FUSELAGES AND ENGINE MOUNTS

§ 4a.289 Proof of fuselages and engine

mounts.

Structural analyses of fuselages and engine mounts will be accepted as complete proof of compliance with load requirements only when the structure conforms with conventional types for which reliable analytical methods are available.

§ 4a.290 Critical column loads.

The end fixity coefficient used in determining critical column loads shall in no case exceed 2.0. A value of 1.0 shall be used for all members in the engine mount. In doubtful cases, tests are required to substantiate the degree of restraint assumed.

§ 4a.291 Baggage compartments.

The ability of baggage compartments to sustain the maximum authorized baggage loads under all required flight and landing conditions shall be demonstrated.

FITTINGS AND PARTS

§ 4a.297 Proof of fittings and parts.

Proof of strength of all fittings and joints of the primary structure is required. Where applicable, structural analysis methods may be used. When such methods are inadequate, a load test is required. Compliance with the multiplying factor of safety requirements for fittings (§§ 4a.207-4a.216) shall be demonstrated.

§ 4a.298 Fittings and attaching members.

Since the system of forces which designs a fitting does not necessarily include the forces which design the attaching members, all the forces acting in all the specified conditions shall be considered for every fitting. The strength of each part of a built-up fitting shall be investigated and proper allowance shall be made for the effects of eccentric loading when initially present or when introduced by deflection of the structure under load.

The primary structure and all mechanisms essential to the safe operation of the airplane shall not incorporate design details which experience has shown to be unreliable or otherwise unsatisfactory. The suitability of all design details shall be established to the satisfaction of the Administrator. Certain design features which have been found to be essential to the airworthiness of an airplane are specified in this subpart and shall be observed.

MATERIALS, WORKMANSHIP, AND FABRICATION METHODS

§ 4a.302 Materials and workmanship.

The primary structure shall be made from materials which experience or conclusive tests have proved to be uniform in quality and strength and to be otherwise suitable for airplane construction. Workmanship shall be of sufficiently high grade as to insure proper continued functioning of all parts.

§ 4a.303 Fabrication methods.

The methods of fabrication employed in constructing the primary structure shall be such as to produce a uniformly sound structure which shall also be reliable with respect to maintenance of the original strength under reasonable service conditions.

Torch welding of primary structural parts may be used only for ferrous materials and for such other materials shown to be suitable therefor.

§ 4a.306 Electric welding.

Electric arc, spot, or seam welding may be used in the primary structure when specifically approved by the Administrator for the application involved. Requests for approval of the use of electric

welding shall be accompanied by information as to the extent to which such welding is to be used, drawings of the parts involved, apparatus employed, general methods of control and inspection, and references to test data substantiating the strength and suitability of the welds obtained.

§ 4a.307 Brazing and soldering.

The use of brazing and soldering in joining parts of the primary structure is prohibited except that brazing may be used in special cases when the suitability of the method and application can be definitely established to the satisfaction of the Administrator.

All members of the primary structure shall be suitably protected against deterioration or loss of strength in service due to corrosion, abrasion, vibration, or other causes. This applies particularly to design details and small parts. In seaplanes special precautions shall be taken against corrosion from salt water, particularly where parts made from different metals are in close proximity. All exposed wood structural members shall be given at least two protective coatings of varnish or approved equivalent. Builtup box spars and similar structures shall be protected on the interior by at least one coat of varnish or approved equivalent and adequate provisions for drainage shall be made. Due care shall be taken to prevent coating of the gluing surfaces.

§ 4a.309 Inspection.

Inspection openings of adequate size shall be provided for such vital parts of the aircraft as require periodic inspection.

JOINTS, FITTINGS, AND CONNECTING PARTS § 4a.312 Joints, fittings, and connecting parts.

In each joint of the primary structure the design details shall be such as to minimize the possibility of loosening of the joint in service, progressive failure due to stress concentration, and damage caused by normal servicing and field operations.

CROSS REFERENCE: For multiplying factors of safety required, see § 4a.208.

§ 4a.313 Bolts, pins, and screws.

All bolts and screws in the structure shall be of uniform material of high

TIE-RODS AND WIRES

§ 4a.319 Tie-rods and wires.

