placard flap speed or limit load factor, it is permissible to retract the flaps during recovery to avoid exceeding these limits. [Supp. 10, 16 F. R. 3284, Apr. 14, 1951]

§3.124-2 Spin tests for category A airplanes (FAA interpretations which apply to §3.124(c)).

If during recovery from a one-turn flaps-down spin the airplane exceeds the placard flap speed or limit load factor, it is permissible to retract the flaps during recovery to avoid exceeding these limits. In addition the airplane is to be placarded "International spins with flaps down prohibited." [Supp. 10, 16 F. R. 3284, Apr. 14, 19511

GROUND AND WATER CHARACTERISTICS §3.143 Requirements.

All airplanes shall comply with the requirements of §§ 3.144 to 3.147. §3.144 Longitudinal stability and control.

There shall be no uncontrollable tendency for landplanes to nose over in any operating condition reasonably expected for the type, or when rebound occurs during landing or take-off. Wheel brakes shall operate smoothly and shall exhibit no undue tendency to induce nosing over. Seaplanes shall exhibit no dangerous or uncontrollable porpoising at any speed at which the airplane is normally operated on the water.

§3.145 Directional stability and control.

(a) There shall be no uncontrollable looping tendency in 90-degree cross winds up to a velocity equal to 0.2 Vo at any speed at which the aircraft may be expected to be operated upon the ground or water.

(b) All landplanes shall be demonstrated to be satisfactorily controllable with no exceptional degree of skill or alertness on the part of the pilot in power-off landings at normal landing speed and during which brakes or engine power are not used to maintain a straight path.

(c) Means shall be provided for adequate directional control during taxying.

§3.146 Shock absorption.

The shock-absorbing mechanism shall not produce damage to the structure

when the airplane is taxied on the roughest ground which it is reasonable to expect the airplane to encounter in normal operation.

§3.147 Spray characteristics.

For seaplanes, spray during taxiing, takeoff, and landing shall at no time dangerously obscure the vision of the pilots nor produce damage to the propeller or other parts of the airplane.

FLUTTER AND VIBRATION

§3.159 Flutter and vibration.

All parts of the airplane shall be demonstrated to be free from flutter and excessive vibration under all speed and power conditions appropriate to the operation of the airplane up to at least the minimum value permitted for Va in §3.184. There shall also be no buffeting condition in any normal flight condition severe enough to interfere with the satisfactory control of the airplane or to cause excessive fatigue to the crew or result in structural damage. However, buffeting as stall warning is considered desirable and discouragement of this type of buffeting is not intended.

Subpart C-Strength Requirements GENERAL

(a) Strength requirements are specified in terms of limit and ultimate loads. Limit loads are the maximum loads anticipated in service. Ultimate loads are equal to the limit loads multiplied by the factor of safety. Unless otherwise described, loads specified are limit loads.

(b) 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 airplane. All such loads shall be distributed in a manner conservatively approximating closely representing actual conditions. If deflections under load would change significantly the distribution of external or internal loads, such redistribution shall be taken into account.

or

(c) Simplified structural design criteria shall be acceptable if the Administrator finds that they result in design loads not less than those prescribed in §§ 3.181 through 3.265.

§3.171-1 Design criteria (FAA policies which apply to § 3.171(c)).

The Administrator finds that the simplified structural design criteria contained in Appendix A1 to Civil Aeronautics Manual 3, result in design loads not less than those prescribed in §§ 3.181 through 3.265.

[Supp. 16, 17 F. R. 11786, Dec. 30, 1952]

§3.171-2 Design loads and load distributions (FAA policies which apply to §3.171(b)).

The simplified method in Appendix D1 to Civil Aeronautics Manual 3 may be used to determine the air loads and air load distributions resulting from the use of tip stores for low speed, low altitude (design Mach number less than 0.4; design altitude less than 15,000 ft.) airplanes with small amounts of sweep (i.e., mid-chord angles of sweep less than 15 degrees).

[Supp. 30, 22 F. R. 10016, Dec. 13, 1957]

§3.172 Factor of safety.

The factor of safety shall be 1.5 unless otherwise specified.

§3.173 Strength and deformations.

The structure shall be capable of supporting limit loads without suffering detrimental permanent deformations. At all loads up to limit loads, the deformation shall be such as not to interfere with safe operation of the airplane. The structure shall be capable of supporting ultimate loads without failure for at least 3 seconds, except that when proof of strength is demonstrated by dynamic tests simulating actual conditions of load application, the 3-second limit does not apply.

§3.173-1 Dynamic tests (FAA policies

which apply to § 3.173).

