Part V - Airworthiness Manual Chapter 529 - Transport Category Rotorcraft

Canadian Aviation Regulations (CARs) 2017-2

Preamble

SUBCHAPTERS 

  • A (529.1-529.20), 
  • B (529.21-529.300), 
  • C (529.301-529.600), 
  • D (529.601-529.900), 
  • E (529.901-529.1300), 
  • F (529.1301-529.1500), 
  • G (529.1501-529.1589)

APPENDICES  

ABCDE

SUBCHAPTER C STRENGTH REQUIREMENTS - GENERAL

529.301 Loads

  1. (a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads.
  2. (b) Unless otherwise provided, the specified air, ground, and water loads shall be placed in equilibrium with inertia forces, considering each item of mass in the rotorcraft. These loads shall be distributed to closely approximate or conservatively represent actual conditions.
  3. (c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution shall be taken into account.

529.302 Reserved

529.303 Factor of Safety

Unless otherwise provided, a factor of safety of 1.5 shall be used. This factor applies to external and inertia loads unless its application to the resulting internal stresses is more conservative.

529.304 Reserved

529.305 Strength and Deformation

  1. (a) The structure shall be able to support limit loads without detrimental or permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation.
  2. (b) The structure shall be able to support ultimate loads without failure. This shall be demonstrated by:
    1. (1) applying ultimate loads to the structure in a static test for at least 3 seconds; or
    2. (2) dynamic tests simulating actual load application.

529.306 Reserved

529.307 Proof of Structure

  1. (a) Compliance with the strength and deformation requirements of this subchapter shall be demonstrated for each critical loading condition accounting for the environment to which the structure will be exposed in operation. Structural analysis (static or fatigue) may be used only if the structure conforms to those for which experience has demonstrated this method to be reliable. In other cases, substantiating load tests shall be made.
  2. (b) Proof of compliance with the strength requirements of this subchapter shall include:
    1. (1) dynamic and endurance tests of rotors, rotor drives, and rotor controls;
    2. (2) limit load tests of the control system, including control surfaces;
    3. (3) operation tests of the control system;
    4. (4) flight stress measurement tests;
    5. (5) landing gear drop tests; and
    6. (6) any additional tests required for new or unusual design features.

529.308 Reserved

529.309 Design Limitations

The following values and limitations shall be established to demonstrate compliance with the structural requirements of this subchapter:

  1. (a) the design maximum and design minimum weights;
  2. (b) the main rotor r.p.m. ranges, power-on and power-off;
  3. (c) the maximum forward speeds for each main rotor r.p.m. within the ranges determined under (b) of this section;
  4. (d) the maximum rearward and sideward flight speeds;
  5. (e) the centre of gravity limits corresponding to the limitations determined under (b), (c), and (d) of this section;
  6. (f) the rotational speed ratios between each powerplant and each connected rotating component; and
  7. (g) the positive and negative limit manoeuvring load factors.

529.310 to 529.320 Reserved

Flight Loads

529.321 General

  1. (a) The flight load factor shall be assumed to act normal to the longitudinal axis of the rotorcraft, and to be equal in magnitude and opposite in direction to the rotorcraft inertia load factor at the centre of gravity.
  2. (b) Compliance with the flight load requirements of this subchapter: shall be demonstrated:
    1. (1) at each weight from the design minimum weight to the design maximum weight; and
    2. (2) with any practical distribution of disposable load within the operating limitations in the Rotorcraft Flight Manual.

529.322 to 529.336 Reserved

529.337 Limit Manoeuvring Load Factor

The rotorcraft shall be designed for:

  1. (a) a limit manoeuvring load factor ranging from a positive limit of 3.5 to a negative limit of -1.0; or
  2. (b) any positive limit manoeuvring load factor not less than 2.0 and any negative limit manoeuvring load factor of not less than -0.5, for which:
    1. (1) the probability of being exceeded is demonstrated by analysis and flight tests to be extremely remote; and
    2. (2) the selected values are appropriate to each weight condition between the design maximum and design minimum weights.

529.338 Reserved

529.339 Resultant Limit Manoeuvring Loads

The loads resulting from the application of limit manoeuvring load factors are assumed to act at the centre of each rotor hub and at each auxiliary lifting surface, and to act in directions and with distributions of load among the rotors and auxiliary lifting surfaces, so as to represent each critical manoeuvring condition, including power-on and power-off flight with the maximum design rotor tip speed ratio. The rotor tip speed ratio is the ratio of the rotorcraft flight velocity component in the plane of the rotor disc to the rotational tip speed of the rotor blades, and is expressed as follows:

Where:

V = The airspeed along the flight path (f.p.s.);

a = The angle between the projection, in the plane of symmetry, of the axis of no feathering and a line perpendicular to the flight path (radians, positive when axis is pointing aft);

W =The angular velocity of rotor (radians per second); and

R = The rotor radius (ft.).

529.340 Reserved

529.341 Gust Loads

Each rotorcraft shall be designed to withstand, at each critical airspeed including hovering, the loads resulting from vertical and horizontal gusts of 30 feet per second.

529.342 to 529.350 Reserved

529.351 Yawing Conditions

  1. (a) Each rotorcraft shall be designed for the loads resulting from the manoeuvres specified in (b) and (c) of this section, with:
    1. (1) unbalanced aerodynamic moments about the centre of gravity which the aircraft reacts to in a rational or conservative manner considering the principal masses furnishing the reacting inertia forces; and
    2. (2) maximum main rotor speed.
  2. (b) To produce the load required in (a) of this section, in unaccelerated flight with zero yaw, at forward speeds from zero up to 0.6 VNE:
    1. (1) displace the cockpit directional control suddenly to the maximum deflection limited by the control stops or by the maximum pilot force specified in section 529.397 (a);
      (amended 1998/10/29)
    2. (2) attain a resulting sideslip angles or 90°, whichever is less; and
    3. (3) return the directional control suddenly to neutral.
  3. (c) To produce the load required in (a) of this section, in unaccelerated flight with zero yaw, at forward speeds from 0.6 VNE up to VNE or VH, whichever is less:
    1. (1) displace the cockpit directional control suddenly to the maximum deflection limited by the control stops or by the maximum pilot force specified in section 529.397 (a);
      (amended 1998/10/29)
    2. (2) attain a resulting sideslip angle or 15°, whichever is less, at the lesser speed of VNE or VH;
    3. (3) vary the sideslip angles of (b)(2) and (c)(2) of this section directly with speed; and
    4. (4) return the directional control suddenly to neutral.

