The camber line of an airfoil of chord c is defined as: z()=c(a-x)1-4r2) (2) where e<<1 and the airfoil located at sc/2.Obtain a in order to have the center of pressure at the same location than the aerodynamic center. = cos0 x= / x-ch2 Figure 4:Camber line centered at the origin Note that the camber line is centered at the origin(see figure)thus the transformation will be cos6 PROBLEM7(15%) velocity along the span direction is 0=-1++4g.=2c< 1.Obtain the value of the lift Lof the wing. 2.Determine the induced drag D in the wing 3.Calculate the roll moment about the axis M.PROBLEM 7 (15%)