The following is about aerodynamics. I included formulas that might be useful to solve this problem Consider using the source panel method to solve a uniform flow V_ with an angle of attack a (alpha) over a straight airfoil with zero thickness (check figure). The source panel method seeks to find X(X) such that the boundary conditions for an arbitrary point on both the upper surface (point A) and the lower surface (point B) are satisfied. Follow the steps below to prove that there is no solution for X(x) for this zero-thickness airfoil. a) Start form the result given that... Vn V2 = \(2) where V_n1 and V_n2 are normal velocity components on the upper and lower sides of the airfoil surface respectively, show that Vn1 = V sina + A(z) 2 X(x) V2 Vaino 2 b) Prove that there is no solution for X(x) unless the angle of attack is zero. 50 Source strength: x(x) A 7 BA Vnz C Fig. 5. Source panel method for a uniform flow over a plat