[fletc900]

Greetings.

I'm trying to model a NACA 0006 airfoil to compare it experimental data (from Theory of Wing Sections, I. Abbott & V. Doenhoff). I'm trying to replicate in this case the flow for angle of attack of 4 degrees (so low alpha), and a Reynolds number of 3e6. The flow is incompressible and moves along the +x direction, and I'm running it using simpleFOAM and the Spalart-Allmaras model for turbulence.

The lift coefficient matches the experimantal data (~0.4), but the predicted drag coefficient is negative (i.e, producing thrust), and the solution seems to have converged after 500 iterations. The wierd thing is that I get a positive Cd value (as expected) when the angle of attack is zero, but as soon as I give it an angle of attack, the Cd becomes negative. I must be doing something wrong on the settings. I'm including the entire validation case as *.gz file. Any clues anyone?

Regards,
Bruno

[fletc900]

Solved. Dumb mistake. I had set the angle of attack by specifying <u,v> components of velocity (horizontal and vertical components) while the airfoil was perfectly horizontal. I still had my liftDir and dragDir pointing along the (0,1,0) and (1,0,0) directions which is wrong. This will give you the normal and axial forces on the airfoil, not the lift and drag. Two fixes would be:

1 - Just rotate liftDir and dragDir axial components by whatever angle of attack you are using.

2 - Or use the following equation to calculate it directly:


Cl=Cn cos(alpha) - Ca sin(alpha)
Cd=Ca cos(alpha) + Cn sin(alpha)

Where Ca and Cn are the axial and normal forces that are computed when using dragDir along (1,0,0) and liftDir (0,1,0) with libforces.

Bruno

[aamer]

Hello Bruno.....

lets say that i am translating an airfoil in x direction and the airfoil has an angle of attack of 40 degree. lets assume that instead of using u and v components of velocity, i have rotated my whole grid by 40 degree (and thus airfoil orients by 40 degree), in such case what should be used for Cl and Cd.... should it be (0,1,0) and (1,0,0) respectively now?

[fletc900]

Good afternoon Aamer. Yes, if you rotate the mesh by 40 deg, and maintain the flow only along the x-direction, then the lift will be along (0 1 0) and the drag will be along (1 0 0). Just remember, by definition, that the lift is always perpendicular to the flow and the a drag is always in the same direction of the flow. Let me know if this is not clear...

Cheers
Bruno

[aamer]

Dear Bruno.....

if the airfoil is moving in still air, then there is no flow around it.... so how can we find out the direction of flow then.....
if airfoil is moving in -x direction (in still air), do we have to assume that the flow direction will be in +x direction....?

[fletc900]

Good morning Aamer.

It doesn't matter if the airfoil is moving or if it is static with wind blowning at the same velocity upon it. The same relative wind velocity is encountered for the airfoil in both cases. Perhaps this link will help you better understand: http://www.allstar.fiu.edu/aero/flight33.htm


Cheers
Bruno

[aamer]

Thanks Bruno.... for the explanation and the link

[Rophys]

Hi everyone,

I am trying to calculate the drag, lift and pitching-moment coefficients for a flow over a probe. The code structure below works fine when the flow is perpendicular to the probe. However, when I changed the angle of attack I got a rude difference for lift and pitching-moment when compared with experimental data (the drag values is OK).

I guess, this difference is related to the vectors (liftDir (0 1 0); dragDir (1 0 0); pitchAxis (0 0 1)). I this this vectors values is just valid for a parallel flow, isn't it ?

Anybody knows how to change this vectors according to a certain angle of attack ? My mesh still the same and I just rotate my geometry in 10 degree.

Thank you very much.

Rophys

forces
{
type forces;
enabled true;
functionObjectLibs ( "libforces.so" );
outputControl outputTime;
patches (probe);
directForceDensity true;
fDName fDMean;
CofR (0.02 0 0);
log on;
}

forceCoeffs
{
type forceCoeffs;
functionObjectLibs ("libforces.so");
patches (probe);
outputControl outputTime;
fDName fDMean;
rhoInf 1.73e-5;
CofR (0.02 0 0);
liftDir (0 1 0);
dragDir (1 0 0);
pitchAxis (0 0 1);
magUInf 1503.1;
lRef 0.05;
Aref 9.817e-4;
}

[Alhasan]

Originally Posted by fletc900

Solved. Dumb mistake. I had set the angle of attack by specifying <u,v> components of velocity (horizontal and vertical components) while the airfoil was perfectly horizontal. I still had my liftDir and dragDir pointing along the (0,1,0) and (1,0,0) directions which is wrong. This will give you the normal and axial forces on the airfoil, not the lift and drag. Two fixes would be:

1 - Just rotate liftDir and dragDir axial components by whatever angle of attack you are using.

2 - Or use the following equation to calculate it directly:


Cl=Cn cos(alpha) - Ca sin(alpha)
Cd=Ca cos(alpha) + Cn sin(alpha)

Where Ca and Cn are the axial and normal forces that are computed when using dragDir along (1,0,0) and liftDir (0,1,0) with libforces.

