Calculation of Moment Coefficient
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.
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)?
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.
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