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)))
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
could you pls provide me with the settings in fluent for the same airfoil. I too work on NACA 23015. Its very urgent.
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...
As for evaluating lift force in inviscid case, it is better to take a sum
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.
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....
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