Abstract
The present study focuses on the combustion of liquid propellants used in engines for space propulsion applications (satellites and launch vehicles). The combustion of the following propellant combinations was studied: hydrogen/oxygen, hydrocarbon fuel/oxygen, hydrocarbon fuel/hydrogen peroxide, hydrazine/nitrogen tetroxide, Mono-methyl hydrazine/nitrogen tetroxide and Unsymmetrical dimethyl hydrazine/nitrogen tetroxide. The purpose of this paper is to determine the combustion flame temperature and the other thermochemical parameters (combustion products politropic parameter and molecular mass) as function of mixture ratio by mass of oxidizer to fuel (O/F). Furthermore, the vacuum specific impulse was also calculated at assumed pressure conditions of propulsion engines to show the effect of mixture ratio of the propellants on the performances. For the determination of the equilibrium composition at assumed temperature, nonlinear systems of equations were solved numerically using Lieberstein’s method. Some results were presented and compared with previous research in this area, the comparison shows good agreement between the results and the difference is less than 5%. As example, three cases of bi-propellant engines were studied using the programs of combustion and frozen fluid flow approximation: Space Shuttle Main Engine (SSME), Zenit second stage engine (RD-120) and Ariane-5 upper stage engine (Aestus).
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