529.352 to 529.360 Reserved

529.361 Engine Torque

The limit engine torque may not be less than the following:

  1. (a) for turbine engines, the highest of:
    1. (1) the mean torque for maximum continuous power multiplied by 1.25;
    2. (2) the torque required by section 529.923;
    3. (3) the torque required by section 529.927; or
    4. (4) the torque imposed by sudden engine stoppage due to malfunction or structural failure (such as compressor jamming);
  2. (b) for reciprocating engines, the mean torque for maximum continuous power multiplied by:
    1. (1) 1.33, for engines with five or more cylinders; and
    2. (2) two, three, and four, for engines with four, three, and two cylinders, respectively.

529.362 to 529.390 Reserved

Control Surface and System Loads

529.391 General

Each auxiliary rotor, each fixed or movable stabilizing or control surface, and each system operating any flight control shall meet the requirements of sections 529.395 through 529.399529.411, and 529.427.
(amended 1998/10/29)

529.392 to 529.394 Reserved

529.395 Control System

  1. (a) The reaction to the loads prescribed in section 529.397 shall be provided by:
    1. (1) the control stops only;
    2. (2) the control locks only;
    3. (3) the irreversible mechanism only (with the mechanism locked and with the control surface in the critical positions for the effective parts of the system within its limit of motion);
    4. (4) the attachment of the control system to the rotor blade pitch control horn only (with the control in the critical positions for the affected parts of the system within the limits of its motion); and
    5. (5) the attachment of the control system to the control surface horn (with the control in the critical positions for the affected parts of the system within the limits of its motion).
  2. (b) Each primary control system, including its supporting structure, shall be designed as follows:
    1. (1) the system shall withstand loads resulting from the limit pilot forces prescribed in section 529.397;
    2. (2) notwithstanding (b)(3) of this section, when power-operated actuator controls or power boost controls are used, the system shall also withstand the loads resulting from the limit pilot forces prescribed in section 529.397 in conjunction with the forces output of each normally energized power device, including any single power boost or actuator system failure;
    3. (3) if the system design or the normal operating loads are such that a part of the system cannot react to the limit pilot forces prescribed in section 529.397, that part of the system shall be designed to withstand the maximum loads that can be obtained in normal operation. The minimum design loads shall, in any case, provide a rugged system for service use, including consideration of fatigue, jamming, ground gusts, control inertia, and friction loads. In the absence of a rational analysis, the design loads resulting from 0.60 of the specified limit pilot forces are acceptable minimum design loads; and
    4. (4) if operational loads may be exceeded through jamming, ground gusts, control inertia, or friction, the system shall withstand the limit pilot forces specified in section 529.397, without yielding.

529.396 Reserved

529.397 Limit Pilot Forces and Torques

  1. (a) Except as provided in (b) of this section, the limit pilot forces are as follows:
    1. (1) for foot controls, 130 pounds; and
    2. (2) for stick controls, 100 pounds fore and aft, and 67 pounds laterally.
  2. (b) For flap, tab, stabilizer, rotor brake, and landing gear operating controls, the following apply (R = radius in inches):
    1. (1) crank, wheel, and lever controls, (1 + R)/3 x 50 pounds, but not less than 50 pounds nor more than 100 pounds for hand operated controls or 130 pounds for foot operated controls, applied at any angle within 20 degrees of the plane of motion of the control; and
    2. (2) twist controls, 80R inch-pounds.

529.398 Reserved

529.399 Dual Control System

Each dual primary flight control system shall be able to withstand the loads that result when pilot forces not less than 0.75 times those obtained under section 529.395 are applied:

  1. (a) in opposition; and
  2. (b) in the same direction.

529.400 Reserved

529.401 Removed

529.402 Reserved

529.403 Removed

529.404 to 529.410 Reserved

529.411 Ground Clearance: Tail Rotor Guard

  1. (a) It shall be impossible for the tail rotor to contact the landing surface during a normal landing.
  2. (b) If a tail rotor guard is required to demonstrate compliance with (a) of this section:
    1. (1) suitable design loads shall be established for the guard; and
    2. (2) the guard and its supporting structure shall be designed to withstand those loads.

529.412 Reserved

529.413 Removed

529.414 to 529.426 Reserved

529.427 Unsymmetrical Loads

  1. (a) Horizontal tail surfaces and their supporting structure shall be designed for unsymmetrical loads arising from yawing and rotor wake effects in combination with the prescribed flight conditions.
  2. (b) To meet the design criteria of (a) of this section, in the absence of more rational data, both of the following shall be met:
    1. (1) one hundred percent of the maximum loading from the symmetrical flight conditions acts on the surface on one side of the plane of symmetry, and no loading acts on the other side; and
    2. (2) fifty percent of the maximum loading from the symmetrical flight conditions acts on the surface on each side of the plane of symmetry but in opposite directions.
  3. (c) For empennage arrangements where the horizontal tail surfaces are supported by the vertical tail surfaces, the vertical tail surfaces and supporting structure shall be designed for the combined vertical and horizontal surface loads resulting from each prescribed flight condition, considered separately. The flight conditions shall be selected so that the maximum design loads are obtained on each surface. In the absence of more rational data, the unsymmetrical horizontal tail surface loading distributions described in this section must be assumed.

529.428 to 529.470 Reserved

Ground Loads

529.471 General

  1. (a) Loads and equilibrium. For limit ground loads:
    1. (1) the limit ground loads obtained in the landing conditions in this Chapter shall be considered to be external loads that would occur in the rotorcraft structure if it were acting as a rigid body; and
    2. (2) in each specified landing condition, the external loads shall be placed in equilibrium with linear and angular inertia loads in a rational or conservative manner.
  2. (b) Critical centres of gravity. The critical centres of gravity within the range for which certification is requested shall be selected so that the maximum design loads are obtained in each landing gear element.

