[andromeda91]

Dear All,

I have tried to check the simple aerodynamic analysis for an airfoil with different angle of attack. I used NACA2415 airfoil. The Cm results seems not correct for me. The Cm was to be found positive for alpha bigger than -4 deg. Moreover, there is too much difference in Cm values for different alphas. You can see the results in the attachment.

The airfoil length is 1 m. I used this line to define my point for calculating pitching moment coefficient.
% Reference origin for moment computation
REF_ORIGIN_MOMENT= ( 0.25, 0.00, 0.00 )

Is there anything that I did it wrong in the analysis?
Any help would be appreciated.

[EMolina]

Hi.

Do you have the experimental values for this airfoil? Can you plot in the figure?

Your mesh have a suitable discretization (y+ value if you are running a NS solution)?

Regards

[andromeda91]

Originally Posted by EMolina

Hi.

Do you have the experimental values for this airfoil? Can you plot in the figure?

Your mesh have a suitable discretization (y+ value if you are running a NS solution)?

Regards

Hi Eduardo,

I don't have any experimental data for this airfoil. However, I think that Cm values should not vary too much like the one I got in here.
I think my mesh have suitable discretization for running NS solution.