(ii) Power off.

(iii) Landing gear both retracted and extended, and

(iv) Trim at the speed for minimum rate of descent.

(4) Hovering. In the case of helicopters the stick position curve shall have a stable slope between the maximum approved rearward speed and a forward speed of 20 mph, with:

(i) Critical weight and center of gravity,

(ii) Power required for hovering in still air,

(iii) Landing gear retracted, and
(iv) Trim for hovering.

NOTE: It is considered acceptable for the stick position versus speed curve to have a negative slope within the speed range specified for each of the conditions in subparagraphs (1) through (3) of this paragraph, provided the negative stick displacement required is not greater than 10 percent of the total stick travel.

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-4, 24 F.R. 7073, Sept. 1, 1959]

GROUND AND WATER HANDLING
CHARACTERISTICS

§ 6.130

General.

The rotorcraft shall be demonstrated to have satisfactory ground and water handling characteristics. There shall be no uncontrollable tendencies in any operating condition reasonably expected for the type.

§ 6.131 Ground resonance.

There shall be no uncontrollable tendency for the rotorcraft to oscillate when

the rotor is turning and the rotorcraft is on the ground.

§ 6.132

Spray characteristics.

For rotorcraft equipped with floats, the spray characteristics during taxying take-off, and landing shall be such as not to obscure the vision of the pilot nor produce damage to the rotors, propellers. or other parts of the rotorcraft.

MISCELLANEOUS FLIGHT REQUIREMENTS § 6.140 Flutter and vibration

All parts of the rotorcraft shall be demonstrated to be free from flutter and excessive vibration under all speed and power conditions appropriate to the operation of the type of rotorcraft. (See § 6.711.)

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-4, 24 F.R. 7073, Sept. 1, 1959] Subpart C-Structure

(a) Strength requirements of this subpart are specified in terms of limit and ultimate loads. Unless otherwise stated. the specified loads shall be considered as limit loads. In determining compliance with these requirements the provisions set forth in paragraphs (b) through (e) of this section shall apply.

(b) The factor of safety shall be 1.5 unless otherwise specified, and shall apply to the external and inertia loads, unless its application to the resulting internal stresses is more conservative.

(c) Unless otherwise provided, the specified air, ground, and water loads shall be placed in equilibrium with inertia forces, considering all items of mass in the rotorcraft.

(d) All loads shall be distributed in a manner closely approximating or conservatively representing actual conditions.

(e) If deflections under load significantly change the distribution of external or internal loads, the redistribution shall be taken into account.

§ 6.201 Strength and deformation.

(a) The structure shall be capable of supporting limit loads without suffering detrimental permanent deformations.

(b) At all loads up to limit loads the deformation shall not be such as to interfere with safe operation of the rotorcraft.

(c) The structure shall be capable of supporting ultimate loads without failure. It shall support the load during a static test for at least 3 seconds, unless proof of strength is demonstrated by dynamic tests simulating actual conditions of load application.

§ 6.202 Proof of structure.

(a) Proof of compliance of the structure with the strength and deformation requirements of § 6.201 shall be made for all critical loading conditions.

(b) Proof of compliance by means of structural analysis shall be acceptable only when the structure conforms to types for which experience has shown such methods to be reliable. In all other cases substantiating tests shall be required.

(c) In all cases certain portions of the structure shall be tested as specified in § 6.203. § 6.203

Structural and dynamic tests.

At least the following structural tests shall be conducted to show compliance with the strength criteria:

(a) Dynamic and endurance tests of rotors and rotor drives, including controis (see § 6.412),

(b) Control surface and system limit load tests (see § 6.323).

(c) Control system operation tests (see § 6.324).

(d) Flight stress measurements (see §§ 6.221 and 6.250).

(e) Landing gear drop tests (see § 6.237).

(f) Such additional tests as may be found necessary by the Administrator to substantiate new and unusual features of the design.

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-4, 24 F.R. 7073, Sept. 1, 1959] § 6.203-1 Fixed or ground adjustable stabilizing surfaces (FAA policies which apply to § 6.10 and to § 6.203 (b)).