529.472 Reserved

529.473 Ground Loading Conditions and Assumptions

  1. (a) For specified landing conditions, a design maximum weight shall be used that is not less than the maximum weight. A rotor lift may be assumed to act through the centre of gravity throughout the landing impact. This lift may not exceed two-thirds of the design maximum weight.
  2. (b) Unless otherwise prescribed, for each specified landing condition, the rotorcraft shall be designed for a limit load factor of not less than the limit inertia load factor substantiated under section 529.725.
  3. (c) Triggering or actuating devices for additional or supplementary energy absorption may not fail under loads established in the tests prescribed in sections 529.725 and 529.727, but the factor of safety prescribed in section 529.303 need not be used.

529.474 Reserved

529.475 Tires and Shock Absorbers

Unless otherwise prescribed, for each specified landing condition, the tires shall be assumed to be in their static position and the shock absorbers to be in their most critical position.

529.476 Reserved

529.477 Landing Gear Arrangement

Sections 529.235529.479 through 529.485, and 529.493 apply to landing gear with two wheels aft, and one or more wheels forward, of the centre of gravity.

529.478 Reserved

529.479 Level Landing Conditions

  1. (a) Attitudes. Under each of the loading conditions prescribed in (b) of this section, the rotorcraft is assumed to be in each of the following level landing attitudes:
    1. (1) an attitude in which each wheel contacts the ground simultaneously; and
    2. (2) an attitude in which the aft wheels contact the ground with the forward wheels just clear of the ground.
  2. (b) Loading conditions. The rotorcraft shall be designed for the following landing loading conditions:
    1. (1) vertical loads applied under section 529.471;
    2. (2) the loads resulting from a combination of the loads applied under (b)(1) of this section with drag loads at each wheel of not less than 25 percent of the vertical load at that wheel;
    3. (3) the vertical load at the instant of peak drag load combined with a drag component simulating the forces required to accelerate the wheel rolling assembly up to the specified ground speed, with:
      1. (i) the ground speed for determination of the spin-up loads being at least 75 percent of the optimum forward flight speed for minimum rate of descent in autorotation, and
      2. (ii) the loading conditions of (b) (3) of this section applied to the landing gear and its attaching structure only;
    4. (4) if there are two wheels forward, a distribution of the loads applied to those wheels under (b)(1) and (b)(2) of this section in a ratio of 40:60.
  3. (c) Pitching moments. Pitching moments are assumed to be resisted by:
    1. (1) in the case of the attitude in (a)(1) of this section, the forward landing gear; and
    2. (2) in the case of the attitude in (a)(2) of this section, the angular inertia forces.

529.480 Reserved

529.481 Tail-down Landing Conditions

  1. (a) The rotorcraft is assumed to be in the maximum nose-up attitude allowing ground clearance by each part of the rotorcraft.
  2. (b) In this attitude, ground loads are assumed to act perpendicular to the ground.

529.482 Reserved

529.483 One-wheel Landing Conditions

For the one-wheel landing condition, the rotorcraft is assumed to be in the level attitude and to contact the ground on one aft wheel. In this attitude:

  1. (a) the vertical load shall be the same as that obtained on that side under section 529.479(b)(1); and
  2. (b) the unbalanced external loads shall be reacted by rotorcraft inertia.

529.484 Reserved

529.485 Lateral Drift Landing Conditions

  1. (a) The rotorcraft is assumed to be in the level landing attitude, with:
    1. (1) side loads combined with one half of the maximum ground reactions obtained in the level landing conditions of section 529.479 (b)(1); and
    2. (2) the loads obtained under (a)(1) of this section applied:
      1. (i) at the ground contact point, or
      2. (ii) for full-swivelling gear, at the centre of the axle.
  2. (b) The rotorcraft shall be designed to withstand, at ground contact:
    1. (1) when only the aft wheels contact the ground, side loads of 0.8 times the vertical reaction acting inward on one side and 0.6 times the vertical reaction acting outward on the other side, all combined with the vertical loads specified in (a) of this section; and
    2. (2) when the wheels contact the ground simultaneously:
      1. (i) for the aft wheels, the side loads specified in (b)(1) of this section, and
      2. (ii) for the forward wheels, a side load of 0.8 times the vertical reaction combined with the vertical load specified in (a) of this section.

529.486 to 529.492 Reserved

529.493 Braked Roll Conditions

Under braked roll conditions with the shock absorbers in their static positions:

  1. (a) the limit vertical load shall be based on a load factor of at least:
    1. (1) 1.33, for the attitude specified in section 529.479 (a)(1); and
    2. (2) 1.0, for the attitude specified in section 529.479 (a)(2);
  2. (b) the structure shall be designed to withstand, at the ground contact point of each wheel with brakes, a drag load of at least the lesser of:
    1. (1) the vertical load multiplied by a coefficient of friction of 0.8; and
    2. (2) the maximum value based on limiting brake torque.

