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Combustion tests were conducted to investigate the high-frequency combustion instability of liquid-propellant rocket engines. An experimental apparatus was designed to examine the rapid and large pressure increase that occurs near the propellant injector, based on the mechanism of the off-design combustion model. The propellants were 2-propanol and gaseous oxygen. The propanol jet created a semi-enclosed base space near the injection port. Pressure in the semi-enclosed space due to combustion was 1.5–3 times higher than the pressure prior to combustion. The duration of the pressure increase ranged from 0.03 to 2 ms. The pressure increase and its duration agreed with the features of the high-frequency combustion instability of liquid-propellant rocket engines. The model temperatures affected the pressure increase and the duration of pressure change.

To clarify the loss mechanism in supersonic turbines, the authors experimentally and numerically studied the flow condition between the rotor blades of a supersonic turbine for liquid rocket engines. The entrance Mach number was 1.94, and the turning angle of the blades was 120 deg. Shock waves were created at the leading edge of the blade. The Mach number and total pressure decreased in the passage between the blades, where the flow condition was restricted by the centrifugal force. Such degradation, which has also been reported in past studies, is an inherent feature of high-speed, large-turning-angle blades. Here, it was found that a convergent–divergent configuration of the passage between the blades suppressed the performance degradation of the supersonic turbine.

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2018-01-19, 資料番号: PA1810077000

形態: 図版あり, Physical characteristics: Original contains illustrations, Accepted: 2016-04-12, 資料番号: PA1710084000

形態: 図版あり, Physical characteristics: Original contains illustrations, Accepted: 2016-06-11, 資料番号: PA1710041000

Conceptual design procedures and design models of HATS are revised and renewed. The results calculated using the revised method are compared with the operating conditions of HATS at the JAXA Kakuda Space Center. The previous physical method of test chamber pressure and that of ejector suction are adopted in the present model. The suction model adopts the inviscid momentum exchange mechanism. The deceleration process in the supersonic diffuser is revised using the pseudo-shock model. Physical and thermodynamic models are constructed for condensation and steam saturated flow conditions. The results calculated are in reasonable agreement with the measured values (e.g., pressure of secondary flow at the ejector section and pressure changes during engine shut down). The effect of ejector steam condensation on the operating conditions of HATS is quantitatively presented., 形態: 図版あり, Physical characteristics: Original contains illustrations, 資料番号: PA1610066000

形態: 図版あり, Physical characteristics: Original contains illustrations, Accepted: 2016-07-20, 資料番号: PA1710080000

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

The 43rt JSASS Annual Meeting (April 12-13, 2012. Sanjo Conference Hall, The University of Tokyo), Tokyo, Japan, 資料番号: PA1210039000

2010 Asian Joint Conference on Propulsion and Power (March 4-6, 2010.), Miyazaki, Japan, 資料番号: AC1000137000, レポート番号: AJCPP2010-003

7th International Symposium on Special Topics in Chemical Propulsion(7-ISICP) (September 17-21, 2007.), Kyoto, Japan, 資料番号: AC1000142000

著者人数: 4名, 資料番号: AC1000038000

資料番号: ARDS06125000

著者人数: 3名, 資料番号: AC1000036000

DOI: 10.2514/1.19969 Multidisciplinary designoptimization isanimportant approachfor the conceptualdesignof space planesbecause these planes are characterized by disciplines that interact with one another. A multidisciplinary design optimization problemof asingle-stage-to-orbit space plane isformulated andsolved inthis study.Themodeling andoptimization ofrigidbodycharacteristics,suchastrimandstability,arefocusedonbecauseasingle-stage-to-orbitspaceplanehas a tendency for considerable shift of both the aerodynamic center and the center of gravity. Moreover, the design of the air-breathing engines are integrated with the airframe and its effect on the rigid body characteristics are also modeled in the framework of multidisciplinary design optimization. Using the all-at-once-based multidisciplinary design optimization approach, which incorporates sparse nonlinear programming and metamodeling, the design of the vehicle and its flight trajectory are successfully optimized. Finally, the characteristics of the optimal solution are investigated, especially the relationships among the airframe-engine integration, rigid body characteristics, and payload transportation capability.

