Physics, 17.07.2019 17:20 QueenNerdy889
Consider a double wedge airfoil with a maximum thickness at 0.5c operating at mach numbers of 2, 4 and 6.
a. calculate the values of lift (cl) and drag coefficients for the 5% and 10% thick airfoils at angle of attack = 1 degrees. use these values to calculate the lift curve slope (dcl/dalpha) and lift-to-drag ratio (l/d).
b. discuss the effects of mach number and maximum thickness on lift curve slope l/d.
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Consider a double wedge airfoil with a maximum thickness at 0.5c operating at mach numbers of 2, 4 a...
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