to 5 able operating ranges must be adequate minutes duration each at 60 to 80 per- to allow for possible variations be- cent of maximum continuous speed.

tween the computed and actual values.

(2) Acceleration and deceleration [Amdt. 29–12, 41 FR 55472, Dec. 20, 1976] must be accomplished in a period not longer than 10 seconds (except where § 29.935 Shafting joints.

maximum engine acceleration rate will require more than 10 seconds), and the Each universal joint, slip joint, and time for changing speeds may not be other shafting joints whose lubrication deducted from the specified time for is necessary for operation must have the overspeed runs. provision for lubrication.

14 CFR Ch. I (1–1–25 Edition) § 29.939 § 29.939 Turbine engine operating § 29.952 Fuel system crash resistance.

characteristics.

Unless other means acceptable to the (a) Turbine engine operating charac- Administrator are employed to mini- teristics must be investigated in flight mize the hazard of fuel fires to occu- pants following an otherwise surviv- to determine that no adverse charac- able impact (crash landing), the fuel teristics (such as stall, surge, of flame- systems must incorporate the design out) are present, to a hazardous degree, features of this section. These systems during normal and emergency oper- must be shown to be capable of sus- ation within the range of operating taining the static and dynamic decel- limitations of the rotorcraft and of the eration loads of this section, consid- engine.

ered as ultimate loads acting alone, (b) The turbine engine air inlet sys- measured at the system component’s tem may not, as a result of airflow dis- center of gravity without structural tortion during normal operation, cause damage to the system components, fuel vibration harmful to the engine.

tanks, or their attachments that would (c) For governor-controlled engines, leak fuel to an ignition source.

it must be shown that there exists no (a) Drop test requirements. Each tank, hazardous torsional instability of the or the most critical tank, must be drive system associated with critical drop-tested as follows: combinations of power, rotational (1) The drop height must be at least speed, and control displacement.

50 feet.

[Amdt. 29–2, 32 FR 6914, May 5, 1967, as (2) The drop impact surface must be amended by Amdt. 29–12, 41 FR 55473, Dec. 20, nondeforming.

1976] (3) The tanks must be filled with water to 80 percent of the normal, full F UEL S YSTEM capacity.

(4) The tank must be enclosed in a § 29.951 General.

surrounding structure representative (a) Each fuel system must be con- of the installation unless it can be es- tablished that the surrounding struc- structed and arranged to ensure a flow ture is free of projections or other de- of fuel at a rate and pressure estab- sign features likely to contribute to lished for proper engine and auxiliary upture of the tank.

power unit functioning under any like- (5) The tank must drop freely and im- ly operating conditions, including the pact in a horizontal position ± 10 ° .

maneuvers for which certification is (6) After the drop test, there must be requested and during which the engine no leakage.

or auxiliary power unit is permitted to (b) Fuel tank load factors. Except for be in operation.

fuel tanks located so that tank rupture (b) Each fuel system must be ar- with fuel release to either significant ranged so that— ignition sources, such as engines, heat- (1) No engine or fuel pump can draw ers, and auxiliary power units, or occu- fuel from more than one tank at a pants is extremely remote, each fuel time; or tank must be designed and installed to (2) There are means to prevent intro- retain its contents under the following ducing air into the system.

ultimate inertial load factors, acting (c) Each fuel system for a turbine en- alone.

gine must be capable of sustained oper- (1) For fuel tanks in the cabin: ation throughout its flow and pressure (i) Upward—4g.

range with fuel initially saturated with (ii) Forward—16g.

water at 80 degrees F. and having 0.75cc (iii) Sideward—8g.

of free water per gallon added and (iv) Downward—20g.

cooled to the most critical condition (2) For fuel tanks located above or for icing likely to be encountered in behind the crew or passenger compart- operation.

ment that, if loosened, could injure an occupant in an emergency landing: [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (i) Upward—1.5g.

amended by Amdt. 29–10, 39 FR 35462, Oct. 1, 1974; Amdt. 29–12, 41 FR 55473, Dec. 20, 1976] (ii) Forward—8g.

Federal Aviation Administration, DOT § 29.952 (iii) Sideward—2g. uation requirements of § 29.571 without (iv) Downward—4g. leaking.

(3) For fuel tanks in other areas: (3) Alternate, equivalent means to (i) Upward—1.5g. the use of breakaway couplings must (ii) Forward—4g. not create a survivable impact-induced (iii) Sideward—2g. load on the fuel line to which it is in- (iv) Downward—4g. stalled greater than 25 to 50 percent of (c) Fuel line self-sealing breakaway the ultimate load (strength) of the couplings. Self-sealing breakaway cou- weakest component in the line and plings must be installed unless haz- must comply with the fatigue require- ardous relative motion of fuel system ments of § 29.571 without leaking.

components to each other or to local (d) Frangible or deformable structural rotorcraft structure is demonstrated to attachments. Unless hazardous relative be extremely improbable or unless motion of fuel tanks and fuel system other means are provided. The cou- components to local rotorcraft struc- plings or equivalent devices must be ture is demonstrated to be extremely installed at all fuel tank-to-fuel line improbable in an otherwise survivable connections, tank-to-tank intercon- impact, frangible or locally deformable nects, and at other points in the fuel attachments of fuel tanks and fuel sys- system where local structural deforma- tem components to local rotorcraft tion could lead to the release of fuel. structure must be used. The attach- (1) The design and construction of ment of fuel tanks and fuel system self-sealing breakaway couplings must components to local rotorcraft struc- incorporate the following design fea- ture, whether frangible or locally de- tures: formable, must be designed such that (i) The load necessary to separate a its separation or relative local defor- breakaway coupling must be between mation will occur without rupture or 25 to 50 percent of the minimum ulti- local tear-out of the fuel tank or fuel mate failure load (ultimate strength) system component that will cause fuel of the weakest component in the fluid- leakage. The ultimate strength of fran- carrying line. The separation load gible or deformable attachments must must in no case be less than 300 pounds, be as follows: regardless of the size of the fluid line. (1) The load required to separate a (ii) A breakaway coupling must sepa- frangible attachment from its support rate whenever its ultimate load (as de- structure, or deform a locally deform- fined in paragraph (c)(1)(i) of this sec- able attachment relative to its support tion) is applied in the failure modes structure, must be between 25 and 50 most likely to occur. percent of the minimum ultimate load (iii) All breakaway couplings must (ultimate strength) of the weakest incorporate design provisions to vis- component in the attached system. In ually ascertain that the coupling is no case may the load be less than 300 locked together (leak-free) and is open pounds.

during normal installation and service. (2) A frangible or locally deformable (iv) All breakaway couplings must in- attachment must separate or locally corporate design provisions to prevent deform as intended whenever its ulti- uncoupling or unintended closing due mate load (as defined in paragraph to operational shocks, vibrations, or (d)(1) of this section) is applied in the accelerations. modes most likely to occur.

(v) No breakaway coupling design (3) All frangible or locally deformable may allow the release of fuel once the attachments must comply with the fa- coupling has performed its intended tigue requirements of § 29.571.

function. (e) Separation of fuel and ignition (2) All individual breakaway cou- sources. To provide maximum crash re- plings, coupling fuel feed systems, or sistance, fuel must be located as far as equivalent means must be designed, practicable from all occupiable areas tested, installed, and maintained so in- and from all potential ignition sources.

advertent fuel shutoff in flight is im- (f) Other basic mechanical design cri- probable in accordance with § 29.955(a) teria. Fuel tanks, fuel lines, electrical and must comply with the fatigue eval- wires, and electrical devices must be 14 CFR Ch. I (1–1–25 Edition) § 29.953 designed, constructed, and installed, as within the limits specified by the en- far as practicable, to be crash resist- gine type certificate data sheet.

ant.

(2) The fuel level in the tank may not (g) Rigid or semirigid fuel tanks. Rigid exceed that established as the unusable or semirigid fuel tank or bladder walls fuel supply for that tank under § 29.959, must be impact and tear resistant.

plus that necessary to conduct the test.

[Doc. No. 26352, 59 FR 50387, Oct. 3, 1994] (3) The fuel head between the tank § 29.953 Fuel system independence.

and the engine must be critical with respect to rotorcraft flight attitudes.

(a) For category A rotorcraft— (4) The fuel flow transmitter, if in- (1) The fuel system must meet the re- stalled, and the critical fuel pump (for quirements of § 29.903(b); and pump-fed systems) must be installed to (2) Unless other provisions are made produce (by actual or simulated fail- to meet paragraph (a)(1) of this section, ure) the critical restriction to fuel flow the fuel system must allow fuel to be to be expected from component failure.

supplied to each engine through a sys- tem independent of those parts of each (5) Critical values of engine rota- system supplying fuel to other engines. tional speed, electrical power, or other (b) Each fuel system for a multien- sources of fuel pump motive power gine category B rotorcraft must meet must be applied.

the requirements of paragraph (a)(2) of (6) Critical values of fuel properties this section. However, separate fuel which adversely affect fuel flow are ap- tanks need not be provided for each en- plied during demonstrations of fuel gine.

flow capability.

(7) The fuel filter required by § 29.997 § 29.954 Fuel system lightning protec- is blocked to the degree necessary to tion.

simulate the accumulation of fuel con- The fuel system must be designed tamination required to activate the in- and arranged to prevent the ignition of dicator required by § 29.1305(a)(18).

fuel vapor within the system by— (b) Fuel transfer system. If normal op- (a) Direct lightning strikes to areas eration of the fuel system requires fuel having a high probability of stroke at- to be transferred to another tank, the tachment; transfer must occur automatically via (b) Swept lightning strokes to areas a system which has been shown to where swept strokes are highly prob- maintain the fuel level in the receiving able; and tank within acceptable limits during (c) Corona and streamering at fuel flight or surface operation of the rotor- vent outlets.

craft.

[Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] (c) Multiple fuel tanks. If an engine can be supplied with fuel from more § 29.955 Fuel flow.

than one tank, the fuel system, in addi- (a) General. The fuel system for each tion to having appropriate manual engine must provide the engine with at switching capability, must be designed least 100 percent of the fuel required to prevent interruption of fuel flow to under all operating and maneuvering that engine, without attention by the conditions to be approved for the rotor- flightcrew, when any tank supplying craft, including, as applicable, the fuel fuel to that engine is depleted of usable required to operate the engines under fuel during normal operation and any the test conditions required by § 29.927.

other tank that normally supplies fuel Unless equivalent methods are used, to that engine alone contains usable compliance must be shown by test dur- fuel.

ing which the following provisions are met, except that combinations of con- [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988, as ditions which are shown to be improb- amended by Amdt. 29–59, 88 FR 8739, Feb. 10, able need not be considered.

2023] (1) The fuel pressure, corrected for accelerations (load factors), must be Federal Aviation Administration, DOT § 29.965 (c) Each integral fuel tank must have § 29.957 Flow between interconnected tanks. facilities for inspection and repair of its interior.

(a) Where tank outlets are inter- (d) The maximum exposed surface connected and allow fuel to flow be- temperature of all components in the tween them due to gravity or flight ac- fuel tank must be less by a safe margin celerations, it must be impossible for than the lowest expected autoignition fuel to flow between tanks in quan- temperature of the fuel or fuel vapor in tities great enough to cause overflow the tank. Compliance with this re- from the tank vent in any sustained quirement must be shown under all op- flight condition.

erating conditions and under all nor- (b) If fuel can be pumped from one mal or malfunction conditions of all tank to another in flight— components inside the tank.

(1) The design of the vents and the (e) Each fuel tank installed in per- fuel transfer system must prevent sonnel compartments must be isolated structural damage to tanks from over- by fume-proof and fuel-proof enclosures filling; and that are drained and vented to the ex- (2) There must be means to warn the terior of the rotorcraft. The design and crew before overflow through the vents construction of the enclosures must occurs.

provide necessary protection for the § 29.959 Unusable fuel supply.

tank, must be crash resistant during a survivable impact in accordance with The unusable fuel supply for each § 29.952, and must be adequate to with- tank must be established as not less stand loads and abrasions to be ex- than the quantity at which the first pected in personnel compartments.

evidence of malfunction occurs under the most adverse fuel feed condition [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as occurring under any intended oper- amended by Amdt. 29–26, 53 FR 34217, Sept. 2, ations and flight maneuvers involving 1988; Amdt. 29–35, 59 FR 50388, Oct. 3, 1994] that tank.

§ 29.965 Fuel tank tests.

§ 29.961 Fuel system hot weather oper- (a) Each fuel tank must be able to ation.

withstand the applicable pressure tests Each suction lift fuel system and in this section without failure or leak- other fuel systems conducive to vapor age. If practicable, test pressures may formation must be shown to operate be applied in a manner simulating the satisfactorily (within certification lim- pressure distribution in service.

its) when using fuel at the most crit- (b) Each conventional metal tank, ical temperature for vapor formation each nonmetallic tank with walls that under critical operating conditions in- are not supported by the rotorcraft cluding, if applicable, the engine oper- structure, and each integral tank must ating conditions defined by § 29.927(b)(1) be subjected to a pressure of 3.5 p.s.i.

and (b)(2).

unless the pressure developed during [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] maximum limit acceleration or emer- gency deceleration with a full tank ex- § 29.963 Fuel tanks: general.

ceeds this value, in which case a hydro- static head, or equivalent test, must be (a) Each fuel tank must be able to applied to duplicate the acceleration withstand, without failure, the vibra- loads as far as possible. However, the tion, inertia, fluid, and structural loads pressure need not exceed 3.5 p.s.i. on to which it may be subjected in oper- surfaces not exposed to the accelera- ation.

tion loading.

(b) Each flexible fuel tank bladder or (c) Each nonmetallic tank with walls liner must be approved or shown to be supported by the rotorcraft structure suitable for the particular application must be subjected to the following and must be puncture resistant. Punc- tests: ture resistance must be shown by meeting the TSO-C80, paragraph 16.0, (1) A pressure test of at least 2.0 p.s.i.

requirements using a minimum punc- This test may be conducted on the ture force of 370 pounds. tank alone in conjunction with the test 14 CFR Ch. I (1–1–25 Edition) § 29.967 specified in paragraph (c)(2) of this sec- (5) During the test, the tank assem- tion. bly must be rocked at the rate of 16 to (2) A pressure test, with the tank 20 complete cycles per minute through mounted in the rotorcraft structure, an angle of 15 degrees on both sides of equal to the load developed by the re- the horizontal (30 degrees total), about action of the contents, with the tank the most critical axis, for 25 hours. If full, during maximum limit accelera- motion about more than one axis is tion or emergency deceleration. How- likely to be critical, the tank must be ever, the pressure need not exceed 2.0 rocked about each critical axis for 12 ⁄2 p.s.i. on surfaces faces not exposed to hours.

the acceleration loading.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 (d) Each tank with large unsupported U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 or unstiffened flat areas, or with other U.S.C. 1655 (c)) features whose failure or deformation [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as could cause leakage, must be subjected amended by Amdt. 29–13, 42 FR 15046, Mar. 17, to the following test or its equivalent: 1977] (1) Each complete tank assembly and its supports must be vibration tested § 29.967 Fuel tank installation.

while mounted to simulate the actual (a) Each fuel tank must be supported installation.

so that tank loads are not con- (2) The tank assembly must be vi- centrated on unsupported tank sur- brated for 25 hours while two-thirds faces. In addition— full of any suitable fluid. The ampli- (1) There must be pads, if necessary, tude of vibration may not be less than to prevent chafing between each tank one thirty-second of an inch, unless and its supports; otherwise substantiated.

(2) The padding must be non- (3) The test frequency of vibration absorbent or treated to prevent the ab- must be as follows: sorption of fuel; (i) If no frequency of vibration result- (3) If flexible tank liners are used, ing from any r.p.m. within the normal they must be supported so that they operating range of engine or rotor sys- are not required to withstand fluid tem speeds is critical, the test fre- loads; and quency of vibration, in number of cy- (4) Each interior surface of tank com- cles per minute, must, unless a fre- partments must be smooth and free of quency based on a more rational anal- projections that could cause wear of ysis is used, be the number obtained by the liner, unless— averaging the maximum and minimum (i) There are means for protection of power-on engine speeds (r.p.m.) for re- the liner at those points; or ciprocating engine powered rotorcraft (ii) The construction of the liner or 2,000 c.p.m. for turbine engine pow- itself provides such protection.

ered rotorcraft.

(ii) If only one frequency of vibration (b) Any spaces adjacent to tank sur- resulting from any r.p.m. within the faces must be adequately ventilated to normal operating range of engine or avoid accumulation of fuel or fumes in rotor system speeds is critical, that those spaces due to minor leakage. If frequency of vibration must be the test the tank is in a sealed compartment, frequency. ventilation may be limited to drain (iii) If more than one frequency of vi- holes that prevent clogging and that bration resulting from any r.p.m. with- prevent excessive pressure resulting in the normal operating range of en- from altitude changes. If flexible tank gine or rotor system speeds is critical, liners are installed, the venting ar- the most critical of these frequencies rangement for the spaces between the must be the test frequency. liner and its container must maintain (4) Under paragraph (d)(3)(ii) and (iii), the proper relationship to tank vent the time of test must be adjusted to ac- pressures for any expected flight condi- complish the same number of vibration tion.

cycles as would be accomplished in 25 (c) The location of each tank must hours at the frequency specified in meet the requirements of § 29.1185(b) paragraph (d)(3)(i) of this section. and (c).

Federal Aviation Administration, DOT § 29.975 (d) No rotorcraft skin immediately charges clear of the entire rotorcraft; adjacent to a major air outlet from the and engine compartment may act as the (3) Each filler cap must provide a wall of an integral tank.

fuel-tight seal under the fluid pressure expected in normal operation and in a [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as survivable impact.

amended by Amdt. 29–26, 53 FR 34217, Sept. 2, 1988; Amdt. 29–35, 59 FR 50388, Oct. 3, 1994] (b) Each filler cap or filler cap cover must warn when the cap is not fully § 29.969 Fuel tank expansion space.

locked or seated on the filler connec- Each fuel tank or each group of fuel tion.

tanks with interconnected vent sys- [Doc. No. 26352, 59 FR 50388, Oct. 3, 1994] tems must have an expansion space of not less than 2 percent of the combined § 29.975 Fuel tank vents and carbu- tank capacity. It must be impossible to retor vapor vents.

fill the fuel tank expansion space inad- (a) Fuel tank vents. Each fuel tank vertently with the rotorcraft in the must be vented from the top part of the normal ground attitude.

expansion space so that venting is ef- [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] fective under normal flight conditions.

In addition— § 29.971 Fuel tank sump.

(1) The vents must be arranged to (a) Each fuel tank must have a sump avoid stoppage by dirt or ice forma- with a capacity of not less than the tion; greater of— (2) The vent arrangement must pre- (1) 0.10 per cent of the tank capacity; vent siphoning of fuel during normal or operation; (2) ⁄ 16 gallon.

(3) The venting capacity and vent (b) The capacity prescribed in para- pressure levels must maintain accept- graph (a) of this section must be effec- able differences of pressure between tive with the rotorcraft in any normal the interior and exterior of the tank, attitude, and must be located so that during— the sump contents cannot escape (i) Normal flight operation; through the tank outlet opening.

(ii) Maximum rate of ascent and de- (c) Each fuel tank must allow drain- scent; and age of hazardous quantities of water (iii) Refueling and defueling (where from each part of the tank to the sump applicable); with the rotorcraft in any ground atti- (4) Airspaces of tanks with inter- tude to be expected in service.

connected outlets must be inter- (d) Each fuel tank sump must have a connected; drain that allows complete drainage of (5) There may be no point in any vent the sump on the ground.

line where moisture can accumulate [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as with the rotorcraft in the ground atti- amended by Amdt. 29–12, 41 FR 55473, Dec. 20, tude or the level flight attitude, unless 1976; Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] drainage is provided; § 29.973 Fuel tank filler connection. (6) No vent or drainage provision may end at any point— (a) Each fuel tank filler connection (i) Where the discharge of fuel from must prevent the entrance of fuel into the vent outlet would constitute a fire any part of the rotorcraft other than hazard; or the tank itself during normal oper- (ii) From which fumes could enter ations and must be crash resistant dur- personnel compartments; and ing a survivable impact in accordance with § 29.952(c). In addition— (7) The venting system must be de- signed to minimize spillage of fuel (1) Each filler must be marked as pre- through the vents to an ignition source scribed in § 29.1557(c)(1); (2) Each recessed filler connection in the event of a rollover during land- that can retain any appreciable quan- ing, ground operations, or a survivable tity of fuel must have a drain that dis- impact.

14 CFR Ch. I (1–1–25 Edition) § 29.977 (b) Carburetor vapor vents. Each car- from the maximum pressure, including buretor with vapor elimination connec- surge, that is likely to occur during tions must have a vent line to lead va- fueling. The maximum surge pressure pors back to one of the fuel tanks. In must be established with any combina- addition— tion of tank valves being either inten- (1) Each vent system must have tionally or inadvertently closed.

means to avoid stoppage by ice; and (d) The rotorcraft defueling system (2) If there is more than one fuel (not including fuel tanks and fuel tank tank, and it is necessary to use the vents) must withstand an ultimate tanks in a definite sequence, each load that is 2.0 times the load arising vapor vent return line must lead back from the maximum permissible to the fuel tank used for takeoff and defueling pressure (positive or nega- landing. tive) at the rotorcraft fueling connec- tion.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–26, 53 FR 34217, Sept. 2, [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as 1988; Amdt. 29–35, 59 FR 50388, Oct. 3, 1994; amended by Amdt. 29–12, 41 FR 55473, Dec. 20, Amdt. 29–42, 63 FR 43285, Aug. 12, 1998] 1976] § 29.977 Fuel tank outlet.

F UEL S YSTEM C OMPONENTS (a) There must be a fuel strainer for § 29.991 Fuel pumps.

the fuel tank outlet or for the booster (a) Compliance with § 29.955 must not pump. This strainer must— be jeopardized by failure of— (1) For reciprocating engine powered (1) Any one pump except pumps that rotorcraft, have 8 to 16 meshes per are approved and installed as parts of a inch; and type certificated engine; or (2) For turbine engine powered rotor- (2) Any component required for pump craft, prevent the passage of any object operation except the engine served by that could restrict fuel flow or damage that pump.

any fuel system component.

(b) The following fuel pump installa- (b) The clear area of each fuel tank tion requirements apply: outlet strainer must be at least five (1) When necessary to maintain the times the area of the outlet line.

proper fuel pressure— (c) The diameter of each strainer (i) A connection must be provided to must be at least that of the fuel tank transmit the carburetor air intake outlet.

static pressure to the proper fuel pump (d) Each finger strainer must be ac- relief valve connection; and cessible for inspection and cleaning.

(ii) The gauge balance lines must be [Amdt. 29–12, 41 FR 55473, Dec. 20, 1976, as independently connected to the carbu- amended by Amdt. 29–59, 88 FR 8739, Feb. 10, retor inlet pressure to avoid incorrect 2023] fuel pressure readings.

(2) The installation of fuel pumps § 29.979 Pressure refueling and fueling having seals or diaphragms that may provisions below fuel level.

leak must have means for draining (a) Each fueling connection below the leaking fuel.

fuel level in each tank must have (3) Each drain line must discharge means to prevent the escape of haz- where it will not create a fire hazard.

ardous quantities of fuel from that tank in case of malfunction of the fuel [Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] entry valve.

