الفهرس 
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Abstract
The effect of mixture ratio, injection pressure and nozzle
area ratio on the performance of the liquid hydrogen/liquid
oxygen propellant rocket engine is studied. Chemical
equilibrium was assumed during combustion in the
combustion chamber while transient-one dimensional frozen
flow was assumed in the nozzle. Euler method was used to
solve the differential equations of the combustion chamber
while the method of characteristics was used to solve the
partial differential equations of the flow through the nozzle.
A computer code was constructed to setup the numerical
solution. For Space Shuttle Main Engine configuration, results
showed that : steady-state specific impulse (thrust per unit
propellant weight flow rate) and thrust are 358 sec and 1.55
MN respectively. These results are within 1.5 to 7 %
difference compared with the most recent published data.
Also, the model indicated that for optimum rocket
performance the mixture ratios (oxidizer flow rate per unit
fuel flow rate) are 6 for maximum specific impulse and 8 for
maximum thrust and the optimum nozzle area ratio is 77.5. KEYWORDS: Rocket, Transient, Combustion, Flow, Simulation, Model