Learning Forum

[dd1367]

I have been trying to simulate flow over a wing of 1m chord length and 5m span(the airfoil of the wing is goe 464) and inlet velocity of 36.11m/s.I am getting a cl of approx 2.6 which is very different from xflr5 result of 0.6(approx) for the same condition I am using k-omega sst for simulation and I have also provided the inflation layer with help of the y+ value.

Why is there so much difference in the cl value and what can i do to correct the value of my cl in ansys

[Rob]

Have you corrected for area? Have a look at the various NACA tutorials/examples in Help. The coefficients are defined relative to a reference, and that reference is rarely the same as the Fluent default. 

[dd1367]

Hello,thank you very much for the comment,can you pls share any link regarding this and why does the refrence value effect the cl this much??(as in the mathematical/code reason)

[Rob]

Maybe start by clicking on Help and searching for lift coefficient and then looking for the tutorials? Learning on here may also have some relevent material. 

[dd1367]

ok thank you very much for your help