I have got the Cp Plot on a wing, now I am trying to calculate Cl and Cd. And the angle of attack is 0 degrees, the wing has been split into N Panels, and the normal vector ni for every Panel is already known, The airfoil(AQUILA) is like the picture:
If the friction can be ignored.
For Cl, I integrate Cp with x from 0 to 1;
For Cd, I integrate Cp with y from 0 to breadth of the airfoil.
I have tried several ways:
-Here L(i) is the length of the i ‘th panel, Cp(i) ist the pressure of center point of the i ‘th panel and n(i) is the normal vector for the i ‘th panel.
With these two methods, I can get the same Answer, Cl = 0.09, Cd = 0.006
But actually Cl should be 0.2~0.4.
Could someone give me some advice and explain, I haven’t studied aerodynamics before, but there are a few parts about that in my paper
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