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| section_id ▼ | title_number | title_name | chapter | subchapter | part_number | part_name | subpart | subpart_name | section_number | section_heading | agency | authority | source_citation | amendment_citations | full_text |
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| 14:14:1.0.1.3.15.1.161.1 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | A | Subpart A—General | § 27.1 Applicability. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-33, 61 FR 21906, May 10, 1996; Amdt. 27-37, 64 FR 45094, Aug. 18, 1999] | (a) This part prescribes airworthiness standards for the issue of type certificates, and changes to those certificates, for normal category rotorcraft with maximum weights of 7,000 pounds or less and nine or less passenger seats. (b) Each person who applies under Part 21 for such a certificate or change must show compliance with the applicable requirements of this part. (c) Multiengine rotorcraft may be type certified as Category A provided the requirements referenced in appendix C of this part are met. | |||
| 14:14:1.0.1.3.15.1.161.2 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | A | Subpart A—General | § 27.2 Special retroactive requirements. | FAA | [Doc. No. 26078, 56 FR 41051, Aug. 16, 1991, as amended by Amdt. 27-37, 64 FR 45094, Aug. 18, 1999] | (a) For each rotorcraft manufactured after September 16, 1992, each applicant must show that each occupant's seat is equipped with a safety belt and shoulder harness that meets the requirements of paragraphs (a), (b), and (c) of this section. (1) Each occupant's seat must have a combined safety belt and shoulder harness with a single-point release. Each pilot's combined safety belt and shoulder harness must allow each pilot, when seated with safety belt and shoulder harness fastened, to perform all functions necessary for flight operations. There must be a means to secure belts and harnesses, when not in use, to prevent interference with the operation of the rotorcraft and with rapid egress in an emergency. (2) Each occupant must be protected from serious head injury by a safety belt plus a shoulder harness that will prevent the head from contacting any injurious object. (3) The safety belt and shoulder harness must meet the static and dynamic strength requirements, if applicable, specified by the rotorcraft type certification basis. (4) For purposes of this section, the date of manufacture is either— (i) The date the inspection acceptance records, or equivalent, reflect that the rotorcraft is complete and meets the FAA-Approved Type Design Data; or (ii) The date the foreign civil airworthiness authority certifies that the rotorcraft is complete and issues an original standard airworthiness certificate, or equivalent, in that country. (b) For rotorcraft with a certification basis established prior to October 18, 1999— (1) The maximum passenger seat capacity may be increased to eight or nine provided the applicant shows compliance with all the airworthiness requirements of this part in effect on October 18, 1999. (2) The maximum weight may be increased to greater than 6,000 pounds provided— (i) The number of passenger seats is not increased above the maximum number certificated on October 18, 1999, or (ii) The applicant shows compliance with all of the airworthiness requirements of this part in effect o… | |||
| 14:14:1.0.1.3.15.2.161.1 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.21 Proof of compliance. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 49 FR 44432, Nov. 6, 1984] | Each requirement of this subpart must be met at each appropriate combination of weight and center of gravity within the range of loading conditions for which certification is requested. This must be shown— (a) By tests upon a rotorcraft of the type for which certification is requested, or by calculations based on, and equal in accuracy to, the results of testing; and (b) By systematic investigation of each required combination of weight and center of gravity if compliance cannot be reasonably inferred from combinations investigated. | |||
| 14:14:1.0.1.3.15.2.161.2 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.25 Weight limits. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 29, 1964, as amended by Amdt. 27-11, 41 FR 55468, Dec. 20, 1976; Amdt. 25-42, 43 FR 2324, Jan. 16, 1978; Amdt. 27-36, 64 FR 43019, Aug. 6, 1999; Amdt. 27-44, 73 FR 10998, Feb. 29, 2008; 73 FR 33876, June 16, 2008] | (a) Maximum weight. The maximum weight (the highest weight at which compliance with each applicable requirement of this part is shown) must be established so that it is— (1) Not more than— (i) The highest weight selected by the applicant; (ii) The design maximum weight (the highest weight at which compliance with each applicable structural loading condition of this part is shown); (iii) The highest weight at which compliance with each applicable flight requirement of this part is shown; or (iv) The highest weight in which the provisions of §§ 27.87 or 27.143(c)(1), or combinations thereof, are demonstrated if the weights and operating conditions (altitude and temperature) prescribed by those requirements cannot be met; and (2) Not less than the sum of— (i) The empty weight determined under § 27.29; and (ii) The weight of usable fuel appropriate to the intended operation with full payload; (iii) The weight of full oil capacity; and (iv) For each seat, an occupant weight of 170 pounds or any lower weight for which certification is requested. (b) Minimum weight. The minimum weight (the lowest weight at which compliance with each applicable requirement of this part is shown) must be established so that it is— (1) Not more than the sum of— (i) The empty weight determined under § 27.29; and (ii) The weight of the minimum crew necessary to operate the rotorcraft, assuming for each crewmember a weight no more than 170 pounds, or any lower weight selected by the applicant or included in the loading instructions; and (2) Not less than— (i) The lowest weight selected by the applicant; (ii) The design minimum weight (the lowest weight at which compliance with each applicable structural loading condition of this part is shown); or (iii) The lowest weight at which compliance with each applicable flight requirement of this part is shown. (c) Total weight with jettisonable external load. A total weight for the rotorcraft with a jettisonable external load attached that is greater than the maximum weight e… | |||
| 14:14:1.0.1.3.15.2.161.3 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.27 Center of gravity limits. | FAA | [Amdt. 27-2, 33 FR 962, Jan. 26, 1968] | The extreme forward and aft centers of gravity and, where critical, the extreme lateral centers of gravity must be established for each weight established under § 27.25. Such an extreme may not lie beyond— (a) The extremes selected by the applicant; (b) The extremes within which the structure is proven; or (c) The extremes within which compliance with the applicable flight requirements is shown. | |||
| 14:14:1.0.1.3.15.2.161.4 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.29 Empty weight and corresponding center of gravity. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 43 FR 2324, Jan. 16, 1978] | (a) The empty weight and corresponding center of gravity must be determined by weighing the rotorcraft without the crew and payload, but with— (1) Fixed ballast; (2) Unusable fuel; and (3) Full operating fluids, including— (i) Oil; (ii) Hydraulic fluid; and (iii) Other fluids required for normal operation of roto-craft systems, except water intended for injection in the engines. (b) The condition of the rotorcraft at the time of determining empty weight must be one that is well defined and can be easily repeated, particularly with respect to the weights of fuel, oil, coolant, and installed equipment. | |||
| 14:14:1.0.1.3.15.2.161.5 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.31 Removable ballast. | FAA | Removable ballast may be used in showing compliance with the flight requirements of this subpart. | ||||
| 14:14:1.0.1.3.15.2.161.6 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.33 Main rotor speed and pitch limits. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR 962, Jan. 26, 1968; Amdt. 27-14, 43 FR 2324, Jan. 16, 1978] | (a) Main rotor speed limits. A range of main rotor speeds must be established that— (1) With power on, provides adequate margin to accommodate the variations in rotor speed occurring in any appropriate maneuver, and is consistent with the kind of governor or synchronizer used; and (2) With power off, allows each appropriate autorotative maneuver to be performed throughout the ranges of airspeed and weight for which certification is requested. (b) Normal main rotor high pitch limits (power on). For rotocraft, except helicopters required to have a main rotor low speed warning under paragraph (e) of this section. It must be shown, with power on and without exceeding approved engine maximum limitations, that main rotor speeds substantially less than the minimum approved main rotor speed will not occur under any sustained flight condition. This must be met by— (1) Appropriate setting of the main rotor high pitch stop; (2) Inherent rotorcraft characteristics that make unsafe low main rotor speeds unlikely; or (3) Adequate means to warn the pilot of unsafe main rotor speeds. (c) Normal main rotor low pitch limits (power off). It must be shown, with power off, that— (1) The normal main rotor low pitch limit provides sufficient rotor speed, in any autorotative condition, under the most critical combinations of weight and airspeed; and (2) It is possible to prevent overspeeding of the rotor without exceptional piloting skill. (d) Emergency high pitch. If the main rotor high pitch stop is set to meet paragraph (b)(1) of this section, and if that stop cannot be exceeded inadvertently, additional pitch may be made available for emergency use. (e) Main rotor low speed warning for helicopters. For each single engine helicopter, and each multiengine helicopter that does not have an approved device that automatically increases power on the operating engines when one engine fails, there must be a main rotor low speed warning which meets the following requirements: (1) The warning must be furnished to the pilo… | |||
| 14:14:1.0.1.3.15.2.162.10 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.65 Climb: all engines operating. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 43 FR 2324, Jan. 16, 1978; Amdt. 27-33, 61 FR 21907, May 10, 1996] | (a) For rotorcraft other than helicopters— (1) The steady rate of climb, at V Y, must be determined— (i) With maximum continuous power on each engine; (ii) With the landing gear retracted; and (iii) For the weights, altitudes, and temperatures for which certification is requested; and (2) The climb gradient, at the rate of climb determined in accordance with paragraph (a)(1) of this section, must be either— (i) At least 1:10 if the horizontal distance required to take off and climb over a 50-foot obstacle is determined for each weight, altitude, and temperature within the range for which certification is requested; or (ii) At least 1:6 under standard sea level conditions. (b) Each helicopter must meet the following requirements: (1) V Y must be determined— (i) For standard sea level conditions; (ii) At maximum weight; and (iii) With maximum continuous power on each engine. (2) The steady rate of climb must be determined— (i) At the climb speed selected by the applicant at or below V NE ; (ii) Within the range from sea level up to the maximum altitude for which certification is requested; (iii) For the weights and temperatures that correspond to the altitude range set forth in paragraph (b)(2)(ii) of this section and for which certification is requested; and (iv) With maximum continuous power on each engine. | |||
| 14:14:1.0.1.3.15.2.162.11 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.67 Climb: one engine inoperative. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 53 FR 34210, Sept. 2, 1988] | For multiengine helicopters, the steady rate of climb (or descent), at V y (or at the speed for minimum rate of descent), must be determined with— (a) Maximum weight; (b) The critical engine inoperative and the remaining engines at either— (1) Maximum continuous power and, for helicopters for which certification for the use of 30-minute OEI power is requested, at 30-minute OEI power; or (2) Continuous OEI power for helicopters for which certification for the use of continuous OEI power is requested. | |||
