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aircraft. We studied elements of the fiuid flow over wings and bodies at subsonic,
REVIEW OF BASIC AERODYNAMIC PRINCIPLES 65
transonic, and supersonic speeds and understood how lift, drag, and pitching mo-
ments vary with angle of attack and Mach number. This aerodynamic background
slriould be useful in the study of performance, stability, and control of the aircraft,
which will be discussed in the next chapters of this text
References
iSchLichting, H., Boundary Layer 7heory, 6th ed., McGraw-Hill, New York, 1968.
2Abbott, I. H., and Von Doenhoff, A. E., Theory of Wing Sections, Dover, New York,
1959.
.3Perkins, C; D., and Hage, R. E*,Airpkme Perj'ormance, Stability cmd Control, 10th ed.,
Wiley, New York, 1965.
4Jacobs,E. W., and Sherman, A.,l*Airfoil Section Characteristics as Affected by Variations
of the Reynolds Number:' Rept. No. 586, 1936.
sCooke, J. C., and Brebner, C. G., "The Nature of Separation and its Prevention by
Geometric Design in a Wholly Subsonic Flow:'Boundary Layer and Flow Control, Vol. l,
Edited by G. V. Lachmann, Pergamon, New York, 1961, pp. 145-185.
6McCormick, B; W.,Aerodynamics, Aeronautics and Flight Mechanics, Wiley, New York,
1979.
7Kucthc, A. M., and Schetzelr, J. D., Foundations of Aerodynamics, 2nd ed., Wiley, New
York, 1961.
8Multhopp, H.,IIMethods for Calculating the Lift Distribution ofWings (Subsonic Lifting
Surface Theory):' ARC R and M 2884, 1950(U).
9Fallmer, V. M., "The Calculation ofthe Acrodynamic Loading on Surface of Any Shape:'
ARC R and M 1910, 1943GD.
loDe Young, J., and Harper, C. W., "Theoretical Symmetric Spanwise Loading at Subsoruc
Speeds for Wrngs Having Arbitrary Planform:' NACA TR 921, 1948(U).
iiNeumark, S.,"'Critical Mach Numbers for Thin Untapered' Swept Wings at Zero
Incidence:' ARC R and M 2821, 1954(U).
12Wh:itcomb, R. E., "A Design Approach and Selected W. ind Tcrnnel Results at High Sub-
sonic Speed for Wing jfip Mounted Winglets," NASA TND-8260, 1976.
13Wlutcomb, R. E., "Review of NASA Supefcritical Airfoils:' The Ninth Congress of
ICAS, Haifa, Israel, 1974.
14Rao, D. 1VL, and Kariya, T. T., "Boundary-Layer Submerged Vortex Generators for
Separation Control-An Exploratory Study:' Proceedings of lst National Flui'd Dynamic
Conference, Cincinnatti, OH, 1988, pp. 839-846.
tsMaltby, R. L., "The Development of the Slender Delta Concept," Aircralt Engineer:ing,
March 1968, pp. 12-17.
16Polhamus, E. C., "Vortex Lift Research, Early Contributions and Some Current
Challenges:' NASA CP-2416, Vol. 1, 1992.
17Moore, M., and Frei, D., "X-29 Forward-Swept Wing Aerodynamic Overview:' AJ[AA
Paper 83-1834, 1983.
18Murri, D. G., Nguyen, L. T., and Grafton, S. B., "Wind Tunnel Free-Flight Investigations
of a Model of Forward-Swept Wing Fighter Configuration:' NASA TP 2230, Feb. 1984.
19Jones, R. T., "Flying-Wing SST for the Pacific:' Aerospace America, Nov. 1986, pp.
32, 33.
:zoWhitcomb, R. E., IIResearch Methods for Reducing Acrodynamic Drag at Transonic
Speeds:' The InaguralEastman Jacobs Lecture, NASA Langley Research Center, Hampton,
VA, Nov. 14, 1994.
66 PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL
2t Whitcomb, R. E., and Silver, J. R., k., "A Supersonic Area Rule and an Application to the
Design ofa Wing-Body Combination with High Lift-to-Drag Ratio:' NASA TR R-72, 1960.
Problems
1.1 A fl wing weighing 20,000 N has a NACA 653-418 airfoil section, area
of30l:2fla/my dg wiaspect ratio 5.0. Determine the angle of attack for a sealevel fiight
at 60 m/s. What is the lift-to-drag ratio? Assume ao = 0.106/deg, Cdo - 0.0043,
and e - 0.92.
1.2 For a certain airfoil section, the pitching-moment coefficient about 0.33
chord behind the leading edge varies with CL as shown in Table P2.1.
Table P2.1
-
Cl Cm
0.20 -0.02
0.40 0
0.60 0.02
0.80 0.04
Determine the locations of the aerodynamic center and the center of pressure
for Ci = 0.5.
1.3 For an aircraft wing with a leading-edge sweep of 45 deg, root chord of
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