14 C.F.R. Subpart B—Flight
Title 14 - Aeronautics and Space
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. [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–21, 49 FR 44432, Nov. 6, 1984] (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 (the highest weight at which compliance with each applicable structural loading condition of this part is shown); or (iii) The highest weight at which compliance with each applicable flight requirement of this part is shown; 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 established under paragraph (a) of this section may be established for any rotorcraft-load combination if— (1) The rotorcraft-load combination does not include human external cargo, (2) Structural component approval for external load operations under either §27.865 or under equivalent operational standards is obtained, (3) The portion of the total weight that is greater than the maximum weight established under paragraph (a) of this section is made up only of the weight of all or part of the jettisonable external load, (4) Structural components of the rotorcraft are shown to comply with the applicable structural requirements of this part under the increased loads and stresses caused by the weight increase over that established under paragraph (a) of this section, and (5) Operation of the rotorcraft at a total weight greater than the maximum certificated weight established under paragraph (a) of this section is limited by appropriate operating limitations under §27.865(a) and (d) of this part. (Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. 1655(c))) [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] 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. [Amdt. 27–2, 33 FR 962, Jan. 26, 1968] (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. (Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. 1655(c))) [Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–14, 43 FR 2324, Jan. 16, 1978] Removable ballast may be used in showing compliance with the flight requirements of this subpart. (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 pilot in all flight conditions, including power-on and power-off flight, when the speed of a main rotor approaches a value that can jeopardize safe flight. (2) The warning may be furnished either through the inherent aerodynamic qualities of the helicopter or by a device. (3) The warning must be clear and distinct under all conditons, and must be clearly distinguishable from all other warnings. A visual device that requires the attention of the crew within the cockpit is not acceptable by itself. (4) If a warning device is used, the device must automatically deactivate and reset when the low-speed condition is corrected. If the device has an audible warning, it must also be equipped with a means for the pilot to manually silence the audible warning before the low-speed condition is corrected. (Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. 1655(c))) [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]
Title 14: Aeronautics and Space
PART 27—AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT
Subpart B—Flight
General
§ 27.21 Proof of compliance.
§ 27.25 Weight limits.
§ 27.27 Center of gravity limits.
§ 27.29 Empty weight and corresponding center of gravity.
§ 27.31 Removable ballast.
§ 27.33 Main rotor speed and pitch limits.
Performance
§ 27.45 General.
(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 atmopheric 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.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 27–14, 43 FR 2324, Jan. 16, 1978, as amended by Amdt. 27–21, 49 FR 44432, Nov. 6, 1984]
§ 27.51 Takeoff.
(a) The takeoff, with takeoff power and r.p.m., and with the extreme forward center of gravity—
(1) May not require exceptional piloting skill or exceptionally favorable conditions; and
(2) 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.
(b) Paragraph (a) of this section must be met throughout the ranges of—
(1) Altitude, from standard sea level conditions to the maximum altitude capability of the rotorcraft, or 7,000 feet, whichever is less; and
(2) Weight, from the maximum weight (at sea level) to each lesser weight selected by the applicant for each altitude covered by paragraph (b)(1) of this section.
§ 27.65 Climb: all engines operating.
(a) For rotorcraft other than helicopters—
(1) The steady rate of climb, at V
(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
(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
(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.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. 1655(c)))
[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]
§ 27.67 Climb: one engine inoperative.
For multiengine helicopters, the steady rate of climb (or descent), at V
(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.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. 1655(c)))
[Doc. No 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–23, 53 FR 34210, Sept. 2, 1988]
§ 27.71 Glide performance.
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.
[Amdt. 27–21, 49 FR 44433, Nov. 6, 1984]
§ 27.73 Performance at minimum operating speed.
(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; or
(ii) For turbine engine powered helicopters, 2,500 feet pressure altitude at maximum weight at a temperature of standard +40 degrees F.
(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.
§ 27.75 Landing.
(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 glide 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
(ii) For multiengine rotocraft, one engine inoperative and with each operating engine within approved operating limitations; and
(3) The approach and landing entered from steady autorotation.
(b) Multiengine rotorcraft must be able to be landed safely after complete power failure under normal operating conditions.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–14, 43 FR 2324, Jan. 16, 1978]
§ 27.79 Limiting height—speed envelope.
(a) If there is any combination of height and forward speed (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-speed 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 7,000 feet, whichever is less; and
(2) Weight, from the maximum weight (at sea level) to the lesser 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 which is lower.
(b) The applicable power failure conditions are—
(1) For single-engine helicopters, full autorotation;
(2) For multiengine helicopters, one engine inoperative (where engine isolation features insure continued operation of the remaining engines), and the remaining engines at the greatest power for which certification is requested, and
(3) For other rotocraft, conditions appropriate to the type.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C. 1655(c)))
[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]
Flight Characteristics
§ 27.141 General.
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.
[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]
§ 27.143 Controllability and maneuverability.
(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) Glide;
(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
(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 (e) of this section) and power on.
(c) A wind velocity of not less than 17 knots 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), with—
(1) Critical weight;
(2) Critical center of gravity;
(3) Critical rotor r.p.m.; and
(4) Altitude, from standard sea level conditions to the maximum altitude capability of the rotorcraft or 7,000 feet, whichever is less.
(d) 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, must be controllable over the range of speeds and altitudes for which certification is requested when such power failure occurs with maximum continuous power and critical weight. No corrective action time delay for any condition following power failure may be less than—
(i) For the cruise condition, one second, or normal pilot reaction time (whichever is greater); and
(ii) For any other condition, normal pilot reaction time.
(e) For helicopters for which a V
(1) The helicopter must be safely slowed to V
(2) At a speed of 1.1 V
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. 1655(c)))
[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]
§ 27.151 Flight controls.
(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.
[Amdt. 27–21, 49 FR 44433, Nov. 6, 1984]
§ 27.161 Trim control.
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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27–21, 49 FR 44433, Nov. 6, 1984]
§ 27.171 Stability: general.
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.
§ 27.173 Static longitudinal stability.
(a) The longitudinal control must be designed so that a rearward movement of the control is necessary to obtain a speed less than the trim speed, and a forward movement of the control is necessary to obtain a speed more than the trim speed.
(b) With the throttle and collective pitch held constant during the maneuvers specified in §27.175 (a) through (c), the slope of the control position versus speed curve must be positive throughout the full range of altitude for which certification is requested.
(c) During the maneuver specified in §27.175(d), the longitudinal control position versus speed curve may have a negative slope within the specified speed range if the negative motion is not greater than 10 percent of total control travel.
[Amdt. 27–21, 49 FR 44433, Nov. 6, 1984]
§ 27.175 Demonstration of static longitudinal stability.
(a) Climb. Static longitudinal stability must be shown in the climb condition at speeds from 0.85 V
(1) Critical weight;
(2) Critical center of gravity;
(3) Maximum continuous power;
(4) The landing gear retracted; and
(5) The rotorcraft trimmed at V
(b) Cruise. Static longitudinal stability must be shown in the cruise condition at speeds from 0.7 V
(1) Critical weight;
(2) Critical center of gravity;
(3) Power for level flight at 0.9 V
(4) The landing gear retracted; and
(5) The rotorcraft trimmed at 0.9 V
(c) Autorotation. Static longitudinal stability must be shown in autorotation at airspeeds from 0.5 times the speed for minimum rate of descent to V
(1) Critical weight;
(2) Critical center of gravity;
(3) Power off;
(4) The landing gear—
(i) Retracted; and
(ii) Extended; and
(5) The rotorcraft trimmed at appropriate speeds found necessary by the Administrator to demonstrate stability throughout the prescribed speed range.
(d) Hovering. For helicopters, the longitudinal cyclic control must operate with the sense and direction of motion prescribed in §27.173 between the maximum approved rearward speed and a forward speed of 17 knots with—
(1) Critical weight;
(2) Critical center of gravity;
(3) Power required to maintain an approximate constant height in ground effect;
(4) The landing gear extended; and
(5) The helicopter trimmed for hovering.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of Transportation Act (49 U.S.C. 1655(c)))
[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]
§ 27.177 Static directional stability.
Static directional stability must be positive with throttle and collective controls held constant at the trim conditions specified in §27.175 (a) and (b). This must be shown by steadily increasing directional control deflection for sideslip angles up to ±10° from trim. Sufficient cues must accompany sideslip to alert the pilot when approaching sideslip limits.
[Amdt. 27–21, 49 FR 44433, Nov. 6, 1984]
Ground and Water Handling Characteristics
§ 27.231 General.
The rotorcraft must have satisfactory ground and water handling characteristics, including freedom from uncontrollable tendencies in any condition expected in operation.
§ 27.235 Taxiing condition.
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.
§ 27.239 Spray characteristics.
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.
§ 27.241 Ground resonance.
The rotorcraft may have no dangerous tendency to oscillate on the ground with the rotor turning.
Miscellaneous Flight Requirements
§ 27.251 Vibration.
Each part of the rotorcraft must be free from excessive vibration under each appropriate speed and power condition.
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