• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.2
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• pp.68-73
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• 2011

For the launch preparation of KSLV-I, gaseous nitrogen with various level of pressure and cryogenic liquid nitrogen are required. Nitrogen Supply System on launch complex has been developed to perform the production of high pressure gaseous nitrogen, the production of gaseous nitrogen with temperature of 273 ${\pm}$ 2K for protection purge of launch vehicle after loading of propellant and the supply of cryogenic liquid nitrogen for cooling of fuel (kerosene) and oxidizer (liquid oxygen). The operational instability of vaporizer mainly caused by its heat transfer characteristics which sensitively depends on the atmospheric conditions was removed by introducing parallel installation of two vaporizer and their switching operation. The developed Nitrogen Supply System carried out its function successfully in preparation of KSLV-I flight tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.752-757
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• 2010

For the launch preparation of KSLV-I, gaseous nitrogen with various level of pressure and cryogenic liquid nitrogen are required. Nitrogen Supply System on launch complex has been developed to perform the production of high pressure gaseous nitrogen, the production of gaseous nitrogen with temperature of $273{\pm}2K$ for protection purge of launch vehicle after loading of propellant and the supply of cryogenic liquid nitrogen for cooling of fuel (kerosene) and oxidizer (liquid oxygen). The operational instability of vaporizer mainly caused by its heat transfer characteristics which sensitively depends on the atmospheric conditions was removed by introducing parallel installation of two vaporizer and their switching operation. The developed Nitrogen Supply System carried out its function successfully in preparation of KSLV-I flight tests.

• Proceedings of the Korea Institute of Fire Science and Engineering Conference
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• 2013.04a
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• pp.163-163
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• 2013

최근의 수많은 산업 현장에서 취급하고 있는 각종 화학물질은 잠재적 위험성이 크므로 보관, 수송 및 취급할 때 특별한 주의가 필요하다. 공정 설계 시 정확하지 않은 폭발한계를 사용함으로서 사고가 유발되는 경우가 많다. 따라서 사업장에서 사용되고 있는 화학물질의 화재 및 폭발 특성치인 인화점, 폭발한계 등을 정확히 파악하는 것은 중요하다. 인화점은 하부인화점과 상부인화점으로 나누고 있고 있으며, 인화점은 가연성 액체의 화재 위험성을 나타내는 지표로써, 가연성액체의 액면 가까이서 인화할 때 필요한 증기를 발산하는 액체의 최저온도 또는 점화원 존재시 인화가 일어날 수 있는 최저온도, 그리고 가연성증기의 포화증기압이 공기와 혼합기체의 폭발한계 하한농도와 같게 되는 온도로 정의한다. 폭발한계는 발화원이 존재할 때 가연성가스와 공기가 혼합하여 일정 농도범위 내에서만 연소가 이루어지는 혼합범위를 말한다. 본 연구에서는 실제 공정에서 사용되고 있는 n-Hexadecane의 인화점을 측정하여 이를 기존 문헌값과 비교 하였고, 측정된 인화점을 이용하여 폭발한계를 예측하였다. 예측된 폭발한계를 여러 문헌에 제시된 자료과 비교하여 공정안전에 타당한 자료를 제시하였다. 본 연구는 n-Hexadecane을 취급하는 공정에서 안전 확보의 중요한 지침 마련과 MSDS D/B의 최신화에 유용한 정보를 제공하는데 목적이 있다.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.2
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• pp.25-33
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• 2012

A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Pressure of the main combustion chamber, nozzle expansion ratio and O/F ratio have been selected as design variables. The target engine has the open gas generator cycle using the LO2/RP-1 propellant. The gas properties of the combustion chamber have been obtained from CEA2 and the mass has been estimated using reference data. The objective function has been set as multi-objective function with the specific impulse and thrust to weight ratio using the weight method. The result shows about 4% improvement of the specific impulse and 23% increase of the thrust to weight ratio. The Pareto frontier line has been also obtained for various thrust requirements.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.85-92
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• 2006

A design program has been developed to predict film cooling performance of a liquid rocket engine. A thermal protecting effect of low mixture ratio gas layer has been analysed by CFD. A one-dimensional film cooling model based on the CFD results has been implemented to the previously developed design program of regenerative cooling. Satisfactory agreement has been achieved by comparing the predicted maximum heat flux at the throat of a subscale chamber and the average measured value, and the predicted nozzle average heat flux and the measured value for a full scale chamber with film cooling. It is ascertained that the film cooling is effective to reduce the throat heat flux in rocket engine chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.877-880
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• 2011

A coaxial swirl/shear injector using GCH4/LOx as propellants was degisned and manufactured. Flow analysis by Fluent was performed to decide the number of orifice and the rear shapes of inlet orifice etc. Flow rate of the injector was measured according to differential pressure and uniformity of injector's spray pattern was confirmed by a patternator. The results showed that the difference of flow rate was around 10% and the spray angle of oxidizer was $66^{\circ}$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.127-134
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• 2011

A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Pressure of the main combustion chamber, nozzle expansion ratio and O/F ratio have been selected as design variables. The target engine has the open gas generator cycle using the LO2/RP-1 propellant. The gas properties of the combustion chamber have been obtained from CEA2 and the mass has been estimated using reference data. The objective function has been set as multi-objective function with the specific impulse and thrust to weight ratio using the weight method. The result shows about 4% improvement of the specific impulse and 23% increase of the thrust to weight ratio. The Pareto frontier line has been also obtained for various thrust requirements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.658-661
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• 2011

A coaxial swirl/shear injector using GCH4/LOx as propellants was degisned and manufactured. Flow analysis by Fluent was performed to decide the number of orifice and the rear shapes of inlet orifice etc. Flow rate of the injector was measured according to differential pressure and uniformity of injector's spray pattern was confirmed by a patternator. The results showed that the difference of flow rate was around 10% and the spray angle of oxidizer was $66^{\circ}$.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.1
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• pp.43-49
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• 2016

An experimental study was performed to investigate the combustion characteristics of the liquid hydrocarbon fuel atomized by an ultrasonic oscillator. Configuration of the flame was caught by the DSLR camera, and images were analyzed in detail through a post-processing. Temperature of the flame zone was measured using thermocouple. It is resulted that the flame area is proportional to the mass flow rate of carrier gas as well as the voltage applied to the ultrasonic oscillator. Temperature of the flame zone is measured and analyzed according to variation in the operating condition of the burner, too.