Airfoil With Viscous Effects Use MFoil (or XFoil) to compute the aerodynamic characteristics of a NACA 23009. Your output should include: Polar plots for Re = 106 and Re = 107, with a ranging from -5 to 14 Plots of Cp versus x for the two Reynolds numbers at a = = 0,7, 14 Now comment on your results, making any extra plots you need to support your answers: How do the upper- and lower-surface transition point locations depend on angle of attack? How do they depend on Reynolds number? At what angle of attack does separation occur for each Reynolds number? What is an appropriate curve fit for the drag polar at each Reynolds number (only try to fit to the unstalled part of the polar)? Finally, pick a different airfoil of your own choosing, and look at the effects of a and Re on the flow. Include plots and comments.