An 8 symmetrical double-wedge airfoil is placed at an angle of attack of 5 in a uniform air stream = at Ma.. = 2 and poo 100 kPa, as shown below. Determine the static pressure and Mach number in regions 1, 2, 3, and 4 and calculate the lift and drag coefficients. Ma = 2 P = 100 kPa -8 (1) (2) Problem 6 (15 pts) For the airfoil flow given in Problem 5 above (i.e., at 5 angle of attack), analyze the flow interactions at the trailing edge, i.e., find the Mach numbers, static pressures, and flow directions in the two flow regions just past the trailing edge (i.e., regions 5 and 6 adjacent to 2 and 4, respectively).