CL/CM - α charts for tube "airfoil" tails?


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Eduardo Fadul

Well-Known Member
Jun 21, 2010
São Carlos, SP - Brazil
Hi everyone

I am working with a plane that owns a tail like the RANS s12, however I do not have any kind of chart for those "airfoils" made of bended tube.

Does any one have or knows how to calculate de CL and CM Vs Alpha charts for those "airfoils"?

Best regards


Well-Known Member
Jun 12, 2013
Novi Sad, Vojvodina
as StarJar said its a flat plate, symmetric airfoil and as such it has Cm=0 when elevator in neutral. I would advice you to play with JavaFoil, it has "flat plate" airfoils, do not trust CL max but dCL/d alpha should be ok. You can set flap there (or in your case elevator). That for the very preliminary results. However, tails are surfaces of low aspect ratio, and flow will be absolutely influenced by 3D effects. So, if your tail is like in your avatar, that is different than Rans s12. Your tail seems to have higher AR, and has that fins at tips; all that makes it more efficient (steeper lift curve, higher lift coeffic, lower max angle of attack) than in case of rans tail.

as a first approximation you can use 2d airfoil data reduced to 3d surface by multiplying with this factor: AR/(AR+2,5)

means if your airfoil (2d) has Cl=1, your tail with AR=4 will have roughly CL=1*4/(4+2,5)=0,61

NACA had performed many wind tunnel tests of various tail configurations in pre-WWII period, and the results are published in their Technical Reports and Notes, if you are persistent you will eventually find what you are looking for HERE