The airfoil on the F-5 supersonic fighter is a thin symmetric airfoil with a thickness ratio of 4.5%. The airfoil is immersed in a Mach 1.5 flow at an angle of attack of 40. The airplane is cruising at 25,000 feet. Assume the wing can be approximated by a thin, flat plate with a 6.0-ft chord and a 25-foot span. The pressure drag coefficient on the wing is equal to 1.5 counts of drag; one count is equal to .001. Determine the total drag on the wing consisting of skin friction and pressure drag, i.e., the profile drag. Assume the flow over the wing is a combination of laminar flow followed by turbulent flow, and transition occurs at a Reynolds number of 500,000. Since the Mach number is greater than 0.3, compressibility needs to be taken into account. Also determine the location of boundary layer transition from the leading edge and the thickness of the boundary layer at the trailing edge of the wing assuming turbulent flow over the entire wing. ?= u= 3.7373(10)-7 sl/f/s. Solve this problem in the British system of units.