Subpart B — Flight

GENERAL

ED Decision 2003/16/RM

Each requirement of this Subpart 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:

(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 centre of gravity, if compliance cannot be reasonably inferred from combinations investigated.

ED Decision 2007/014/R

(a)Maximum weight. The maximum weight (the highest weight at which compliance with each applicable requirement of this CS-29 is shown) or, at the option of the applicant, the highest weight for each altitude and for each practicably separable operating condition, such as take-off, en-route operation, and landing, must be established so that it is not more than:

(1)The highest weight selected by the applicant;

(2)The design maximum weight (the highest weight at which compliance with each applicable structural loading condition of this CS-29 is shown); or

(3)The highest weight at which compliance with each applicable flight requirement of this CS-29 is shown.

(4)For Category B rotorcraft with 9 or less passenger seats, the maximum weight, altitude, and temperature at which the rotorcraft can safely operate near the ground with the maximum wind velocity determined under CS 29.143(c) and may include other demonstrated wind velocities and azimuths. The operating envelopes must be stated in the Limitations section of the Rotorcraft Flight Manual.

(b)Minimum weight. The minimum weight (the lowest weight at which compliance with each applicable requirement of this CS-29 is shown) must be established so that it is not less than:

(1)The lowest weight selected by the applicant;

(2)The design minimum weight (the lowest weight at which compliance with each structural loading condition of this CS-29 is shown); or

(3)The lowest weight at which compliance with each applicable flight requirement of this CS-29 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 sub-paragraph (a) 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 CS 29.865, or under equivalent operational standards is obtained,

(3)The portion of the total weight that is greater than the maximum weight established under sub-paragraph (a) 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 CS-29 under the increased loads and stresses caused by the weight increase over that established under sub-paragraph (a), and

(5)Operation of the rotorcraft at a total weight greater than the maximum certificated weight established under sub-paragraph (a) is limited by appropriate operating limitations under CS 29.865(a) and (d).

[Amdt. No.: 29/1]

ED Decision 2003/16/RM

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 CS 29.25. 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.

ED Decision 2003/16/RM

(a)The empty weight and corresponding centre 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 rotorcraft 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.

ED Decision 2003/16/RM

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

ED Decision 2003/16/RM

(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 manoeuvre, and is consistent with the kind of governor or synchroniser used; and

(2)With power off, allows each appropriate autorotative manoeuvre to be performed throughout the ranges of airspeed and weight for which certification is requested.

(b)Normal main rotor high pitch limit (power-on). For rotorcraft, except helicopters required to have a main rotor low speed warning under sub-paragraph (e), 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 limit (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 sub-paragraph (b)(1), 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 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)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 jeopardise 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 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)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

ED Decision 2003/16/RM

(a)The performance prescribed in this subpart must be determined:

(1)With normal piloting skill; and

(2)Without exceptionally favourable conditions.

(b)Compliance with the performance requirements of this subpart must be shown:

(1)For still air at sea-level with a standard atmosphere; and

(2)For the approved range of atmospheric variables.

(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 at the values for which certification is requested and approved.

(d)For reciprocating engine-powered rotorcraft, the performance, as affected by engine power, must be based on a relative humidity of 80% 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%, at and below standard temperature; and

(2)34%, at and above standard temperature plus 28°C (50°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 take-off that each engine is capable of developing the power necessary to achieve the applicable rotorcraft performance prescribed in this subpart.

ED Decision 2003/16/RM

(a)For each Category A helicopter, the hovering performance must be determined over the ranges of weight, altitude and temperature for which take-off data are scheduled:

(1)With not more than take-off power;

(2)With the landing gear extended; and

(3)At a height consistent with the procedure used in establishing the take-off, climbout and rejected take-off paths.

(b)For each Category B helicopter, the hovering performance must be determined over the ranges of weight, altitude and temperature for which certification is requested, with:

(1)Take-off power;

(2)The landing gear extended; and

(3)The helicopter in ground effect at a height consistent with normal take-off procedures.

