Consider an NACA 2412 airfoil with the following tabulation of the lift, drag, and moment coefficients about the quarter chord for this airfoil, as a function of angle of attack. From this table: Calculate the maximum lift to drag ratio and find the corresponding angle of attack. Calculate the location of the aerodynamic center. Calculate the coefficient of the moment about the aerodynamic center. Plot on graph paper the variation of the center of pressure location Xcp/c as a function of a (Comment on the results) a (degrees) -2.0 0 2.0 4.0 6.0 8.0 10.0 12.0 14.0 CI 0.05 0.25 0.44 0.64 0.85 1.08 1.26 1.43 1.56 Cd 0.006 0.006 0.006 0.007 0.0075 0.0092 0.0115 0.0150 0.0186 Cm.c/4 -0.042 -0.040 -0.038 -0.036 -0.036 -0.036 -0.034 -0.030 -0.025