| 14:14:1.0.1.3.15.2.162.12 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.71 Autorotation performance. | FAA | [Amdt. 27-21, 49 FR 44433, Nov. 6, 1984] | For single-engine helicopters and multiengine helicopters that do not meet the Category A engine isolation requirements of Part 29 of this chapter, the minimum rate of descent airspeed and the best angle-of-glide airspeed must be determined in autorotation at— (a) Maximum weight; and (b) Rotor speed(s) selected by the applicant. | |||
| 14:14:1.0.1.3.15.2.162.13 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.75 Landing. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 43 FR 2324, Jan. 16, 1978; Amdt. 27-44, 73 FR 10999, Feb. 29, 2008] | (a) The rotorcraft must be able to be landed with no excessive vertical acceleration, no tendency to bounce, nose over, ground loop, porpoise, or water loop, and without exceptional piloting skill or exceptionally favorable conditions, with— (1) Approach or autorotation speeds appropriate to the type of rotorcraft and selected by the applicant; (2) The approach and landing made with— (i) Power off, for single engine rotorcraft and entered from steady state autorotation; or (ii) One-engine inoperative (OEI) for multiengine rotorcraft, with each operating engine within approved operating limitations, and entered from an established OEI approach. (b) Multiengine rotorcraft must be able to be landed safely after complete power failure under normal operating conditions. | |||
| 14:14:1.0.1.3.15.2.162.14 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.87 Height-velocity envelope. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 43 FR 2324, Jan. 16, 1978; Amdt. 27-21, 49 FR 44433, Nov. 6, 1984; Amdt. 27-44, 73 FR 10999, Feb. 29, 2008; Amdt. 27-51, 88 FR 8737, Feb. 10, 2023] | (a) If there is any combination of height and forward velocity (including hover) under which a safe landing cannot be made under the applicable power failure condition in paragraph (b) of this section, a limiting height-velocity envelope must be established (including all pertinent information) for that condition, throughout the ranges of— (1) Altitude, from standard sea level conditions to the maximum altitude capability of the rotorcraft, or 7000 feet density altitude, whichever is less; and (2) Weight, from the maximum weight at sea level to the weight selected by the applicant for each altitude covered by paragraph (a)(1) of this section. For helicopters, the weight at altitudes above sea level may not be less than the maximum weight or the highest weight allowing hovering out-of-ground effect, whichever is lower. (b) The applicable power failure conditions are— (1) For single-engine helicopters, full autorotation; (2) For multiengine helicopters, OEI (where engine isolation features ensure continued operation of the remaining engines), and the remaining engine(s) within approved limits and at the minimum installed specification power available for the most critical combination of approved ambient temperature and pressure altitude resulting in 7000 feet density altitude or the maximum altitude capability of the helicopter, whichever is less, and (3) For other rotorcraft, conditions appropriate to the type. | |||
| 14:14:1.0.1.3.15.2.162.7 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.45 General. | FAA | [Amdt. 27-14, 43 FR 2324, Jan. 16, 1978, as amended by Amdt. 27-21, 49 FR 44432, Nov. 6, 1984] | (a) Unless otherwise prescribed, the performance requirements of this subpart must be met for still air and a standard atmosphere. (b) The performance must correspond to the engine power available under the particular ambient atmospheric conditions, the particular flight condition, and the relative humidity specified in paragraphs (d) or (e) of this section, as appropriate. (c) The available power must correspond to engine power, not exceeding the approved power, less— (1) Installation losses; and (2) The power absorbed by the accessories and services appropriate to the particular ambient atmospheric conditions and the particular flight condition. (d) For reciprocating engine-powered rotorcraft, the performance, as affected by engine power, must be based on a relative humidity of 80 percent in a standard atmosphere. (e) For turbine engine-powered rotorcraft, the performance, as affected by engine power, must be based on a relative humidity of— (1) 80 percent, at and below standard temperature; and (2) 34 percent, at and above standard temperature plus 50 degrees F. Between these two temperatures, the relative humidity must vary linearly. (f) For turbine-engine-powered rotorcraft, a means must be provided to permit the pilot to determine prior to takeoff that each engine is capable of developing the power necessary to achieve the applicable rotorcraft performance prescribed in this subpart. | |||
| 14:14:1.0.1.3.15.2.162.8 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.49 Performance at minimum operating speed. | FAA | [Amdt. 27-44, 73 FR 10998, Feb. 29, 2008] | (a) For helicopters— (1) The hovering ceiling must be determined over the ranges of weight, altitude, and temperature for which certification is requested, with— (i) Takeoff power; (ii) The landing gear extended; and (iii) The helicopter in-ground effect at a height consistent with normal takeoff procedures; and (2) The hovering ceiling determined under paragraph (a)(1) of this section must be at least— (i) For reciprocating engine powered helicopters, 4,000 feet at maximum weight with a standard atmosphere; (ii) For turbine engine powered helicopters, 2,500 feet pressure altitude at maximum weight at a temperature of standard plus 22 °C (standard plus 40 °F). (3) The out-of-ground effect hovering performance must be determined over the ranges of weight, altitude, and temperature for which certification is requested, using takeoff power. (b) For rotorcraft other than helicopters, the steady rate of climb at the minimum operating speed must be determined over the ranges of weight, altitude, and temperature for which certification is requested, with— (1) Takeoff power; and (2) The landing gear extended. | |||
| 14:14:1.0.1.3.15.2.162.9 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.51 Takeoff. | FAA | [Amdt. 27-44, 73 FR 10999, Feb. 29, 2008] | The takeoff, with takeoff power and r.p.m. at the most critical center of gravity, and with weight from the maximum weight at sea level to the weight for which takeoff certification is requested for each altitude covered by this section— (a) May not require exceptional piloting skill or exceptionally favorable conditions throughout the ranges of altitude from standard sea level conditions to the maximum altitude for which takeoff and landing certification is requested, and (b) Must be made in such a manner that a landing can be made safely at any point along the flight path if an engine fails. This must be demonstrated up to the maximum altitude for which takeoff and landing certification is requested or 7,000 feet density altitude, whichever is less. | |||
| 14:14:1.0.1.3.15.2.163.15 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.141 General. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR 962, Jan. 26, 1968; Amdt. 27-11, 41 FR 55468, Dec. 20, 1976; Amdt. 27-19, 48 FR 4389, Jan. 31, 1983; Amdt. 27-21, 49 FR 44433, Nov. 6, 1984] | The rotorcraft must— (a) Except as specifically required in the applicable section, meet the flight characteristics requirements of this subpart— (1) At the altitudes and temperatures expected in operation; (2) Under any critical loading condition within the range of weights and centers of gravity for which certification is requested; (3) For power-on operations, under any condition of speed, power, and rotor r.p.m. for which certification is requested; and (4) For power-off operations, under any condition of speed and rotor r.p.m. for which certification is requested that is attainable with the controls rigged in accordance with the approved rigging instructions and tolerances; (b) Be able to maintain any required flight condition and make a smooth transition from any flight condition to any other flight condition without exceptional piloting skill, alertness, or strength, and without danger of exceeding the limit load factor under any operating condition probable for the type, including— (1) Sudden failure of one engine, for multiengine rotorcraft meeting Transport Category A engine isolation requirements of Part 29 of this chapter; (2) Sudden, complete power failure for other rotorcraft; and (3) Sudden, complete control system failures specified in § 27.695 of this part; and (c) Have any additional characteristic required for night or instrument operation, if certification for those kinds of operation is requested. Requirements for helicopter instrument flight are contained in appendix B of this part. | |||
| 14:14:1.0.1.3.15.2.163.16 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.143 Controllability and maneuverability. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR 963, Jan. 26, 1968; Amdt. 27-14, 43 FR 2325, Jan. 16, 1978; Amdt. 27-21, 49 FR 44433, Nov. 6, 1984; Amdt. 27-44, 73 FR 10999, Feb. 29, 2008] | (a) The rotorcraft must be safely controllable and maneuverable— (1) During steady flight; and (2) During any maneuver appropriate to the type, including— (i) Takeoff; (ii) Climb; (iii) Level flight; (iv) Turning flight; (v) Autorotation; (vi) Landing (power on and power off); and (vii) Recovery to power-on flight from a balked autorotative approach. (b) The margin of cyclic control must allow satisfactory roll and pitch control at V NE with— (1) Critical weight; (2) Critical center of gravity; (3) Critical rotor r.p.m.; and (4) Power off (except for helicopters demonstrating compliance with paragraph (f) of this section) and power on. (c) Wind velocities from zero to at least 17 knots, from all azimuths, must be established in which the rotorcraft can be operated without loss of control on or near the ground in any maneuver appropriate to the type (such as crosswind takeoffs, sideward flight, and rearward flight)— (1) With altitude, from standard sea level conditions to the maximum takeoff and landing altitude capability of the rotorcraft or 7000 feet density altitude, whichever is less; with— (i) Critical Weight; (ii) Critical center of gravity; (iii) Critical rotor r.p.m.; (2) For takeoff and landing altitudes above 7000 feet density altitude with— (i) Weight selected by the applicant; (ii) Critical center of gravity; and (iii) Critical rotor r.p.m. (d) Wind velocities from zero to at least 17 knots, from all azimuths, must be established in which the rotorcraft can be operated without loss of control out-of-ground-effect, with— (1) Weight selected by the applicant; (2) Critical center of gravity; (3) Rotor r.p.m. selected by the applicant; and (4) Altitude, from standard sea level conditions to the maximum takeoff and landing altitude capability of the rotorcraft. (e) The rotorcraft, after (1) failure of one engine in the case of multiengine rotorcraft that meet Transport Category A engine isolation requirements, or (2) complete engine failure in the case of other rotorcraft, mus… | |||
| 14:14:1.0.1.3.15.2.163.17 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.151 Flight controls. | FAA | [Amdt. 27-21, 49 FR 44433, Nov. 6, 1984] | (a) Longitudinal, lateral, directional, and collective controls may not exhibit excessive breakout force, friction, or preload. (b) Control system forces and free play may not inhibit a smooth, direct rotorcraft response to control system input. | |||
| 14:14:1.0.1.3.15.2.163.18 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.161 Trim control. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 49 FR 44433, Nov. 6, 1984] | The trim control— (a) Must trim any steady longitudinal, lateral, and collective control forces to zero in level flight at any appropriate speed; and (b) May not introduce any undesirable discontinuities in control force gradients. | |||
| 14:14:1.0.1.3.15.2.163.19 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.171 Stability: general. | FAA | The rotorcraft must be able to be flown, without undue pilot fatigue or strain, in any normal maneuver for a period of time as long as that expected in normal operation. At least three landings and takeoffs must be made during this demonstration. | ||||
