Dear all

I am having a problem with the pressure distribution over the surface of an airfoil. I have my own experimental data for lift, drag, and pressure distribution. The flow over the airfoil is separated in real life, and I have it separated in fluent. I have the lift accurate and the drag accurate but the pressure distribution constructed through xy plots is well off.

I don't understand how the lift, and drag can be accurate but the pressure is wrong. I am plotting static pressure against the distance along he airfoil. Has anybody had this problem? What was the remedy? Do I need to use surface integrals to work out the pressure distribution? Could it be to with reference pressures? I have set the gauge pressure on the pressure outlet boundary to zero. The wind tunnel discharges to atmosphere so this is correct, or is it? The operating pressure is 1 atmosphere. But the reference location (0,0) happens to coincide with the leading edge of the airfoil, does this matter?

Need help, a lot of it please!!!

Cheers

Andy

I hope you might be looking at the pressure coefficient values rather than the static pressure values. Did the experimental xy plot has the nondimensional units, if so they are pressure coefficients.

Try out the pressure coefficients for xy plot.

Thanks, but that's not the answer the experimental data is my own from a wind tunnel and all the pressure are just static.

Cheers anyway

Andy

if you know,please tell me.

just see whether you are using cell values or node values. try by changing the node values on and off in xy -plot good luck