• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.13-19
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• 2003

A supersonic ground test facility to develop Ramjet and SCRamjet(Supersonic Combustion Ramjet) engine should be able to simulate high altitude and high Mach number conditions including air total pressure, oxygen level and specific heat ratio at the combustion chamber entrance. The test facility also should simulate the effect of oblique shock wave caused by the flight vehicle. The test facility developed in this study is supersonic free-jet blowdown type, which consists of high pressure air supply source(maximum pressure=32MPa), air heater(vitiation type), supersonic diffuser, ejector, and test chamber(nozzle exit dimension=200mm$\times$200mm).

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.4
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• pp.53-62
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• 2003

A supersonic ground test facility to develop Ramjet and SCRamjet(Supersonic Combustion Ramjet) engine should be able to simulate high altitude and high Mach number conditions including air total pressure, oxygen level and specific heat ratio at the combustion chamber entrance. The test facility also should simulate the effect of oblique shock wave caused by the flight vehicle. The test facility developed in this study is supersonic free-jet blow down type, which consists of high pressure air supply source(maximum pressure=32MPa), air heater(vitiation type), supersonic diffuser, ejector, and test chamber(nozzle exit dimension=200mm${\times}$200mm).

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.12
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• pp.1195-1201
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• 2010

Performance tests of a supersonic exhaust diffuser were conducted by using a liquid rocket engine for simulating high-altitude environment. The experimental setup consisted of a combustion chamber, a vacuum chamber and a diffuser. The combustion tests for simulating high-altitude environment were carried out at three cases by chamber pressure variation(26, 29, 32barg). The test results showed that the diffuser was started at all case and vacuum chamber pressures were approximately 140torr. The starting pressure using combustion gas was similar with that of cold gas, but the vacuum chamber pressure was relatively high because of high temperature in the vacuum chamber. The results of this test can be used as an essential database for the design of real-scale high-altitude simulation test facility in the future.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.379-380
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• 2009

한국항공우주연구원 추진기관팀은 1999년 10월에 3,000 lbf 급 고공환경 엔진시험 설비를 갖추고 소형 가스터빈 엔진의 고공환경 성능시험에 이를 활용하고 있다. 하지만 새롭게 2008년부터 고공환경 성능시험을 진행하고 있는 엔진은 1,000 lbf 미만의 초소형 엔진으로써 기존 추력측정 시스템을 이용하여서는 정확한 추력의 측정을 보장할 수 없다. 본 논문에서는 초소형 엔진의 고공환경 성능시험 수행을 위한 추력대의 구축 과정을 다루고 있다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.75-81
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• 2006

The propulsion system of KSLV-I second stage is engine with high expansion ratio and its starting altitude is high. To verify the performance of engine before the launch in the ground, high altitude test facility to simulate its operating condition is necessary. This material is about the concept design of high altitude simulation test facility for second stage engine. And it will be the basis for the construction of test facility and the test of engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.309-312
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• 2005

High altitude ignition test was performed to understand high altitude ignition characteristics of slinger combustor. To verify ignition limits, test was carried out with variation of altitude and fuel nozzle rotational speed using AETF(Altitude Engine Test Facility) in KARI(Korea Aerospace Research Institute). From the result, the effect of major factors which affect on ignition characteristics was observed. The reduction of ignition limit with increasing altitude and expansion of ignition limit with increasing rotational speed of fuel nozzle was verified. Also minimum rotational speed of fuel nozzle at high altitude must be greater than that of seal level condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.185-188
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• 2009

The rocket motor of the space launch vehicle offers thrust for satellite to enter into the orbit. Characters of the solid propellant for rocket motors are affected by the space conditions such as vacuum and space radiation. The solid propellant used for such a purpose should not undergo physical, internal ballistic and energetic changes when exposed to vacuum and space radiation. This study describes the development of the solid propellant composition for the rocket motor of the space launch vehicle. Also, experimental study was conducted on supersonic diffuser in order to verify the performance of the solid propellant composition which was applied to standard motor on the ground in the vacuum condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.627-630
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• 2009

Ejector system can induce the secondary flow or affect the secondary chamber pressure by both shear stress and pressure drop which are generated in the primary jet boundary. Ejectors are widely used in a range of applications such as a turbine-based combined-cycle propulsion system and a high altitude test facility for rocket engine, pressure recovery system, desalination plant and ejector ramjet etc. The primary interest of this study is to set up an configuration and operating conditions for an ejector in the condition of sonic and subsonic. Experimental and theoretical investigation on the sonic and subsonic ejectors with a converging-diverging diffuser was carried out. Numerical simulation was adopted for an optimal geometry design and satisfying the required performance. Also, some ejectors with a various of nozzle throat and mixing chamber diameter were manufactured precisely and tested for the comparison with the calculation results.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.41 no.9
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• pp.563-570
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• 2017

Experimental and numerical studies were performed to investigate the performance and internal flow characteristics of a supersonic second throat exhaust diffuser (STED) with back pressure ($P_a$). An ejector system was used to vary the back pressure ($P_a$) conditions. The operating gas for the STED and the ejector was high pressure nitrogen at room temperature. When the back pressure ($P_a$) at a constant nozzle inlet pressure $P_0$) decreases, the pressure recovery location moves downstream. If the pressure ratio $P_0/P_a$) is the same, even if the nozzle inlet pressures $P_0$) are different, the diffuser's internal flow pattern and starting pressure ratio ($(P_0/P_a)_{st}$) are almost the same.