2 days ago · A thin symmetric aerofoil is shown in Figure below The angle of attack is a = 5° and the free stream Mach number is M=2.0. (a) Assuming zero thickness and using the results of the linearised supersonic flow theory, calculate the variation of the pressure coefficient and sketch it. (b) Calculate the lift coefficient. (c) If the thickness of the aerofoil is finite but still small, how would you calculate the pressure distribution? Thus C Lα is a function of wing aspect ratio, mid-chord sweep angle Λ c/2, Mach number, and airfoil section (defined parallel to the free stream) lift curve slope.The factor κ in Equation (5.18) is the ratio of the experimental two-dimensional (i.e., airfoil) lift curve slope (per radian) at the appropriate Mach number (c la) M to the theoretical value at that Mach number, 2π/β, or κ ...