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形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-07-09, 資料番号: PA2110004000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2018-01-19, 資料番号: PA1810076000

To reduce the cost of space transportation, air-breathing engines are considered to be candidates for propulsion. However, to cover a wide range of flight speeds, the propulsion system has to operate in various modes to be efficient under incoming atmospheric-air conditions. The Japan Aerospace Exploration Agency is proposing a rocket-based combined cycle engine for operation under various condition, an ejector-jet mode being adopted for the low-speed regime. The suction performance ejector-jets has long been studied experimentally and numerically at JAXA, and little success has been achieved in explaining the deterioration of suction performance with high-temperature gas or light gas such as helium. In the present study, based on former models, a simple one-dimensional model was introduced incorporating the mixing effects of the primary flow (rocket flow) and secondary flow (induced air flow). The results were compared using several experimental and numerical data to check the plausibility of the model. It was found that if greater mixing occurs, suction performance is degraded, explaining the actual phenomena of the experiments., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1610077000

資料番号: ARDS06125000

2005 Asian Joint Conference on Propulsion and Power (January 27-29, 2005.), Kitakyushu, Fukuoka, Japan, 資料番号: AC1000138000, レポート番号: AJCPP2005-22132

We conducted a sensitivity analysis of engine performance on measurements of gas sampling, pitot pressure and static pressure. It indicates that the heat loss in engine and the insufficient quenching of reactions in the gas sampling probes results in the mismatch in the mass flux through the engines. An analytical study of the reactive flow in probes showed that the introduction of a small heat of reaction with radical recombination to the energy equation yielded a saddle-type singularity in a phase plane for temperature and radical. The separatrix running from the initial point to the saddle point gave the criterion dividing the thermal runaway and the chemical quenching. A one-dimensional reaction code with a full-kinetics confirmed this analytical results.

Engine system analysis upon a Rocket-Based Combined Cycle (RBCC) engine for the booster stage of a TSTO launch vehicle was conducted, to evaluate the effects of hydrocarbon fuel type upon the performance of the engine and the whole launch vehicle. A higher thrust and cooling performance with methane was shown, while some technical challenges were also noted. Combination of the scramjet flow pass and the rocket engines, i.e., rocket engine being embedded into the scramjet flow pass or separately mounted to the airframe, was varied to show that the separated configuration mitigated cooling problem, and resulted in a better performance for mid-speed regime., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1610081000

A series of experiments was conducted for studying scramjet nozzles under Mach 8 flight conditions using a high temperature gas flow produced by combustion of monomethyl-hydrazine (MMH) and nitrogen tetroxide (NTO). The stagnation temperature was designed to be 3170K, and the stagnation pressure to be 1MPa. Scramjet nozzles with expansion ratios of 3 (EN3) and 5 (EN5) were examined. Delivered specific impulse (Isp) was found to be 17.3s (EN3) and 26.1s (EN5). Performances of nozzles were predicted using numerical codes, and energy release loss, kinetic loss and two-dimensional loss were identified. The two-dimensional loss was predicted to be 16.5% in the EN3 nozzle and to decrease to 7.0% in the EN5 nozzle. The kinetic loss due to chemical freezing was found to be about 4%. A friction loss evaluated from thrusts suggested laminar flow in the scramjet nozzle with the MMH/NTO experiments. Measured heat flux on nozzles also supported the laminar boundary layer. Experiments using a cold gas flow with the nozzles were also conducted to compare with the hot flow.

43rd AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, Jan. 10-13, 2005, 資料番号: AC1000080000, レポート番号: AIAA-2005-1045

39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Huntsville, Alabama, July 20-23, 2003, 資料番号: AC1000096000, レポート番号: AIAA-2003-4737

AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 34th, Cleveland, OH, July 13-15, 1998, 資料番号: AC1000105000, レポート番号: AIAA-1998-3134

27th International Symposium on Space Technology and Science, Note: Selected papers from the 27th International Symposium on Space Technology and Science, 資料番号: PA0910006000

資料番号: PA0810076000