§ 29.993 Fuel system lines and fittings.

(b) For systems intended for pressure refueling, a means in addition to the (a) Each fuel line must be installed normal means for limiting the tank and supported to prevent excessive vi- content must be installed to prevent bration and to withstand loads due to damage to the tank in case of failure of fuel pressure, valve actuation, and ac- the normal means. celerated flight conditions.

(c) The rotorcraft pressure fueling (b) Each fuel line connected to com- system (not fuel tanks and fuel tank ponents of the rotorcraft between vents) must withstand an ultimate which relative motion could exist must load that is 2.0 times the load arising have provisions for flexibility.

Federal Aviation Administration, DOT § 29.1001 (c) Each flexible connection in fuel ponents required for proper rotorcraft lines that may be under pressure or or engine fuel system operation.

subjected to axial loading must use [Amdt. 29–10, 39 FR 35462, Oct. 1, 1974, as flexible hose assemblies.

amended by Amdt. 29–22, 49 FR 6850, Feb. 23, (d) Flexible hose must be approved. 1984; Amdt. 29–26, 53 FR 34217, Sept. 2, 1988] (e) No flexible hose that might be ad- § 29.999 Fuel system drains.

versely affected by high temperatures may be used where excessive tempera- (a) There must be at least one acces- tures will exist during operation or sible drain at the lowest point in each after engine shutdown. fuel system to completely drain the system with the rotorcraft in any § 29.995 Fuel valves.

ground attitude to be expected in serv- ice.

In addition to meeting the require- (b) Each drain required by paragraph ments of § 29.1189, each fuel valve (a) of this section including the drains must— prescribed in § 29.971 must— (a) [Reserved] (1) Discharge clear of all parts of the (b) Be supported so that no loads re- rotorcraft; sulting from their operation or from (2) Have manual or automatic means accelerated flight conditions are trans- to ensure positive closure in the off po- mitted to the lines attached to the sition; and valve.

(3) Have a drain valve— (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 (i) That is readily accessible and U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 which can be easily opened and closed; U.S.C. 1655 (c)) and (ii) That is either located or pro- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as tected to prevent fuel spillage in the amended by Amdt. 29–13, 42 FR 15046, Mar. 17, event of a landing with landing gear re- 1977] tracted.

§ 29.997 Fuel strainer or filter.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55473, Dec. 20, There must be a fuel strainer or filter 1976; Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] between the fuel tank outlet and the inlet of the first fuel system compo- § 29.1001 Fuel jettisoning.

nent which is susceptible to fuel con- If a fuel jettisoning system is in- tamination, including but not limited stalled, the following apply: to the fuel metering device or an en- (a) Fuel jettisoning must be safe dur- gine positive displacement pump, ing all flight regimes for which jetti- whichever is nearer the fuel tank out- soning is to be authorized.

let. This fuel strainer or filter must— (b) In showing compliance with para- (a) Be accessible for draining and graph (a) of this section, it must be cleaning and must incorporate a screen shown that— or element which is easily removable; (1) The fuel jettisoning system and (b) Have a sediment trap and drain, its operation are free from fire hazard; except that it need not have a drain if (2) No hazard results from fuel or fuel the strainer or filter is easily remov- vapors which impinge on any part of able for drain purposes; the rotorcraft during fuel jettisoning; (c) Be mounted so that its weight is and not supported by the connecting lines (3) Controllability of the rotorcraft or by the inlet or outlet connections of remains satisfactory throughout the the strainer or filter inself, unless ade- fuel jettisoning operation.

quate strengh margins under all load- (c) Means must be provided to auto- ing conditions are provided in the lines matically prevent jettisoning fuel and connections; and below the level required for an all-en- (d) Provide a means to remove from gine climb at maximum continuous the fuel any contaminant which would power from sea level to 5,000 feet alti- jeopardize the flow of fuel through tude and cruise thereafter for 30 min- rotorcraft or engine fuel system com- utes at maximum range engine power.

14 CFR Ch. I (1–1–25 Edition) § 29.1011 (d) The controls for any fuel jetti- (b) Expansion space. Oil tank expan- soning system must be designed to sion space must be provided so that— allow flight personnel (minimum crew) (1) Each oil tank used with a recipro- to safely interrupt fuel jettisoning dur- cating engine has an expansion space of ing any part of the jettisoning oper- not less than the greater of 10 percent ation.

of the tank capacity or 0.5 gallon, and (e) The fuel jettisoning system must each oil tank used with a turbine en- be designed to comply with the power- gine has an expansion space of not less plant installation requirements of than 10 percent of the tank capacity; § 29.901(c).

(2) Each reserve oil tank not directly (f) An auxiliary fuel jettisoning sys- connected to any engine has an expan- tem which meets the requirements of sion space of not less than two percent paragraphs (a), (b), (d), and (e) of this of the tank capacity; and section may be installed to jettison ad- (3) It is impossible to fill the expan- ditional fuel provided it has separate sion space inadvertently with the and independent controls.

rotorcraft in the normal ground atti- tude.

[Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] (c) Filler connections. Each recessed O IL S YSTEM oil tank filler connection that can re- tain any appreciable quantity of oil § 29.1011 Engines: general.

must have a drain that discharges clear of the entire rotorcraft. In addition— (a) Each engine must have an inde- (1) Each oil tank filler cap must pro- pendent oil system that can supply it vide an oil-tight seal under the pres- with an appropriate quantity of oil at a sure expected in operation; temperature not above that safe for continuous operation. (2) For category A rotorcraft, each (b) The usable oil capacity of each oil tank filler cap or filler cap cover system may not be less than the prod- must incorporate features that provide uct of the endurance of the rotorcraft a warning when caps are not fully under critical operating conditions and locked or seated on the filler connec- the maximum allowable oil consump- tion; and tion of the engine under the same con- (3) Each oil filler must be marked ditions, plus a suitable margin to en- under § 29.1557(c)(2).

sure adequate circulation and cooling.

(d) Vent. Oil tanks must be vented as Instead of a rational analysis of endur- follows: ance and consumption, a usable oil ca- (1) Each oil tank must be vented pacity of one gallon for each 40 gallons from the top part of the expansion of usable fuel may be used for recipro- space to that venting is effective under cating engine installations.

all normal flight conditions.

(c) Oil-fuel ratios lower than those (2) Oil tank vents must be arranged prescribed in paragraph (c) of this sec- so that condensed water vapor that tion may be used if they are substan- might freeze and obstruct the line can- tiated by data on the oil consumption not accumulate at any point; of the engine.

(e) Outlet. There must be means to (d) The ability of the engine and oil prevent entrance into the tank itself, cooling provisions to maintain the oil or into the tank outlet, of any object temperature at or below the maximum that might obstruct the flow of oil established value must be shown under through the system. No oil tank outlet the applicable requirements of §§ 29.1041 may be enclosed by a screen or guard through 29.1049.

that would reduce the flow of oil below a safe value at any operating tempera- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ture. There must be a shutoff valve at amended by Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] the outlet of each oil tank used with a turbine engine unless the external por- § 29.1013 Oil tanks.

tion of the oil system (including oil tank supports) is fireproof.

(a) Installation. Each oil tank instal- lation must meet the requirements of (f) Flexible liners. Each flexible oil § 29.967. tank liner must be approved or shown Federal Aviation Administration, DOT § 29.1023 to be suitable for the particular instal- the oil is contaminated to a degree lation. (with respect to particle size and den- sity) that is greater than that estab- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as lished for the engine under Part 33 of amended by Amdt. 29–10, 39 FR 35462, Oct. 1, this chapter.

1974] (3) The oil strainer or filter, unless it § 29.1015 Oil tank tests.

is installed at an oil tank outlet, must Each oil tank must be designed and incorporate a means to indicate con- installed so that— tamination before it reaches the capac- (a) It can withstand, without failure, ity established in accordance with any vibration, inertia, and fluid loads paragraph (a)(2) of this section.

to which it may be subjected in oper- (4) The bypass of a strainer or filter ation; and must be constructed and installed so (b) It meets the requirements of that the release of collected contami- § 29.965, except that instead of the pres- nants is minimized by appropriate lo- sure specified in § 29.965(b)— cation of the bypass to ensure that col- (1) For pressurized tanks used with a lected contaminants are not in the by- turbine engine, the test pressure may pass flow path.

not be less than 5 p.s.i. plus the max- (5) An oil strainer or filter that has imum operating pressure of the tank; no bypass, except one that is installed and at an oil tank outlet, must have a (2) For all other tanks, the test pres- means to connect it to the warning sure may not be less than 5 p.s.i.

system required in § 29.1305(a)(19).

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (b) Each oil strainer or filter in a amended by Amdt. 29–10, 39 FR 35462, Oct. 1, powerplant installation using recipro- 1974] cating engines must be constructed and § 29.1017 Oil lines and fittings. installed so that oil will flow at the normal rate through the rest of the (a) Each oil line must meet the re- system with the strainer or filter ele- quirements of § 29.993.

ment completely blocked.

(b) Breather lines must be arranged so that— [Amdt. 29–10, 39 FR 35463, Oct. 1, 1974, as (1) Condensed water vapor that might amended by Amdt. 29–22, 49 FR 6850, Feb. 23, freeze and obstruct the line cannot ac- 1984; Amdt. 29–26, 53 FR 34218, Sept. 2, 1988; cumulate at any point; Amdt. 29–59, 88 FR 8739, Feb. 10, 2023] (2) The breather discharge will not constitute a fire hazard if foaming oc- § 29.1021 Oil system drains.

curs, or cause emitted oil to strike the A drain (or drains) must be provided pilot’s windshield; and to allow safe drainage of the oil sys- (3) The breather does not discharge tem. Each drain must— into the engine air induction system.

(a) Be accessible; and § 29.1019 Oil strainer or filter. (b) Have manual or automatic means for positive locking in the closed posi- (a) Each turbine engine installation tion.

must incorporate an oil strainer or fil- ter through which all of the engine oil [Amdt. 29–22, 49 FR 6850, Feb. 23, 1984] flows and which meets the following re- quirements: § 29.1023 Oil radiators.

(1) Each oil strainer or filter that has (a) Each oil radiator must be able to a bypass must be constructed and in- withstand any vibration, inertia, and stalled so that oil will flow at the nor- oil pressure loads to which it would be mal rate through the rest of the sys- subjected in operation.

tem with the strainer or filter com- (b) Each oil radiator air duct must be pletely blocked.

located, or equipped, so that, in case of (2) The oil strainer or filter must fire, and with the airflow as it would be have the capacity (with respect to op- erating limitations established for the with and without the engine operating, engine) to ensure that engine oil sys- flames cannot directly strike the radi- tem functioning is not impaired when ator.

14 CFR Ch. I (1–1–25 Edition) § 29.1025 flow of lubricant from the outlet to the § 29.1025 Oil valves.

filter required by paragraph (b)(1) of (a) Each oil shutoff must meet the re- this section. The requirements of para- quirements of § 29.1189.

graph (b)(1) of this section do not apply (b) The closing of oil shutoffs may to screens installed at lubricant tank not prevent autorotation.

or sump outlets.

(c) Each oil valve must have positive (c) Splash type lubrication systems stops or suitable index provisions in for rotor drive system gearboxes must the ‘‘on’’ and ‘‘off’’ positions and must comply with §§ 29.1021 and 29.1337(d).

be supported so that no loads resulting from its operation or from accelerated [Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] flight conditions are transmitted to the lines attached to the valve. C OOLING § 29.1027 Transmission and gearboxes: § 29.1041 General.

general.

(a) The powerplant and auxiliary (a) The oil system for components of power unit cooling provisions must be the rotor drive system that require able to maintain the temperatures of continuous lubrication must be suffi- powerplant components, engine fluids, ciently independent of the lubrication and auxiliary power unit components systems of the engine(s) to ensure— and fluids within the temperature lim- (1) Operation with any engine inoper- its established for these components ative; and and fluids, under ground, water, and (2) Safe autorotation.

flight operating conditions for which (b) Pressure lubrication systems for certification is requested, and after transmissions and gearboxes must normal engine or auxiliary power unit comply with the requirements of shutdown, or both.

§§ 29.1013, paragraphs (c), (d), and (f) (b) There must be cooling provisions only, 29.1015, 29.1017, 29.1021, 29.1023, and to maintain the fluid temperatures in 29.1337(d). In addition, the system must any power transmission within safe have— values under any critical surface (1) An oil strainer or filter through (ground or water) and flight operating which all the lubricant flows, and conditions.

must— (c) Except for ground-use-only auxil- (i) Be designed to remove from the iary power units, compliance with lubricant any contaminant which may paragraphs (a) and (b) of this section damage transmission and drive system must be shown by flight tests in which components or impede the flow of lu- the temperatures of selected power- bricant to a hazardous degree; and plant component and auxiliary power (ii) Be equipped with a bypass con- unit component, engine, and trans- structed and installed so that— mission fluids are obtained under the (A) The lubricant will flow at the conditions prescribed in those para- normal rate through the rest of the graphs.

system with the strainer or filter com- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as pletely blocked; and amended by Amdt. 29–26, 53 FR 34218, Sept. 2, (B) The release of collected contami- 1988] nants is minimized by appropriate lo- cation of the bypass to ensure that col- § 29.1043 Cooling tests.

lected contaminants are not in the by- pass flowpath; (a) General. For the tests prescribed (iii) Be equipped with a means to in- in § 29.1041(c), the following apply: dicate collection of contaminants on (1) If the tests are conducted under the filter or strainer at or before open- conditions deviating from the max- ing of the bypass; imum ambient atmospheric tempera- (2) For each lubricant tank or sump ture specified in paragraph (b) of this outlet supplying lubrication to rotor section, the recorded powerplant tem- drive systems and rotor drive system peratures must be corrected under components, a screen to prevent en- paragraphs (c) and (d) of this section, trance into the lubrication system of unless a more rational correction any object that might obstruct the method is applicable.

Federal Aviation Administration, DOT § 29.1045 (2) No corrected temperature deter- barrel temperature recorded during the mined under paragraph (a)(1) of this cooling test.

section may exceed established limits.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- (3) The fuel used during the cooling eral Aviation Act of 1958 (49 U.S.C. 1354(a), tests must be of the minimum grade 1421, 1423, 1424, and 1425); and sec. 6(c) of the approved for the engines, and the mix- Dept. of Transportation Act (49 U.S.C.

1655(c))) ture settings must be those used in normal operation.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–12, 41 FR 55473, Dec. 20, (4) The test procedures must be as 1976; Amdt. 29–15, 43 FR 2327, Jan. 16, 1978; prescribed in §§ 29.1045 through 29.1049.

Amdt. 29–26, 53 FR 34218, Sept. 2, 1988] (5) For the purposes of the cooling tests, a temperature is ‘‘stabilized’’ § 29.1045 Climb cooling test proce- when its rate of change is less than 2 ° F dures.

per minute.

(a) Climb cooling tests must be con- (b) Maximum ambient atmospheric tem- ducted under this section for— perature. A maximum ambient atmos- (1) Category A rotorcraft; and pheric temperature corresponding to (2) Multiengine category B rotorcraft sea level conditions of at least 100 de- for which certification is requested grees F. must be established. The as- under the category A powerplant in- sumed temperature lapse rate is 3.6 de- stallation requirements, and under the grees F. per thousand feet of altitude requirements of § 29.861(a) at the steady above sea level until a temperature of rate of climb or descent established ¥ 69.7 degrees F. is reached, above under § 29.67(b).

which altitude the temperature is con- (b) The climb or descent cooling tests sidered constant at ¥ 69.7 degrees F.

must be conducted with the engine in- However, for winterization installa- operative that produces the most ad- tions, the applicant may select a max- verse cooling conditions for the re- imum ambient atmospheric tempera- maining engines and powerplant com- ture corresponding to sea level condi- ponents.

tions of less than 100 degrees F.

(c) Each operating engine must— (c) Correction factor (except cylinder (1) For helicopters for which the use barrels). Unless a more rational correc- of 30-minute OEI power is requested, be at 30-minute OEI power for 30 minutes, tion applies, temperatures of engine and then at maximum continuous fluids and powerplant components (ex- power (or at full throttle when above cept cylinder barrels) for which tem- the critical altitude); perature limits are established, must (2) For helicopters for which the use be corrected by adding to them the dif- of continuous OEI power is requested, ference between the maximum ambient be at continuous OEI power (or at full atmospheric temperature and the tem- throttle when above the critical alti- perature of the ambient air at the time tude); and of the first occurrence of the maximum (3) For other rotorcraft, be at max- component or fluid temperature re- imum continuous power (or at full corded during the cooling test.

throttle when above the critical alti- (d) Correction factor for cylinder barrel tude).

temperatures. Cylinder barrel tempera- (d) After temperatures have sta- tures must be corrected by adding to bilized in flight, the climb must be— them 0.7 times the difference between (1) Begun from an altitude not great- the maximum ambient atmospheric er than the lower of— temperature and the temperature of (i) 1,000 feet below the engine critcal the ambient air at the time of the first altitude; and occurrence of the maximum cylinder (ii) 1,000 feet below the maximum al- titude at which the rate of climb is 150 f.p.m; and (2) Continued for at least five min- utes after the occurrence of the highest temperature recorded, or until the 14 CFR Ch. I (1–1–25 Edition) § 29.1047 rotorcraft reaches the maximum alti- (ii) At least 5 minutes after the oc- tude for which certification is re- currence of the highest temperature re- quested. corded, if continuous OEI power or (e) For category B rotorcraft without maximum continuous power is used.

a positive rate of climb, the descent (5) The speeds must be those used in must begin at the all-engine-critical determining the takeoff flight path altitude and end at the higher of— under § 29.59.

(b) Category B. For each category B (1) The maximum altitude at which rotorcraft, cooling must be shown dur- level flight can be maintained with one ing takeoff and subsequent climb as engine operative; and follows: (2) Sea level.

(1) Each temperature must be sta- (f) The climb or descent must be con- bilized while hovering in ground effect ducted at an airspeed representing a with— normal operational practice for the (i) The power necessary for hovering; configuration being tested. However, if (ii) The appropriate cowl flap and the cooling provisions are sensitive to shutter settings; and rotorcraft speed, the most critical air- (iii) The maximum weight.

speed must be used, but need not ex- (2) After the temperatures have sta- ceed the speeds established under bilized, a climb must be started at the § 29.67(a)(2) or § 29.67(b). The climb cool- lowest practicable altitude with take- ing test may be conducted in conjunc- off power.

tion with the takeoff cooling test of (3) Takeoff power must be used for § 29.1047.

the same time interval as takeoff [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as power is used in determining the take- amended by Amdt. 29–26, 53 FR 34218, Sept. 2, off flight path under § 29.63.

1988] (4) At the end of the time interval prescribed in paragraph (a)(3) of this § 29.1047 Takeoff cooling test proce- section, the power must be reduced to dures.

maximum continuous power and the (a) Category A. For each category A climb must be continued for at least rotorcraft, cooling must be shown dur- five minutes after the occurance of the ing takeoff and subsequent climb as highest temperature recorded.

follows: (5) The cooling test must be con- (1) Each temperature must be sta- ducted at an airspeed corresponding to bilized while hovering in ground effect normal operating practice for the con- with— figuration being tested. However, if the (i) The power necessary for hovering; cooling provisions are sensitive to (ii) The appropriate cowl flap and rotorcraft speed, the most critical air- shutter settings; and speed must be used, but need not ex- (iii) The maximum weight. ceed the speed for best rate of climb with maximum continuous power.

(2) After the temperatures have sta- bilized, a climb must be started at the [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as lowest practicable altitude and must be amended by Amdt. 29–1, 30 FR 8778, July 13, conducted with one engine inoperative.

1965; Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] (3) The operating engines must be at § 29.1049 Hovering cooling test proce- the greatest power for which approval dures.

is sought (or at full throttle when above the critical altitude) for the The hovering cooling provisions must same period as this power is used in de- be shown— termining the takeoff climbout path (a) At maximum weight or at the under § 29.59.

greatest weight at which the rotorcraft (4) At the end of the time interval can hover (if less), at sea level, with prescribed in paragraph (b)(3) of this the power required to hover but not section, the power must be changed to more than maximum continuous that used in meeting § 29.67(a)(2) and power, in the ground effect in still air, the climb must be continued for— until at least five minutes after the oc- (i) Thirty minutes, if 30-minute OEI currence of the highest temperature re- power is used; or corded; and Federal Aviation Administration, DOT § 29.1093 (b) With maximum continuous power, free of visible moisture at a tempera- maximum weight, and at the altitude ture of 30 ° F., and with the engines at resulting in zero rate of climb for this 60 percent of maximum continuous configuration, until at least five min- power— utes after the occurrence of the highest (1) Each rotorcraft with sea level en- temperature recorded. gines using conventional venturi car- buretors has a preheater that can pro- I NDUCTION S YSTEM vide a heat rise of 90 ° F.; (2) Each rotorcraft with sea level en- § 29.1091 Air induction.

gines using carburetors tending to pre- (a) The air induction system for each vent icing has a preheater that can engine and auxiliary power unit must provide a heat rise of 70 ° F.; supply the air required by that engine (3) Each rotorcraft with altitude en- and auxiliary power unit under the op- gines using conventional venturi car- erating conditions for which certifi- buretors has a preheater that can pro- cation is requested.

vide a heat rise of 120 ° F.; and (b) Each engine and auxiliary power (4) Each rotorcraft with altitude en- unit air induction system must provide gines using carburetors tending to pre- air for proper fuel metering and mix- vent icing has a preheater that can ture distribution with the induction provide a heat rise of 100 ° F.

system valves in any position.

(b) Turbine engines. (1) It must be (c) No air intake may open within shown that each turbine engine and its the engine accessory section or within air inlet system can operate through- other areas of any powerplant compart- out the flight power range of the en- ment where emergence of backfire gine (including idling)— flame would constitute a fire hazard.

(i) Without accumulating ice on en- (d) Each reciprocating engine must gine or inlet system components that have an alternate air source.

would adversely affect engine oper- (e) Each alternate air intake must be ation or cause a serious loss of power located to prevent the entrance of rain, under the icing conditions specified in ice, or other foreign matter.

appendix C of this Part; and (f) For turbine engine powered rotor- (ii) In snow, both falling and blowing, craft and rotorcraft incorporating aux- without adverse effect on engine oper- iliary power units— ation, within the limitations estab- (1) There must be means to prevent lished for the rotorcraft.

hazardous quantities of fuel leakage or (2) Each turbine engine must idle for overflow from drains, vents, or other 30 minutes on the ground, with the air components of flammable fluid systems bleed available for engine icing protec- from entering the engine or auxiliary tion at its critical condition, without power unit intake system; and adverse effect, in an atmosphere that is (2) The air inlet ducts must be lo- at a temperature between 15 ° and 30 ° F cated or protected so as to minimize (between ¥ 9 ° and ¥ 1 ° C) and has a liq- the ingestion of foreign matter during uid water content not less than 0.3 takeoff, landing, and taxiing.

grams per cubic meter in the form of drops having a mean effective diameter (Secs. 313(a), 601, 603, 604, Federal Aviation not less than 20 microns, followed by Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 momentary operation at takeoff power U.S.C. 1655(c))) or thrust. During the 30 minutes of idle operation, the engine may be run up [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as periodically to a moderate power or amended by Amdt. 29–3, 33 FR 969, Jan. 26, 1968; Amdt. 29–17, 43 FR 50601, Oct. 30, 1978] thrust setting in a manner acceptable to the Administrator.

