Consider a wing with a rectangular planform. The wingspan is b = 11 m and the chord is c = 1.45 m. The wing is composed of cambered airfoil sections with a constant lift slope of m 0 = 6.5 per radian and a zero-lift angle of attack of -2 degrees. The wing has twist with the geometric angle of attack equal to 5 degrees at the root (y = 0) and 1 degree at each wing tip (y = +/- b/2). Use lifting line theory to estimate the wing coefficients of lift CL and induced drag CDi. To simplify your analysis, only two terms in the Fourier expression were included and evaluated at y = 0 and y = b/4. The resulting Fourier coefficients are A1 = 0.01963 and A3 = -0.00144. You may use this result as a starting point in your solution