Span is the wingspan of the airplane. It is the distance from one wing tip to other. It is measured in a straight line and is independent of the wing shape and sweep.  Lift per unit span produced by a wing is given as ${L}’ = \frac{1}{2} \rho _{\infty}V_{\infty}^{2} c\left ( 1 \right )c_{l}$. Here, $\frac{1}{2} \rho_{\infty}V_{\infty}^{2}$ is called the dynamic pressure , $q_{\infty}$.  Coffeicent of lift, $c_{l}$, for this airfoil will be $$c_{l} = \frac{{L}’}{q_{\infty}c\left ( 1 \right )}$$ $$\Rightarrow c_{l} = \frac{1400}{\left (\frac{1}{2} \times 1.225 \times 70^{2} \right )\left ( 2.3 \right )\left ( 1 \right )}$$ $$\Rightarrow c_{l} = 0.203$$ Therefore, coefficient of lift for this airfoil is $0.203$. For, NACA $0012$ airfoil from the lift curve -slope for this airfoil, the angle of attack is $2^{\circ}$.