FIG. 03-5(a) GUST LOADINGS ON HORIZONTAL TAIL SURFACE 20 FIG.03-5(b) -GUST LOADINGS ON HORIZONTAL TAIL SURFACE W/S, •GROSS WEIGHT/AREA OF VERTICAL TAIL SURFACE FIGS. 03-7 TO 03-10-CONTROL SURFACE LOAD DISTRIBUTIONS U V a a pounds nominal gust intensity in fps airplane velocity in mph = tail surface area in sq. ft. slope of lift curve of tail surface, CL per degree slope of lift curve of wing, CL per degree Raspect ratio of the wing § 03.2214 Unsymmetrical loads. The maximum horizontal tail surface loading (that is, load per unit area), as determined by the preceding subsections, shall be applied to the horizontal surfaces on one side of the plane of symmetry and the following percentages of that loading shall be applied on the opposite side: 100-10 (n-1) %, but not greater than 80%, where n is the specified positive maneuvering load factor. § 03.222 Vertical tail surfaces. § 03.2221 Maneuvering loads. At all speeds up to Vp: (a) With the airplane in unaccelerated flight at zero yaw, assume a sudden displacement of the rudder control to the maximum deflection as limited by the control stops or pilot effort, whichever is critical. The average loading of Fig. 03-3 and the distribution of Fig. 03-8 may be used. (b) Assume that the airplane yaws to a sideslip angle of 15°, while the rudder control is maintained at full deflection (except as limited by pilot effort) in the direction tending to increase the sideslip. The average loading of Fig. 03-3 and the distribution of Fig. 03-7 may be used. (c) Assume that the airplane yaws to a sideslip angle of 15° while the rudder control is maintained in the neutral position (except as limited by pilot effort). The average loading of Fig. 03-3 and the distribution of Fig. 03-9 may be used. The assumed sideslip angles may be reduced if it is shown that the value chosen for a particular speed cannot be exceeded in the cases of steady slips, uncoordinated rolls from a steep bank, and sudden failure of the critical engine with delayed corrective action. less than 1.0. A value of K obtained by rational U nominal gust intensity in fps S, a-slope of lift curve of vertical surface in radians corrected for aspect ratio vertical surface area in sq. ft. W design weight in lbs. This loading applies only to that portion of the vertical surfaces having a well-defined leading edge. The average loading of Figure 03-6 and the distribution of Figure 03-9 may be used. § 03.2223 Outboard fins. 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. When outboard fins extend above and below the horizontal surface, the maximum vertical surface loading (load per unit area) as determined by § 03.222 shall be applied to the portion of the vertical surfaces above (or below) the horizontal surface, and 80% below (or above) the horizontal surface. § 03.223 Ailerons. In the symmetrical flight conditions (see § 03.211), the ailerons shall be designed for all loads to which they are subjected while in the neutral position. In unsymmetrical flight conditions (see 03.2131), the ailerons shall be designed for the loads resulting from the following deflections except as limited by pilot effort: (1) At speed, Vp, assume a sudden displacement of the aileron control to the maximum deflection. (2) When Vc is greater than Vp, the aileron deflection at Ve shall be that required to produce a rate of roll not less than that obtained in condition (1). (3) At speed, Va, the aileron deflection shall be that required to produce a rate of roll not less than 3 of that which would be obtained at the speed and aileron deflection specified in condition (1). NOTE: For conventional ailerons, the deflections for conditions (2) and (3) may be computed from: = total aileron deflection (sum of both aileron deflections in condition (1)). = total aileron deflection in condition (2). 83 total deflection in condition (3). In the equation for 83, the 0.5 factor is used instead of 0.33 to allow for wing torsional flexibility. The critical loading on the ailerons should occur in conditions (2) if Va is less than 2 Ve and the wing meets the torsional stiffness criteria. The normal force coefficient, Ca, for the ailerons may be taken as 0.04 8 where & is the deflection of the individual aileron in degrees. The critical condition for wing torsional loads will depend upon the basic airfoil moment coefficient as well as the speed and may be determined as follows: T3/T is the ratio of wing torsion in condition (3) to that in condition (2) 82d and 83d are the down deflections of the individual aileron in conditions (2) and (3) respectively. When T/T, is greater than 1.0, condition (3) is critical; when T/T, is less than 1.0, condition (2) is critical. In lieu of the above rational conditions, the average loading of Figure 03-3 and the distribution of Figure 03-10 may be used. § 03.224 Wing flaps. Wing flaps, their operating mechanism, and supporting structure shall be designed for critical loads occurring in the Flap Extended Flight Conditions (see § 03.212) with the flaps extended to any position from fully retracted to fully extended. The effects of propeller slipstream corresponding to take-off power shall be taken into account at an airplane speed of not less than 1.4 Vs where Vs is the computed stalling speed with flaps fully retracted at the design weight. § 03.225 Tabs. Control surface tabs shall be designed for the most severe combination of airspeed and