Reasonable error in lift and drag coefficients? -- CFD Online Discussion Forums

May 17, 2011, 07:54	Reasonable error in lift and drag coefficients?	#1
xTamx420 New Member Tom Join Date: May 2011 Posts: 14 Rep Power: 14	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.

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