The purpose of § 6.203 is to require the testing of certain components which in the details of their construction, operational characteristics, or loading, do not lend themselves to established and reliable methods of analysis. In this regard, proof testing such items as fixed or ground adjustable stabilizing surfaces is not considered a minimum requirement and will not be necessary provided suffi

cient experience has been accumulated from previous satisfactory designs, methods of analysis and tests to justify acceptance of these components on the basis of structural analysis. Therefore, these components may be regarded structurally the same as any other part of the basic airframe.

[Supp. 9, 18 F. R. 2877, May 19, 1953]

§ 6.204 Design limitations.

The following values shall be established by the applicant for purposes of showing compliance with the structural requirements specified in this subpart:

(a) Maximum design weight,

(b) Power-on and power-off main rotor rpm ranges (see §§ 6.103 and 6.713 through 6.714 (b)),

(c) Maximum forward speeds for the power-on and power-off rotor rpm ranges established in accordance with paragraph (b) of this section (see § 6.711),

(d) Maximum rearward and sideward flight speeds,

(e) Extreme positions of rotorcraft center of gravity to be used in conjunction with the limitations of paragraphs (b), (c), and (d) of this section,

(f) Rotational speed ratios between the powerplant and all connected rotating components, and

(g) Positive and negative limit maneuvering load factors.

Flight load requirements shall be complied with at all weights from the design minimum weight to the design maximum weight, with any practicable distribution of disposable load within prescribed operating limitations stated in the Rotorcraft Flight Manual. (See § 6.741.) § 6.211 Flight load factors.

The flight load factors shall represent rotor load factors. The net load factor acting at the center of gravity of the rotorcraft shall be obtained by proper consideration of balancing loads acting in the specific flight conditions. § 6.212 Maneuvering conditions.

The rotorcraft structure shall be designed for a positive maneuvering limit load factor of 3.5 and for a negative maneuvering limit load factor of 1.0, except that lesser values shall be allowed

if the manufacturer shows by analytical study and flight demonstrations that the probability of exceeding the values selected is extremely remote. In no case shall the limit load factors be less than 2.0 positive and 0.5 negative. The resultant loads shall be assumed to be applied at the center(s) of the rotor hub(s) and to act in such directions as necessary to represent all critical maneuvering motions of the rotorcraft applicable to the particular type, including flight at the maximum design rotor tip speed ratio under power-on and power-off conditions.

The structure of all auxiliary rotors (antitorque and control), fixed or movable stabilizing and control surfaces, and all systems operating any flight controls shall be designed to comply with the provisions of §§ 6.221 through 6.225.

§ 6.221 Auxiliary rotor assemblies.

Auxiliary rotor assemblies shall be tested in accordance with the provisions of § 6.412 for rotor drives. In addition, auxiliary rotor assemblies with detachable blades shall be substantiated for centrifugal loads resulting from the maximum design rotor rpm. In the case of auxiliary rotors with highly stressed metal components, the vibration stresses shall be determined in flight, and it shall be demonstrated that these stresses do not exceed safe values for continuous operation.

§ 6.221-1 Service life or auxiliary rotor assemblies (FAA interpretations which apply to § 6.221).

The requirement in § 6.221 that vibration stresses in highly stressed metal components of auxiliary rotors must not exceed safe values for continuous operation is interpreted to mean that the service life of such components should be determined by fatigue tests or by other methods found acceptable by the Administrator. The methods of service life determination for main rotor structure outlined under § 6.250-1 are

considered to be acceptable in showing compliance with the pertinent portion of § 6.221.

[Supp. 6, 16 F. R. 3405, Apr. 19, 1951] § 6.222 Auxiliary rotor attachment

structure.

The attachment structure for the auxiliary rotors shall be designed to withstand a limit load equal to the maximum loads in the structure occurring under the flight and landing conditions.

§ 6.223 Tail rotor guard.

When a tail rotor is provided on a rotorcraft it shall not be possible for the tail rotor to contact the landing medium during a normal landing. If a tail rotor guard is provided which will contact the landing medium during landings and thus prevent tail rotor contact, suitable design loads for the guard shall be established, and the guard and its supporting structure shall be designed to withstand the established loads.

§ 6.224 Stabilizing and control surfaces.

