Gentlemen, I am trying to converge the value of the numerical result of the lift coefficient for the experimental result of the airfoil S1223. I've tried it in several ways and am getting values between Cl = 0.95 and Cl = 0.98, but the experimental value is Cl = 2.23. This for conditions of approximately Re = 380000 and an angle of attack alpha= 10 degrees. In my opinion the problem is the mesh, particularly at the trailing edge of the airfoil ... How can I improve the mesh in this region? I think that if a simulation was done in the edX course for a validation case of the NACA0012 airfoil with a relatively simple mesh and cells similar, we can see in below image, in the trailing edge region, I can also obtain a numerical result close to the result experimental, performing knit procedure in the same way ...
Or does anyone know what can be done so that the value approaches the experimental value?
Gentlemans, thanks for help and attention!
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