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thanos February 11, 2010 19:38

problem with lift curve in airfoil

I tried to make the diagram a.o.a v.s Cl for a NACA 0012 airfoil but i have a problem. Even though i have accurate results for angles of attack between 0 and 13, then for greater angles my results are not correct.
More specifically, in a Re=2*10^6 NACA 0012 stalls in a=14 degrees, that means that from that point Cl starts to decrease. However, in my simulations as the angle of attack increases so does the Cl. Does anyone have any idea about what is going wrong?

chord=1m, Re=2*10^6, grid=60000 elements, material=air, u=40m/s, incompressible flow, Spalart-Allamaras viscous model, y+<10

Thanks in advance.

ahmmed04 February 12, 2010 01:42

try with different mesh for angle 14 or up. As you can increase your boundary distance from the foil,or try with increasing or decreasing the gauge pressure bcos gauge pressure has strong effect on the lift force.

thanos February 12, 2010 08:09

Yes but, in the experiment's report the author doesn't refer to gauge presure, so i guess this should be zero.

ahmmed04 February 14, 2010 01:35

you can increase the boundary of luck

thanos February 14, 2010 08:26

You mean to expand my fluid domain in the y and x directions? This is because flow may have gradients of pressure far away from the airfoil, strong enough to change my exact value of Cl in these angles of attack?

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