A staged combustion rocket combustor using LOX/methane propellants was experimentally evaluated using a small scale thrust chamber. The fuel rich preburner and the main chamber were tested individually, and preburner/main chamber assembly tests were conducted at main combustion chamber pressures ranging from 7 to 9.6 MPa, oxidant to fuel ratios of 3.1 to 3.7, preburner combustion gas temperatures of 760 to 1070 K, and thrusts ranging from 4.6 to 6.4 kN. Combustion performance of the preburner and main chamber, heat flux distribution in the main chamber and carbon deposition on the chamber walls including injector faces were evaluated for different types of main injectors and preburner injectors. An empirical correlation was deduced for the combustion efficiency of the main chamber. Significant carbon deposition was not observed anywhere in the assembly, and significant reduction of heat flux caused by carbon deposition on the chamber walls was also not detected.
LOX/methane staged combustion rocket combustor investigation
Untersuchung ueber Raketenbrennkammern fuer die stufenweise Verbrennung von Fluessigsauerstoff/Methan-Gemischen
AIAA-Papers ; Jul ; 1-6
1987
6 Seiten, 13 Bilder, 3 Tabellen, 11 Quellen
Aufsatz (Konferenz)
Englisch
Preliminary Study for Staged Combustion Hybrid Rocket
British Library Conference Proceedings | 2000
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