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┃- (c  ┃  ---.    . -��                                  Thickness Ratio
                                         c) Variation of trailing-edge angle ~rE
Fig.3.13 Sectional(two-dimensional)lift-curveslopeofwings,corttinued.l
several NACA family airfoils. In general, ~rE +- ~/n. However, in the absence of
actual geometrical data, we may assume ~;z- = 4rE and approximately determine
the value of q>rE using the data given in Fig. 3.13c.
    The midchord sweep Ac/2 iS given by the following expression:
       tanAc/2=tanA,E-( b )     (3.19)
 where ALE i-s the leading-edge sweep angle, Cr iS the root chord, Cr iS the tip chord,
and b is the wing span.
   For high-aspect ratio rectangular wings at 'very low speeds (incompressible
flow), the following simple formulais quite often used:
aw = ~+                           (3.20)
It may be observed that for such cases, Eq. (3.16) reduces to Eq. (3.20).
      The estimation oflift-curve slope of straight tapered wings at supersonic speeds
can be done using the Datcom datal presented jn Figs. 3.14 and 3.15. At high
speeds, the slope of normal-force coefficient is usually given instead of the lift-
curve slope. Approximately, CNu - CUt. The procedure of estimating CNa iS as
follows:
     1) For the given value of taper ratio A, find (CNa)meory from Figs. 3.14a-3.14f.
Use intenlolation if the given value of A is not covered in Figs. 3.14a-3.14f.
STATIC STABILITY AND CONTROL
                                                                                                                                                         
                                                                                                                                                         
　
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