High CL lift coefficient in Stall Angle of attack airfoil
 

I did a flow simulation over airfoil NACA 0015 using ansys fluent. Everything is ok in values of CL and CD for AoA 0 degree till 10 degree. But when i did simulation in AoA 15 degree values of CL is very high arround 1, it should be the value of CL in this AoA is low arround 0.4 cause 15 degree is near the stall angle (due to experiment result).
Whats wrong with my simulation?
Thanks before