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   rHoak, D. E., et al., *^The USAF Stability and Control Datcom:' Air Force Wright
Aeronautical Laboratories, TR-83-3048, OcL 1960 (Revised 1978).
     2Munk, M. M., "Aerodynamic Forces of Airship Hulls:' NASA TR 184, 1924.
      3Multhopp, H., isAerodynamics of Fuselage:' NASA TM-1036,  1942.
     4Abbott, I. H., and Von Doenhoff, A. E., 7heory of dWng Sections, Dover, New York,
1958.
     s Civil Airworthiness Spec~cations, Parts 23 and 25, Federal Aviarron ReguLations, U.S.
Govemment Printing Office, Washington, DC, 1991.
     6British C/wL Airworthiness Requirements, Section D, Air Registration Board, England.
    7Roskam, J.,Airplane Flight Dynamics and Automatic Flight Control, Part I, Roskam
Aviation and Engineering, Lawrence, KS, 1979.
      gSova, G., and Divan, P., "Aerodynamic Preliminary Analysis System II:' Part 2, User's
Manual, NASA CR-182077, 1990.
    9Seckel, E., Stability and Control of Airplanes and Helicopters, Academic, New York,
1964.
    loPamadi, B. N., and Pai, T. G., "A Note on the Yawing Moment Due to Sideslip for
Swept-Back Wings:' Journal of Aircra[r, Vol.  17, No. 5,  1980, pp. 378-380.
   Problems
3.1   For an airplane configuration of the type shown in Fig. 3.57, deternune the
low-speed slope.of pitching-moment-coefficient curve using Multhopp's method
STATIC STABILITY AND CONTROL
313
and the following data: Cre - 3.6 m, cr - 2.0 m, c - 3.0 m, S - 43.5 m2,b = 15 m,
A = 5.17, Ac/4 = 3.5 deg,lh = 5.0 m, hH = 0.05bh, and CLcr.WB - 0.06ldeg. The
geometrical data are given in Table P3.1.
TableP3.1 Geometricalparameters
        of the :urplane in Exercise 3.1
_- _~--_---
Secoon  Ax,m  bj,m  xi,m
 1      0.70    0.30   4.45
 2      0.60   0.50   3.90
 3      0.60    0.65   3.30
 4      0.60    0.70   2.70
 5      1.20   0.80   1.80
6      1.20   0.85   1.20
 7      1.20   0.85   0,60
 8      1.20   0.80   1.80
9      0.60   0.72   2.70
10      0.60   0.62   3.30
11      0.60   0.50   3.90
12      0.60   0.37   4.50
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.
　
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