Based on the author's recent study on detailed modeling of the atomization characteristics of a liquid/gas phase coaxial injector, an improved computation model for predicting the combustion performance of a LOX/hydrogen engine is proposed. The features of this model are that it allows calculation of the local rate of atomization of the LOX jet and deals with droplet size distribution of the LOX spray. Additionally the model applies the burning rate constant of LOX/hydrogen combustion derived by the author's former experiment. Using this model, evaluations of the design criteria for the LE-5 engine, which was equipped on the H-2 launcher, and its derivative engines, which have been used with the H-2A launcher, were conducted. Furthermore, optimization of the derivative engine's injector design to improve combustion stability and combustion performance was discussed.

The ability to efficiently develop ram and scramjet engines can be influenced by ground facilities. The National Aerospace Laboratory of Japan completed a ramjet engine test facility (RJTF) in 1994. It can duplicate engine test conditions in the range of flight Mach numbers from 4 to 8. The facility can supply non-vitiated air for M4 and M6 to identify the effects of contamination in vitiated air, providing a basis for evaluating engine performance in the M8 flight condition. In this paper, the unique features and operating characteristics of the RJTF are outlined. The quality of air stream obtained during facility calibration and the facility-engine interaction are described. Finally, we review tests of a H2-fueled scramjet and Ctfy-fueled ramjet engines currently underway.

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An optimum method for design of a liquid hydrogen regenerative cooling combustor for the LOX/hydrogen engine was constructed using the author's previous empirical correlation of C Ã efficiency and calculation model for combustion characteristics, and the present calculation model for the heat load characteristics for LOX/hydrogen combustion. Using this method, the atomization characteristics of the injected LOX jet, the combustion performance including combustion stability, and the heat load on the combustor were evaluated for LOX/hydrogen upper-stage engines such as the LE-5, RL-10 and HM-7. This method was then applied to the LE-5B engine, which is the derivative engine of the LE-5 and has been used as the second stage of the H-2A launcher, to improve combustion stability and to optimize configuration of the injector and combustor. A reduction of about 30% in chamber length of it with sufficient combustion performance was achieved by such optimization.

A comprehensive design method for a LOX/Liquid-Methane (L-CH 4 ) rocket engine combustor with a coaxial injector and the preliminary design of the regenerative cooling combustor with 100-kN thrust in vacuum at a combustion pressure of a 3.43 MPa are presented. Reasonable dimensions for the combustor that satisfy the targeted C Ã efficiency of more than 98% and combustion stability are obtained.