I use a 2D naca2412 , 200000Re, 0 deg.I set refernce area as the airfoil chord.Mesh type is C type.The result of cd is 0.02 but according to xfoil and airfoiltools , the cd should be about 0.01 .
Is there any tutorial source or recommended video whose fluent result coefficient data is accurate?Give me a link plz! Thanks a lot!
The attachment is my project document。
Hello, Xu... Going with the assumption that your setup is correct, the result may be actually acceptable. You should test for other AOA to see if there's also a difference.
Goodluck!
a lil more point... you shouldn't expect CFD results and Xfoil results to be 100% thesame (CFD is based on NS-equations discretizations and Xfoil is based on BEM). Slight variations is allowed. Also, you should do a MEsh sensitivity test to be sure that the mesh size doesn't have an effect .
If you have doubled checked this and still have significant deviation then, we can talk about other possible causes.
Please check your mesh, your first layer of the mesh near the airfoil should be well within viscous sublayer region. The Y+ value criteria must be met for the turbulent flows. Also, Please check your reference values, the value of the area. if you used scaled model, accordingly you should vary area values. Cd=Fd/(1\2*rho*A*V*V)
1. Check your Y+ value in trailing edge. It should have the same value on the wall. Also shape points > 300 is better.
2. Check boundary conditions.
3. Check correct turbulent intensity of experiment case.
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