# lift coefficient

 Register Blogs Members List Search Today's Posts Mark Forums Read

 December 29, 2010, 07:07 lift coefficient #1 Senior Member     Morteza Join Date: May 2010 Location: Iran,Islamic Republic of Posts: 159 Rep Power: 7 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

 November 23, 2011, 14:02 #2 New Member   Shibin Mohamed Join Date: Nov 2011 Posts: 8 Rep Power: 5 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.

 November 23, 2011, 17:10 #3 Senior Member     Morteza Join Date: May 2010 Location: Iran,Islamic Republic of Posts: 159 Rep Power: 7 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

November 24, 2011, 05:39
#4
Senior Member

Join Date: Jan 2011
Posts: 235
Blog Entries: 5
Rep Power: 8
Quote:
 Originally Posted by morteza08 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.

 November 24, 2011, 05:51 #5 Senior Member     Morteza Join Date: May 2010 Location: Iran,Islamic Republic of Posts: 159 Rep Power: 7 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

 Thread Tools Display Modes Linear Mode

 Posting Rules You may not post new threads You may not post replies You may not post attachments You may not edit your posts BB code is On Smilies are On [IMG] code is On HTML code is OffTrackbacks are On Pingbacks are On Refbacks are On Forum Rules

 Similar Threads Thread Thread Starter Forum Replies Last Post vinz OpenFOAM Running, Solving & CFD 98 October 27, 2008 09:43 MH FLUENT 0 February 25, 2007 12:48 Rola FLUENT 1 November 12, 2006 14:29 Noé CD-adapco 5 July 13, 2004 10:21 Jan Jagrik FLUENT 0 September 21, 2000 10:21

All times are GMT -4. The time now is 03:10.