Control surface tabs shall be designed for the most severe combination of air speed and tab deflection likely to be obtained within the limit V-n diagram (Fig. 3-1) for any usable loading condition of the airplane.

§ 3.224-1 Trim tab design (FAA policies which apply to § 3.224).

(a) To provide ruggedness and for emergency use of tabs, it is recommended that trim tabs, their attachments and actuating mechanism be designed for loads corresponding to full tab deflection at speed V. with main surface neutral; except that the tab deflection need not exceed that which would produce a hinge moment on the main surface corresponding to maximum pilot effort.

(b) A trapezoidal chord load distribution with the loading of the leading edge twice that of the trailing edge is acceptable.

[Supp. 10, 16 F.R. 3288, Apr. 14, 1951] § 3.225

Special devices.

loads corresponding to 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in §§ 3.211 to 3.225, subject to the following maxima and minima:

(1) The system limit loads need not exceed those which can be produced by the pilot and automatic devices operating the controls.

(2) The loads shall in any case be sufficient to provide a rugged system for service use, including consideration of jamming, ground gusts, taxying tail to wind, control inertia, and friction.

(b) Acceptable maximum and minimum pilot loads for elevator, aileron, and rudder controls are shown in Figure 3-11. 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 of the control system to the control surface horn.

§ 3.231-1 Hinge moments (FAA policies which apply to § 3.231(a)).

The 125 percent factor on computed hinge moments provided in § 3.231(a) need be applied only to elevator, aileron and rudder systems. A factor as low as 1.0 may be used when hinge moments are based on test data; however, the exact reduction will depend to an extent upon the accuracy and reliability of the data. Small scale wind tunnel data are generally not reliable enough to warrant elimination of the factor. If accurate flight test data are used, the factor may be reduced to 1.0.

[Supp. 10, 16 F. R. 3288, Apr. 14, 1951]

§ 3.231-2 System limit loads (FAA poli

cies which apply to § 3.231(a)(1)). (a) When the autopilot is acting in conjunction with the human pilot, the autopilot effort need not be added to human pilot effort, but the autopilot effort should be used for design if it alone can produce greater control system loads than the human pilot.

sys

(b) When the human pilot acts in opposition to the autopilot, that portion of the system between them should be designed for the maximum effort of human pilot or autopilot, whichever is the lesser. [Supp. 10, 16 F. R. 3288, Apr. 14, 1951] § 3.231-3 Interconnected control tems on two-control airplanes (FAA policies which apply to § 3.231). (a) With respect to interconnected control systems such as in two control airplanes, the following is recommended in showing the "same level of safety" specified in § 3.10.

(1) If, in the case of two or more interconnected control systems, the control wheel or stick forces due to combined control system loads resulting from air loads on the control surfaces are less than the minimum prescribed in Figure 3-11 of this part, each control system from the interconnection to the control surface should be designed for minimum pilot effort on the control wheel or stick in order that sufficient ruggedness be incorporated into the system.

(2) If the control wheel or stick forces due to combined control system loads resulting from air loads on the control surfaces are in excess of the maximum forces prescribed in Figure 3-11 of this part, it is considered permissible to divide the maximum pilot effort loads in the control systems from the point of interconnection to the control surfaces in proportion to the control surface air loads. However, the load in each such control system should be increased 25 percent to allow for any error in the determination of the control surface loads, but in no case need the resulting loads in any one system exceed the total pilot effort, if the pilot effort were applied to that system alone. In any case, the minimum load in any one system should be no less than that specified in subparagraph (1) of this paragraph. [Supp. 10, 16 F. R. 3288, Apr. 14, 1951] § 3.232 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 percent of those obtained in accordance with § 3.231, except that the individual pilot loads shall not be less than the minimum loads specified in Figure 3-11.

§ 3.233 Ground gust conditions.

(a) The following ground gust conditions shall be investigated in cases where a deviation from the specific values for minimum control forces listed in Figure 3-11 is applicable. The following conditions are intended to simulate the loadings on control surfaces due to ground gusts and when taxiing with the wind. (b) The limit hinge moment H shall be obtained from the following formula: H=K&Sq

(c) As used in paragraph (b) 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. § 3.233-1

Ground gust loads (FAA policies which apply to § 3.233). Section 3.233 requires ground gust loads to be investigated when a reduction in minimum pilot effort loads is desired. In such cases the entire system shall be investigated for ground gust loads. However, in instances where the designer desires to investigate ground gust loads without intending to reduce pilot effort loads, the ground gust load

- need be carried only from the control surface horn to the nearest stops or gust locks, including the stops or locks and their supporting structures.

[Supp. 1, 12 F. R. 3436, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5, 19491

§ 3.234 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.

§ 3.241-1

Four-wheel

type alighting gears (FAA policies which apply to § 3.241).

At present, little operational data or other information are available on which to base requirements for airplanes equipped with four wheel type alighting gears. The following is suggested for applying the requirements of this part to aircraft equipped with four wheel type alighting gears.

(a) The provisions of §§ 3.241 through 3.256, except for the following, should be considered applicable: §§ 3.245 (a), 3.246 (a), and 3.250 through 3.252.