(a) Section 3.173 permits dynamic testing in lieu of stress analysis or static testing in the proof of compliance of the structure with strength and deformation requirements. In demonstrating, by dynamic tests, proof of strength of landing gears for the stipulated landing conditions contained in §§ 3.245, 3.246, and 3.247, it is necessary to employ a procedure which will not result in the accepting of landing gears weaker than those

1 Not filed for publication in the Office of the Federal Register.

qualified for acceptance under present procedures, i.e., stress analysis or static testing.

(b) The Administrator will accept, as an adequate procedure for this purpose, the following dynamic tests:

The structure shall be dropped a minimum of 10 times from the limit drop height, and at least one time from the be ultimate drop height, for each basic design condition for which proof of strength is being made by drop tests.

(c) With regard to the extent to which the structure can be proved by dynamic tests, such dynamic tests shall be accepted as proof of strength for only those elements of the structure for which it can be shown that the critical limit and ultimate loads have been reproduced. [Supp. 1, 12 F. R. 3435, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5, 1949]

Proof of compliance of the structure with the strength and deformation requirements of § 3.173 shall be made for all critical loading conditions. Proof of compliance by means of structural ( analysis will be accepted only when the structure conforms with types for which experience has shown such methods to be reliable. In all other cases substantiating load tests are required. Dynamic tests including structural flight tests shall be acceptable, provided that it is demonstrated that the design load conditions have been simulated. In all cases certain portions of the structure must be subjected to tests as specified in Subpart D of this part.

§ 3.174-1 Material correction factors (FAA policies which apply to § 3.174).

(a) In tests conducted for the purpose of establishing allowable strengths of structural elements such as sheet, sheet stringer combinations, riveted joints. etc., test results should be reduced to values which would be met by elements of the structure if constructed of materials having properties equal to design allowable values. Material correction factors in this case may be omitted, however, if sufficient test data are obtained to permit a probability analysis showing that 90 percent or more of the elements will either equal or exceed in strength the

18

The

selected design allowable values. number of individual test specimens needed to form a basis of "probability values" cannot be definitely stated but must be decided on the basis of consistency of results; i. e., "spread of results", deviations from mean value, and range of sizes, dimensions of specimens, etc., to be covered. This item should therefore be a matter for decision between the manufacturer and the FAA. (Sections 1.654 and 1.655 of ANC-5a 1949 edition outline two means of accomplishing material corrections in element tests; these methods, however, are by no means considered the only methods available.)

2

(b) In cases of static or dynamic tests of structural components, no material correction factor is required. The manufacturer, however, should use care to see that the strength of the component tested conservatively represents the strength of subsequent similar components to be used on aircraft to be presented for certification. The manufacturer should, in addition, include in his report of tests of major structural components, a statement substantially as follows:

The strength properties of materials and dimensions of parts used in the structural component(s) tested are such that subsequent components of these types used in aircraft presented for certification will have strengths substantially equal to or exceeding the strengths of the components tested. [Supp. 6, 15 F. R. 619, Feb. 4, 1950] §3.174-2 Structural testing of new projects (FAA policies which apply to § 3.174).

(a) The following is a general procedure that may be followed for determining the extent of required structural testing of a new project:

(1) As the initial step to determine the structural testing of a new project, a meeting between representatives of the manufacturer, the Federal Aviation Agency project engineer, and (if practicable) the pertinent Branch Chief of the Aircraft Division should be ar

'ANC-5a, "Strength of Aircraft Elements" is published by the Army-Navy-Civil Committee on Aircraft Design Criteria and may be obtained from the Superintendent of Documents, Government Printing Office, Washington 25, D. C.

ranged. The question of minimum tests should be reviewed first. This will include generally such tests as proof and operation tests of control surfaces and systems, drop tests of landing gear, vibration tests, and wing torsional stiffness tests.

(2) If the structure is of a type on which the manufacturer has a thorough background of experience, analysis and proof tests can usually be considered acceptable. If, in addition, the analysis has a high degree of conservatism, proof tests other than those specifically required by regulation may be omitted at the discretion of the FAA.

(b) If the structure or parts thereof are definitely outside the manufacturer's previous experience, the manufacturer may be requested to establish a strength test program. In the case of a wing, this will usually involve a 100 percent ultimate load test for PHAA. In cases of this type, it should be suggested to the manufacturer that he carry the PHAA test to destruction. If a comparison of the effects of inverted and normal types of loading can be carried out, some of the above tests, such as ILAA test, can be omitted and a test made for one condition only.

(c) When ultimate load static tests are made, the limit load need not be removed provided that continuous readings of deflections of the structure are measured at an adequate number of points, and also provided that a close examination of the structure is maintained throughout the tests with particular emphasis being placed upon close observation of the structure at limit load for any indications of local distress, yielding buckles, etc.

(d) In the case of small airplanes of other than two spar and steel tube construction, the manufacturer should be encouraged to strength test his product and reduce formal analysis to а minimum.

[Supp. 10, 16 F. R. 3284, Apr. 14, 1951]

§ 3.174-3 Allowable bending moments of stable sections in the plastic range (FAA policies which apply to § 3.174).