The minimum size of tie-rod which may be used in primary structure is No. 6-40. The corresponding minimum allowable size of single-strand hard wire is No. 13 (0.072-inch diameter). § 4a.320 Wire terminals.

The assumed terminal efficiency of single-strand hard wire shall not be greater than 85 percent.

§ 4a.321 Wire anchorages.

A fitting attached to a wire or cable up to and including the 3,400-pound size shall have at least the rated strength of the wire or cable, and the multiplying factor of safety for fitting (§ 4a.208) is not required in such cases. In the case of fittings to which several tie-rods or wires are attached, this requirement applies separately to each portion of the fitting to which a tie-rod or wire is at

tached, but does not require simultaneous application of rated wire loads. The end connections of brace wires shall be such as to minimize restraint against ending or vibration.

4a.322 Counter wire sizes.

(See also §§ 4a.211, 4a.212.) In a wirebraced structure the wire sizes shall be such that any wire can be rigged to at least 10 percent of its rated strength without causing any other wire to be baded to more than 20 percent of its rated strength. As used here "rated trength" refers to the wire proper, not the terminal.

FLUTTER PREVENTION

4a.326 General flutter prevention

measures.

When he deems it necessary in the nterest of safety, the Administrator may equire special provisions against flutter. or specific requirements, see §§ 4a.264, a.336, 4a.449, 4a.450, 4a.451, 4a.452, a.465, 4a.466 and 4a.680.

Amdt. 75, 5 FR. 3946, Oct. 8, 1940, as mended by Amdt. 04-2, 8 F.R. 13999, Oct. 4,1943]

DETAIL DESIGN OF WINGS

4a.329 External bracing.

When streamline wires are used for xternal lift bracing they shall be double nless the design complies with the liftire-cut condition specified in § 48.95. See also § 4a.210.)

4a.330

Wire-braced monoplanes.

If monoplane wings are externally raced by wires only, the right and left des of the bracing shall be independent I each other so that an unsymmetrical ad from one side will not be carried rough the opposite wires before being unteracted, unless the design complies Ith the following conditions:

(a) The minimum true angle between y external brace wire and a spar is 14 egrees.

(b) The counter (landing) wires are signed to remain in tension at least to the limit load.

(c) The landing and flying wires are uble.

§ 4a.332 Jury struts.

When clamps are used for attachment of jury struts to lift struts, the design shall be such as to prevent misalignment or local crushing of the lift strut. § 4a.333 Wing beams.

Provisions shall be made to reinforce wing beams against torsional failure, especially at the point of attachment of lift struts, brace wires, and aileron hinge brackets.

§ 4a.334 Wing beam joints.

Joints in metal beams (except pinned joints) and joints in mid-bays of wood beams shall maintain 100 percent efficiency of the beam with respect to bending, shear, and torsion.

(a) Fabric-covered wing structures having a cantilever length of overhang such that the ratio of span of overhang to chord at root of overhang is greater than 1.75 shall have a double system of internal drag trussing spaced as far apart as possible, or other means of providing equivalent torsional stiffness. In the former case counter wires shall be of the same size as the drag wires. (See also § 4a.212.)

(b) Multiple-strand cable shall not be used in drag trusses unless such use is substantiated to the satisfaction of the Administrator.

§ 4a.336 Aileron and flap attachments.

Aileron and flap attachment ribs or brackets shall be rigidly constructed and firmly attached to the main wing structure in order to reduce wing flutter tendencies.

All control systems and operating devices shall be so designed and installed as to provide reasonable ease of operation by the crew and so as to preclude the probability of inadvertent operation, jamming, chafing, interference by cargo, passengers, or loose objects, and the slapping of cables against parts of the airplane. All pulleys shall be provided with satisfactory guards.

[Amdt. 56, 5 F. R. 2100, June 1, 1940] § 4a.460 Stops.

All control systems shall be provided with stops which positively limit the range of motion of the control surfaces. Stops shall be capable of withstanding the loads corresponding to the design conditions for the control system.

