Abstract
The effects of the momentum-flux ratio of propellant upon the combustion efficiency of a gas-centered-swirl-coaxial (GCSC) injector used in the combustion chamber of a full-scale 9-tonf staged-combustion-cycle engine were studied experimentally. In the combustion experiment, liquid oxygen was used as an oxidizer, and kerosene was used as fuel. The liquid oxygen and kerosene burned in the preburner drive the turbine of the turbopump under the oxidizer-rich hot-gas condition before flowing into the GCSC injector of the combustion chamber. The oxidizer-rich hot gas is mixed with liquid kerosene passed through combustion chamber’s cooling channel at the injector outlet. This mixture has a dimensionless momentum-flux ratio that depends upon the dispensing speed of the two fluids. Combustion tests were performed under varying mixture ratios and combustion pressures for different injector shapes and numbers of injectors, and the characteristic velocities and performance efficiencies of the combustion were compared. It was found that, for 61 gas-centered swirl-coaxial injectors, as the moment flux ratio increased from 9 to 23, the combustion-characteristic velocity increased linearly and the performance efficiency increased from 0.904 to 0.938. In addition, excellent combustion efficiency was observed when the combustion chamber had a large number of injectors at the same momentum-flux ratio.
Highlights
Turbopump-driven liquid-rocket engines can generally be divided into gas-generator-cycle, expander-cycle, and staged-combustion-cycle types
The tests were performed while varying the Combustion tests were conducted s up to 114 s for combustor headsengine-start applying with mixture ratio and combustion pressurefrom for 2different injector shapes
Chamber and the flow rate of kerosene used for regenerative cooling were not optimized, we focused on comparing the relative performances of different combustor heads, rather than comparing absolute combustion-characteristic values
Summary
Turbopump-driven liquid-rocket engines can generally be divided into gas-generator-cycle, expander-cycle, and staged-combustion-cycle types. In a gas-generator system, a portion of the propellant is burned within the gas generator to drive the turbine and is discharged to the outside; here, the propellant supply pressure required for turbopump is relatively low, making it easy to develop and manufacture. Such systems have the disadvantage of a propellant efficiency that is several percentage points lower than that in other systems, as the propellant burned inside the gas generator is dumped outside the turbine [1,2].
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