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Old   December 29, 2010, 07:07
Default lift coefficient
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Morteza
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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
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Old   November 23, 2011, 14:02
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Shibin Mohamed
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DEAR morteza,

could you pls provide me with the settings in fluent for the same airfoil. I too work on NACA 23015. Its very urgent.
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Old   November 23, 2011, 17:10
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Dear Friend
Although we are using the same airfoils, but the approachs are different. what are wreote above is aboute an old problem in a panel method code,
But i can not understand what you really want???
fluent settings while modelling an airefoil can be found here:
ask me if you have another question...

http://www.google.com/url?sa=t&rct=j...wKhOdQ&cad=rja


regards
morteza_a195@yahoo.com
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Old   November 24, 2011, 05:39
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Quote:
Originally Posted by morteza08 View Post
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..

max cl is nearly 6..
any body can tell me the problem
Are you using panel method coupled with a boundary layer? I suppose you are not talking about inviscid panel methods, since stalling does not occur in inviscid flow. Anyway, to predict stalling your code should be a very complicated one taking into account flow separation/reattachment as well as transition to turbulence.

As for evaluating lift force in inviscid case, it is better to take a sum

Fy=sum cp(i)*(x(i+1)-x(i))

clockwise along whole contour and not separately for top and bottom, as for thick airfoils some LE panels which were at bottom at small aoa could go on top at big aoa.

In viscous case, you should add an y-component of tangential friction forces for each panel to the above pressure force.
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Old   November 24, 2011, 05:51
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dear truffaldino
At the time I wrote this post, I was working in an inviscid panel method.
But, after that I extended my work to couple the viscous effect to my code.
And It is already done...
Anyway, I thank you very much for your response....

Good luck
S.M.A
Iran
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