[ht2017]

Hi everyone,

I have a problem while I simulate with the airfoil. My airfoil is A_airfoil with inlet condition is U =51.9 m/s,length =1, attack angle =13.3, M = 0.15.

When I complete my simulation. I see that Cl = 1.56 and Cd = 0.74 but in the paper which I am referring the Cd = 0.0263.
This is my set up, you can see it and give me advice. Why is my Drag coefficient too hight?

hope that you can help me.

sincerely thank you.
A-airfoil.tar.gz