(c) Other inspection procedures and variability factors may be used if found satisfactory by the Administrator.

§ 3.304-1 Casting factors (FAA policies which apply to § 3.304).

With reference to paragraphs (b) and (c) of § 3.304, the Administrator has approved specific proposals which permit the use of lower casting factors as specified in (b), with 100 percent radiographic inspection on initial runs, but with radiographic inspection gradually reduced on production lots as it becomes evident that adequate quality control has been established. All such procedures require the submittal and execution of a satisfactory process specification and statistical proof that adequate quality control has been achieved.

[Supp. 1, 12 F. R. 3437, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5, 1949] § 3.305 Bearing factors.

(a) The factor of safety in bearing at bolted or pinned joints shall be suitably increased to provide for the following conditions:

(1) Relative motion in operation (control surface and system joints are covered in §§ 3.327-3.347).

(2) Joints with clearance (free fit) I subject to pounding or vibration.

(b) Bearing factors need not be applied when covered by other special factors.

§ 3.306 Fitting factor.

This

Fittings are defined as parts such as end terminals used to join one structural member to another. A multiplying factor of safety of at least 1.15 shall be used in the analysis of all fittings the strength of which is not proved by limit and ultimate load tests in which the actual stress conditions are simulated in the fitting and the surrounding structure. factor applies to all portions of the fitting, the means of attachment, and bearing on the members joined. In the case of integral fittings, the part shall be treated as a fitting up to the point where the section properties become typical of the member. The fitting factor need not be applied where a type of joint design based on comprehensive test data is used. The following are examples: continuous joints in metal plating, welded joints, and scarf joints in wood, all made in accordance with approved practices.

§ 3.307 Fatigue strength.

The structure shall be designed, insofar as practicable, to avoid points of stress concentration where variable stresses above the fatigue limit are likely to occur in normal service.

FLUTTER AND VIBRATION

§ 3.311 Flutter and vibration prevention

measures.

Wings, tail, and control surfaces shall be free from flutter, airfoil divergence, and control reversal from lack of rigidity, for all conditions of operation within the limit V-n envelope, and the following detail requirements shall apply:

(a) Adequate wing torsional rigidity shall be demonstrated by tests or other methods found suitable by the Administrator.

(b) The mass balance of surfaces shall be such as to preclude flutter.

(c) The natural frequencies of all main structural components shall be determined by vibration tests or other methods found satisfactory by the Administrator.

§ 3.311-1 Simplified flutter prevention criteria (FAA policies which apply to § 3.311 (a) and (b)).

Compliance with the rigidity and mass balance criteria presented on pages 4 through 12 of CAA Airframe and Equipment Engineering Report No. 45, as corrected February 1952, "Simplified Flutter Prevention Criteria for Personal Type Aircraft", is considered to be an acceptable means of meeting the flutter prevention requirements of § 3.311 (a) and (b) with the following limitations:

(a) The wing and aileron flutter prevention criteria as represented by the wing torsional stiffness and aileron balance criteria are limited to aircraft which do not have large mass concentrations located along their wing span. For example, these criteria are not applicable for wings carrying engines, floats, fuel in outer-panels, etc.

(b) The elevator and rudder balance criteria are limited in application to tail surface configurations which include a

1 Airframe and Equipment Engineering Report No. 45 has been distributed under Aviation Safety Release No. 330, dated December 2, 1949. Copies of the report are available from the Federal Aviation Agency.

Various procedures are available for demonstrating freedom from flutter in complying with the requirements of § 3.311 (a) and (b), namely: performance of a flutter analysis, application of the simplified criteria of § 3.311-1 when applicable, and performance of a flight flutter test by the applicant. It is not recommended that flight flutter tests be used as a general procedure for substantiating freedom from flutter. The performance of a flutter analysis or the application of the simplified flutter criteria are considered preferred procedures due to the hazards involved in flight flutter tests.

(a) Acceptability. Flight flutter tests will be acceptable as substantiation of freedom from flutter when it can be demonstrated by such tests that proper and adequate attempts to induce flutter have been made within the speed range up to Va, and the vibratory response of the structure during the tests indicates freedom from flutter.

(b) Records. Flight recording instrumentation of either the electrical or photographic type should be installed to provide a permanent record of the control surface and/or fixed surface response to the applied flutter exciting forces, as well as a record of the associated test airspeed.

(c) Test procedures.

The following

is an outline of an acceptable procedure for demonstrating freedom from flutter by flight tests in which rapid control surface deflections are applied to induce flutter:

(1) The tests should cover the flight speed range with excitation applied at small incremental increases of airspeed up to Va. The incremental speed increases between 0.8 Va and Va should be not more than 5 m. p. h. At lower flight speeds, larger increments or airspeed may be used.

(2) The controls should be deflected to attempt to excite flutter as follows: (i) Aileron control to induce wing and aileron flutter.

(ii) Rudder control to induce rudder and vertical tail flutter.