Stabilizing and control surface shall be designed to withstand the critical loading from maneuvers or from combined maneuver and gust. In no case shall the limit load be less than 15 pounds per square foot or a load due to CN=0.55 at the maximum design speed. The load distribution shall simulate closely the actual pressure distribution conditions.

§ 6.225 Primary control systems.

Manual control systems shall comply with the provisions of paragraphs (a) and (b) of this section.

(a) From the pilot compartment to the stops which limit the range of motion of the pilots' controls, the controls shall be designed to withstand the limit pilot applied forces as set forth in subparagraphs (1) through (3) of this paragraph, unless it is shown that the pilot is unable to apply such loads to the system. In the latter event the system shall be designed for the maximum loads which the pilot is able to apply, except that in any case values less than 0.60 of those specified shall not be employed.

(1) Foot type controls-130 pounds, (2) Stick type controls-fore and aft 100 pounds—laterally 67 pounds,

(3) Wheel type controls-fore and aft 100 pounds-laterally 53-pound couple

applied on opposite sides of the control wheel.

(b) From the stops to the attachment of the control system to the rotor blades (or control areas) the control system shall be designed to withstand the maximum loads which can be obtained in normal operation of the rotorcraft, except that where jamming, ground gusts, control inertia, or friction can cause loads exceeding operational loads, the system shall be capable of supporting without yielding 0.60 of the loads specified in paragraph (a) (1), (2), and (3) of this section.

(a) Loads and equilibrium. The limit loads obtained in the landing conditions shall be considered as external loads which would occur in a rotorcraft structure if it were acting as a rigid body. In each of the conditions the external loads shall be placed in equilibrium with the linear and angular inertia loads in a rational or conservative manner. In applying the specified conditions the provisions of paragraphs (b) through (e) of this section shall be complied with.

(b) Center of gravity positions. The critical center of gravity positions within the certification limits shall be selected so that the maximum design loads in each of the landing gear elements are obtained.

(c) Design weight. The design weight used in the landing conditions shall not be less than the maximum weight of the rotorcraft. It shall be acceptable to assume a rotor lift, equal to one-half the design maximum weight, to exist throughout the landing impact and to act through the center of gravity of the rotorcraft. Higher values of rotor lift shall be acceptable if substantiated by tests and/or data which are applicable to the particular rotorcraft.

(d) Load factor. The structure shall be designed for a limit load factor, selected by the applicant, of not less than the value of the limit inertia load factor substantiated in accordance with the provisions of § 6.237, except in conditions in which other values of load factor are prescribed.

(e) Landing gear position. The tires shall be assumed to be in their static position, and the shock absorbers shall be assumed to be in the most critical position, unless otherwise prescribed.

The

(f) Landing gear arrangement. provisions of §§ 6.231 through 6.236 shall be applicable to landing gear arrangements where two wheels are located aft and one or more wheels are located forward of the center of gravity.

§ 6.231 Level landing conditions.

(a) Under loading conditions prescribed in paragraph (b) of this section, the rotorcraft shall be assumed to be in the following two level landing attitudes: (1) All wheels contacting the ground simultaneously, and

(2) The aft wheels contacting the ground while the forward wheel(s) being just clear of the ground.

(b) The following two level landing loading conditions shall be considered. Where the forward portion of the landing gear has two wheels, the total load applied to the forward wheels shall be divided between the two wheels in a 40:60 proportion.

(1) Vertical loads shall be applied in accordance with the provisions of § 6.230.

(2) The vertical loads specified in subparagraph (1) of this paragraph shall be combined with a drag load at each wheel. The drag loads shall not be less than 25 percent of the respective vertical loads. For the attitude prescribed in paragraph (a) (1) of this section the resulting pitching moment shall be assumed resisted by the forward gear, while for the attitude prescribed in paragraph (a) (2) of this section the resulting pitching moment shall be assumed resisted by angular inertia forces.

§ 6.231-1

Distribution of vertical ground reaction loads and determination of angular inertia loads (FAA interpretations which apply to § 6.231(b) (2)).

(a) Although § 6.231(b) (2) states that the vertical loads are those specified in § 6.231(b) (1), the distribution of the vertical loads among the ground reaction points is not necessarily the same for the two subparagraphs since the requirements of § 6.230 must be met. Section 6.230(a) states, in part, that the external loads shall be placed in equilibrium with the linear and angular inertia loads in a rational or conservative manner.

