曝光臺 注意防騙 網曝天貓店富美金盛家居專營店坑蒙拐騙欺詐消費者

in Fig. 3.90. Addition of a suitably sized dorsal fin helps to delay the vertical tail
stall to higher sideslip (Fig. 3.91a) and minimizes the possibility of rudder lock.
Also, the dorsal fin makes the pedal forces vary monotonically with sideslip as
indicated in Fig. 3.91b.
                                      Example 3.8
      For the tailless'aircraft of Example 3.2, determine the static directional stability
parameter Cnp at M -0.7 and altitude of 8500 m (c   = 0.3881) and at M = 2.0
and 18,000 m (cr = 0.094).
    Solution.   Let us calculate Cnp for M = 0.7 (subsonic) at 8500 m altitude.
    The aspect ratio of the given wing is 2.6893. Hence, the strip theory cannot be
used. We use the empirical relation in Eq. (3.271) to determine approximately the
wing contribution due to dihedral as follows:
                              (C,,p)r.w = -0.075 FCL/rad
where F is in radians. We have F = 3.5 deg so that
                       (C,p)r.w = -0.075 (53753) C~/rad
                              - -0.0046CLlrad
                              - -O.OOOICL/deg
      ' ' :':
:r,
;:-.:"i.
{
288           PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL
With Dorsal
a)
  Fp    :oucdkd "            
┏━━┳━━━━━━━━━━┓
┃    ┃                    ┃
┣━━╋━━━━━━━━━━┫
┃ O  ┃~--..-   -------- - ┃
┃    ┣━━━━━━━━━━┫
┃    ┃              -     ┃
┗━━┻━━━━━━━━━━┛
in VertiCal
I
,stall
ain Verticat
il
ith Dorsal
Fig.3.91    Effect ofdorsal f/n.
The wing contribution caused by sweep is given by the Eq. (3.299):
      (C,,ph.w  1     tanAc/4
                 CZ - = 47rA   ~A~A+4cosA/4~
        A A2 +6-a~/)
                     x  cosAc/4 - 2 -  8cosA~4+
We have Ac/4 = 36.3704 deg and A = 2.6893. Here, xa is the horizontal distance
between the wing aerodynamic center ofthe exposed wing and the center of gravity
in ter.ms of mean aerodynar/uic chord and xa > 0 if the center of gravity is aft of
the aerodynamic center. From the calculations done for Example 3.2 at M - 0.7,
we have
                       3367 M3 _ 0.4810 M2+0.2214 M +0.4918       (0 < M < 0.80)
   -,,,) w = 0.33,
For M = 0.7, we get
                         -                             (-,,,) w = 0.5266
The center of gravity is located 15,9334 m from the leading edge of the fuselage.
From Fig. 3.58, we note that the leading edge of the exposed wing is located
STATIC STABILITY AND CONTROL
289
   14.1275 m from the leading edge of the fuselage. We have Cre - 8.94 m and c -
  6.8072 m. Then,
                                  xcg - 15.9334 -14.1275
                              Crt            8 94
　
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