Part V - Airworthiness Manual Chapter 527 - Normal Category Aircraft

Canadian Aviation Regulations (CARs) 2017-2

Content last revised: 2009/12/01

Preamble

SUBCHAPTERS

  • A (527.1-527.2), 
  • B (527.21-527.251), 
  • C (527.301-527.571), 
  • D (527.601-527.873), 
  • E (527.901-527.1195), 
  • F (527.1301-527.1461), 
  • G (527.1501-527.1589)

APPENDICES

ABCD

(2002/06/01; no previous version)

SUBCHAPTER A GENERAL

527.1 Applicability

  1. (a) This Chapter sets out airworthiness standards for the issue of type certificates and changes to those type certificates, for normal category rotorcraft with maximum mass (weights) of 3,175 kg (7,000 lbs.) or less and nine or less passenger seats.
    (amended 2009/12/01)
  2. (b) Reserved.
    (amended 2009/12/01)
  3. (c) Multi-engine rotorcraft may be type certificated as Category A provided the requirements referenced in Appendix C of this chapter are met.

    (Change 527-2 (92-02-01))
    (Change 527-4)

527.2 Special Retroactive Requirements

  1. [(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. [(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. [(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. [(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. [(4) For purposes of this section, the date of manufacture is either:
      1. [(i) The date the statement of conformity or equivalent inspection acceptance records, reflects that the rotorcraft is complete and meets the type design data approved by the Minister; or
      2. [(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.
  2. (b) [For rotorcraft with a certification basis established prior to November 23, 1999:
    1. [(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 chapter in effect on November 23, 1999.
    2. [(2) The maximum weight may be increased to greater than 2,720 kg (6,000 lbs.) provided:
      1. [(i) The number of passenger seats is not increased above the maximum number certificated on November 23, 1999, or
      2. [(ii) The applicant shows compliance with all of the airworthiness requirements of this chapter in effect on November 23, 1999.]

Information Note:

The underlined effective dates above are different from the effective dates stated in the same requirements of FAR section 27.2(b). The change in the requirement of section 527.2(b) above is FAR amendment 27-37. Consequently, the difference in dates exists because FAR amendment 27-37 is not effective in Canada until November 23, 1999 as per NPA 1999-168.

(Change 527-2 (92-02-01))

(Change 527-4)

SUBCHAPTER B FLIGHT- GENERAL

527.21 Proof of Compliance

Each requirement of this subchapter must be met at each appropriate combination of weight and centre of gravity within the range of loading conditions for which certification is requested. This must be shown:

  1. (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
  2. (b) By systematic investigation of each required combination of weight and centre of gravity if compliance cannot be reasonably inferred from combinations investigated.

527.25 Weight Limits

  1. (a) Maximum weight. The maximum weight (the highest weight at which compliance with each applicable requirement of this chapter is shown) must be established so that it is:
    1. (1) Not more than:
      1. (i) The highest weight selected by the applicant;
      2. (ii) The design maximum weight (the highest weight at which compliance with each applicable structural loading condition of this chapter is shown);
      3. (iii) The highest weight at which compliance with each applicable flight requirement of this chapter is shown; or
        (amended 2009/05/11)
      4. (iv) The highest weight in which the provisions of 527.87 or 527.143(c)(1), or combinations thereof, are demonstrated if the weights and operating conditions (altitude and temperature) prescribed by those requirements cannot be met; and
        (amended 2009/05/11; no previous version)
    2. (2) Not less than the sum of:
      1. (i) The empty weight determined under 527.29;
      2. (ii) The weight of usable fuel appropriate to the intended operation with full payload;
      3. (iii) The weight of full oil capacity; and
      4. (iv) For each seat, an occupant weight of 170 pounds or any lower weight for which certification is requested.
  2. (b) Minimum weight. The minimum weight (the lowest weight at which compliance with each applicable requirement of this chapter is shown) must be established so that it is:
    1. (1) Not more than the sum of:
      1. (i) The empty weight determined under 527.29; and
      2. (ii) The weight of the minimum crew necessary to operate the rotorcraft, assuming for each crew member a weight no more than 170 pounds, or any lower weight selected by the applicant, or included in the loading instructions; and
    2. (2) Not less than:
      1. (i) The lowest weight selected by the applicant;
      2. (ii) The design minimum weight (the lowest weight at which compliance with each applicable structural loading condition of this chapter is shown); or
      3. (iii) The lowest weight at which compliance with each applicable flight requirement of this chapter is shown.
  3. (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. (1) [The rotorcraft-load combination does not include human external cargo;
    2. (2) [Structural component approval for external load operations under either 527.865 or under equivalent operational standards is obtained;
    3. (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. [(4) Structural components of the rotorcraft are shown to comply with the applicable structural requirements of this chapter under the increased loads and stresses caused by the weight increase over that established under paragraph (a) of this section; and
    5. [(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 527.865 (a) and (d) of this chapter.]

    (Change 527-4)

527.27 Centre of Gravity Limits

The extreme forward and aft centres of gravity and, where critical, the extreme lateral centres of gravity must be established for each weight established under 527.25. Such an extreme may not lie beyond:

  1. (a) The extremes selected by the applicant;
  2. (b) The extremes within which the structure is proven; or
  3. (c) The extremes within which compliance with the applicable flight requirements is shown.

527.29 Empty Weight and Corresponding Centre of Gravity

  1. (a) The empty weight and corresponding centre of gravity must be determined by weighing the rotorcraft without the crew and payload, but with:
    1. (1) Fixed ballast;
    2. (2) Unusable fuel; and
    3. (3) Full operating fluids, including:
      1. (i) Oil;
      2. (ii) Hydraulic fluid; and
      3. (iii) Other fluids required for normal operation of rotorcraft systems, except water intended for injection in the engines.
  2. (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.

527.31 Removable Ballast

Removable ballast may be used in showing compliance with the flight requirements of this subchapter.

527.33 Main Rotor Speed and Pitch Limits

  1. (a) Main rotor speed limits. A range of main rotor speeds must be established that:
    1. (1) With power-on, provides adequate margin to accommodate the variations in rotor speed occurring in any appropriate manoeuvre, and is consistent with the kind of governor or synchronizer used; and
    2. (2) With power-off, allows each appropriate autorotative manoeuvre to be performed throughout the ranges of airspeed and weight for which certification is requested.
  2. (b) Normal main rotor high pitch limits (power-on). For rotorcraft, 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. (1) Appropriate setting of the main rotor high pitch stop;
    2. (2) Inherent rotorcraft characteristics that make unsafe low main rotor speeds unlikely; or
    3. (3) Adequate means to warn the pilot of unsafe main rotor speeds.
  3. (c) Normal main rotor low pitch limits (power-off). It must be shown, with power-off, that:
    1. (1) The normal main rotor low pitch limit provides sufficient rotor speed, if any autorotative condition, under the most critical combinations of weight and airspeed; and
    2. (2) It is possible to prevent over speeding of the rotor without exceptional piloting skill.
  4. (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.
  5. (e) Main rotor low speed warning for helicopters. For each single engine helicopter, and each multi-engine 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. (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. (2) The warning may be furnished either through the inherent aerodynamic qualities of the helicopter or by a device.
    3. (3) The warning must be clear and distinct under all conditions, 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. (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.

Performance

527.45 General

  1. (a) Unless otherwise prescribed, the performance requirements of this subchapter must be met for still air and a standard atmosphere.
  2. (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.
  3. (c) The available power must correspond to engine power, not exceeding the approved power, less:
    1. (1) Installation losses; and
    2. (2) The power absorbed by the accessories and services appropriate to the particular ambient atmospheric conditions and the particular flight condition.
  4. (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.
  5. (e) For turbine engine-powered rotorcraft, the performance, as affected by engine power, must be based on a relative humidity of:
    1. (1) 80 percent, at and below standard temperature; and
    2. (2) 34 percent, at and above standard temperature plus 50° F. Between these two temperatures, the relative humidity must vary linearly.
  6. (f) For turbine-engine-powered rotorcraft, a means must be provided to permit the pilot to determine prior to take-off that each engine is capable of developing the power necessary to achieve the applicable rotorcraft performance prescribed in this subchapter.

527.49 Performance at Minimum Operating Speed

(amended 2009/05/11; no previous version)

  1. (a) For helicopters:
    (amended 2009/05/11; no previous version)
    1. (1) The hovering ceiling must be determined over the ranges of weight, altitude, and temperature for which certification is requested, with:
      (amended 2009/05/11; no previous version)
      1. (i) Take-off power;
        (amended 2009/05/11; no previous version)
      2. (ii) The landing gear extended; and
        (amended 2009/05/11; no previous version)
      3. (iii) The helicopter in-ground effect at a height consistent with normal take-off procedures; and
        (amended 2009/05/11; no previous version)
    2. (2) The hovering ceiling determined under paragraph (a)(1) of this section must be at least:
      (amended 2009/05/11; no previous version)
      1. (i) For reciprocating engine powered helicopters, 4,000 feet at maximum weight with a standard atmosphere;
        (amended 2009/05/11; no previous version)
      2. (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).
        (amended 2009/05/11; no previous version)
    3. (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 take-off power.
      (amended 2009/05/11; no previous version)
  2. (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:
    (amended 2009/05/11; no previous version)
    1. (1) Take-off power; and
      (amended 2009/05/11; no previous version)
    2. (2) The landing gear extended.
      (amended 2009/05/11; no previous version)

527.51 Take-off

The take-off, with take-off power and r.p.m., at the most critical centre of gravity, and with weight from the maximum weight at sea level to the weight for which take-off certification is requested for each altitude covered by this section:
(amended 2009/05/11)

  1. (a) May not require exceptional piloting skill or exceptionally favourable conditions throughout the ranges of altitude from standard sea level conditions to the maximum altitude for which take-off and landing certification is requested,
    and
    (amended 2009/05/11)
  2. (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 take-off and landing certification is requested or 7,000 feet density altitude, whichever is less.
    (amended 2009/05/11)

527.65 Climb:All Engines Operating

  1. (a) For rotorcraft other than helicopters:
    1. (1) The steady rate of climb, at VY, must be determined:
      1. (i) With maximum continuous power on each engine;
      2. (ii) With the landing gear retracted; and
      3. (iii) For the weights, altitudes, and temperatures for which certification is requested; and
    2. (2) [The climb gradient, at the rate of climb determined in accordance with paragraph (a)(1) of this section, must be either:]
      1. (i) At least 1:10 if the horizontal distance required to takeoff and climb over a 50-foot obstacle is determined for each weight, altitude, and temperature within the range for which certification is requested; or
      2. (ii) [At least 1:6 under standard sea level conditions;]
  2. (b) Each helicopter must meet the following requirements:
    1. (1) VY must be determined:
      1. (i) For standard sea level conditions;
      2. (ii) At maximum weight; and
      3. (iii) With maximum continuous power on each engine.
    2. (2) [The steady rate of climb must be determined:
      1. (i) At the climb speed selected by the applicant at or below VNE;
      2. (ii) Within the range from sea level up to the maximum altitude for which certification is requested;
      3. (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
      4. (iv) With maximum continuous power on each engine.

    (Change 527-4)

527.67 Climb:One Engine Inoperative

For multi-engine helicopters, the steady rate of climb (or descent), at VY (or at the speed for minimum rate of descent), must be determined with:

  1. (a) Maximum weight;
  2. (b) The critical engine inoperative and the remaining engines at either:
    1. (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. (2) Continuous OEI power for helicopters for which certification for the use of continuous OEI power is requested.

    (Change 527-1 (89-01-01))

527.71 Autorotation Performance

(amended 2009/05/11)

For single-engine helicopters and multi-engine helicopters that do not meet the Category A engine isolation requirements of Chapter 529 of this Manual, the minimum rate of descent airspeed and the best angle-of-glide airspeed must be determined in autorotation at:

  1. (a) Maximum weight; and
  2. (b) Rotor speed(s) selected by the applicant.

527.73 Reserved

(amended 2009/05/11)

527.75 Landing

  1. (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 favourable conditions, with:
    1. (1) Approach or autorotation speeds appropriate to the type of rotorcraft and selected by the applicant;
      (amended 2009/05/11)
    2. (2) The approach and landing made with:
      1. (i) Power off, for single engine rotorcraft and entered from steady state autorotation; or
        (amended 2009/05/11)
      2. (ii) One-engine inoperative (OEI) for multiengine rotorcraft, with each operating engine within approved operating limitations, and entered from an established OEI approach.
        (amended 2009/05/11)
    3. (3) The approach and landing entered from steady autorotation.
  2. (b) Multi-engine rotorcraft must be able to be landed safely after complete power failure under normal operating conditions.

527.79 Reserved

(amended 2009/05/11)

527.87 Height-Speed Envelope

(amended 2009/05/11; no previous version)

  1. (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:
    (amended 2009/05/11; no previous version)
    1. (1) Altitude, from standard sea level conditions to the maximum altitude capability of the rotorcraft, or 7,000 feet density altitude, whichever is less; and
      (amended 2009/05/11; no previous version)
    2. (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.
      (amended 2009/05/11; no previous version)
  2. (b) The applicable power failure conditions are:
    (amended 2009/05/11; no previous version)
    1. (1) For single-engine helicopters, full autorotation;
      (amended 2009/05/11; no previous version)
    2. (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 7,000 feet density altitude or the maximum altitude capability of the helicopter, whichever is less, and
      (amended 2009/05/11; no previous version)
    3. (3) For other rotorcraft, conditions appropriate to the type.
      (amended 2009/05/11; no previous version)

Flight Characteristics

527.141 General

The rotorcraft must:

  1. (a) Except as specifically required in the applicable section, meet the flight characteristics requirements of this subchapter:
    1. (1) At the altitudes and temperatures expected in operation;
    2. (2) Under any critical loading condition within the range of weights and centres of gravity for which certification is requested;
    3. (3) For power-on operations, under any condition of speed, power, and rotor r.p.m. for which certification is requested; and
    4. (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;
  2. (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. (1) Sudden failure of one engine, for multi-engine rotorcraft meeting transport Category A engine isolation requirements of Chapter 529 of this Manual;
    2. (2) Sudden, complete power failure, for other rotorcraft; and
    3. (3) Sudden, complete control system failures specified in 527.695 of this chapter; and
  3. (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 chapter.

527.143 Controllability and Manoeuvrability

  1. (a) The rotorcraft must be safely controllable and manoeuvrable:
    1. (1) During steady flight; and
    2. (2) During any manoeuvre appropriate to the type including:
      1. (i) Takeoff;
      2. (ii) Climb;
      3. (iii) Level flight;
      4. (iv) Turning flight;
      5. (v) Autorotation;
        (amended 2009/05/11)
      6. (vi) Landing (power on and power off); and
      7. (vii) Recovery to power-on flight from a balked autorotative approach.
  2. (b) The margin of cyclic control must allow satisfactory roll and pitch control at VNE with:
    1. (1) Critical weight;
    2. (2) Critical centre of gravity;
    3. (3) Critical rotor r.p.m.; and
    4. (4) Power off (except for helicopters demonstrating compliance with paragraph (f) of this section) and power-on.
      (amended 2009/05/11)
  3. (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 manoeuvre appropriate to the type (such as crosswind take offs, sideward flight, and rearward flight):
    (amended 2009/05/11)
    1. (1) With altitude, from standard sea level conditions to the maximum take-off and landing altitude capability of the rotorcraft or 7,000 feet density altitude, whichever is less; with:
      (amended 2009/05/11)
      1. (i) Critical weight;
        (amended 2009/05/11)
      2. (ii) Critical centre of gravity; and
        (amended 2009/05/11)
      3. (iii) Critical rotor r.p.m.; and
        (amended 2009/05/11)
    2. (2) For take-off and landing altitudes above 7,000 feet density altitude with:
      (amended 2009/05/11)
      1. (i) Weight selected by the applicant;
        (amended 2009/05/11)
      2. (ii) Critical centre of gravity; and
        (amended 2009/05/11)
      3. (iii) Critical rotor r.p.m.
        (amended 2009/05/11)
    3. (3) Critical rotor r.p.m.; and
    4. (4) Altitude, from standard sea level conditions to the maximum altitude capability of the rotorcraft or 7,000 feet, whichever is less.
  4. (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:
    (amended 2009/05/11)
    1. (1) Weight selected by the applicant;
      (amended 2009/05/11)
    2. (2) Critical centre of gravity;
      (amended 2009/05/11)
    3. (3) Rotor r.p.m. selected by the applicant; and
      (amended 2009/05/11)
    4. (4) Altitude, from standard sea level conditions to the maximum take-off and landing altitude capability of the rotorcraft.
      (amended 2009/05/11)
  5. (e) The rotorcraft, after:
    1. (1) Failure of one engine in the case of multi-engine rotorcraft that meet transport Category A engine isolation requirements, or
    2. (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:
      1. (i) For the cruise condition, one second, or normal pilot reaction time (whichever is greater); and
      2. (ii) For any other condition, normal pilot reaction time.
  6. (f) For helicopters for which a VNE (power-off) is established under 527.1505 (c), compliance must be demonstrated with the following requirements with critical weight, critical centre of gravity, and critical rotor r.p.m.:
    1. (1) The helicopter must be safely slowed to VNE (power-off), without exceptional pilot skill, after the last operating engine is made inoperative at power-on VNE.
    2. (2) At a speed of 1.1 VNE (power-off), the margin of cyclic control must allow satisfactory roll and pitch control with power off.

527.151 Flight Controls

  1. (a) Longitudinal, lateral, directional, and collective controls may not exhibit excessive breakout force, friction, or preload.
  2. (b) Control system forces and free play may not inhibit a smooth, direct rotorcraft response to control system input.

527.161 Trim Control

The trim control:

  1. (a) Must trim any steady longitudinal, lateral, and collective control forces to zero in level flight at any appropriate speed; and
  2. (b) May not introduce any undesirable discontinuities in control force gradients.

