الفهرس 
CHAPTER 1Only 14 pages are availabe for public view
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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.