§ 29.1093 Induction system icing pro- (c) Supercharged reciprocating engines.

tection.

For each engine having a supercharger (a) Reciprocating engines. Each recip- to pressurize the air before it enters rocating engine air induction system the carburetor, the heat rise in the air must have means to prevent and elimi- caused by that supercharging at any nate icing. Unless this is done by other altitude may be utilized in determining means, it must be shown that, in air compliance with paragraph (a) of this

14 CFR Ch. I (1–1–25 Edition) § 29.1101 section if the heat rise utilized is that and entering any other compartment which will be available, automatically, or area of the rotorcraft in which a for the applicable altitude and oper- hazard would be created resulting from ation condition because of super- the entry of hot gases. The materials charging.

used to form the remainder of the in- duction system duct and plenum cham- (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 ber of the auxiliary power unit must be U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 U.S.C. 1655 (c)) capable of resisting the maximum heat conditions likely to occur.

[Amdt. 29–3, 33 FR 969, Jan. 26, 1968, as amended by Amdt. 29–12, 41 FR 55473, Dec. 20, (f) Each auxiliary power unit induc- 1976; Amdt. 29–13, 42 FR 15046, Mar. 17, 1977; tion system duct must be constructed Amdt. 29–22, 49 FR 6850, Feb. 23, 1984; Amdt.

of materials that will not absorb or 29–26, 53 FR 34219, Sept. 2, 1988] trap hazardous quantities of flammable fluids that could be ignited in the § 29.1101 Carburetor air preheater de- event of a surge or reverse flow condi- sign.

tion.

Each carburetor air preheater must be designed and constructed to— (Secs. 313(a), 601, 603, 604, Federal Aviation (a) Ensure ventilation of the pre- Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), heater when the engine is operated in sec. 6(c), Dept. of Transportation Act (49 cold air; U.S.C. 1655(c))) (b) Allow inspection of the exhaust [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as manifold parts that it surrounds; and amended by Amdt. 29–17, 43 FR 50602, Oct. 30, (c) Allow inspection of critical parts 1978] of the preheater itself.

§ 29.1105 Induction system screens.

§ 29.1103 Induction systems ducts and air duct systems. If induction system screens are used— (a) Each induction system duct up- (a) Each screen must be upstream of stream of the first stage of the engine the carburetor; supercharger and of the auxiliary (b) No screen may be in any part of power unit compressor must have a the induction system that is the only drain to prevent the hazardous accu- mulation of fuel and moisture in the passage through which air can reach ground attitude. No drain may dis- the engine, unless it can be deiced by charge where it might cause a fire haz- heated air; ard.

(c) No screen may be deiced by alco- (b) Each duct must be strong enough hol alone; and to prevent induction system failure (d) It must be impossible for fuel to from normal backfire conditions.

strike any screen.

(c) Each duct connected to compo- nents between which relative motion § 29.1107 Inter-coolers and after-cool- could exist must have means for flexi- ers.

bility.

Each inter-cooler and after-cooler (d) Each duct within any fire zone for must be able to withstand the vibra- which a fire-extinguishing system is re- tion, inertia, and air pressure loads to quired must be at least— which it would be subjected in oper- (1) Fireproof, if it passes through any ation.

firewall; or (2) Fire resistant, for other ducts, ex- § 29.1109 Carburetor air cooling.

cept that ducts for auxiliary power units must be fireproof within the aux- It must be shown under § 29.1043 that iliary power unit fire zone.

each installation using two-stage su- (e) Each auxiliary power unit induc- perchargers has means to maintain the tion system duct must be fireproof for air temperature, at the carburetor a sufficient distance upstream of the inlet, at or below the maximum estab- auxiliary power unit compartment to lished value.

prevent hot gas reverse flow from burn- ing through auxiliary power unit ducts Federal Aviation Administration, DOT § 29.1125 E XHAUST S YSTEM § 29.1123 Exhaust piping.

(a) Exhaust piping must be heat and § 29.1121 General.

corrosion resistant, and must have pro- For powerplant and auxiliary power visions to prevent failure due to expan- unit installations the following apply: sion by operating temperatures.

(a) Each exhaust system must ensure (b) Exhaust piping must be supported safe disposal of exhaust gases without to withstand any vibration and inertia fire hazard or carbon monoxide con- loads to which it would be subjected in tamination in any personnel compart- operation.

ment.

(c) Exhaust piping connected to com- (b) Each exhaust system part with a ponents between which relative motion surface hot enough to ignite flammable could exist must have provisions for fluids or vapors must be located or flexibility.

shielded so that leakage from any sys- tem carrying flammable fluids or va- § 29.1125 Exhaust heat exchangers.

pors will not result in a fire caused by For reciprocating engine powered impingement of the fluids or vapors on rotorcraft the following apply: any part of the exhaust system includ- (a) Each exhaust heat exchanger ing shields for the exhaust system.

must be constructed and installed to (c) Each component upon which hot withstand the vibration, inertia, and exhaust gases could impinge, or that other loads to which it would be sub- could be subjected to high tempera- jected in operation. In addition— tures from exhaust system parts, must be fireproof. Each exhaust system com- (1) Each exchanger must be suitable ponent must be separated by a fire- for continued operation at high tem- proof shield from adjacent parts of the peratures and resistant to corrosion rotorcraft that are outside the engine from exhaust gases; and auxiliary power unit compart- (2) There must be means for inspect- ments.

ing the critical parts of each ex- (d) No exhaust gases may discharge changer; so as to cause a fire hazard with re- (3) Each exchanger must have cooling spect to any flammable fluid vent or provisions wherever it is subject to drain.

contact with exhaust gases; and (e) No exhaust gases may discharge (4) No exhaust heat exchanger or where they will cause a glare seriously muff may have stagnant areas or liquid affecting pilot vision at night.

traps that would increase the prob- (f) Each exhaust system component ability of ignition of flammable fluids must be ventilated to prevent points of or vapors that might be present in case excessively high temperature.

of the failure or malfunction of compo- (g) Each exhaust shroud must be ven- nents carrying flammable fluids.

tilated or insulated to avoid, during (b) If an exhaust heat exchanger is normal operation, a temperature high used for heating ventilating air used by enough to ignite any flammable fluids personnel— or vapors outside the shroud.

(1) There must be a secondary heat (h) If significant traps exist, each exchanger between the primary ex- turbine engine exhaust system must haust gas heat exchanger and the ven- have drains discharging clear of the tilating air system; or rotorcraft, in any normal ground and (2) Other means must be used to pre- flight attitudes, to prevent fuel accu- vent harmful contamination of the mulation after the failure of an at- ventilating air.

tempted engine start.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (Secs. 313(a), 601, and 603, 72 Stat. 752, 755, 49 amended by Amdt. 29–12, 41 FR 55473, Dec. 20, U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 1976; Amdt. 29–41, 62 FR 46173, Aug. 29, 1997] U.S.C. 1655 (c)) [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–3, 33 FR 970, Jan. 26, 1968; Amdt. 29–13, 42 FR 15046, Mar. 17, 1977] 14 CFR Ch. I (1–1–25 Edition) § 29.1141 P OWERPLANT CONTROLS AND (b) Power controls must be arranged A CCESSORIES to allow ready synchronization of all engines by— § 29.1141 Powerplant controls: general.

(1) Separate control of each engine; and (a) Powerplant controls must be lo- (2) Simultaneous control of all en- cated and arranged under § 29.777 and gines.

marked under § 29.1555.

(c) Each power control must provide (b) Each control must be located so a positive and immediately responsive that it cannot be inadvertently oper- means of controlling its engine.

ated by persons entering, leaving, or (d) Each fluid injection control other moving normally in the cockpit.

than fuel system control must be in (c) Each flexible powerplant control the corresponding power control. How- must be approved.

ever, the injection system pump may (d) Each control must be able to have a separate control.

maintain any set position without— (e) If a power control incorporates a (1) Constant attention; or fuel shutoff feature, the control must (2) Tendency to creep due to control have a means to prevent the inad- loads or vibration.

vertent movement of the control into (e) Each control must be able to the shutoff position. The means must— withstand operating loads without ex- (1) Have a positive lock or stop at the cessive deflection.

idle position; and (f) Controls of powerplant valves re- (2) Require a separate and distinct quired for safety must have— operation to place the control in the (1) For manual valves, positive stops shutoff position.

or in the case of fuel valves suitable (f) For rotorcraft to be certificated index provisions, in the open and closed for a 30-second OEI power rating, a position; and means must be provided to automati- (2) For power-assisted valves, a cally activate and control the 30-sec- means to indicate to the flight crew ond OEI power and prevent any engine when the valve— from exceeding the installed engine (i) Is in the fully open or fully closed limits associated with the 30-second position; or OEI power rating approved for the (ii) Is moving between the fully open rotorcraft.

and fully closed position.

[Amdt. 29–26, 53 FR 34219, Sept. 2, 1988, as (Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 amended by Amdt. 29–34, 59 FR 47768, Sept.

U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 16, 1994] U.S.C. 1655(c)) § 29.1145 Ignition switches.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–13, 42 FR 15046, Mar. 17, (a) Ignition switches must control 1977; Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] each ignition circuit on each engine.

(b) There must be means to quickly § 29.1142 Auxiliary power unit con- shut off all ignition by the grouping of trols.

switches or by a master ignition con- Means must be provided on the flight trol.

deck for starting, stopping, and emer- (c) Each group of ignition switches, gency shutdown of each installed auxil- except ignition switches for turbine en- iary power unit.

gines for which continuous ignition is not required, and each master ignition (Secs. 313(a), 601, 603, 604, Federal Aviation control must have a means to prevent Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 its inadvertent operation.

U.S.C. 1655(c))) (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 [Amdt. 29–17, 43 FR 50602, Oct. 30, 1978] U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 U.S.C. 1655 (c)) § 29.1143 Engine controls.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (a) There must be a separate power amended by Amdt. 29–13, 42 FR 15046, Mar. 17, control for each engine. 1977] Federal Aviation Administration, DOT § 29.1165 vided for accessory drives located on § 29.1147 Mixture controls.

any component of the transmission and (a) If there are mixture controls, rotor drive system to prevent damage each engine must have a separate con- to these components from excessive ac- trol, and the controls must be arranged cessory load.

to allow— (1) Separate control of each engine; [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as and amended by Amdt. 29–22, 49 FR 6850, Feb. 23, (2) Simultaneous control of all en- 1984; Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] gines.

(b) Each intermediate position of the § 29.1165 Engine ignition systems.

mixture controls that corresponds to a (a) Each battery ignition system normal operating setting must be iden- must be supplemented with a generator tifiable by feel and sight.

that is automatically available as an § 29.1151 Rotor brake controls. alternate source of electrical energy to allow continued engine operation if (a) It must be impossible to apply the any battery becomes depleted.

rotor brake inadvertently in flight.

(b) The capacity of batteries and gen- (b) There must be means to warn the erators must be large enough to meet crew if the rotor brake has not been completely released before takeoff. the simultaneous demands of the en- gine ignition system and the greatest § 29.1157 Carburetor air temperature demands of any electrical system com- controls.

ponents that draw from the same There must be a separate carburetor source.

air temperature control for each en- (c) The design of the engine ignition gine.

system must account for— (1) The condition of an inoperative § 29.1159 Supercharger controls.

generator; Each supercharger control must be (2) The condition of a completely de- accessible to— pleted battery with the generator run- (a) The pilots; or ning at its normal operating speed; and (b) (If there is a separate flight engi- (3) The condition of a completely de- neer station with a control panel) the pleted battery with the generator oper- flight engineer.

ating at idling speed, if there is only § 29.1163 Powerplant accessories.

one battery.

(a) Each engine mounted accessory (d) Magneto ground wiring (for sepa- must— rate ignition circuits) that lies on the (1) Be approved for mounting on the engine side of any firewall must be in- engine involved; stalled, located, or protected, to mini- (2) Use the provisions on the engine mize the probability of the simulta- for mounting; and neous failure of two or more wires as a (3) Be sealed in such a way as to pre- result of mechanical damage, electrical vent contamination of the engine oil fault, or other cause.

system and the accessory system.

(e) No ground wire for any engine (b) Electrical equipment subject to may be routed through a fire zone of arcing or sparking must be installed, another engine unless each part of that to minimize the probability of igniting wire within that zone is fireproof.

flammable fluids or vapors.

(f) Each ignition system must be (c) If continued rotation of an engine- independent of any electrical circuit driven cabin supercharger or any re- that is not used for assisting, control- mote accessory driven by the engine will be a hazard if they malfunction, ling, or analyzing the operation of that there must be means to prevent their system.

hazardous rotation without interfering (g) There must be means to warn ap- with the continued operation of the en- propriate crewmembers if the malfunc- gine.

tioning of any part of the electrical (d) Unless other means are provided, torque limiting means must be pro- 14 CFR Ch. I (1–1–25 Edition) § 29.1181 system is causing the continuous dis- nents must be shielded or located so as charge of any battery necessary for en- to safeguard against the ignition of gine ignition. leaking flammable fluid. An integral oil sump of less than 25-quart capacity [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as on a reciprocating engine need not be amended by Amdt. 29–12, 41 FR 55473, Dec. 20, fireproof nor be enclosed by a fireproof 1976] shield.

P OWERPLANT F IRE P ROTECTION (b) Paragraph (a) of this section does not apply to— § 29.1181 Designated fire zones: re- (1) Lines, fittings, and components gions included.

which are already approved as part of a (a) Designated fire zones are— type certificated engine; and (1) The engine power section of recip- (2) Vent and drain lines, and their fit- rocating engines; tings, whose failure will not result in (2) The engine accessory section of or add to, a fire hazard.

reciprocating engines; [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (3) Any complete powerplant com- amended by Amdt. 29–2, 32 FR 6914, May 5, partment in which there is no isolation 1967; Amdt. 29–10, 39 FR 35463, Oct. 1, 1974; between the engine power section and Amdt. 29–22, 49 FR 6850, Feb. 23, 1984] the engine accessory section, for recip- § 29.1185 Flammable fluids.

rocating engines; (4) Any auxiliary power unit com- (a) No tank or reservoir that is part partment; of a system containing flammable (5) Any fuel-burning heater and other fluids or gases may be in a designated combustion equipment installation de- fire zone unless the fluid contained, the scribed in § 29.859; design of the system, the materials (6) The compressor and accessory sec- used in the tank and its supports, the tions of turbine engines; and shutoff means, and the connections, (7) The combustor, turbine, and tail- lines, and controls provide a degree of pipe sections of turbine engine instal- safety equal to that which would exist lations except sections that do not con- if the tank or reservoir were outside tain lines and components carrying such a zone.

flammable fluids or gases and are iso- (b) Each fuel tank must be isolated lated from the designated fire zone pre- from the engines by a firewall or scribed in paragraph (a)(6) of this sec- shroud.

tion by a firewall that meets § 29.1191.

(c) There must be at least one-half (b) Each designated fire zone must inch of clear airspace between each meet the requirements of §§ 29.1183 tank or reservoir and each firewall or through 29.1203.

shroud isolating a designated fire zone, unless equivalent means are used to [Amdt. 29–3, 33 FR 970, Jan. 26, 1968, as prevent heat transfer from the fire amended by Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] zone to the flammable fluid.

(d) Absorbent material close to flam- § 29.1183 Lines, fittings, and compo- mable fluid system components that nents.

might leak must be covered or treated (a) Except as provided in paragraph to prevent the absorption of hazardous (b) of this section, each line, fitting, quantities of fluids.

and other component carrying flam- § 29.1187 Drainage and ventilation of mable fluid in any area subject to en- fire zones.

gine fire conditions and each compo- nent which conveys or contains flam- (a) There must be complete drainage mable fluid in a designated fire zone of each part of each designated fire must be fire resistant, except that zone to minimize the hazards resulting flammable fluid tanks and supports in from failure or malfunction of any a designated fire zone must be fireproof component containing flammable or be enclosed by a fireproof shield un- fluids. The drainage means must be— less damage by fire to any non-fire- (1) Effective under conditions ex- proof part will not cause leakage or pected to prevail when drainage is spillage of flammable fluid. Compo- needed; and Federal Aviation Administration, DOT § 29.1193 (2) Arranged so that no discharged (e) Each shutoff valve and its control fluid will cause an additional fire haz- must be designed, located, and pro- ard. tected to function properly under any (b) Each designated fire zone must be condition likely to result from fire in a ventilated to prevent the accumulation designated fire zone.

of flammable vapors.

(f) Except for ground-use-only auxil- (c) No ventilation opening may be iary power unit installations, there where it would allow the entry of flam- must be means to prevent inadvertent mable fluids, vapors, or flame from operation of each shutoff and to make other zones.

it possible to reopen it in flight after it (d) Ventilation means must be ar- has been closed.

ranged so that no discharged vapors [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as will cause an additional fire hazard.

amended by Amdt. 29–12, 41 FR 55473, Dec. 20, (e) For category A rotorcraft, there 1976; Amdt. 29–22, 49 FR 6850, Feb. 23, 1984; must be means to allow the crew to Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] shut off the sources of forced ventila- tion in any fire zone (other than the § 29.1191 Firewalls.

engine power section of the powerplant (a) Each engine, including the com- compartment) unless the amount of ex- bustor, turbine, and tailpipe sections of tinguishing agent and the rate of dis- turbine engine installations, must be charge are based on the maximum air- isolated by a firewall, shroud, or equiv- flow through that zone.

alent means, from personnel compart- § 29.1189 Shutoff means.

ments, structures, controls, rotor mechanisms, and other parts that are— (a) There must be means to shut off (1) Essential to controlled flight and or otherwise prevent hazardous quan- landing; and tities of fuel, oil, de-icing fluid, and other flammable fluids from flowing (2) Not protected under § 29.861.

into, within, or through any designated (b) Each auxiliary power unit, com- fire zone, except that this means need bustion heater, and other combustion not be provided— equipment to be used in flight, must be (1) For lines, fittings, and compo- isolated from the rest of the rotorcraft nents forming an integral part of an by firewalls, shrouds, or equivalent engine; means.

(2) For oil systems for turbine engine (c) Each firewall or shroud must be installations in which all components constructed so that no hazardous quan- of the system, including oil tanks, are tity of air, fluid, or flame can pass fireproof or located in areas not subject from any engine compartment to other to engine fire conditions; or parts of the rotorcraft.

(3) For engine oil systems in category (d) Each opening in the firewall or B rotorcraft using reciprocating en- shroud must be sealed with close-fit- gines of less than 500 cubic inches dis- ting fireproof grommets, bushings, or placement.

firewall fittings.

(b) The closing of any fuel shutoff (e) Each firewall and shroud must be valve for any engine may not make fireproof and protected against corro- fuel unavailable to the remaining en- sion.

gines.

(f) In meeting this section, account (c) For category A rotorcraft, no haz- must be taken of the probable path of ardous quantity of flammable fluid a fire as affected by the airflow in nor- may drain into any designated fire mal flight and in autorotation.

zone after shutoff has been accom- plished, nor may the closing of any fuel [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–3, 33 FR 970, Jan. 26, shutoff valve for an engine make fuel 1968] unavailable to the remaining engines.

(d) The operation of any shutoff may § 29.1193 Cowling and engine compart- not interfere with the later emergency ment covering.

operation of any other equipment, such as the means for declutching the en- (a) Each cowling and engine compart- gine from the rotor drive. ment covering must be constructed and 14 CFR Ch. I (1–1–25 Edition) § 29.1194 supported so that it can resist the vi- zones, other than tail surfaces not sub- bration, inertia, and air loads to which ject to heat, flames, or sparks ema- it may be subjected in operation. nating from a designated fire zone or (b) Cowling must meet the drainage engine compartment, must be at least and ventilation requirements of fire resistant.

§ 29.1187.

[Amdt. 29–3, 33 FR 970, Jan. 26, 1968] (c) On rotorcraft with a diaphragm isolating the engine power section from § 29.1195 Fire extinguishing systems.

the engine accessory section, each part (a) Each turbine engine powered of the accessory section cowling sub- rotorcraft and Category A recipro- ject to flame in case of fire in the en- cating engine powered rotorcraft, and gine power section of the powerplant each Category B reciprocating engine must— (1) Be fireproof; and powered rotorcraft with engines of (2) Meet the requirements of § 29.1191. more than 1,500 cubic inches must have (d) Each part of the cowling or engine a fire extinguishing system for the des- compartment covering subject to high ignated fire zones. The fire extin- temperatures due to its nearness to ex- guishing system for a powerplant must haust system parts or exhaust gas im- be able to simultaneously protect all pingement must be fireproof. zones of the powerplant compartment (e) Each rotorcraft must— for which protection is provided.

(1) Be designated and constructed so (b) For multiengine powered rotor- that no fire originating in any fire zone craft, the fire extinguishing system, can enter, either through openings or the quantity of extinguishing agent, by burning through external skin, any and the rate of discharge must— other zone or region where it would (1) For each auxiliary power unit and create additional hazards; combustion equipment, provide at least (2) Meet the requirements of para- one adequate discharge; and graph (e)(1) of this section with the (2) For each other designated fire landing gear retracted (if applicable); zone, provide two adequate discharges.

and (c) For single engine rotorcraft, the (3) Have fireproof skin in areas sub- quantity of extinguishing agent and ject to flame if a fire starts in or burns the rate of discharge must provide at out of any designated fire zone.

least one adequate discharge for the (f) A means of retention for each engine compartment.

openable or readily removable panel, (d) It must be shown by either actual cowling, or engine or rotor drive sys- or simulated flight tests that under tem covering must be provided to pre- critical airflow conditions in flight the clude hazardous damage to rotors or discharge of the extinguishing agent in critical control components in the each designated fire zone will provide event of— an agent concentration capable of ex- (1) Structural or mechanical failure tinguishing fires in that zone and of of the normal retention means, unless minimizing the probability of reigni- such failure is extremely improbable; tion.

or (2) Fire in a fire zone, if such fire (Secs. 313(a), 601, 603, 604, Federal Aviation could adversely affect the normal Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 means of retention.

U.S.C. 1655(c))) (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as U.S.C. 1655(c)) amended by Amdt. 29–3, 33 FR 970, Jan. 26, 1968; Amdt. 29–13, 42 FR 15047, Mar. 17, 1977; [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as Amdt. 29–17, 43 FR 50602, Oct. 30, 1978] amended by Amdt. 29–3, 33 FR 970, Jan. 26, 1968; Amdt. 29–13, 42 FR 15046, Mar. 17, 1977; § 29.1197 Fire extinguishing agents.

Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] (a) Fire extinguishing agents must— § 29.1194 Other surfaces.

(1) Be capable of extinguishing All surfaces aft of, and near, engine flames emanating from any burning of compartments and designated fire fluids or other combustible materials Federal Aviation Administration, DOT § 29.1203 in the area protected by the fire extin- § 29.1201 Fire extinguishing system guishing system; and materials.

(2) Have thermal stability over the (a) No materials in any fire extin- temperature range likely to be experi- guishing system may react chemically enced in the compartment in which with any extinguishing agent so as to they are stored.

create a hazard.

