ZOOM + 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 a = 0 and a = 40 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 40 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 Inviscid calculation from PANEL -1.50 -."**:"" NACA 4412, a = 0 NACA 4412, a = 40 -1.00 Cp -0.50 0.00 0.50 Note: For a comparison of cambered and uncambered presuure distributions at the same lift, see Fig. 18. 1.090.1 0.1 0.3 0.5 0.7 0.9 1.1 x/c