The inviscid, incompressible pressure distributions for a NACA 4412 airfoil for two different angles of attack are shown on the below figure. Determine the lift coefficient for = 0 and = 4 angle of attack by determining the area between the upper and lower surface curves for pressure distribution as a function of percent of chord. You may use the planimeter or numerical integration to integrate for the area. If the airfoil is flying at 4 angle of attack and at Mach = 0.7, determine the pressure coefficient on the upper surface at an x/c of 0.5 using the Prandt- Glauret rule. Highly Cambered Airfoil Pressure Distribution NACA 4412- - -2.00 -1.50 -1.00 -0.50 0.00 0.50 1.00 0.1 0.1 Inviscid calculation from PANEL NACA 4412, = 0 NACA 4412, = 4 Note: For a comparison of cambered and uncambered presuure distributions at the same lift, see Fig. 18. 0.3 0.5 x/c 0.7 0.9 1.1 ZOOM +