(c)For each helicopter, the out-of ground- effect hovering performance must be determined over the ranges of weight, altitude and temperature for which certification is requested, with take-off power.

(d)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)Take-off power; and

(2)The landing gear extended.

ED Decision 2003/16/RM

(a)The take-off data required by CS 29.53, 29.55, 29.59, 29.60, 29.61, 29.62, 29.63 and 29.67 must be determined:

(1)At each weight, altitude, and temperature selected by the applicant; and

(2)With the operating engines within approved operating limitations.

(b)Take-off data must:

(1)Be determined on a smooth, dry, hard surface; and

(2)Be corrected to assume a level take-off surface.

(c)No take-off made to determine the data required by this paragraph may require exceptional piloting skill or alertness, or exceptionally favourable conditions.

ED Decision 2003/16/RM

The take-off performance must be determined and scheduled so that, if one engine fails at any time after the start of take-off, the rotorcraft can:

(a)Return to and stop safely on, the take-off area; or

(b)Continue the take-off and climb-out, and attain a configuration and airspeed allowing compliance with CS 29.67(a)(2).

ED Decision 2003/16/RM

(a)The take-off decision point (TDP) is the first point from which a continued take-off capability is assured under CS 29.59 and is the last point in the take-off path from which a rejected take-off is assured within the distance determined under CS 29.62.

(b)The TDP must be established in relation to the take-off path using no more than two parameters, such as airspeed and height, to designate the TDP.

(c)Determination of the TDP must include the pilot recognition time interval following failure of the critical engine.

ED Decision 2003/16/RM

(a)The take-off path extends from the point of commencement of the take-off procedure to a point at which the rotorcraft is 305 m (1000 ft) above the take-off surface and compliance with CS 29.67(a)(2) is shown. In addition:

(1)The take-off path must remain clear of the height-velocity envelope established in accordance with CS 29.87;

(2)The rotorcraft must be flown to the engine failure point at which point the critical engine must be made inoperative and remain inoperative for the rest of the take-off;

(3)After the critical engine is made inoperative, the rotorcraft must continue to the TDP, and then attain VTOSS.

(4)Only primary controls may be used while attaining VTOSS and while establishing a positive rate of climb. Secondary controls that are located on the primary controls may be used after a positive rate of climb and VTOSS are established but in no case less than 3 seconds after the critical engine is made inoperative; and

(5)After attaining VTOSS and a positive rate of climb, the landing gear may be retracted.

(b)During the take-off path determination made in accordance with sub-paragraph (a) and after attaining VTOSS and a positive rate of climb, the climb must be continued at a speed as close as practicable to, but not less than, VTOSS until the rotorcraft is 61 m (200 ft) above the take-off surface. During this interval, the climb performance must meet or exceed that required by CS 29.67(a)(1).

(c)During the continued take-off the rotorcraft shall not descend below 4.6 m (15 ft) above the take-off surface when the TDP is above 4.6 m (15 ft).

(d)From 61 m (200 ft) above the take-off surface, the rotorcraft take-off path must be level or positive until a height 305 m (1000 ft) above the take-off surface is attained with not less than the rate of climb required by CS 29.67(a)(2). Any secondary or auxiliary control may be used after attaining 61 m (200 ft) above the take-off surface.

(e)Take-off distance will be determined in accordance with CS 29.61.

ED Decision 2003/16/RM

(a)The elevated heliport take-off path extends from the point of commencement of the take-off procedure to a point in the take-off path at which the rotorcraft is 305 m (1 000 ft) above the take-off surface and compliance with CS 29.67(a)(2) is shown. In addition:

(1)The requirements of CS 29.59(a) must be met;

(2)While attaining VTOSS and a positive rate of climb, the rotorcraft may descend below the level of the take-off surface if, in so doing and when clearing the elevated heliport edge, every part of the rotorcraft clears all obstacles by at least 4.6 m (15 ft);

(3)The vertical magnitude of any descent below the take-off surface must be determined; and

(4)After attaining VTOSS and a positive rate of climb, the landing gear may be retracted.

