[fatihcoban]

Hello,

I have been trying to analyze a wing in fluent and trying to get the best cl and cd values at different aoa. I create my model in catia, for mesh i used fluent meshing tool. I create my geometry and mesh model according to cornell's 3d wing analysis tutorial. My problem is that I haven't got a desired value for cl and cd yet. I analyzed it at 0,2,4,6 aoa. At 0 aoa i saw 0.45 cl and 0.035 cd and at 6 aoa i saw 0.61 cl and 0.077 cd values. I know that these values are wrong. When I look at my airfoil's cl and cd values at the same aoa they are very different. Also I analyzed the airfoil using Xfoil and wing using XFLR5 modules then compared the results. See the results below.

Airfoil using xfoil
0 aoa cl 0.54 - cd 0.009
6 aoa cl 1.19 - cd 0.01

wing using xflr5
0 aoa cl 0.6 - cd 0.013
6 aoa cl 1.14 - cd 0.045

wing using fluent
0 aoa cl 0.45 - cd 0.035
6 aoa cl 0.61 - cd 0.077

fluent solution case;
model: Spalart allmaras
material: constant-air-1.007 kg/m^3-1.726 e-5 viscousity
velocity 26 m/s
outlet pressure 0
reference values from inlet-reference area 1.13 m^2-reference lenght is wing chord
solution method coupled-second order

Could you please help me the solve this problem?