Kakuda Branch Office, National Aerospace Laboratory(NAL)
Kakuda Branch Office, National Aerospace Laboratory(NAL)
Kakuda Branch Office, National Aerospace Laboratory(NAL)
Kakuda Branch Office, National Aerospace Laboratory(NAL)
出版者
航空宇宙技術研究所
出版者（英）
National Aerospace Laboratory(NAL)
雑誌名
航空宇宙技術研究所報告
雑誌名(英)
Technical Report of National Aerospace Laboratory TR-473
巻
473
ページ
26
発行年
1976-09
抄録（英）
An experimental investigation of combustion performance was conducted in a LOX/GH2 rocket combustor with coaxial type injectors at a thrust level of 300 kgf with a nominal chamber pressure of 20 atm. The oxidant-fuel ratios ranged from approximately 2 to 10. The combustion chambers were of the segmented, heat-sink type. The chamber geometry varied in lengths from 11 to 26 cm and contraction rations of 2.03, 3.41 and 4.46, with a constant nozzle throat area. The injector had twelve coaxial elements with no recess and three replaceable faceplates having different hydrogen port areas. The attendant hydrogen-oxygen injection velocity ratio ranged from approximately 10 to 50. Observed C* efficiencies were correlated as a function of an equivalent chamber length, Le, in which the injection velocity ratio, VR, and the chamber length, L, are expressed as Le=L+αVR where α was considered nearly constant throughout the experiment. The correlation takes the form of ηc*=1-A exp(-BLe) where the constants A and B depend on the chamber contraction ratio. The variations of all experimental data from the relation were within approximately ± 1.8 percent.
液体酸素・ガス水素ロケットの燃焼性能(Ⅰ)
naltr00473.pdf
(4.83MB)
