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Reasonable error in lift and drag coefficients?

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Old   May 17, 2011, 07:54
Default Reasonable error in lift and drag coefficients?
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Tom
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Hi, I am a new user to Fluent and CFD in general. I have been running simulations on NACA 0012 foils and I was wondering what a reasonable Cd and Cl are? I get around 0.009 for Cd at 0 angle of attack with a reynolds number of 3000000. The data for NACA 0012 in Theory of Wing Sections has Cd around 0.006. Is this a reasonable error or am I doing something wrong? Cl us usually between 0.55 and 0.65 depending on the viscosity model for an angle of attack of 6 degrees which matches the data in Theory of Wing Sections pretty well.
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Old   May 18, 2011, 08:12
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hi,

did ur solution converge?

have u set proper reference values?

Cheers
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Old   May 18, 2011, 08:32
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The solution did converge. I got around 0.009 for Spalart-Allmaras model at a couple different Reynolds Numbers. In Theory of Wing Sections the Cd at 0 angle of attack also is the same number for different reynolds numbers its just that mine is off by 50%. I was just looking at the Fluent manual and I think my mesh might be too coarse. I have the same mesh shape as the one in a tutorial I used with 40000 cells. Is that too coarse?


This is the tutorial I used:

https://confluence.cornell.edu/displ...+Specification
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Old   May 18, 2011, 12:44
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Hi,

1. Check the y+ of the grid. Try to refine the grid to reasonably low y+.

2. Try running with a different turbulence model.

Come back if the results still dont improve.

Cheers
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Old   May 18, 2011, 13:40
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How do you check y+?
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Old   May 19, 2011, 05:00
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Read this http://my.fit.edu/itresources/manual...e517.htm#76840
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