(a) At 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.

NOTE: The average loading of Figure 3-3 and the distribution of Figure 3-8 may be used unless the Administrator finds it results in unrealistic loads. 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 V, 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 Report No. 688, "Aerodynamic Characteristics of Horizontal Tail Surfaces," or other applicable NACA reports.

(b) Same as case (a) except that the elevator deflection is downward.

NOTE: The average loading of Figure 8-3 and the distribution of Figure 3-8 may be used unless the Administrator finds it results in unrealistic loads.

(c) At all speeds above Vp the horizontal tail shall be designed for the

nm=positive limit maneuvering load factor used in the design of the airplane.

Vinitial speed in miles per hour.

(d) The total tail load for the conditions specified in (c) shall be the sum of: (1) The balancing tail load corresponding with the condition at speed V and the specified value of the normal load factor n, plus (2) the maneuvering load in crement due to the specified value of the angular acceleration.

NOTE: The maneuvering load increment of Figure 3-4 and the distributions of Figure 3-8 (for down loads) and Figure 3-9 (for up loads) may be used unless the Administrator finds it results in unrealistic loads. These distributions apply to the total tail load.

[21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-3, 23 F.R. 2589, Apr. 19, 1958]

§ 3.216-2 Unchecked pull-up maneuver (FAA policies which apply to § 3.216(a)).

(a) The condition given in § 3.216(a) represents what may occur in an "unchecked" pull-up maneuver. The basic assumption is that while the airplane is flying in steady level flight at the speed V, the pilot suddenly pulls the elevator control back and holds it in the full back position.

(1) An example of the time history of a typical case of an unchecked pull-up maneuver is shown in Figure 1 (a) (see § 3.216-1). It will be noted from this figure that approximately full elevator deflection was applied in roughly 0.1 second and that the elevator was held in the full up position until after the peak c. g. acceleration was obtained. It will also be noted that the maximum down tail load was attained before the airplane had a chance to pitch appreciably since the c. g. acceleration corresponding to maximum down tail load was approximately 1.5.

(2) This condition is intended to represent the condition obtained at the instant of maximum down tail load in an unchecked pull-up as shown on the Figure 1 (a) (see § 3.216-1) at the time of approximately 0.15 seconds.

(b) For purposes of simplifying analysis procedure the download applied to the horizontal tail surface may be carried forward to the wing attachment points, assuming that the fuselage load factor is equal to zero. The moment at the wing due to the above described loads need not be balanced out as a couple at the wing attachment points. However, the linear and angular inertia forces may be taken into account if desired.

[Supp. 10, 16 F.R. 3287, Apr. 14, 1951] § 3.216-3 Unchecked push-down maneuvering load (FAA policies which apply to § 3.216(b)).

The condition given in § 3.216(b) represents an "unchecked" push-down and is identical to § 3.216(a) in principle, except that sudden application of full forward stick is assumed. To simplify the analysis the up load applied to the horizontal tail surfaces may be carried through the attachment of the horizontal tail surfaces to the fuselage, and local fuselage members. No other structure need be investigated for this condition. [Supp. 10, 16 F. R. 3287, Apr. 14, 1951]

§ 3.216 4 Checked maneuvering load condition (FAA policies which apply to § 3.216(c)).

(a) The condition given in § 3.216(c) involves a down load and up load corresponding to what may occur in a "checked maneuver."

(b) A "checked maneuver" is defined as one in which the pitching control is suddenly displaced in one direction and then suddenly moved in the opposite direction, the deflections and timing being such as to avoid exceeding the limit maneuvering load factor.

(c) A typical case of a fully checked pull-up maneuver is shown for the DC-3 airplane in Figure 1 (c) (see § 3.216-1). This figure will be briefly reviewed as it contains all of the information essential to explaining the down load and up load cases required by § 3.216 (c).

(1) It will be noted that 8 degrees of up elevator was obtained in approximately 0.2 second. This 0.2 second time is the time at which the critical down load case occurs. It will be noted that a maximum down tail load of approximately 2,500 pounds is obtained at this point; further, that the airplane load factor is only slightly over 1 g. (The requirements specify a load factor of 1.0 for simplicity.) As time increases, it will be noted that the load factor begins to build up but that, when the load factor had been built up to approximately 2.7 g, the pilot started to push forward rapidly on the elevator control. This pushing forward is called "checking" and at speeds above the maneuvering speed such "checking" is required in order to prevent the airplane from exceeding the limit maneuvering factor. It will be noted that at the end of one second, the elevator has been completely "checked" back to zero deflection and that the maximum up tail load was obtained at this point concurrent with the maximum load factor of 3.2 g. The condition occurring at this time (1.0 second) represents the critical up tail load condition of § 3.216 (c).