tab deflection likely to be obtained within the limit V-n diagram, Figure 03-1, for any usable loading condition of the airplane. § 03.226 Special devices. The loading for special devices employing aerodynamic surfaces, such as slots and spoilers, shall be based on test data. § 03.23 Control system loads. § 03.230 Primary flight controls and systems. Flight control systems and supporting structures shall be designed for loads corresponding to 125% of the computed hinge moments of the movable control surface in the conditions prescribed in § 03.22, subject to the following maxima and minima: (a) The system limit loads need not exceed those which can be produced by the pilot or pilots and automatic devices operating the controls. (b) The loads shall in any case be sufficient to provide a rugged system for service use, including consideration of jamming, ground gusts, taxiing tail to wind, control inertia, and friction. Acceptable maximum and minimum pilot loads for elevator, aileron, and rudder controls are shown in Figure 03-11. These pilot loads shall be assumed to act at the appropriate control grips or pads in a manner simulating flight conditions and to be reacted at the attachments to the control system to the control surface horn. § 03.2300 Dual controls. When dual controls are provided, the systems shall be designed for the pilots operating in opposition, using individual pilot loads equal to 75% of those obtained in accordance with § 03.230, except that the individual pilot loads shall not be less than the minimum loads specified in figure 03-11. § 03.231 Ground gust conditions. The following ground gust conditions are not mandatory except that they shall be investigated in cases where a deviation from the minimum control forces of figure 03-11 is desired. The following conditions are intended to simulate the loadings on control surfaces due to ground gusts and taxiing tail to wind. For design weights, W, greater than 5,000 lbs., the above specified maximum values shall be increased linearly with weight to 1.5 times the specified values, at a design weight of 25,000 lbs. 2 In cases where the ground gust conditions of § 03.231 would give rise to lower control forces than the minimums specified in this figure, such lower values may be used except that they shall not be less than 0.6 of the above stated minimums. The critical portions of the aileron control system shall also be designed for a single tangential force having a limit value equal to 1.25 times the couple force determined from the above criteria. D= wheel diameter. As used above in connection with ailerons and elevators, a positive value of K indicates a moment tending to depress the surface while a negative value of K indicates a moment tending to raise the surface. § 03.232 Secondary controls and systems. Secondary controls, such as wheel brakes, spoilers, and tab controls, shall be designed for the loads based on the maximum which a pilot is likely to apply to the control in question. §03.24 Ground loads. The loads specified in the following conditions shall be considered as the external loads and inertia forces that would occur in an airplane structure acting as a rigid body. In each of the ground loads conditions speci fied the external reactions shall be placed in equilibrium with the linear and angular inertia forces in a rational or conservative manner. § 03.240 Design weight. The design weight used in the landing conditions shall not be less than the maximum weight for which certification is desired: Provided, however, That for multi-engine airplanes meeting the one-engine inoperative climb requirement of § 03.123 (b), the airplane may be designed for a design landing weight which is less than the maximum design weight, if compliance is shown with the following sections of Part 04 in lieu of the corresponding requirements of this part: the ground load requirements of § 04.24, the shock absorption requirements of § 04.361 and its subsections, the wheel and tire requirements of §§ 04.363 and 04.364, and the fuel jettisoning system requirements of § 04.428. § 03.241 Load factor for landing conditions. In the following landing conditions the limit vertical inertia load factor at the center of gravity of the airplane shall be chosen by the designer but shall not be less than the value which would be obtained when landing the airplane with a descent velocity, in fps, equal to the following value: V (fps)=4.4 (W/S) except that the descent velocity need not exceed 10 fps and shall not be less than 7 fps. Wing lift not exceeding 3 of the weight of the airplane may be assumed to exist throughout the landing impact and may, if desired, be assumed to act through the airplane c. g. When such wing lift is assumed, the ground reaction load factor may be taken equal to the inertia load factor minus the ratio of the assumed wing lift to the airplane weight. (See § 03.3612 for requirements concerning the energy absorption tests which determine the limit load factor corresponding to the required limit descent velocities.) In no case, however, shall the inertia load factor used for design purposes be less than 2.67, nor shall the limit ground reaction load factor be less than 2.0, unless it is demonstrated that lower values of limit load factor will not be exceeded in taxiing the airplane over terrain having the maximum degree of roughness to be expected under intended |