Hi everyone! I have posted a message previously about airfoil. Finally,i have achieved 8.7% error for my lift coefficient in Fluent compared to DesignFoil. But,the drag coefficient was getting smaller,around 0.0025. I think it's too small for my NACA 9410 with its angle attack 0 degree. Can anyone help me about this? What should i do to make right or at least nearly right? Thanks,

Both drag and lift coefficients is dependant on a referenced area. For drag, it is usually the projected frontal area. For lift, it is usually the projected horizontal area. Is this how you calculated your drag and lift coefficients to be compared to the FLUENT results?

In the Report menu, under Referenced Values... there's only one value for reference area. Please use the reference area for lift, when calculating the lift coefficients, and use the reference area for drag when calculating the drag coefficient. Try this out, and see if it improve your results or not.

Hope that helps.

[thanos]

Originally Posted by zongtwi
;155679

Both drag and lift coefficients is dependant on a referenced area. For drag, it is usually the projected frontal area.

How can you calculate the projected area if you have an AOA not 0 ?