= =gust alleviation factor; airplane mass ratio; Ude derived gust velocities referred to in subparagraphs (1) through (3) of this paragraph (fps); p=density of air (slugs/cu. ft.); W/S=wing loading (psf); C=mean geometric chord (ft.); V = airplane equivalent speed (knots); coefficient curve CNA per radian if the gust loads are applied to the wings and horizontal tail surfaces simultaneously by a rational method. It shall be acceptable to use the wing lift curve slope CL per radian when the gust load is applied to the wings only and the horizontal tail gust loads treated as a separate condition. [15 F. R. 3543, June 8, 1950; 15 F. R. 4171, June 29, 1950, as amended by Amdt. 4b-6, 17 F. R. 1089, Feb. 5, 1952; Amdt. 4b-3, 21 F. R. 990, Feb. 11, 1956] § 4b.212 Effect of high lift devices. are (1) Maneuvering to a positive limit load factor of 2.0, (2) Positive and negative 25 fps derived gusts acting normal to the flight path in level flight. (b) When flaps or similar high lift devices are intended for use in en route conditions (e. g., as speed brakes) the airplane shall be assumed to be subjected to symmetrical maneuvers and gusts, with flaps in the appropriate position at speeds up to the VFE speed established in accordance with § 4b.714 (c), resulting in limit load factors, within the range determined by the following conditions: (1) Maneuvering to a positive limit load factor of 2.5, (2) Positive and negative derived gusts as prescribed in § 4b.211 (b) acting normal to the flight path in level flight. (c) In designing flaps and supporting structure on tractor type airplanes, slipstream effects shall be taken into account as specified in § 4b.221. For other than tractor type airplanes, a head-on gust equivalent to the intensity prescribed in § 4b.211 (b) (3) with no alleviations acting along the flight path shall be considered. (d) When automatic flap operation is provided, the airplane shall be designed for the speeds and the corresponding flap positions which the mechanism permits. (See § 4b.323.) [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-6, 17 F. R. 1089, Feb. 5, 1952; Amdt. 4b-2, 20 F. R. 5305, July 26, 1955; Amdt. 4b-3, 21 F.R. 991, Feb. 11, 1956; Amdt. 4b-6, 22 F.R. 5564, July 16, 1957] § 4b.213 Symmetrical flight conditions. (a) Procedure of analysis. In the analysis of symmetrical flight conditions at least those specified in paragraphs (b), (c), and (d) of this section shall be considered. The following procedure of analysis shall be applicable: (1) A sufficient number of points on the maneuvering and gust envelopes shall be investigated to insure that the maximum load for each part of the airplane structure is obtained. It shall be acceptable to use a conservative combined envelope for this purpose. (2) All significant forces acting on the airplane shall be placed in equilibrium in a rational or a conservative manner. The linear inertia forces shall be considered in equilibrium with wing and horizontal tail surface loads, while the angular (pitching) inertia forces shall be considered in equilibrium with wing and fuselage aerodynamic moments and horizontal tail surface loads. (3) Where sudden displacement of a control is specified, the assumed rate of displacement need not exceed that which actually could be applied by the pilot. (4) In determining elevator angles and chordwise load distribution in the maneuvering conditions of paragraphs (b) and (c) of this section in turns and pullups, account shall be taken of the effect of corresponding pitching velocities. (b) Maneuvering balanced conditions. The maneuvering conditions A through I on the maneuvering envelope (fig. 4b2) shall be investigated, assuming the airplane to be in equilibrium with zero pitching acceleration. (c) Maneuvering pitching conditions. The following conditions involving pitching acceleration shall be investigated (see figure 4b-2): (1) Maximum elevator displacement at speed VA. The airplane shall be assumed to be flying in steady level flight (point A, on figure 4b-2) and the pitching control suddenly moved to obtain extreme positive pitching (nose up) except as limited by pilot effort in accordance with § 4b.220(a). (2) Checked maneuver at speeds between VA and VD. The airplane shall be assumed to be subjected to a checked maneuver from steady level flight (points A1 to D1 on figure 4b-2) and from the positive load factor (points A, to D2 on figure 4b-2) as follows: (i) A positive pitching acceleration (nose up), equal to at least the following value, shall be assumed to be attained concurrently with the airplane load factor of unity (points A1 to D1 on figure 4b-2) unless it is shown that lesser values could not be exceeded: where n is the positive