529.494 to 529.496 Reserved

529.497 Ground Loading Conditions: Landing Gear with Tail Wheels

  1. (a) General. Rotorcraft with landing gear with two wheels forward and one wheel aft of the centre of gravity shall be designed for loading conditions as prescribed in this section.
  2. (b) Level landing attitude with only the forward wheels contacting the ground. In this attitude:
    1. (1) the vertical loads shall be applied under sections 529.471 through 529.475;
    2. (2) the vertical load at each axle shall be combined with a drag load at that axle of not less than 25 percent of that vertical load; and
    3. (3) unbalanced pitching moments are assumed to be resisted by angular inertia forces.
  3. (c) Level landing attitude with all wheels contacting the ground simultaneously. In this attitude, the rotorcraft shall be designed for landing loading conditions as prescribed in (b) of this section.
  4. (d) Maximum nose-up attitude with only the rear wheel contacting the ground. The attitude for this condition shall be the maximum nose-up attitude expected in normal operation, including autorotative landings. In this attitude:
    1. (1) the appropriate ground loads specified in (b)(1) and (b)(2) of this section shall be determined and applied, using a rational method to account for the moment arm between the rear wheel ground reaction and the rotorcraft centre of gravity; or
    2. (2) the probability of landing with initial contact on the rear wheel shall be demonstrated to be extremely remote.
  5. (e) Level landing attitude with only one forward wheel contacting the ground. In this attitude, the rotorcraft shall be designed for ground loads as specified in (b)(1) and (b)(3) of this section.
  6. (f) Side loads in the level landing attitude. In the attitudes specified in (b) and (c) of this section, the following apply:
    1. (1) the side loads shall be combined at each wheel with one-half of the maximum vertical ground reactions obtained for that wheel under (b) and (c) of this section. In this condition, the side loads shall be:
      1. (i) for the forward wheels, 0.8 times the vertical reaction (on one side) acting inward, and 0.6 times the vertical reaction (on the other side) acting outward, and
      2. (ii) for the rear wheel, 0.8 times the vertical reaction;
    2. (2) the loads specified in (f)(1) of this section shall be applied:
      1. (i) at the ground contact point with the wheel in the trailing position (for non-full swivelling landing gear or for full swivelling landing gear with a lock, steering device, or shimmy damper to keep the wheel in the trailing position), or
      2. (ii) at the centre of the axle (for full swivelling landing gear without a lock, steering device, or shimmy damper).
  7. (g) Braked roll conditions in the level landing attitude. In the attitudes specified in (b) and (c) of this section, and with the shock absorbers in their static positions, the rotorcraft shall be designed for braked roll loads as follows:
    1. (1) the limit vertical load shall be based on a limit vertical load factor of not less than:
      1. (i) 1.0, for the attitude specified in (b) of this section, and
      2. (ii) 1.33, for the attitude specified in (c) of this section;
    2. (2) For each wheel with brakes, a drag load shall be applied, at the ground contact point, of not less than the lesser of:
      1. (i) 0.8 times the vertical load, and
      2. (ii) the maximum based on limiting brake torque.
  8. (h) Rear wheel turning loads in the static ground attitude. In the static ground attitude, and with the shock absorbers and tires in their static positions, the rotorcraft shall be designed for rear wheel turning loads as follows:
    1. (1) a vertical ground reaction equal to the static load on the rear wheel shall be combined with an equal side load;
    2. (2) the load specified in (h)(1) of this section shall be applied to the rear landing gear:
      1. (i) through the axle, if there is a swivel (the rear wheel being assumed to be swivelled 90° to the longitudinal axis of the rotorcraft), or
      2. (ii) at the ground contact point if there is a lock, steering device or shimmy damper (the rear wheel being assumed to be in the trailing position).
  9. (i) Taxiing condition. The rotorcraft and its landing gear shall be designed for the loads that would occur when the rotorcraft is taxied over the roughest ground that may reasonably be expected in normal operation.

529.498 to 529.500 Reserved

529.501 Ground Loading Conditions: Landing Gear with Skids

  1. (a) General. Rotorcraft with landing gear with skids shall be designed for the loading conditions specified in this section. In demonstrating compliance with this section, the following apply:
    1. (1) the design maximum weight, centre of gravity, and load factor shall be determined under sections 529.471 through 529.475;
    2. (2) structural yielding of elastic spring members under limit loads is acceptable;
    3. (3) design ultimate loads for elastic spring members need not exceed those obtained in a drop test of the gear with:
      1. (i) a drop height of 1.5 times that specified in section 529.725, and
      2. (ii) an assumed rotor lift of not more than 1.5 times that used in the limit drop tests prescribed in section 529.725;
    4. (4) compliance with (b) through (e) of this section shall be demonstrated with:
      1. (i) the gear in its most critically deflected position for the landing condition being considered, and
      2. (ii) the ground reactions rationally distributed along the bottom of the skid tube.
  2. (b) Vertical reactions in the level landing attitude. In the level attitude, and with the rotorcraft contacting the ground along the bottom of both skids, the vertical reactions shall be applied as prescribed in (a) of this section.
  3. (c) Drag reactions in the level landing attitude. In the level attitude, and with the rotorcraft contacting the ground along the bottom of both skids, the following apply:
    1. (1) the vertical reactions shall be combined with horizontal drag reactions of 50 percent of the vertical reaction applied at the ground; and
    2. (2) the resultant ground loads shall equal the vertical load specified in (b) of this section.
  4. (d) Sideloads in the level landing attitude. In the level attitude, and with the rotorcraft contacting the ground along the bottom of both skids, the following apply:
    1. (1) the vertical ground reaction shall be:
      1. (i) equal to the vertical loads obtained in the condition specified in (b) of this section, and
      2. (ii) divided equally among the skids;
    2. (2) the vertical ground reactions shall be combined with a horizontal sideload of 25 percent of their value;
    3. (3) the total sideload shall be applied equally between skids and along the length of the skids;
    4. (4) the unbalanced moments are assumed to be resisted by angular inertia;
    5. (5) the skid gear shall be investigated for:
      1. (i) inward acting sideloads, and
      2. (ii) outward acting sideloads.
  5. (e) One-skid landing loads in the level attitude. In the level attitude, and with the rotorcraft contacting the ground along the bottom of one skid only, the following apply:
    1. (1) the vertical load on the ground contact side shall be the same as that obtained on that side in the condition specified in (b) of this section; and
    2. (2) the unbalanced moments are assumed to be resisted by angular inertia.
  6. (f) Special conditions. In addition to the conditions specified in (b) and (c) of this section, the rotorcraft shall be designed for the following ground reactions:
    1. (1) a ground reaction load acting up and aft at an angle of 45° to the longitudinal axis of the rotorcraft. This load shall be:
      1. (i) equal to 1.33 times the maximum weight,
      2. (ii) distributed symmetrically among the skids,
      3. (iii) concentrated at the forward end of the straight part of the skid tube, and
      4. (iv) applied only to the forward end of the skid tube and its attachment to the rotorcraft;
    2. (2) with the rotorcraft in the level landing attitude, a vertical ground reaction load equal to one-half of the vertical load determined under (b) of this section. This load shall be:
      1. (i) applied only to the skid tube and its attachment to the rotorcraft, and
      2. (ii) distributed equally over 33.3 percent of the length between the skid tube attachments and centrally located midway between the skid tube attachments.