| 14:14:1.0.1.3.15.2.163.20 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.173 Static longitudinal stability. | FAA | [Amdt. 27-21, 49 FR 44433, Nov. 6, 1984, as amended by Amdt. 27-44, 73 FR 10999, Feb. 29, 2008] | (a) The longitudinal control must be designed so that a rearward movement of the control is necessary to obtain an airspeed less than the trim speed, and a forward movement of the control is necessary to obtain an airspeed more than the trim speed. (b) Throughout the full range of altitude for which certification is requested, with the throttle and collective pitch held constant during the maneuvers specified in § 27.175(a) through (d), the slope of the control position versus airspeed curve must be positive. However, in limited flight conditions or modes of operation determined by the Administrator to be acceptable, the slope of the control position versus airspeed curve may be neutral or negative if the rotorcraft possesses flight characteristics that allow the pilot to maintain airspeed within ±5 knots of the desired trim airspeed without exceptional piloting skill or alertness. | |||
| 14:14:1.0.1.3.15.2.163.21 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.175 Demonstration of static longitudinal stability. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR 963, Jan. 26, 1968; Amdt. 27-11, 41 FR 55468, Dec. 20, 1976; Amdt. 27-14, 43 FR 2325, Jan. 16, 1978; Amdt. 27-21, 49 FR 44433, Nov. 6, 1984; Amdt. 27-34, 62 FR 46173, Aug. 29, 1997; Amdt. 27-44, 73 FR 10999, Feb. 29, 2008] | (a) Climb. Static longitudinal stability must be shown in the climb condition at speeds from Vy − 10 kt to Vy + 10 kt with— (1) Critical weight; (2) Critical center of gravity; (3) Maximum continuous power; (4) The landing gear retracted; and (5) The rotorcraft trimmed at V Y. (b) Cruise. Static longitudinal stability must be shown in the cruise condition at speeds from 0.8 V NE − 10 kt to 0.8 V NE + 10 kt or, if V H is less than 0.8 V NE , from V H −10 kt to V H + 10 kt, with— (1) Critical weight; (2) Critical center of gravity; (3) Power for level flight at 0.8 V NE or V H , whichever is less; (4) The landing gear retracted; and (5) The rotorcraft trimmed at 0.8 V NE or V H , whichever is less. (c) V NE. Static longitudinal stability must be shown at speeds from V NE − 20 kt to V NE with— (1) Critical weight; (2) Critical center of gravity; (3) Power required for level flight at V NE −10 kt or maximum continuous power, whichever is less; (4) The landing gear retracted; and (5) The rotorcraft trimmed at V NE − 10 kt. (d) Autorotation. Static longitudinal stability must be shown in autorotation at— (1) Airspeeds from the minimum rate of descent airspeed−10 kt to the minimum rate of descent airspeed + 10 kt, with— (i) Critical weight; (ii) Critical center of gravity; (iii) The landing gear extended; and (iv) The rotorcraft trimmed at the minimum rate of descent airspeed. (2) Airspeeds from best angle-of-glide airspeed−10 kt to the best angle-of-glide airspeed + 10 kt, with— (i) Critical weight; (ii) Critical center of gravity; (iii) The landing gear retracted; and (iv) The rotorcraft trimmed at the best angle-of-glide airspeed. | |||
| 14:14:1.0.1.3.15.2.163.22 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.177 Static directional stability. | FAA | [Amdt. 27-44, 73 FR 11000, Feb. 29, 2008] | (a) The directional controls must operate in such a manner that the sense and direction of motion of the rotorcraft following control displacement are in the direction of the pedal motion with the throttle and collective controls held constant at the trim conditions specified in § 27.175(a), (b), and (c). Sideslip angles must increase with steadily increasing directional control deflection for sideslip angles up to the lesser of— (1) ±25 degrees from trim at a speed of 15 knots less than the speed for minimum rate of descent varying linearly to ±10 degrees from trim at V NE ; (2) The steady state sideslip angles established by § 27.351; (3) A sideslip angle selected by the applicant, which corresponds to a sideforce of at least 0.1g; or (4) The sideslip angle attained by maximum directional control input. (b) Sufficient cues must accompany the sideslip to alert the pilot when the aircraft is approaching the sideslip limits. (c) During the maneuver specified in paragraph (a) of this section, the sideslip angle versus directional control position curve may have a negative slope within a small range of angles around trim, provided the desired heading can be maintained without exceptional piloting skill or alertness. | |||
| 14:14:1.0.1.3.15.2.164.23 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.231 General. | FAA | The rotorcraft must have satisfactory ground and water handling characteristics, including freedom from uncontrollable tendencies in any condition expected in operation. | ||||
| 14:14:1.0.1.3.15.2.164.24 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.235 Taxiing condition. | FAA | The rotorcraft must be designed to withstand the loads that would occur when the rotorcraft is taxied over the roughest ground that may reasonably be expected in normal operation. | ||||
| 14:14:1.0.1.3.15.2.164.25 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.239 Spray characteristics. | FAA | If certification for water operation is requested, no spray characteristics during taxiing, takeoff, or landing may obscure the vision of the pilot or damage the rotors, propellers, or other parts of the rotorcraft. | ||||
| 14:14:1.0.1.3.15.2.164.26 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.241 Ground resonance. | FAA | The rotorcraft may have no dangerous tendency to oscillate on the ground with the rotor turning. | ||||
| 14:14:1.0.1.3.15.2.165.27 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | B | Subpart B—Flight | § 27.251 Vibration. | FAA | Each part of the rotorcraft must be free from excessive vibration under each appropriate speed and power condition. | ||||
| 14:14:1.0.1.3.15.3.166.1 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.301 Loads. | FAA | (a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads. (b) Unless otherwise provided, the specified air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the rotorcraft. These loads must be distributed to closely approximate or conservatively represent actual conditions. (c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account. | ||||
| 14:14:1.0.1.3.15.3.166.2 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.303 Factor of safety. | FAA | Unless otherwise provided, a factor of safety of 1.5 must be used. This factor applies to external and inertia loads unless its application to the resulting internal stresses is more conservative. | ||||
| 14:14:1.0.1.3.15.3.166.3 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.305 Strength and deformation. | FAA | (a) The structure must be able to support limit loads without detrimental or permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation. (b) The structure must be able to support ultimate loads without failure. This must be shown by— (1) Applying ultimate loads to the structure in a static test for at least three seconds; or (2) Dynamic tests simulating actual load application. | ||||
| 14:14:1.0.1.3.15.3.166.4 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.307 Proof of structure. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-3, 33 FR 14105, Sept. 18, 1968; Amdt. 27-26, 55 FR 7999, Mar. 6, 1990] | (a) Compliance with the strength and deformation requirements of this subpart must be shown for each critical loading condition accounting for the environment to which the structure will be exposed in operation. Structural analysis (static or fatigue) may be used only if the structure conforms to those structures for which experience has shown this method to be reliable. In other cases, substantiating load tests must be made. (b) Proof of compliance with the strength requirements of this subpart must include— (1) Dynamic and endurance tests of rotors, rotor drives, and rotor controls; (2) Limit load tests of the control system, including control surfaces; (3) Operation tests of the control system; (4) Flight stress measurement tests; (5) Landing gear drop tests; and (6) Any additional test required for new or unusual design features. | |||
| 14:14:1.0.1.3.15.3.166.5 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.309 Design limitations. | FAA | The following values and limitations must be established to show compliance with the structural requirements of this subpart: (a) The design maximum weight. (b) The main rotor r.p.m. ranges power on and power off. (c) The maximum forward speeds for each main rotor r.p.m. within the ranges determined under paragraph (b) of this section. (d) The maximum rearward and sideward flight speeds. (e) The center of gravity limits corresponding to the limitations determined under paragraphs (b), (c), and (d) of this section. (f) The rotational speed ratios between each powerplant and each connected rotating component. (g) The positive and negative limit maneuvering load factors. | ||||
| 14:14:1.0.1.3.15.3.167.10 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.351 Yawing conditions. | FAA | [Amdt. 27-26, 55 FR 7999, Mar. 6, 1990, as amended by Amdt. 27-34, 62 FR 46173, Aug. 29, 1997] | (a) Each rotorcraft must be designed for the loads resulting from the maneuvers specified in paragraphs (b) and (c) of this section with— (1) Unbalanced aerodynamic moments about the center of gravity which the aircraft reacts to in a rational or conservative manner considering the principal masses furnishing the reacting inertia forces; and (2) Maximum main rotor speed. (b) To produce the load required in paragraph (a) of this section, in unaccelerated flight with zero yaw, at forward speeds from zero up to 0.6 V NE — (1) Displace the cockpit directional control suddenly to the maximum deflection limited by the control stops or by the maximum pilot force specified in § 27.397(a); (2) Attain a resulting sideslip angle or 90°, whichever is less; and (3) Return the directional control suddenly to neutral. (c) To produce the load required in paragraph (a) of this section, in unaccelerated flight with zero yaw, at forward speeds from 0.6 V NE up to V NE or V H , whichever is less— (1) Displace the cockpit directional control suddenly to the maximum deflection limited by the control stops or by the maximum pilot force specified in § 27.397(a); (2) Attain a resulting sideslip angle or 15°, whichever is less, at the lesser speed of V NE or V H ; (3) Vary the sideslip angles of paragraphs (b)(2) and (c)(2) of this section directly with speed; and (4) Return the directional control suddenly to neutral. | |||
| 14:14:1.0.1.3.15.3.167.11 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.361 Engine torque. | FAA | [Amdt. 27-23, 53 FR 34210, Sept. 2, 1988] | (a) For turbine engines, the limit torque may not be less than the highest of— (1) The mean torque for maximum continuous power multiplied by 1.25; (2) The torque required by § 27.923; (3) The torque required by § 27.927; or (4) The torque imposed by sudden engine stoppage due to malfunction or structural failure (such as compressor jamming). (b) For reciprocating engines, the limit torque may not be less than the mean torque for maximum continuous power multiplied by— (1) 1.33, for engines with five or more cylinders; and (2) Two, three, and four, for engines with four, three, and two cylinders, respectively. | |||
| 14:14:1.0.1.3.15.3.167.6 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.321 General. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR 55468, Dec. 20, 1976] | (a) The flight load factor must be assumed to act normal to the longitudinal axis of the rotorcraft, and to be equal in magnitude and opposite in direction to the rotorcraft inertia load factor at the center of gravity. (b) Compliance with the flight load requirements of this subpart must be shown— (1) At each weight from the design minimum weight to the design maximum weight; and (2) With any practical distribution of disposable load within the operating limitations in the Rotorcraft Flight Manual. | |||
| 14:14:1.0.1.3.15.3.167.7 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.337 Limit maneuvering load factor. | FAA | [Amdt. 27-26, 55 FR 7999, Mar. 6, 1990] | The rotorcraft must be designed for— (a) A limit maneuvering load factor ranging from a positive limit of 3.5 to a negative limit of −1.0; or (b) Any positive limit maneuvering load factor not less than 2.0 and any negative limit maneuvering load factor of not less than −0.5 for which— (1) The probability of being exceeded is shown by analysis and flight tests to be extremely remote; and (2) The selected values are appropriate to each weight condition between the design maximum and design minimum weights. | |||
| 14:14:1.0.1.3.15.3.167.8 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.339 Resultant limit maneuvering loads. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR 55469, Dec. 20, 1976] | The loads resulting from the application of limit maneuvering load factors are assumed to act at the center of each rotor hub and at each auxiliary lifting surface, and to act in directions, and with distributions of load among the rotors and auxiliary lifting surfaces, so as to represent each critical maneuvering condition, including power-on and power-off flight with the maximum design rotor tip speed ratio. The rotor tip speed ratio is the ratio of the rotorcraft flight velocity component in the plane of the rotor disc to the rotational tip speed of the rotor blades, and is expressed as follows: where— V = The airspeed along flight path (f.p.s.); a = The angle between the projection, in the plane of symmetry, of the axis of no feathering and a line perpendicular to the flight path (radians, positive when axis is pointing aft); omega = The angular velocity of rotor (radians per second); and R = The rotor radius (ft). where— V = The airspeed along flight path (f.p.s.); a = The angle between the projection, in the plane of symmetry, of the axis of no feathering and a line perpendicular to the flight path (radians, positive when axis is pointing aft); omega = The angular velocity of rotor (radians per second); and R = The rotor radius (ft). | |||
| 14:14:1.0.1.3.15.3.167.9 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.341 Gust loads. | FAA | The rotorcraft must be designed to withstand, at each critical airspeed including hovering, the loads resulting from a vertical gust of 30 feet per second. | ||||
| 14:14:1.0.1.3.15.3.168.12 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.391 General. | FAA | [Amdt. 27-26, 55 FR 7999, Mar. 6, 1990, as amended by Amdt. 27-34, 62 FR 46173, Aug. 29, 1997] | Each auxiliary rotor, each fixed or movable stabilizing or control surface, and each system operating any flight control must meet the requirements of §§ 27.395, 27.397, 27.399, 27.411, and 27.427. | |||
| 14:14:1.0.1.3.15.3.168.13 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.395 Control system. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-26, 55 FR 7999, Mar. 6, 1990] | (a) The part of each control system from the pilot's controls to the control stops must be designed to withstand pilot forces of not less than— (1) The forces specified in § 27.397; or (2) If the system prevents the pilot from applying the limit pilot forces to the system, the maximum forces that the system allows the pilot to apply, but not less than 0.60 times the forces specified in § 27.397. (b) Each primary control system, including its supporting structure, must be designed as follows: (1) The system must withstand loads resulting from the limit pilot forces prescribed in § 27.397. (2) Notwithstanding paragraph (b)(3) of this section, when power-operated actuator controls or power boost controls are used, the system must also withstand the loads resulting from the force output of each normally energized power device, including any single power boost or actuator system failure. (3) If the system design or the normal operating loads are such that a part of the system cannot react to the limit pilot forces prescribed in § 27.397, that part of the system must be designed to withstand the maximum loads that can be obtained in normal operation. The minimum design loads must, in any case, provide a rugged system for service use, including consideration of fatigue, jamming, ground gusts, control inertia, and friction loads. In the absence of rational analysis, the design loads resulting from 0.60 of the specified limit pilot forces are acceptable minimum design loads. (4) If operational loads may be exceeded through jamming, ground gusts, control inertia, or friction, the system must withstand the limit pilot forces specified in § 27.397, without yielding. | |||
| 14:14:1.0.1.3.15.3.168.14 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.397 Limit pilot forces and torques. | FAA | [Amdt. 27-11, 41 FR 55469, Dec. 20, 1976, as amended by Amdt. 27-40, 66 FR 23538, May 9, 2001] | (a) Except as provided in paragraph (b) of this section, the limit pilot forces are as follows: (1) For foot controls, 130 pounds. (2) For stick controls, 100 pounds fore and aft, and 67 pounds laterally. (b) For flap, tab, stabilizer, rotor brake, and landing gear operating controls, the follows apply (R = radius in inches): (1) Crank, wheel, and lever controls, [1 + R]/3 × 50 pounds, but not less than 50 pounds nor more than 100 pounds for hand operated controls or 130 pounds for foot operated controls, applied at any angle within 20 degrees of the plane of motion of the control. (2) Twist controls, 80R inch-pounds. | |||
| 14:14:1.0.1.3.15.3.168.15 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.399 Dual control system. | FAA | Each dual primary flight control system must be designed to withstand the loads that result when pilot forces of 0.75 times those obtained under § 27.395 are applied— (a) In opposition; and (b) In the same direction. | ||||
| 14:14:1.0.1.3.15.3.168.16 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.411 Ground clearance: tail rotor guard. | FAA | (a) It must be impossible for the tail rotor to contact the landing surface during a normal landing. (b) If a tail rotor guard is required to show compliance with paragraph (a) of this section— (1) Suitable design loads must be established for the guard; and (2) The guard and its supporting structure must be designed to withstand those loads. | ||||
| 14:14:1.0.1.3.15.3.168.17 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.427 Unsymmetrical loads. | FAA | [Amdt. 27-26, 55 FR 7999, Mar. 6, 1990, as amended by Amdt. 27-27, 55 FR 38966, Sept. 21, 1990] | (a) Horizontal tail surfaces and their supporting structure must be designed for unsymmetrical loads arising from yawing and rotor wake effects in combination with the prescribed flight conditions. (b) To meet the design criteria of paragraph (a) of this section, in the absence of more rational data, both of the following must be met: (1) One hundred percent of the maximum loading from the symmetrical flight conditions acts on the surface on one side of the plane of symmetry, and no loading acts on the other side. (2) Fifty percent of the maximum loading from the symmetrical flight conditions acts on the surface on each side of the plane of symmetry but in opposite directions. (c) For empennage arrangements where the horizontal tail surfaces are supported by the vertical tail surfaces, the vertical tail surfaces and supporting structure must be designed for the combined vertical and horizontal surface loads resulting from each prescribed flight condition, considered separately. The flight conditions must be selected so the maximum design loads are obtained on each surface. In the absence of more rational data, the unsymmetrical horizontal tail surface loading distributions described in this section must be assumed. | |||
| 14:14:1.0.1.3.15.3.169.18 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.471 General. | FAA | (a) Loads and equilibrium. For limit ground loads— (1) The limit ground loads obtained in the landing conditions in this part must be considered to be external loads that would occur in the rotorcraft structure if it were acting as a rigid body; and (2) In each specified landing condition, the external loads must be placed in equilibrium with linear and angular inertia loads in a rational or conservative manner. (b) Critical centers of gravity. The critical centers of gravity within the range for which certification is requested must be selected so that the maximum design loads are obtained in each landing gear element. | ||||
| 14:14:1.0.1.3.15.3.169.19 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.473 Ground loading conditions and assumptions. | FAA | [Amdt. 27-2, 33 FR 963, Jan. 26, 1968] | (a) For specified landing conditions, a design maximum weight must be used that is not less than the maximum weight. A rotor lift may be assumed to act through the center of gravity throughout the landing impact. This lift may not exceed two-thirds of the design maximum weight. (b) Unless otherwise prescribed, for each specified landing condition, the rotorcraft must be designed for a limit load factor of not less than the limit inertia load factor substantiated under § 27.725. | |||
| 14:14:1.0.1.3.15.3.169.20 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.475 Tires and shock absorbers. | FAA | Unless otherwise prescribed, for each specified landing condition, the tires must be assumed to be in their static position and the shock absorbers to be in their most critical position. | ||||
| 14:14:1.0.1.3.15.3.169.21 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.477 Landing gear arrangement. | FAA | Sections 27.235, 27.479 through 27.485, and 27.493 apply to landing gear with two wheels aft, and one or more wheels forward, of the center of gravity. | ||||
| 14:14:1.0.1.3.15.3.169.22 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.479 Level landing conditions. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964; 29 FR 17885, Dec. 17, 1964] | (a) Attitudes. Under each of the loading conditions prescribed in paragraph (b) of this section, the rotorcraft is assumed to be in each of the following level landing attitudes: (1) An attitude in which all wheels contact the ground simultaneously. (2) An attitude in which the aft wheels contact the ground with the forward wheels just clear of the ground. (b) Loading conditions. The rotorcraft must be designed for the following landing loading conditions: (1) Vertical loads applied under § 27.471. (2) The loads resulting from a combination of the loads applied under paragraph (b)(1) of this section with drag loads at each wheel of not less than 25 percent of the vertical load at that wheel. (3) If there are two wheels forward, a distribution of the loads applied to those wheels under paragraphs (b)(1) and (2) of this section in a ratio of 40:60. (c) Pitching moments. Pitching moments are assumed to be resisted by— (1) In the case of the attitude in paragraph (a)(1) of this section, the forward landing gear; and (2) In the case of the attitude in paragraph (a)(2) of this section, the angular inertia forces. | |||
| 14:14:1.0.1.3.15.3.169.23 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.481 Tail-down landing conditions. | FAA | (a) The rotorcraft is assumed to be in the maximum nose-up attitude allowing ground clearance by each part of the rotorcraft. (b) In this attitude, ground loads are assumed to act perpendicular to the ground. | ||||
| 14:14:1.0.1.3.15.3.169.24 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.483 One-wheel landing conditions. | FAA | For the one-wheel landing condition, the rotorcraft is assumed to be in the level attitude and to contact the ground on one aft wheel. In this attitude— (a) The vertical load must be the same as that obtained on that side under § 27.479(b)(1); and (b) The unbalanced external loads must be reacted by rotorcraft inertia. | ||||
| 14:14:1.0.1.3.15.3.169.25 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.485 Lateral drift landing conditions. | FAA | (a) The rotorcraft is assumed to be in the level landing attitude, with— (1) Side loads combined with one-half of the maximum ground reactions obtained in the level landing conditions of § 27.479 (b)(1); and (2) The loads obtained under paragraph (a)(1) of this section applied— (i) At the ground contact point; or (ii) For full-swiveling gear, at the center of the axle. (b) The rotorcraft must be designed to withstand, at ground contact— (1) When only the aft wheels contact the ground, side loads of 0.8 times the vertical reaction acting inward on one side, and 0.6 times the vertical reaction acting outward on the other side, all combined with the vertical loads specified in paragraph (a) of this section; and (2) When all wheels contact the ground simultaneously— (i) For the aft wheels, the side loads specified in paragraph (b)(1) of this section; and (ii) For the forward wheels, a side load of 0.8 times the vertical reaction combined with the vertical load specified in paragraph (a) of this section. | ||||
| 14:14:1.0.1.3.15.3.169.26 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.493 Braked roll conditions. | FAA | Under braked roll conditions with the shock absorbers in their static positions— (a) The limit vertical load must be based on a load factor of at least— (1) 1.33, for the attitude specified in § 27.479(a)(1); and (2) 1.0 for the attitude specified in § 27.479(a)(2); and (b) The structure must be designed to withstand at the ground contact point of each wheel with brakes, a drag load at least the lesser of— (1) The vertical load multiplied by a coefficient of friction of 0.8; and (2) The maximum value based on limiting brake torque. | ||||
| 14:14:1.0.1.3.15.3.169.27 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.497 Ground loading conditions: landing gear with tail wheels. | FAA | (a) General. Rotorcraft with landing gear with two wheels forward, and one wheel aft, of the center of gravity must be designed for loading conditions as prescribed in this section. (b) Level landing attitude with only the forward wheels contacting the ground. In this attitude— (1) The vertical loads must be applied under §§ 27.471 through 27.475; (2) The vertical load at each axle must be combined with a drag load at that axle of not less than 25 percent of that vertical load; and (3) Unbalanced pitching moments are assumed to be resisted by angular inertia forces. (c) Level landing attitude with all wheels contacting the ground simultaneously. In this attitude, the rotorcraft must be designed for landing loading conditions as prescribed in paragraph (b) of this section. (d) Maximum nose-up attitude with only the rear wheel contacting the ground. The attitude for this condition must be the maximum nose-up attitude expected in normal operation, including autorotative landings. In this attitude— (1) The appropriate ground loads specified in paragraphs (b)(1) and (2) of this section must be determined and applied, using a rational method to account for the moment arm between the rear wheel ground reaction and the rotorcraft center of gravity; or (2) The probability of landing with initial contact on the rear wheel must be shown to be extremely remote. (e) Level landing attitude with only one forward wheel contacting the ground. In this attitude, the rotorcraft must be designed for ground loads as specified in paragraphs (b)(1) and (3) of this section. (f) Side loads in the level landing attitude. In the attitudes specified in paragraphs (b) and (c) of this section, the following apply: (1) The side loads must be combined at each wheel with one-half of the maximum vertical ground reactions obtained for that wheel under paragraphs (b) and (c) of this section. In this condition, the side loads must be— (i) For the forward wheels, 0.8 times the vertical reaction (on one side) acting inward, and … | ||||
| 14:14:1.0.1.3.15.3.169.28 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.501 Ground loading conditions: landing gear with skids. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR 963, Jan. 26, 1968; Amdt. 27-26, 55 FR 8000, Mar. 6, 1990] | (a) General. Rotorcraft with landing gear with skids must be designed for the loading conditions specified in this section. In showing compliance with this section, the following apply: (1) The design maximum weight, center of gravity, and load factor must be determined under §§ 27.471 through 27.475. (2) Structural yielding of elastic spring members under limit loads is acceptable. (3) Design ultimate loads for elastic spring members need not exceed those obtained in a drop test of the gear with— (i) A drop height of 1.5 times that specified in § 27.725; and (ii) An assumed rotor lift of not more than 1.5 times that used in the limit drop tests prescribed in § 27.725. (4) Compliance with paragraphs (b) through (e) of this section must be shown with— (i) The gear in its most critically deflected position for the landing condition being considered; and (ii) The ground reactions rationally distributed along the bottom of the skid tube. (b) Vertical reactions in the level landing attitude. In the level attitude, and with the rotorcraft contacting the ground along the bottom of both skids, the vertical reactions must be applied as prescribed in paragraph (a) of this section. (c) Drag reactions in the level landing attitude. In the level attitude, and with the rotorcraft contacting the ground along the bottom of both skids, the following apply: (1) The vertical reactions must be combined with horizontal drag reactions of 50 percent of the vertical reaction applied at the ground. (2) The resultant ground loads must equal the vertical load specified in paragraph (b) of this section. (d) Sideloads in the level landing attitude. In the level attitude,and with the rotorcraft contacting the ground along the bottom of both skids, the following apply: (1) The vertical ground reaction must be— (i) Equal to the vertical loads obtained in the condition specified in paragraph (b) of this section; and (ii) Divided equally among the skids. (2) The vertical ground reactions must be combined with a horizontal… | |||
| 14:14:1.0.1.3.15.3.169.29 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.505 Ski landing conditions. | FAA | If certification for ski operation is requested, the rotorcraft, with skis, must be designed to withstand the following loading conditions (where P is the maximum static weight on each ski with the rotorcraft at design maximum weight, and n is the limit load factor determined under § 27.473(b). (a) Up-load conditions in which— (1) A vertical load of Pn and a horizontal load of Pn/ 4 are simultaneously applied at the pedestal bearings; and (2) A vertical load of 1.33 P is applied at the pedestal bearings. (b) A side-load condition in which a side load of 0.35 Pn is applied at the pedestal bearings in a horizontal plane perpendicular to the centerline of the rotorcraft. (c) A torque-load condition in which a torque load of 1.33 P (in foot pounds) is applied to the ski about the vertical axis through the centerline of the pedestal bearings. | ||||
| 14:14:1.0.1.3.15.3.170.30 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.521 Float landing conditions. | FAA | If certification for float operation is requested, the rotorcraft, with floats, must be designed to withstand the following loading conditions (where the limit load factor is determined under § 27.473(b) or assumed to be equal to that determined for wheel landing gear): (a) Up-load conditions in which— (1) A load is applied so that, with the rotorcraft in the static level attitude, the resultant water reaction passes vertically through the center of gravity; and (2) The vertical load prescribed in paragraph (a)(1) of this section is applied simultaneously with an aft component of 0.25 times the vertical component. (b) A side-load condition in which— (1) A vertical load of 0.75 times the total vertical load specified in paragraph (a)(1) of this section is divided equally among the floats; and (2) For each float, the load share determined under paragraph (b)(1) of this section, combined with a total side load of 0.25 times the total vertical load specified in paragraph (b)(1) of this section, is applied to that float only. | ||||
| 14:14:1.0.1.3.15.3.171.31 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.547 Main rotor structure. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-3, 33 FR 14105, Sept. 18, 1968] | (a) Each main rotor assembly (including rotor hubs and blades) must be designed as prescribed in this section. (b) [Reserved] (c) The main rotor structure must be designed to withstand the following loads prescribed in §§ 27.337 through 27.341: (1) Critical flight loads. (2) Limit loads occurring under normal conditions of autorotation. For this condition, the rotor r.p.m. must be selected to include the effects of altitude. (d) The main rotor structure must be designed to withstand loads simulating— (1) For the rotor blades, hubs, and flapping hinges, the impact force of each blade against its stop during ground operation; and (2) Any other critical condition expected in normal operation. (e) The main rotor structure must be designed to withstand the limit torque at any rotational speed, including zero. In addition: (1) The limit torque need not be greater than the torque defined by a torque limiting device (where provided), and may not be less than the greater of— (i) The maximum torque likely to be transmitted to the rotor structure in either direction; and (ii) The limit engine torque specified in § 27.361. (2) The limit torque must be distributed to the rotor blades in a rational manner. | |||
| 14:14:1.0.1.3.15.3.171.32 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.549 Fuselage, landing gear, and rotor pylon structures. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-3, 33 FR 14105, Sept. 18, 1968] | (a) Each fuselage, landing gear, and rotor pylon structure must be designed as prescribed in this section. Resultant rotor forces may be represented as a single force applied at the rotor hub attachment point. (b) Each structure must be designed to withstand— (1) The critical loads prescribed in §§ 27.337 through 27.341; (2) The applicable ground loads prescribed in §§ 27.235, 27.471 through 27.485, 27.493, 27.497, 27.501, 27.505, and 27.521; and (3) The loads prescribed in § 27.547 (d)(2) and (e). (c) Auxiliary rotor thrust, and the balancing air and inertia loads occurring under accelerated flight conditions, must be considered. (d) Each engine mount and adjacent fuselage structure must be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque. | |||
| 14:14:1.0.1.3.15.3.172.33 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.561 General. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-25, 54 FR 47318, Nov. 13, 1989; Amdt. 27-30, 59 FR 50386, Oct. 3, 1994; Amdt. 27-32, 61 FR 10438, Mar. 13, 1996] | (a) The rotorcraft, although it may be damaged in emergency landing conditions on land or water, must be designed as prescribed in this section to protect the occupants under those conditions. (b) The structure must be designed to give each occupant every reasonable chance of escaping serious injury in a crash landing when— (1) Proper use is made of seats, belts, and other safety design provisions; (2) The wheels are retracted (where applicable); and (3) Each occupant and each item of mass inside the cabin that could injure an occupant is restrained when subjected to the following ultimate inertial load factors relative to the surrounding structure: (i) Upward—4g. (ii) Forward—16g. (iii) Sideward—8g. (iv) Downward—20g, after intended displacement of the seat device. (v) Rearward—1.5g. (c) The supporting structure must be designed to restrain, under any ultimate inertial load up to those specified in this paragraph, any item of mass above and/or behind the crew and passenger compartment that could injure an occupant if it came loose in an emergency landing. Items of mass to be considered include, but are not limited to, rotors, transmissions, and engines. The items of mass must be restrained for the following ultimate inertial load factors: (1) Upward—1.5g. (2) Forward—12g. (3) Sideward—6g. (4) Downward—12g. (5) Rearward—1.5g (d) Any fuselage structure in the area of internal fuel tanks below the passenger floor level must be designed to resist the following ultimate inertial factors and loads and to protect the fuel tanks from rupture when those loads are applied to that area: (i) Upward—1.5g. (ii) Forward—4.0g. (iii) Sideward—2.0g. (iv) Downward—4.0g. | |||