load factor (see § 4b.211(a) (1)), at the speed under consideration, and V is the airplane equivalent speed, knots. (ii) A negative pitching acceleration (nose down) equal to at least the following value shall be assumed to be attained concurrently with the airplane positive maneuvering load factor (points A, to D2 on figure 4b-2) unless it is shown that lesser values could not be exceeded: 26 where n is the positive load factor (see § 4b.211(a) (1)), at the speed under consideration, and V is the airplane equivalent speed, knots. (3) Specified control displacement. In lieu of subparagraph (2) of this paragraph, a checked maneuver based on a rational pitching control motion vs. time profile may be established such that the design limit load factor as defined in § 4b.211(a) (1) will not be exceeded. The airplane response shall result in pitching accelerations not less than those specified in subparagraph (2) unless it is shown that lesser values cannot be exceeded. (d) Gust conditions. The gust conditions B' through J' on figure 4b-3 shall be investigated. The following provisions shall apply: (1) The air load increment due to a specified gust shall be added to the initial balancing tail load corresponding with steady level flight. (2) It shall be acceptable to include the alleviating effect of wing down-wash and of the airplane's motion in response to the gust in computing the tail gust load increment. (3) In lieu of a rational investigation of the airplane response it shall be acceptable to apply the gust factor K, (see § 4b.211 (b)) to the specified gust intensity for the horizontal tail. [15 F. R. 3543, June 8, 1950; 15 F. R. 4171, June 29, 1950, as amended by Amdt. 4b-3, 21 F.R. 991, Feb. 11, 1956; Amdt. 4b-11, 24 F.R. 7069, Sept. 1, 1959] The airplane shall be designed for rolling loads resulting from the conditions specified in paragraphs (a) and (b) of this section. Unbalanced aerodynamic moments about the center of gravity shall be reacted in a rational or a conservative manner considering the principal masses furnishing the reacting inertia forces. (a) Maneuvering. The following conditions, aileron deflection, and speeds, except as the deflections may be limited by pilot effort (see § 4b.220 (a)), shall be considered in combination with an airplane load factor of zero and of twothirds of the positive maneuvering factor used in the design of the airplane. In determining the required aileron deflections, the torsional flexibility of the wing shall be taken into account in accordance with § 4b.200 (d). with (1) Conditions corresponding steady rolling velocity shall be investigated. In addition, conditions corresponding with maximum angular acceleration shall be investigated for airplanes having engines or other weight concentrations outboard of the fuselage. For the angular acceleration conditions, it shall be acceptable to assume zero rolling velocity in the absence of a rational time history investigation of the maneuver. A (2) At speed V, a sudden deflection of the aileron to the stop shall be assumed. (3) At speed Vo the aileron deflection shall be that required to produce a rate of roll not less than that obtained in condition (2) of this paragraph. (4) At speed VD the aileron deflection shall be that required to produce a rate of roll not less than one-third of that in condition (2) of this paragraph. (b) Unsymmetrical gusts. The condition of unsymmetrical gusts shall be considered by modifying the symmetrical flight conditions B' or C' of figure 4b-3, whichever produces the greater load factor. It shall be assumed that 100 percent of the wing air load acts on one side of the airplane, and 80 percent acts on the other side. [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-6, 17 F. R. 1089, Feb. 5, 1952] § 4b.215 Yawing conditions. The airplane shall be designed for loads resulting from the conditions specified in paragraphs (a) and (b) of this section. Unbalanced aerodynamic moments about the center of gravity shall be reacted in a rational or a conservative manner considering the principal masses furnishing the reacting inertia forces. (a) Maneuvering. At all speeds from VMC to VA the following maneuvers shall be considered. In computing the tail loads it shall be acceptable to assume the yawing velocity to be zero. (1) With the airplane in unaccelerated flight at zero yaw, it shall be assumed that the rudder control is suddenly displaced to the maximum deflection as limited by the control stops or by a 300 lb. rudder pedal force, whichever is critical. (2) With the rudder deflected as specified in subparagraph (1) of this paragraph it shall be assumed that the airplane yaws to the resulting sideslip angle. (3) With the airplane