529.502 to 529.504 Reserved

529.505 Ski Landing Conditions

If certification for ski operation is requested, the rotorcraft, with skis, shall be designed to withstand the following loading conditions (where P is the maximum static weight on each ski with the rotorcraft at design maximum weight, and n is the limit load factor determined under section 529.473(b):

  1. (a) up-load conditions in which:
    1. (1) a vertical load of Pn and a horizontal load of Pn/4 are simultaneously applied at the pedestal bearings; and
    2. (2) a vertical load of 1.33 P is applied at the pedestal bearings;
  2. (b) a side load condition in which a side load of 0.35 Pn is applied at the pedestal bearings in a horizontal plane perpendicular to the centreline of the rotorcraft; and
  3. (c) a torque-load condition in which a torque load of 1.33 P (in foot-pounds) is applied to the ski about the vertical axis through the centreline of the pedestal bearings.

529.506 to 529.510 Reserved

529.511 Ground Load: Unsymmetrical Loads on Multiple-Wheel Units

  1. (a) In dual-wheel gear units, 60 percent of the total ground reaction for the gear unit shall be applied to one wheel and 40 percent to the other.
  2. (b) To provide for the case of one deflated tire, 60 percent of the specified load for the gear unit shall be applied to either wheel except that the vertical ground reaction may not be less than the full static value.
  3. (c) In determining the total load on a gear unit, the transverse shift in the load centroid, due to unsymmetrical load distribution on the wheels, may be neglected.

529.512 to 529.518 Reserved

Water Loads

529.519 Hull Type Rotorcraft: Water-based and Amphibian

  1. (a) General. For hull type rotorcraft, the structure shall be designed to withstand the water loading set forth in (b), (c), and (d) of this section considering the most severe wave heights and profiles for which approval is desired. The loads for the landing conditions of (b) and (c) of this section shall be developed and distributed along and among the hull and auxiliary floats, if used, in a rational and conservative manner, assuming a rotor lift not exceeding two-thirds of the rotorcraft weight to act throughout the landing impact.
  2. (b) Vertical landing conditions. The rotorcraft shall initially contact the most critical wave surface at zero forward speed in likely pitch and roll attitudes which result in critical design loadings. The vertical descent velocity may not be less than 6.5 f.p.s. relative to the mean water surface.
  3. (c) Forward speed landing conditions. The rotorcraft shall contact the most critical wave at forward velocities from 0 up to 30 knots in likely pitch, roll, and yaw attitudes and with a vertical descent velocity of not less than 6.5 f.p.s. relative to the mean water surface. A maximum forward velocity of less than 30 knots may be used in design if it can be demonstrated that the forward velocity selected would not be exceeded in a normal one-engine-out landing.
  4. (d) Auxiliary float immersion condition. In addition to the loads from the landing conditions, the auxiliary float and its support and attaching structure in the hull, shall be designed for the load developed by a fully immersed float unless it can be demonstrated that full immersion of the float is unlikely, in which case the highest likely float buoyancy load shall be applied that considers loading of the float immersed to create restoring moments compensating for upsetting moments caused by side wind, asymmetrical rotorcraft loading, water wave action, and rotorcraft inertia.

529.520 Reserved

529.521 Float Landing Conditions

If certification for float operation (including float amphibian operation) is requested, the rotorcraft, with floats, shall be designed to withstand the following loading conditions (where the limit load factor is determined under section 529.473(b) or assumed to be equal to that determined for wheel landing gear):

  1. (a) up-load conditions in which:
    1. (1) a load is applied so that, with the rotorcraft in the static level attitude, the resultant water reaction passes vertically through the centre of gravity; and
    2. (2) the vertical load prescribed in (a)(1) of this section is applied simultaneously with an aft component of 0.25 times the vertical component;
  2. (b) a side load condition in which:
    1. (1) a vertical load of 0.75 times the total vertical load specified in (a)(1) of this section is divided equally among the floats; and
    2. (2) for each float, the load share determined under (b)(1) of this section combined with a total side load of 0.25 times the total vertical load specified in (b)(1) of this section, is applied to that float only.

529.522 to 529.546 Reserved

Main Component Requirements

529.547 Main and Tail Rotor Structure

(amended 1997/04/07)

  1. (a) A rotor is an assembly of rotating components, which includes the rotor, hub, blades, blade dampers, the pitch control mechanisms, and all other parts that rotate with the assembly.
    (amended 1997/04/07)
  2. (b) Each rotor assembly shall be designed as prescribed in this section and shall function safely for the critical flight load and operating conditions. A design assessment shall be performed, including a detailed failure analysis to identify all failures that will prevent continued safe flight or safe landing, and shall identify the means to minimize the likelihood of their occurrence.
    (amended 1997/04/07)
  3. (c) The rotor structure shall be designed to withstand the following loads prescribed in sections 529.337 through 529.341, and 529.351:
    (amended 1997/04/07)
    1. (1) critical flight loads;
    2. (2) limit loads occurring under normal conditions of autorotation.
  4. (d) The rotor structure shall be designed to withstand loads simulating:
    (amended 1997/04/07)
    1. (1) for the rotor blades, hubs, and flapping hinges, the impact force of each blade against its stop during ground operation; and
    2. (2) any other critical condition expected in normal operation.
  5. (e) The rotor structure shall be designed to withstand the limit torque at any rotational speed, including zero. In addition:
    (amended 1997/04/07)
    1. (1) the limit torque need not be greater than the torque defined by a torque limiting device (where provided), and may not be less than the greater of:
      1. (i) the maximum torque likely to be transmitted to the rotor structure, in either direction, by the rotor drive or by sudden application of the rotor brake, and
      2. (ii) for the main rotor, the limit engine torque specified in section 529.361; and
        (amended 1997/04/07)
    2. (2) the limit torque shall be equally and rationally distributed to the rotor blades.

