• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.1-1
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• 2005

Japan Aerospace Exploration Agency(JAXA) has been conducting research and development of the Scramjet engines and their derivative combined cycle engines as hypersonic propulsion system for space access. Its history will be introduced first, and its recent advances, focusing on the engine performance progress, will follow. Finally, future plans for a flight test of scramjet and ground test of combined cycle engine will be introduced. Two types of test facilities for testing those hypersonic engines. namely, the 'Ramjet Engine Test Facility (RJTF)' and the 'High Enthalpy Shock Tunnel (HIEST)' were designed and fabricated during 1988 through 1996. These facilities can test engines under simulated flight Mach numbers up to 8 for the former, whereas beyond 8 for the latter, respectively. Several types of hydrogen-fueled scramjet engines have been designed, fabricated and tested under flight conditions of Mach 4, 6 and 8 in the RJTF since 1996. Initial test results showed that the thrust was insufficient because of occurrence of flow separation caused by combustion in the engines. These difficulty was later eliminated by boundary-layer bleeding and staged fuel injection. Their results were compared with theory to quantify achieved engine performances. The performances with regards to combustion, net thrust are discussed. We have reached the stage where positive net thrust can be attained for all the test coditions. Results of these engine tests will be discussed. We are also intensively attempting the improvement of thrust performance at high speed condition of Mach 8 to 15 in High Enthalpy Shock Tunnel (HIEST). Critical issues for this purposemay be air/fuel mixing enhancement, and temperature control of combustion gas to avoid thermal dissociation. To overcome these issues we developed the Hypermixier engine which applies stream-wise vortices for mixing enhancement, and the M12-engines which optimizes combustor entrance temperature. Moreover, we are going to conduct the flight experiment of the Hypermixer engine by utilizing flight test infrastructure (HyShot) provided by the University of Queensland in fall of 2005 for comparison with the HIEST result. The plan of the flight experiment is also presented.

• Journal of the Korean Society for Precision Engineering
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• v.22 no.7 s.172
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• pp.102-111
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• 2005

In this paper, we address a position control for a parallel stage, which is levitated and driven by electric magnetic force. This consists of a levitating object (called platen) with 4 permanent magnetic linear synchronous motors in parallel. Each motor generates vertical force for suspension against gravity and propulsion force horizontally as well. This stage can generate six degrees of freedom motion by the vertical and horizontal force. A dynamic equation of the stage system is derived based on Newton-Euler method and it's special Jacobian matrix describing a relation between the limited velocity and Cartesian velocity is done. There are proposed two control methods for positioning which are Cartesian space controller and Actuator space controller. The control performance of the Cartesian space controller is better than the Actuator space controller in task space trajectory while the Actuator space controller is simpler than the Cartesian space controller in controller realization.

• The Transactions of The Korean Institute of Electrical Engineers
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• v.62 no.11
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• pp.1640-1646
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• 2013

Conventional power conversion unit that is a major part of the propulsion system has applied GTO thyristor as a switching semiconductor device of main circuit since introduction of the 8200 electric locomotive. But problem that quick maintenance is difficult and its cost is increasing occurs because major components of the power conversion unit are slowly discontinued. To solve these, in this paper, it was analyzed the response characteristic of the propulsion system modeling of the 8200 electric locomotive using IGBT which is applied recently to ensure propulsion control technology. As results of response for a Propulsion system modeling, it show that a power conversion unit is controlled by PLL(Phase-locked loop) and SVPWM(Space Voltage PWM) respectively.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.66-68
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• 2011

Chollian is a geostationary satellite, and its bipropellant propulsion system is mainly composed of one main engine for orbit transfer and fourteen thrusters for on-station operations. The Chollian was launched successfully at Kourou Space Center in French Guiana. After it separated from the launcher, the propulsion system was initialised automatically. Then three times of main engine firing were successfully performed, and the target obit insertion was accomplished. This paper details the major CPS events during LEOP phase for the Chollian satellite.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.168-171
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• 2007

A technical review of current slurry and gelled propellants is presented. In advanced countries, it is confirmed that these propellants have high specific impulse, density, excellent handling, safety characteristics and thrust controllability through research since 1950s. Substantial researches have been pursued to characterize the rheological properties, spray/combustion phenomena and propulsion system design for the gel propellant characteristics. Slurry and gelled propellants are developing actively to applicate both military and space propulsion fields such as tactical missile, air-breathing ramjet, pulse detonation engine, and combined cycle engine of future propulsion mode.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.795-796
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• 2011

During launch process, ground support facilities perform its duty in established processes by communications with launch vehicle. All ground support systems are operated independently or organically. This paper studied algorithm of propellant filling process and method for liquid oxygen filling system in launch operation in Naro space complex.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.753-762
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• 2004

Conceptual studies of a combined-cycle engine have been conducted. Herein, the results are presented. The engine is composed of ejector-jet, ramjet, scramjet and rocket modes, and will be mounted on the Single-Stage-to-Orbit aerospace plane. Propellants are hydrogen and oxygen. Calculated engine thrust performances and cooling requirement of the engine are presented. Pitching moment of the plane with the engine will be balanced even in the vacuum condition. The experimental results of the inlet and the ejector-jet, ramjet and scramjet modes are presented. The effect of the airframe configuration on the engine performance and the thermal environment in the in-side of the plane are also presented. Through the investigations, possibilities of the combined-cycle engine and the aerospace plane are being made clear now.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.892-897
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• 2011

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility(CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The fuel supply system modeling using AMESim was performed based on the results of the detailed design, and the fuel supply characteristics was analyzed in this paper.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.3
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• pp.58-64
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• 2003

Pogo is the instability resulting from the interaction between rocket structure and propulsion system of liquid propellant rocket. The coupling of structure and propulsion system can lead to severe problem in rocket. For the analysis of pogo, a time-invariant linearized mathematical model is developed for a selected flight time. Propulsion system is modeled using element representations for each components. Rocket structure is modeled using FEM. Form the results of modal analysis of structure, the behavior of structure can be represented. System equations for coupling structure and propulsion system are composed. The stability in obtained by the eigen solution of system matrix. The optimization of the design variables such as size, place of accumulator for suppressing pogo instability in carried out. This article of study can be used to determine the degree of stability, and guide the design of pogo suppression system.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.92-97
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• 2012

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility (CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The oxidizer supply system modeling using AMESim was performed based on the results of the detailed design, and the oxidizer supply characteristics was analyzed in this paper.