HW26solution - Nov 22 2011 AAE 334 Fall 2011 Homework 26...

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Nov. 22, 2011 AAE 334 Fall 2011, Homework 26 Due Friday, November 18 at the beginning of class. Use the shock-expansion 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 m-file 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. 3 1 4 2 OS OS PM y x
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Nov. 22, 2011 Flow from 1 to 2 is no change for 0 and is a compression corner of turning angle for 0
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