ble that angular inertia forces initially resist most of the unbalanced couple added by the gust, while in a more or less steady pull-up condition the balancing tail load may consist entirely of a balancing air load from the tail surfaces. Compliance Suggestion BALANCING THE GLIDER a. The following considerations are involved in balancing the glider: 1. It is assumed that the limit load factors specified for the basic flight conditions Chap. 1, p. 9 are wing load factors. A solution is therefore made for the net load factor acting on the whole glider. The value so determined can then be used in connection with each item of weight (or with each group of items) in analyzing the fuselage. For balancing purposes the net factor is assumed to act at the center of gravity of the glider. 2. Assuming that it is possible for a load to be acting in the opposite direction on the elevator, it is recommended that the center of pressure of the horizontal tail be placed at 20 percent of the mean chord of the entire tail surface. This arbitrary location may also be considered as the point of application of inertia forces resulting from angular acceleration, thus simplifying the balancing process. 3. In fig. 1-X the external forces are assumed to be acting at three points only. The assumption can generally be made that the fuselage drag acts at the center of gravity. When more accurate data are available, the resultant fuselage drag force can of course be computed and applied at the proper point. Where large independent items having considerable drag are present it is advisable to extend the arrangement of forces shown in fig. 1-X to include the additional external forces. b. As shown in fig. 1-X, a convenient reference axis is the basic chord line of the mean aerodynamic wing chord. (The basic chord line is usually specified along with the dimensions of the airfoil section.) The determination of the size and location of the mean aerodynamic chord is outlined in Chap. 1, p. 15. c. A tabular form will simplify the computations required to obtain the balancing loads for various flight conditions. A typical form for this purpose is shown in table 1-II on p. 80. In using fig. 1-X and table 1-II the following assumptions and conventions should be employed: 1. Careful attention should be paid to those footnotes of fig. 1-X which pertain to the sign convention that has been adopted. If known forces are opposite in direction from those shown in fig. 1-X, a negative sign should be prefixed before inserting in the computations. In particular, it should be noted that the vertical distance, h2, is negative when the wing aerodynamic center is above the CG, and positive when the wing aerodynamic center is below the CG. The direction of unknown forces will be indicated by the sign of the value obtained from the equations. A negative value of nz will usually be determined from the balancing process, indicating a down load on the tail. For conditions of positive acceleration the solution should give a negative value for nz, as the inertia load will be acting downward. The convention for mi corresponds to that used for moment coefficients; that is, when the value of Cm is negative mi should also be negative, indicating a diving moment. 2. All distances should be divided by the mean aerodynamic chord before being used in the computations. 3. The chord load acting at the tail surfaces may be neglected. d. In table 1-II the computation of balancing loads is indicated for typical flight conditions. The equations are based on the fact that the use of the average force coefficients in connection with the design wing area, mean aerodynamic chord, and mean aerodynamic center will give resultant forces and moments of the proper magnitude, direction and location. Provision is made in the table for obtaining the balancing loads for different design weights. The table may be expanded to include computations for several loading conditions, special flight conditions, or conditions involving the use of auxiliary devices. It should be noted that a change in the location of the CG will require a corresponding change in the values of X2 and hz on fig. 1-X. (Condition IV will usually result in the largest balancing tail load.) • When the full-load center of gravity position varies appreciably the glider should be balanced for both extreme positions unless it is apparent that only one is critical. In general, only one center of gravity need be considered for single-place gliders. In special cases, it may also be necessary to check the balancing tail loads required for the loading conditions which produce the most forward and most rearward center of gravity positions for which approval is desired. e. The following explanatory notes refer by number to items appearing in table 1-II: (4) The wing loadings should be based on the design wing area. (5) ni = limit load factor required for the condition being in vestigated. (See Chap. 1, p. 3, “Load factors" and Chap. 1, p.8i.) See also Chap. 1, p.15, in cases involving wing flaps. ments about point (2) of fig. 1-X from which the following equation is obtained: = 1 mi-nqi hz +ni X2 13 (23 – 22) NOTE.—The above explanatory notes apply only when the force arrangement shown in figure 1-X is used. If a different distribution of external loads or a different system of measuring distances is employed, the computations should be correspondingly modified. CONTROL SURFACE LOADS In addition to the flight loads specified in the beginning of this chapter, the primary structure should meet the requirements described herein to account for the loads acting on the control surfaces. The following loading conditions include the application of balancing loads derived from the symmetrical flight conditions and also cover the possibility of loading the control surfaces in operating the glider and by encountering gusts. A minimum limit factor of safety of 1.0 and a minimum ultimate factor of safety of 1.5 should be used unless otherwise specified. Also, see table for multiplying factors of safety required in certain cases. Compliance Suggestion UNIT AND CHORD LOADING OF SURFACES a. The recommendations for the design of control surfaces, as outlined above, are based on the two separate functions of control surfaces: balancing and maneuvering. The requirements also are specified to account for the effects of gust loads. 6. The average unit loading normal to any surface is determined from the design gliding speed V, as shown by figs. 1-XIII, 1-XIV, and 1-XVI. When dealing with tail surfaces, it is customary to specify the value of the average unit loading for the entire surface, including both the fixed and movable surfaces. The total load, which is equal to the average unit loading multiplied by the area of the entire surface, is then distributed so as to simulate the conditions which exist in flight. In the case of ailerons, flaps or tabs, the value of the average unit loading is usually determined only for the particular surface, without reference to the surface to which it is attached. c. The average unit loading is assumed to be constant over the span. Also, as shown in figs. 1-XI, and 1-XII, the unit loading at the hinge line is constant over the span. d. Although there are no specific chord loading conditions for control surfaces specified above, such surfaces should be designed to withstand a reasonable amount of chord load in either direction. A total chord load equal to 20 percent of the maximum normal load may be used as a separate design condition. The distribution along the span may be made proportional to the chord, if desired. Tests for this condition are not required unless the structure is such as to indicate the advisability of such tests. |