Alockheed f-104 supersonic fighter has a wing with a thin symmetric airfoil with a thickness ratio of 3.6 %. consider this airfoil in a flow at an angle of attack of 70. if the lift coefficient for incompressible flow, i. e., m < 0.3, is given by the following expression, cl = 2, where  is the angle of attack in radians, determine the lift coefficient for the wing of the f-104. if the lift coefficient for the wing in transonic flow, i. e. 0.3 < m < 0.9, is 0.9 and based on the incompressible value determined above, determine the exact mach number of the flow. if the lift coefficient for the wing in supersonic flow, i. e., 1.1 < m < 5.0, is 0.3 , determine the exact mach number of the flow.