You have a slim airfoil with the camber line displayed in the diagram beneath. Find as a) The angle of attack at zero lift and the lift slope b) The lift coefficient as a function of the angle of attack c) The moment coefficient about the leading edge as a function of angle of attack d) The center of pressure

2. The table below shows experimental data for the shape and pressure distribution on the upper and lower surface of an airfoil at zero angle of attack. X 0 0.25 0.5 0.75 1 Yu 0 0.0952 0.0922 0.0588 0 Fx = dYu dx 5 Y₁ 0 -0.0254 -0.0144 -0.0052 0 PL Pd) d: dx Y = Yu (x) Calculate the drag and lift force on the airfoil by numerically evaluating the integrals = S₁² ( P₂₁ = ['(P₁ - F (P₁ - Pu)dx 0 y = Y₂ (x) Pu 1.000 -1.640 -0.786 -0.212 0 Fy x Pi 1.000 0.589 0.426 0.322 0 Use finite difference approximations of the derivatives and the composite Simpson's rule to evaluate the integrals. Hint: Use central differences to estimate the derivatives wherever possible since they are more accurate than forward or backward differences.

A straight tapered wing has a wingspan of 40 ft and a tip chord of 2 ft. The aspect ratio is 7.5. Calculate the mean aerodynamic chord in feet. (Round off answer with no decimal places, No units, No commas)