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O.Oa      0.11      0.16      O.ZO
 Iluckncss Ratio
  b)
Fig. 3.15  Sonic leading-edge correction factor for supersonic Iift-curve slope of
wings.l
STATIC STABILITY AND CONTROL
183
    2) Apply the correction for sonic leading-edge effects using the data given in
Fig. 3.15a and the following relation:
CN_(CNcr)rheory
CNtr = ~CNa tt.ory
(3.21)
Datain Fig. 3.15a are ploaed as functions ofthe parameter Ayi, which is defined as
AYi
Ay
cos ALE
(3.22)
where Ay is the difference between the upper-surface coordinates expressed in
percent chord at the 6u/0 and 0.15o/o chord stations as given in Fig. 3.15b for vari-
ous types of airfoil sections. For double wedge and biconvex airfoils, AYi can
also be obtained using the following relation:
Ayi = 5.85 tan 81
(3.23)
where 81 is the leading-edge, semiwedge angle. The 81 values superposed in
Fig. 3.15a apply only for double wedge and biconvex airfoils and are relatecl by
Eq. (3.23)- For other wing sections, ignore the values of 81 in Fig. 3.15a. Use th. e
curves on either side depending on whether p/ tan ALE is greater or smaller than
unity.
   For estimation of the lift-curve slope of more complex wing planforms such
as delta wings with leading-edge extensions or strakes, the reader may refer to
Datcom.l
3.3.3  Wing-Fuselage Contribution
   For configurations with relatively large values of wing-span-to-body-diameter
ratios, the mutual interference effects between the wing and the body are small
and can be ignored. For such cases, the contributions of the wing and fuselage
can be individually determined as above and added together to give the wing-
body contribution. However, for configurations with small wing-span-to-body-
diameter ratio, the mutual interference effects between wing and fuselage are
quite significant. Therefore, for such configurations, it is preferable to evaluate
the combined wing-body contribution using Datcom methods] discussed in the
following sections.
     Estimation of /ift curve s/ope.    The lift-curve slope of Lhe combined wing-
body is given by
           Sexp
CLa.WB - [KN + KWcB) + KBcW)]C,a..e  S
(3.24)
where KN, K WcB), and KBcW) represent the ratios of nose lift, the wing lift in pres-
ence of the body, and body lift in presence of the wing to wing-alone lift. We have
KN = ( CC,.., ) SS
(3.25)
Here, CLa,N iS the lift-curve slope of the isolated nose, CLct.e iS the lift-curve
slope of the exposed wing, Sexp iS the exposed wing area, and S is the reference
  CNa .8
(CNa)theory
       .7
       3
Ay
. CHOID
184           PERFORMANCE, STABILtTY, DYNAMICS, AND CONTROL
}- L, -.{
Theoretical Wing Area
Fig 3.16   Schematic def/cnition ofbody and wing nomenclature.l
(wing) area. The exposed wing is that part of the wing that lies outboard the fuse-
lage on either side. Usually, the reference wing area is equal to the theoretical
wing area, which is obtained by extending the leading and trailing edges of both
the right and left wings to meet on the centerline. Thus, the theoretical wing area
is the sum of exposed wing area and the area of that part of the fuselage obtained
by an extension of the leading and trailing edges of the wing up to the fuselage
centerline as shown in Fig. 3,16.
    For subsonic speeds,
              2(k2 - kI)SB.max
CLor.N -- ~-                            (3.26)
             S
where k2 - ki is the apparent mass constant (Fig. 3.6) and SB.m" iS the maximum
cross-sectional area of the fuselage. The value of CLa,N given by Eq. (3.26) is per
radian. For supersonic speeds, as said before, such simple and generalized methods
are not available. For cone-cylinders and ogive -cylinders, the Datconil  data given
　
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