Cd and Cl values
 

hi,
I am working on an airfoil, naca2412, with 7.25 degrees angle of attack. I want to calculate its drag and lift coofficient on an airfoil in fluent, 3D model. I cant converge these values. I am suspecting of referance values. Where should I calculate it from inlet or an airfoil?? the airfoil is twisting.Attachment 31159 the velocty is in y direction in fluid domain.Attachment 31160 What should I write referance values?? I know the project area of an airfoil in y direction but what is its lenght?? Its chord lenght is changing in every station(in x direction).
thank you all for answers

help :(
 

Is there anyone to help me??

hey,

for the calculation of the reference values you will need the dynamic pressure (calculated by the freestream velocity). so it´s the best way to calculate the reference values from the values you specified at your inlet !
in addition to that force coefficients in your 3d setup only need the projected area. the length is just required for monitoring moment coefficients.

Can you specify what you mean with the phrase "I cant converge these values"?, so we can help you with more precision. Regards.
Gonzalo

Hi Traction, You mean lenght is not important to calculate Cd and Cl in 3D model?? Are you certainly sure that?? So if it is, why drag and lift is changing with lenght in my model?? I mean, If I change only referance lenght value in my model, I get different cl and cd.
Hi Gfoam, I meant that my Cd and Cl values are not equal that I want to get. According to this site http://airfoiltools.com/polar/detail...a2412-il-50000 I should get Cl=0.98, Cd=0.03. But the values that I get Cl is around 1.38 Cd is around 0.88. I am also not sure the values in that site is for 2D or 3D
Thank you for answers...

hey,
regarding the reference values you may look at the fluent user guide (search for task page reference guide and reference values task page, in my guide its chapter 36.10)

regarding your incorrect results - have you checked different meshes or done a mesh independence study ? what kind of mesh do you use ?

as far as i can see the data is for 2d airfoils with an infinite aspect ratio (my xfoil course was some years ago, i cannot remember every detail).
maybe you should take into consideration that your "wing" has no infinite aspect ratio.
in addition to that here will definitely be a (more or less large) difference between the 2d and 3d values because your wing is twisted.

why do you use the projected area in y direction ? if i understand your case correct you should use the area in z-direction (from what i can see in your pictures this area should be larger than in y-direction, this will lower your cl and cd)

Hi Tranction, I will consider your advices. Btw; the air flows in y-direction in fluid domain, so I should use the projected area in y-direction, right?
Thank you...

Dear erayeisik you cant't compare results for a twisted wing (3D) with the results of a 2D simulation of a wing foil section. That's totally wrong. May be you should read some books of Aerodynamics of 2D and 3D foil before proceeding doing simulations with FLUENT. If you want to compare results for your 3D model, may be you should search for experiment or something similar that you can use as a reference to do the validation. Good luck with that.
Gonzalo

Cd and Cl values for an airfoil
 

Hi, I am simulating an airfoil and want to verify the values of Cd mean for different angles of attack. As values for Cl mean are compared and are verified but I did not find the values for mean Cd. Does anyone has the graph of Cd values to compare. please share it. I will highly appreciate.