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13      0.60    0.30   5.10
14      0.45   0.20   5.60
~_-r  ----7---r---__
3.2     Estimate the lift-curve slope of a wing having a NACA 64009 airfoil section,
Ieading-edge sweep of 4~ deg, root chord 3.5 m, tip chord 2.0 m, span  15 m, and
Reynolds number 6 x l06 at Mach numbers 0.5 and 2.0.
3.6    Estimate the dynamic pressure ratio at the horizontal tail ofExercise 3.5 for
M - 0.3.
3.7   Estimate the low-speed elevator effectiveness r and hinge-moment coeffi-
cients Cha and Ch8 using the following data.
3.4     For a generic wing-body configuration shown in Fig. P3.4, determine CLa,IW
and Cmcr.WB at M  -- 0.3 and 2.0. Assume ao - 0.085ldeg (for low subsonic
speeds) and Ay = 2.2. (Hint: you may ignore the hemispherical nose cap in lift
and pitching moment calculations.)
3.5    Estimate the subsonic downwash gradient with respect to angle of attack at
the aerodynamic center of a horizontal tail located 2.5 root chords downstream of
the wing and 5cYo span above the wing root chordline. For the wing, use the data
of Exercise 3.2.
314           PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL
H
N
l-,;gJ
Fig.P3.4   Generic wing-body configuration.
n
      Horizontal tail: root chord 1.2 m, tip chord 0.8 m, span 5 m,leading-edge sweep
30 deg, and NACA 65A009 airfoil section.
    Elevator: ratio of chord length ahead of the hingeline to that aft of hingeline
0.085, ratio of chord length aft of the hingeline to the horizontal tail chord 0.165,
hingeline sweep 15 deg, and tc]2c f -. 0.06.
      Assume that the elevator extends from 20 to 80% semispan ofthe horizontal tail.
3.8   Show that
                       &~. )-=  Cm~(8e,R -cSe./)
                 CL
3.9     The aerodynamic center (a.c.) of wing-fuselage is located at 0.2 mac and the
center ofgravity is at 0.25 mac. Using the following data, determine the horizontal
tail area to give a minimum static margin of 0.08 mac. Cm.f  = O.ICL,aw =
O.l/degr: 5.5=m9.08/deg, e  - 0.45a, lr/c -. 2.5, r7r  - 0.95, and S  = 25 fl12.
[Answer:  c12,]
3.10  A flying wing employs elevons for pitch/roll control.     imetrically de-
fiected. elevons give pitch control and. when asymme .,~I" d, /m2~,ed. give ro,
control. Assumin  Cm.ac = -0.02 - 0.018e, W/S - 289C     12, and a static
margin of 79to, determine the symmetric elevon defiection for trim in level flight
at 154 m/s. [Answer: -3.3860;eg.]
3.11  An airplane with an all-movable horizontal tail has the following data:
 W/S = 1500 N/m2, S = 30 m2, CL.muL = 1.5, c = 3 m, xac = 0.25, Cmw : - 0.05,
STATIC STABILITY AND CONTROL
315
aoL.w - -2.5 deg, /u,  - -2.Odeg, aw  = O.I/deg, Cm.f  - 0-05 + 0.12CL, ar -
0.08/deg, e -.0.4 rx,tt -2.5 c, and r7t =0.8. Assuming that the most forward
and aft permissible center of gravity (c.g.) locations are 0.20 c and 0.35 c, de-
termine (a) the tail area and (b) tail setting angle. [Answer: (a) 6.87 IT12 and
(b) -5.6475 deg.J
3.12   A canard aircraft has the following data: aw - O.lO/deg, ac = 0.08/deg,
(:          -. -0.02, Xa  = -0.5 (c.g. ahead of wing a.c.), Cm. f  - O.lC/, Cc  - 0.3 c,
(.mac,zn -
C      -. -0.015, and Sc - 0.3 S. Determine the distance between the canard
( mac.c -
aerodynamic center and the center of gravity in terms of c for a static margin of
O.lO.Ignore upwash/downwash effects for canard and wing. [Answer: 1.6667.]
3.13    An aircraft with a swept-back wing has the following data: wing loading =
2000 N/m2, wing area= 30 m2, xa  = -0.15 (c,g..ahead of wing a.c.), Cm. f  =
0.05 +0.08 CL, Cmac.w  - -0.05, Cmac., - 0,  7r - 0.9, S,/S - 0.3, and /r/c = 2.5.
(a) Determine the tail-lift coefficient for steady level flight at a speed of 250 m/s
and at sea level. (b) If CD = 0.025 + 0.05CZ:ywhat is the approximate trim drag
penalty? [Answer: -0.0903, 7350 N.]
3.14    A light aircraft has the following data: CL  = 0.085(a + 1.5), iw = -1 deg,
/r = - 2 deg, Cmac.w = - 0.012, Cm. f  = 0.05 +O.lOCL, ar  = 0.075ldeg, r7r  -  0.9,
　
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