(b)The scheduled take-off weight must be such that the climb requirements of CS 29.67(a)(1) and CS 29.67(a)(2) will be met.

(c)Take-off distance will be determined in accordance with CS 29.61.

ED Decision 2003/16/RM

(a)The normal take-off distance is the horizontal distance along the take-off path from the start of the take-off to the point at which the rotorcraft attains and remains at least 11 m (35 ft) above the take-off surface, attains and maintains a speed of at least VTOSS; and establishes a positive rate of climb, assuming the critical engine failure occurs at the engine failure point prior to the TDP.

(b)For elevated heliports, the take-off distance is the horizontal distance along the take-off path from the start of the take-off to the point at which the rotorcraft attains and maintains a speed of at least VTOSS and establishes a positive rate of climb, assuming the critical engine failure occurs at the engine failure point prior to the TDP.

ED Decision 2003/16/RM

The rejected take-off distance and procedures for each condition where take-off is approved will be established with:

(a)The take-off path requirements of CS 29.59 and 29.60 being used up to the TDP where the critical engine failure is recognised, and the rotorcraft landed and brought to a stop on the take-off surface;

(b)The remaining engines operating within approved limits;

(c)The landing gear remaining extended throughout the entire rejected take-off; and

(d)The use of only the primary controls until the rotorcraft is on the ground. Secondary controls located on the primary control may not be used until the rotorcraft is on the ground. Means other than wheel brakes may be used to stop the rotorcraft if the means are safe and reliable and consistent results can be expected under normal operating conditions.

ED Decision 2003/16/RM

The horizontal distance required to take-off and climb over a 15 m (50-foot) obstacle must be established with the most unfavourable centre of gravity. The take-off may be begun in any manner if –

(a)The take-off surface is defined;

(b)Adequate safeguards are maintained to ensure proper centre of gravity and control positions; and

(c)A landing can be made safely at any point along the flight path if an engine fails.

ED Decision 2003/16/RM

Compliance with the requirements of CS 29.65 and 29.67 must be shown at each weight, altitude and temperature within the operational limits established for the rotorcraft and with the most unfavourable centre of gravity for each configuration. Cowl flaps, or other means of controlling the engine-cooling air supply, will be in the position that provides adequate cooling at the temperatures and altitudes for which certification is requested.

ED Decision 2003/16/RM

(a)The steady rate of climb must be determined:

(1)With maximum continuous power;

(2)With the landing gear retracted; and

(3)At VY for standard sea-level conditions and at speeds selected by the applicant for other conditions.

(b)For each Category B rotorcraft except helicopters, the rate of climb determined under sub-paragraph (a) must provide a steady climb gradient of at least 1:6 under standard sea-level conditions.

ED Decision 2003/16/RM

(a)For Category A rotorcraft, in the critical take-off configuration existing along the take-off path, the following apply:

(1)The steady rate of climb without ground effect, 61 m (200 ft) above the take-off surface, must be at least 30 m (100 ft) per minute, for each weight, altitude, and temperature for which take-off data are to be scheduled with:

(i)The critical engine inoperative and the remaining engines within approved operating limitations, except that for rotorcraft for which the use of 30-second/2-minute OEI power is requested, only the 2-minute OEI power may be used in showing compliance with this paragraph;

(ii)The landing gear extended; and

(iii)The take-off safety speed selected by the applicant.

(2)The steady rate of climb without ground effect, 305 m (1 000 ft) above the take- off surface, must be at least 46 m (150 ft) per minute, for each weight, altitude, and temperature for which take-off data are to be scheduled with:

(i)The critical engine inoperative and the remaining engines at maximum continuous power including continuous OEI power, if approved, or at 30-minute OEI power for rotorcraft for which certification for use of 30-minute OEI power is requested;

(ii)The landing gear retracted; and

(iii)The speed selected by the applicant.