Bruno

Hello Everyone,
Can someone tell me what are the normal and axial forces he is talking about here the Ca and Cn ??? is it the Cd and Cl computed from when using dragDir along (1,0,0) and liftDir (0,1,0) or something else because when I use the Cd and Cl calculated from computations in the place of Ca and Cn to get the actual Cd and Cl I am getting a list of negative and positive values for a range of angle of attacks from -5 to 20

Edit: Answer: yes the Ca and Cl are the computed lift and drag using dragDir along (1,0,0) and liftDir (0,1,0) and the reason I was getting a mixture of positive and negative values is because I was calculating in Excel and I did not use Radians. you have to use RADIANS when using sin and cos to get the right answer.


Thanks,
Hasan K.J

[sangeet]

Originally Posted by fletc900

Solved. Dumb mistake. I had set the angle of attack by specifying <u,v> components of velocity (horizontal and vertical components) while the airfoil was perfectly horizontal. I still had my liftDir and dragDir pointing along the (0,1,0) and (1,0,0) directions which is wrong. This will give you the normal and axial forces on the airfoil, not the lift and drag. Two fixes would be:

1 - Just rotate liftDir and dragDir axial components by whatever angle of attack you are using.

2 - Or use the following equation to calculate it directly:


Cl=Cn cos(alpha) - Ca sin(alpha)
Cd=Ca cos(alpha) + Cn sin(alpha)

Where Ca and Cn are the axial and normal forces that are computed when using dragDir along (1,0,0) and liftDir (0,1,0) with libforces.

Bruno

This was a life-saver. Was completely lost

[Muhammad Bilal Anwar]

Originally Posted by fletc900

Solved. Dumb mistake. I had set the angle of attack by specifying <u,v> components of velocity (horizontal and vertical components) while the airfoil was perfectly horizontal. I still had my liftDir and dragDir pointing along the (0,1,0) and (1,0,0) directions which is wrong. This will give you the normal and axial forces on the airfoil, not the lift and drag. Two fixes would be:

1 - Just rotate liftDir and dragDir axial components by whatever angle of attack you are using.

2 - Or use the following equation to calculate it directly:


Cl=Cn cos(alpha) - Ca sin(alpha)
Cd=Ca cos(alpha) + Cn sin(alpha)

Where Ca and Cn are the axial and normal forces that are computed when using dragDir along (1,0,0) and liftDir (0,1,0) with libforces.

Bruno

Hey Bruno!
I am working on the sweeping motion of the wing. I have encountered a similar problem.
i have a query regarding the force vectors for drag calculation....

Initially, The chord of wing is in X direction and the span is in Z direction. So for drag calculation, in SOLVE>> mONITOR>> Force>> Drag, i have given X=1, Y=0, Z=0.

NOW the wing initially starts from rest, it rotates about Y axis and achieves constant angular velocity after some time. .Now as the wing rotates, the drag components of X=1, Y=0 and Z=0 will not give correct drag anymore. For instance After wing rotation of 90 degree, The chord will be along Z axis (i.e no more along X axis) and span along X -axis .
How is it possible to ensure that fluent keep calculating drag force, and changes force vectors of drag itself in order to calculate correct drag force through out the rotation.

Thanks in advance

[sjsivabharathy]

Hi Bruno,

I have a few questions regarding the domain size and initial condition.
1) What is your domain size? How far are the upper and lower surface boundary. Because when you simulate the flow by changing the angle
of attack, the v component of velocity is zero and to simulate the angle of attack flow, one has to specify u = U cos(alpha) and
v = U sin(alpha) on the inflow boundary. For large angle of attack
does the wake interfere with the surface boundary.
2) What initial condition would you specify to start the simulations?

Thanks,
Jawahar.

[vinutha]

i have already gven the angle of attack in the model itself. i'm still getting negative cd values after 500 iterations. lifrDIR (0 1 0) and dragDIR(1 0 0). velocity i have given only in x direction. Would be grateful is someone cleared this problem as i'm not getting it.

[sjsivabharathy]

Hi Vinutha,

Is your flow converged and what about the value of lift coefficient?

[mehulsasvadiya1996@gmail.]

Originally Posted by Alhasan

Hello Everyone,
Can someone tell me what are the normal and axial forces he is talking about here the Ca and Cn ??? is it the Cd and Cl computed from when using dragDir along (1,0,0) and liftDir (0,1,0) or something else because when I use the Cd and Cl calculated from computations in the place of Ca and Cn to get the actual Cd and Cl I am getting a list of negative and positive values for a range of angle of attacks from -5 to 20

Edit: Answer: yes the Ca and Cl are the computed lift and drag using dragDir along (1,0,0) and liftDir (0,1,0) and the reason I was getting a mixture of positive and negative values is because I was calculating in Excel and I did not use Radians. you have to use RADIANS when using sin and cos to get the right answer.


Thanks,
Hasan K.J

Hello


I am doing airfoil2D simulation in openFOAM. My freestreamvelocity come from downward Y-direction to the Up-ward Y-direction and airfoil is movingWallVelocity. So my free stram velocity is perpendicullar to the drag orce and parallel to the lift force So Is it possible to get -ve drag force? Bcz I got -ve drag force . Please help out of this problem. Thanks.....