Consider flow over a thin aerofoil at Mach number,M_{\infty}=0.5 at an angle of attack of \alpha. Using the Prandtl-Glauert rule for compressibility correction find a formula for lift coefficient C_{L}.
Consider flow over a thin aerofoil at Mach number, M_{\infty}=0.5 at an angle of attack of \alpha. Using the Prandtl-Glauert rule for compressibility correction find a formula for lift coefficient C_{L}.
Prandtl-Glauert Compressibility correction for a low speed flow is c_{L}=\frac{c_{l,0}}{\sqrt{1-M_{\infty}^{2}}}
We will get c_{l,0} by using thin airfoil theory. Thin airfoil theory is c_{l,0}=2\pi\alpha
\Rightarrow c_{L}=\frac{2\pi\alpha}{\sqrt{1-\left ( 0.5 \right )^{2}}} \\\Rightarrow c_{L}=7.255\alpha\approx 7.26\alpha
