Wrong lift coefficient and Cp on supercritical airfoil
 

Hello,

I have an issue running a simulation of an euler flow around a SC(2)-714 Supercritical airfoil (in 2D).

I used the following .cfg file : inv_NACA0012.cfg of the "euler" folder in the SU2 test cases and changed it to fit with another mesh and it runs well.

Nevertheless, the results i get from this simulation seems wrong to me as it differs from a NASA technical record regarding the lift coeff (1.5 instead of 0.67 according to airfoiltools.com) and I have a strong shock (much stronger than for a NACA0012) whereas a supercritical profile is supposed to reduce this shock.

As I am a beginner in CFD, I sadly have no idea where does these differences come from.

Does anyone here has ever had the same kind of issue and could someone give me some clue about where I could have made a mistake?

I join my.cgf file, I am pretty sure my mesh is all right.

Thank you for your help.

Lucas


https://www.cfd-online.com/Forums/da...EsIDAuOTUgKQo=

Same problem
 

Hey!

Even I am facing a similar problem, although I am analysing in ANSYS Fluent.
I am doing the analysis on the GA(W)-2 airfoil at a Reynolds number of 5.7 million at mean sea level atmospheric conditions. Using the data in the atmosphere tables, I derived my Mach number.

I used ANSYS Meshing to generate a structured C grid about the airfoil with a radius of the semicircle = 12.5c and the width of the rectangle = 15c (c= chord length of airfoil). I gave boundary condition for the domain boundaries as 'Pressure Far Field' with Gauge Pressure 0, Mach number as derived by the process mentioned above and direction components as per angle of attack. For 0 alpha,my Cl value from simulation matches with that from experimental data. But for higher angle of attacks, simulation values are well below the experimental values.

I have tried every permutation of the boundary conditions which are valid, but that only deteriorates the result. Someone please help. :confused: