ED Decision 2003/15/RM
The rotorcraft must:
(a) Except as specifically required in the applicable paragraph, 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 centres of gravity for which certification is requested;
(3) For power-on operations, under any condition of speed, power, and rotor rpm for which certification is requested; and
(4) For power-off operations, under any condition of speed and rotor rpm 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 multi-engine rotorcraft meeting category A engine isolation requirements of CS-29;
(2) Sudden, complete power failure for other rotorcraft; and
(3) Sudden, complete control system failures specified in CS 27.695; 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.
Appendix B – Airworthiness Criteria for Helicopter Instrument Flight
ED Decision 2007/013/R
I. General. A small helicopter may not be type certificated for operation under the instrument flight rules (IFR) unless it meets the design and installation requirements contained in this appendix.
II. Definitions
(a) VYI means instrument climb speed, utilised instead of VY for compliance with the climb requirements for instrument flight.
(b) VNEI means instrument flight never exceed speed, utilised instead of VNE for compliance with maximum limit speed requirements for instrument flight.
(c) VMINI means instrument flight minimum speed, utilised in complying with minimum limit speed requirements for instrument flight.
III. Trim. It must be possible to trim the cyclic, collective, and directional control forces to zero at all approved IFR airspeeds, power settings, and configurations appropriate to the type.
IV. Static longitudinal stability
(a) General. The helicopter must possess positive static longitudinal control force stability at critical combinations of weight and centre of gravity at the conditions specified in paragraphs IV (b) or (c) of this Appendix. The stick force must vary with speed so that any substantial speed change results in a stick force clearly perceptible to the pilot. For single pilot approval the airspeed must return to within 10% of the trim speed when the control force is slowly released for each trim condition specified in paragraph IV(b) of this Appendix.
(b) For single-pilot approval
(1) Climb. Stability must be shown in climb throughout the speed range 37 km/h (20 knots) either side of trim with –
(i) The helicopter trimmed at VYI;
(ii) Landing gear retracted (if retractable); and
(iii) Power required for limit climb rate (at least 5 m/s (1000 fpm)) at VYI or maximum continuous power, whichever is less.
(2) Cruise. Stability must be shown throughout the speed range from 0.7 to 1.1 VH or VNEI, whichever is lower, not to exceed ±37 km/h (±20 knots) from trim with –
(i) The helicopter trimmed and power adjusted for level flight at 0.9 VH or 0.9 VNEI, whichever is lower; and
(ii) Landing gear retracted (if retractable).
(3) Slow cruise. Stability must be shown throughout the speed range from 0.9 VMINI to 1.3 VMINI or 37 km/h (20 knots) above trim speed, whichever is greater, with –
(i) The helicopter trimmed and power adjusted for level flight at 1.1 VMINI; and
(ii) Landing gear retracted (if retractable).
(4) Descent. Stability must be shown throughout the speed range 37 km/h (20 knots) either side of trim with –
(i) The helicopter trimmed at 0.8 VH or 0.8 VNEI (or 0.8 VLE for the landing gear extended case), whichever is lower;
(ii) Power required for 1000 fpm descent at trim speed; and
(iii) Landing gear extended and retracted, if applicable.
(5) Approach. Stability must be shown throughout the speed range from 0.7 times the minimum recommended approach speed to 37 km/h (20 knots) above the maximum recommended approach speed with –
(i) The helicopter trimmed at the recommended approach speed or speeds;
(ii) Landing gear extended and retracted, if applicable; and
(iii) Power required to maintain a 3° glide path and power required to maintain the steepest approach gradient for which approval is requested.
(c) Helicopters approved for a minimum crew of two pilots must comply with the provisions of paragraphs IV(b)(2) and IV(b)(5) of this Appendix.
V. Static lateral-directional stability
(a) Static directional stability must be positive throughout the approved ranges of airspeed, power, and vertical speed. In straight and steady sideslips up to ± 10° from trim, directional control position must increase without discontinuity with the angle of sideslip, except for a small range of sideslip angles around trim. At greater angles up to the maximum sideslip angle appropriate to the type, increased directional control position must produce increased angle of sideslip. It must be possible to maintain balanced flight without exceptional pilot skill or alertness.
(b) During sideslips up to ± 10° from trim throughout the approved ranges of airspeed, power, and vertical speed there must be no negative dihedral stability perceptible to the pilot through lateral control motion or force. Longitudinal cyclic movement with sideslip must not be excessive.
VI. Dynamic stability
(a) For single-pilot approval –
(1) Any oscillation having a period of less than 5 seconds must damp to ½ amplitude in not more than one cycle.