(b) Compliance with § 6.231 (b) (2) is interpreted to require that a vertical inertia load of nW and a horizontal inertia load of 0.25 nW be applied at the center of gravity. For the level landing

with drag on all wheels, the vertical ground reaction loads should be distributed between the forward and rear wheels to place the ground reaction loads in equilibrium with the rotorcraft linear inertia loads. For the level landing with drag on main wheels only, the pitching moments arising from the vertical and horizontal ground reactions should be placed in equilibrium with an angular inertia load about the c. g.

(c) The drag load at each wheel, in both cases, is required to be equal to 0.25 times the respective wheel vertical load. [Supp. 7, 17 F. R. 8323, Sept. 17, 1952] § 6.232 Nose-up landing condition.

The rotorcraft shall be assumed in the maximum nose-up attitude permitting clearance of the ground by all parts of the rotorcraft. The ground loads shall be applied perpendicularly to the ground. § 6.233 One-wheel landing condition.

The rotorcraft shall be assumed in the level attitude to contact the ground on one of the wheels located aft of the center of gravity. The vertical load shall be the same as that obtained on the one side in the condition specified in § 6.231 (b) (1). The unbalanced external loads shall be reacted by the inertia of the rotorcraft.

§ 6.234 Lateral-drift landing condition.

(a) The rotorcraft shall be assumed in the level landing attitude. Side loads shall be combined with one-half the maximum ground reactions obtained in the level landing conditions of § 6.231 (b) (1). These loads shall be applied at the ground contact point, unless the landing gear is of the full-swiveling type in which case the loads shall be applied at the center of the axle. The conditions set forth in paragraphs (b) and (c) of this section shall be considered.

(b) Only the wheels aft of the center of gravity shall be assumed to contact the ground. Side loads equal to 0.8 of the vertical reaction acting inward (on one side) and 0.6 of the vertical reaction acting outward (on the other side) shall be combined with the vertical loads specified in paragraph (a) of this section.

(c) The forward and aft wheels shall be assumed to contact the ground simultaneously. Side loads on the wheels aft of the center of gravity shall be applied in accordance with paragraph (b) of this section. A side load at the forward

gear equal to 0.8 of the vertical reaction shall be combined with the vertical load specified in paragraph (a) of this section. § 6.235 Brake roll conditions.

The rotocraft attitudes shall be assumed to be the same as those prescribed for the level landing conditions in § 6.231 (a), with the shock absorbers deflected to their static position. The limit vertical load shall be based upon a load factor of 1.33 when the rotorcraft attitude is as specified in § 6.231 (a) (1); the limit vertical load factor may be reduced to 1.0 when the attitude is as specified in § 6.231 (a)(2). A drag load equal to the vertical load multiplied by a coefficient of friction of 0.8 shall be applied at the ground contact point of each wheel equipped with brakes, except that the drag load need not exceed the maximum value based on limiting brake torque.

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 604, 24 F.R. 7073, Sept. 1, 1959] § 6.236 Taxiing condition.

The rotorcraft and its landing gear shall be designed for loads which occur when the rotorcraft is taxied over the roughest ground which it is reasonable to expect in normal operation.

§ 6.237

Shock absorption tests.

Drop tests shall be conducted in accordance with paragraphs (a) and (b) of this section to substantiate the landing limit inertia load factor (see § 6.230(d)) and to demonstrate the reserve energy absorption capacity of the landing gear. The drop tests shall be conducted with the complete rotorcraft or on units consisting of wheel, tire, and shock absorber in their proper relation.

(a) Limit drop test. The drop height in the limit drop test shall be 13 inches measured from the lowest point of the landing gear to the ground. A lesser drop height shall be permissible if it results in a drop test contact velocity found by the Administrator to be equal to the greatest probable sinking speed of the rotorcraft at ground contact in power-off landings likely to be made in normal operation of the rotorcraft. In no case shall the drop height be less than 8 inches. If rotor lift is considered (see § 6.230 (c)), it shall be introduced in the drop test by the use of appropriate energy absorbing devices or by the use of an effective mass. The attitude in which the landing gear unit is tested shall be such as to simulate the land