(3)The steady rate of climb (or descent), in feet per minute, at each altitude and temperature at which the rotorcraft is expected to operate and at each weight within the range of weights for which certification is requested, must be determined with:

(i)The critical engine inoperative and the remaining engines at maximum continuous power including continuous OEI power, if approved, and at 30-minute OEI power for rotorcraft for which certification for the use of 30-minute OEI power is requested;

(ii)The landing gear retracted; and

(iii)The speed selected by the applicant.

(b)For multi-engine Category B rotorcraft meeting the Category A engine isolation requirements, the steady rate of climb (or descent) must be determined at the speed for best rate of climb (or minimum rate of descent) at each altitude, temperature, and weight at which the rotorcraft is expected to operate, with the critical engine inoperative and the remaining engines at maximum continuous power including continuous OEI power, if approved, and at 30-minute OEI power for rotorcraft for which certification for the use of 30-minute OEI power is requested.

ED Decision 2003/16/RM

For each Category B helicopter, except multi-engine helicopters meeting the requirements of CS 29.67(b) and the powerplant installation requirements of Category A, the steady angle of glide must be determined in autorotation:

(a)At the forward speed for minimum rate of descent as selected by the applicant;

(b)At the forward speed for best glide angle;

(c)At maximum weight; and

(d)At the rotor speed or speeds selected by the applicant.

ED Decision 2003/16/RM

(a)For each rotorcraft:

(1)The corrected landing data must be determined for a smooth, dry, hard and level surface;

(2)The approach and landing must not require exceptional piloting skill or exceptionally favourable conditions; and,

(3)The landing must be made without excessive vertical acceleration or tendency to bounce, nose over, ground loop, porpoise, or water loop.

(b)The landing data required by CS 29.77, 29.79, 29.81, 29.83 and 29.85 must be determined:

(1)At each weight, altitude and temperature for which landing data are approved:

(2)With each operating engine within approved operating limitations: and

(3)With the most unfavourable centre of gravity.

ED Decision 2003/16/RM

(a)The landing decision point (LDP) is the last point in the approach and landing path from which a balked landing can be accomplished in accordance with CS 29.85.

(b)Determination of the LDP must include the pilot recognition time interval following failure of the critical engine.

ED Decision 2003/16/RM

(a)For Category A rotorcraft:

(1)The landing performance must be determined and scheduled so that if the critical engine fails at any point in the approach path, the rotorcraft can either land and stop safely or climb out and attain a rotorcraft configuration and speed allowing compliance with the climb requirement of CS 29.67(a)(2);

(2)The approach and landing paths must be established with the critical engine inoperative so that the transition between each stage can be made smoothly and safely;

(3)The approach and landing speeds must be selected for the rotorcraft and must be appropriate to the type of rotorcraft; and

(4)The approach and landing path must be established to avoid the critical areas of the height-velocity envelope determined in accordance with CS 29.87.

(b)It must be possible to make a safe landing on a prepared landing surface after complete power failure occurring during normal cruise.

ED Decision 2003/16/RM

The horizontal distance required to land and come to a complete stop (or to a speed of approximately 5.6 km/h (3 knots) for water landings) from a point 15 m (50 ft) above the landing surface must be determined from the approach and landing paths established in accordance with CS 29.79.

ED Decision 2003/16/RM

(a)For each Category B rotorcraft, the horizontal distance required to land and come to a complete stop (or to a speed of approximately 5.6 km/h (3 knots) for water landings) from a point 15 m (50 ft) above the landing surface must be determined with:

(1)Speeds appropriate to the type of rotorcraft and chosen by the applicant to avoid the critical areas of the height-velocity envelope established under CS 29.87; and

(2)The approach and landing made with power on and within approved limits.

