Moment Coefficient of 2D Airfoil
I am currently trying to find how to get the leading edge moment coefficient of a one meter NACA 16006 and I am not getting the right results. I know that I have set up the physics model correctly because I am getting the correct lift and drag coefficients based on a comparison to wind tunnel data from NTRS. It seems as I have set up the moment coefficient report data correctly and the only things that I can see being the source of my error are the reference radius and the axis origin. I have set the reference radius to 1m and the axis origin to [0,0,0] with the reference coordinate system located at the leading edge. Any insight to this problem would be greatly appreciated.

Hi,
What are the errors of your Cl and Cd values due to the wind tunnel data? The error of Cl should be around %2%4. However I suppose that your Cd error is higher. Pitch moment is generally calculated at x=0.25c. Is your wind tunnel Cm calculated at x=0? 
Yes, my Cl is within that range. As for the moment coefficient, I do not have wind tunnel data for that range but I know it is incorrect because it is positive when there is a positive angle of attack when I was taking it about the quarter chord and the values were about .4.5. I switched to the leading edge because it seemed simpler to set up since the coordinate system was already located there and it is simple to convert to the Cm at the quarter chord.

Interesting,
As a result I think your Cd error is too high to calculate Cm. It is normal. Cd evaluating is very challenging. When I need Cm value, I prefer 2D codes like Xfoil. 
Drag shouldn't have that much influence on the moment coefficient.
Are you sure, you've defined your axis the right direction? In my testcase, it's behaving exactly as one would expect. The only thing I'm unsure about is the reference radius, but there's only a proportional correlation to Cm, so it will not change a positive to a negative number. 
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