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I need to calculate CL, CD, CM coefficients for NACA 0012 airfoil at angle of attack 6 degrees and Re=7.5e5, Mach=0.05. I used Xfoil for getting the CP distribution, Cp vs. x/c graph and Cp-x-y data. With Cp-x and Cp-y data points I tried to calculate CL and CD values with a matlab code with equations given below; enter image description here

I integrate the CP-x, CP-y points which executed with XFoil respectively to getting Cx and Cy. I used trapz command in matlab to integrate these points. However at the end of my calculation matlab code was found that CL=1.1936e-04 and CD=0.7249. However at the graph screen of XFoil CL and CD was different. enter image description here

I can not understand that if the problem was about my solution method or my matlab code? How can I get the coefficients with Cp distribution?

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You have done it correctly. Only mistake appears to be swapping the CL and CD value. Being said that, for the info, this CD mearly is a numerical error due to ds being finite as invisvid flow drag is zero.

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