Cl = 1400 kg * 9.8 m/s2 / (0.5 * 1.225 kg/m^3 * (95 / 1.94 m/s) * 11.9 m2) = 0.79
That's quite a high lift coefficient, though on the only aerofoil for which I have a polar that, in effect, includes thickness as a parameter (NACA 64 series), it's still in the regime where thicker is better. So I think we'd need more detailed data specific to the aerofoil.