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] 8 3.174–8 Unusual test situations (FAA policies which apply to 8 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, 1961] FLIGHT LOADS cluded in the test loads. In an interediate case, a factor between 1.0 and 25 may be employed in wing section sts if it is suitably established that a duction from 1.25 is warranted by the rticular conditions of the test. upp. 10, 16 F. R. 3286, Apr. 14, 1961) 3.174–7 Establishment of material strength properties and design values by static test (FAA policies which apply to 8 3.174). (a) There are several types of maial design allowables, all of which are ived from test data. These are: 1) Minimum acceptable values based a minimum value already in an apcable materials procurement specifiion. 2) Minimum non-specification values ived from tests of a series of standard cimens. 3) Ninety percent probability values ich are the lowest strength values exted in 90 percent of the specimens ed. !) Values based on "premium selec" of the material. b) Where testing is used to determine of these types of allowables, procees outlined in existing Government or istry specifications, e. g. QQ-M–151, M's, etc., should be used although ir procedures if approved by the FAA, be used. No clear-cut rules as to the nt of testing to be done can be estab:d in this section, as this usually va with the case. It is therefore a ter for joint discussion between the ufacturer and the FAA. The re\, however, should be based on a sufatly large number of tests of the maI to establish minimum acceptable robability values on a statistical $ 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 8.$ 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) Design values pertinent to the sin paragraphs (a) (1), (2) and (3) Lis section are presented in ANC-5 ANC-18 for commonly used ma Is. | With reference to paragraph (a) f this section, some manufacturers indicated a desire to use values er than the established minimum table values even in cases where the use of minimum acceptable 3 is indicated. Such increases will reptable provided that specimens of § 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. LOAD FACTOR, + SPEED V, MPH E(N) |-CNA MAX~ JE (UA) 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. 3-1-(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 by the designer except that they shall subjected to symmetrical maneuvers I not be less than the following values: sulting in the following limit load factor Vc (design cruising speed) except where limited by maximu =38 VW/S (NU) (static) lift coefficients: =42 VW/S (A) (a) The positive maneuvering lo factor specified in § 3.186 at all speel except that for values of W/S greater up to Va, than 20, the above numerical multiply (b) The negative maneuvering lo ing factors shall be decreased linearly with W/S to a value of 33 at W/S=100: factor specified in g 3.186 at speed And further provided, That the required and factors varying linearly with spe minimum value need be no greater than from the specified value at Vc to 0.0 Va for the N category and -1.0 at Val 0.9 Vn actually obtained at sea level. the A and U categories. Va (design dive speed) =1.40 Vc min (N) § 3.186 Maneuvering load factors. =1.50 Vc min (0) =1.55 Vc min (A) (a) The positive limit maneuveri load factors shall not be less than 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+ Categor 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 3.8 and shall not be less than 2.5. design cruising speed specified above.) n=4.4. Categor Vo (design manuevering speed) n=6.0. Categor =V, Vn where: V,=a computed stalling speed with (b) The negative limit maneuver flaps fully retracted at the de load factors shall not be less than sign weight, normally based on the maximum airplane times the positive load factor for th normal force coefficient, CNA: and U categories, and shall not be n=limit maneuvering load factor than -0.5 times the positive load fax used in design, for the A category. except that the value of Vp need not ex. (c) Lower values of maneuvering ceed the value of Vc used in design. factor may be employed only if it 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 Va. 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) 2.67 =1.33 S16 ): 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 lit curve, CL per radian, corrected for aspect ratio. W/S=wing loading, p. 8. I. 8 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 $3.216 and 3.217) shall be reacted by angular inertia of the complete airplane in a rational or conservative manner. $ 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. Vi 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 V sr=the computed stalling speed with flaps 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. $ 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 $ 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) UNSYMMETRICAL FLIGHT CONDITIONS $ 3.191 Unsymmetrical flight conditions. The airplane shall be assumed to be subjected to rolling and yawing maneuvers as described in the following conditions. Unbalanced aerodynamic moments about the center of gravity shall be reacted in a rational or conservative manner considering the principal masses furnishing the reacting inertia forces. (a) Rolling conditions. The airplane shall be designed for (1) unsymmetrical wing loads appropriate to the category, and (2) the loads resulting from the aileron deflections and speeds specified in $ 3.222, in combination with an airplane load factor of at least two-thirds of the positive maneuvering factor used in the design of the airplane. Only the wing and wing bracing need be investigated for this condition. NOTE: Unless the Administrator finds such data result in unrealistic loads, these conditions may be covered as follows: (a) Rolling accelerations may be obtained by modifying the symmetrical flight conditions shown in Figure 3–1 as follows: (1) Acrobatic category. In conditions A and F assume 100 percent of the wing air load acting on one side of the plane of symmetry and 60 percent on the other. (2) Normal and utility categories. In condition A, assume 100 percent of the wing air load acting on one side of the airplane and 70 percent on the other. For airplanes over 1,000 pounds design weight, the latter percentage may be increased linearly with weight up to 80 percent at 25,000 pounds. (b) The