(b) The conditions as specified in §§ 3.245 (b) (2), 3.246 (b), 3.247 and 3.249 should be considered applicable to four wheel type gear without modification, the rear wheels being considered the main gear.

(c) The landing conditions specified in §3.245 (b) (1) should be modified by dividing the total required load on the forward gear between the two wheels, 60 percent to one wheel and 40 percent to the other.

(d) The requirements of § 3.253 should be modified by applying the required loads simultaneously to the two front wheels, 120 percent to one wheel and 80 percent to the other. (Note that this gives an 80-40 percent distribution of the total load on the front gear.)

(e) It is believed that the method of applying the requirements of this part for single nose wheel type alighting gear to four wheel type alighting gear should result in a satisfactory design. It is suggested, however, that sufficient landing and taxiing tests be conducted to determine the suitability of the landing gear design and configuration. Since higher speed turns should be possible with a four wheel aircraft than with one having a conventional tricycle gear, it is believed that provision should be made to include high speed turns in the taxiing test programs of all four wheel aircraft.

(f) If an aircraft with four wheel type alighting gear is also designed for roadability, i. e. for use as an automobile, which is usually the case, the design of the alighting gear in accordance with applicable motor vehicle design requirements is acceptable, provided it can be shown that these requirements fully cover the airworthiness requirements of the regulations in this subchapter. [Supp. 10, 16 F. R. 3288, Apr. 14, 1951] § 3.241-2 Ground load evaluation for aircraft with wing tip-tanks (FAA policies which apply to § 3.241). The assumption of the aircraft structure as a rigid body in applying the ground load conditions of §§ 3.241 through 3.243 is not considered directly applicable to aircraft which incorporate tip-tanks. When such a design feature is present, the dynamic response of the wing to the short-period landing load impulse may induce inertia loadings of the wing which are significantly higher than the rigid body inertia loading and which create critical wing loadings greater than all other wing design conditions. Accordingly, neglect of the wing inertia loadings due to dynamic response of the wing structure under the landing loads may render the aircraft unsafe. Therefore, the dynamic inertia loading of the airplane wing structure should be considered in evaluating the design loads under the ground load conditions when tip-tanks are present in accordance with the following:

(a) Only the two-wheel level landing condition, § 3.245 (a) or (b) (2), need be considered in substantiating the structural strength of the wing, wing tip-tank and wing fuselage attaching structure for dynamic loads.

(b) The spanwise inertia loading should be determined by either of the following methods:

(1) An engineering evaluation which conservatively provides for the effect of dynamic response. One acceptable method of dynamic landing load analysis is given in CAA Engineering Report No. 52, entitled "Outline of an Acceptable Method of Determining Dynamic Landing Loads".

(2) Dynamic tests of the complete aircraft consisting of static drop tests or in-flight landing tests wherein suitable test instrumentation is used to evaluate the design variation of the vertical inertia load factor from the aircraft centerline to the wing tip under the landing impact.

[Supp. 20, 19 F. R. 8653, Dec. 17, 1954]

§ 3.242 Design weight.

The design landing weight shall not be less than the maximum weight for which the airplane is to be certificated, except as provided in paragraph (a) or (b) of this section.

(a) A design landing weight equal to not less than 95 percent of the maximum weight shall be acceptable if it is demonstrated that the structural limit load values at the maximum weight are not exceeded when the airplane is operated over terrain having the degree of roughness to be expected in service at all speeds up to the take-off speed. In addition, the following shall apply:

(1) The minimum fuel capacity shall not be less than the total of the capacity prescribed in § 3.440 and of the capacity equivalent to the weight of fuel equal in amount to that by which the maximum weight exceeds the design landing weight.

(2) The operating limitations shall limit the take-off weight in such a manner as to assure that landings in normal operation would not exceed the design landing weight.

(b) A design landing weight equal to less than 95 percent of the maximum weight shall be acceptable for multiengine airplanes, meeting the one-engine-inoperative climb requirement of § 3.85 (b) or § 3.85a (b) if compliance is

1 Not filled for publication in the Office of the Federal Register.

shown with the following sections of Part 4b of this subchapter in lieu of the corresponding requirement of this part: The ground load requirements of § 4b.230, the landing gear requirements of §§ 4b.331 through 4b.336, and the fuel jettisoning system requirements of § 4b.437.

§ 3.243 Load factor for landing conditions.

In the following landing conditions the limit vertical inertia load factor at t 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 feet per second, equal to the following value:

V=4.4 (W/S)\

except that the descent velocity need not exceed 10 feet per second and shall not be less than 7 feet per second. Wing lift not exceeding two-thirds of the weight of the airplane may be assumed to exist throughout the landing impact and may be assumed to act through the airplane center of gravity. 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 3.354 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 taxying the airplane over terrain having the maximum degree of roughness to be expected under intended service use at all speeds up to take-off speed.

LANDING CASES AND ATTITUDES

§ 3.244 Landing cases and attitudes. For conventional arrangements of main and nose, or main and tail wheels, the airplane shall be assumed to contact the ground at the specified limit vertical velocity in the attitudes described in §§ 3.245-3.247. (See Figs. 3-12 (a) and 3-12 (b) for acceptable landing conditions which are considered to conform with §§ 3.245-3.247.)