Liquid hydrogen (at 20.27 K) and liquid oxygen (at 90.17 K) are burned in the combustion chamber (with an infinite area ratio) of a rocket.
The combustion products contain H
2, O
2, H, O, OH, HO
2 and H
2O
2
This application will calculate
- the adiabatic flame temperature and composition of combustion products
- the pressures and temperatures in the throat and exit
- and the theoretical rocket performance, including the ideal specific impulse, characteristic velocity, sonic velocity and more
Assumptions:
- The combustion chamber is large compared to the throat, hence the assumption of an infinite area ratio
- The flow composition does not change through the nozzle expansion (i.e. reaction rate is slow compared to flowrate). This is also known as "frozen" flow
