Consider a NACA 1408 airfoil in low speed flow at zero degrees angle of attack and a Reynolds number of 8.9 ´106. Calculate the percentage of drag that is due to pressure drag due to flow separation (form drag). Assume a transition Reynolds number of 500,000. Assume that the airfoil is thin enough that the skin-friction drag may be estimated by the flat plate results discussed in the section on viscous flow. Repeat this analysis for an angle of attack of 6 degree. What does this tell you about the rapid increase in Cd as the angle of attack of the airfoil is increased? What is the total skin friction drag over the laminar, and turbulent flow?