Hi all
I wrote the vortex panel method code..
it gives me a very nice pressure distribution ...but i have problem in calculating lift coefficient..
i do this:
cpl= (cp at each point above airfoil) *abs( (x(i+1)-x(i)))
cpu= (cp at each point below airfoil) * abs((x(i+1)-x(i)))
cl=cpl-cpu

it gives unreal values fof cl.
the common shape of cl-alpha diagram is true..
but it goes up linearly to angle 55 then the stall occures. and the max cl is nearly 6..
any body can tell me the problem
here is that part of the code
i work on naca 23015
i read coordinate firstly for pressure side from TE to LE then coordinates from LE to TE for suction side

cpl=0
cpu=0

do i=1,m/2-1
cpl=cpl+cp(i)*abs((x(i+1)-x(i)))
end do

do i=m/2+1,m-1
cpu=cpu+cp(i)*abs(x(i+1)-x(i))
end do

print*,'cpu=',cpu,'cpl=',cpl
c_lift=cpl-cpu