Text: Information for Problems 1
The exit Mach number and ratio of specific heats for a rocket nozzle are 3.0 and 1.24, respectively. The chamber temperature and pressure are 3,600 K and 880 kPa (assume these to be stagnation conditions). The molecular weight of the gas in the nozzle flow is 22.0 kg/kmol. The throat area is At = A* = 7.854x10^-3 m^2 (FYI, throat diameter 10 cm, approximately 4 inches).
Calculate:
1. Nozzle expansion area ratio, A2/A*.
2. Exit temperature ratio, T0/T2, and exit temperature, T2.
3. Exit pressure ratio, P0/P2, using the temperature ratio, and the exit pressure, P2.
4. Velocity at the nozzle exit using the pressure ratio.
5. Velocity at the nozzle exit.
I need help with finding the throat pressure, please.