(a) General. The following policies will govern the Federal Aviation Agency in determining circumstances under which a wing flap position indicator is required:

(1) Flap installations incorporating only the fully extended and fully retracted positions. An indicator is required except:

(i) Where a direct operating mechanism provides a sense of feel and position such as when a mechanical linkage is employed, or,

(ii) Where the flap position is readily determined without seriously detracting from other piloting duties under all conditions of flight, either day or night.

(2) Flap installations incorporating an intermediate flap position. An indicator is required except when the installation complies with subparagraph (1) (i) of this paragraph.

[Supp. 5, 14 F. R. 5743, Sept. 20, 1949]

§ 3.339 Flap interconnection.

(a) The motion of flaps on opposite sides of the plane of symmetry shall be synchronized by a mechanical interconnection, unless the airplane is demonstrated to have safe flight characteristics while the flaps are retracted on one side and extended on the other.

(b) Where an interconnection is used, in the case of multiengine airplanes, it shall be designed to account for the unsymmetrical loads resulting from flight with the engines on one side of the plane of symmetry inoperative and the remaining engines at take-off power. For single-engine airplanes, it may be assumed that 100 percent of the critical air load acts on one side and 70 percent on the other.

§ 3.340 Stops.

All control systems shall be provided with stops which positively limit the range of motion of the control surfaces.

Stops shall be so located in the system that wear, slackness, or take-up adjustments will not appreciably affect the range of surface travel. Stops shall be capable of withstanding the loads corresponding to the design conditions for the control system.

§ 3.341 Control system locks.

When a device is provided for locking a control surface while the airplane is on the ground or water:

(a) The locking device shall be so installed as to provide unmistakable warning to the pilot when it is engaged, and

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

§ 3.342 Proof of strength.

Tests shall be conducted to prove compliance with limit load requirements. The direction of test loads shall be such as to produce the most severe loading of the control system structure. The tests shall include all fittings, pulleys, and brackets used to attach the control system to the primary structure. Analyses or individual load tests shall be conducted to demonstrate compliance with the multiplying factor of safety requirements specified for control system joints subjected to angular motion.

§ 3.343 Operation test.

An operation test shall be conducted by operating the controls from the pilot compartment with the entire system so loaded as to correspond to the limit air loads on the surface. For secondary control systems including adjustable stabilizers, the system loading in the operation test need not exceed that corresponding to the maximum pilot effort established in accordance with § 3.234. In this test there shall be no jamming, excessive friction, or excessive deflection. [21 F. R. 3339, May 22, 1956, as amended by Amdt. 3-2, 22 F. R. 5561, July 16, 1957]

shall be made to prevent the slapping of cables or tubes against parts of the airplane. The elements of the flight control system shall incorporate design features or shall be distinctively and permanently marked to minimize the possibility of incorrect assembly which could result in malfunctioning of the control system. § 3.345 Cable systems.

Cables, cable fittings, turnbuckles, splices, and pulleys shall be in accordance with approved specifications. Cables smaller than -inch diameter shall not be used in primary control systems. The design of cable systems shall be such that there will not be hazardous change in cable tension throughout the range of travel under operating conditions and temperature variations. Pulley types and sizes shall correspond to the cables with which they are used, as specified on the pulley specification. All pulleys shall be provided with satisfactory guards which shall be closely fitted to prevent the cables becoming misplaced or fouling, even when slack. The pulleys shall lie in the plane passing through the cable within such limits that the cable does not rub against the pulley flange. Fairleads shall be so installed that they are not required to cause a change in cable direction of more than 3 degrees. Clevis pins (excluding those not subject to load or moion) retained only by cotter pins shall not be employed in the control system. Turnbuckles shall be attached to parts having angular motion in such a manner is to prevent positively binding throughbut the range of travel. Provisions for visual inspection shall be made at all fairleads, pulleys, terminals, and turnbuckles.

33.345-1 Cables in primary control systems (FAA interpretations which apply to § 3.345).

# Section 3.345 provides that "cables maller than 8-inch diameter shall not >e used in primary control systems." Primary control systems are normally considered to be the aileron, rudder, and elevator control systems. Hence this ninimum of 1⁄2 inch need not be aplied to tab control cables having high trength margins. However, in cases vhere the airplane would not be safely ontrollable in flight and landing with abs in the most adverse positions required for the various critical trim, veight, and center of gravity conditions,

the Administrator will require that tab systems be so designed as to provide reliability equivalent to that required for primary systems. Examples are pulley sizes, guards, use of fairleads, inspection provisions, etc.

[Supp. 1, 12 F. R. 3437, May 28, 1947, as amended by Amdt. 1, 14 F. R. 35, Jan. 5, 1949] § 3.345-2 Special aircraft turnbuckle assemblies and/or turnbuckle safetying devices (FAA rules which apply to § 3.345).

The minimum safety requirements for special aircraft turnbuckle assemblies and/or turnbuckle safetying devices which are intended for use in civil aircraft have been established by the Administrator in Technical Standard Order No. TSO-C21 effective October 1, 1949, "Special Aircraft Turnbuckle Assemblies and/or Turnbuckle Safetying Devices" (§ 514.21 of this title).

[Supp. 10, 16 F. R. 3289, Apr. 14, 1951] § 3.345-3 Approval of pulleys for control systems (FAA policies which apply to § 3.345).

The Administrator does not issue specific approval as such for pulleys for general use on aircraft. Approval is limited to its use as a part of a specific airplane design. Conformance with the Army-Navy-Aeronautical Standards, National Aircraft Standards, or with standards established for pulleys previously approved for use in civil aircraft, or adequate substantiation of the manufacturer's own design through stress analysis or tests are the procedures utilized in complying with the "approved specifications" of CAR 3.345.

[Supp. 22, 20 F. R. 8809, Dec. 1, 1955] § 3.346

Joints.

Shock absorbing elements in main, nose, and tail wheel units shall be substantiated by the tests specified in the following section. In addition, the shock absorbing ability of the landing gear in taxiing must be demonstrated in the operational tests of § 3.146.

§ 3.352 Shock absorption tests.

(a) It shall be demonstrated by energy absorption tests that the limit load factors selected for design in accordance with § 3.243 will not be exceeded in landings with the limit descent velocity specified in that section.

(b) In addition, a reserve of energy absorption shall be demonstrated by a test in which the descent velocity is at least 1.2 times the limit descent velocity. In this test there shall be no failure of the shock absorbing unit, although yielding of the unit will be permitted. Wing lift equal to the weight of the airplane may be assumed for purposes of this test. § 3.352-1 Landing gear drop tests (FAA policies which apply to § 3.352).

(a) The following method has been approved by the Administrator for determining the effective mass to be dropped in drop tests of nose wheel landing gear assemblies pursuant to § 3.352 (a): For aircraft with nose wheel type gear, the effective mass to be used in free drop test of the nose wheel shall be determined from the formula for We (§§ 3.353 and 3.355) using W=Wn where Wn is equal to the vertical components of the resultant force acting on the nose wheel, computed under the following assumptions: (1) the mass of the airplane concentrated at the center of gravity and exerting a force of 1.0 g downward and 0.33 g forward, (2) the nose and the main gears and tires in static position, and (3) the resultant reactions at the main and nose gears acting through the axles and parallel to the resultant force at the airplane center of gravity.

NOTE: By way of explanation, the use of an inclined reactions condition as the basis for determining the mass to be dropped with a nose wheel unit is based on rational dynamic investigation of the landing condition, assuming the landing is made with simultaneous three-point contact, zero pitching

velocity, and a drag component representin the average wheel spin-up reactions duri the landing impact. Although spin-up loa on small airplanes may be less than the val implied by the formula, such airplanes a more likely to be landed with a nosing dow pitching velocity, or in soft ground. T vertical component of the ground reaction specified above because the method of d fining the direction of the inertia force the center of gravity gives a resultant effe tive mass greater than that of the airplan

(b) The following procedure has bee approved by the Administrator for d termining the attitude in which th landing gear unit should be droppe pursuant to § 3.352 (a): The attitude which a landing gear unit it dropp shall be that which simulates the ai plane landing condition which is critic from the standpoint of energy to absorbed by the particular unit, thus: ( For nose wheel type landing gear, th nose wheel gear shall be drop tested an attitude which simulates the thr point landing inclined reaction cond tion; (2) the attitude selected for ma gear drop tests shall be that which sim lates the two-wheel level landing wi inclined reactions condition.

NOTE: In addition, it is recommended th the main gear be dropped in an attitu simulating the tail-down landing with ve tical reactions condition if the geometry the gear is such that this condition is like to result in shock strut action appreciab different from that obtained in level at tude drop tests; for example, when a can lever shock strut has a large inclination wi respect to the direction of the grou reaction.

(3) Tail wheel units shall be tested such a manner as to simulate the ta down landing condition (three-poi contact). Drag components may covered separately by the tail whe "obstruction" condition.

(c) The Administrator has accept the following procedure for determini slopes of inclined platforms when su are used in drop tests: When the ar trary drag components given on F 3-12 (a) of this part are used for t design of the landing gear in the le landing conditions, the drag loads the drop tests for these conditions m be simulated by dropping the units or inclined platforms so arranged as obtain the proper direction of the sultant ground reactions in relation the landing gear. (If wheel spinloads for these conditions are det mined by rational methods and found

be more severe than the arbitrary drag loads, it is suggested that the spin-up loads be simulated by dropping the gear onto a level platform with wheel spinning.) In at least one limit drop test the platform should simulate the friction characteristics of paved runways and the rotational speed of the wheel just prior to contact should correspond to an airplane ground speed of 1.2 Vs. It is suggested that additional limit drops be made onto surfaces of lower friction coefficient and at several wheel rotational speeds; coefficients for example, corresponding to 0.6, 0.8 and 1.0 Vε0.