yawed to the static sideslip angle corresponding with the rudder deflection specified in subparagraph (1) of this paragraph, it shall be assumed that the rudder is returned to neutral. (b) Lateral gusts. The airplane shall be assumed to encounter derived gusts normal to the plane of symmetry while in unaccelerated flight. The derived gusts and airplane speeds corresponding with conditions B' through J' on Figure 4b-3 as determined by §§ 4b.211 (b) and 4b.212 (a) (2) or § 4b.212 (b) (2) shall be investigated. The shape of the gust shall be as specified in § 4b.211 (b). In the absence of a rational investigation of the airplane's response to a gust, it shall be acceptable to compute the gust loading on the vertical tail surfaces by the following formula: Vairplane equivalent speed (knots). [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-6, 17 F. R. 1089, Feb. 5, 1952; Amdt. 4b-3, 21 F. R. 991, Feb. 11, 1956; 21 F. R. 1088, Feb. 17, 1956] § 4b.216 Supplementary flight conditions. (a) Engine torque effects. Engine mounts and their supporting structures shall be designed for engine torque effects combined with basic flight conditions as described in subparagraphs (1) through (3) of this paragraph. 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. (1) The limit torque corresponding with take-off power and propeller speed shall act simultaneously with 75 percent of the limit loads from flight condition A (see fig. 4b-2). (2) The limit torque corresponding with maximum continuous power and propeller speed shall act simultaneously with the limit loads from flight condition A (see fig. 4b-2). (3) For turbine propeller installations, in addition to the conditions specified in subparagraphs (1) and (2) of this paragraph, the limit torque corresponding with takeoff power and propeller speed multiplied by a factor of 1.6 shall be considered to act simultaneously with lg level flight loads. (4) For turbine engine installations, the limit torque load imposed by sudden engine stoppage due to malfunction or structural failure (e. g., compressor jamup) shall be considered in the design of the engine mounts and supporting structure. For turbine propeller installations the limit torque shall be obtained by multiplying the mean torque by a factor of 1.25. (b) Side load on engine mount. The limit load factor in a lateral direction for this condition shall be equal to the maximum obtained in the yawing conditions, but shall not be less than either 1.33 or one-third the limit load factor for flight condition A (see fig. 4b-2). Engine mounts and their supporting structure shall be designed for this condition which may be assumed independent of other flight conditions. When (c) Pressurized cabin loads. pressurized compartments are provided for the occupants of the airplane, the following requirements shall be met. (See § 4b.373.) (1) The airplane structure shall have sufficient strength to withstand the flight loads combined with pressure differential loads from zero up to the maximum relief valve setting. Account shall be taken of the external pressure distribution in flight. Stress concentration and fatigue effects shall be accounted for in the design of pressure cabins (see § 4b.270). (2) 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 to be permitted during landing. (3) 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 omit all other loads in this case. (4) 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. (d) Unsymmetrical tail load. The airplane shall be designed for unsymmetrical loads resulting from failure of one engine. (e) Gyroscopic loads. The structure supporting the engines shall be designed for gyroscopic loads associated with the conditions specified in §§ 4b.213 through 4b.215 with the engines operating at maximum continuous rpm. [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-2, 20 F. R. 5305, July 26, 1955; Amdt. 4b-3, 21 F.R. 991, Feb. 11, 1956; Amdt. 4b-6, 22 F.R. 5564, July 16, 1957; Amdt. 4b-11, 24 F.R. 7069, Sept. 1, 1959] § 4b.217 Speed control devices. When speed control devices (e.g., spoilers, drag flaps, etc.) are incorporated for use in en route conditions, the following conditions shall apply: (a) The airplane shall be designed for the symmetrical maneuvers and gusts prescribed in § 4b.211 and the yawing maneuvers and lateral gusts in § 4b.215 with the device extended at all speeds up to the placard device extended speed. (b) When the speed control device incorporates automatic operation or load limiting features, the airplane shall be designed for the maneuver and gust conditions prescribed in paragraph (a) of this section, at the speeds and corresponding device positions which the mechanism permits. [Amdt. 