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level flight without encountering excessive loss of altitude, uncontrollable rolling characteristics, or uncontrollable spinning tendencies. These qualities shall be demonstrated by performing the following maneuver: After a steady curvilinear level coordinated flight condition in a 30-degree bank is established and while maintaining the 30-degree bank, the airplane shall be stalled by steadily and progressively tightening the turn with the elevator control until the airplane is stalled or until the elevator has reached its stop. When the stall has fully developed, recovery to level flight shall be made with normal use of the controls. $ 3.123 One-engine-inoperative stalls.

Multiengine airplanes shall not display any undue spinning tendency and shall be safely recoverable without applying power to the inoperative engine when stalled with:

(a) The critical engine inoperative,
(b) Flaps and landing gear retracted,

(c) The remaining engines operating at up to 75 percent of maximum continuous power, except that the power need not be greater than that at which the use of maximum control travel just holds the wings laterally level in approaching the stall. The operating engines may be throttled back during the recovery from the stall.

incapable of spinning" in which case they shall be so designated. (See paragraph (d) of this section.)

(b) Category U. Airplanes in this category shall comply with either the entire requirements of paragraph (a) of this section or the entire requirements of paragraph (c) of this section.

(C) Category A. All airplanes in this category shall be capable of spinning and shall comply with the following:

(1) At any permissible combination of weight and center of gravity position obtainable with all or part of the design useful load, the airplane shall recover from a six-turn spin, or from any point in a six-turn spin, in not more than 142 additional turns after the application of the controls in the manner normally used for recovery.

(2) It shall be possible to recover from the maneuver prescribed in subparagraph (1) of this paragraph without exceeding either the limiting air speed or the limit positive maneuvering load factor of the airplane.

(3) It shall not be possible to obtain uncontrollable spins by means of any possible use of the controls.

(4) A placard shall be placed in the cockpit of the airplane setting forth the use of the controls required for recovery from spinning maneuvers.

(d) Category NU. When it is desired to designate an airplane as a type "characteristically incapable of spinning,” the flight tests to demonstrate this characteristic shall also be conducted with:

(1) A maximum weight 5 percent in excess of the weight for which approval is desired,

(2) A center of gravity at least 3 percent aft of the rearmost position for which approval is desired,

(3) An available up-elevator travel 4 degrees in excess of that to which the elevator travel is to be limited by appropriate stops.

(4) An available rudder travel 7 degrees, in both directions, in excess of that to which the rudder travel is to be limited by appropriate stops. $ 3.124–1 Spin tests for category N air

planes (FAA interpretations which

apply to 3.124(a)). If during recovery from a one-turn flaps-down spin the airplane exceeds the


$ 3.124 Spinning.

(a) Category N. All airplanes of 4,000 lbs. or less maximum weight shall recover from a one-turn spin with the controls applied normally for recovery in not more than one additional turn and without exceeding either the limiting air speed or the limit positive maneuvering load factor for the airplane. In addition, there shall be no excessive back pressure either during the spin or in the recovery. It shall not be possible to obtain uncontrollable spins by means of any possible use of the controls. Compliance with these requirements shall be demonstrated at any permissible combination of weight and center of gravity positions obtainable with all or any part of the designed useful load. All airplanes in category N, regardless of weight, shall be placarded against spins or demonstrated to be "characteristically

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 air.

planes (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, 1961)


All airplanes shall comply with the requirements of $8 3.144 to 3.147. § 3.144 Longitudinal stability and con

trol. 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 V80 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 8 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 & 3.171 Loads.

(a) Strength requirements are speci. fied 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 or 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.

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

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 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)

8 3.171-1 Design criteria (FAA policies

which apply to g 3.171(c)). The Administrator finds that the simplified structural design criteria contained in Appendix A' 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 dis

tributions (FAA policies which apply

to 8 3.171(b)). The simplified method in Appendix Di 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) 8 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. 8 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

§ 3.174 Proof of structure.

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

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

elected design allowable values. The lumber of individual test specimens leeded to form a basis of "probability alues" cannot be definitely stated but aust be decided on the basis of consistncy of results; i. e., “spread of results”, leviations from mean value, and range f sizes, dimensions of specimens, etc., o be covered. This item should thereore be a matter for decision between he manufacturer and the FAA. (Secions 1.654 and 1.655 of ANC-5a 1949 ediion outline two means of accomplishng material corrections in element ests; these methods, however, are by no neans considered the only methods vailable.)

(b) In cases of static or dynamic tests f structural components, no material orrection factor is required. The manfacturer, however, should use care to ee that the strength of the component ested conservatively represents the trength of subsequent similar compolents to be used on aircraft to be resented for certification. The manuacturer should, in addition, include in is report of tests of major structural omponents, a statement substantially s follows:

The strength properties of materials and imensions of parts used in the structural omponent(s) tested are such that subseuent components of these types used in ircraft presented for certification will have trengths substantially equal to or exceeding he strengths of the components tested. Supp. 6, 15 F. R. 619, Feb. 4, 1950) 3.174-2 Structural testing of

projects (FAA policies which apply

to § 3.174). (a) The following is a general proedure that may be followed for deteruining the extent of required structural esting of a new project:

(1) As the initial step to determine he structural testing of a new project,

meeting between representatives of he manufacturer, the Federal Aviation gency project engineer, and (if pracicable) the pertinent Branch Chief of he Aircraft Division should be ar

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 a 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


* ANC-5a, "Strength of Alrcraft Elements"

published by the Army-Navy-Civil Comaittee on Aircraft Design Criteria and may e obtained from the Superintendent of Documents, Government Printing Office, Vashington 25, D. C.

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) § 3.1744 Acceptability of static and/or

dynamic tests in lieu of stress anal. yses (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. 3286, 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 8.8 3.343 and 3.358. (Supp. 10, 16 F. R. 3286, Apr. 14, 1951) & 3.174–6 Material correction factors,

fitting factors, and other factors; their effect on test loads (FAA poli

cies which apply to 8 3.174). (a) Use of factors to establish design and test loads. This part specifies cer

tain factors which must be taken into account in establishing design and test loads for structural components. These factors are to be found in the following sections of this part and are discussed in paragraphs (b) through (g) of this section:

(1) $ 3.172 Factor of safety.

(2) $3.301 Material strength properties and design values.

(3) $ 3.304 Castings.
(4) $ 3.305 Bearing factors.
(5) $ 3.318 Ribs.
(6) $ 3.329 Hinges.
(7) $ 3.346 Joints.

(b) Factor of safety of 1.50. In all cases of ultimate load testing the factor of safety of 1.50 should be included in the test load,

(c) Material correction factors. (See $ 3.174–1).

(d) Fitting factor. The additional multiplying factor of safety of 1.15 specified in $ 3.306 need not be included in test loads in which the actual stress conditions are simulated in the fitting and the surrounding structure. Also, these factors are considered to be included in and covered by the other special factors specified in $ 3.302.

(e) Casting factors. Casting factors should be included in all tests in the substantiation of castings. (See $ 3.304-1.)

(f) Hinge and bearing factors. Hinge and bearing factors specified shall be included in tests unless the appropriate portions of the parts are substantiated 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 & completely rational load analysis and distribution, a factor of 1.25 should be

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