# Consider a thin, symmetric airfoil at \(1.5^{\circ}\) angle of attack. From the results of thin airfoil theory, calculate the lift coefficient and the moment coefficient about the leading edge.

From the results of thin airfoil theory \[c_{l}=2\pi\alpha=2\pi\left ( \frac{1.5}{57.3} \right )=0.164\;\mathrm{radians}\]

\[c_{m,le}=-\frac{c_{l}}{4}=-\frac{0.164}{4}=-0.041\]