Cl = Clo + (Cla * alpha)
where:
Cl = coefficient of lift
Clo = coefficient of lift at zero angle of attack
Cla = lift curve slope (the slope of Cl versus alpha)
alpha = angle of attack
From the lift formula Cl = L/(0.5*rho*V^2*S), the only two variables for a certain value of Cl at 0 alpha (producing a certain L) are V and S. Where S depends on the chord and V on the true wind over the airfoil.