著者人数: 5名, 資料番号: AC1000034000

著者人数: 4名, 資料番号: AC1000033000

著者人数: 3名, 資料番号: AC1000035000

著者人数: 5名, 資料番号: AC1000032000

著者人数: 5名, 資料番号: AC1000031000

24th Congress of International Council of the Aeronautical Sciences (August 29-September 3, 2004.), Yokohama, Kanagawa, Japan, 資料番号: AC1000135000, レポート番号: ICAS 2004-1.11.1

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

著者人数: 4名, 資料番号: AC1000029000

著者人数: 5名, 資料番号: AC1000030000

著者人数: 2名, 資料番号: AC1000027000

著者人数: 5名, 資料番号: AC1000026000

著者人数: 5名, 資料番号: AC1000025000

AIAA International Space Planes and Hypersonic Systems and Technologies Conference, 9th, Norfolk, VA, Nov. 1-5, 1999, 資料番号: AC1000062000, レポート番号: AIAA-1999-4871

著者人数: 4名, 資料番号: AC1000023000

著者人数: 5名, 資料番号: AC1000022000

Computations of internal viscous flowfields of scramjet models were conducted at inflow Mach numbers of 3.4 and 5.45. A hybrid unstructured grid method was used to compute the scramjet models with and without a short strut. The numerical method to solve the Navier-Stokes equations on the hybrid unstructured grid was developed using a finite volume cell vertex scheme and the lower-upper symmetric Gauss-Seidel (LU-SGS) implicit time integration algorithm. The computational results revealed that a thick subsonic region did not exist in the combustor near the top wall at Mach number 5.45. It was a desirable feature to avoid unstarting the engine. With the use of a strut, the relatively low velocity regions increased and the downward flow toward the cowl behind the step became strong. An overconcentration of fuel toward the top wall during weak combustion was observed in the experiment. The reason for this was that the airflow near the injector was turned to the top wall due to the small influence of combustion in the experiment.

著者人数: 5名, 資料番号: AC1000021000

著者人数: 5名, 資料番号: AC1000037000

著者人数: 6名, 資料番号: AC1000020000

著者人数: 5名, 資料番号: AC1000019000

ASME, SAE, and ASEE, Joint Propulsion Conference and Exhibit, 31st, San Diego, CA, July 10-12, 1995, 資料番号: AC1000101000, レポート番号: AIAA-1995-3141

著者人数: 4名, 資料番号: AC1000028000

A combined-cycle engine has two or more operating modes with a single flowpath. The motivation behind a combined-cycle engine is to extend the engine's operating envelope and attain higher average performance. Several combined-cycle engines and a combined engine system are presented. The rocket-based combined-cycle (RBCC) engine is a combination of a rocket engine and a ramjet engine. RBCC has wide operating range and structural simplicity, since there are no rotating parts. The drawback is a relatively lower specific impulse in the subsonic and transonic range. RBCC is suitable for the single-stage-to-orbit (SSTO) vehicle due to its operability in low ambient pressure conditions and in space. The liquid air cycle engine (LACE) liquefies air using cryogenic propellants to reduce the onboard oxygen requirement. The advantages of LACE are reduced engine weight and wide operating conditions. However, owing to the low liquefaction efficiency of air, LACE suffers from low specific impulse and high fuel consumption. The turbo-ramjet (TRE) and the air-turbo-rocket (ATR) engines employ turbocompressors to compress the airflow like in turbojet engines. Despite the advantage of increased specific impulse, both engines have low thrust to weight ratios. The dual-mode scramjet engine like the RBCC has no moving parts. The combustion is subsonic in the ramjet mode and supersonic in the scramjet mode. The combination engine system, which mounts several engines on a vehicle, extends the operating range of the vehicle by operating appropriate engines according to the flight conditions. Keywords: rocket-based combined cycle; liquid air cycle; turbo-ramjet; air-turbo-rocket; dual-mode scramjet