3. Wing lift effectiveness ${{C}_{{{L}_{\alpha }}}}$ at Mach number ${{M}_{\infty }}=0.2$ Wing lift curve slope ${{C}_{{{L}_{\alpha }}}}$ in subsonic flow → lifting line theory를 수정해서 만들어짐 \[{{C}_{{{L}_{\alpha }}}}=\frac{2\pi A}{2+\sqrt{\frac{{{A}^{2}}{{\beta }^{2}}}{{{\kappa }^{2}}}\left( 1+\frac{{{\tan }^{2}}{{\Lambda }_{c/2}}}{{{\beta }^{2}}} \right)+4}}\left( /rad \right)\] \[\beta =\sqrt{1-M..