| 14:14:1.0.1.3.15.3.172.34 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.562 Emergency landing dynamic conditions. | FAA | [Amdt. 27-25, 54 FR 47318, Nov. 13, 1989] | (a) The rotorcraft, although it may be damaged in an emergency crash landing, must be designed to reasonably protect each occupant when— (1) The occupant properly uses the seats, safety belts, and shoulder harnesses provided in the design; and (2) The occupant is exposed to the loads resulting from the conditions prescribed in this section. (b) Each seat type design or other seating device approved for crew or passenger occupancy during takeoff and landing must successfully complete dynamic tests or be demonstrated by rational analysis based on dynamic tests of a similar type seat in accordance with the following criteria. The tests must be conducted with an occupant, simulated by a 170-pound anthropomorphic test dummy (ATD), as defined by 49 CFR 572, subpart B, or its equivalent, sitting in the normal upright position. (1) A change in downward velocity of not less than 30 feet per second when the seat or other seating device is oriented in its nominal position with respect to the rotorcraft's reference system, the rotorcraft's longitudinal axis is canted upward 60° with respect to the impact velocity vector, and the rotorcraft's lateral axis is perpendicular to a vertical plane containing the impact velocity vector and the rotorcraft's longitudinal axis. Peak floor deceleration must occur in not more than 0.031 seconds after impact and must reach a minimum of 30g's. (2) A change in forward velocity of not less than 42 feet per second when the seat or other seating device is oriented in its nominal position with respect to the rotorcraft's reference system, the rotorcraft's longitudinal axis is yawed 10° either right or left of the impact velocity vector (whichever would cause the greatest load on the shoulder harness), the rotorcraft's lateral axis is contained in a horizontal plane containing the impact velocity vector, and the rotorcraft's vertical axis is perpendicular to a horizontal plane containing the impact velocity vector. Peak floor deceleration must occur in not more than 0.071 seconds after impa… | |||
| 14:14:1.0.1.3.15.3.172.35 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.563 Structural ditching provisions. | FAA | [Amdt. 27-26, 55 FR 8000, Mar. 6, 1990] | If certification with ditching provisions is requested, structural strength for ditching must meet the requirements of this section and § 27.801(e). (a) Forward speed landing conditions. The rotorcraft must initially contact the most critical wave for reasonably probable water conditions at forward velocities from zero up to 30 knots in likely pitch, roll, and yaw attitudes. The rotorcraft limit vertical descent velocity may not be less than 5 feet per second relative to the mean water surface. Rotor lift may be used to act through the center of gravity throughout the landing impact. This lift may not exceed two-thirds of the design maximum weight. A maximum forward velocity of less than 30 knots may be used in design if it can be demonstrated that the forward velocity selected would not be exceeded in a normal one-engine-out touchdown. (b) Auxiliary or emergency float conditions —(1) Floats fixed or deployed before initial water contact. In addition to the landing loads in paragraph (a) of this section, each auxiliary or emergency float, of its support and attaching structure in the airframe or fuselage, must be designed for the load developed by a fully immersed float unless it can be shown that full immersion is unlikely. If full immersion is unlikely, the highest likely float buoyancy load must be applied. The highest likely buoyancy load must include consideration of a partially immersed float creating restoring moments to compensate the upsetting moments caused by side wind, unsymmetrical rotorcraft loading, water wave action, rotorcraft inertia, and probable structural damage and leakage considered under § 27.801(d). Maximum roll and pitch angles determined from compliance with § 27.801(d) may be used, if significant, to determine the extent of immersion of each float. If the floats are deployed in flight, appropriate air loads derived from the flight limitations with the floats deployed shall be used in substantiation of the floats and their attachment to the rotorcraft. For this purpose, the desig… | |||
| 14:14:1.0.1.3.15.3.173.36 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.571 Fatigue evaluation of flight structure. | FAA | [Amdt. 27-3, 33 FR 14106, Sept. 18, 1968, as amended by Amdt. 27-12, 42 FR 15044, Mar. 17, 1977; Amdt. 27-18, 45 FR 60177, Sept. 11, 1980; Amdt. 27-26, 55 FR 8000, Mar. 6, 1990] | (a) General. Each portion of the flight structure (the flight structure includes rotors, rotor drive systems between the engines and the rotor hubs, controls, fuselage, landing gear, and their related primary attachments), the failure of which could be catastrophic, must be identified and must be evaluated under paragraph (b), (c), (d), or (e) of this section. The following apply to each fatigue evaluation: (1) The procedure for the evaluation must be approved. (2) The locations of probable failure must be determined. (3) Inflight measurement must be included in determining the following: (i) Loads or stresses in all critical conditions throughout the range of limitations in § 27.309, except that maneuvering load factors need not exceed the maximum values expected in operation. (ii) The effect of altitude upon these loads or stresses. (4) The loading spectra must be as severe as those expected in operation including, but not limited to, external cargo operations, if applicable, and ground-air-ground cycles. The loading spectra must be based on loads or stresses determined under paragraph (a)(3) of this section. (b) Fatigue tolerance evaluation. It must be shown that the fatigue tolerance of the structure ensures that the probability of catastrophic fatigue failure is extremely remote without establishing replacement times, inspection intervals or other procedures under section A27.4 of appendix A. (c) Replacement time evaluation. it must be shown that the probability of catastrophic fatigue failure is extremely remote within a replacement time furnished under section A27.4 of appendix A. (d) Fail-safe evaluation. The following apply to fail-safe evaluation: (1) It must be shown that all partial failures will become readily detectable under inspection procedures furnished under section A27.4 of appendix A. (2) The interval between the time when any partial failure becomes readily detectable under paragraph (d)(1) of this section, and the time when any such failure is expected to reduce the remai… | |||
| 14:14:1.0.1.3.15.3.173.37 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | C | Subpart C—Strength Requirements | § 27.573 Damage Tolerance and Fatigue Evaluation of Composite Rotorcraft Structures. | FAA | [Doc. No. FAA-2009-0660, Amdt. 27-47, 76 FR 74663, Dec. 1, 2011] | (a) Each applicant must evaluate the composite rotorcraft structure under the damage tolerance standards of paragraph (d) of this section unless the applicant establishes that a damage tolerance evaluation is impractical within the limits of geometry, inspectability, and good design practice. If an applicant establishes that it is impractical within the limits of geometry, inspectability, and good design practice, the applicant must do a fatigue evaluation in accordance with paragraph (e) of this section. (b) The methodology used to establish compliance with this section must be submitted to and approved by the Administrator. (c) Definitions: (1) Catastrophic failure is an event that could prevent continued safe flight and landing. (2) Principal Structural Elements (PSEs) are structural elements that contribute significantly to the carrying of flight or ground loads, the failure of which could result in catastrophic failure of the rotorcraft. (3) Threat Assessment is an assessment that specifies the locations, types, and sizes of damage, considering fatigue, environmental effects, intrinsic and discrete flaws, and impact or other accidental damage (including the discrete source of the accidental damage) that may occur during manufacture or operation. (d) Damage Tolerance Evaluation: (1) Each applicant must show that catastrophic failure due to static and fatigue loads, considering the intrinsic or discrete manufacturing defects or accidental damage, is avoided throughout the operational life or prescribed inspection intervals of the rotorcraft by performing damage tolerance evaluations of the strength of composite PSEs and other parts, detail design points, and fabrication techniques. Each applicant must account for the effects of material and process variability along with environmental conditions in the strength and fatigue evaluations. Each applicant must evaluate parts that include PSEs of the airframe, main and tail rotor drive systems, main and tail rotor blades and hubs, rotor controls, fixed … | |||
| 14:14:1.0.1.3.15.4.174.1 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.601 Design. | FAA | (a) The rotorcraft may have no design features or details that experience has shown to be hazardous or unreliable. (b) The suitability of each questionable design detail and part must be established by tests. | ||||
| 14:14:1.0.1.3.15.4.174.10 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.619 Special factors. | FAA | (a) The special factors prescribed in §§ 27.621 through 27.625 apply to each part of the structure whose strength is— (1) Uncertain; (2) Likely to deteriorate in service before normal replacement; or (3) Subject to appreciable variability due to— (i) Uncertainties in manufacturing processes; or (ii) Uncertainties in inspection methods. (b) For each part to which §§ 27.621 through 27.625 apply, the factor of safety prescribed in § 27.303 must be multiplied by a special factor equal to— (1) The applicable special factors prescribed in §§ 27.621 through 27.625; or (2) Any other factor great enough to ensure that the probability of the part being understrength because of the uncertainties specified in paragraph (a) of this section is extremely remote. | ||||
| 14:14:1.0.1.3.15.4.174.11 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.621 Casting factors. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-34, 62 FR 46173, Aug. 29, 1997] | (a) General. The factors, tests, and inspections specified in paragraphs (b) and (c) of this section must be applied in addition to those necessary to establish foundry quality control. The inspections must meet approved specifications. Paragraphs (c) and (d) of this section apply to structural castings except castings that are pressure tested as parts of hydraulic or other fluid systems and do not support structural loads. (b) Bearing stresses and surfaces. The casting factors specified in paragraphs (c) and (d) of this section— (1) Need not exceed 1.25 with respect to bearing stresses regardless of the method of inspection used; and (2) Need not be used with respect to the bearing surfaces of a part whose bearing factor is larger than the applicable casting factor. (c) Critical castings. For each casting whose failure would preclude continued safe flight and landing of the rotorcraft or result in serious injury to any occupant, the following apply: (1) Each critical casting must— (i) Have a casting factor of not less than 1.25; and (ii) Receive 100 percent inspection by visual, radiographic, and magnetic particle (for ferromagnetic materials) or penetrant (for nonferromagnetic materials) inspection methods or approved equivalent inspection methods. (2) For each critical casting with a casting factor less than 1.50, three sample castings must be static tested and shown to meet— (i) The strength requirements of § 27.305 at an ultimate load corresponding to a casting factor of 1.25; and (ii) The deformation requirements of § 27.305 at a load of 1.15 times the limit load. (d) Noncritical castings. For each casting other than those specified in paragraph (c) of this section, the following apply: (1) Except as provided in paragraphs (d)(2) and (3) of this section, the casting factors and corresponding inspections must meet the following table: (2) The percentage of castings inspected by nonvisual methods may be reduced below that specified in paragraph (d)(1) of this section when an approved qual… | |||
| 14:14:1.0.1.3.15.4.174.12 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.623 Bearing factors. | FAA | (a) Except as provided in paragraph (b) of this section, each part that has clearance (free fit), and that is subject to pounding or vibration, must have a bearing factor large enough to provide for the effects of normal relative motion. (b) No bearing factor need be used on a part for which any larger special factor is prescribed. | ||||
