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July 25, 2013, 11:58 
Airfoil lift and drag understimated

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Hello everybody, I am quite new in CFD and in particular in OpenFoam.
I use the 2.1.1 version and I try to simulate the behaviour of a 4415 airfoil at different angle of attack. There were no problem to obtain a "correct" trend of Cp (so pressure) according to experimental values, but lift and drag are understimate of 50%. I guess the error can be associated to the BC of k and omega in kw model and nuTilda in SA, because I choose them according to the parameters used in OpenFoam tutorial. I have a freestream velocity U=100 m/s, solver sonicFoam, k=150 and omega=462 at the inlet fot komega, and nuTilda=0,14 for SA, wall_function for patches type wall. Another question is about the function for calculate forces in controlDict file, I set this for angle of attack=0: functions { forces { type forceCoeffs; functionObjectLibs ( "libforces.so" ); outputControl timeStep; outputInterval 1; patches (profile_wall) pName p; UName U; log true; rhoInf 1.225; CofR ( 0 0 0 ); liftDir ( 0 1 0 ); dragDir (1 0 0 ); pitchAxis ( 0 0 1 ); magUInf 100.0; lRef 0.1254; Aref 0.01292; } } For angle of attack >0 I have to set the correct directions for lift and drag or the function calculates the force resulting on the airfoil and from that split lift and drag? Thanks everybody 

July 25, 2013, 13:54 

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something that could help to understand the problem with k and epsilon (or omega, with all models are the same..)


July 27, 2013, 03:42 

#3  
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Phillip
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Quote:
is your case 2D or 3D, because there are different definitions for lift and drag coefficients? Just set the "lref" and "Aref" values both to 1 and after simulating calculate your coefficients with your definition. For angle of attack > 0, i think you don't have to set new directions for liftDir and dragDir, beacause you want the forces in global coordinates and not in local. To your boundary conditions for the turbulence values:
Greetz, Phil. 

July 27, 2013, 04:42 

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Hi, sorry I forgot it. It is a 2D case, the turbulence intensity is 10% of the velocity and turbulence lenght scale is 0.0965 .
Here the case setup for the kepsilon model. thanks to all 

July 27, 2013, 09:27 

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Quote:


July 27, 2013, 13:20 

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Yes, the chord length is 0.1254 m. The values are experimental, from a report NASA. I check that the flow conditions are the same.


July 29, 2013, 13:45 

#7 
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Hello Kio,
You did not provide the /constant folder. I faced a similar kind of problem like you a few months ago. The solution of my problem was quiet obvious. I made mistake in system/forceCoeff file. I did not take into account the length in zdirection. What I am meaning is that you have to find out the reference area like this: Chord_length * Height of extrusion in zdirection. Anyway check the small detail and keep us posted. Normally a tiny misplaced setting plays the main role in this kind of error. Good luck.
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July 30, 2013, 03:51 

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No the reference area is correct, I calculated area in the way you tell me.
The flow simulated is correct, I cannot understand. The cd and cl coefficient are non dimensional, so I could compare with the coefficients from the report also if the airfoil scale is different I cannot provide you the constant folder because is too big 

September 11, 2013, 04:18 

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Hello everybody, after one month I do not find the problem.
I try different meshes and turbulence models but the problem is still the same, pressure field is correct but cl and cd are understimate. I do not understand if the direction of lift and drag are simply (0 1 0) and (1 0 0) or inclinated with the flux. However also with an angle of attack of 0° cl is understimate. There is another problem with an angle of attack > 0° with the coefficient cd that is negative. I attach an image to show you the reference system, how can I set the lift and drag direction? I think (0 1 0) and (1 0 0) but the cl is positive and understimate of 2030% and cd is negative and in any case too high. I attach also the log file of a case with angle=6° and komega model. Please help me! I have to present the case within two weeks! 

September 11, 2013, 07:15 

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A little doubt: I read all the threads about airfoil simulation on the forum and I observed that in all simulations the airfoil mesh is divided in two patches, top and bottom, but only one coefficient is shown in the log file. How the library choose on which patch calculate forces?


September 11, 2013, 07:57 

#11 
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Hello kio. I am uploading my casefiles. I have validated the results with real data. Although I have simulated for incompressible flow using simpleFOAM. But still I think you might get an idea. Good luck.
https://dl.dropboxusercontent.com/u/...8/NACA0015.rar
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September 11, 2013, 09:56 

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thank you very much Naruto, I tried also with simpleFoam and you boundary conditions (my original solver was sonicFoam) but the result remains the same, the cl is too low and cd too high.
Are you sure your results are right? I find a forceCoeffs.dat in your case uploaded, but with an angle of attack of 13° specified in the file 0/U also your cl was too low! For example the coefficient with xfoil are here: http://airfoiltools.com/airfoil/deta...il=naca0015il 

September 11, 2013, 12:45 

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Yes I am sure that my data are ok. I checked them.
But I am saying again they are for incompressible flow.
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September 11, 2013, 13:37 

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ok, thanks again. Probably the problem is related to the compressible flow


September 13, 2013, 06:41 

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I try with a mesh generated automatically by a script found here
https://www.hpc.ntnu.no/display/hpc/...l+Calculations Although the pressure field is not accurate like that of the previous mesh, cl and cd are calculated exactly! y+ is between 1000 and 5000, I think my previous mesh was better because y+ was between 10 and 30, but the forces were not right. I cannot explain that... 

October 1, 2016, 13:48 

#16 
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jiyou
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Hello Kio, do you have any insight to this problem?
I'm doing a validation for NACA 0012 and 4412, i am facing the same problem as well. Hope you may give you advise... 

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