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selected design allowable values. The 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.)

(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

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 & 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" is published by the Army-Navy-Civil Committee on Alrcraft Design Criteria and may be obtained from the Superintendent of Documents, Government Printing Office, Washington 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 icular 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. 3286, 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 analyses (FAA policies which apply to

8 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 g 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 poli.

cies which apply to $ 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) 8 3.301 Material strength properties and design values.

(3) $ 3.304 Castings.
(4) 83.305 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 in. cluded in tests unless the appropriate portions of the parts are substantiated otherwise,

(6) 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 lover 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 a completely rational load analysis and distribution, a factor of 1.25 should be 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)

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. 3286, 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 established in this section, as this usually varies with the case. It is therefore a matter for joint discussion between the manufacturer and the FAA. The results, however, should be based on a sufficiently large number of tests of the material to establish minimum acceptable or probability values on a statistical basis.

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

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

8 3.174–8 Unusual test situations (FAA

policies which apply to g 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 suficient to substantiate the structural strength of the aircraft or its component. (Supp. 10, 16 F. R. 3286, Apr. 14, 1961)

FLIGHT LOADS 8 3.181 General.

Flight load requirements shall be complied with at critical altitudes within the range in which the airplane may be expected to operate and at all weights between the minimum design weight and the maximum design weight, with any practicable distribution of disposable load within prescribed operating limitations stated in $.$ 3.777-3.780. § 3.182 Definition of flight load factor.

The flight load factors specified represent the acceleration component (in terms of the gravitational constant g) normal to the assumed longitudinal axis of the airplane, and equal in magnitude and opposite in direction to the airplane inertia load factor at the center of gravity.

SYMMETRICAL FLIGHT CONDITIONS (FLAPS

RETRACTED)

$ 3.183 General.

The strength requirements shall be met at all combinations of air speed and load factor on and within the boundaries of a pertinent V-n diagram, constructed similarly to the one shown in Figure 3-1, which represents the envelope of the flight loading conditions specified by the maneuvering and gust criteria of 8.8 3.185 and 3.187. This diagram will also be used in determining the airplane structural operating limitations as specified in Subpart G of this part.

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|-CNAMAX

JE(UEA)
- MANEUVER
G*

F

* NOTE LIMIT MANEUVER ENVELOPES

POINT G NEED NOT BE INVESTIGATED LIMIT GUST ENVELOPE

WHEN SUPPLEMENTARY CONDITION LIMIT COMBINED ENVELOPE

SPECIFIED IN $ 3.194 IS INVESTIGATED. FIG. 3-14(V-n) DIAGRAM (FLIGHT ENVELOPE) & 3.184 Design air speeds.

$ 3.185 Maneuvering envelope. The design air speeds shall be chosen The airplane shall be assumed to be by the designer except that they shall subjected to symmetrical maneuvers renot be less than the following values: sulting in the following limit load factors, Vc (design cruising speed)

except where limited by maximum =38 VW/S (NU)

(static) lift coefficients: =42 VW/S (A)

(a) The positive maneuvering load

factor specified in $ 3.186 at all speeds except that for values of W/S greater

up to Va, than 20, the above numerical multiply

(b) The negative maneuvering load ing factors shall be decreased linearly with W/S to a value of 33 at W/S=100:

factor specified in § 3.186 at speed Ve; And further provided, That the required

and factors varying linearly with speed minimum value need be no greater than

from the specified value at Vc to 0.0 at 0.9 Vn actually obtained at sea level.

Va for the N category and -1.0 at Va for

the A and U categories.
Va (design dive speed)
=1.40 Vc min (N)

8 3.186 Maneuvering load factors.
=1.50 Vc min (u
=1.55 Vc min

(a) The positive limit maneuvering

load factors shall not be less than the except that for values of W/S greater

following values: than 20, the above numerical multiply

24,000 ing factors shall be decreased linearly n=2.1+

Category N

W + 10,000 with W/S to a value of 1.35 at W/S=100. (Vc min is the required minimum value of except that n need not be greater than

3.8 and shall not be less than 2.5. design cruising speed specified above.)

n=4.4.