527.171 Stability:General

The rotorcraft must be able to be flown, without undue pilot fatigue or strain, in any normal manoeuvre for a period of time as long as that expected in normal operation. At least three landings and take-offs must be made during this demonstration.

527.173 Static Longitudinal Stability

  1. (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.
    (amended 2009/05/11)
  2. (b) Throughout the full range of altitude for which certification is required, with the throttle and collective pitch held constant during the manoeuvres specified in 527.175 (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 Minister 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.
    (amended 2009/05/11)
  3. (c) During the manoeuvres specified in 527.175 (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.

527.175 Demonstration of Static Longitudinal Stability

  1. (a) Climb. Static longitudinal stability must be shown in the climb condition at speeds from VY–10 kt to VY+10 kt with:
    (amended 2009/05/11)
    1. (1) Critical weight;
    2. (2) Critical centre of gravity;
    3. (3) Maximum continuous power;
    4. (4) The landing gear retracted; and
    5. (5) The rotorcraft trimmed at VY.
  2. (b) Cruise. Static longitudinal stability must be shown in the cruise condition at speeds from 0.8 VNE –10 kt to 0.8 VNE +10 kt or, if VH is less than 0.8 VNE , from V –10 kt to V +10 kt, with:
    (amended 2009/05/11)
    1. (1) Critical weight;
    2. (2) Critical centre of gravity;
    3. (3) Power for level flight at 0.8 VNE or VNH, whichever is less;
      (amended 2009/05/11)
    4. (4) The landing gear retracted; and
    5. (5) The rotorcraft trimmed at 0.8 VNE or VNH, whichever is less.
      (amended 2009/05/11)
  3. (c) VNE. Static longitudinal stability must be shown at speeds from VNE -20 kt to VNE  with:
    (amended 2009/05/11)
    1. (1) Critical weight;
      (amended 2009/05/11)
    2. (2) Critical centre of gravity;
      (amended 2009/05/11)
    3. (3) Power required for level flight at VNE -10 kt or maximum continuous power, whichever is less;
      (amended 2009/05/11)
    4. (4) The landing gear retracted; and
      (amended 2009/05/11)
    5. (5) The rotorcraft trimmed at VNE -10 kt.
      (amended 2009/05/11)
  4. (d) Autorotation. Static longitudinal stability must be shown in autorotation at:
    (amended 2009/05/11)
    1. (1) Airspeeds from the minimum rate of descent airspeed -10 kt to the minimum rate of descent airspeed +10 kt, with:
      (amended 2009/05/11)
      1. (i) Critical weight;
      2. (ii) Critical centre of gravity;
      3. (iii) The landing gear extended; and
        (amended 2009/05/11)
      4. (iv) The rotorcraft trimmed at the minimum rate of descent airspeed.
        (amended 2009/05/11)
    2. (2) Airspeeds from best angle-of-glide airspeed -10 kt to the best angle-of-glide airspeed +10 kt, with:
      (amended 2009/05/11)
      1. (i) Critical weight;
        (amended 2009/05/11)
      2. (ii) Critical centre of gravity;
        (amended 2009/05/11)
      3. (iii) The landing gear retracted; and
        (amended 2009/05/11)
      4. (iv) The rotorcraft trimmed at the best angle-of-glide airspeed.
        (amended 2009/05/11)

527.177 Static Directional Stability

  1. (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 527.175(a), (b), and (c). Sideslip angles must increase with steadily increasing directional control deflection for sideslip angles up to the lesser of:
    (amended 2009/05/11)
    1. (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 VNE;
      (amended 2009/05/11)
    2. (2) The steady state sideslip angles established by 527.351;
      (amended 2009/05/11)
    3. (3) A sideslip angle selected by the applicant, which corresponds to a sideforce of at least 0.1g; or
      (amended 2009/05/11)
    4. (4) The sideslip angle attained by maximum directional control input.
      (amended 2009/05/11)
  2. (b) Sufficient cues must accompany the sideslip to alert the pilot when the aircraft is approaching the sideslip limits.
    (amended 2009/05/11)
  3. (c) During the manoeuvre 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.
    (amended 2009/05/11)

Ground and Water Handling Characteristics 

527.231 General

The rotorcraft must have satisfactory ground and water handling characteristics, including freedom from uncontrollable tendencies in any condition expected in operation.

527.235 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.

527.239 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.

527.241 Ground Resonance

The rotorcraft may have no dangerous tendency to oscillate on the ground with the rotor turning.

Miscellaneous Flight Requirements

527.251 Vibration

Each part of the rotorcraft must be free from excessive vibration under each appropriate speed and power condition.

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