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problem with lift curve in airfoil

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Old   February 11, 2010, 19:38
Default problem with lift curve in airfoil
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thanos
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Hello,

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.
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Old   February 12, 2010, 01:42
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shakil ahmmed
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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.
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Old   February 12, 2010, 08:09
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Yes but, in the experiment's report the author doesn't refer to gauge presure, so i guess this should be zero.
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Old   February 14, 2010, 01:35
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you can increase the boundary distance....best of luck
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Old   February 14, 2010, 08:26
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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?
Cheers!
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