The airfoil on the Lockheed F-104 straight-wing supersonic fighter is a thin symmetric airfoil with a thickness ratio of 3.5 percent. Consider this airfoil in a flow at an angle of attack of 5°. The incompressible lift coefficient for the airfoil is given approximately by cl = 2πα, where a is the angle of attack in radians. Estimate the airfoil lift coefficient for (a) M = 0.2, (b) M = 0.7, and (c) M = 2.0.
We need at least 10 more requests to produce the solution.
0 / 10 have requested this problem solution
The more requests, the faster the answer.