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Old   November 5, 2009, 20:14
Default How to get corect Cd
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Hi
I want to calculate Cl and Cd on wing...
Gambit:
wing wit NACA2412 Crs-sec
fine structured C type mesh wit 200k cells

Fluent: SA model, vorticity based
density>ideal-gas
viscosity>sutherlands
BC: pre-farfield
Discritization: 2nd order fro all
pre-vel coupling: simplec

using abv conditions fluent predicted Cl and Cd as 1.113 and 0.0164 respectuvely at AoA- 8deg.

Cl seems to be reasonable but, Cd is 100% more, it has to be 0.008.
I am unable to figure out why it is, plz help me how to get corect Cd value.
few pics are attached...
Thnx a lot.
Attached Images
File Type: jpg grid .jpg (42.2 KB, 15 views)
File Type: jpg Pre-cont.jpg (93.6 KB, 14 views)
File Type: jpg vel_vec.jpg (98.5 KB, 14 views)
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Old   November 9, 2009, 09:14
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Hello makaero
My problem is same as yours. When i get my Cl close to handbook but my cd far from hand book. In left and right side there is no boundary so horse tail effect on the airfoil tip cannot be figured. Do you active gravitation property?? I think you should choose SIMPLE as press-vel couple. because it there is contain turbulence.
by the way, what is the property of your BC??
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Old   November 9, 2009, 19:14
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Hi teguhtf

Bc's for side faces are symmetry and abt Cd i dint figered it out yet!
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Old   November 9, 2009, 21:06
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Do you have report of your simulation?? Can you Post here??

Last edited by teguhtf; November 9, 2009 at 22:45.
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Old   November 9, 2009, 23:49
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hi
i dnt knw how to dotht, i hav mentioned in my frst post, wht i hav done...

can u help me wit how does y+ values affect the solution in prediction of drag?

in my case most of the points on wing are 30<y+<80

plz see the fig...

thnq.
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Old   November 10, 2009, 00:13
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Here is the step to write report.
1. Choose report menu in FLUENT window
2. Click summary
3. in report option choose all
4. click print

I'll try to check it.
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3dwing, airfoil, cd drag, naca2412

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