(b)Each multi-engine Category B rotorcraft that meets the powerplant installation requirements for Category A must meet the requirements of:

(1)CS 29.79 and 29.81; or

(2)Sub-paragraph (a).

(c)It must be possible to make a safe landing on a prepared landing surface if complete power failure occurs during normal cruise.

ED Decision 2003/16/RM

For Category A rotorcraft, the balked landing path with the critical engine inoperative must be established so that:

(a)The transition from each stage of the manoeuvre to the next stage can be made smoothly and safely;

(b)From the LDP on the approach path selected by the applicant, a safe climbout can be made at speeds allowing compliance with the climb requirements of CS 29.67(a)(1) and (2); and

(c)The rotorcraft does not descend below 4.6 m (15 ft) above the landing surface. For elevated heliport operations, descent may be below the level of the landing surface provided the deck edge clearance of CS 29.60 is maintained and the descent (loss of height) below the landing surface is determined.

ED Decision 2003/16/RM

(a)If there is any combination of height and forward velocity (including hover) under which a safe landing cannot be made after failure of the critical engine and with the remaining engines (where applicable) operating within approved limits, a height-velocity envelope must be established for:

(1)All combinations of pressure altitude and ambient temperature for which take-off and landing are approved; and

(2)Weight, from the maximum weight (at sea-level) to the highest weight approved for take-off and landing at each altitude. For helicopters, this weight need not exceed the highest weight allowing hovering out of ground effect at each altitude.

(b)For single engine or multi-engine rotorcraft that do not meet the Category A engine isolation requirements, the height-velocity envelope for complete power failure must be established.

FLIGHT CHARACTERISTICS

ED Decision 2003/16/RM

The rotorcraft must:

(a)Except as specifically required in the applicable paragraph, meet the flight characteristics requirements of this Subpart:

(1)At the approved operating altitudes and temperatures;

(2)Under any critical loading condition within the range of weights and centres of gravity for which certification is requested; and

(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;

(2)Sudden, complete power failure, for other rotorcraft; and

(3)Sudden, complete control system failures specified in CS 29.695; and

(c)Have any additional characteristics 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.

ED Decision 2007/014/R

I.General. A large 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.

(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 sub-paragraphs IV (b) to (f) 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. The airspeed must return to within 10% of the trim speed when the control force is slowly released for each trim condition specified in sub-paragraphs IV (b) to (f) of this appendix.

(b)Climb. Stability must be shown in climb throughout the speed range 37 km/h (20 knots) either side of trim with:

(1)The helicopter trimmed at VYI;

(2)Landing gear retracted (if retractable); and

(3)Power required for limit climb rate (at least 5.1 m/s (1000 fpm)) at VYI or maximum continuous power, whichever is less.

(c)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:

(1)The helicopter trimmed and power adjusted for level flight at 0.9 VH or 0.9 VNEI, whichever is lower; and

(2)Landing gear retracted (if retractable).

(d)Slow cruise. Stability must be shown throughout the speed range from 0.9 VMINI to 1.3VMINI or 37 km/h (20 knots) above trim speed, whichever is greater, with:

(1)The helicopter trimmed and power adjusted for level flight at 1.1 VMINI; and

(2)Landing gear retracted (if retractable).

(e)Descent. Stability must be shown throughout the speed range 37 km/h (20 knots) either side of trim with:

(1)The helicopter trimmed at 0.8 VH or 0.8 VNEI (or 0.8 VLE for the landing gear extended case), whichever is lower;

(2)Power required for 5.1 m/s (1000 fpm) descent at trim speed; and

(3)Landing gear extended and retracted, if applicable.

(f)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:

(1)The helicopter trimmed at the recommended approach speed or speeds;

(2)Landing gear extended and retracted, if applicable; and

(3)Power required to maintain a 3° glide path and power required to maintain the steepest approach gradient for which approval is requested.

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)Any oscillation having a period of less than 5 seconds must damp to ½ amplitude in not more than one cycle.

(b)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.

(c)Any oscillation having a period of 10 seconds or more but less than 20 seconds must be damped.

(d)Any oscillation having a period of 20 seconds or more may not achieve double amplitude in less than 20 seconds.

(e)Any aperiodic response may not achieve double amplitude in less than 9 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 that 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-29must 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;

(iii)For Category A helicopters, the dynamic stability requirements of Subpart B of CS-29 must also be met throughout a practical flight envelope; and

(iv)The static longitudinal and static directional stability requirements of Subpart B of CS-29 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 Subpart F of CS-29 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). If standby batteries are provided, they may be charged from the aircraft electrical system if adequate isolation is incorporated. The system must be designed so that the standby batteries may not be used for engine starting.

(b)Miscellaneous requirements

(1)Instrument systems and other systems essential for IFR flight that could be adversely affected by icing must be provided with adequate ice protection 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 crew member action, after any single failure or combination of failures that is not shown to be extremely improbable; and

(iv)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.

(6)In determining compliance with the requirements of CS 29.1351(d)(2), the supply of electrical power to all systems necessary for flight under IFR must be included in the evaluation.

(c)Thunderstorm lights. In addition to the instrument lights required by CS 29.1381(a), thunderstorm lights which provide high intensity white flood lighting to the basic flight instruments must be provided. The thunderstorm lights must be installed to meet the requirements of CS 29.1381(b).

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 flightcrew 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.: 29/1]

ED Decision 2008/010/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; and

(vi)Landing (power on and power off).

(b)The margin of cyclic control must allow satisfactory roll and pitch control a 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), with:

(1)Critical weight;

(2)Critical centre of gravity;

(3)Critical rotor rpm; and

(4)Altitude from standard sea-level conditions to the maximum take-off and landing altitude capability of the rotorcraft.

(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 failure of one engine, in the case of multi-engine rotorcraft that meet Category A engine isolation requirements, or complete power 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)For the cruise condition, one second, or normal pilot reaction time (whichever is greater); and

(2)For any other condition, normal pilot reaction time.

(f)For helicopters for which a VNE (power-off) is established under CS 29.1505(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.: 29/1]

[Amdt. No. 29/2]

ED Decision 2003/16/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 2003/16/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 2003/16/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.

ED Decision 2007/014/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 29.175(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.: 29/1]

ED Decision 2007/014/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 0.8 VNE - 19 km/h (10 knots) to 0.8 VNE + 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 – 37 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.: 29/1]

ED Decision 2007/014/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 29.175(a), (b), (c) and (d). 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 29.351;

(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 approaching 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.: 29/1]

ED Decision 2003/16/RM

Any short period oscillation occurring at any speed from VY to VNE must be positively damped with the primary flight controls free and in a fixed position.

GROUND AND WATER HANDLING CHARACTERISTICS

ED Decision 2003/16/RM

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

ED Decision 2003/16/RM

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.

ED Decision 2003/16/RM

If certification for water operation is requested, no spray characteristics during taxying, take-off, or landing may obscure the vision of the pilot or damage the rotors, propellers, or other parts of the rotorcraft.

ED Decision 2003/16/RM

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

MISCELLANEOUS FLIGHT REQUIREMENTS

ED Decision 2003/16/RM

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

ED Decision 2023/001/R

This AMC supplements FAA AC 29-2C, § AC 29.251 and should be used in conjunction with that AC when demonstrating compliance with CS 29.251.

The applicant should investigate each individual installation of the rotorcraft for compliance with CS 29.251. The absence of coupling with the rotors vibration frequencies has to be demonstrated by a combination of analysis, vibration and flight tests.

Qualitative and quantitative flight tests should be performed depending on the extent of the change. For any installation, the failure of which or its attachment would have a catastrophic consequence, a fatigue evaluation should be performed when the vibrations are likely to affect the fatigue strength.

[Amdt No: 29/11]