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Lift and drag coefficient with strange values for NACA airfoil 

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December 15, 2008, 06:07 
Hi everybody.
I'm trying to

#1 
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antonio segalini
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Hi everybody.
I'm trying to simulate the flow around a NACA airfoil with the simpleFoam solver using the Spalart Allmaras model with Re=1e6. Strangely, the boundary layer is very thick resembling flow separation. Also, the drag coefficient is 0.2 which, for an airfoil at 0 AOA, is quite high and the lift coefficient is 0.07 while xfoil gives me 0.5. I can understand that the drag should be higher using a turbulence model, but why the lift coeff is so low?. I'm posting here all the files in the system directory just in case  /** 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 4000; stopAt endTime; endTime 8000; deltaT 1; writeControl timeStep; writeInterval 100; purgeWrite 0; writeFormat ascii; writePrecision 6; writeCompression uncompressed; timeFormat general; timePrecision 6; graphFormat raw; 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 100.0; lRef 0.1; Aref 0.01; } ); // ************************************************** *********************** //  /** 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 fvSchemes; } // * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * // ddtSchemes { default steadyState; } gradSchemes { default Gauss linear; grad(p) Gauss linear; grad(U) Gauss linear; } divSchemes { default none; div(phi,U) Gauss upwind; div(phi,k) Gauss upwind; div(phi,epsilon) Gauss upwind; div(phi,R) Gauss upwind; div(R) Gauss linear; div(phi,nuTilda) Gauss upwind; div((nuEff*dev(grad(U).T()))) Gauss linear; } laplacianSchemes { default none; laplacian(nuEff,U) Gauss linear limited 0.7; laplacian(nu,U) Gauss linear limited 0.7; laplacian((1A(U)),p) Gauss linear limited 1; laplacian(DkEff,k) Gauss linear corrected; laplacian(DepsilonEff,epsilon) Gauss linear corrected; laplacian(DREff,R) Gauss linear corrected; laplacian(DnuTildaEff,nuTilda) Gauss linear corrected; } interpolationSchemes { default linear; interpolate(U) linear; } snGradSchemes { default corrected; } fluxRequired { default no; p; } // ************************************************** *********************** //  /** 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 fvSolution; } // * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * // solvers { p PCG { preconditioner DIC; tolerance 1e08; relTol 0; }; U PBiCG { preconditioner DILU; tolerance 1e08; relTol 0; }; k PBiCG { preconditioner DILU; tolerance 1e08; relTol 0.1; }; epsilon PBiCG { preconditioner DILU; tolerance 1e08; relTol 0.1; }; R PBiCG { preconditioner DILU; tolerance 1e08; relTol 0.1; }; nuTilda PBiCG { preconditioner DILU; tolerance 1e08; relTol 0.1; }; } /* k BICCG 1e06 0; epsilon BICCG 1e06 0; R BICCG 1e06 0; nuTilda BICCG 1e06 0; */ SIMPLE { nNonOrthogonalCorrectors 2; } PISO { nCorrectors 1; nNonOrthogonalCorrectors 1; /* pRefCell 0; pRefValue 0;*/ } relaxationFactors { p 0.3; U 0.7; k 0.5; epsilon 0.5; /* R 0.7;*/ nuTilda 0.7; } // ************************************************** *********************** // 

December 15, 2008, 09:21 
I found the porblem. Close to

#2 
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antonio segalini
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I found the porblem. Close to the Trailing edge the mesh was not so strong. I have increased there the number of cells there and now everything is better...!


December 16, 2008, 04:51 
what did you use to mesh it?

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Leonardo Honfi Camilo
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what did you use to mesh it?


December 17, 2008, 08:00 
gmsh. It is nice and powerful.

#4 
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antonio segalini
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gmsh. It is nice and powerful. I wrote the mesh file through matlab and then compiled it with gmsh.


