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.

Lower values of maneuvering load factor may be employed only if it be proven that the airplane embodies features of design which make it impossible to ex(See also ceed such values in flight. § 03.131.)

§ 03.2112 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 fps nominal intensity at all speeds up to Ve, (b) positive and negative 15 fps gusts at Va. Gust load factors shall be assumed to vary linearly between Ve and Va.

§ 03.21120 Gust load factors. In applying the gust requirements, the gust load factors shall be computed by the following formula:

lar inertia of the complete airplane in a rational or conservative manner.

§ 03.212 Flaps extended flight conditions. 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: (a) Maneuvering to positive limit load factor of 2.0, (b) positive and negative 15 fps gusts acting normal to the flight path in level flight. The gust load factors shall be computed by the formula of § 03.21120.

Vr 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. Var the computed stalling speed with flaps fully extended at the design weight.

In determining the external loads on the airplane as a whole, the thrust, slipstream, and pitching acceleration may be assumed equal to zero. In designing the flaps and supporting structure, slipstream effects must be taken into account as specified in § 03.224.

§ 03.213 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.

§ 03.2131 Rolling conditions. The airplane shall be designed for (a) unsymmetrical wing loads appropriate to the category, and (b) the loads resulting from the aileron deflections and speeds specified in § 03.223, in combination with an airplane load factor of at least 2/3 of the positive maneuvering factor used in the design of the airplane.

NOTE: These conditions may be covered as noted below.

(a) Rolling accelerations may be obtained by modifying the symmetrical flight conditions shown on Figure 03-1 as follows:

Acrobatic category. In conditions A and F, assume 100% of the wing airload acting on one side of the plane of symmetry and 60% on the other.

Normal and utility categories. In condition A, assume 100% of the wing airload

as a limit condition, a CL value of -0.8 may be assumed and a linear downward acting pressure distribution peaking at the trailing edge and dropping off to zero at the leading edge may be used.

§ 03.2142 Engine torque effects. Engine mounts and their supporting structures shall be designed for engine torque effects combined with certain basic flight conditions as described in (a) and (b) below. Engine torque may be neglected in the other flight conditions.

(a) The limit torque corresponding to take-off power and propeller speed acting simultaneously with 75% of the limit loads from flight condition A. (See Figure 03-1.)

(b) The limit torque corresponding to maximum continuous power and propeller speed, acting simultaneously with the limit loads from flight condition A. (See Figure 03-1.)

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.

§ 03.2143 Side load on engine mount. The limit load factor in a lateral direction for this condition shall be at least equal to 3 of the limit load factor for flight condition A (see Figure 03-1) 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.

§ 03.22 Control surface loads.

§ 03.220 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 §§ 03.2113, 03.212, and 03.213. See Figures 03-3 to 03-10 for acceptable values of control surface loadings which are considered as conforming to the following detailed rational requirements.

§ 03.2201 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 03-11. In applying this criterion, proper consideration shall be given to the effects of servo mechanisms, tabs, and automatic pilot systems in assisting the pilot.

§ 03.2202 Trim tab effects. The effects of trim tabs on the control surface design conditions need be taken into account only in cases where the surface loads are limited on the basis of maximum pilot effort. In such cases the tabs shall be considered to be deflected in the direction which would assist the pilot and the deflection shall correspond to the maximum expected degree of "out of trim" at the speed for the condition under consideration.

§ 03.221 Horizontal tail surfaces. The horizontal tail surfaces shall be designed for the following conditions.

§ 03.2211 Balancing loads. A horizontal tail balancing load is defined as that necessary to maintain the airplane in equilibrium in a specified flight condition with zero pitching acceleration. The horizontal tail surfaces shall be designed for the balancing loads occurring at any point on the limit maneuvering envelope,

Figure 03-1, or in the flap conditions. The distribution of Figure 03-7 may be assumed.

(a) At

§ 03.2212 Maneuvering loads. maneuvering speed, V», assume a sudden deflection of the elevator control to the maximum upward deflection as limited by the control stops or pilot effort, whichever is critical. The average loading of Fig. 03-3 and the distribution of Figure 03-8 may be used.

NOTE: In determining the resultant normal force coefficient for the tail under these conditions, it will be permissible to assume that the angle of attack of the stabilizer with respect to the resultant direction of air flow is equal to that which occurs when the airplane is in steady unaccelerated flight at a flight speed equal to Vp. The maximum elevator deflection can then be determined from the above criteria and the tail normal force coefficient can be obtained from the data given in NACA Restricted Report No. 688, "Aerodynamic Characteristics of Horizontal Tail Surfaces".

(b) Same as case (a) except that the elevator deflection is downward. The average loading of Figure 03-3 and the distribution of Figure 03-8 may be used.

(c) At all speeds above Vp, the horizontal tail shall be designed for the maneuvering loads resulting from a sudden upward deflection of the elevator, followed by a downward deflection of the elevator such that the following limit conditions of normal acceleration and angular acceleration are obtained:

and the distributions of Fig. 03-8 (for downloads) and Fig. 03-9 (for uploads) may be used.

§ 03.2213 Gust loads. The horizontal tail surfaces shall be designed for loads occurring in the following conditions: (a) Positive and negative gusts of 30 fps nominal intensity at speed, V., corresponding to flight condition § 03.2112 (a) with flaps retracted. The average loadings of Figure 03-5 and the distribution of Figure 03-9 may be used for the total tail loading in this condition. (b) Positive and negative gusts of 15 fps nominal intensity at speed, Vr, corresponding to flight condition § 03.212 (b) with flaps extended. In determining the total load on the horizontal tail for these conditions, the initial balancing tail loads shall first be determined for steady unaccelerated flight at the pertinent design speed, Ve or Vt. The incremental tail load resulting from the gust shall then be added

ACCEPTABLE VALUES OF LIMIT AVERAGE MANEUVERING CONTROL SURFACE LOADINGS CAN BE OBTAINED FROM FIGURE 03-3 (B) AS FOLLOWS:

HORIZONTAL TAIL SURFACES

(1) Condition § 03.2212 (a): Obtain W as function of W/S and surface deflection; Use Curve C for deflection 10° or less Use Curve B for deflection 20°

Use Curve A for deflection 30° or more. (Interpolate for other deflections)

Use distribution of Figure 03-8.

(2) Condition § 03.2212 (b): Obtain Ŵ from curve B. Use distribution of Figure 03-8.

VERTICAL TAIL SURFACES

(3) Condition § 03.2221 (a): Obtain W as function of W/S and surface deflection in same manner as outlined in (1) above, use distribution of Figure 03-8.

(4) Condition 03.2221 (b): Obtain W from Curve C, use distribution of Figure 03-7.

(5) Condition § 03.2221 (c): Obtain W from Curve A, use distribution of Figure 03-9. (Note that condition § 03.2222 will generally be more critical than this condition.)

AILERONS

(6) In lieu of conditions (1), (2), and (3) of § 03.223, obtain W from Curve B, acting in both up and down directions. Use distribution of Figure 03-10.

FIGURE 03-3 (a).-Limit average maneuvering control surface loadings, PSF.