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)
Kakuda Branch Office, National Aerospace Laboratory(NAL)
Kakuda Branch Office, National Aerospace Laboratory(NAL)
Kakuda Branch Office, National Aerospace Laboratory(NAL)
National Space Development Agency of Japan(NASDA)
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-679
An experimental investigation of the combustion and the heat transfer characteristics of a liquid oxygen-liquid hydrogen rocket combustor was conducted. The liquid hydrogen cooled chamber with a sl otted wall liner made of OFHC copper for coolant passage was designed for use at a thrust level of 300 kgf with a nominal combustion chamber pressure of 35 atm. In order to obtain the liquid hydrogen cooling characteristics at near and super-critical conditions, the combustion experiments were performed using both the independent cooling method and the regenerative cooling methods. And to obtain data on the combustion performance of coaxial type injectors, the number of injector elements, the oxidant-fuel ratios and the hydrogen injection temperature were varied over a wide range. The following results were noted. The regenerative cooling combustion tests at the design thrust and the combustion pressure were successfully performed. The slotted wall chamber was revealed to have a sufficiently high heat transfer rate and enough cooling margin even in this small thrust level combustor. In the independent cooling experiments, a new thermal design equation was obtained, which should be recommended as the design equation for regeneratively cooled rocket engines. As a result, many design equations previously obtained from Joule heated tube experiments proved to be not suitable for the design of slotted wall combustors. Observed C* efficiencies were correlated as a function of the hydrogen-oxygen injection velocity ratio and were found to be the same as the results obtained by the previous studies on the liquid oxygen-gaseous hydrogen rocket combustion performance.
小型溝構造液水冷却燃焼器の研究(1)(液水独立冷却および再生冷却燃焼実験)
naltr00679.pdf
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