Question: Examine the flow over the supersonic airfoil illustrated in Fig., using shock-expansion theory. Assume that the Mach number =2.5, the chord length c = 1 m, the camber ratio h/c = 0.05, and the thickness ratio t/c = 0.02
(a) What are the lift and drag coefficients on this airfoil at angles of attack α, from -5 to 5 degrees? Display your results graphically.
(b) Compare your results with those predicted by applying the linearized theory.
(c) What is the lift coefficient at a zero angle of attack for this airfoil?