(b) If any toxic extinguishing agent is (b) Each system component in an en- used, it must be shown by test that gine compartment must be fireproof.

entry of harmful concentrations of fluid or fluid vapors into any personnel § 29.1203 Fire detector systems.

compartment (due to leakage during (a) For each turbine engine powered normal operation of the rotorcraft, or rotorcraft and Category A recipro- discharge on the ground or in flight) is cating engine powered rotorcraft, and prevented, even though a defect may for each Category B reciprocating en- exist in the extinguishing system.

gine powered rotorcraft with engines of (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 more than 900 cubic inches displace- U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 ment, there must be approved, quick- U.S.C. 1655(c)) acting fire detectors in designated fire [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as zones and in the combustor, turbine, amended by Amdt. 29–12, 41 FR 55473, Dec. 20, and tailpipe sections of turbine instal- 1976; Amdt. 29–13, 42 FR 15047, Mar. 17, 1977] lations (whether or not such sections are designated fire zones) in numbers § 29.1199 Extinguishing agent con- and locations ensuring prompt detec- tainers.

tion of fire in those zones.

(a) Each extinguishing agent con- (b) Each fire detector must be con- tainer must have a pressure relief to structed and installed to withstand any prevent bursting of the container by vibration, inertia, and other loads to excessive internal pressures.

which it would be subjected in oper- (b) The discharge end of each dis- ation.

charge line from a pressure relief con- (c) No fire detector may be affected nection must be located so that dis- by any oil, water, other fluids, or charge of the fire extinguishing agent fumes that might be present.

would not damage the rotorcraft. The line must also be located or protected (d) There must be means to allow to prevent clogging caused by ice or crewmembers to check, in flight, the other foreign matter.

functioning of each fire detector sys- (c) There must be a means for each tem electrical circuit.

fire extinguishing agent container to (e) The writing and other components indicate that the container has dis- of each fire detector system in an en- charged or that the charging pressure gine compartment must be at least fire is below the established minimum nec- resistant.

essary for proper functioning.

(f) No fire detector system compo- (d) The temperature of each con- nent for any fire zone may pass tainer must be maintained, under in- through another fire zone, unless— tended operating conditions, to prevent (1) It is protected against the possi- the pressure in the container from— bility of false warnings resulting from (1) Falling below that necessary to fires in zones through which it passes; provide an adequate rate of discharge; or or (2) The zones involved are simulta- (2) Rising high enough to cause pre- neously protected by the same detector mature discharge.

and extinguishing systems.

(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 amended by Amdt. 29–3, 33 FR 970, Jan. 26, U.S.C. 1655 (c)) 1968] [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–13, 42 FR 15047, Mar. 17, 1977]

14 CFR Ch. I (1–1–25 Edition) § 29.1301 (6) Is located on the instrument panel Subpart F—Equipment in a position acceptable to the Admin- G ENERAL istrator that will make it plainly visi- ble to and useable by any pilot at his § 29.1301 Function and installation.

station; and (7) Is appropriately lighted during all Each item of installed equipment phases of operation.

must— (a) Be of a kind and design appro- (h) A gyroscopic direction indicator.

priate to its intended function; (i) A rate-of-climb (vertical speed) in- (b) Be labeled as to its identification, dicator.

function, or operating limitations, or (j) For Category A rotorcraft, a speed any applicable combination of these warning device when V is less than NE factors; the speed at which unmistakable over- (c) Be installed according to limita- speed warning is provided by other tions specified for that equipment; and pilot cues. The speed warning device (d) Function properly when installed.

must give effective aural warning (dif- fering distinctively from aural warn- § 29.1303 Flight and navigation instru- ings used for other purposes) to the pi- ments.

lots whenever the indicated speed ex- The following are required flight and ceeds V plus 3 knots and must oper- NE navigational instruments: ate satisfactorily throughout the ap- (a) An airspeed indicator. For Cat- proved range of altitudes and tempera- egory A rotorcraft with V less than a NE tures.

speed at which unmistakable pilot cues (Secs. 313(a), 601, 603, 604, and 605 of the Fed- provide overspeed warning, a maximum eral Aviation Act of 1958 (49 U.S.C. 1354(a), allowable airspeed indicator must be 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

provided. If maximum allowable air- of Transportation Act (49 U.S.C. 1655(c))) speed varies with weight, altitude, [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as temperature, or r.p.m., the indicator amended by Amdt. 29–12, 41 FR 55474, Dec. 20, must show that variation.

1976; Amdt. 29–14, 42 FR 36972, July 18, 1977; (b) A sensitive altimeter.

Amdt. 29–24, 49 FR 44438, Nov. 6, 1984; 70 FR (c) A magnetic direction indicator.

2012, Jan. 12, 2005] (d) A clock displaying hours, min- utes, and seconds with a sweep-second § 29.1305 Powerplant instruments.

pointer or digital presentation.

The following are required power- (e) A free-air temperature indicator.

plant instruments: (f) A non-tumbling gyroscopic bank (a) For each rotorcraft— and pitch indicator.

(1) A carburetor air temperature indi- (g) A gyroscopic rate-of-turn indi- cator for each reciprocating engine; cator combined with an integral slip- (2) A cylinder head temperature indi- skid indicator (turn-and-bank indi- cator for each air-cooled reciprocating cator) except that only a slip-skid indi- engine, and a coolant temperature indi- cator is required on rotorcraft with a cator for each liquid-cooled recipro- third attitude instrument system cating engine; that— (3) A fuel quantity indicator for each (1) Is usable through flight attitudes fuel tank; of ± 80 degrees of pitch and ± 120 degrees (4) A low fuel warning device for each of roll; fuel tank which feeds an engine. This (2) Is powered from a source inde- device must— pendent of the electrical generating (i) Provide a warning to the crew system; when approximately 10 minutes of usa- (3) Continues reliable operation for a ble fuel remains in the tank; and minimum of 30 minutes after total fail- (ii) Be independent of the normal fuel ure of the electrical generating system; quantity indicating system.

(4) Operates independently of any other attitude indicating system; (5) A means to indicate manifold (5) Is operative without selection pressure for each altitude engine; after total failure of the electrical gen- (6) An oil pressure indicator for each erating system; pressure-lubricated gearbox.

Federal Aviation Administration, DOT § 29.1305 (7) An oil pressure warning device for not employ any pumps, filters, or other each pressure-lubricated gearbox to in- components subject to degradation or dicate when the oil pressure falls below failure which may adversely affect fuel a safe value; pressure at the engine; (8) An oil quantity indicator for each (22) A means to indicate to the oil tank and each rotor drive gearbox, flightcrew the failure of any fuel pump if lubricant is self-contained; installed to show compliance with (9) An oil temperature indicator for § 29.955; each engine; (23) Warning or caution devices to (10) An oil temperature warning de- signal to the flightcrew when ferro- vice to indicate unsafe oil tempera- magnetic particles are detected by the tures in each main rotor drive gearbox, chip detector required by § 29.1337(e); including gearboxes necessary for rotor and phasing; (24) For auxiliary power units, an in- (11) A means to indicate the gas tem- dividual indicator, warning or caution perature for each turbine engine; device, or other means to advise the (12) A means to indicate the gas pro- flightcrew that limits are being exceed- ducer speed for each turbine engine; ed, if exceeding these limits can be haz- (13) A tachometer for each engine ardous, for— that, if combined with the applicable (i) Gas temperature; instrument required by paragraph (ii) Oil pressure; and (a)(14) of this section, indicates rotor (iii) Rotor speed.

r.p.m. during autorotation.

(25) For rotorcraft for which a 30-sec- (14) At least one tachometer to indi- ond/2-minute OEI power rating is re- cate, as applicable— (i) The r.p.m. of the single main quested, a means must be provided to rotor; alert the pilot when the engine is at (ii) The common r.p.m. of any main the 30-second and 2-minute OEI power rotors whose speeds cannot vary appre- levels, when the event begins, and ciably with respect to each other; and when the time interval expires.

(iii) The r.p.m. of each main rotor (26) For each turbine engine utilizing whose speed can vary appreciably with 30-second/2-minute OEI power, a device respect to that of another main rotor; or system must be provided for use by (15) A free power turbine tachometer ground personnel which— for each turbine engine; (i) Automatically records each usage (16) A means, for each turbine engine, and duration of power at the 30-second to indicate power for that engine; and 2-minute OEI levels; (17) For each turbine engine, an indi- (ii) Permits retrieval of the recorded cator to indicate the functioning of the data; powerplant ice protection system; (iii) Can be reset only by ground (18) An indicator for the filter re- maintenance personnel; and quired by § 29.997 to indicate the occur- (iv) Has a means to verify proper op- rence of contamination of the filter to eration of the system or device.

the degree established in compliance (b) For category A rotorcraft— with § 29.955; (1) An individual oil pressure indi- (19) For each turbine engine, a warn- cator for each engine, and either an ing means for the oil strainer or filter independent warning device for each required by § 29.1019, if it has no bypass, engine or a master warning device for to warn the pilot of the occurrence of the engines with means for isolating contamination of the strainer or filter the individual warning circuit from the before it reaches the capacity estab- master warning device; lished in accordance with § 29.1019(a)(2); (2) An independent fuel pressure (20) An indicator to indicate the func- warning device for each engine or a tioning of any selectable or control- master warning device for all engines lable heater used to prevent ice clog- with provision for isolating the indi- ging of fuel system components; vidual warning device from the master (21) An individual fuel pressure indi- warning device; and cator for each engine, unless the fuel system which supplies that engine does (3) Fire warning indicators.

14 CFR Ch. I (1–1–25 Edition) § 29.1307 (4) For each Category A rotorcraft (1) The occurrence of any cata- for which OEI Training Mode is re- strophic failure condition is extremely quested, a means must be provided to improbable; indicate to the pilot the simulation of (2) The occurrence of any major fail- ure condition is no more than improb- an engine failure, the annunciation of able; and that simulation, and a representation (3) For the occurrence of any other of the OEI power being provided.

failure condition in between major and (c) For category B rotorcraft— catastrophic, the probability of the (1) An individual oil pressure indi- failure condition must be inversely cator for each engine; and (2) Fire warning indicators, when fire proportional to its consequences.

detection is required. (c) A means to alert the crew in the event of a failure must be provided [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as when an unsafe system operating con- amended by Amdt. 29–3, 33 FR 970, Jan. 26, dition exists and to enable them to 1968; Amdt. 29–10, 39 FR 35463, Oct. 1, 1974; take corrective action. Systems, con- Amdt. 29–26, 53 FR 34219, Sept. 2, 1988; Amdt.

trols, and associated monitoring and 29–34, 59 FR 47768, Sept. 16, 1994; Amdt. 29–40, 61 FR 21908, May 10, 1996; 61 FR 43952, Aug. 27, crew alerting means must be designed 1996; Amdt. 29–59, 88 FR 8739, Feb. 10, 2023] to minimize crew errors that could cre- ate additional hazards.

§ 29.1307 Miscellaneous equipment.

(d) Compliance with the require- The following is required miscella- ments of this section must be shown by neous equipment: analysis and, where necessary, by (a) An approved seat for each occu- ground, flight, or simulator tests. The pant. analysis must account for: (b) A master switch arrangement for (1) Possible modes of failure, includ- electrical circuits other than ignition.

ing malfunctions and misleading data (c) Hand fire extinguishers.

and input from external sources; (d) A windshield wiper or equivalent (2) The effect of multiple failures and device for each pilot station.

latent failures; (e) A two-way radio communication (3) The resulting effects on the rotor- system. craft and occupants, considering the stage of flight and operating condi- [Amdt. 29–12, 41 FR 55473, Dec. 20, 1976] tions; and (4) The crew alerting cues and the § 29.1309 Equipment, systems, and in- stallations. corrective action required.

The equipment, systems, and instal- [Amdt. 29–59, 88 FR 8739, Feb. 10, 2023] lations whose functioning is required § 29.1316 Electrical and electronic sys- by this subchapter must be designed tem lightning protection.

and installed to ensure that they per- form their intended functions under (a) Each electrical and electronic any foreseeable operating condition.

system that performs a function, for For any item of equipment or system which failure would prevent the contin- whose failure has not been specifically ued safe flight and landing of the rotor- addressed by another requirement in craft, must be designed and installed so this chapter, the following require- that— ments also apply: (1) The function is not adversely af- (a) The design of each item of equip- fected during and after the time the ment, system, and installation must be rotorcraft is exposed to lightning; and analyzed separately and in relation to (2) The system automatically recov- other rotorcraft systems and installa- ers normal operation of that function tions to determine and identify any in a timely manner after the rotorcraft failure that would affect the capability is exposed to lightning.

of the rotorcraft or the ability of the (b) Each electrical and electronic crew to perform their duties in all op- system that performs a function, for erating conditions. which failure would reduce the capa- (b) Each item of equipment, system, bility of the rotorcraft or the ability of and installation must be designed and the flightcrew to respond to an adverse installed so that: operating condition, must be designed Federal Aviation Administration, DOT § 29.1321 and installed so that the function re- versely affected when the equipment covers normal operation in a timely providing these functions is exposed to manner after the rotorcraft is exposed equipment HIRF test level 3, as de- to lightning. scribed in appendix E to this part.

(d) Before December 1, 2012, an elec- [Doc. No. FAA–2010–0224, Amdt. 29–53, 76 FR trical or electronic system that per- 33136, June 8, 2011] forms a function whose failure would § 29.1317 High-intensity Radiated prevent the continued safe flight and Fields (HIRF) Protection.

landing of a rotorcraft may be designed and installed without meeting the pro- (a) Except as provided in paragraph visions of paragraph (a) provided— (d) of this section, each electrical and (1) The system has previously been electronic system that performs a func- shown to comply with special condi- tion whose failure would prevent the continued safe flight and landing of the tions for HIRF, prescribed under § 21.16, rotorcraft must be designed and in- issued before December 1, 2007; stalled so that— (2) The HIRF immunity characteris- (1) The function is not adversely af- tics of the system have not changed fected during and after the time the since compliance with the special con- rotorcraft is exposed to HIRF environ- ditions was demonstrated; and ment I, as described in appendix E to (3) The data used to demonstrate this part; compliance with the special conditions (2) The system automatically recov- is provided.

ers normal operation of that function, [Doc. No. FAA–2006–23657, 72 FR 44027, Aug. 6, in a timely manner, after the rotor- 2007] craft is exposed to HIRF environment I, as described in appendix E to this I NSTRUMENTS : I NSTALLATION part, unless this conflicts with other operational or functional requirements § 29.1321 Arrangement and visibility.

of that system; (a) Each flight, navigation, and pow- (3) The system is not adversely af- erplant instrument for use by any pilot fected during and after the time the must be easily visible to him from his rotorcraft is exposed to HIRF environ- station with the minimum practicable ment II, as described in appendix E to deviation from his normal position and this part; and line of vision when he is looking for- (4) Each function required during op- ward along the flight path.

eration under visual flight rules is not (b) Each instrument necessary for adversely affected during and after the safe operation, including the airspeed time the rotorcraft is exposed to HIRF indicator, gyroscopic direction indi- environment III, as described in appen- cator, gyroscopic bank-and-pitch indi- dix E to this part.

cator, slip-skid indicator, altimeter, (b) Each electrical and electronic rate-of-climb indicator, rotor tachom- system that performs a function whose eters, and the indicator most rep- failure would significantly reduce the resentative of engine power, must be capability of the rotorcraft or the abil- grouped and centered as nearly as prac- ity of the flightcrew to respond to an ticable about the vertical plane of the adverse operating condition must be pilot’s forward vision. In addition, for designed and installed so the system is rotorcraft approved for IFR flight— not adversely affected when the equip- (1) The instrument that most effec- ment providing these functions is ex- tively indicates attitude must be on posed to equipment HIRF test level 1 the panel in the top center position; or 2, as described in appendix E to this part. (2) The instrument that most effec- (c) Each electrical and electronic sys- tively indicates direction of flight tem that performs such a function must be adjacent to and directly below whose failure would reduce the capa- the attitude instrument; bility of the rotorcraft or the ability of (3) The instrument that most effec- the flightcrew to respond to an adverse tively indicates airspeed must be adja- operating condition must be designed cent to and to the left of the attitude and installed so the system is not ad- instrument; and 14 CFR Ch. I (1–1–25 Edition) § 29.1322 (4) The instrument that most effec- ard atmosphere) with a minimum prac- tively indicates altitude or is most fre- ticable instrument calibration error quently utilized in control of altitude when the corresponding pitot and stat- must be adjacent to and to the right of ic pressures are applied.

the attitude instrument. (b) Each system must be calibrated (c) Other required powerplant instru- to determine system error excluding ments must be closely grouped on the airspeed instrument error. This cali- instrument panel.

bration must be determined— (d) Identical powerplant instruments (1) In level flight at speeds of 20 for the engines must be located so as to knots and greater, and over an appro- prevent any confusion as to which en- priate range of speeds for flight condi- gine each instrument relates.

tions of climb and autorotation; and (e) Each powerplant instrument vital (2) During takeoff, with repeatable to safe operation must be plainly visi- and readable indications that ensure— ble to appropriate crewmembers.

(i) Consistent realization of the field (f) Instrument panel vibration may lengths specified in the Rotorcraft not damage, or impair the readability Flight Manual; and or accuracy of, any instrument.

(ii) Avoidance of the critical areas of (g) If a visual indicator is provided to the height-velocity envelope as estab- indicate malfunction of an instrument, lished under § 29.87.

it must be effective under all probable (c) For Category A rotorcraft— cockpit lighting conditions.

(1) The indication must allow con- sistent definition of the takeoff deci- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), sion point; and 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

(2) The system error, excluding the of Transportation Act (49 U.S.C. 1655(c))) airspeed instrument calibration error, may not exceed— [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–14, 42 FR 36972, July 18, (i) Three percent or 5 knots, which- 1977; Amdt. 29–21, 48 FR 4391, Jan. 31, 1983] ever is greater, in level flight at speeds above 80 percent of takeoff safety § 29.1322 Warning, caution, and advi- speed; and sory lights.

(ii) Ten knots in climb at speeds from If warning, caution or advisory lights 10 knots below takeoff safety speed to are installed in the cockpit they must, 10 knots above V .

Y unless otherwise approved by the Ad- (d) For Category B rotorcraft, the ministrator, be— system error, excluding the airspeed (a) Red, for warning lights (lights in- instrument calibration error, may not dicating a hazard which may require exceed 3 percent or 5 knots, whichever immediate corrective action); is greater, in level flight at speeds (b) Amber, for caution lights (lights above 80 percent of the climbout speed indicating the possible need for future attained at 50 feet when complying corrective action); with § 29.63.

(c) Green, for safe operation lights; (e) Each system must be arranged, so and far as practicable, to prevent malfunc- (d) Any other color, including white, tion or serious error due to the entry of for lights not described in paragraphs moisture, dirt, or other substances.

(a) through (c) of this section, provided (f) Each system must have a heated the color differs sufficiently from the pitot tube or an equivalent means of colors prescribed in paragraphs (a) preventing malfunction due to icing.

through (c) of this section to avoid pos- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as sible confusion.

amended by Amdt. 29–3, 33 FR 970, Jan. 26, [Amdt. 29–12, 41 FR 55474, Dec. 20, 1976] 1968; Amdt. 29–24, 49 FR 44439, Nov. 6, 1984; Amdt. 29–39, 61 FR 21901, May 10, 1996; Amdt.

§ 29.1323 Airspeed indicating system.

29–44, 64 FR 45338, Aug. 19, 1999] For each airspeed indicating system, § 29.1325 Static pressure and pressure the following apply: altimeter systems.

(a) Each airspeed indicating instru- ment must be calibrated to indicate (a) Each instrument with static air true airspeed (at sea level with a stand- case connections must be vented to the Federal Aviation Administration, DOT § 29.1329 outside atmosphere through an appro- static pressure source being open or priate piping system. blocked.

(b) Each vent must be located where (Secs. 313(a), 601, 603, 604, and 605 of the Fed- its orifices are least affected by airflow eral Aviation Act of 1958 (49 U.S.C. 1354(a), variation, moisture, or foreign matter. 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

of Transportation Act (49 U.S.C. 1655(c))) (c) Each static pressure port must be designed and located in such manner [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–14, 42 FR 36972, July 18, that the correlation between air pres- 1977; Amdt. 29–24, 49 FR 44439, Nov. 6, 1984] sure in the static pressure system and true ambient atmospheric static pres- § 29.1327 Magnetic direction indicator.

sure is not altered when the rotorcraft (a) Each magnetic direction indicator encounters icing conditions. An anti- must be installed so that its accuracy icing means or an alternate source of is not excessively affected by the static pressure may be used in showing rotorcraft’s vibration or magnetic compliance with this requirement. If fields.

the reading of the altimeter, when on (b) The compensated installation the alternate static pressure system, may not have a deviation, in level differs from the reading of altimeter flight, greater than 10 degrees on any when on the primary static system by heading.

more than 50 feet, a correction card must be provided for the alternate § 29.1329 Automatic pilot and flight guidance system.

static system.

(d) Except for the vent into the at- For the purpose of this subpart, an mosphere, each system must be air- automatic pilot and flight guidance system may consist of an autopilot, tight.

flight director, or a component that (e) Each pressure altimeter must be interacts with stability augmentation approved and calibrated to indicate or trim.

pressure altitude in a standard atmos- (a) Each automatic pilot and flight phere with a minimum practicable guidance system must be designed so calibration error when the cor- that it: responding static pressures are applied.

(1) Can be overpowered by one pilot (f) Each system must be designed and to allow control of the rotorcraft; installed so that an error in indicated (2) Provides a means to disengage the pressure altitude, at sea level, with a system, or any malfunctioning compo- standard atmosphere, excluding instru- nent of the system, by each pilot to ment calibration error, does not result prevent it from interfering with the in an error of more than ± 30 feet per 100 control of the rotorcraft; and knots speed. However, the error need (3) Provides a means to indicate to not be less than ± 30 feet.

the flight crew its current mode of op- (g) Except as provided in paragraph eration. Selector switch position is not (h) of this section, if the static pressure acceptable as a means of indication.

(b) Unless there is automatic syn- system incorporates both a primary chronization, each system must have a and an alternate static pressure source, means to readily indicate to the pilot the means for selecting one or the the alignment of the actuating device other source must be designed so in relation to the control system it op- that— erates.

(1) When either source is selected, the (c) Each manually operated control other is blocked off; and for the system’s operation must be (2) Both sources cannot be blocked readily accessible to the pilots.

off simultaneously.

(d) The system must be designed so (h) For unpressurized rotorcraft, that, within the range of adjustment paragraph (g)(1) of this section does not available to the pilot, it cannot apply if it can be demonstrated that produce hazardous loads on the rotor- the static pressure system calibration, craft, or create hazardous deviations in when either static pressure source is the flight path, under any flight condi- selected, is not changed by the other tion appropriate to its use or in the 14 CFR Ch. I (1–1–25 Edition) § 29.1331 event of a malfunction, assuming that first pilot may be connected to that op- corrective action begins within a rea- erating system.

sonable period of time.

(b) The equipment, systems, and in- (e) If the automatic pilot and flight stallations must be designed so that guidance system integrates signals one display of the information essen- from auxiliary controls or furnishes tial to the safety of flight which is pro- signals for operation of other equip- vided by the flight instruments re- ment, there must be a means to pre- mains available to a pilot, without ad- vent improper operation.

ditional crewmember action, after any (f) If the automatic pilot system can single failure or combination of fail- be coupled to airborne navigation ures that are not shown to be ex- equipment, means must be provided to tremely improbable.

indicate to the pilots the current mode (c) Additional instruments, systems, of operation. Selector switch position or equipment may not be connected to is not acceptable as a means of indica- the operating system for a second pilot tion.

unless provisions are made to ensure [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as the continued normal functioning of amended by Amdt. 29–24, 49 FR 44439, Nov. 6, the required flight instruments in the 1984; Amdt. 29–24, 49 FR 47594, Dec. 6, 1984; event of any malfunction of the addi- Amdt. 29–42, 63 FR 43285, Aug. 12, 1998; Amdt.

29–59, 88 FR 8739, Feb. 10, 2023] tional instruments, systems, or equip- ment which is not shown to be ex- § 29.1331 Instruments using a power tremely improbable.

supply.

[Amdt. 29–24, 49 FR 44439, Nov. 6, 1984, as For category A rotorcraft— amended by Amdt. 29–59, 88 FR 8740, Feb. 10, (a) Each required flight instrument 2023] using a power supply must have— (1) Two independent sources of power; § 29.1337 Powerplant instruments.

(2) A means of selecting either power (a) Instruments and instrument lines.

source; and (3) A visual means integral with each (1) Each powerplant and auxiliary instrument to indicate when the power power unit instrument line must meet adequate to sustain proper instrument the requirements of §§ 29.993 and 29.1183.

performance is not being supplied. The (2) Each line carrying flammable power must be measured at or near the fluids under pressure must— point where it enters the instrument.

(i) Have restricting orifices or other For electrical instruments, the power safety devices at the source of pressure is considered to be adequate when the to prevent the escape of excessive fluid voltage is within the approved limits; if the line fails; and and (ii) Be installed and located so that (b) The installation and power supply the escape of fluids would not create a system must be such that failure of hazard.

any flight instrument connected to one (3) Each powerplant and auxiliary source, or of the energy supply from power unit instrument that utilizes one source, or a fault in any part of the flammable fluids must be installed and power distribution system does not located so that the escape of fluid interfere with the proper supply of en- would not create a hazard.

ergy from any other source.

(b) Fuel quantity indicator. There [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as must be means to indicate to the flight amended by Amdt. 29–24, 49 FR 44439, Nov. 6, crew members the quantity, in gallons 1984] or equivalent units, of usable fuel in § 29.1333 Instrument systems.

each tank during flight. In addition— (1) Each fuel quantity indicator must For systems that operate the re- be calibrated to read ‘‘zero’’ during quired flight instruments which are lo- level flight when the quantity of fuel cated at each pilot’s station, the fol- remaining in the tank is equal to the lowing apply: (a) For pneumatic systems, only the unusable fuel supply determined under required flight instruments for the § 29.959; Federal Aviation Administration, DOT § 29.1351 (2) When two or more tanks are close- mission cables, and associated control, ly interconnected by a gravity feed sys- regulation, and protective devices. It tem and vented, and when it is impos- must be designed so that— sible to feed from each tank sepa- (1) Power sources function properly rately, at least one fuel quantity indi- when independent and when connected cator must be installed; in combination; (3) Tanks with interconnected outlets (2) No failure or malfunction of any and airspaces may be treated as one power source can create a hazard or tank and need not have separate indi- impair the ability of remaining sources cators; and to supply essential loads; (4) Each exposed sight gauge used as (3) The system voltage and frequency a fuel quantity indicator must be pro- (as applicable) at the terminals of es- tected against damage.

sential load equipment can be main- (c) Fuel flowmeter system. If a fuel tained within the limits for which the flowmeter system is installed, each equipment is designed, during any metering component must have a probable operating condition; means for bypassing the fuel supply if (4) System transients due to switch- malfunction of that component se- ing, fault clearing, or other causes do verely restricts fuel flow.

not make essential loads inoperative, (d) Oil quantity indicator. There must and do not cause a smoke or fire haz- be a stick gauge or equivalent means ard; to indicate the quantity of oil— (5) There are means accessible in (1) In each tank; and flight to appropriate crewmembers for (2) In each transmission gearbox.

the individual and collective dis- (e) Rotor drive system transmissions connection of the electrical power and gearboxes utilizing ferromagnetic sources from the main bus; and materials must be equipped with chip (6) There are means to indicate to ap- detectors designed to indicate the pres- propriate crewmembers the generating ence of ferromagnetic particles result- system quantities essential for the safe ing from damage or excessive wear operation of the system, such as the within the transmission or gearbox.

voltage and current supplied by each Each chip detector must— generator.

(1) Be designed to provide a signal to (c) External power. If provisions are the indicator required by made for connecting external power to § 29.1305(a)(22); and the rotorcraft, and that external power (2) Be provided with a means to allow can be electrically connected to equip- crewmembers to check, in flight, the ment other than that used for engine function of each detector electrical cir- starting, means must be provided to cuit and signal.

ensure that no external power supply (Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 having a reverse polarity, or a reverse U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 phase sequence, can supply power to U.S.C. 1655(c)) the rotorcraft’s electrical system.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (d) Operation with the normal elec- amended by Amdt. 29–13, 42 FR 15047, Mar. 17, trical power generating system inoper- 1977; Amdt. 29–26, 53 FR 34219, Sept. 2, 1988] ative.

E LECTRICAL S YSTEMS AND E QUIPMENT (1) It must be shown by analysis, tests, or both, that the rotorcraft can § 29.1351 General.

be operated safely in VFR conditions (a) Electrical system capacity. The re- for a period of not less than 5 minutes, quired generating capacity and the with the normal electrical power gen- number and kind of power sources erating system (electrical power must— sources excluding the battery) inoper- (1) Be determined by an electrical ative, with critical type fuel (from the load analysis; and standpoint of flameout and restart ca- (2) Meet the requirements of § 29.1309. pability), and with the rotorcraft ini- (b) Generating system. The generating tially at the maximum certificated al- system includes electrical power titude. Parts of the electrical system sources, main power busses, trans- may remain on if— 14 CFR Ch. I (1–1–25 Edition) § 29.1353 (i) A single malfunction, including a toxic gases, smoke, or fluids that could wire bundle or junction box fire, can- accumulate in hazardous quantities not result in loss of the part turned off within the rotorcraft.

and the part turned on; (c) Corrosive fluids or gases that es- (ii) The parts turned on are elec- cape from the system must not damage trically and mechanically isolated surrounding structures, adjacent equip- from the parts turned off; and ment, or systems necessary for contin- (2) Additional requirements for Cat- ued safe flight and landing.

egory A Rotorcraft. (d) The maximum amount of heat (i) Unless it can be shown that the and pressure that can be generated dur- loss of the normal electrical power gen- ing any operation or under any failure erating system is extremely improb- condition of the energy storage system able, an emergency electrical power or its individual components must not system, independent of the normal result in any hazardous effect on rotor- electrical power generating system, craft structure, equipment, or systems must be provided, with sufficient ca- necessary for continued safe flight and pacity to power all systems necessary landing.

for continued safe flight and landing. (e) Energy storage system installa- (ii) Failures, including junction box, tions required for continued safe flight control panel, or wire bundle fires, and landing of the rotorcraft must which would result in the loss of the have monitoring features and a means normal and emergency systems, must to indicate to the pilot the status of all be shown to be extremely improbable. critical system parameters.

(iii) Systems necessary for imme- [Amdt. 29–59, 88 FR 8740, Feb. 10, 2023] diate safety must continue to operate following the loss of the normal elec- § 29.1355 Distribution system.

trical power generating system, with- (a) The distribution system includes out the need for flight crew action.

the distribution busses, their associ- (e) Electrical equipment, controls, ated feeders, and each control and pro- and wiring must be installed so that tective device.

operation of any one unit or system of (b) If two independent sources of units will not adversely affect the si- electrical power for particular equip- multaneous operation of any other ment or systems are required by this electrical unit or system essential to chapter, in the event of the failure of safe operation.

one power source for such equipment or (f) Cables must be grouped, routed, system, another power source (includ- and spaced so that damage to essential ing its separate feeder) must be pro- circuits will be minimized if there are vided automatically or be manually se- faults in heavy current-carrying ca- lectable to maintain equipment or sys- bles.

tem operation.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

of Transportation Act (49 U.S.C. 1655(c))) of Transportation Act (49 U.S.C. 1655(c))) [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–14, 42 FR 36973, July 18, amended by Amdt. 29–14, 42 FR 36973, July 18, 1977; Amdt. 29–40, 61 FR 21908, May 10, 1996; 1977; Amdt. 29–24, 49 FR 44439, Nov. 6, 1984] Amdt. 29–42, 63 FR 43285, Aug. 12, 1998; Amdt.

29–59, 88 FR 8740, Feb. 10, 2023] § 29.1357 Circuit protective devices.

§ 29.1353 Energy storage systems.

(a) Automatic protective devices Energy storage systems must be de- must be used to minimize distress to signed and installed as follows: the electrical system and hazard to the (a) Energy storage systems must pro- rotorcraft system and hazard to the vide automatic protective features for rotorcraft in the event of wiring faults any conditions that could prevent con- or serious malfunction of the system or tinued safe flight and landing. connected equipment.

(b) Energy storage systems must not (b) The protective and control de- emit any flammable, explosive, or vices in the generating system must be Federal Aviation Administration, DOT § 29.1385 designed to de-energize and disconnect (1) The tests must be performed on a faulty power sources and power trans- mock-up using the same generating mission equipment from their associ- equipment used in the rotorcraft; ated buses with sufficient rapidity to (2) The equipment must simulate the provide protection from hazardous electrical characteristics of the dis- overvoltage and other malfunctioning.

tribution wiring and connected loads to (c) Each resettable circuit protective the extent necessary for valid test re- device must be designed so that, when sults; and an overload or circuit fault exists, it (3) Laboratory generator drives must will open the circuit regardless of the simulate the prime movers on the position of the operating control.

rotorcraft with respect to their reac- (d) If the ability to reset a circuit tion to generator loading, including breaker or replace a fuse is essential to loading due to faults.

safety in flight, that circuit breaker or (b) For each flight condition that fuse must be located and identified so cannot be simulated adequately in the that it can be readily reset or replaced laboratory or by ground tests on the in flight.

rotorcraft, flight tests must be made.

(e) Each essential load must have in- dividual circuit protection. However, L IGHTS individual protection for each circuit in an essential load system (such as § 29.1381 Instrument lights.

each position light circuit in a system) The instrument lights must— is not required.

(a) Make each instrument, switch, (f) If fuses are used, there must be and other device for which they are spare fuses for use in flight equal to at provided easily readable; and least 50 percent of the number of fuses (b) Be installed so that— of each rating required for complete (1) Their direct rays are shielded circuit protection.

from the pilot’s eyes; and (g) Automatic reset circuit breakers (2) No objectionable reflections are may be used as integral protectors for visible to the pilot.

electrical equipment provided there is circuit protection for the cable sup- § 29.1383 Landing lights.

plying power to the equipment.

(a) Each required landing or hovering [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as light must be approved.

amended by Amdt. 29–24, 49 FR 44440, Nov. 6, (b) Each landing light must be in- 1984] stalled so that— § 29.1359 Electrical system fire and (1) No objectionable glare is visible smoke protection.

to the pilot; (a) Components of the electrical sys- (2) The pilot is not adversely affected tem must meet the applicable fire and by halation; and smoke protection provisions of §§ 29.831 (3) It provides enough light for night and 29.863. operation, including hovering and land- (b) Electrical cables, terminals, and ing.

equipment, in designated fire zones, (c) At least one separate switch must and that are used in emergency proce- be provided, as applicable— dures, must be at least fire resistant.

(1) For each separately installed (c) Insulation on electrical wire and landing light; and cable installed in the rotorcraft must (2) For each group of landing lights be self-extinguishing when tested in ac- installed at a common location.

cordance with Appendix F, Part I(a)(3), of part 25 of this chapter. § 29.1385 Position light system installa- tion.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29–42, 63 FR 43285, Aug. 12, (a) General. Each part of each posi- 1998] tion light system must meet the appli- cable requirements of this section and § 29.1363 Electrical system tests.

each system as a whole must meet the (a) When laboratory tests of the elec- requirements of §§ 29.1387 through trical system are conducted— 29.1397.

14 CFR Ch. I (1–1–25 Edition) § 29.1387 (b) Forward position lights. Forward that dihedral angle, if such solid angle position lights must consist of a red is within a cone whose apex is at the and a green light spaced laterally as rear position light and whose elements far apart as practicable and installed make an angle of 30 ° with a vertical forward on the rotorcraft so that, with line passing through the rear position the rotorcraft in the normal flying po- light.

sition, the red light is on the left side, (49 U.S.C. 1655(c)) and the green light is on the right side.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as Each light must be approved.

amended by Amdt. 29–9, 36 FR 21279, Nov. 5, (c) Rear position light. The rear posi- 1971] tion light must be a white light mount- ed as far aft as practicable, and must § 29.1389 Position light distribution be approved.

and intensities.

(d) Circuit. The two forward position (a) General. The intensities prescribed lights and the rear position light must in this section must be provided by new make a single circuit.

equipment with light covers and color (e) Light covers and color filters. Each filters in place. Intensities must be de- light cover or color filter must be at termined with the light source oper- least flame resistant and may not ating at a steady value equal to the av- change color or shape or lose any ap- erage luminous output of the source at preciable light transmission during the normal operating voltage of the normal use.

rotorcraft. The light distribution and § 29.1387 Position light system dihe- intensity of each position light must dral angles.

meet the requirements of paragraph (b) of this section.

(a) Except as provided in paragraph (b) Forward and rear position lights.

(e) of this section, each forward and The light distribution and intensities rear position light must, as installed, of forward and rear position lights show unbroken light within the dihe- must be expressed in terms of min- dral angles described in this section.

imum intensities in the horizontal (b) Dihedral angle L (left) is formed plane, minimum intensities in any by two intersecting vertical planes, the vertical plane, and maximum inten- first parallel to the longitudinal axis of sities in overlapping beams, within di- the rotorcraft, and the other at 110 de- hedral angles, L, R, and A, and must grees to the left of the first, as viewed meet the following requirements: when looking forward along the longi- (1) Intensities in the horizontal plane.

tudinal axis.

Each intensity in the horizontal plane (c) Dihedral angle R (right) is formed (the plane containing the longitudinal by two intersecting vertical planes, the axis of the rotorcraft and perpendicular first parallel to the longitudinal axis of to the plane of symmetry of the rotor- the rotorcraft, and the other at 110 de- craft), must equal or exceed the values grees to the right of the first, as viewed in § 29.1391.

when looking forward along the longi- (2) Intensities in any vertical plane.

tudinal axis.

Each intensity in any vertical plane (d) Dihedral angle A (aft) is formed (the plane perpendicular to the hori- by two intersecting vertical planes zontal plane) must equal or exceed the making angles of 70 degrees to the appropriate value in § 29.1393 where I is right and to the left, respectively, to a the minimum intensity prescribed in vertical plane passing through the lon- § 29.1391 for the corresponding angles in gitudinal axis, as viewed when looking the horizontal plane.

aft along the longitudinal axis.

(e) If the rear position light, when (3) Intensities in overlaps between adja- mounted as far aft as practicable in ac- cent signals. No intensity in any over- cordance with § 29.1385(c), cannot show lap between adjacent signals may ex- unbroken light within dihedral angle A ceed the values in § 29.1395, except that (as defined in paragraph (d) of this sec- higher intensities in overlaps may be tion), a solid angle or angles of ob- used with the use of main beam inten- structed visibility totaling not more sities substantially greater than the than 0.04 steradians is allowable within minima specified in §§ 29.1391 and Federal Aviation Administration, DOT § 29.1401 29.1393 if the overlap intensities in rela- (b) Area B includes all directions in tion to the main beam intensities do the adjacent dihedral angle that pass not adversely affect signal clarity. through the light source and intersect the common boundary plane at more § 29.1391 Minimum intensities in the than 20 degrees.

horizontal plane of forward and rear position lights.

§ 29.1397 Color specifications.

Each position light intensity must Each position light color must have equal or exceed the applicable values in the applicable International Commis- the following table: sion on Illumination chromaticity co- ordinates as follows: Angle from right or left (a) Aviation red — Dihedral angle (light in- of longitudinal axis, Intensity cluded) measured from dead (candles) y is not greater than 0.335; and ahead z is not greater than 0.002.

and green). 10 ° to 20 ° ................... 30 (b) Aviation green — x is not greater than 0.440 ¥ 0.320 y ; x is not greater than y ¥ 0.170; and y is not less than 0.390 ¥ 0.170 x.

§ 29.1393 Minimum intensities in any (c) Aviation white — vertical plane of forward and rear position lights.

x is not less than 0.300 and not greater than 0.540; Each position light intensity must y is not less than x ¥ 0.040 or y c ¥ 0.010, equal or exceed the applicable values in whichever is the smaller; and the following table: y is not greater than x + 0.020 nor 0.636 ¥ 0.400 x ; Angle above or below the horizontal plane Intensity, I Where Y is the y coordinate of the Planck- e 0 ° ......................................................................... 1.00 ian radiator for the value of x considered.

0 ° to 5 ° ................................................................ .90 [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as 5 ° to 10 ° .............................................................. .80 amended by Amdt. 29–7, 36 FR 12972, July 10, 10 ° to 15 ° ............................................................ .70 15 ° to 20 ° ............................................................ .50 1971] 20 ° to 30 ° ............................................................ .30 30 ° to 40 ° ............................................................ .10 § 29.1399 Riding light.

40 ° to 90 ° ............................................................ .05 (a) Each riding light required for water operation must be installed so § 29.1395 Maximum intensities in over- that it can— lapping beams of forward and rear (1) Show a white light for at least position lights.

two miles at night under clear atmos- No position light intensity may ex- pheric conditions; and ceed the applicable values in the fol- (2) Show a maximum practicable un- lowing table, except as provided in broken light with the rotorcraft on the § 29.1389(b)(3).

water.

(b) Externally hung lights may be Maximum intensity used.

Overlaps Area A Area B (candles) (candles) § 29.1401 Anticollision light system.

Green in dihedral angle L ......... 10 1 (a) General. If certification for night Red in dihedral angle R ............ 10 1 operation is requested, the rotorcraft Green in dihedral angle A ......... 5 1 Red in dihedral angle A ............ 5 1 must have an anticollision light sys- Rear white in dihedral angle L .. 5 1 tem that— Rear white in dihedral angle R 5 1 (1) Consists of one or more approved anticollision lights located so that Where— their emitted light will not impair the (a) Area A includes all directions in crew’s vision or detract from the con- the adjacent dihedral angle that pass spicuity of the position lights; and through the light source and intersect (2) Meets the requirements of para- the common boundary plane at more graphs (b) through (f) of this section.

than 10 degrees but less than 20 de- (b) Field of coverage. The system must grees; and consist of enough lights to illuminate 14 CFR Ch. I (1–1–25 Edition) § 29.1411 the vital areas around the rotorcraft, Effective Angle above or below the horizontal plane intensity considering the physical configuration (candles) and flight characteristics of the rotor- craft. The field of coverage must ex- tend in each direction within at least 30 degrees above and 30 degrees below the horizontal plane of the rotorcraft, except that there may be solid angles [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as of obstructed visibility totaling not amended by Amdt. 29–7, 36 FR 12972, July 10, more than 0.5 steradians.

1971; Amdt. 29–11, 41 FR 5290, Feb. 5, 1976] (c) Flashing characteristics. The ar- rangement of the system, that is, the S AFETY E QUIPMENT number of light sources, beam width, § 29.1411 General.

speed of rotation, and other character- istics, must give an effective flash fre- (a) Accessibility. Required safety quency of not less than 40, nor more equipment to be used by the crew in an than 100, cycles per minute. The effec- emergency, such as automatic liferaft tive flash frequency is the frequency at releases, must be readily accessible.

which the rotorcraft’s complete anti- (b) Stowage provisions. Stowage provi- collision light system is observed from sions for required emergency equip- a distance, and applies to each sector ment must be furnished and must— of light including any overlaps that (1) Be arranged so that the equip- exist when the system consists of more ment is directly accessible and its loca- than one light source. In overlaps, tion is obvious; and flash frequencies may exceed 100, but (2) Protect the safety equipment not 180, cycles per minute.

from inadvertent damage.

(d) Color. Each anticollision light (c) Emergency exit descent device. The must be aviation red and must meet stowage provisions for the emergency the applicable requirements of § 29.1397.

exit descent device required by (e) Light intensity. The minimum § 29.809(f) must be at the exits for which light intensities in any vertical plane, they are intended.

measured with the red filter (if used) (d) Liferafts. Liferafts must be stowed and expressed in terms of ‘‘effective’’ near exits through which the rafts can intensities must meet the require- be launched during an unplanned ditch- ments of paragraph (f) of this section.

ing. Rafts automatically or remotely The following relation must be as- released outside the rotorcraft must be sumed: attached to the rotorcraft by the static line prescribed in § 29.1415.

t I t dt ( ) (e) Long-range signaling device. The

∫

t stowage provisions for the long-range I = e t t + − . ( ) 0 2 signaling device required by § 29.1415 2 1 must be near an exit available during where: an unplanned ditching.

I = effective intensity (candles).

e (f) Life preservers. Each life preserver I(t) = instantaneous intensity as a function must be within easy reach of each oc- of time.

cupant while seated.

t ¥ t = flash time interval (seconds).

2 l Normally, the maximum value of effective § 29.1413 Safety belts: passenger warn- intensity is obtained when t and t are cho- 2 1 ing device.

sen so that the effective intensity is equal to the instantaneous intensity at t and t .

2 1 (a) If there are means to indicate to the passengers when safety belts (f) Minimum effective intensities for should be fastened, they must be in- anticollision light. Each anticollision stalled to be operated from either pilot light effective intensity must equal or seat.

exceed the applicable values in the fol- lowing table: Federal Aviation Administration, DOT § 29.1433 (b) Each safety belt must be equipped pheric icing conditions and by one or with a metal to metal latching device. more of the following tests as found necessary to determine the adequacy of (Secs. 313, 314, and 601 through 610 of the Fed- the ice protection system: eral Aviation Act of 1958 (49 U.S.C. 1354, 1355, (1) Laboratory dry air or simulated and 1421 through 1430) and sec. 6(c), Dept. of icing tests, or a combination of both, of Transportation Act (49 U.S.C. 1655(c))) the components or models of the com- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ponents.

amended by Amdt. 29–16 43 FR 46233, Oct. 5, (2) Flight dry air tests of the ice pro- 1978] tection system as a whole, or its indi- § 29.1415 Ditching equipment. vidual components.

(3) Flight tests of the rotorcraft or (a) Emergency flotation and sig- its components in measured simulated naling equipment required by any oper- icing conditions.

ating rule of this chapter must meet (d) The ice protection provisions of the requirements of this section.

this section are considered to be appli- (b) Each liferaft and each life pre- cable primarily to the airframe. Power- server must be approved. In addition— plant installation requirements are (1) Provide not less than two rafts, of contained in Subpart E of this part.

an approximately equal rated capacity (e) A means must be identified or and buoyancy to accommodate the oc- provided for determining the formation cupants of the rotorcraft; and of ice on critical parts of the rotor- (2) Each raft must have a trailing craft. Unless otherwise restricted, the line, and must have a static line de- means must be available for nighttime signed to hold the raft near the rotor- as well as daytime operation. The craft but to release it if the rotorcraft rotorcraft flight manual must describe becomes totally submerged.

the means of determining ice forma- (c) Approved survival equipment tion and must contain information nec- must be attached to each liferaft.

essary for safe operation of the rotor- (d) There must be an approved sur- craft in icing conditions.

vival type emergency locator trans- mitter for use in one life raft.

[Amdt. 29–21, 48 FR 4391, Jan. 31, 1983] [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as M ISCELLANEOUS E QUIPMENT amended by Amdt. 29–8, 36 FR 18722, Sept. 21, 1971; Amdt. 29–19, 45 FR 38348, June 9, 1980; § 29.1431 Electronic equipment.

Amdt. 27–26, 55 FR 8005, Mar. 6, 1990; Amdt.

29–33, 59 FR 32057, June 21, 1994] (a) Radio communication and naviga- tion equipment installations must be § 29.1419 Ice protection.

free from hazards in themselves, in (a) To obtain certification for flight their method of operation, and in their into icing conditions, compliance with effects on other components, under any this section must be shown. critical environmental conditions.

(b) It must be demonstrated that the (b) Radio communication and naviga- rotorcraft can be safely operated in the tion equipment, controls, and wiring continuous maximum and intermittent must be installed so that operation of maximum icing conditions determined any one unit or system of units will under appendix C of this part within not adversely affect the simultaneous the rotorcraft altitude envelope. An operation of any other radio or elec- analysis must be performed to estab- tronic unit, or system of units, re- lish, on the basis of the rotorcraft’s quired by this chapter.

operational needs, the adequacy of the § 29.1433 Vacuum systems.

ice protection system for the various components of the rotorcraft. (a) There must be means, in addition (c) In addition to the analysis and to the normal pressure relief, to auto- physical evaluation prescribed in para- matically relieve the pressure in the graph (b) of this section, the effective- discharge lines from the vacuum air ness of the ice protection system and pump when the delivery temperature of its components must be shown by the air becomes unsafe.

flight tests of the rotorcraft or its com- (b) Each vacuum air system line and ponents in measured natural atmos- fitting on the discharge side of the 14 CFR Ch. I (1–1–25 Edition) § 29.1435 pump that might contain flammable fluid must meet the applicable require- vapors or fluids must meet the require- ments of §§ 29.861, 29.1183, 29.1185, and ments of § 29.1183 if they are in a des- 29.1189.

ignated fire zone.

§ 29.1439 Protective breathing equip- (c) Other vacuum air system compo- ment.

nents in designated fire zones must be at least fire resistant. (a) If one or more cargo or baggage compartments are to be accessible in § 29.1435 Hydraulic systems.

flight, protective breathing equipment must be available for an appropriate (a) Design. Each hydraulic system crewmember.

must be designed as follows: (b) For protective breathing equip- (1) Each element of the hydraulic ment required by paragraph (a) of this system must be designed to withstand, section or by any operating rule of this without detrimental, permanent defor- chapter— mation, any structural loads that may (1) That equipment must be designed be imposed simultaneously with the to protect the crew from smoke, carbon maximum operating hydraulic loads.

(2) Each element of the hydraulic dioxide, and other harmful gases while system must be designed to withstand on flight deck duty; (2) That equipment must include— pressures sufficiently greater than (i) Masks covering the eyes, nose, and those prescribed in paragraph (b) of mouth; or this section to show that the system (ii) Masks covering the nose and will not rupture under service condi- mouth, plus accessory equipment to tions.

protect the eyes; and (3) There must be means to indicate (3) That equipment must supply pro- the pressure in each main hydraulic tective oxygen of 10 minutes duration power system.

per crewmember at a pressure altitude (4) There must be means to ensure of 8,000 feet with a respiratory minute that no pressure in any part of the sys- volume of 30 liters per minute BTPD.

tem will exceed a safe limit above the maximum operating pressure of the § 29.1457 Cockpit voice recorders.

system, and to prevent excessive pres- sures resulting from any fluid volu- (a) Each cockpit voice recorder re- metric change in lines likely to remain quired by the operating rules of this closed long enough for such a change to chapter must be approved, and must be take place. The possibility of detri- installed so that it will record the fol- mental transient (surge) pressures dur- lowing: ing operation must be considered. (1) Voice communications trans- (5) Each hydraulic line, fitting, and mitted from or received in the rotor- component must be installed and sup- craft by radio.

ported to prevent excessive vibration (2) Voice communications of flight and to withstand inertia loads. Each crewmembers on the flight deck.

element of the installation must be (3) Voice communications of flight protected from abrasion, corrosion, and crewmembers on the flight deck, using mechanical damage. the rotorcraft’s interphone system.

(6) Means for providing flexibility (4) Voice or audio signals identifying must be used to connect points, in a navigation or approach aids introduced hydraulic fluid line, between which rel- into a headset or speaker.

ative motion or differential vibration (5) Voice communications of flight exists. crewmembers using the passenger loud- (b) Tests. Each element of the system speaker system, if there is such a sys- must be tested to a proof pressure of 1.5 tem, and if the fourth channel is avail- times the maximum pressure to which able in accordance with the require- that element will be subjected in nor- ments of paragraph (c)(4)(ii) of this sec- mal operation, without failure, mal- tion.

function, or detrimental deformation (6) If datalink communication equip- of any part of the system. ment is installed, all datalink commu- (c) Fire protection. Each hydraulic nications, using an approved data mes- system using flammable hydraulic sage set. Datalink messages must be Federal Aviation Administration, DOT § 29.1457 recorded as the output signal from the (iii) Each microphone on the flight communications unit that translates deck that is used with the rotorcraft’s the signal into usable data.

loudspeaker system if its signals are (b) The recording requirements of not picked up by another channel.

paragraph (a)(2) of this section may be (d) Each cockpit voice recorder must met— be installed so that— (1) By installing a cockpit-mounted (1)(i) It receives its electrical power area microphone, located in the best from the bus that provides the max- position for recording voice commu- imum reliability for operation of the nications originating at the first and cockpit voice recorder without jeopard- second pilot stations and voice commu- izing service to essential or emergency nications of other crewmembers on the loads.

flight deck when directed to those sta- (ii) It remains powered for as long as tions; or possible without jeopardizing emer- (2) By installing a continually ener- gized or voice-actuated lip microphone gency operation of the rotorcraft.

at the first and second pilot stations.

(2) There is an automatic means to simultaneously stop the recorder and The microphone specified in this para- prevent each erasure feature from func- graph must be so located and, if nec- tioning, within 10 minutes after crash essary, the preamplifiers and filters of impact; the recorder must be so adjusted or supplemented, that the recorded com- (3) There is an aural or visual means munications are intelligible when re- for preflight checking of the recorder corded under flight cockpit noise con- for proper operation; ditions and played back. The level of (4) Whether the cockpit voice re- intelligibility must be approved by the corder and digital flight data recorder Administrator. Repeated aural or vis- are installed in separate boxes or in a ual playback of the record may be used combination unit, no single electrical in evaluating intelligibility.

failure external to the recorder may (c) Each cockpit voice recorder must disable both the cockpit voice recorder be installed so that the part of the and the digital flight data recorder; communication or audio signals speci- and fied in paragraph (a) of this section ob- (5) It has an independent power tained from each of the following source— sources is recorded on a separate chan- (i) That provides 10 ± 1 minutes of nel: electrical power to operate both the (1) For the first channel, from each cockpit voice recorder and cockpit- microphone, headset, or speaker used mounted area microphone; at the first pilot station.

(ii) That is located as close as prac- (2) For the second channel, from each microphone, headset, or speaker used ticable to the cockpit voice recorder; at the second pilot station. and (3) For the third channel, from the (iii) To which the cockpit voice re- cockpit-mounted area microphone, or corder and cockpit-mounted area the continually energized or voice-ac- microphone are switched automati- tuated lip microphones at the first and cally in the event that all other power second pilot stations.

to the cockpit voice recorder is inter- (4) For the fourth channel, from— rupted either by normal shutdown or (i) Each microphone, headset, or by any other loss of power to the elec- speaker used at the stations for the trical power bus.

third and fourth crewmembers; or (e) The record container must be lo- (ii) If the stations specified in para- cated and mounted to minimize the graph (c)(4)(i) of this section are not re- probability of rupture of the container quired or if the signal at such a station as a result of crash impact and con- is picked up by another channel, each sequent heat damage to the record microphone on the flight deck that is from fire.

used with the passenger loudspeaker (f) If the cockpit voice recorder has a system if its signals are not picked up by another channel. bulk erasure device, the installation 14 CFR Ch. I (1–1–25 Edition) § 29.1459 must be designed to minimize the prob- disable both the cockpit voice recorder ability of inadvertent operation and ac- and the digital flight data recorder.

tuation of the device during crash im- (b) Each nonejectable recorder con- pact.

tainer must be located and mounted so (g) Each recorder container must be as to minimize the probability of con- either bright orange or bright yellow. tainer rupture resulting from crash im- (h) When both a cockpit voice re- pact and subsequent damage to the corder and a flight data recorder are record from fire.

required by the operating rules, one (c) A correlation must be established combination unit may be installed, between the flight recorder readings of provided that all other requirements of airspeed, altitude, and heading and the this section and the requirements for corresponding readings (taking into ac- flight data recorders under this part count correction factors) of the first pi- are met. lot’s instruments. This correlation must cover the airspeed range over [Amdt. 29–6, 35 FR 7293, May 9, 1970, as which the aircraft is to be operated, amended by Amdt. 29–50, 73 FR 12564, Mar. 7, the range of altitude to which the air- 2008; 74 FR 32800, July 9, 2009; Amdt. 29–52, 75 craft is limited, and 360 degrees of FR 17045, Apr. 5, 2010] heading. Correlation may be estab- § 29.1459 Flight data recorders. lished on the ground as appropriate.

(d) Each recorder container must: (a) Each flight recorder required by (1) Be either bright orange or bright the operating rules of Subchapter G of yellow; this chapter must be installed so that: (2) Have a reflective tape affixed to (1) It is supplied with airspeed, alti- its external surface to facilitate its lo- tude, and directional data obtained cation under water; and from sources that meet the accuracy (3) Have an underwater locating de- requirements of §§ 29.1323, 29.1325, and vice, when required by the operating 29.1327 of this part, as applicable; rules of this chapter, on or adjacent to (2) The vertical acceleration sensor is the container which is secured in such rigidly attached, and located longitu- a manner that it is not likely to be sep- dinally within the approved center of arated during crash impact.

gravity limits of the rotorcraft; (e) When both a cockpit voice re- (3)(i) It receives its electrical power corder and a flight data recorder are from the bus that provides the max- required by the operating rules, one imum reliability for operation of the combination unit may be installed, flight data recorder without jeopard- provided that all other requirements of izing service to essential or emergency this section and the requirements for loads.

cockpit voice recorders under this part (ii) It remains powered for as long as are met.

possible without jeopardizing emer- gency operation of the rotorcraft.

[Amdt. 29–25, 53 FR 26145, July 11, 1988; 53 FR (4) There is an aural or visual means 26144, July 11, 1988, as amended by Amdt. 29– for perflight checking of the recorder 50, 73 FR 12564, Mar. 7, 2008; 74 FR 32800, July 9, 2009; Amdt. 29–52, 75 FR 17045, Apr. 5, 2010] for proper recording of data in the stor- age medium; § 29.1461 Equipment containing high (5) Except for recorders powered sole- energy rotors.

ly by the engine-drive electrical gener- (a) Equipment containing high en- ator system, there is an automatic ergy rotors must meet paragraph (b), means to simultaneously stop a re- (c), or (d) of this section.

corder that has a data erasure feature (b) High energy rotors contained in and prevent each erasure feature from equipment must be able to withstand functioning, within 10 minutes after damage caused by malfunctions, vibra- any crash impact; and tion, abnormal speeds, and abnormal (6) Whether the cockpit voice re- temperatures. In addition— corder and digital flight data recorder are installed in separate boxes or in a (1) Auxiliary rotor cases must be able combination unit, no single electrical to contain damage caused by the fail- failure external to the recorder may ure of high energy rotor blades; and Federal Aviation Administration, DOT § 29.1505 (2) Equipment control devices, sys- (i) 0.9 times the maximum forward tems, and instrumentation must rea- speeds established under § 29.309; sonably ensure that no operating limi- (ii) 0.9 times the maximum speed tations affecting the integrity of high shown under §§ 29.251 and 29.629; or energy rotors will be exceeded in serv- (iii) 0.9 times the maximum speed ice.

substantiated for advancing blade tip (c) It must be shown by test that mach number effects under critical al- equipment containing high energy ro- titude conditions.

tors can contain any failure of a high (b) V may vary with altitude, NE energy rotor that occurs at the highest r.p.m., temperature, and weight, if— speed obtainable with the normal speed (1) No more than two of these vari- control devices inoperative.

ables (or no more than two instru- (d) Equipment containing high en- ments integrating more than one of ergy rotors must be located where these variables) are used at one time; rotor failure will neither endanger the and occupants nor adversely affect contin- (2) The ranges of these variables (or ued safe flight.

of the indications on instruments inte- [Amdt. 29–3, 33 FR 971, Jan. 26, 1968] grating more than one of these vari- ables) are large enough to allow an Subpart G—Operating Limitations operationally practical and safe vari- and Information ation of V .

NE (c) For helicopters, a stabilized § 29.1501 General.

power-off V denoted as V (power- NE NE (a) Each operating limitation speci- off) may be established at a speed less fied in §§ 29.1503 through 29.1525 and than V established pursuant to para- NE other limitations and information nec- graph (a) of this section, if the fol- essary for safe operation must be es- lowing conditions are met: tablished.

(1) V (power-off) is not less than a NE (b) The operating limitations and speed midway between the power-on other information necessary for safe V and the speed used in meeting the NE operation must be made available to requirements of— the crewmembers as prescribed in (i) § 29.67(a)(3) for Category A heli- §§ 29.1541 through 29.1589.

copters; (Secs. 313(a), 601, 603, 604, and 605 of the Fed- (ii) § 29.65(a) for Category B heli- eral Aviation Act of 1958 (49 U.S.C. 1354(a), copters, except multi-engine heli- 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

copters meeting the requirements of of Transportation Act (49 U.S.C. 1655(c))) § 29.67(b); and [Amdt. 29–15, 43 FR 2327, Jan. 16, 1978] (iii) § 29.67(b) for multi-engine Cat- egory B helicopters meeting the re- O PERATING L IMITATIONS quirements of § 29.67(b).

(2) V (power-off) is— NE § 29.1503 Airspeed limitations: general.

(i) A constant airspeed; (a) An operating speed range must be (ii) A constant amount less than established.

power-on V or NE ´ (b) When airspeed limitations are a (iii) A constant airspeed for a portion function of weight, weight distribution, of the altitude range for which certifi- altitude, rotor speed, power, or other cation is requested, and a constant factors, airspeed limitations cor- amount less than power-on V for the NE responding with the critical combina- remainder of the altitude range.

tions of these factors must be estab- lished.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), § 29.1505 Never-exceed speed.

1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

of Transportation Act (49 U.S.C. 1655(c))) (a) The never-exceed speed, V must NE, be established so that it is— [Amdt. 29–3, 33 FR 971, Jan. 26, 1968, as (1) Not less than 40 knots (CAS); and amended by Amdt. 29–15, 43 FR 2327, Jan. 16, (2) Not more than the lesser of— 1978; Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] 14 CFR Ch. I (1–1–25 Edition) § 29.1509 (1) The maximum rotational speed, § 29.1509 Rotor speed.

which may not be greater than— (a) Maximum power-off (autorotation).

(i) The maximum value determined The maximum power-off rotor speed by the rotor design; or must be established so that it does not (ii) The maximum value shown dur- exceed 95 percent of the lesser of— ing the type tests; (1) The maximum design r.p.m. deter- (2) The maximum allowable manifold mined under § 29.309(b); and pressure (for reciprocating engines); (2) The maximum r.p.m. shown dur- (3) The maximum allowable turbine ing the type tests.

inlet or turbine outlet gas temperature (b) Minimum power-off. The minimum (for turbine engines); power-off rotor speed must be estab- (4) The maximum allowable power or lished so that it is not less than 105 torque for each engine, considering the percent of the greater of— power input limitations of the trans- (1) The minimum shown during the mission with all engines operating; type tests; and (5) The maximum allowable power or (2) The minimum determined by de- torque for each engine considering the sign substantiation.

power input limitations of the trans- (c) Minimum power-on. The minimum mission with one engine inoperative; power-on rotor speed must be estab- (6) The time limit for the use of the lished so that it is— power corresponding to the limitations (1) Not less than the greater of— established in paragraphs (b)(1) (i) The minimum shown during the through (5) of this section; and type tests; and (7) If the time limit established in (ii) The minimum determined by de- paragraph (b)(6) of this section exceeds sign substantiation; and 2 minutes— (2) Not more than a value determined (i) The maximum allowable cylinder under § 29.33 (a)(1) and (c)(1).

head or coolant outlet temperature (for reciprocating engines); and § 29.1517 Limiting height-velocity en- (ii) The maximum allowable engine velope.

and transmission oil temperatures.

For Category A rotorcraft, if a range (c) Continuous operation. The contin- of heights exists at any speed, includ- uous operation must be limited by— ing zero, within which it is not possible (1) The maximum rotational speed, to make a safe landing following power which may not be greater than— failure, the range of heights and its (i) The maximum value determined variation with forward speed must be by the rotor design; or established, together with any other (ii) The maximum value shown dur- pertinent information, such as the kind ing the type tests; of landing surface.

(2) The minimum rotational speed [Amdt. 29–21, 48 FR 4391, Jan. 31, 1983, as shown under the rotor speed require- amended by Amdt. 29–59, 88 FR 8739, Feb. 10, ments in § 29.1509(c).

2023] (3) The maximum allowable manifold pressure (for reciprocating engines); § 29.1519 Weight and center of gravity.

(4) The maximum allowable turbine The weight and center of gravity lim- inlet or turbine outlet gas temperature itations determined under §§ 29.25 and (for turbine engines); 29.27, respectively, must be established (5) The maximum allowable power or as operating limitations.

torque for each engine, considering the power input limitations of the trans- § 29.1521 Powerplant limitations.

mission with all engines operating; (a) General. The powerplant limita- (6) The maximum allowable power or tions prescribed in this section must be torque for each engine, considering the established so that they do not exceed power input limitations of the trans- the corresponding limits for which the mission with one engine inoperative; engines are type certificated.

and (b) Takeoff operation. The powerplant (7) The maximum allowable tempera- takeoff operation must be limited by— tures for— Federal Aviation Administration, DOT § 29.1521 (i) The cylinder head or coolant out- (4) The maximum allowable oil tem- let (for reciprocating engines); perature.

(ii) The engine oil; and (h) Continuous OEI power operation.

(iii) The transmission oil. Unless otherwise authorized, the use of (d) Fuel grade or designation. The min- continuous OEI power must be limited imum fuel grade (for reciprocating en- to multiengine, turbine-powered rotor- gines) or fuel designation (for turbine craft for continued flight after failure engines) must be established so that it of an engine. The use of continuous is not less than that required for the OEI power must also be limited by— operation of the engines within the (1) The maximum rotational speed, limitations in paragraphs (b) and (c) of which may not be greater than— this section. (i) The maximum value determined (e) Ambient temperature. Ambient by the rotor design; or temperature limitations (including (ii) The maximum value shown dur- limitations for winterization installa- ing the type tests.

tions if applicable) must be established (2) The maximum allowable gas tem- as the maximum ambient atmospheric perature; temperature at which compliance with (3) The maximum allowable torque; the cooling provisions of §§ 29.1041 and through 29.1049 is shown. (4) The maximum allowable oil tem- (f) Two and one-half minute OEI power perature.

operation. Unless otherwise authorized, (i) Rated 30-second OEI power oper- the use of 2 ⁄ 2 -minute OEI power must ation. Rated 30-second OEI power is be limited to engine failure operation permitted only on multiengine, tur- of multiengine, turbine-powered rotor- bine-powered rotorcraft, also certifi- craft for not longer than 2 ⁄2 minutes cated for the use of rated 2-minute OEI for any period in which that power is power, and can only be used for contin- used. The use of 2 ⁄ 2 -minute OEI power ued operation of the remaining en- must also be limited by— gine(s) after a failure or precautionary (1) The maximum rotational speed, shutdown of an engine. It must be which may not be greater than— shown that following application of 30- (i) The maximum value determined second OEI power, any damage will be by the rotor design; or readily detectable by the applicable in- (ii) The maximum value shown dur- spections and other related procedures ing the type tests; furnished in accordance with Section (2) The maximum allowable gas tem- A29.4 of appendix A of this part and perature; Section A33.4 of appendix A of part 33.

(3) The maximum allowable torque; The use of 30-second OEI power must be and limited to not more than 30 seconds for (4) The maximum allowable oil tem- any period in which that power is used, perature. and by— (g) Thirty-minute OEI power operation. (1) The maximum rotational speed Unless otherwise authorized, the use of which may not be greater than— 30-minute OEI power must be limited (i) The maximum value determined to multiengine, turbine-powered rotor- by the rotor design; or craft for not longer than 30 minutes (ii) The maximum value dem- after failure of an engine. The use of 30- onstrated during the type tests; minute OEI power must also be limited (2) The maximum allowable gas tem- by— perature; and (1) The maximum rotational speed, (3) The maximum allowable torque.

which may not be greater than— (j) Rated 2-minute OEI power oper- (i) The maximum value determined ation. Rated 2-minute OEI power is per- by the rotor design; or mitted only on multiengine, turbine- (ii) The maximum value shown dur- powered rotorcraft, also certificated ing the type tests; for the use of rated 30-second OEI (2) The maximum allowable gas tem- power, and can only be used for contin- perature; ued operation of the remaining en- (3) The maximum allowable torque; gine(s) after a failure or precautionary and shutdown of an engine. It must be 14 CFR Ch. I (1–1–25 Edition) § 29.1522 shown that following application of 2- § 29.1525 Kinds of operations.

minute OEI power, any damage will be The kinds of operations (such as readily detectable by the applicable in- VFR, IFR, day, night, or icing) for spections and other related procedures which the rotorcraft is approved are es- furnished in accordance with Section tablished by demonstrated compliance A29.4 of appendix a of this part and with the applicable certification re- Section A33.4 of appendix A of part 33.

quirements and by the installed equip- The use of 2-minute OEI power must be ment.

limited to not more than 2 minutes for [Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] any period in which that power is used, and by— § 29.1527 Maximum operating altitude.

(1) The maximum rotational speed, The maximum altitude up to which which may not be greater than— operation is allowed, as limited by (i) The maximum value determined flight, structural, powerplant, func- by the rotor design; or tional, or equipment characteristics, (ii) The maximum value dem- must be established.

onstrated during the type tests; (2) The maximum allowable gas tem- (Secs. 313(a), 601, 603, 604, and 605 of the Fed- perature; and eral Aviation Act of 1958 (49 U.S.C. 1354(a), (3) The maximum allowable torque.

1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

of Transportation Act (49 U.S.C. 1655(c))) (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), [Amdt. 29–15, 43 FR 2327, Jan. 16, 1978] 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

§ 29.1529 Instructions for Continued of Transportation Act (49 U.S.C. 1655(c))) Airworthiness.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as The applicant must prepare Instruc- amended by Amdt. 29–1, 30 FR 8778, July 13, 1965; Amdt. 29–3, 33 FR 971, Jan. 26, 1968; tions for Continued Airworthiness in Amdt. 29–15, 43 FR 2327, Jan. 16, 1978; Amdt.

accordance with appendix A to this 29–26, 53 FR 34220, Sept. 2, 1988; Amdt. 29–34, part that are acceptable to the Admin- 59 FR 47768, Sept. 16, 1994; Amdt. 29–41, 62 FR istrator. The instructions may be in- 46173, Aug. 29, 1997] complete at type certification if a pro- gram exists to ensure their completion § 29.1522 Auxiliary power unit limita- prior to delivery of the first rotorcraft tions.

or issuance of a standard certificate of If an auxiliary power unit that meets airworthiness, whichever occurs later.

the requirements of TSO-C77 is in- [Amdt. 29–20, 45 FR 60178, Sept. 11, 1980] stalled in the rotorcraft, the limita- tions established for that auxiliary M ARKINGS AND P LACARDS power unit under the TSO including the categories of operation must be § 29.1541 General.

specified as operating limitations for (a) The rotorcraft must contain— the rotorcraft.

(1) The markings and placards speci- (Secs. 313(a), 601, 603, 604, Federal Aviation fied in §§ 29.1545 through 29.1565; and Act of 1958 (49 U.S.C. 1354(a), 1421, 1423), sec.

(2) Any additional information, in- 6(c), Dept. of Transportation Act (49 U.S.C.

strument markings, and placards re- 1655(c))) quired for the safe operation of the [Amdt. 29–17, 43 FR 50602, Oct. 30, 1978] rotorcraft if it has unusual design, op- erating or handling characteristics.

§ 29.1523 Minimum flight crew.

(b) Each marking and placard pre- The minimum flight crew must be es- scribed in paragraph (a) of this sec- tablished so that it is sufficient for safe tion— operation, considering— (1) Must be displayed in a con- (a) The workload on individual crew- spicuous place; and members; (2) May not be easily erased, dis- (b) The accessibility and ease of oper- figured, or obscured.

ation of necessary controls by the ap- § 29.1543 Instrument markings: gen- propriate crewmember; and eral.

(c) The kinds of operation authorized under § 29.1525. For each instrument— Federal Aviation Administration, DOT § 29.1555 (a) When markings are on the cover § 29.1549 Powerplant instruments.

glass of the instrument there must be For each required powerplant instru- means to maintain the correct align- ment, as appropriate to the type of in- ment of the glass cover with the face of struments— the dial; and (a) Each maximum and, if applicable, (b) Each arc and line must be wide minimum safe operating limit must be enough, and located to be clearly visi- marked with a red line; ble to the pilot.

(b) Each normal operating range must be marked as a green or un- § 29.1545 Airspeed indicator.

marked range; (a) Each airspeed indicator must be (c) Each takeoff and precautionary marked as specified in paragraph (b) of range must be marked with a yellow this section, with the marks located at range or yellow line; the corresponding indicated airspeeds.

(d) Each engine or rotor range that is (b) The following markings must be restricted because of excessive vibra- made: tion stresses must be marked with red ranges or red lines; and (1) A red line: (e) Each OEI limit or approved oper- (i) For rotorcraft other than heli- ating range must be marked to be copters, at V .

NE clearly differentiated from the mark- (ii) For helicopters, at V (power- NE ings of paragraphs (a) through (d) of on).

this section except that no marking is (iii) For helicopters, at V (power- NE normally required for the 30-second off). If V (power-off) is less than V NE NE OEI limit.

(power-on) and both are simulta- neously displayed, the red line at V [Amdt. 29–12, 41 FR 55474, Dec. 20, 1976, as NE amended by Amdt. 29–26, 53 FR 34220, Sept. 2, (power-off) must be clearly distinguish- 1988; Amdt. 29–34, 59 FR 47769, Sept. 16, 1994; able from the red line at V (power- NE Amdt. 29–59, 88 FR 8739, Feb. 10, 2023] on).

(2) [Reserved] § 29.1551 Oil quantity indicator.

(3) For the caution range, a yellow Each oil quantity indicator must be range.

marked with enough increments to in- (4) For the normal operating range, a dicate readily and accurately the quan- green or unmarked range.

tity of oil.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- § 29.1553 Fuel quantity indicator.

eral Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

If the unusable fuel supply for any of Transportation Act (49 U.S.C. 1655(c))) tank exceeds one gallon, or five per- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as cent of the tank capacity, whichever is amended by Amdt. 29–15, 43 FR 2327, Jan. 16, greater, a red arc must be marked on 1978; 43 FR 3900, Jan. 30, 1978; Amdt. 29–17, 43 its indicator extending from the cali- FR 50602, Oct. 30, 1978; Amdt. 29–59, 88 FR brated zero reading to the lowest read- 8740, Feb. 10, 2023] ing obtainable in level flight.

§ 29.1547 Magnetic direction indicator.

§ 29.1555 Control markings.

(a) A placard meeting the require- (a) Each cockpit control, other than ments of this section must be installed primary flight controls or control on or near the magnetic direction indi- whose function is obvious, must be cator.

plainly marked as to its function and (b) The placard must show the cali- method of operation.

bration of the instrument in level (b) For powerplant fuel controls— flight with the engines operating.

(1) Each fuel tank selector valve con- (c) The placard must state whether trol must be marked to indicate the po- the calibration was made with radio re- sition corresponding to each tank and ceivers on or off.

to each existing cross feed position; (d) Each calibration reading must be (2) If safe operation requires the use in terms of magnetic heading in not of any tanks in a specific sequence, more than 45 degree increments. that sequence must be marked on, or 14 CFR Ch. I (1–1–25 Edition) § 29.1557 adjacent to, the selector for those (1) Fuel filler openings must be tanks; and marked at or near the filler cover with— (3) Each valve control for any engine (i) The word ‘‘fuel’’; of a multiengine rotorcraft must be (ii) For reciprocating engine powered marked to indicate the position cor- rotorcraft, the minimum fuel grade; responding to each engine controlled.

(iii) For turbine-engine-powered (c) Usable fuel capacity must be rotorcraft, the permissible fuel des- marked as follows: ignations, except that if impractical, (1) For fuel systems having no selec- this information may be included in tor controls, the usable fuel capacity of the rotorcraft flight manual, and the the system must be indicated at the fuel filler may be marked with an ap- fuel quantity indicator unless it is: propriate reference to the flight man- (i) Provided by another system or ual; and equipment readily accessible to the (iv) For pressure fueling systems, the pilot; and maximum permissible fueling supply (ii) Contained in the limitations sec- pressure and the maximum permissible tion of the rotorcraft flight manual.

defueling pressure.

(2) For fuel systems having selector (2) Oil filler openings must be controls, the usable fuel capacity marked at or near the filler cover with available at each selector control posi- the word ‘‘oil’’.

tion must be indicated near the selec- (d) Emergency exit placards. Each tor control.

placard and operating control for each (d) For accessory, auxiliary, and emergency exit must differ in color emergency controls— from the surrounding fuselage surface (1) Each essential visual position in- as prescribed in § 29.811(f)(2). A placard dicator, such as those showing rotor must be near each emergency exit con- pitch or landing gear position, must be trol and must clearly indicate the loca- marked so that each crewmember can tion of that exit and its method of op- determine at any time the position of eration.

the unit to which it relates; and [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as (2) Each emergency control must be amended by Amdt. 29–3, 33 FR 971, Jan. 26, red and must be marked as to method 1968; Amdt. 29–12, 41 FR 55474, Dec. 20, 1976; of operation.

Amdt. 29–26, 53 FR 34220, Sept. 2, 1988; Amdt.

29–58, 87 FR 75711, Dec. 9, 2022] (e) For rotorcraft incorporating re- tractable landing gear, the maximum § 29.1559 Limitations placard.

landing gear operating speed must be There must be a placard in clear view displayed in clear view of the pilot.

of the pilot that specifies the kinds of [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as operations (VFR, IFR, day, night, or amended by Amdt. 29–12, 41 FR 55474, Dec. 20, icing) for which the rotorcraft is ap- 1976; Amdt. 29–24, 49 FR 44440, Nov. 6, 1984; proved.

Amdt. 29–59, 88 FR 8740, Feb. 10, 2023] [Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] § 29.1557 Miscellaneous markings and placards.

§ 29.1561 Safety equipment.

(a) Baggage and cargo compartments, (a) Each safety equipment control to and ballast location. Each baggage and be operated by the crew in emergency, cargo compartment, and each ballast such as controls for automatic liferaft location must have a placard stating releases, must be plainly marked as to any limitations on contents, including its method of operation.

weight, that are necessary under the (b) Each location, such as a locker or loading requirements.

compartment, that carries any fire ex- (b) Seats. If the maximum allowable tinguishing, signaling, or other life weight to be carried in a seat is less saving equipment, must be so marked.

than 170 pounds, a placard stating the (c) Stowage provisions for required lesser weight must be permanently at- emergency equipment must be con- tached to the seat structure.

spicuously marked to identify the con- (c) Fuel and oil filler openings. The fol- tents and facilitate removal of the lowing apply: equipment.

Federal Aviation Administration, DOT § 29.1585 (d) Each liferaft must have obviously (3) Information necessary for mark- marked operating instructions. ing the instruments required by (e) Approved survival equipment §§ 29.1549 through 29.1553.

must be marked for identification and (c) Weight and loading distribution.

method of operation. The weight and center of gravity limits required by §§ 29.25 and 29.27, respec- § 29.1565 Tail rotor.

tively, must be furnished. If the vari- ety of possible loading conditions war- Each tail rotor must be marked so rants, instructions must be included to that its disc is conspicuous under nor- allow ready observance of the limita- mal daylight ground conditions.

tions.

[Amdt. 29–3, 33 FR 971, Jan. 26, 1968] (d) Flight crew. When a flight crew of more than one is required, the number R OTORCRAFT F LIGHT M ANUAL and functions of the minimum flight crew determined under § 29.1523 must be § 29.1581 General.

furnished.

(a) Furnishing information. A Rotor- (e) Kinds of operation. Each kind of craft Flight Manual must be furnished operation for which the rotorcraft and with each rotorcraft, and it must con- its equipment installations are ap- tain the following: proved must be listed.

(1) Information required by §§ 29.1583 (f) Limiting heights. Enough informa- through 29.1589.

tion must be furnished to allow compli- (2) Other information that is nec- ance with § 29.1517.

essary for safe operation because of de- (g) Maximum allowable wind. For Cat- sign, operating, or handling character- egory A rotorcraft, the maximum al- istics. lowable wind for safe operation near the ground must be furnished.

(b) Approved information. Each part of (h) Altitude. The altitude established the manual listed in §§ 29.1583 through under § 29.1527 and an explanation of 29.1589 that is appropriate to the rotor- the limiting factors must be furnished.

craft, must be furnished, verified, and (i) Ambient temperature. Maximum approved, and must be segregated, and minimum ambient temperature indentified, and clearly distinguished limitations must be furnished.

from each unapproved part of that manual.

(Secs. 313(a), 601, 603, 604, and 605 of the Fed- (c) [Reserved] eral Aviation Act of 1958 (49 U.S.C. 1354(a), (d) Table of contents. Each Rotorcraft 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

of Transportation Act (49 U.S.C. 1655(c))) Flight Manual must include a table of contents if the complexity of the man- [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as ual indicates a need for it.

amended by Amdt. 29–3, 33 FR 971, Jan. 26, 1968; Amdt. 29–15, 43 FR 2327, Jan. 16, 1978; (Secs. 313(a), 601, 603, 604, and 605 of the Fed- Amdt. 29–17, 43 FR 50602, Oct. 30, 1978; Amdt.

eral Aviation Act of 1958 (49 U.S.C. 1354(a), 29–24, 49 FR 44440, Nov. 6, 1984] 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

of Transportation Act (49 U.S.C. 1655(c))) § 29.1585 Operating procedures.

[Amdt. 29–15, 43 FR 2327, Jan. 16, 1978] (a) The parts of the manual con- taining operating procedures must § 29.1583 Operating limitations.

have information concerning any nor- (a) Airspeed and rotor limitations. In- mal and emergency procedures, and formation necessary for the marking of other information necessary for safe airspeed and rotor limitations on or operation, including the applicable pro- near their respective indicators must cedures, such as those involving min- be furnished. The significance of each imum speeds, to be followed if an en- limitation and of the color coding must gine fails.

be explained.

(b) For multiengine rotorcraft, infor- (b) Powerplant limitations. The fol- mation identifying each operating con- lowing information must be furnished: dition in which the fuel system inde- (1) Limitations required by § 29.1521. pendence prescribed in § 29.953 is nec- (2) Explanation of the limitations, essary for safety must be furnished, to- when appropriate. gether with instructions for placing 14 CFR Ch. I (1–1–25 Edition) § 29.1587 the fuel system in a configuration used data was determined, and must con- to show compliance with that section. tain— (c) For helicopters for which a V (1) The indicated airspeeds cor- NE (power-off) is established under responding with those determined for § 29.1505(c), information must be fur- takeoff, and the procedures to be fol- nished to explain the V (power-off) NE lowed if the critical engine fails during and the procedures for reducing air- takeoff; speed to not more than the V (power- NE (2) The airspeed calibrations; off) following failure of all engines.

(3) The techniques, associated air- (d) For each rotorcraft showing com- speeds, and rates of descent for auto- pliance with § 29.1353 (c)(6)(ii) or rotative landings; (c)(6)(iii), the operating procedures for (4) The rejected takeoff distance de- disconnecting the battery from its termined under § 29.62 and the takeoff charging source must be furnished.

distance determined under § 29.61; (e) If the unusable fuel supply in any (5) The landing data determined tank exceeds 5 percent of the tank ca- under § 29.81 and § 29.85; pacity, or 1 gallon, whichever is great- (6) The steady gradient of climb for er, information must be furnished each weight, altitude, and temperature which indicates that when the fuel for which takeoff data are to be sched- quantity indicator reads ‘‘zero’’ in uled, along the takeoff path deter- level flight, any fuel remaining in the mined in the flight conditions required fuel tank cannot be used safely in flight. in § 29.67(a)(1) and (a)(2): (f) Information on the total quantity (i) In the flight conditions required in of usable fuel for each fuel tank must § 29.67(a)(1) between the end of the be furnished. takeoff distance and the point at which (g) For Category B rotorcraft, the the rotorcraft is 200 feet above the airspeeds and corresponding rotor takeoff surface (or 200 feet above the speeds for minimum rate of descent lowest point of the takeoff profile for and best glide angle as prescribed in elevated heliports); § 29.71 must be provided.

(ii) In the flight conditions required in § 29.67(a)(2) between the points at (Secs. 313(a), 601, 603, 604, and 605 of the Fed- eral Aviation Act of 1958 (49 U.S.C. 1354(a), which the rotorcraft is 200 and 1000 feet 1421, 1423, 1424, and 1425); and sec. 6(c), Dept.

above the takeoff surface (or 200 and of Transportation Act (49 U.S.C. 1655(c))) 1000 feet above the lowest point of the takeoff profile for elevated heliports); [Amdt. 29–2, 32 FR 6914, May 5, 1967, as amended by Amdt. 29–15, 43 FR 2328, Jan. 16, and 1978; Amdt. 29–17, 43 FR 50602, Oct. 30, 1978; (7) Out-of-ground effect hover per- Amdt. 29–24, 49 FR 44440, Nov. 6, 1984] formance determined under § 29.49 and the maximum weight for each altitude § 29.1587 Performance information.

and temperature condition at which Flight manual performance informa- the rotorcraft can safely hover out-of- tion which exceeds any operating limi- ground effect in winds of not less than tation may be shown only to the extent 17 knots from all azimuths. These data necessary for presentation clarity or to must be clearly referenced to the ap- determine the effects of approved op- propriate hover charts.

tional equipment or procedures. When (b) Category B. For each category B data beyond operating limits are rotorcraft, the Rotorcraft Flight Man- shown, the limits must be clearly indi- ual must contain— cated. The following must be provided: (1) The takeoff distance and the (a) Category A. For each category A climbout speed together with the perti- rotorcraft, the Rotorcraft Flight Man- nent information defining the flight ual must contain a summary of the path with respect to autorotative land- performance data, including data nec- ing if an engine fails, including the cal- essary for the application of any oper- culated effects of altitude and tempera- ating rule of this chapter, together ture; with descriptions of the conditions, such as airspeeds, under which this Federal Aviation Administration, DOT Pt. 29, App. A (2) The steady rates of climb and in- A PPENDIX A TO P ART 29—I NSTRUCTIONS ground-effect hovering ceiling, to- FOR C ONTINUED A IRWORTHINESS gether with the corresponding air- a29.1 General speeds and other pertinent informa- (a) This appendix specifies requirements tion, including the calculated effects of for the preparation of Instructions for Con- altitude and temperature; tinued Airworthiness as required by § 29.1529.

(3) The landing distance, appropriate (b) The Instructions for Continued Air- worthiness for each rotorcraft must include airspeed, and type of landing surface, the Instructions for Continued Airworthiness together with all pertinent information for each engine and rotor (hereinafter des- that might affect this distance, includ- ignated ‘‘products’’), for each appliance re- ing the effects of weight, altitude, and quired by this chapter, and any required in- temperature; formation relating to the interface of those (4) The maximum safe wind for oper- appliances and products with the rotorcraft.

ation near the ground; If Instructions for Continued Airworthiness are not supplied by the manufacturer of an (5) The airspeed calibrations; appliance or product installed in the rotor- (6) The height-velocity envelope ex- craft, the Instructions for Continued Air- cept for rotorcraft incorporating this worthiness for the rotorcraft must include as an operating limitation; the information essential to the continued (7) Glide distance as a function of al- airworthiness of the rotorcraft.

(c) The applicant must submit to the FAA titude when autorotating at the speeds a program to show how changes to the In- and conditions for minimum rate of de- structions for Continued Airworthiness made scent and best glide angle, as deter- by the applicant or by the manufacturers of mined in § 29.71; products and appliances installed in the (8) Out-of-ground effect hover per- rotorcraft will be distributed.

formance determined under § 29.49 and a29.2 Format the maximum safe wind demonstrated (a) The Instructions for Continued Air- under the ambient conditions for data worthiness must be in the form of a manual presented. In addition, the maximum or manuals as appropriate for the quantity weight for each altitude and tempera- of data to be provided.

(b) The format of the manual or manuals ture condition at which the rotorcraft must provide for a practical arrangement.

can safely hover out-of-ground-effect in a29.3 Content winds of not less than 17 knots from all The contents of the manual or manuals azimuths. These data must be clearly must be prepared in the English language.

referenced to the appropriate hover The Instructions for Continued Airworthi- charts; and ness must contain the following manuals or (9) Any additional performance data sections, as appropriate, and information: necessary for the application of any op- (a) Rotorcraft maintenance manual or section.

erating rule in this chapter.

(1) Introduction information that includes an explanation of the rotorcraft’s features and [Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as data to the extent necessary for mainte- amended by Amdt. 29–21, 48 FR 4392, Jan. 31, nance or preventive maintenance.

1983; Amdt. 29–24, 49 FR 44440, Nov. 6, 1984; (2) A description of the rotorcraft and its Amdt. 29–39, 61 FR 21901, May 10, 1996; Amdt.

systems and installations including its en- 29–40, 61 FR 21908, May 10, 1996; Amdt. 29–44, gines, rotors, and appliances.

64 FR 45338, Aug. 19, 1999; Amdt. 29–51, 73 FR (3) Basic control and operation information 11001, Feb. 29, 2008; Amdt. 29–59, 88 FR 8740, describing how the rotorcraft components Feb. 10, 2023] and systems are controlled and how they op- erate, including any special procedures and § 29.1589 Loading information.

limitations that apply.

(4) Servicing information that covers de- There must be loading instructions tails regarding servicing points, capacities of for each possible loading condition be- tanks, reservoirs, types of fluids to be used, tween the maximum and minimum pressures applicable to the various systems, weights determined under § 29.25 that location of access panels for inspection and servicing, locations of lubrication points, the can result in a center of gravity beyond lubricants to be used, equipment required for any extreme prescribed in § 29.27, as- servicing, tow instructions and limitations, suming any probable occupant weights.

mooring, jacking, and leveling information.

(b) Maintenance Instructions. (1) Scheduling information for each part of the rotorcraft 14 CFR Ch. I (1–1–25 Edition) Pt. 29, App. B and its engines, auxiliary power units, ro- proved and specifies maintenance required tors, accessories, instruments, and equip- under §§ 43.16 and 91.403 of the Federal Avia- ment that provides the recommended periods tion Regulations unless an alternative pro- at which they should be cleaned, inspected, gram has been FAA approved.’’ adjusted, tested, and lubricated, and the de- [Amdt. 29–20, 45 FR 60178, Sept. 11, 1980, as gree of inspection, the applicable wear toler- amended by Amdt. 29–27, 54 FR 34330, Aug. 18, ances, and work recommended at these peri- 1989; Amdt. 29–54, 76 FR 74664, Dec. 1, 2011] ods. However, the applicant may refer to an accessory, instrument, or equipment manu- A PPENDIX B TO P ART 29—A IRWORTHI - facturer as the source of this information if NESS C RITERIA FOR H ELICOPTER I N- the applicant shows that the item has an ex- STRUMENT F LIGHT ceptionally high degree of complexity requir- ing specialized maintenance techniques, test I. General. A transport category helicopter equipment, or expertise. The recommended may not be type certificated for operation overhaul periods and necessary cross ref- under the instrument flight rules (IFR) of erences to the Airworthiness Limitations this chapter unless it meets the design and section of the manual must also be included.

installation requirements contained in this In addition, the applicant must include an appendix.

inspection program that includes the fre- II. Definitions. (a) V YI means instrument quency and extent of the inspections nec- climb speed, utilized instead of V Y for com- essary to provide for the continued air- pliance with the climb requirements for in- worthiness of the rotorcraft.

strument flight.

(2) Troubleshooting information describing means instrument flight never ex- (b) V NEI probable malfunctions, how to recognize for com- ceed speed, utilized instead of V NE those malfunctions, and the remedial action pliance with maximum limit speed require- for those malfunctions.

ments for instrument flight.

(3) Information describing the order and (c) V means instrument flight min- MINI method of removing and replacing products imum speed, utilized in complying with min- and parts with any necessary precautions to imum limit speed requirements for instru- be taken.

ment flight.

(4) Other general procedural instructions III. Trim. It must be possible to trim the including procedures for system testing dur- cyclic, collective, and directional control ing ground running, symmetry checks, forces to zero at all approved IFR airspeeds, weighing and determining the center of grav- power settings, and configurations appro- ity, lifting and shoring, and storage limita- priate to the type.

tions.

IV. Static longitudinal stability. (a) General.

(c) Diagrams of structural access plates The helicopter must possess positive static and information needed to gain access for in- longitudinal control force stability at crit- spections when access plates are not pro- ical combinations of weight and center of vided.

gravity at the conditions specified in para- (d) Details for the application of special in- graphs IV (b) through (f) of this appendix.

spection techniques including radiographic The stick force must vary with speed so that and ultrasonic testing where such processes any substantial speed change results in a are specified.

stick force clearly perceptible to the pilot.