Bolts with castellated nuts safetied with cotter pins or with an approved type of self-locking nut shall be used throughout the control system, except that the use of clevis pins in standard cable ends, thimbles, and shackles is satisfactory for light airplanes.

[CAR, May 31, 1938, as amended by Amdt 04-2, 8 F. R. 13999, Oct. 14, 1943] § 4a.462 Welds.

Welds shall not be employed in control systems to carry tension withou reinforcement from rivets or bolts.

§ 4a.463 Flap controls.

The flap operating mechanism shal be such as to prevent sudden, inad vertent, or automatic opening of the flap at speeds above the design speed for the extended flap conditions. time required to fully extend or re tract flaps shall not be less than 1 seconds, unless it can be demonstrated to the satisfaction of the Administra

The

tor that the operation of the flaps in a lesser time does not result in unsatisfactory flight characteristics. Means shall be provided to retain flaps in their fully retracted position and to indicate such position to the pilot.

§ 4a.464–T_Flap controls.

(a) For transport category airplanes, the flap control shall provide means for bringing the flaps from any position within the operating range to any one of three positions, designated as landing, approach, and take-off positions, or to the fully retracted position, by placing the primary flap control in a single setting marked as corresponding to each such flap position, the flaps thereupon moving directly to the desired position without requiring further attention. If any extension of the flaps beyond the landing position is possible, the flap control shall be clearly marked to identify such range of extension.

(b) The landing position, approach position, and take-off position, or any of them, may be made variable with altitude or weight by means of a secondary flap control provided for that purpose. Such a secondary control, if provided, shall operate independently of the primary control and in such manner that when it has been adjusted (for the effect of weight or altitude), the necessary flap position can thereafter be obtained by placing the primary flap control in the desired position. The secondary control shall be so designed and marked as to be readily operable by the crew.

(c) The rate of flap retraction shall be such as to permit compliance with 4a.752-T.

[Amdt. 04-4, 7 F. R. 984, Feb. 14, 1942]

§ 4a.465 Tab controls.

(a) Tab controls shall be irreversible and nonflexible, unless the tab is statically balanced about its hinge line. Proper precautions shall be taken against the possibility of inadvertent or abrupt tab operation and operation in the wrong direction.

(b) When adjustable elevator tabs are used for the purpose of trimming the airplane, a tab position indicator shall be installed, and means shall be provided for indicating to the pilot a range of adjustment suitable for safe take-off and the directions of motion of the control for nose-up and nose-down motions of the airplane.

§ 4a.466 Spring devices.

The use of springs in the control system either as a return mechanism or as an auxiliary mechanism for assisting the pilot (bungee device) is prohibited except under the following conditions:

(a) The airplane shall be satisfactorily maneuverable and controllable and free from flutter under all conditions with and without the use of the spring device.

(b) In all cases the spring mechanism shall be of a type and design satisfactory to the Administrator.

(c) Rubber cord shall not be used for this purpose.

§ 4a.467 Single-cable controls.

Single-cable controls are prohibited except in special cases in which their use can be proved to be satisfactory. § 4a.468 Control system locks.

When a device is provided for locking a control surface while the aircraft is on the ground or water, compliance with the following requirements shall be shown.

(a) The locking device shall be so installed as to positively prevent taxiing the aircraft faster than 20 miles per hour, either intentionally or inadvertently, while the lock is engaged.

(b) Means shall be provided to preclude the possibility of the lock becoming engaged during flights.

§ 4a.469-T Trim controls.

For transport category airplanes, the trimming devices shall be capable of continued normal operation in spite of the failure of any one connecting or transmitting element in the primary control system. Trim controls shall operate in the plane and with the sense of the motion of the airplane which their operation is intended to produce. [Amdt. 04-5, 7 F. R. 985, Feb. 14, 1942]

DETAIL DESIGN OF LANDING GEAR

§ 4a.475 Shock absorption.

All landing gear (including tail gear installations) shall be provided with shock-absorbing systems which will permit the airplane to be landed under the conditions specified in §§ 4a.148(b), 4a.152 without exceeding the ultimate load used in the analysis of any landing gear member. (See § 4a.278 for proof of absorption capacity.) If the design of the shock-absorbing system is such that the