Learning Forum

[keans0018]

i had a very high lift coefficient and the pressure coefficient does not looks right. This is the fluent setup: 1) i am using C-domain with 40,000 mesh elements with 40,400 nodes. 2) i am using density-based since i am doing a test with high speed flow where the mach number is above one. 3) model: k-epsilon standard. 4) Air properties, density is in ideal-gas. 5) Cell-zone condition: the pressure gauge is 101325 Pa 6) i am using pressure far-field, the inputs are: 0 pressure gauge; Mach = 2 ; AOA = 1 degree. 7) the solution methods and controls remain default. i ran the iteration between 500-1000

This were the method i get the lift coefficient. Under the Reports > Forces > x ( -ve sin 1) ; y (+ve cos 1). This will print out on the console, and the lift force was about 11 KN, i will put this value into excel and calculate the lift coefficient by rearranges the lift equation. which is about Cl= 11.0 ++ 

I am seeking for help from my lecturer, he can't help me either since he is not familiar in supersonic flow conditions. 

 

The Lift force: 11,342.087 N, its very high... really needs help

[Raef.Kobeissi]

Whats the dimensions of the airfoil? Did you make sure you set the case with the right units? How does your residual error plots look like, could you attach a picture of it + velocity contour please.

Best Regards

[keans0018]

what do you mean about the right units ? would you mind to mention more specific for the word " right units" ? the chord length of this 2D biconvex airfoil is 1meter, the thickness is 6%. 

 

[keans0018]

i am not sure this is the residual error plot that you mentioned before. i really need your help, since i tried different way, all i need is how could i find lift,drag and pressure coefficient. i knew this pressure coefficient is completely wrong. 

[keans0018]

ok i got my mistakes now, i forgot to change the velocity from my equation, that's why my lift coefficient went wrong. Thanks anyway.