(2) Any oscillation having a period of 5 seconds or more but less than 10 seconds must damp to ½ amplitude in not more than two cycles.
(3) Any oscillation having a period of 10 seconds or more but less than 20 seconds must be damped.
(4) Any oscillation having a period of 20 seconds or more may not achieve double amplitude in less than 20 seconds.
(5) Any a periodic response may not achieve double amplitude in less than 6 seconds.
(b) For helicopters approved with a minimum crew of two pilots –
(1) Any oscillation having a period of less than 5 seconds must damp to ½ amplitude in not more than two cycles.
(2) Any oscillation having a period of 5 seconds or more but less than 10 seconds must be damped.
(3) Any oscillation having a period of 10 seconds or more may not achieve double amplitude in less than 10 seconds.
VII. Stability augmentation system (SAS)
(a) If a SAS is used, the reliability of the SAS must be related to the effects of its failure. Any SAS failure condition that would prevent continued safe flight and landing must be extremely improbable. It must be shown that, for any failure condition of the SAS which is not shown to be extremely improbable –
(1) The helicopter is safely controllable when the failure or malfunction occurs at any speed or altitude within the approved IFR operating limitations; and
(2) The overall flight characteristics of the helicopter allow for prolonged instrument flight without undue pilot effort. Additional unrelated probable failures affecting the control system must be considered. In addition:
(i) The controllability and manoeuvrability requirements in Subpart B of CS-27 must be met throughout a practical flight envelope;
(ii) The flight control, trim, and dynamic stability characteristics must not be impaired below a level needed to allow continued safe flight and landing; and
(iii) The static longitudinal and static directional stability requirements of Subpart B of CS-27 must be met throughout a practical flight envelope.
(b) The SAS must be designed so that it cannot create a hazardous deviation in flight path or produce hazardous loads on the helicopter during normal operation or in the event of malfunction or failure, assuming corrective action begins within an appropriate period of time. Where multiple systems are installed, subsequent malfunction conditions must be considered in sequence unless their occurrence is shown to be improbable.
VIII. Equipment, systems, and installation. The basic equipment and installation must comply with CS 29.1303, 29.1431 and 29.1433, with the following exceptions and additions:
(a) Flight and navigation instruments
(1) A magnetic gyro-stabilised direction indicator instead of the gyroscopic direction indicator required by CS 29.1303(h); and
(2) A standby attitude indicator which meets the requirements of CS 29.1303(g)(1) to (7), instead of a rate-of-turn indicator required by CS 29.1303(g). For two-pilot configurations, one pilot’s primary indicator may be designated for this purpose. If standby batteries are provided they may be charged from the aircraft electrical system if adequate isolation is incorporated.
(b) Miscellaneous requirements
(1) Instrument systems and other systems essential for IFR flight that could be adversely affected by icing must be adequately protected when exposed to the continuous and intermittent maximum icing conditions defined in appendix C of CS-29, whether or not the rotorcraft is certificated for operation in icing conditions.
(2) There must be means in the generating system to automatically de-energise and disconnect from the main bus any power source developing hazardous overvoltage.
(3) Each required flight instrument using a power supply (electric, vacuum, etc.) must have a visual means integral with the instrument to indicate the adequacy of the power being supplied.
(4) When multiple systems performing like functions are required, each system must be grouped, routed, and spaced so that physical separation between systems is provided to ensure that a single malfunction will not adversely affect more than one system.
(5) For systems that operate the required flight instruments at each pilot’s station –
(i) Only the required flight instruments for the first pilot may be connected to that operating system;
(ii) Additional instruments, systems, or equipment may not be connected to an operating system for a second pilot unless provisions are made to ensure the continued normal functioning of the required instruments in the event of any malfunction of the additional instruments, systems, or equipment which is not shown to be extremely improbable;
(iii) The equipment, systems, and installations must be designed so that one display of the information essential to the safety of flight which is provided by the instruments will remain available to a pilot, without additional crewmember action, after any single failure or combination of failures that is not shown to be extremely improbable; and For single-pilot configurations, instruments which require a static source must be provided with a means of selecting an alternate source and that source must be calibrated.
IX. Rotorcraft flight manual. A rotorcraft flight manual or rotorcraft flight manual IFR supplement must be provided and must contain –
(a) Limitations. The approved IFR flight envelope, the IFR flight crew composition, the revised kinds of operation, and the steepest IFR precision approach gradient for which the helicopter is approved;
(b) Procedures. Required information for proper operation of IFR systems and the recommended procedures in the event of stability augmentation or electrical system failures; and
(c) Performance. If VYI differs from VY, climb performance at VYI and with maximum continuous power throughout the ranges of weight, altitude, and temperature for which approval is requested.
