Lift produced by airfoil varies with angle of attack. Each airfoil has a stall angle or critical angle of attack after which the lift generated decreases.
Every airfoil profile has a unique plot which shows how lift changes with increasing angle of attack. The lift produced by an airfoil is expressed by coefficient of lift or cL. Higher the cL, more will be the lift produced.
Shown in the image is the cL versus alpha (angle of attack) plot of the NACA0012 airfoil (as obtained from AirfoilTools). Here each line is the plot of the airfoil for a different Reynold’s number. As we can see, the cL peaks at an alpha of about 8, and then starts decreasing. Hence we can ascertain that the stall angle of this airfoil is 8 degrees.
MIT’s Xfoil tools says the same thing. After multiple convergence plots for the NACA0012 for alpha varying from 0 – 10, we get the stall angle of 8 degrees, with the maximum cL being 0.84 for a Reynold’s number of 100000.
An interesting fact is that symmetrical airfoils produce zero lift when angle of attack is zero, however asymmetrical airfoils may produce positive or negative lift with zero angle of attack. This is an important factor which is considered while choosing a suitable airfoil for an aircraft.