Hi Im simulating the flow over a forward swept wing using steady state simualtions carried out over various angles of attack. Im trying to validate my simulations with that of a paper and im getting different values for my lift coefficient, just want to make sure im working it out correctly. In the paper it says the force coefficients are non-dimensionalised with the free stream dynamic pressure and the wing planform surface.

I have set wall forces to be printed and have taken the total forces in the x y axis from the .erd file that is generated. The total y force is 422.184

Ive calculated the free stream dynamic pressure as 0.5*rho*V^2 where rho is 2.01398 and V is 103.02m/s (Mach No. is 0.3)

therefore if the wing planform area (of just half the wing as I use a symplane) is 0.17283, I get a lift coeff of 422.184/(0.5*rho*V^2 * Area) = 0.228

Ive also used the acoeff command and it gives me the same result of 0.228

In the paper the cl at this angle of attack (4 deg) is ~0.285.

In the paper the mesh they use extends beyond the wingtip while in my solution the mesh just covers the wing from the root chord to the wing tip. At the root I have a symmetry plane and at the tip I have a riemman boundary plane. Becuase Im not simulating wing tip vortices am I getting a lower cl than the paper? Is the way Im calculating cl correct?

Im also unsure of how to get the lift acting on each section of the wing in order to plot a chart of spanwise lift coefficient. I know that how to calculate lift coeff for each section using the local chord but how are local forces found on the model?

thanks in advance for any help

[arunintn]

Hi Abid,
I just starting my project on forward swept wing but i'm modeling a concept jet. If you like we both can help each other I like to know title of the paper you are referring now.