[antonio_ing]

Hi everybody

i'm trying to calculate the lift coefficient around a NACA0015 airfoil with the spalart-allmaras model. I have created the mesh with gmsh using two surfaces (profile1 and profile2) for the higher and lower profile surface. I have found some post where someone descrived how to calculate the force coefficients. But, if i take the profile at 3 degrees of attack the lift coefficient should be 0.5 (from xfoil) but it is 0.02 from OpenFoam.
The chord is 0.1m, the freestream velocity 10m/s and the Re=1e5. This is the controlDict file:

/*--------------------------------*- C++ -*----------------------------------*\
| ========= | |
| \ / F ield | OpenFOAM: The Open Source CFD Toolbox |
| \ / O peration | Version: 1.5 |
| \ / A nd | Web: http://www.OpenFOAM.org |
| \/ M anipulation | |
\*---------------------------------------------------------------------------*/
FoamFile
{
version 2.0;
format ascii;
class dictionary;
object controlDict;
}
// * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * //

application simpleFoam;

startFrom startTime;

startTime 3000;

stopAt endTime;

endTime 3010;

deltaT 1;

writeControl timeStep;

writeInterval 5;

purgeWrite 0;

writeFormat ascii;

writePrecision 6;

writeCompression uncompressed;

timeFormat general;

timePrecision 6;

runTimeModifiable yes;

functions
(
forces
{
type forces;
functionObjectLibs ("libforces.so"); //Lib to load -> dylib on Mac and so on Linux
patches (profile1 profile2);//Name of patche to integrate forces
rhoInf 1.0; //Reference density for fluid - can be changed later ...
CofR ( 0 0 0);
}
forceCoeffs
{
type forceCoeffs;
functionObjectLibs ("libforces.so");
patches (profile1 profile2);
rhoInf 1.0;
CofR (0 0 0);
liftDir (0 1 0);
dragDir (1 0 0);
pitchAxis (0 0 0);
magUInf 10.0;
lRef 0.1;
Aref 0.1;
} );

// ************************************************** *********************** //

any suggestions?

thanks in advance

[wolle1982]

hi,

- according to your Aref=0.1 and t=0.1m your z-depth is 1m right? check that.

- try to use only one patch on the wing

- if the airfoil is on 0deg and your inlet-stream is 3 degrees angle of attack, then you also need to correct the lift-drag vector in the controlDict (or afterwards divide by cos(alpha) )

- for t=0.1m your deltaT seems a little high. how does ist calculate (what solver do you use)??. why you start from 3000s?

- your rhoinf seems too little. where did you get the cl=0.5? check what rho they used.


greetz

[madad2005]

"- if the airfoil is on 0deg and your inlet-stream is 3 degrees angle of attack, then you also need to correct the lift-drag vector in the controlDict (or afterwards divide by cos(alpha) ) "

Your corrected values will be:

Clnew=Cl*cos(alpha)-Cd*sin(alpha)
Cdnew=Cd*cos(alpha)+Cl*sin(alpha)

I've checked the 0015 at 3deg with my own panel method and it gives a Cl=0.367 with 200 panels, including a compressiblity correction. Lifting-line theory gives Cl=0.328 so we are in the right ballpark.

What solver are you using?

[antonio_ing]

hi guys.

thanks for the quick reply.

Effectivelythe thickness of the wing was not 1 m but 0.1 m so the Cl reached 0.19. So it is still low but not so much :-)
I'm using The solver simpleFoam with the Spalart_Allmaras turbulence model. My grid is just with only 1 cell in the z direction. The profile is twisted of 3 degrees with a freestream velocity direction 0f (1 0 0). The drag coefficient is instead too high (0.24) giving a total force coefficient of 0.311. Do you know what rho simplefoam uses in the calculations? any other ideas?

[antonio_ing]

and, also the minimum pressure coefficient is lower than the xfoil's one.

[antonio_ing]

i mean lower in absolute value :-)

[antonio_ing]

I did another simulation with 6 degrees AOA and the lift coefficient is always lower than the thin airfoil theory. Now the Cl from OF is 0.3845 and the Cl from theory is 0.6579. In respect to the previous case i noted that the ration between the CL_OF and CL_theory is constant and equal to 1.7.
Also, the drag coefficient is still higher than i could expect because now it is 0.25118 and with 3 degrees was 0.2448. any suggestions?

[ruby_nuaa]

Hi ,antonio_ing
Sorry to bother you .I'm new to OF. I 've got a problem caculating the lift coefficient of 2D Naca airfoils. The Cl is usually too low compared with the experiment data,same situation like you. I searched for a solution for this and got this thread. I know this thread was posted a long time ago, so I wonder if you have solved this problem yet? A solution or a theory,or a explaination ,enligten me.
3Q