A wing uses the same airfoil for all sections, and the airfoil's drag polar and lift curve slope is shown below

At a wing angle of attack of $7$ degrees, the lift coefficient along the wing $($across the span$)$ is $0\mathrm{.}75$ at every station. Calculate the viscous drag coefficient for the entire wing. The airplane has a constant chord of $2$ m $,$ and a wingspan of $14$ m $.$