The airfoil on the lockheed f-104 straight-wing supersonic fighter is a thin, symmetric airfoil with a thickness ratio of 3.5%. consider this airfoil in a flow at an angle of attack of 5o. the incompressible lift coefficient for the airfoil is given approximately by cl=2παcl=2πα, where αα is the angle of attack in radians.
estimate the airfoil lift coefficient for
(a) m = 0.2
(b) m = 0.7
(c) m = 2.0.

We have a problem with three different state of the ratio of flow velocity to speed of sound.

That is,

a) Mach number to evaluate is 0.2, that mean we have a subsonic state.

The equation here for lift coefficient is,

where should be expressed in Rad.

So replacing in equation for subsonic state,

b) In this situation we have a transonic state, so we need to use the Prandtl-Glauert rule,

c) For this case we have a supersonic state, so we use that equation,

i think its c if i'm wrong

because if the tool that you are using is not balanced it there would be false mass.