4b-11, 24 F.R. 7069, Sept. 1, 1959] CONTROL SURFACE AND SYSTEM LOADS § 4b.220 Control surface loads; general. The control surfaces shall be designed for the limit loads resulting from the flight conditions prescribed in §§ 4b.213 through 4b.215 and the ground gust conditions prescribed in § 4b.226, taking into account the provisions of paragraphs (a) through (e) of this section. (a) Effect of pilot effort. (1) In the control surface flight loading conditions the air loads 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 fig. 4b-5, except that two-thirds of the maximum values specified for the aileron and elevator shall be acceptable when control surface hinge moments are based on reliable data. In applying this criterion, proper consideration shall be given to the effects of servo mechanisms, tabs, and automatic pilot systems in assisting the pilot. (b) Effect of trim tabs. The effect of trim tabs on the main control surface design conditions need be taken into account only in cases where the surface loads are limited by pilot effort in accordance with the provisions of paragraph (a) of this section. In such cases the trim tabs shall be considered to be deflected in the direction which would assist the pilot, and the deflection shall be as follows: (1) For elevator trim tabs the deflections shall be those required to trim the airplane at any point within the positive portion of the V-n diagram (fig. 4b-2), except as limited by the stops. (2) For aileron and rudder trim tabs the deflections shall be those required to trim the airplane in the critical unsymmetrical power and loading conditions, with appropriate allowance for rigging tolerances. (c) Unsymmetrical loads. Horizontal tail surfaces and the supporting structure shall be designed for unsymmetrical loads arising from yawing and slipstream effects in combination with the prescribed flight conditions. NOTE: In the absence of more rational data, the following assumptions may be made for airplanes which are conventional in regard to location of propellers, wings, tail surfaces, and fuselage shape: 100 percent of the maximum loading from the symmetrical flight conditions acting on the surface on one side of the plane of symmetry and 80 percent of this loading on the other side. Where the design is not conventional (e. g., where the horizontal tail surfaces have appreciable dihedral or are supported by the vertical tail surfaces), the surfaces and supporting structures may be designed for combined vertical and horizontal surface loads resulting from the prescribed maneuvers. (d) Outboard fins. (1) When outboard fins are carried on the horizontal tail surface, the tail surfaces shall be designed for the maximum horizontal surface load in combination with the corresponding loads induced on the vertical surfaces by end plate effects. Such induced effects need not be combined with other vertical surface loads. (2) To provide for unsymmetrical loading when outboard fins extend above and below the horizontal surface, the critical vertical surface loading (load per unit area) as determined by the provisions of this section shall also be applied as follows: (1) 100 percent to the area of the vertical surfaces above (or below) the horizontal surface, and (ii) 80 percent to the area below (or above) the horizontal surface. (e) Loads parallel to hinge line. Control surfaces and supporting hinge brackets shall be designed for inertia loads acting parallel to the hinge line. NOTE: In lieu of a more rational analysis the inertia loads may be assumed to be equal to KW, where: K=24 for vertical surfaces, K=12 for horizontal surfaces, W weight of the movable surfaces. [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-6, 17 F. R. 1089, Feb. 5, 1952] § 4b.221 Wing flaps. (a) Wing flaps, their operating mechanism, and supporting structure shall be designed for critical loads prescribed by § 4b.212 (a) with the flaps extended to any position from fully retracted to the landing position. (b) The effects of propeller slipstream corresponding with take-off power shall be taken into account at an airplane speed of not less than 1.4 V11, where V1 is the stalling speed with flaps as follows: (For automatic flaps see § 4b.212 (d).) (1) Landing and approach settings at the design landing weight, (2) Take-off and en route settings at the design take-off weight. (c) It shall be acceptable to assume the airplane load factor to be equal to 1.0 for investigating the slipstream condition. [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-2, 20 F. R. 5305, July 26, 1955] § 4b.222 Tabs. The following shall apply to tabs and their installations: (a) Trimming tabs. Trimming tabs shall be designed to withstand loads |