(iii) Elevator control to induce elevator and horizontal tail flutter, and symmetric wing flutter.

(iv) Control surface to which the tab is attached to induce tab flutter.

(3) Attempts to induce flutter should be made by abrupt rotational deflections of the respective control surfaces. These deflections should be obtained by striking the corresponding control with the free hand, or foot, and the disturbed control should be allowed to stabilize without restraint by the pilot. The force applied should be sufficient to produce an impulsive deflection of the control surface of at least 3 degrees.

(4) At each test speed, at least 3 attempts to induce flutter should be made for each of the surfaces being investigated.

(5) A permanent record1 at each test speed should be obtained as follows:

(i) In flutter tests of the rudder, elevator and tab surfaces; a time history of the control surface rotational deflection, and the associated airspeed.

(ii) In flutter tests of the wing and aileron; a time history of the aileron rotational deflection, a time history of the wing vibratory response, and the associated test airspeed.

(6) The tests should consider significant variations in mass and rigidity values which might be expected in service. The aileron and tab control systems should be freed to the extent necessary to be representative of what might be expected in service. (See Airframe and Equipment Engineering Report No. 45.) ' In tests of wings incorporating wing-tip

1 These data can be obtained by installing a control surface position indicating device at the control surface, a vibration pickup in the vicinity of one wing tip to detect the wing response, and a suitable pressure transducer connected to the aircraft's pitot-static system to measure the airspeed with the electrical signals from these instruments connected to a recording oscillograph.

Alternatively, photographic methods using cameras may also be used for recording the flight test data. However, the photographic records obtained should be of a nature that will permit satisfactory evaluation of the degree of control surface deflection applied. the flutter stability and a correlation of the airspeed record with the associated flutter test point.

2 Airframe and Equipment Engineering Report No. 45 is available, free of charge, from the Federal Aviation Agency, Aviation Information Office, W-47, Washington 25, D.C.

fuel tanks, the tip-tank weight during the tests should include the weight which results in the most adverse wing flutter characteristics. An engineering evaluation of the ground vibration survey results may be made to determine which value of tip-tank weight is most critical. In some instances, such an engineering evaluation will not permit selection of any single weight as being most critical and flight flutter tests at several tip-tank weights should be performed.

(7) The tests should be conducted at altitudes of approximately 50 to 75 percent of the service ceiling.

(8) The vibration survey required under § 3.311 (c) should be conducted prior to performing the flight flutter tests. The structural frequency and vibration mode data should be evaluated to determine what structural modes are most likely to be flutter critical. An E engineering evaluation of the wing bending and torsion frequencies as well as the ratios of wing bending to wing torsion frequencies should be made for various wing tip-tank weight conditions to determine the critical tip-tank weight(s) which should be considered in the flight flutter tests.

[Supp. 24, 21 F. R. 5188, July 12, 1956]

Control surface hinges, excepting ball and roller bearings, shall incorporate a multiplying factor of safety of not less than 6.67 with respect to the ultimate bearing strength of the softest material used as a bearing. For hinges incorporating ball or roller bearings, the approved rating of the bearing shall not be exceeded. Hinges shall provide sufficient strength and rigidity for loads parallel to the hinge line.

§ 3.330 Mass balance weights.

The supporting structure and the attachment of concentrated mass balance weights which are incorporated on control surfaces shall be designed for the following limit accelerations: 24g normal to the plane of the control surface, 12g fore and aft, and 12g parallel to the hinge line.

§ 3.336 Primary flight controls.

(a) Primary flight controls are defined as those used by the pilot for the immediate control of the pitching, rolling, and yawing of the airplane.

(b) For two-control airplanes the design shall be such as to minimize the likelihood of complete loss of the lateral directional control in the event of failure of any connecting or transmitting element in the control system.

§ 3.336-1

Aileron controls for two-control airplanes (FAA interpretations which apply to § 3.336(b)).

In the case of two-control airplanes having side by side control wheels, the aileron controls in the right wing should be independent of those in the left wing; however, they may be connected to a common bell-crank or lever in the fuselage. From the point of common connection to the control wheels, the margins of safety should be large and the detail design adequate to minimize possibility of failure of any part. [Supp. 10, 16 F. R. 3289, Apr. 14, 1951]

§ 3.337 Trimming controls.

Proper precautions shall be taken against the possibility of inadvertent, improper, or abrupt tab operations. Means shall be provided to indicate to the pilot the direction of control movement relative to airplane motion and the position of the trim device with respect to the range of adjustment. The means used to indicate the direction of the control movement shall be adjacent to the control, and the means used to indicate the position of the trim device shall be easily visible to the pilot and so located and operated as to preclude the possibility of confusion. Longitudinal trimming devices for single-engine airplanes and longitudinal and directional trimming devices for multiengine airplanes shall be capable of continued normal operation notwithstanding the failure of any one connecting or transmitting element in the primary flight control system. Tab controls shall be irreversible unless the tab is properly balanced and possesses no unsafe flutter characteristics. Irreversible tab systems shall provide adequate rigidity and reliability in the portion of the system from the tab to the attachment of the irreversible unit to the airplane structure.

