§3.184 Design air speeds. The design air speeds shall be chosen by the designer except that they shall not be less than the following values: Ve (design cruising speed) =38 VW/S (NU) =42 VW/S (4) except that for values of W/S greater than 20, the above numerical multiplying factors shall be decreased linearly with W/S to a value of 33 at W/S=100: And further provided, That the required minimum value need be no greater than 0.9 Vn actually obtained at sea level. Va (design dive speed) except that for values of W/S greater §3.185 Maneuvering envelope. The airplane shall be assumed to be subjected to symmetrical maneuvers resulting in the following limit load factors, except where limited by maximum (static) lift coefficients: (a) The positive maneuvering load factor specified in §3.186 at all speeds up to Va, (b) The negative maneuvering load factor specified in §3.186 at speed Ve;) and factors varying linearly with speed from the specified value at Ve to 0.0 at Va for the N category and -1.0 at Va for the A and U categories. §3.186 Maneuvering load factors. (a) The positive limit maneuvering load factors shall not be less than the following values: n=4.4. n=6.0__ (b) The negative limit maneuvering load factors shall not be less than -0.4' times the positive load factor for the N and U categories, and shall not be less than -0.5 times the positive load factor for the A category. (c) Lower values of maneuvering load factor may be employed only if it be a 720 3 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: For purposes of gust load computations ✓ as required in § 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. FLAPS EXTENDED FLIGHT CONDITIONS §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 following 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. V1 shall be assumed not less than 1.4 Vs or 1.8 Vst, whichever is greater, where: V, the computed stalling speed with flaps fully retracted at the design weight V1f 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 V., 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 aileron 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 8=down aileron deflection in degrees in critical condition (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] §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 C value, corresponding to 3 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 C 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 % 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 % of the symmetrical maneuvering load factor. The resulting 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] The 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 Cz is equal to -0.8 and the chordwise distribution 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 acting simultaneously with 75 percent of the limit loads from flight condition A. (See Fig. 3-1.) (2) The limit torque corresponding to maximum continuous power and propeller 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 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. SUPPLEMENTARY CONDITIONS §3.197 Pressurized cabin loads. The provisions of paragraphs (a) through (d) of this section shall apply to pressurized compartments. (a) The airplane structure shall have sufficient strength to withstand the flight loads combined with pressure differen (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. 5561, July 16, 1957] CONTROL SURFACE LOADS The control surface loads specified in the following sections shall be assumed to occur in the symmetrical and unsymmetrical flight conditions as described in §§ 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 serious control-surface load problems will develop on the seaplane. [21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-2, 22 F.R. 5561, July 16, 1957] § 3.211-1 Control surface loads for design of "Vee" type tail assemblies (FAA policies which apply to § 3.211). (a) "Vee" type tail assemblies require special design criteria in order to show "the same level of safety" under § 3.10. Thus, for "Vee" type tail assemblies, all the tail load requirements as set forth in this part are considered acceptable to this type tail design. It will be necessary, however, to increase the unit loads on each side of the tail surface to account for the tail surface dihedral, since air loads act normal to the surface only. Thus the unit loads, based on the projected area, on each side of the tail surface due to vertical loads on the tail assembly should be increased by a factor equal to 1/cos Ø, while the unit horizontal loads on the tail assembly should be increased by a factor equal to 1/sin 0, where is the dihedral angle, or the angle between each side of the tail surface and the horizontal. (b) The following supplementary conditions should also be investigated: (1) A ±30 fps gust, acting normal to the chord plane of one side of the tail surface at Vc, should be combined with a one "g" balancing tail load. Reduction for downwash is acceptable. It is evident that this condition will be unsymmetrical, since one side of the "Vee" tail will not be highly loaded by the gust. (2) Combined rudder and elevator maneuvering condition. (i) In order to obtain the full one way travel of the ruddervator, it is desirable to have full elevator travel in conjunction with full rudder travel. The limiting factor for this configuration is 3 elevator load for one pilot, and 3 rudder load for one pilot applied simultaneously. (ii) When it can be shown that the lateral gust condition (reference: § 3.220) is less critical than the condition in subparagraph (1) of this paragraph, no analysis for the lateral gust need be made. [Supp. 10, 16 F. R. 3286, Apr. 14, 1951] § 3.212 Pilot effort. In the control surface loading conditions described, the airloads on the movable surfaces and the corresponding deflections need not exceed those which could be obtained in flight by employing the maximum pilot control forces specified in Figure 3-11. In applying this criterion, proper consideration shall be given to the effects of control system boost and servo mechanisms, tabs, and automatic pilot systems in assisting the pilot. § 3.212-1 Automatic pilot systems (FAA policies which apply to § 3.212). The Administrator will accept the following procedure as giving proper consideration of automatic pilot systems in assisting the pilot under § 3.212: The autopilot effort need not be added to human pilot effort but the autopilot effort shall be used for design if it alone can produce greater control surface loads than the human pilot. [Supp. 1, 12 F. R. 3435, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5. 1949] |