The di

rection of wheel rotation in the drop test should be opposite to that which would occur in landing the airplane. Spin-up loads which are slightly greater than the arbitrary drag loads can probably be simulated satisfactorily by inclined platforms, but platforms having greater inclinations may not simulate spin-up loads correctly and are not recommended.

[Supp. 1, 12 F. R. 3437, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5, 1949] § 3.353 Limit drop tests.

(a) If compliance with the specified limit landing conditions of § 3.352(a) is demonstrated by free drop tests, these shall be conducted on the complete airplane, or on units consisting of wheel, tire, and shock absorber in their proper relation, from free drop heights not less than the following:

h (inches) 3.6 (W/S) 0.5

except that the free drop height shall not be less than 9.2 inches and need not be greater than 18.7 inches.

(b) In simulating the permissible wing lift in free drop tests, the landing gear unit shall be dropped with an effective mass equal to:

where:

h+ (1—L) d h+d

d]

We the effective weight to be used in the drop test.

h specified height of drop in inches. d:= deflection under impact of the tire

(at the approved inflation pressure) plus the vertical component of the axle travel relative to the drop mass. The value of d used in the computation of We shall not exceed the value actually obtained in the drop tests.

designed for the maximum load factors in the flight conditions when the gear is in the retracted position. It shall also be designed for the combination of friction, inertia, brake torque, and air loads occurring during retraction at any air speed up to 1.6 V11, flaps retracted and any load factors up to those specified for the flaps extended condition, § 3.190. The landing gear and retracting mechanism, including the wheel well doors, shall withstand flight loads with the landing gear extended at any speed up to at least 1.6 V1, flaps retracted. Positive means shall be provided for the purpose of maintaining the wheels in the extended position. § 3.356-1 Retracting mechanism (FAA policies which apply to § 3.356).

(a) In order to provide for adequate strength of the landing gear doors, landing gear, etc., in yawed attitude, it will be satisfactory to show compliance with § 3.356 at the maximum yaw angle as determined by the flight characteristic requirements of § 3.105 and at speeds up to 1.6V., flaps retracted.

(b) To meet the requirement that a positive means be provided for maintaining wheels in the extended position, a positive mechanized lock or latch should be provided that can be released directly or sequentially only by some specific manual actuation by the pilot. In this regard, the use of hydraulic pressure is not considered a positive means of down lock.

[Supp. 10, 16 F. R. 3289, Apr. 14, 1951]

§ 3.357 Emergency operation.

When other than manual power for the operation of the landing gear is employed, an auxiliary means of extending the landing gear shall be provided.

§ 3.358 Operation test.

Proper functioning of the landing gear retracting mechanism shall be demonstrated by operation tests.

(a) The means to be provided for in § 3.359 to indicate to the pilot when the wheels are secured in the extreme positions may consist of lights. For example, a green light for the gear down and locked position is considered satisfactory, provided a placard indicates that this is the down position. "All lights

out" is considered satisfactory for intermediate gear positions. However, there should then be another light indicating gear up and locked. "All lights out" is not considered desirable for either extreme gear locked position, since such a system would not "fail safe" if a lamp burned out.

(b) The regulations do not require an aural warning device for amphibian aircraft. A two-light warning system similar to the following would be considered sufficient and satisfactory:

Gear Up------- Light #1.
Gear Down............. Light #2..

Water Land

When light #1 is on, the gear would be in the extreme up position and locked and when light #2 is on, the gear would be in the extreme down position and locked.

[Supp. 10, 16 F. R. 3289, Apr. 14, 1951] § 3.359-2 Position indicator and warning device (FAA policies which apply to § 3.359).

A throttle stop is not considered an acceptable alternative to an aural landing gear warning device.

[Supp. 10, 16 F.R. 3289, Apr. 14, 1951]

§ 3.359-3 Landing gear position indicator switches (FAA interpretations which apply to § 3.359).

The phrase "means shall be provided for indicating to the pilot" includes a landing gear position indicator as well as the switches necessary to actuate such indicator. The switches must be so located and coupled to the landing gear mechanical system as to preclude the possibility of an erroneous indication of 'down and locked" if the landing gear is not in a fully extended position, or "up and locked" if the landing gear is not in the completely retracted position. Location of the switches so that they are operated by the actual landing gear locking latch or device is an acceptable method of compliance with the requirements of this section.

[Supp. 23, 21 F.R. 3002, May 5, 1956]