effect of alleron displacement on wing torsion may be accounted for by adding the following increment to the basic airfoil moment coefficient over the aileron portion of the span in the critical condition as determined by the note under $ 3.222: Acm= –.018 where: Acm=moment coefficient increment grees in critical condition $ 3.191-1 Aileron rolling conditions (FAA policies which apply to $ 3.191(a)). In determining whether airplanes of small to medium size and speed comply with $ 31.91(a), the Administrator will accept the following simplified procedure: (a) Steady roll. Determine the Cm value, corresponding to 23 the symmetrical maneuvering load factor. The Cn distribution over the span may be assumed the same as that for the symmetrical flight conditions. Modify the wing movement coefficient over the aileron portions of the span, as described in the “Note" under $ 3.191 (a), corresponding to the required aileron deflections. The wing may be critical in torsion on the up, as well as on the down aileron side, depending upon airfoil section, elastic axis location, aileron differential, etc. (For the up aileron, the moment coefficient increment will be positive.) The above assumption concerning Co distribution implies that the aerodynamic damping forces have exactly the same distribution as the rolling forces, which is not strictly correct. However, since the load factor in the rolling conditions is only 23 of that in the symmetrical conditions, the errors involved in this assumption are not likely to be significant. (b) Maximum angular acceleration. This condition need be investigated only for wings carrying large mass items outboard. In such cases instantaneous aileron deflection (zero rolling velocity) may be assumed and the local value of Cn and Cm over the aileron portions of the span modified accordingly to obtain the spanwise airload distribution. The average Cn of the entire wing should correspond to 33 of the symmetrica) maneuvering load factor. The resultin rolling moment should be resisted by the rolling inertia of the entire airplane This procedure is, in general, conservative, and a more rational investigation based on the time history of the contro movement and response of the airplane may be used if desired. (Supp. 10, 16 F. R. 3285, Apr. 14, 1951) (b) Yawing conditions. The airplane shall be designed for the yawing loads resulting from the vertical surface loads specified in $$ 3.219 to 3.221. (21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-3, 23 F.R. 2589, Apr. 19, 1958) SUPPLEMENTARY CONDITIONS $ 3.194 Special condition for rear lift truss. When a rear lift truss is employed, it shall be designed for conditions of reversed airflow at a design speed of: V=10VW/S +10 (m. p. h.) NOTE: It may be assumed that the value of Cų is equal to -0.8 and the chordwise distributior is triangular between a peak at the trailing edge and zero at the leading edge. $ 3.195 Engine torque effects. (a) Engine mounts and their supporting structures shall be designed for engine torque effects combined with certain basic flight conditions as described in subparagraphs (1) and (2) of this paragraph. Engine torque may be neglected in the other flight conditions. (1) The limit torque corresponding to take-off power and propeller speed act ing simultaneously with 75 percent of į tbe limit loads from flight condition A. (See Fig. 3–1.) (2) The limit torque corresponding to i maximum continuous power and propel ler speed, acting simultaneously with the limit loads from flight condition A. (See Fig. 3-1.) (b) The limit torque shall be obtained by multiplying the mean torque by a factor of 1.33 in the case of engines having 5 or more cylinders. For 4-, 3-, and 2-cylinder engines, the factors shall be 2, 3, and 4, respectively. $ 3.196 Side load on engine mount. $ The limit load factor in a lateral direction for this condition shall be at least pe equal to one-third of the limit load factor for flight condition A (see Fig. 3–1) except that it shall not be less than 1.33. Engine mounts and their supporting structure shall be designed for this condition which may be assumed independent of other flight conditions. tial loads from zero up to the maximum relief valve setting. Account shall be taken of the external pressure distribution in flight. Stress concentrations shall be taken into account in the design of the pressurized structure. (See $ 3.270.) (b) If landings are to be permitted with the cabin pressurized, landing loads shall be combined with pressure differential loads from zero up to the maximum permitted during landing. (c) The airplane structure shall have sufficient strength to withstand the pressure differential loads corresponding with the maximum relief valve setting multiplied by a factor of 1.33. It shall be acceptable to eliminate all other loads in this case. (d) Where a pressurized cabin is separated into two or more compartments by bulkheads or floor, the primary structure shall be designed for the effects of sudden release of pressure in any compartment having external doors or windows. This condition shall be investigated for the effects resulting from the failure of the largest opening in a compartment. Where intercompartment venting is provided, it shall be acceptable to take into account the effects of such venting. (Amdt. 3-2, 22 F. R. 6561, July 16, 1957) CONTROL SURFACE LOADS $ 3.211 General. The control surface loads specified in the following sections shall be assumed to occur in the symmetrical and unsymmetrical flight conditions as described in 88 3.189–3.191. See Figures 3-3 to 3-10 for acceptable values of control surface loadings which are considered as conforming to the following detailed rational requirements. NOTE: For a seaplane version of a landplane, it is normally acceptable to use the wing loading of the landplane in determining the limit maneuvering control surface loadings from Figure 3-3 (b) provided: the power of the engines and the placard maneuver speed of the seaplane do not exceed those established for the landplane; the maximum certificated weight of the seaplane does not exceed the corresponding weight of the landplane by more than 10 percent; and service experience with the landplane is such that no evidence of any serious control-surface load problems is indicated and is such that the service experience is of sufficient scope to deduce with reasonable accuracy that no |