pressure distribution over an airfoil
Hello everybody, I am trying ti simulate an airfoil using the solver simpleFoam. I want to have a pressure distribution as if the flux is real. The airfoil I'm using is a NASA/Langley GAW-1, it's used on the piper Tomahawk, the velocity is the cruise velocity =45m/s , and the ambiant pressure should be 797.952hPa at an altitude of 2000m. The file p at the 0 directory is modified in this way Quote:
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It appears that you are imposing the same fixed value pressure boundary condition on the inlet and the outlet. Do you think it is a safe assumption that there is no pressure drop across the airfoil? This is a rather high Re flow so it might be more or less inviscid and you might be right (I'm not too familiar with aerodynamic flows).
Also, simpleFoam is an incompressible solver so the absolute value of pressure is somewhat meaningless, only the gradient participates in the flow equations. |
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what you are defining in the p file is the dynamic pressure, while you are using the values for the static pressure.
if you want to see the actual pressure numbers on your plot, just perform the simulation with the right bc for pressure and add your static pressure value during the postprocessing in paraview using the calculator. |
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Dear appa,
I am a student,now I want to find the accurate coordinates of GAW-1. Thank you very much. |
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