[Supp. 10, 16 F. R. 3287, Apr. 14, 1951]

§ 3.216-5 Principles applicable to detailed analysis of conditions given in § 3.216 (FAA policies which apply to § 3.216).

(a) The basic principles underlying detailed analysis for the conditions covered in § 3.216 (a), (b) and (c) are described below:

(1) For the down load case, a normal acceleration of 1.0 is specified, concurrent with a specified positive value of angular acceleration. The forces acting on the airplane should therefore satisfy the following conditions:

(i) The algebraic sum of the upload on the wing and down load on the tail should equal the weight of the airplane. (For analysis purposes, a reasonable approximation to this condition is satisfactory.)

(ii) The summation of wing, fuselage and tail moments about the center of gravity of the airplane should be equal to the pitching moment of inertia of the airplane multiplied by the specified angular acceleration.

(2) The analysis of the upload condition may be carried out in the same manner, except that "nm" times the weight of the airplane is used in subparagraph (1) (i) of this paragraph.

(b) In all of the conditions covered in § 3.216 (c), the thrust may be assumed zero for simplicity. There are many computation procedures by which these conditions can be satisfied. An example of a typical method is that given in Navy Specification SS-1A. In Figure 3-4 of this part, the maneuvering tail load increment has been based on average values of the ratio of airplane pitching inertia to overall length.

(c) Conditions specified by this requirement are likely to be critical only at speeds Vp and Va. Investigation has shown that at Vp the specified down load condition is adequately taken care of by § 3.216 (a) and that the specified upload condition is adequately taken care of by § 3.216 (b). For these reasons, the conditions of § 3.216 (c) need not be investigated at the speed Vp.

[Supp. 10, 16 F. R. 3287, Apr. 14, 1951]

§ 3.216-6 Maneuvering control surface loading Figure 3-3 (b) in this part (FAA policies which apply § 3.216).

to

(a) The curves on Figure 3-3 (b) in this part were derived as follows:

(1) The three curves A, B and C of Figure 3-3 (b) giving control surfaces loading vs. W/S correspond to normal force coefficients of 0.80, 0.70, and 0.55 respectively. These curves represent psf loading obtained with the above normal force coefficients acting at a design speed of Vp based on the assumption of Cmax equals 1.5.

The horizontal tail surfaces shall be designed for loads occurring in the conditions specified in paragraphs (a) and (b) of this section.

(a) Positive and negative gusts of 30 feet per second nominal intensity at speed Vc corresponding with the flight condition specified in §3.187 (a) with flaps retracted.

NOTE: The average loadings of Figures 3-5 (a) and (b) and the distribution of Figure 3-9 may be used for the total tail loading in this condition unless the Administrator finds it results in unrealistic loads.

(b) Positive and negative gusts of 15 feet per second nominal intensity at speed V corresponding with the flight condition specified in § 3.190 (b) with flaps extended and at speed Va corresponding with the flight condition specified in §3.187 (b) with flaps retracted.

(c) In determining the total load on the horizontal tail for the conditions specified in paragraphs (a) and (b) of this section, the initial balancing tail loads shall first be determined for steady unaccelerated flight at the pertinent design speeds V1, Vc, and Va. The incremental tail load resulting from the gust shall be added to the initial balancing tail load to obtain the total tail load.

NOTE: The incremental tail load due to the gust may be computed by the following formula:

Acceptable values of limit average maneuvering control surface loadings can be obtained from Figure 3-3 (b) as follows:

Horizontal Tail Surfaces

(1) Condition § 3.216 (a):

Obtain was 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 3-8.

(2) Condition § 3.216 (b):

Obtain w from Curve B. Use distribution of Figure 3-8.

a=slope of lift curve of wing, Cz per degree, and

Raspect ratio of the wing.

Vertical Tail Surfaces

(3) Condition § 3.219 (a):

Obtain was function of W/S and surface deflection in same manner as outlined in (1) above, use distribution of Figure 3-8;

(4) Condition § 3.219 (b):

Obtain w from Curve C, use distribution of Figure 3-7;

(5) Condition § 3.219 (c):

Obtain w from Curve A, use distribution of Figure 3-9. (Note that condition § 3.220 generally will be more critical than this condition.)

Ailerons

(6) In lieu of conditions § 3.222 (b): Obtain w from Curve B, acting in both up and down directions.

Use distribution of Figure 3-10.