• Proceedings of the KIEE Conference
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• 2006.10c
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• pp.407-409
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• 2006

우주발사체I 상단 시스템은 비행 중에 킥 모터 점화, 위성분리와 같은 HBW 점화기 기폭이 순차적으로 발생 한다. HBW 점화기의 기폭시에는 일반적으로 전도성 구조물을 통한 단락전류가 일시적으로 발생한다. 이러한 단락전류는 구동 전원 측으로 최대 전류 값 및 루프 면적에 비례한 일시적인 자기장을 형성시키고 near 필드(${\lambda}/2{\pi}$) 내의 인접하게 위치한 하니스에 자계 결합을 통한 역기전력 발생의 원인이 될 수 있다. 이러한 인접 하니스에 자계 노이즈 결합은 여러 자료를 통해 우주시스템 환경에서 일시적인 데이터 손실의 원인이 되는 것으로 분석되고 있다. 본 논문은 우주발사체의 HBW 점화기 기폭시 발생되는 전도성 구조물 단락전류 귀환현상으로 인한 자계노이즈 분석 및 감쇄방안에 대해 논의하고자 한다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.111-114
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• 2009

A catalyst ignitor of small thrust engine using hydrogen peroxide and kerosene was designed and fabricated, which confirmed mass flow rate for design pressure through the water cold-flow test in this study. In order to investigate ignition performance, it was changed that mixture ratio for kerosene mass flow rate in a position which heat of hydrogen peroxide decomposition comes to a steady state. And we confirmed stable ignition property in a wide range of mixture ratio.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.225-228
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• 2007

Micro solid propellant thruster is the most feasible for development with current MEMS. Basic components of micro solid propellant thruster are diverging nozzle, micro igniter, combustion chamber, and solid propellant. Micro nozzles and micro chambers were fabricated using photosensitive glass by anisotropic wet etching technique. Micro Pt heaters on glass membrane which ignited solid propellant were developed. Components of thruster were integrated. Successful ignition was observed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.353-358
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• 2008

Inha Rocket Research Institute has made the igniter that is combination of black powder and PVA polymer for ignition small rocket. But recent igniter is not satisfy because of the performance of igniter is not identified. So, we confirmed requirement of igniter by comparing of ratio of black powder and PVA through experimental method. Especially we studied with ignition temperature for propellant and stable combustion pressure that is requirements of propellant. We can know the tendency of combustion properties by ratio of oxidizer and combustion catalyst through changing of temperature and pressure of exhaust gas of igniter.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.433-436
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• 2005

The combustion behavior of initiators and can-type ignitor, acting on vacuum pressure, was invested. It was found that if the ignited propellants of a can-type igniter were exposed at certain level of low pressure, it would be burnt improperly or even be extinguished. When we design the can-type igniters for some rocket motor, it is most important that the pressure built-up by the gases which were already burnt inside of the igniter case keep as high as possible. For this purpose, it is recommended that the ignitor case have some strength to produce as much gases as possible before breaking out and the free volume be kept as small as possible.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.2
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• pp.96-107
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• 2018

Flow analysis using computational fluid dynamics was conducted to investigate the effect of the igniter flame caused by the gap between the igniter and the propellant grain in a solid rocket motor. Two propellant grain types were assumed; namely cylinder type (1 mm, 3 mm, and 5 mm gap) and the slot type. The slot type had two igniter hole locations. One was located at the small gap of the propellant grain, and the other one was located at the large gap. In the case of the cylinder type, the pressure in the igniter zone was higher with a thinner gap. Additionally, in the case of the cylinder type, the pressure difference between the igniter installed zone and the free volume was also higher as the gap became lower. The cylinder types were affected by the gap distance, but the slot types were not. Moreover, the results of the slot types were similar to the 5-mm gap case of the cylinder type.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.6
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• pp.56-62
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• 2011

An experimental study on design of a catalytic ignitor was performed to use an ignition source for a small bi-propellant liquid rocket engine which use hydrogen peroxide and kerosene as propellants. In the catalytic ignitor, hot gas of hydrogen peroxide which was decomposed by a catalyst induced autoignition of kerosene. Mass flow rate and O/F ratio for the ignitor were calculated by CEA code. A combustion chamber which had a quartz window and thermocouples was manufactured to determine whether the ignition is successful. Ignition performance was investigated according to exit area of fixed rings and mixture ratio. Results showed that reliable ignition performance was achieved at non-choking exit area of fixed ring and O/F ratio of 6~8.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.814-819
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• 2017

The objective of this paper is to observe effects of an ignition position and time on the ignition transition. A gaseous oxygen and liquid kerosene are used for propellants with the shear-coaxial injector. In order to study the ignition delay time and combustion instability intensity, the pressure transducer was used. The ignition position was changed with the injector spacer. Sequences except the igniter operation time were fixed to compare the ignition time only. Initial pressure peak and ignition delay time increased as the ignition time was delayed. Also, the unstable flame development zone was detected as the igniter was away from the injector.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.202-205
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• 2010

Auto ignition of hybrid rocket using $N_2O$ catalytic decomposition was studied in the present study. The hybrid rocket consists of catalytic igniter, solid fuel, combustor, and nozzle. The Ru/$Al_2O_3$ catalyst for $N_2O$ decomposition was synthesized by an impregnation method, and $N_2O$ conversion as reaction temperatures was measured. The temperature change of the catalytic ignitor was measured at the operating condition, and the possibility for the auto ignition of hybrid rocket was validated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.730-733
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• 2017

An ignitor for the initiation of burning propellants becomes one of the most critical components for the operation of liquid rocket engines. The important phenomenon of the igniter is thermoacoustic resonance. This paper mainly includes experimental results on thermoacoustic resonance phenomenon from distance between nozzle and resonance tube. The ultimate goal is to develop an ignitor capable of multiple ignitions based on research results about the usage of the thermoacoustic resonance phenomenon.