[Amdt. No.: 27/1]
CS 27.143 Controllability and manoeuvrability
ED Decision 2007/013/R
(a) The rotorcraft must be safely controllable and manoeuvrable:
(1) During steady flight; and
(2) During any manoeuvre appropriate to the type, including:
(i) Take-off;
(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 VNE with:
(1) Critical weight;
(2) Critical centre of gravity;
(3) Critical rotor rpm; and
(4) Power off, except for helicopters demonstrating compliance with sub-paragraph (f), and power on.
(c) Wind velocities from zero to at least 31 km/h (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:
(1) With altitude, from standard sea-level conditions to the maximum take-off and landing altitude capability of the rotorcraft or 2134 m (7000 ft) density altitude, whichever is less; with:
(i) Critical weight;
(ii) Critical centre of gravity; and
(iii) Critical rotor rpm.
(2) For take-off and landing altitudes above 2134 m (7000 ft) density altitude with:
(i) Weight selected by the applicant;
(ii) Critical center of gravity; and
(iii) Critical rotor rpm.
(d) Wind velocities from zero to at least 31 km/h (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 rpm selected by the applicant; and
(4) Altitude, from standard sea-level conditions to the maximum take-off and landing altitude capability of the rotorcraft.
(e) The rotorcraft, after
(1) failure of one engine in the case of multi-engine rotorcraft that meet 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.
(f) For helicopters for which a VNE (power-off) is established under CS 27.1505(c), compliance must be demonstrated with the following requirements with critical weight, critical centre of gravity, and critical rotor rpm:
(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) At a speed of 1.1 VNE (power-off), the margin of cyclic control must allow satisfactory roll and pitch control with power off.
[Amdt. No.: 27/1]
ED Decision 2003/15/RM
(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.
ED Decision 2003/15/RM
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.
ED Decision 2003/15/RM
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.
CS 27.173 Static longitudinal stability
ED Decision 2007/013/R
(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 manoeuvres specified in CS 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 Agency 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 ±9 km/h (±5 knots) of the desired trim airspeed without exceptional piloting skill or alertness.
[Amdt. No.: 27/1]
CS 27.175 Demonstration of static longitudinal stability
ED Decision 2007/013/R
(a) Climb. Static longitudinal stability must be shown in the climb condition at speeds from VY - 19 km/h (10 knots) to VY + 19 km/h (10 knots), with:
(1) Critical weight;
(2) Critical centre of gravity;
(3) Maximum continuous power;
(4) The landing gear retracted; and
(5) The rotorcraft trimmed at VY.
(b) Cruise. Static longitudinal stability must be shown in the cruise condition at speeds from 0.8 VNE – 19 km/h (10 knots) to 0.8 VNE + 19 km/h (10 knots) or, if VH is less than 0.8 VNE, from VH – 19 km/h (10 knots) to VH + 19 km/h (10 knots), with:
(1) Critical weight;
(2) Critical centre of gravity;
(3) Power for level flight at 0.8 VNE or VH, whichever is less;
(4) The landing gear retracted; and
(5) The rotorcraft trimmed at 0.8 VNE or VH, whichever is less.
(c) VNE. Static longitudinal stability must be shown at speeds from VNE – 28 km/h (20 knots) to VNE with:
(1) Critical weight;
(2) Critical center of gravity;
(3) Power required for level flight at VNE – 19 km/h (10 knots) or maximum continuous power, whichever is less;
(4) The landing gear retracted; and
(5) The rotorcraft trimmed at VNE – 19 km/h (10 knots)
(d) Autorotation. Static longitudinal stability must be shown in autorotation at:
(1) Airspeeds from the minimum rate of descent airspeed – 19 km/h (10 knots) to the minimum rate of descent airspeed + 19 km/h (10 knots), 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 the best angle-of-glide airspeed – 19 km/h (10 knots) to the best angle-of-glide airspeed + 19 km/h (10 knots), 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.
[Amdt. No.: 27/1]
CS 27.177 Static directional stability
ED Decision 2007/013/R
(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 throttle and collective controls held constant at the trim conditions specified in CS 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 28 km/h (15 knots) less than the speed for minimum rate of descent varying linearly to ±10 degrees from trim at VNE;
(2) The steady state sideslip angles established by CS 27.351;
(3) A sideslip angle selected by the applicant which corresponds to a sideforce of at least 0.1 g; 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 manoeuvre specified in sub-paragraph (a) of this paragraph, 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.
[Amdt. No.: 27/1]