(e) Information needed to apply protective The airspeed must return to within 10 per- treatments to the structure after inspection.

cent of the trim speed when the control force (f) All data relative to structural fasteners is slowly released for each trim condition such as identification, discard recommenda- specified in paragraphs IV (b) through (f) of tions, and torque values.

this appendix.

(g) A list of special tools needed.

(b) Climb. Stability must be shown in climb a29.4 Airworthiness Limitations Section thoughout the speed range 20 knots either The Instructions for Continued Airworthi- side of trim with— ness must contain a section titled Airworthi- (1) The helicopter trimmed at V ; YI ness Limitations that is segregated and (2) Landing gear retracted (if retractable); clearly distinguishable from the rest of the and document. This section must set forth each (3) Power required for limit climb rate (at mandatory replacement time, structural in- least 1,000 fpm) at V or maximum contin- YI spection interval, and related structural in- uous power, whichever is less.

spection procedure required for type certifi- (c) Cruise. Stability must be shown cation. If the Instructions for Continued Air- throughout the speed range from 0.7 to 1.1 V H worthiness consist of multiple documents, or V , whichever is lower, not to exceed ± 20 NEI the section required by this paragraph must knots from trim with— be included in the principal manual. This (1) The helicopter trimmed and power ad- section must contain a legible statement in justed for level flight at 0.9 V or 0.9 V , H NEI a prominent location that reads: ‘‘The Air- whichever is lower; and worthiness Limitations section is FAA ap- (2) Landing gear retracted (if retractable).

Federal Aviation Administration, DOT Pt. 29, App. B (d) Slow cruise. Stability must be shown (e) Any aperiodic response may not achieve throughout the speed range from 0.9 V to double amplitude in less than 9 seconds.

MINI 1.3 V or 20 knots above trim speed, which- VII. Stability Augmentation System (SAS) MINI ever is greater, with— (a) If a SAS is used, the reliability of the (1) The helicopter trimmed and power ad- SAS must be related to the effects of its fail- justed for level flight at 1.1 V ; and MINI ure. Any SAS failure condition that would (2) Landing gear retracted (if retractable).

prevent continued safe flight and landing (e) Descent. Stability must be shown must be extremely improbable. It must be throughout the speed range 20 knots either shown that, for any failure condition of the side of trim with— SAS that is not shown to be extremely im- (1) The helicopter trimmed at 0.8 V or 0.8 probable— H V (or 0.8 V for the landing gear extended (1) The helicopter is safely controllable NEI LE case), whichever is lower; when the failure or malfunction occurs at (2) Power required for 1,000 fpm descent at any speed or altitude within the approved trim speed; and IFR operating limitations; and (3) Landing gear extended and retracted, if (2) The overall flight characteristics of the applicable. helicopter allow for prolonged instrument (f) Approach. Stability must be shown flight without undue pilot effort. Additional throughout the speed range from 0.7 times unrelated probable failures affecting the con- the minimum recommended approach speed trol system must be considered. In addi- to 20 knots above the maximum rec- tion— ommended approach speed with— (i) The controllability and maneuver- ability requirements in Subpart B must be (1) The helicopter trimmed at the rec- met throughout a practical flight envelope; ommended approach speed or speeds; (2) Landing gear extended and retracted, if (ii) The flight control, trim, and dynamic applicable; and stability characteristics must not be im- (3) Power required to maintain a 3 ° glide paired below a level needed to allow contin- path and power required to maintain the ued safe flight and landing; steepest approach gradient for which ap- (iii) For Category A helicopters, the dy- proval is requested. namic stability requirements of Subpart B V. Static Lateral Directional Stability must also be met throughout a practical (a) Static directional stability must be flight envelope; and positive throughout the approved ranges of (iv) The static longitudinal and static di- airspeed, power, and vertical speed. In rectional stability requirements of Subpart straight and steady sideslips up to ± 10 ° from B must be met throughout a practical flight trim, directional control position must in- envelope.

crease without discontinuity with the angle (b) The SAS must be designed so that it of sideslip, except for a small range of side- cannot create a hazardous deviation in flight slip angles around trim. At greater angles up path or produce hazardous loads on the heli- to the maximum sideslip angle appropriate copter during normal operation or in the to the type, increased directional control po- event of malfunction or failure, assuming sition must produce an increased angle of corrective action begins within an appro- sideslip. It must be possible to maintain bal- priate period of time. Where multiple sys- anced flight without exceptional pilot skill tems are installed, subsequent malfunction or alertness. conditions must be considered in sequence unless their occurrence is shown to be im- (b) During sideslips up to ± 10 ° from trim probable.

throughout the approved ranges of airspeed, VIII. Equipment, systems, and installation.

power, and vertical speed there must be no The basic equipment and installation must negative dihedral stability perceptible to the comply with §§ 29.1303, 29.1431, and 29.1433, pilot through lateral control motion or with the following exceptions and additions: force. Longitudinal cyclic movement with sideslip must not be excessive. (a) Flight and navigation instruments. (1) A VI. Dynamic stability. (a) Any oscillation magnetic gyro-stabilized direction indicator having a period of less than 5 seconds must instead of the gyroscopic direction indicator damp to ⁄2 amplitude in not more than one required by § 29.1303(h); and cycle. (2) A standby attitude indicator which (b) Any oscillation having a period of 5 sec- meets the requirements of §§ 29.1303(g)(1) onds or more but less than 10 seconds must through (7), instead of a rate-of-turn indi- damp to ⁄2 amplitude in not more than two cator required by § 29.1303(g). If standby bat- cycles. teries are provided, they may be charged (c) Any oscillation having a period of 10 from the aircraft electrical system if ade- seconds or more but less than 20 seconds quate isolation is incorporated. The system must be damped. must be designed so that the standby bat- (d) Any oscillation having a period of 20 teries may not be used for engine starting.

seconds or more may not achieve double am- (b) Miscellaneous requirements. (1) Instru- plitude in less than 20 seconds. ment systems and other systems essential 14 CFR Ch. I (1–1–25 Edition) Pt. 29, App. C for IFR flight that could be adversely af- revised kinds of operation, and the steepest fected by icing must be provided with ade- IFR precision approach gradient for which quate ice protection whether or not the the helicopter is approved; rotorcraft is certificated for operation in (b) Procedures. Required information for icing conditions.

proper operation of IFR systems and the rec- (2) There must be means in the generating ommended procedures in the event of sta- system to automatically de-energize and dis- bility augmentation or electrical system connect from the main bus any power source failures; and developing hazardous overvoltage.

(c) Performance. If V differs from V , YI Y (3) Each required flight instrument using a climb performance at V and with maximum YI power supply (electric, vacuum, etc.) must continuous power throughout the ranges of have a visual means integral with the instru- weight, altitude, and temperature for which ment to indicate the adequacy of the power approval is requested.

being supplied.

(4) When multiple systems performing like [Amdt. 29–21, 48 FR 4392, Jan. 31, 1983, as functions are required, each system must be amended by Amdt. 29–31, 55 FR 38967, Sept.

grouped, routed, and spaced so that physical 21, 1990; 55 FR 41309, Oct. 10, 1990; Amdt. 29– separation between systems is provided to 40, 61 FR 21908, May 10, 1996; Amdt. 29–51, 73 ensure that a single malfunction will not ad- FR 11002, Feb. 29, 2008; Amdt. 29–59, 88 FR versely affect more than one system.

8740, Feb. 10, 2023] (5) For systems that operate the required flight instruments at each pilot’s station— A PPENDIX C TO P ART 29—I CING (i) For pneumatic systems, only the re- C ERTIFICATION quired flight instruments for the first pilot may be connected to that operating system; (a) Continuous maximum icing. The max- (ii) Additional instruments, systems, or imum continuous intensity of atmospheric equipment may not be connected to an oper- icing conditions (continuous maximum ating system for a second pilot unless provi- icing) is defined by the variables of the cloud sions are made to ensure the continued nor- liquid water content, the mean effective di- mal functioning of the required instruments ameter of the cloud droplets, the ambient air in the event of any malfunction of the addi- temperature, and the interrelationship of tional instruments, systems, or equipment these three variables as shown in Figure 1 of which is not shown to be extremely improb- this appendix. The limiting icing envelope in able; terms of altitude and temperature is given in (iii) The equipment, systems, and installa- Figure 2 of this appendix. The interrelation- tions must be designed so that one display of ship of cloud liquid water content with drop the information essential to the safety of diameter and altitude is determined from flight which is provided by the instruments Figures 1 and 2. The cloud liquid water con- will remain available to a pilot, without ad- tent for continuous maximum icing condi- ditional crew-member action, after any sin- tions of a horizontal extent, other than 17.4 gle failure or combination of failures that is nautical miles, is determined by the value of not shown to be extremely improbable; and liquid water content of Figure 1, multiplied (iv) For single-pilot configurations, instru- by the appropriate factor from Figure 3 of ments which require a static source must be this appendix.

provided with a means of selecting an alter- (b) Intermittent maximum icing. The inter- nate source and that source must be cali- mittent maximum intensity of atmospheric brated.

(6) In determining compliance with the re- icing conditions (intermittent maximum quirements of § 29.1351(d)(2), the supply of icing) is defined by the variables of the cloud electrical power to all systems necessary for liquid water content, the mean effective di- flight under IFR must be included in the ameter of the cloud droplets, the ambient air evaluation. temperature, and the interrelationship of (c) Thunderstorm lights. In addition to the these three variables as shown in Figure 4 of instrument lights required by § 29.1381(a), this appendix. The limiting icing envelope in thunderstorm lights which provide high in- terms of altitude and temperature is given in tensity white flood lighting to the basic Figure 5 of this appendix. The interrelation- flight instruments must be provided. The ship of cloud liquid water content with drop thunderstorm lights must be installed to diameter and altitude is determined from meet the requirements of § 29.1381(b). Figures 4 and 5. The cloud liquid water con- IX. Rotorcraft Flight Manual. A Rotorcraft tent for intermittent maximum icing condi- Flight Manual or Rotorcraft Flight Manual tions of a horizontal extent, other than 2.6 IFR Supplement must be provided and must nautical miles, is determined by the value of contain— cloud liquid water content of Figure 4 multi- (a) Limitations. The approved IFR flight en- plied by the appropriate factor in Figure 6 of velope, the IFR flightcrew composition, the this appendix.

Federal Aviation Administration, DOT Pt. 29, App. C 14 CFR Ch. I (1–1–25 Edition) Pt. 29, App. C Federal Aviation Administration, DOT Pt. 29, App. C 14 CFR Ch. I (1–1–25 Edition) Pt. 29, App. C Federal Aviation Administration, DOT Pt. 29, App. C 14 CFR Ch. I (1–1–25 Edition) Pt. 29, App. C [Amdt. 29–21, 48 FR 4393, Jan. 31, 1983] Federal Aviation Administration, DOT Pt. 29, App. D (i) No passenger may be assigned a specific A PPENDIX D TO P ART 29—C RITERIA FOR seat except as the Administrator may re- D EMONSTRATION OF E MERGENCY quire. Except as required by paragraph (1) of E VACUATION P ROCEDURES U NDER this appendix, no employee of the applicant § 29.803 may be seated next to an emergency exit, ex- (a) The demonstration must be conducted cept as allowed by the Administrator.

either during the dark of the night or during (j) Seat belts and shoulder harnesses (as re- daylight with the dark of night simulated. If quired) must be fastened.

the demonstration is conducted indoors dur- (k) Before the start of the demonstration, ing daylight hours, it must be conducted in- approximately one-half of the total average side a darkened hangar having doors and amount of carry-on baggage, blankets, pil- windows covered. In addition, the doors and lows, and other similar articles must be dis- windows of the rotorcraft must be covered if tributed at several locations in the aisles the hangar illumination exceeds that of a and emergency exit access ways to create moonless night. Illumination on the floor or minor obstructions.

ground may be used, but it must be kept low (l) No prior indication may be given to any and shielded against shining into the rotorcraft’s windows or doors. crewmember or passenger of the particular (b) The rotorcraft must be in a normal at- exits to be used in the demonstration.

titude with landing gear extended.

(m) The applicant may not practice, re- (c) Safety equipment such as mats or in- hearse, or describe the demonstration for the verted liferafts may be placed on the floor or participants nor may any participant have ground to protect participants. No other taken part in this type of demonstration equipment that is not part of the rotorcraft’s within the preceding 6 months.

emergency evacuation equipment may be (n) A pretakeoff passenger briefing may be used to aid the participants in reaching the given. The passengers may also be advised to ground.

follow directions of crewmembers, but not be (d) Except as provided in paragraph (a) of instructed on the procedures to be followed this appendix, only the rotorcraft’s emer- in the demonstration.

gency lighting system may provide illumina- (o) If safety equipment, as allowed by para- tion.

graph (c) of this appendix, is provided, either (e) All emergency equipment required for the planned operation of the rotorcraft must all passenger and cockpit windows must be be installed.

blacked out or all emergency exits must (f) Each external door and exit and each in- have safety equipment to prevent disclosure ternal door or curtain must be in the takeoff of the available emergency exits.

configuration.

(p) Not more than 50 percent of the emer- (g) Each crewmember must be seated in gency exits in the sides of the fuselage of a the normally assigned seat for takeoff and rotorcraft that meet all of the requirements must remain in that seat until receiving the applicable to the required emergency exits signal for commencement of the demonstra- for that rotorcraft may be used for dem- tion. For compliance with this section, each onstration. Exits that are not to be used for crewmember must be— the demonstration must have the exit handle (1) A member of a regularly scheduled line deactivated or must be indicated by red crew; or lights, red tape, or other acceptable means (2) A person having knowledge of the oper- placed outside the exits to indicate fire or ation of exits and emergency equipment.

(h) A representative passenger load of per- other reasons why they are unusable. The sons in normal health must be used as fol- exits to be used must be representative of all lows: the emergency exits on the rotorcraft and (1) At least 25 percent must be over 50 must be designated by the applicant, subject years of age, with at least 40 percent of these to approval by the Administrator. If in- being females.

stalled, at least one floor level exit (Type I; (2) The remaining, 75 percent or less, must § 29.807(a)(1)) must be used as required by be 50 years of age or younger, with at least § 29.807(c).

30 percent of these being females.

(q) All evacuees must leave the rotorcraft (3) Three life-size dolls, not included as by a means provided as part of the part of the total passenger load, must be car- rotorcraft’s equipment.

ried by passengers to simulate live infants 2 (r) Approved procedures must be fully uti- years old or younger, except for a total pas- lized during the demonstration.

senger load of fewer than 44 but more than 19, one doll must be carried. A doll is not re- (s) The evacuation time period is com- quired for a 19 or fewer passenger load. pleted when the last occupant has evacuated (4) Crewmembers, mechanics, and training the rotorcraft and is on the ground.

personnel who maintain or operate the rotor- [Amdt. 27–26, 55 FR 8005, Mar. 6, 1990] craft in the normal course of their duties may not be used as passengers.

14 CFR Ch. I (1–1–25 Edition) Pt. 29, App. E A PPENDIX E TO P ART 29—HIRF E NVI - T ABLE III.—HIRF E NVIRONMENT III RONMENTS AND E QUIPMENT HIRF Field strength T EST L EVELS (volts/meter) Frequency Peak Average This appendix specifies the HIRF environ- ments and equipment HIRF test levels for 10 kHz–100 kHz ................................ 150 150 electrical and electronic systems under 100 kHz–400 MHz ............................. 200 200 § 29.1317. The field strength values for the 400 MHz–700 MHz ........................... 730 200 HIRF environments and laboratory equip- 700 MHz–1 GHz ................................ 1,400 240 ment HIRF test levels are expressed in root- 1 GHz–2 GHz .................................... 5,000 250 mean-square units measured during the peak 2 GHz–4 GHz .................................... 6,000 490 4 GHz–6 GHz .................................... 7,200 400 of the modulation cycle.

6 GHz–8 GHz .................................... 1,100 170 (a) HIRF environment I is specified in the 8 GHz–12 GHz .................................. 5,000 330 following table: 12 GHz–18 GHz ................................ 2,000 330 18 GHz–40 GHz ................................ 1,000 420 T ABLE I.—HIRF E NVIRONMENT I In this table, the higher field strength applies at the fre- quency band edges.

Field strength (volts/meter) Frequency (d) Equipment HIRF Test Level 1. (1) From 10 Peak Average kilohertz (kHz) to 400 megahertz (MHz), use conducted susceptibility tests with contin- 10 kHz–2 MHz ................................... 50 50 uous wave (CW) and 1 kHz square wave mod- 2 MHz–30 MHz ................................. 100 100 ulation with 90 percent depth or greater. The 30 MHz–100 MHz ............................. 50 50 conducted susceptibility current must start 100 MHz–400 MHz ........................... 100 100 at a minimum of 0.6 milliamperes (mA) at 10 400 MHz–700 MHz ........................... 700 50 kHz, increasing 20 decibel (dB) per frequency 700 MHz–1 GHz ................................ 700 100 decade to a minimum of 30 mA at 500 kHz.

1 GHz–2 GHz .................................... 2,000 200 (2) From 500 kHz to 40 MHz, the conducted 2 GHz–6 GHz .................................... 3,000 200 susceptibility current must be at least 30 6 GHz–8 GHz .................................... 1,000 200 mA.

8 GHz–12 GHz .................................. 3,000 300 (3) From 40 MHz to 400 MHz, use conducted 12 GHz–18 GHz ................................ 2,000 200 susceptibility tests, starting at a minimum 18 GHz–40 GHz ................................ 600 200 of 30 mA at 40 MHz, decreasing 20 dB per fre- In this table, the higher field strength applies at the fre- quency decade to a minimum of 3 mA at 400 quency band edges.

MHz.

(4) From 100 MHz to 400 MHz, use radiated (b) HIRF environment II is specified in the susceptibility tests at a minimum of 20 volts following table: per meter (V/m) peak with CW and 1 kHz square wave modulation with 90 percent T ABLE II.—HIRF E NVIRONMENT II depth or greater.

(5) From 400 MHz to 8 gigahertz (GHz), use Field strength (volts/meter) radiated susceptibility tests at a minimum Frequency of 150 V/m peak with pulse modulation of 4 Peak Average percent duty cycle with a 1 kHz pulse repeti- tion frequency. This signal must be switched 10 kHz–500 kHz ................................ 20 20 on and off at a rate of 1 Hz with a duty cycle 500 kHz–2 MHz ................................. 30 30 of 50 percent.

2 MHz–30 MHz ................................. 100 100 30 MHz–100 MHz ............................. 10 10 (e) Equipment HIRF Test Level 2. Equipment 100 MHz–200 MHz ........................... 30 10 HIRF test level 2 is HIRF environment II in 200 MHz–400 MHz ........................... 10 10 table II of this appendix reduced by accept- 400 MHz–1 GHz ................................ 700 40 able aircraft transfer function and attenu- 1 GHz–2 GHz .................................... 1,300 160 ation curves. Testing must cover the fre- 2 GHz–4 GHz .................................... 3,000 120 quency band of 10 kHz to 8 GHz.

4 GHz–6 GHz .................................... 3,000 160 (f) Equipment HIRF Test Level 3. (1) From 10 6 GHz–8 GHz .................................... 400 170 kHz to 400 MHz, use conducted susceptibility 8 GHz–12 GHz .................................. 1,230 230 tests, starting at a minimum of 0.15 mA at 10 12 GHz–18 GHz ................................ 730 190 kHz, increasing 20 dB per frequency decade 18 GHz–40 GHz ................................ 600 150 to a minimum of 7.5 mA at 500 kHz.

(2) From 500 kHz to 40 MHz, use conducted In this table, the higher field strength applies at the fre- quency band edges. susceptibility tests at a minimum of 7.5 mA.

(3) From 40 MHz to 400 MHz, use conducted (c) HIRF environment III is specified in the susceptibility tests, starting at a minimum following table: of 7.5 mA at 40 MHz, decreasing 20 dB per fre- quency decade to a minimum of 0.75 mA at 400 MHz.

Federal Aviation Administration, DOT § 31.12 (4) From 100 MHz to 8 GHz, use radiated APPENDIX A TO P ART 31—I NSTRUCTIONS FOR susceptibility tests at a minimum of 5 V/m. C ONTINUED A IRWORTHINESS [Doc. No. FAA–2006–23657, 72 FR 44028, Aug. 6, AUTHORITY : 49 U.S.C. 106(g), 40113, 44701– 2007] 44702, 44704.

S OURCE : Docket No. 1437, 29 FR 8258, July 1, PART 31—AIRWORTHINESS STAND- 1964, as amended by Amdt. 31–1, 29 FR 14563, Oct. 24, 1964, unless otherwise noted.

ARDS: MANNED FREE BAL- LOONS Subpart A—General Subpart A—General § 31.1 Applicability.

Sec.

(a) This part prescribes airworthiness 31.1 Applicability.

standards for the issue of type certifi- cates and changes to those certificates, Subpart B—Flight Requirements for manned free balloons.

31.12 Proof of compliance.

(b) Each person who applies under 31.14 Weight limits.

Part 21 for such a certificate or change 31.16 Empty weight.

must show compliance with the appli- 31.17 Performance: Climb.

cable requirements of this part.

31.19 Performance: Uncontrolled descent.

(c) For purposes of this part— 31.20 Controllability.

(1) A captive gas balloon is a balloon Subpart C—Strength Requirements that derives its lift from a captive lighter-than-air gas; 31.21 Loads.

(2) A hot air balloon is a balloon that 31.23 Flight load factor.

derives its lift from heated air; 31.25 Factor of safety.

31.27 Strength. (3) The envelope is the enclosure in which the lifting means is contained; Subpart D—Design Construction (4) The basket is the container, sus- pended beneath the envelope, for the 31.31 General.

balloon occupants; 31.33 Materials.

(5) The trapeze is a harness or is a 31.35 Fabrication methods.

31.37 Fastenings. seat consisting of a horizontal bar or 31.39 Protection. platform suspended beneath the enve- 31.41 Inspection provisions.

lope for the balloon occupants; and 31.43 Fitting factor.

(6) The design maximum weight is 31.45 Fuel cells.

the maximum total weight of the bal- 31.46 Pressurized fuel systems.

loon, less the lifting gas or air.

31.47 Burners.

31.49 Control systems.

[Doc. No. 1437, 29 FR 8258, July 1, 1964, as 31.51 Ballast.

amended by Amdt. 31–3, 41 FR 55474, Dec. 20, 31.53 Drag rope.

1976] 31.55 Deflation means.

31.57 Rip cords.

Subpart B—Flight Requirements 31.59 Trapeze, basket, or other means pro- vided for occupants.

§ 31.12 Proof of compliance.

31.61 Static discharge.

31.63 Safety belts.

(a) Each requirement of this subpart 31.65 Position lights.

must be met at each weight within the range of loading conditions for which Subpart E—Equipment certification is requested. This must be shown by— 31.71 Function and installation.

(1) Tests upon a balloon of the type Subpart F—Operating Limitations and for which certification is requested or Information by calculations based on, and equal in accuracy to, the results of testing; and 31.81 General.

(2) Systematic investigation of each 31.82 Instructions for Continued Airworthi- weight if compliance cannot be reason- ness.

ably inferred from the weights inves- 31.83 Conspicuity.

31.85 Required basic equipment. tigated.