An thin symmetric airfoil is immersed in a freestream of air at 20°C and 20Kpa and the Mach number is 3.5. Determine the wall shear stress at X=1.5m and the wall heat transfer. Repeat the calculation for Mach number 2.
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2. Air flows around a flat-bottomed airfoil shown in the figure. The air velocity at the edge of the boundary layer at the top is, u(y), and at the bottom it is the free stream velocity, U. Determine the drag coefficient, CD, for this flow if u2 (y) = U2 [1 + 1.3 (y/T)] in the front side of the airfoil and u2 (y) = U2 [2.3 (y/T)] in the back side, where T is the maximum thickness of the airfoil. You can assume that the wall shear stress at the top surface can be neglected, and the wall shear stress at the bottom surface of the airfoil is given by tw = two ( 2 - x/c ) , where two is a known constant.
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