§ 3.337-1

Independence of bungee trim system from primary control system (FAA interpretations which apply to § 3.337).

The sentence "Trimming devices shall be capable of continued normal operation notwithstanding the failure of any one connecting or transmitting element in the primary flight control system" permits a bungee system acting directly on the control surface horn in such a manner that a failure of the primary control system will not adversely affect operation of the trimming device. A bungee system which actuates a control surface through elements of the primary control system does not meet these requirements. Each trim control system will be reviewed on the basis of its individual merits.

[Supp. 25, 22 F. R. 135, Jan. 5, 1957]

§ 3.337-2 Electrical trim tab systems (FAA policies which apply §§ 3.337 and 3.681).

to

(a) General. Electrical trim tab systems (electrically actuated and manually controlled) must conform to the general provisions in §§ 3.337 and 3.681. In showing compliance with these general provisions, the flight and electrical standards in paragraphs (b) and (c) of this section should be applied.

(b) Flight standards. For each applicable malfunction described in paragraph (c) of this section, the following should be demonstrated, at critical weights and center of gravity positions. without danger of exceeding any of the structural limits or placard speeds. Malfunctions which do not affect the operation of the tab need not be considered.

(1) When corrective action (confined to operation of the primary surface controls) is delayed two seconds after the malfunction has been detected, dangerous deviations from any normal flight condition, including take-off and landing, should not be produced.

(2) Assuming the malfunction to occur during any normal flight conditions, it should be possible to:

(i) Control the airplane readily and easily for a prolonged period of time. without requiring undue effort or concentration on the part of the pilot.

(ii) Perform all the maneuvers and operations necessary in effecting a safe landing.

(c) Electrical standards. Means should be provided in the system design to comply with the flight standards in paragraph (b) of this section after the occurrence of any single malfunction described in this paragraph:

(1) Wiring malfunctions. (1) A ground fault on any electric cable in the system.

(ii) An open circuit in any electric cable in the system.

= (iii) Inadvertent application of electric power to any cable, unless this cable is physically isolated from all other cables or equipment capable of supplying electric power, or unless this cable is covered with a grounded shield.

(2) Switch and relay malfunctions. (i) Internal failure of any switch such that the contacts remain in the open or closed position, or such that any contact is grounded.

(ii) Internal failure of any relay, such that normal actuation or release of the relay contacts cannot be accomplished, or such that any contact is grounded.

(iii) Inadvertent application of electric power to any single exposed contact, or its extension, within any relay unless the contact is physically isolated from adjacent contacts or other equipment capable of supplying electric power.

(iv) Inadvertent application of electric power to any single external terminal of any switch or relay, unless the terminal is physically isolated from adjacent terminals or other equipment capable of supplying electric power.

(3) Connector and terminal strip malfunctions. (i) Inadvertent application of electric power to any single pin of any connector, unless all other pins in the connector are incapable of supplying electric power.

(ii) Inadvertent application of electric power to any single terminal on any terminal strip, unless the terminal is physically isolated from adjacent terminals or other equipment capable of supplying electric power.

(4) Motor malfunctions. (i) Internal electrical or mechanical failure of the trim tab drive motor, such that the trim tab cannot be driven in one or both directions under normal load.

(5) Electric power system malfunctions. (i) Failure of electric power for any reason, or opening of the master switch due to some other emergency.

(6) Other malfunctions similar to those listed, affecting parts of the system not specifically mentioned in this paragraph.

(d) Maintenance check procedures. In complying with paragraphs (b) and (c) of this section, designers may resort to the use of duplicate equipment or other design features of such nature that a single malfunction can remain undetected in normal operation of the aircraft. One example of this condition would be a system using two control switches in series and actuated by the same toggle. Failure of one switch in the closed position would not affect the operation of the trim tab and could therefore remain undetected. The system is then vulnerable to subsequent malfunction of the other switch. Therefore, when the system design is such that a single malfunction would not be detected in normal operation of the aircraft, a maintenance check procedure should be established to assure that the system is free of such malfunction before each flight.

[Supp. 15, 18 F. R. 5646, Sept. 22, 1953] § 3.337-3

Trim device indications (FAA policies which apply to § 3.337).

In addition to providing means to indicate to the pilot the position of the trim device with respect to the range of adjustment, provisions should also be incorporated to indicate when the trimming surfaces are in the neutral position with respect to the primary control surfaces.

[Supp. 26, 22 F. R. 1025, Feb. 20, 1957] § 3.338 Wing flap controls.

The controls shall be such that when the flap has been placed in any position upon which compliance with the performance requirements is based, the flap will not move from that position except upon further adjustment of the control or the automatic operation of a flap load limiting device. Means shall be provided to indicate the flap position to the pilot. If any flap position other than fully retracted or extended is used to show compliance with the performance requirements, such means shall indicate