Reasonable error in lift and drag coefficients?
 

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.

hi,

did ur solution converge?

have u set proper reference values?

Cheers

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

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

How do you check y+?

Read this http://my.fit.edu/itresources/manual...e517.htm#76840