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AAE 334 Fall 2011, Homework 26
Due Friday, November 18 at the beginning of class.
Use the shockexpansion method to compute the lift and drag coefficients for the airfoil
sketched below in a Mach 1.8 flow.
Chord length is c and maximum thickness is at the quarter
chord point and equal to 8% of the chord.
Angle of attack varies from 0 to 10 degrees.
Use
Matlab compressible flow files to save time.
Hand in:
1.
Written explanation of your method
2.
Your mfile
3.
Plot of the lift curve
4.
Plot of the drag polar
Geometry first:
Label freestream as 1, lower surface 2, forward upper surface 3, and aft upper surface 4.
Surface slopes are
0
2
dx
dy
,
32
.
0
25
.
0
08
.
0
3
dx
dy
, and
107
.
0
75
.
0
08
.
0
4
dx
dy
.
Leading edge angle is
74
.
17
25
.
0
08
.
0
arctan
LE
.
Similarly, trailing edge angle is
090
.
6
75
.
0
08
.
0
arctan
TE
.
Thus the internal angle at the apex of the upper surface is 156.2
o
.
The turning angle then from 3 to 4 is 180
o
156.2
o
= 23.8
o
.
Now find the pressures.
x
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 Fall '09
 COLLICOTT

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