Hi,

I am simulating some simple performances of NACA 0012 airfoil, but I have some problems with the results I get: they don't fit the theoretical values. I've studied them and I have found out that the source of my problems are the Cl and the CD coefficient (there's a difference between the results of Fluent and the theoretical values I have). Maybe I have wrong theoretical NACA 0012 curves and values.

Can anybody tell me where I can find the NACA 0012 polar curve, the Cl - angle of attack curve, and some other theoretical information like these?

Thank you.

Luis

Hi, carry out example file from cornhill university web-site, you will get a clear idea of how to carry out aerofoil CFD problems.

I am carrying out airfoil example file of Fluent. would you like share the concept of theoretical calculation of lift and Drag force.

It will be very helpful for me.

with regards san

Hi, I have the same problems, I am simulating the performance of a naca0012 with trapped vortex. how many cell are in your mesh?

Hi, please what's the internet address of cornhill university, I've not found it.