A triangular airfoil is placed in supersonic flow with angle of attack a = 5. 0 = a as shown in the figure. A leading edge oblique shock wave is generated in the bottom of the airfoil and the shock wave angle is 28. 5. (a) (2 points)Determine the income flow Mach number (M). (b) (5 points) Calculate c, c using the shock-expansion theory. (c) (3 points) Calculate C, Ca using linearized theory (neglect the camber and thickness). Moo = ? B= 28.5 0 = 0 = 5