How calculated the drag coefficient lift coefficient with pressure coefficient?

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:

1. (CD, CL) = sum(Cp(i) * L(i)* n(i))/(reference length)

-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.

2. like the below equation in the screenshots:

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