| 14:14:1.0.1.3.15.4.174.13 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.625 Fitting factors. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-35, 63 FR 43285, Aug. 12, 1998] | For each fitting (part or terminal used to join one structural member to another) the following apply: (a) For each fitting whose strength is not proven by limit and ultimate load tests in which actual stress conditions are simulated in the fitting and surrounding structures, a fitting factor of at least 1.15 must be applied to each part of— (1) The fitting; (2) The means of attachment; and (3) The bearing on the joined members. (b) No fitting factor need be used— (1) For joints made under approved practices and based on comprehensive test data (such as continuous joints in metal plating, welded joints, and scarf joints in wood); and (2) With respect to any bearing surface for which a larger special factor is used. (c) For each integral fitting, the part must be treated as a fitting up to the point at which the section properties become typical of the member. (d) Each seat, berth, litter, safety belt, and harness attachment to the structure must be shown by analysis, tests, or both, to be able to withstand the inertia forces prescribed in § 27.561(b)(3) multiplied by a fitting factor of 1.33. | |||
| 14:14:1.0.1.3.15.4.174.14 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.629 Flutter. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-26, 55 FR 8000, Mar. 6, 1990] | Each aerodynamic surface of the rotorcraft must be free from flutter under each appropriate speed and power condition. | |||
| 14:14:1.0.1.3.15.4.174.2 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.602 Critical parts. | FAA | [Doc. No. 29311, 64 FR 46232, Aug. 24, 1999] | (a) Critical part. A critical part is a part, the failure of which could have a catastrophic effect upon the rotocraft, and for which critical characteristics have been identified which must be controlled to ensure the required level of integrity. (b) If the type design includes critical parts, a critical parts list shall be established. Procedures shall be established to define the critical design characteristics, identify processes that affect those characteristics, and identify the design change and process change controls necessary for showing compliance with the quality assurance requirements of part 21 of this chapter. | |||
| 14:14:1.0.1.3.15.4.174.3 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.603 Materials. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR 55469, Dec. 20, 1976; Amdt. 27-16, 43 FR 50599, Oct. 30, 1978] | The suitability and durability of materials used for parts, the failure of which could adversely affect safety, must— (a) Be established on the basis of experience or tests; (b) Meet approved specifications that ensure their having the strength and other properties assumed in the design data; and (c) Take into account the effects of environmental conditions, such as temperature and humidity, expected in service. | |||
| 14:14:1.0.1.3.15.4.174.4 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.605 Fabrication methods. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 43 FR 50599, Oct. 30, 1978] | (a) The methods of fabrication used must produce consistently sound structures. If a fabrication process (such as gluing, spot welding, or heat-treating) requires close control to reach this objective, the process must be performed according to an approved process specification. (b) Each new aircraft fabrication method must be substantiated by a test program. | |||
| 14:14:1.0.1.3.15.4.174.5 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.607 Fasteners. | FAA | [Amdt. 27-4, 33 FR 14533, Sept. 27, 1968] | (a) Each removable bolt, screw, nut, pin, or other fastener whose loss could jeopardize the safe operation of the rotorcraft must incorporate two separate locking devices. The fastener and its locking devices may not be adversely affected by the environmental conditions associated with the particular installation. (b) No self-locking nut may be used on any bolt subject to rotation in operation unless a nonfriction locking device is used in addition to the self-locking device. | |||
| 14:14:1.0.1.3.15.4.174.6 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.609 Protection of structure. | FAA | Each part of the structure must— (a) Be suitably protected against deterioration or loss of strength in service due to any cause, including— (1) Weathering; (2) Corrosion; and (3) Abrasion; and (b) Have provisions for ventilation and drainage where necessary to prevent the accumulation of corrosive, flammable, or noxious fluids. | ||||
| 14:14:1.0.1.3.15.4.174.7 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.610 Lightning and static electricity protection. | FAA | [Amdt. 27-21, 49 FR 44433, Nov. 6, 1984, as amended by Amdt. 27-37, 64 FR 45094, Aug. 18, 1999; Amdt. 27-46, 76 FR 33135, June 8, 2011] | (a) The rotorcraft must be protected against catastrophic effects from lightning. (b) For metallic components, compliance with paragraph (a) of this section may be shown by— (1) Electrically bonding the components properly to the airframe; or (2) Designing the components so that a strike will not endanger the rotorcraft. (c) For nonmetallic components, compliance with paragraph (a) of this section may be shown by— (1) Designing the components to minimize the effect of a strike; or (2) Incorporating acceptable means of diverting the resulting electrical current so as not to endanger the rotorcraft. (d) The electrical bonding and protection against lightning and static electricity must— (1) Minimize the accumulation of electrostatic charge; (2) Minimize the risk of electric shock to crew, passengers, and service and maintenance personnel using normal precautions; (3) Provide an electrical return path, under both normal and fault conditions, on rotorcraft having grounded electrical systems; and (4) Reduce to an acceptable level the effects of static electricity on the functioning of essential electrical and electronic equipment. | |||
| 14:14:1.0.1.3.15.4.174.8 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.611 Inspection provisions. | FAA | There must be means to allow the close examination of each part that requires— (a) Recurring inspection; (b) Adjustment for proper alignment and functioning; or (c) Lubrication. | ||||
| 14:14:1.0.1.3.15.4.174.9 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.613 Material strength properties and design values. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 43 FR 50599, Oct. 30, 1978; Amdt. 27-26, 55 FR 8000, Mar. 6, 1990] | (a) Material strength properties must be based on enough tests of material meeting specifications to establish design values on a statistical basis. (b) Design values must be chosen to minimize the probability of structural failure due to material variability. Except as provided in paragraphs (d) and (e) of this section, compliance with this paragraph must be shown by selecting design values that assure material strength with the following probability— (1) Where applied loads are eventually distributed through a single member within an assembly, the failure of which would result in loss of structural integrity of the component, 99 percent probability with 95 percent confidence; and (2) For redundant structure, those in which the failure of individual elements would result in applied loads being safely distributed to other load-carrying members, 90 percent probability with 95 percent confidence. (c) The strength, detail design, and fabrication of the structure must minimize the probability of disastrous fatigue failure, particularly at points of stress concentration. (d) Design values may be those contained in the following publications (available from the Naval Publications and Forms Center, 5801 Tabor Avenue, Philadelphia, Pennsylvania 19120) or other values approved by the Administrator: (1) MIL-HDBK-5, “Metallic Materials and Elements for Flight Vehicle Structure”. (2) MIL-HDBK-17, “Plastics for Flight Vehicles”. (3) ANC-18, “Design of Wood Aircraft Structures”. (4) MIL-HDBK-23, “Composite Construction for Flight Vehicles”. (e) Other design values may be used if a selection of the material is made in which a specimen of each individual item is tested before use and it is determined that the actual strength properties of that particular item will equal or exceed those used in design. | |||
| 14:14:1.0.1.3.15.4.175.15 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.653 Pressure venting and drainage of rotor blades. | FAA | [Amdt. 27-2, 33 FR 963, Jan. 26, 1968] | (a) For each rotor blade— (1) There must be means for venting the internal pressure of the blade; (2) Drainage holes must be provided for the blade; and (3) The blade must be designed to prevent water from becoming trapped in it. (b) Paragraphs (a)(1) and (2) of this section does not apply to sealed rotor blades capable of withstanding the maximum pressure differentials expected in service. | |||
| 14:14:1.0.1.3.15.4.175.16 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.659 Mass balance. | FAA | [Amdt. 27-2, 33 FR 963, Jan. 26, 1968] | (a) The rotors and blades must be mass balanced as necessary to— (1) Prevent excessive vibration; and (2) Prevent flutter at any speed up to the maximum forward speed. (b) The structural integrity of the mass balance installation must be substantiated. | |||
| 14:14:1.0.1.3.15.4.175.17 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.661 Rotor blade clearance. | FAA | [Amdt. 27-2, 33 FR 963, Jan. 26, 1968] | There must be enough clearance between the rotor blades and other parts of the structure to prevent the blades from striking any part of the structure during any operating condition. | |||
| 14:14:1.0.1.3.15.4.175.18 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.663 Ground resonance prevention means. | FAA | [Amdt. 27-2, 33 FR 963, Jan. 26, 1968, as amended by Amdt. 27-26, 55 FR 8000, Mar. 6, 1990] | (a) The reliability of the means for preventing ground resonance must be shown either by analysis and tests, or reliable service experience, or by showing through analysis or tests that malfunction or failure of a single means will not cause ground resonance. (b) The probable range of variations, during service, of the damping action of the ground resonance prevention means must be established and must be investigated during the test required by § 27.241. | |||
| 14:14:1.0.1.3.15.4.176.19 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.671 General. | FAA | (a) Each control and control system must operate with the ease, smoothness, and positiveness appropriate to its function. (b) Each element of each flight control system must be designed, or distinctively and permanently marked, to minimize the probability of any incorrect assembly that could result in the malfunction of the system. | ||||
| 14:14:1.0.1.3.15.4.176.20 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.672 Stability augmentation, automatic, and power-operated systems. | FAA | [Amdt. 27-21, 49 FR 44433, Nov. 6, 1984; 49 FR 47594, Dec. 6, 1984] | If the functioning of stability augmentation or other automatic or power-operated systems is necessary to show compliance with the flight characteristics requirements of this part, such systems must comply with § 27.671 of this part and the following: (a) A warning which is clearly distinguishable to the pilot under expected flight conditions without requiring the pilot's attention must be provided for any failure in the stability augmentation system or in any other automatic or power-operated system which could result in an unsafe condition if the pilot is unaware of the failure. Warning systems must not activate the control systems. (b) The design of the stability augmentation system or of any other automatic or power-operated system must allow initial counteraction of failures without requiring exceptional pilot skill or strength by overriding the failure by movement of the flight controls in the normal sense and deactivating the failed system. (c) It must be shown that after any single failure of the stability augmentation system or any other automatic or power-operated system— (1) The rotorcraft is safely controllable when the failure or malfunction occurs at any speed or altitude within the approved operating limitations; (2) The controllability and maneuverability requirements of this part are met within a practical operational flight envelope (for example, speed, altitude, normal acceleration, and rotorcraft configurations) which is described in the Rotorcraft Flight Manual; and (3) The trim and stability characteristics are not impaired below a level needed to permit continued safe flight and landing. | |||
