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§ 03.141 Longitudinal stability and control. There shall be no uncontrollable tendency for landplane to nose over in any operating condition reasonably expected for the type, or when rebound occurs during landing or take-off. Wheel brakes shall operate smoothly and shall exhibit no undue tendency to induce nosing over. Seaplanes shall exhibit no dangerous or uncontrollable porpoising at any speed at which the airplane is normally operated on the water.

§ 03.142 Directional stability and control. (a) There shall be no uncontrollable looping tendency in 90° crosswinds up to a velocity equal to 0.2 V so at any speed at which the aircraft may be expected to be operated upon the ground or water.

(b) All landplanes shall be demonstrated to be satisfactorily controllable with no exceptional degree of skill or alertness on the part of the pilot in power-off landings at normal landing speed and during which brakes or engine power are not used to maintain a straight path.

(c) Means shall be provided for adequate directional control during taxiing.

03.143 Shock absorption. The shock absorbing mechanism shall not produce damage to the structure when the airplane is taxied on the roughest ground which it is reasonable to expect the airplane to encounter in normal operation.

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§ 03.15 Flutter and vibration. parts of the airplane shall be demonstrated to be free from flutter and excessive vibration under all speed and power conditions appropriate to the operation of the airplane up to at least the minimum value permitted for Va in § 03.2110. There shall also be no buffeting condition in any normal flight condition severe enough to interfere with the satisfactory control of the airplane or to cause excessive fatigue to the crew or result in structural damage. However, buffeting as stall warning is considered desirable and discouragement of this type of buffeting is not intended.

§ 03.2 Strength requirements.

§ 03.20 General.

§ 03.200 Loads. Strength requirements are specified in terms of limit and ultimate loads. Limit loads are the maximum loads anticipated in service. Ultimate loads are equal to the limit loads multiplied by the factor of safety. Unless otherwise described, loads specified are limit loads.

Unless otherwise provided, the specified air, ground, and water loads shall be placed in equilibrium with inertia forces, considering all items of mass in the airplane. All such loads shall be distributed in a manner conservatively approximating or closely representing actual conditions. If deflections under load would change significantly the distribution of external or internal loads, such redistribution shall be taken into account.

§ 03.201 Factor of safety. The factor of safety shall be 1.5 unless otherwise specified.

§ 03.202 Strength and deformations. The structure shall be capable of supporting limit loads without suffering detrimental permanent deformations. At all loads up to limit loads, the deformation shall be such as not to interfere with safe operation of the airplane. The structure shall be capable of supporting ultimate loads without failure for at least 3 seconds, except that when proof of strength is demonstrated by dynamic tests simulating actual conditions of load application, the 3-second limit does not apply.

§ 03.203 Proof of structure. Proof of compliance of the structure with the strength and deformation requirements of § 03.202 shall be made for all critical loading conditions. Proof of compliance by means of structural analysis will be accepted only when the structure conforms with types for which experience has shown such methods to be reliable. In all other cases substantiating load tests are required. In all cases certain portions of the structure must be subjected to tests as specified in § 03.3.

§ 03.21 Flight loads.

§ 03.210 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 § 03.63.

§ 03.2101 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.

§ 03.211 Symmetrical flight conditions (flaps retracted). 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 03-1, which represents the envelope of the flight loading conditions specified by the maneuvering and gust criteria of §§ 03.2111 and 03.2112. This diagram will also be used in determining the airplane structural operating limitations as specified in § 03.6.

§ 03.2110 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 (A)

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except that the value of V, need not exceed the value of Vc used in design.

§ 03.2111 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 § 03.21110 at all speeds up to Va,

(b) The negative maneuvering load factor specified in § 03.21110 at speed Vc; 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.

§ 03.21110 Maneuvering load factors. The positive limit maneuvering load factors shall not be less than the following values (see fig. 03-2):

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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 exceed such values in flight. (See also § 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 Vc,

(b) Positive and negative 15 fps gusts at Va. Gust load factors shall be assumed to vary linearly between Ve and Va.

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(for W/S>16 psf)

2.67 (W/S) U-nominal gust velocity, fps. (Note that the "effective sharp-edged gust" equals KU.)

V airplane speed, mph.

m-slope of lift curve, CL per radian, corrected for aspect ratio. W/S=wing loading, psf.

§ 03.2113 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 03.2211) shall be taken into account in a rational or conservative manner.

Incremental horizontal tail loads due to maneuvering and gusts (see §§ 03.2212 and 03.2213) shall be reacted by angular 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 a 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.

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 § 03.353.)

In designing the flaps and supporting structure, slipstream effects shall be taken into account as specified in § 03.224.

NOTE: In determining the external loads on the airplane as a whole, the thrust, slipstream, and pitching acceleration may be assumed equal to zero.

§ 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 twothirds 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 in Figure 03-1 as follows:

(1) Acrobatic category. In conditions A and F, assume 100% of the wing air load acting on one side of the plane of symmetry and 60% on the other.

(2) Normal and utility categories. In condition A, assume 100% of the wing air load acting on one side of the airplane and 70% on the other. For airplanes over 1,000 lbs. design weight, the latter percentage may be increased linearly with weight up to 80% at 25,000 lbs.

(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 § 03.223.

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§ 03.2132 Yawing conditions. The airplane shall be designed for the yawing loads resulting from the vertical surface loads specified in § 03.222.

S 03.214 Supplementary conditions. § 03.2141 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 (mph)

NOTE: It may be assumed that the value of C is equal to -0.8 and the chord wise distribution is triangular between a peak at the trailing edge and zero at the leading edge.

§ 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 paragraphs (a) and (b) of this section. 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 percent 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 control system boost and 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, and in the flap conditions. (See § 03.212.)

NOTE: The distribution of Figure 03-7 may be used.

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