200           PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL
Horizonta
 Tail
Fig. 3.28    Jet-induced flow field at horizontal taiL
higher local lift and drag forces. The effect of this change in local lift and drag
forces on pitching moment is usually small and can be ignored.
      For ajet aircraft, the direct effects caused by thrust and intake normal force are
similar to those of propeller aircraft. The indirect effects due to the jet-induced
flow field may affect the horizontal tail as schematically shown in Fig. 3.28.
    In the following analysis, we ignore the contribution of power effects to static
longitudinal stability. The interested reader may refer to Datcoml for additional
information.
     Having evaluated the contribution of each of the components, we are now in a
position to analyze the static longitudinal stability of the airplane. In the following
analysis, we have assumed that the contribution of the fuselage, wing, and hori-
zontal tail can be separately represented. This approach, as mentioned earlier, is
justified for general aviation- type aircraft configurations with high values of wing
span to body diameter. For configurations with lower values of these parameters,
the contributions of wing and fuselage have to be determined as wing-body contri-
bution as discussed in Section 3.3.3. For such cases, the terms representing the
fuselage and wing contributions should be replaced by the combined wing-body
contribution. This exercise, however, is left to the reader.
       Consider the forces and moments acting on a typicallow-speed general aviation-
type airplane as shown in Fig. 3.29. Such an airplane usually has the wing aerody-
namic center ahead of the center of gravity.ln this chapter, we will refer to such a
configuration as a conventional air)lane.
     We assume that the thrust vector passes through the center of gravity. Further-
more, we ignore power effects. Summing up the pitching-moment contributions
caused by the wing, fuselage, and horizontal tail, we have
Cm g = C,,w + Cmf + Cmt
(3.64)
- CL.wXa + Cmac.w + Cmj - CL.r17t Vl                   (3.65)
STATIC STABILITY AND CONTROL
201
Fig.3.29     Aerodynamic forces and moment acting on an airplanein steadyle'veltlight.
   Let us drop the suffix cg with an understanding that C,rl denotes the pitching-
moment coefficient about the center of gravity. Then,
Then,
Cm - CL.wXa + Crnac.Lu + Cm f  ~ CLt V i rlt
(3.66)
dC,,, ' -
d~ =  a + G~i  )f - -.. (i - :l  ) V,,7t                 (3.67)
   It is usual to assume CL -. CL.w because the lift coefficient of the
more or less equal to the wing-lift coefficient. The tail-lift coefficient
small compared to wing-lift coefficient.
     With xa  = Xc8 ~ Xac, we have
   dCm _      :,'--,.)
         dC =X,g-Xa,+(:g ),--..(l   -)V,,7,
aircraft is
is usually
(3.68)
                         0 = No -  a,+ (:g  )f - :. (, - -,.) " ",
or
                             No =',,, - &C/ )f + -.. (i - :  ) V,,7,
(3.69)
(3.70)
202            PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL
+
     C mo
Cm
     0
+
E
┏━━━━━━━━━━━━━━━━━━━━━━━━━━━━━━━━━┓
┃                                                         :   ,P-- ┃
┃ -------                                                          ┃
　
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