| 14:14:1.0.1.3.15.4.176.21 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.673 Primary flight control. | FAA | [Amdt. 27-21, 49 FR 44434, Nov. 6, 1984] | Primary flight controls are those used by the pilot for immediate control of pitch, roll, yaw, and vertical motion of the rotorcraft. | |||
| 14:14:1.0.1.3.15.4.176.22 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.674 Interconnected controls. | FAA | [Amdt. 27-26, 55 FR 8001, Mar. 6, 1990] | Each primary flight control system must provide for safe flight and landing and operate independently after a malfunction, failure, or jam of any auxiliary interconnected control. | |||
| 14:14:1.0.1.3.15.4.176.23 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.675 Stops. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 43 FR 50599, Oct. 30, 1978] | (a) Each control system must have stops that positively limit the range of motion of the pilot's controls. (b) Each stop must be located in the system so that the range of travel of its control is not appreciably affected by— (1) Wear; (2) Slackness; or (3) Takeup adjustments. (c) Each stop must be able to withstand the loads corresponding to the design conditions for the system. (d) For each main rotor blade— (1) Stops that are appropriate to the blade design must be provided to limit travel of the blade about its hinge points; and (2) There must be means to keep the blade from hitting the droop stops during any operation other than starting and stopping the rotor. | |||
| 14:14:1.0.1.3.15.4.176.24 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.679 Control system locks. | FAA | If there is a device to lock the control system with the rotorcraft on the ground or water, there must be means to— (a) Give unmistakable warning to the pilot when the lock is engaged; and (b) Prevent the lock from engaging in flight. | ||||
| 14:14:1.0.1.3.15.4.176.25 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.681 Limit load static tests. | FAA | (a) Compliance with the limit load requirements of this part must be shown by tests in which— (1) The direction of the test loads produces the most severe loading in the control system; and (2) Each fitting, pulley, and bracket used in attaching the system to the main structure is included. (b) Compliance must be shown (by analyses or individual load tests) with the special factor requirements for control system joints subject to angular motion. | ||||
| 14:14:1.0.1.3.15.4.176.26 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.683 Operation tests. | FAA | It must be shown by operation tests that, when the controls are operated from the pilot compartment with the control system loaded to correspond with loads specified for the system, the system is free from— (a) Jamming; (b) Excessive friction; and (c) Excessive deflection. | ||||
| 14:14:1.0.1.3.15.4.176.27 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.685 Control system details. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR 55469, Dec. 20, 1976; Amdt. 27-26, 55 FR 8001, Mar. 6, 1990; 69 FR 18803, Apr. 9, 2004; Doc. No. FAA-2018-0119, Amdt. 27-49, 83 FR 9170, Mar. 5, 2018] | (a) Each detail of each control system must be designed to prevent jamming, chafing, and interference from cargo, passengers, loose objects or the freezing of moisture. (b) There must be means in the cockpit to prevent the entry of foreign objects into places where they would jam the system. (c) There must be means to prevent the slapping of cables or tubes against other parts. (d) Cable systems must be designed as follows: (1) Cables, cable fittings, turnbuckles, splices, and pulleys must be of an acceptable kind. (2) The design of the cable systems must prevent any hazardous change in cable tension throughout the range of travel under any operating conditions and temperature variations. (3) No cable smaller than three thirty-seconds of an inch diameter may be used in any primary control system. (4) Pulley kinds and sizes must correspond to the cables with which they are used. The pulley cable combinations and strength values which must be used are specified in Military Handbook MIL-HDBK-5C, Vol. 1 & Vol. 2, Metallic Materials and Elements for Flight Vehicle Structures, (Sept. 15, 1976, as amended through December 15, 1978). This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. section 552(a) and 1 CFR part 51. Copies may be obtained from the Naval Publications and Forms Center, 5801 Tabor Avenue, Philadelphia, Pennsylvania, 19120. Copies may be inspected at the National Archives and Records Administration (NARA). For information on the availability of this material at NARA, call 202-741-6030, or go to: http://www.archives.gov/federal-register/cfr/ibr-locations.html (5) Pulleys must have close fitting guards to prevent the cables from being displaced or fouled. (6) Pulleys must lie close enough to the plane passing through the cable to prevent the cable from rubbing against the pulley flange. (7) No fairlead may cause a change in cable direction of more than 3°. (8) No clevis pin subject to load or motion and retained only by cotter pins may … | |||
| 14:14:1.0.1.3.15.4.176.28 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.687 Spring devices. | FAA | (a) Each control system spring device whose failure could cause flutter or other unsafe characteristics must be reliable. (b) Compliance with paragraph (a) of this section must be shown by tests simulating service conditions. | ||||
| 14:14:1.0.1.3.15.4.176.29 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.691 Autorotation control mechanism. | FAA | Each main rotor blade pitch control mechanism must allow rapid entry into autorotation after power failure. | ||||
| 14:14:1.0.1.3.15.4.176.30 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.695 Power boost and power-operated control system. | FAA | (a) If a power boost or power-operated control system is used, an alternate system must be immediately available that allows continued safe flight and landing in the event of— (1) Any single failure in the power portion of the system; or (2) The failure of all engines. (b) Each alternate system may be a duplicate power portion or a manually operated mechanical system. The power portion includes the power source (such as hydraulic pumps), and such items as valves, lines, and actuators. (c) The failure of mechanical parts (such as piston rods and links), and the jamming of power cylinders, must be considered unless they are extremely improbable. | ||||
| 14:14:1.0.1.3.15.4.177.31 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.723 Shock absorption tests. | FAA | The landing inertia load factor and the reserve energy absorption capacity of the landing gear must be substantiated by the tests prescribed in §§ 27.725 and 27.727, respectively. These tests must be conducted on the complete rotorcraft or on units consisting of wheel, tire, and shock absorber in their proper relation. | ||||
| 14:14:1.0.1.3.15.4.177.32 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.725 Limit drop test. | FAA | The limit drop test must be conducted as follows: (a) The drop height must be— (1) 13 inches from the lowest point of the landing gear to the ground; or (2) Any lesser height, not less than eight inches, resulting in a drop contact velocity equal to the greatest probable sinking speed likely to occur at ground contact in normal power-off landings. (b) If considered, the rotor lift specified in § 27.473(a) must be introduced into the drop test by appropriate energy absorbing devices or by the use of an effective mass. (c) Each landing gear unit must be tested in the attitude simulating the landing condition that is most critical from the standpoint of the energy to be absorbed by it. (d) When an effective mass is used in showing compliance with paragraph (b) of this section, the following formula may be used instead of more rational computations: where: W e = the effective weight to be used in the drop test (lbs.); W = W M for main gear units (lbs.), equal to the static reaction on the particular unit with the rotorcraft in the most critical attitude. A rational method may be used in computing a main gear static reaction, taking into consideration the moment arm between the main wheel reaction and the rotorcraft center of gravity. W = W N for nose gear units (lbs.), equal to the vertical component of the static reaction that would exist at the nose wheel, assuming that the mass of the rotorcraft acts at the center of gravity and exerts a force of 1.0 g downward and 0.25 g forward. W = W T for tailwheel units (lbs.), equal to whichever of the following is critical: (1) The static weight on the tailwheel with the rotorcraft resting on all wheels; or (2) The vertical component of the ground reaction that would occur at the tailwheel, assuming that the mass of the rotorcraft acts at the center of gravity and exerts a force of l g downward with the rotorcraft in the maximum nose-up attitude considered in the nose-up landing conditions. h = specified free drop height (inch… | ||||
| 14:14:1.0.1.3.15.4.177.33 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.727 Reserve energy absorption drop test. | FAA | [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-26, 55 FR 8001, Mar. 6, 1990] | The reserve energy absorption drop test must be conducted as follows: (a) The drop height must be 1.5 times that specified in § 27.725(a). (b) Rotor lift, where considered in a manner similar to that prescribed in § 27.725(b), may not exceed 1.5 times the lift allowed under that paragraph. (c) The landing gear must withstand this test without collapsing. Collapse of the landing gear occurs when a member of the nose, tail, or main gear will not support the rotorcraft in the proper attitude or allows the rotorcraft structure, other than the landing gear and external accessories, to impact the landing surface. | |||
| 14:14:1.0.1.3.15.4.177.34 | 14 | Aeronautics and Space | I | C | 27 | PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT | D | Subpart D—Design and Construction | § 27.729 Retracting mechanism. | FAA | [Amdt. 27-21, 49 FR 44434, Nov. 6, 1984] | For rotorcraft with retractable landing gear, the following apply: (a) Loads. The landing gear, retracting mechansim, wheel-well doors, and supporting structure must be designed for— (1) The loads occurring in any maneuvering condition with the gear retracted; (2) The combined friction, inertia, and air loads occurring during retraction and extension at any airspeed up to the design maximum landing gear operating speed; and (3) The flight loads, including those in yawed flight, occurring with the gear extended at any airspeed up to the design maximum landing gear extended speed. (b) Landing gear lock. A positive means must be provided to keep the gear extended. (c) Emergency operation. When other than manual power is used to operate the gear, emergency means must be provided for extending the gear in the event of— (1) Any reasonably probable failure in the normal retraction system; or (2) The failure of any single source of hydraulic, electric, or equivalent energy. (d) Operation tests. The proper functioning of the retracting mechanism must be shown by operation tests. (e) Position indicator. There must be a means to indicate to the pilot when the gear is secured in the extreme positions. (f) Control. The location and operation of the retraction control must meet the requirements of §§ 27.777 and 27.779. (g) Landing gear warning. An aural or equally effective landing gear warning device must be provided that functions continuously when the rotorcraft is in a normal landing mode and the landing gear is not fully extended and locked. A manual shutoff capability must be provided for the warning device and the warning system must automatically reset when the rotorcraft is no longer in the landing mode. |
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title_number INTEGER,
title_name TEXT,
chapter TEXT,
subchapter TEXT,
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