Category U V, (design manuevering speed)

n=6.0.

Category A =V, VT where: V,=a computed stalling speed with (b) The negative limit maneuvering flaps fully retracted at the de

load factors shall not be less than -0.4 sign weight, normally based the maximum airplane

times the positive load factor for the N normal force coefficient, CNA.

and U categories, and shall not be less n=llmit maneuvering load factor than -0.5 times the positive load factor used in design,

for the A category. except that the value of Vo need not ex. (c) Lower values of maneuvering load ceed the value of Vc used in design. factor may be employed only if it be

on

FLAPS EXTENDED FLIGHT CONDITIONS

proven that the airplane embodies features of design which make it impossible to exceed such values in flight. (See also $ 3.106.) $ 3.187

Gust envelope. The airplane shall be assumed to encounter symmetrical vertical gusts as specified below while in level flight and the resulting loads shall be considered limit loads:

(a) Positive (up) and negative (down) gusts of 30 feet per second nominal intensity at all speeds up to Vc,

(b) Positive and negative 15 feet per second gusts at Vd. Gust load factors shall be assumed to vary linearly between Vc and Va. § 3.188 Gust load factors.

In applying the gust requirements, the gust load factors shall be computed by the following formula:

KUVm n=1+

575 (W/S) where: K=(W/S)1/4 (for W/S<16 p. s. 1.)

2

2.67 =1.33

§ 3.190 Flaps extended flight conditions.

(a) When flaps or similar high lift devices intended for use at the relatively low air speeds of approach, landing, and take-off are installed, the airplane shall be assumed to be subjected to symmetrical maneuvers and gusts with the flaps fully deflected at the design flap speed V, resulting in limit load factors within the range determined by the fol. lowing conditions:

(1) Maneuvering, to a positive limit load factor of 2.0.

(2) Positive and negative 15-feet-persecond gusts acting normal to the flight path in level flight. The gust load factors shall be computed by the formula of § 3.188.

V, shall be assumed not less than 1.4 V, or 1.8 Vst, whichever is greater, where: Ve=the computed stalling speed with flaps

fully retracted at the design weight Vor=the computed stalling speed with iaps

fully extended at the design weight except that when an automatic flap load limiting device is employed, the airplane may be designed for critical combinations of air speed and flap position permitted by the device. (See also $ 3.338.)

(b) In designing the flaps and supporting structure, slipstream effects shall be taken into account as specified in $ 3.223.

NOTE: In determining the external loads on the airplane as a whole, the thrust, slipstream, and pitching acceleration may be assumed equal to zero.

(W/S):7/(for W/S> 16p.8.I.)

U=nominal gust velocity, 1. p. 8.

(Note that the "effective sharp

edged gust” equals KU.) V=airplane speed, m. p. h. m=slope of it curve, CL per radian,

corrected for aspect ratio. W/S=wing loading, p. 8. 1. § 3.188-1 "Slope of lift curve” (FAA interpretations which apply

to $ 3.188). For purposes of gust load computations as required in g 3.188 the slope of the lift curve may be assumed equal to that of the wing alone. (Supp. 1, 12 F. R. 3435, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5, 1949) $ 3.189 Airplane equilibrium.

In determining the wing loads and linear inertia loads corresponding to any of the above specified flight conditions, the appropriate balancing horizontal tail load (see $ 3.215) shall be taken into account in a rational or conservative manner.

Incremental horizontal tail loads due to maneuvering and gusts (see $8 3.216 and 3.217) shall be reacted by angular inertia of the complete airplane in a retional or conservative manner.

$ 3.190–1 Design flap speed V (FAA

interpretations which apply to

$ 3.190(a)). (a) The minimum permissible speed of 1.8 Vor is specified in order to cover power-off flight tests as required by § 3.115(a). Section 3.223 requires that slipstream effects be considered in the design of the flaps and operating mechanism up to a speed of at least 1.4 V, in order to cover the power on flight tests of g 3.109(b) (5).

(b) The designer may treat the foregoing conditions as two separate cases, or he may combine them if he so desires. (Supp. 10, 16 F. R. 3285, Apr. 14, 1951)

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