[Turbomachine]

Dear All,

I need to plot the isentropic Mach number distribution on a compressor blade. I defined an expression using the formula:

http://www.cfd-online.com/Wiki/Isentropic_Mach_number

I am using total temperature in relative frame and static pressure so the formula in cfx post looks like this:

sqrt(((Total Pressure in Rel Frame/Pressure)^(0.4/1.4)-1)*2/0.4)

Even though the plots show high pressure gradients where the shock wave is, the Mach number in front of it is subsonic, which is meaningless.

The problem must be with how the isentropic mach number is defined and used in this case. Can someone help me out with this??

Thank you very much

[rafr894]

Hi, I know it's been a while since you posted this question, but have you figured a way out to your problem?

[ghorrocks]

The equation in the post does not look correct to me.
1) The posted equation uses Total Pressure in Rel Frame (which is a local variable), but the linked definition equation uses total pressure in the freestream outside of the boundary layer, and usually the inlet total pressure is used. These are not local variables.
2) You will need to be careful with the reference pressure in this equation. It would be safer to use absolute pressures in the equation.

[rafr894]

Thanks for that note!