529.548 Reserved

529.549 Fuselage and Rotor Pylon Structures

  1. (a) Each fuselage and rotor pylon structure shall be designed to withstand:
    1. (1) the critical loads prescribed in sections 529.337 through 529.341, and 529.351;
    2. (2) the applicable ground loads prescribed in sections 529.235529.471 through 529.485, 529.493529.497529.505, and 529.521; and
    3. (3) the loads prescribed in section 529.547 (d)(1) and (e)(1)(i).
  2. (b) Auxiliary rotor thrust, the torque reaction of each rotor drive system, and the balancing air and inertia loads occurring under accelerated flight conditions, shall be considered.
  3. (c) Each engine mount and adjacent fuselage structure shall be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque.
  4. (d) Removed
  5. (e) If approval for the use of 2 1/2-minute OEI power is requested, each engine mount and adjacent structure shall be designed to withstand the loads resulting from a limit torque equal to 1.25 times the mean torque for 2 1/2-minute OEI power combined with 1 g flight loads.

529.550 Reserved

529.551 Auxiliary Lifting Surfaces

Each auxiliary lifting surface shall be designed to withstand:

  1. (a) the critical flight loads in sections 529.337 through 529.341, and 529.351;
  2. (b) the applicable ground loads in sections 529.235529.471 through 529.485529.493529.505, and 529.521; and
  3. (c) any other critical condition expected in normal operation.

529.552 to 529.660 Reserved

Emergency Landing Conditions

529.561 General

  1. (a) The rotorcraft, although it may be damaged in emergency landing conditions on land or water, shall be designed as prescribed in this section to protect the occupants under those conditions.
  2. (b) The structure shall be designed to give each occupant every reasonable chance of escaping serious injury in a crash landing when:
    1. (1) proper use is made of seats, belts, and other safety design provisions;
    2. (2) the wheels are retracted (where applicable); and
    3. (3) each occupant and each item of mass inside the cabin that could injure an occupant is restrained when subjected to the following ultimate inertial load factors relative to the surrounding structure:
      1. (i) upward - 4g,
      2. (ii) forward - 16g,
      3. (iii) sideward - 8g,
      4. (iv) downward - 20g after the intended displacement of the seat device, and
      5. (v) rearward - 1.5g.
        (amended 1997/04/07)
  3. (c) The supporting structure shall be designed to restrain under any ultimate inertial load factor up to those specified in (c) of this section, any item of mass above and/or behind the crew and passenger compartment that could injure an occupant if it came loose in an emergency landing. Items of mass to be considered include, but are not limited to, rotors, transmission, and engines. The items of mass shall be restrained for the following ultimate inertial load factors:
    1. (1) upward - 1.5g;
    2. (2) forward - 12g;
      (amended 1997/04/07)
    3. (3) sideward - 6g;
      (amended 1997/04/07)
    4. (4) downward -12g; and
      (amended 1997/04/07)
    5. (5) rearward - 1.5g.
      (amended 1997/04/07)
  4. (d) Any fuselage structure in the area of internal fuel tanks below the passenger floor level shall be designed to resist the following ultimate inertial factors and loads, and to protect the fuel tanks from rupture, if rupture is likely when those loads are applied to that area:
    1. (1) upward - 1.5g;
    2. (2) forward - 4.0g;
    3. (3) sideward - 2.0g; and
    4. (4) downward - 4.0g.

529.562 Emergency landing dynamic conditions

  1. (a) The rotorcraft, although it may be damaged in a crash landing, shall be designed to reasonably protect each occupant when:
  2. (1) the occupant properly uses the seats, safety belts, and shoulder harnesses provided in the design; and
  3. (2) the occupant is exposed to loads equivalent to those resulting from the conditions prescribed in this section.
  4. (b) Each seat type design or other seating device approved for crew or passenger occupancy during take-off and landing shall successfully complete dynamic tests or be demonstrated by rational analysis based on dynamic tests of a similar type seat in accordance with the following criteria. The tests shall be conducted with an occupant, simulated by a 170-pound anthropomorphic test dummy (ATD), as defined by 49 CFR 572, Subpart B, or its equivalent, sitting in the normal upright position.
  5. (1) a change in downward velocity of not less than 30 feet per second when the seat or other seating device is oriented in its nominal position with respect to the rotorcraft's reference system, the rotorcraft's longitudinal axis is canted upward 60° with respect to the impact velocity vector, and the rotorcraft's lateral axis is perpendicular to a vertical plane containing the impact velocity vector and the rotorcraft's longitudinal axis. Peak floor deceleration shall occur in not more than 0.031 seconds after impact and shall reach a minimum of 30g's.
  6. (2) a change in forward velocity of not less than 42 feet per second when the seat or other seating device is oriented in its nominal position with respect to the rotorcraft's reference system, the rotorcraft's longitudinal axis is yawed 10° either right or left of the impact velocity vector (whichever would cause the greatest load on the shoulder harness), the rotorcraft's lateral axis is contained in a horizontal plane containing the impact velocity vector, and the rotorcraft's vertical axis is perpendicular to a horizontal plane containing the impact velocity vector. Peak floor deceleration shall occur in not more than 0.071 seconds after impact and shall reach a minimum of 18.4g's.
  7. (3) where floor rails or floor or [sidewall attachment] devices are used to attach the seating devices to the airframe structure for the conditions of this section, the rails or devices shall be misaligned with respect to each other by at least 10° vertically (i.e., pitch out of parallel) and by at least a 10° lateral roll, with the directions optional, to account for possible floor warp.
  8. (c) Compliance with the following shall be demonstrated:
    1. (1) the seating device system shall remain intact although it may experience separation intended as part of its design;
    2. (2) the attachment between the seating device and the airframe structure shall remain intact, although the structure may have exceeded its limit load;
    3. (3) the ATD's shoulder harness strap or straps shall remain on or in the immediate vicinity of the ATD's shoulder during the impact;
    4. (4) the safety belt must remain on the ATD's pelvis during the impact;
    5. (5) the ATD's head either does not contact any portion of the crew or passenger compartment, or if contact is made, the head impact does not exceed a head injury criteria (HIC) of 1,000 as determined by this equation.

      Where:

      a(t) is the resultant acceleration at the centre of gravity of the head form expressed as a multiple of g (the acceleration of gravity) and

      t2 - t1 is the time duration, in seconds, of major head impact, not to exceed 0.05 seconds; and

    6. (6) loads in individual shoulder harness straps shall not exceed 1,750 pounds. If dual straps are used for retaining the upper torso, the total harness strap loads shall not exceed 2,000 pounds;
    7. (7) the maximum compressive load measured between the pelvis and the lumbar column of the ATD shall not exceed 1,500 pounds.
  9. (d) An alternate approach that achieves an equivalent or greater level of occupant protection, as required by this section, shall be substantiated on a rational basis.

529.563 Structural ditching provisions

If certification with ditching provisions is requested, structural strength for ditching shall meet the requirements of this section and section 529.801 (e).

  1. (a) Forward speed landing conditions. The rotorcraft shall initially contact the most critical wave for reasonably probable water conditions at forward velocities from zero up to 30 knots in likely pitch, roll, and yaw attitudes. The rotorcraft limit vertical descent velocity may not be less than 5 feet per second relative to the mean water surface. Rotor lift may be used to act through the centre of gravity throughout the landing impact. This lift may not exceed two-thirds of the design maximum weight. A maximum forward velocity of less than 30 knots may be used in design if it can be demonstrated that the forward velocity selected would not be exceeded in a normal one-engine-out touchdown;
  2. (b) Auxiliary or emergency float conditions:
    1. (1) Floats fixed or deployed before initial water contact. In addition to the landing loads in (a) of this section, each auxiliary or emergency float, or its support and attaching structure in the airframe or fuselage, shall be designed for the load developed by a fully immersed float unless it can be demonstrated that full immersion is unlikely. If full immersion is unlikely, the highest likely float buoyancy load shall be applied. The highest likely buoyancy load shall include consideration of a partially immersed float creating restoring moments to compensate the upsetting moments caused by side wind. Unsymmetrical rotorcraft loading, water wave action, rotorcraft inertia, and probable structural damage and leakage considered under section 529.801(d). Maximum roll and pitch angles determined from compliance with 529.801(d) may be used, if significant, to determine the extent of immersion of each float. If the floats are deployed in flight, appropriate air loads derived from the flight limitations with the floats deployed shall be used in substantiation of the floats and their attachment to the rotorcraft. For this purpose, the design airspeed for limit load is the float deployed airspeed operating limit multiplied by 1.11; and
    2. (2) Floats deployed after initial water contact. Each float shall be designed for full or partial immersion prescribed in (b)(1) of this section. In addition, each float shall be designed for combined vertical and drag loads using a relative limit speed of 20 knots between the rotorcraft and the water. The vertical load may not be less than the highest likely buoyancy load determined under b)(1) of this section.

529.564 to 529.570 Reserved

Fatigue Evaluation

529.571 Fatigue Tolerance Evaluation of Metallic Structures

(effective 2014/07/08)

  1. (a) A fatigue tolerance evaluation (FTE) of each principal structural element (PSE) must be performed and appropriate inspections and retirement time or approved equivalent means must be established to avoid catastrophic failure during the operational life of the rotorcraft. The FTE must consider the effects of both fatigue and the damage determined under (e)(4) of this section. Parts to be evaluated include PSEs of the rotors, rotor drive systems between the engines and rotor hubs, controls, fuselage, fixed and movable control surfaces, engine and transmission mountings, landing gear, and their related primary attachments.
    (effective 2014/07/08)
  2. (b) For the purposes of this section, the terms:
    (effective 2014/07/08)
    1. (1) catastrophic failure – means an event that could prevent continued safe flight and landing;
    2. (2) principal structural element (PSE) – means a principal structural element that contributes significantly to the carriage of flight or ground loads, and the fatigue failure of that structural element could result in catastrophic failure of the aircraft.
  3. (c) The methodology used to establish compliance with this section must be submitted to and approved by the Minister.
    (effective 2014/07/08)
  4. (d) Considering all rotorcraft structures, structural elements and assemblies, each PSE must be identified.
    (effective 2014/07/08)
  5. (e) Each fatigue tolerance evaluation required by this section must include:
    (effective 2014/07/08)
    1. (1) in-flight measurements to determine the fatigue loads or stresses for the PSEs identified in (d) of this section in all critical conditions throughout the range of design limitations required by section 529.309 (including altitude effects), except that manoeuvring load factors need not exceed the maximum values expected in operations;
      (effective 2014/07/08)
    2. (2) the loading spectra as severe as those expected in operation based on loads or stresses determined under (e)(1) of this section, including external load operations, if applicable, and other high frequency power cycle operations;
      (effective 2014/07/08)
    3. (3) take-off, landing and taxi loads when evaluating the landing gear and other affected PSEs;
      (effective 2014/07/08)
    4. (4) for each PSE identified in (d) of this section, a threat assessment which includes a determination of the probable locations, types and sizes of damage, taking into account fatigue, environmental effects, intrinsic and discrete flaws or accidental damage that may occur during manufacture or operation;
      (effective 2014/07/08)
    5. (5) a determination of the fatigue tolerance characteristics for the PSE with the damage identified in (e)(4) of this section that supports the inspection and retirement times or other approved equivalent means; and
      (effective 2014/07/08)
    6. (6) analyses supported by test evidence and, if available, service experience.
      (effective 2014/07/08)
  6. (f) A residual strength determination that substantiates the maximum damage size assumed in the FTE is required. In determining inspection intervals based on damage growth, the residual strength evaluation must show that the remaining structure, after damage growth, is able to withstand design limit loads without failure.
    (effective 2014/07/08)
  7. (g) The effect of damage on stiffness, dynamic behaviour, loads and functional performance must be considered.
    (effective 2014/07/08)
  8. (h) Based on the requirements of this section, inspections and retirement times or approved equivalent means must be established to avoid catastrophic failure. The inspections and retirement times or approved equivalent means must be included in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness required by section 529.1529 and section A529.4 of Appendix A of this Chapter.
    (effective 2014/07/08)
  9. (i) If inspections for any of the damage types identified in (e)(4) of this section cannot be established within the limitations of geometry, inspectability or good design practice, then supplemental procedures, in conjunction with the PSE retirement time, must be established to minimize the risk of occurrence of these types of damage that could result in a catastrophic failure during the operational life of the rotorcraft.
    (effective 2014/07/08)