December 19, 2008, 04:49 
Hi Antonio,
did you use tet m

#5 
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Hi Antonio,
did you use tet mesh, or a transfinite one? How did you specifie the b.l. stretching? I'm managing to use gmsh for airfoil meshing, but I have problems with wall resolving.. 

December 29, 2008, 03:29 
you are right ivan. Actually i

#6 
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antonio segalini
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you are right ivan. Actually i have used the transfinite algorithm close to the profile surface and in the first part of the wake. With this method you can easily define the vertical or horizontal stretching of the cells close to the profile. The gmsh tutorials t3.geo and t6.geo helped me a lot in doing that


December 29, 2008, 10:17 
Antonio,
I tryed to do the sa

#7 
Senior Member

Antonio,
I tryed to do the same, and my guess was to use the tet meshing far from the foil in order to reduce the number of elements. Did you do something similar? With which kind of results? 

December 29, 2008, 15:23 
actualy I did the opposite sin

#8 
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antonio segalini
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actualy I did the opposite since it is more important to have a cartesian grid close to the airfoil in order to simulate correctely the boundary layer evolution.
If you give me your mail I can send you some screenshots of my mesh or the code I have used in gmsh. Hopefully this will help you 

December 30, 2008, 00:48 
hi everyone
i could not

#9 
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NAVEEN.K.M
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hi everyone
i could not solve naca0012 case because my vertices are not working correctly...could u please send me naca 0012 vertices.. 

January 20, 2009, 03:52 
hi everybody
i am trying

#10 
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NAVEEN.K.M
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hi everybody
i am trying to work on C type domain of naca 0012 airfoil case in openfoam 1.4.1,but i dont hav the vertices and edges for naca 0012 airfoil,can u please send me the vertices and edges for naca 0012 airfiol.. 

January 20, 2009, 04:22 
hi everybody
i am trying

#11 
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NAVEEN.K.M
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hi everybody
i am trying to work on C type domain of naca 0012 airfoil case in openfoam 1.4.1,but i dont hav the vertices and edges for naca 0012 airfoil,can u please send me the vertices and edges for naca 0012 airfiol.. 

January 20, 2009, 04:38 
Naveen,
http://www.aerospac

#12 
Senior Member

Naveen,
http://www.aerospaceweb.org/question...ls/q0100.shtml here you can find the naca 4 digit equation, but did you ever try to google "naca 0012"? 

January 20, 2009, 04:44 
I have left the code, case and

#13 
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Hrvoje Jasak
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I have left the code, case and results for you on:
http://powerlab.fsb.hr/ped/kturbo/Op...P/naca0012.tgz It is unbelievable that all the work you did on this since December 2008 is to ask the same question over and over again. Which school did you go to? Hrv
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Hrvoje Jasak Providing commercial FOAM/OpenFOAM and CFD Consulting: http://wikki.co.uk 

January 20, 2009, 04:49 
hi
yes i tried that webs

#14 
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NAVEEN.K.M
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hi
yes i tried that website but i need the edges of naca 0012 airfoil to work in openfoam 1.4.1...can u send me the tutorial of naca 0012 airfoil in openfoam 1.4.1...including how to solve and postprocessing.. 

January 20, 2009, 04:58 
Dear Naveen,
do you really ex

#15 
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Christian Winkler
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Dear Naveen,
do you really expect other people to do all the work for you? Have a look here (google is your friend) http://www.basiliscus.com/ProaSectio.../AppendixD.pdf http://www.ppart.de/aerodynamics/profiles/NACA4.html Christian 

January 20, 2009, 05:03 
hi
thanks for replying..

#16 
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NAVEEN.K.M
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hi
thanks for replying.... 

September 13, 2012, 12:21 

#17 
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R. P.
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Hi all,
I'm doing some simulations using the OpenFoam 2.1 and I'd like to measure the following aerodynamic forces: Drag Coefficient, Lift Coefficient, Axialforce coefficient, Normalforce coefficient, Pitchingmoment coefficient and center of pressure. How can I state this in the controlDict ? Many thanks. 

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