(a) The analytical method for determining allowable bending moments of stable sections in the plastic range as

outlined in "Bending Strength in the Plastic Range" by F. P. Cozzone, Journal of Aeronautical Sciences. May 1943, is satisfactory for general use; however, the following should be considered in the application of this method of analysis to particular problems:

(1) The above method may be unconservative and should not be used for sections subject to local failure unless verified by suitable tests. For example, ANC-5 should be used for round tubing.

(2) The method may be unconservative and should be verified by testing representative cross sections for materials having stress-strain curves differing materially from those discussed in the reference article, or for materials whose stress-strain properties in compression differ materially from those in tension.

[Supp. 10, 16 F. R. 3285, Apr. 14, 1951, as amended by Supp. 14, 17 F. R. 9066, Oct. 11, 1952]

§3.174 4 Acceptability of static and/or dynamic tests in lieu of stress analyses (FAA policies which apply to §3.174).

Static testing to ultimate load is considered an adequate substitute for and in some cases superior to formal stress analysis where static loads are critical in the design of the component. In cases where a dynamic loading is critical dynamic load tests are equivalent to formal stress analysis. An example of components on which dynamic loading is usually critical is the landing gear and landing gear structure of an aircraft. (See §3.174–2.) The same yield criteria apply to dynamic tests as to static tests.

[Supp. 10, 16 F. R. 3285, Apr. 14, 1951] §3.174-5 Operation tests (FAA policies

which apply to § 3.174).

Operation tests of structural components are required for mechanisms and linkages in several of the regulations in this subchapter. For this part these are §§ 3.343 and 3.358.

[Supp. 10, 16 F. R. 3285, Apr. 14, 1951] §3.174-6 Material correction factors,

fitting factors, and other factors; their effect on test loads (FAA policies which apply to § 3.174).

(a) Use of factors to establish design and test loads. This part specifies cer

(f) Hinge and bearing factors. Hinge and bearing factors specified shall be in- rie cluded in tests unless the appropriate m portions of the parts are substantiated m otherwise.

(g) Other factors. Test factors for rib, wing, and wing-covering are as follows:

(1) No additional factors of safety need be applied when rational chordwise upper and lower surface pressure distribution is used, provided that the test includes a complete wing or a section of a wing with end conditions and loadings applied in a manner closely simulating the actual wing conditions.

(2) When a rib alone, a section of wing, or small section of the airplane covering is tested without employing s completely rational load analysis and distribution, a factor of 1.25 should be

included in the test loads. In an intermediate case, a factor between 1.0 and 1.25 may be employed in wing section tests if it is suitably established that a reduction from 1.25 is warranted by the particular conditions of the test. [Supp. 10, 16 F. R. 3285, Apr. 14, 1951]

§3.174-7 Establishment

of material strength properties and design values by static test (FAA policies which apply to § 3.174).

(a) There are several types of material design allowables, all of which are derived from test data. These are:

(1) Minimum acceptable values based on a minimum value already in an applicable materials procurement specification.

(2) Minimum non-specification values derived from tests of a series of standard specimens.

(3) Ninety percent probability values which are the lowest strength values expected in 90 percent of the specimens tested.

(4) Values based on "premium selection" of the material.

(b) Where testing is used to determine any of these types of allowables, procedures outlined in existing Government or Industry specifications, e. g. QQ-M-151, ASTM's, etc., should be used although other procedures if approved by the FAA, may be used. No clear-cut rules as to the extent of testing to be done can be estabished in this section, as this usually vacies with the case. It is therefore a natter for joint discussion between the nanufacturer and the FAA. The results, however, should be based on a suficiently large number of tests of the maerial to establish minimum acceptable or probability values on a statistical asis.

(c) Design values pertinent to the tems in paragraphs (a) (1), (2) and (3) of this section are presented in ANC-5 and ANC-18 for commonly used maerials.

(d) With reference to paragraph (a) 4) of this section, some manufacturers have indicated a desire to use values reater than the established minimum cceptable values even in cases where nly the use of minimum acceptable alues is indicated. Such increases will e acceptable provided that specimens of

each individual item of basic material as obtained are tested prior to use, to ascertain that the strength properties of that particular item will equal or exceed the properties to be used in design.

[Supp. 10, 16 F. R. 3285, Apr. 14, 1951, as amended by Supp. 14, 17 F. R. 9066, Oct. 11, 1952]

§3.174-8 Unusual test situations (FAA policies which apply to § 3.174).

It should be borne in mind that in any unusual or different situations a conference between the FAA and the manufacturer should be held to determine if the testing program as proposed by the manufacturer is sufficient to substantiate the structural strength of the aircraft or its component.

[Supp. 10, 16 F. R. 3285, Apr. 14, 1951]