Dear All,

I have been generating mach .85 at the inlet of my suersonic nozzle, but I was wondering if this would cause problems in generating supersonic flow downstream. I was worried about transonic effects. Does anyone know how to calculate the critical mach number for a supersonic nozzle (is there such a thing)...i know you can calculate it for a specific NACA airfoil.

I hope someone can help me, this is extremely important for our experimental design. Hope to hear from someone soon.

-manish

Seems to me you can use the area-Mach number relation, unless there's something I'm missing.

(A/A*)^2 = (1/M^2)*((2/(g+1))*(1+(g-1)*(M^2)/2))^((g+1)/(g-1))

Where A is your area, A* is the throat area, M is mach number of course, and g is gamma. Be careful, this is an isentropic calculation. You will back out a subsonic and supersonic number. The subsonic number is your "critical" mach number... at that mach number you will get Mach 1.0 in the throat... any Mach number above that will result in normal shocks, no longer isentropic flow...

John Anderson's book "Modern Compressible Flow" is a good resource.

Hope this helps, Jason