A semi-analytical calculation of combustion gas parameters and the characteristic velocity performance, based on the experimental pressure distribution of LOX/GH2 rocket combustion chambers with constant cross-sectional areas and coaxial type injectors. The thrust level was 300 kgf with a nominal chamber pressure of 20 atm. The oxidant-fuel mixture ratios ranged from approximately 2 to 10. The combustion chamber geometry varied in lengths from 11 to 26 cm and contraction ratios of 2.03, 3.41 and 4.46, with a constant nozzle throat area. The injector has twelve coaxial elements with three replaceable faceplates having different hydrogen port areas.
An Experimental-Analytical Calculation of Combustion Performance of LOX/GH2 Rocket with Coaxial Injectors
1976
22 pages
Report
No indication
English
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