529.572   Reserved

(effective 2014/07/08)

529.573   Damage Tolerance and Fatigue Evaluation of Composite Rotorcraft Structures 

(effective 2014/07/08)

  1. (a) Each applicant shall evaluate the composite rotorcraft structure under the damage tolerance standards in (d) of this section unless the applicant establishes that a damage tolerance evaluation is impractical within the limits of geometry, inspectability, and good design practice. If an applicant establishes that it is impractical within the limits of geometry, inspectability, and good design practice, the applicant shall do a fatigue evaluation in accordance with (e) of this section.
  2. (b) The methodology used to establish compliance with this section must be submitted to and approved by the Minister.
  3. (c) Definitions:
    1. (1) Catastrophic failure is an event that could prevent continued safe flight and landing.
    2. (2) Principal Structural Elements (PSEs) are structural elements that contribute significantly to the carrying of flight or ground loads, the failure of which could result in catastrophic failure of the rotorcraft.
    3. (3) Threat Assessment is an assessment that specifies the locations, types, and sizes of damage, considering fatigue, environmental effects, intrinsic and discrete flaws, and impact or other accidental damage (including the discrete source of the accidental damage) that may occur during manufacture or operation.
  4. (d) Damage Tolerance Evaluation
    1. (1) Each applicant shall show that catastrophic failure due to static and fatigue loads, considering the intrinsic or discrete manufacturing defects or accidental damage, is avoided throughout the operational life or prescribed inspection intervals of the rotorcraft by performing damage tolerance evaluations of the strength of composite PSEs and other parts, detail design points, and fabrication techniques. Each applicant shall account for the effects of material and process variability along with environmental conditions in the strength and fatigue evaluations. Each applicant shall evaluate parts that include PSEs of the airframe, main and tail rotor drive systems, main and tail rotor blades and hubs, rotor controls, fixed and movable control surfaces, engine and transmission mountings, landing gear, other parts, detail design points, and fabrication techniques deemed critical by the Minister. Each damage tolerance evaluation must include:
      1. (i) the identification of all PSEs;
      2. (ii) in-flight and ground measurements for determining the loads or stresses for all PSEs for all critical conditions throughout the range of limits in 529.309 (including altitude effects), except that manoeuvring load factors need not exceed the maximum values expected in service;
      3. (iii) the loading spectra as severe as those expected in service based on loads or stresses determined in (d)(1)(ii) of this section, including external load operations, if applicable, and other operations including high-torque events;
      4. (iv) a threat assessment for all PSEs that specifies the locations, types, and sizes of damage, considering fatigue, environmental effects, intrinsic and discrete flaws, and impact or other accidental damage (including the discrete source of the accidental damage) that may occur during manufacture or operation; and
      5. (v) an assessment of the residual strength and fatigue characteristics of all PSEs that supports the replacement times and inspection intervals established in (d)(2) of this section.
    2. (2) Each applicant shall establish replacement times, inspections, or other procedures for all PSEs to require the repair or replacement of damaged parts before a catastrophic failure. These replacement times, inspections, or other procedures must be included in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness required by section 529.1529.
      1. (i) Replacement times for PSEs must be determined by tests, or by analysis supported by tests, and must show that the structure is able to withstand the repeated loads of variable magnitude expected in service. In establishing these replacement times, the following items must be considered:
        1. (A) damage identified in the threat assessment required by (d)(1)(iv) of this section;
        2. (B) maximum acceptable manufacturing defects and in-service damage (i.e., those that do not lower the residual strength below ultimate design loads and those that can be repaired to restore ultimate strength); and
        3. (C) ultimate load strength capability after applying repeated loads.
      2. (ii) Inspection intervals for PSEs must be established to reveal any damage identified in the threat assessment required by (d)(1)(iv) of this section that may occur from fatigue or other in-service causes before such damage has grown to the extent that the component cannot sustain the required residual strength capability. In establishing these inspection intervals, the following items must be considered:
        1. (A) the growth rate, including no-growth, of the damage under the repeated loads expected in service determined by tests or analysis supported by tests;
        2. (B) the required residual strength for the assumed damage established after considering the damage type, inspection interval, detectability of damage, and the techniques adopted for damage detection. The minimum required residual strength is limit load; and
        3. (C) whether the inspection will detect the damage growth before the minimum residual strength is reached and restored to ultimate load capability, or whether the component will require replacement.
    3. (3) Each applicant shall consider the effects of damage on stiffness, dynamic behaviour, loads, and functional performance on all PSEs when substantiating the maximum assumed damage size and inspection interval.
  5. (e) Fatigue Evaluation:  If an applicant establishes that the damage tolerance evaluation described in (d) of this section is impractical within the limits of geometry, inspectability, or good design practice, the applicant shall do a fatigue evaluation of the particular composite rotorcraft structure and:
    1. (1) identify all PSEs considered in the fatigue evaluation;
    2. (2) identify the types of damage for all PSEs considered in the fatigue evaluation;
    3. (3) establish supplemental procedures to minimize the risk of catastrophic failure associated with the damages identified in (d) of this section; and
    4. (4) include these supplemental procedures in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness required by section 529.1529.

529.574 to 529.600 Reserved

(effective 2014/07/08)

Date modified: