• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.2
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• pp.46-53
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• 2010

The purpose of this paper is to study combustion characteristics with the different phase of oxidizer in hybrid rocket combustion. HDPE(High Density Polyethylene) as fuel and $GN_2O$(Gas $N_2O$), $LN_2O$(Liquid $N_2O$) as oxidizer were used to perform the experiments. An investigation was performed for a change of the regression rate, pressure of combustion chamber and combustion efficiency according to the variation of oxidizer phase. In case of using $LN_2O$ as oxidizer, the regression rate is not significantly different from using $GN_2O$ as oxidizer. It is considered that combustion energy is much larger than latent heat energy which was used in the evaporation of liquid oxidizer. However propulsion performance efficiency for $LN_2O$ showed lower value than for $GN_2O$. By increasing the flow rate of liquid oxidizer, heat transfer needed for vaporization of liquid oxidizer was increased, which resulted in the growth of combustion instability.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.8
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• pp.747-752
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• 2007

Domestic satellite development programme is generally classified into two categories: COMS as GEO satellite and KOMPSAT as LEO one. Each satellite has the on-board propulsion system fulfilling its own mission requirements. The COMS propulsion system provides the thrust and torque required for the insertion into GEO, attitude and orbit control/adjustment of spacecraft. It is the well-known Chemical Propulsion System(CPS) using bipropellants. On the other hand, the monopropellant propulsion system is employed in KOMPSAT, and its main role is on-station attitude control excluding the orbit transfer function. In this study, these two representative propulsion systems are compared and analysed as well, in terms of essential differences and important characteristics.

• Korean Chemical Engineering Research
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• v.54 no.3
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• pp.360-366
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• 2016

The Propulsion Test Facilities for the development of Korea Space Launch Vehicle-II are being built, some test facilities are completed and various combustion tests are running. The Propulsion Test Facilities consists test-stand, which carries out tests for engine development model, and various sub-systems and vessels containing LOX and Jet A-1 as propellant. There are always risks of fire and explosion at the test-stand since engine development model is conducted at test-stand with real combustion test with very high pressure, mixed propellant and high energy. In this paper, in order to establish the consequence analysis and risk reduction measures in the Propulsion Test Facilities, followings are considered. 1) a propellant leak accident scenario is assumed in test-stand. 2) TNT equivalent model equation based on blast wave of the explosion was used to analyze blast overpressure and impacts. Also, technical, systematic and managemental measure is described to ensure risk reduction for propulsion test facility.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.2
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• pp.74-80
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• 2008

The feasibility study into the development of turbine spinners, which start up the turbo-pump, has been carried out and the design requirements and parameters ranges have been presented. Turbine spinners use the solid propellant as such composite propellant based AN compound with high energy plasticizers, coolants, and phase stabilizer which relieves a sensible volume change due to the phase transformation of AN near room temperature. Propellants which have a homing rate of $0.2{\sim}0.3\;mm/s$ and pressure exponent ranged from 0.3 to 0.6, showed stable burn-out in the standard motor tests. Both the magnitude of ignition energy and its thermal transfer mechanism have been proved to have a tangible effect on the ignition of the pyre starter, and the results of this study showed that a flame temperature of 1400K would be quite adequate to get a stable ignition for the AN composite propellant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.159-162
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• 2008

In this study, various cycles of liquid rocket engines were surveyed and specifically gas generator after burning cycle was investigated for upper stage motors. The engines for the upper stage can be categorized into three group based on the cycles and propellants at the diagram. Kerosene engines which adapt the gas generator after burning cycle and are located in the region II, are characterized for high combustion pressure and complexity. This cycle usually needs more than two pumps to use the turbine power efficiently. The fuel line can be divided into the gas generator line and the combustor line, and only the gas generator line is need to be pressured more because the combustion pressure in the gas generator is much higher than that of the combustor. Basically, all the oxidizer goes into the gas generator and than to the combustor, thus the auxiliary LOx pump is not critically necessary. However, for the various reasons, the LOx line requires a booster pump. A gas generator after burning cycle engines produces relatively high specific impuls than that of the open cycle engines. Thus it is suitable for upper stages of launch vehicles.

• Feed Journal
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• v.3 no.5 s.21
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• pp.73-78
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• 2005

FTA지각생으로 불리던 우리나라도 FTA추진 논의가 활발하다. 작년 2월 FTA국회 비준 이후 싱가폴, 일본, ASEAN, EFTA 등 여러나라와 FTA관련협상을 추진 중이며 싱가폴의 경우에는 작년 11월 ASEAN +3 정상회담을 계기로 양국정상이 사실상 협상타결을 선언하였다. 종전에는 대부분의 나라들이 자유무역협정(FTA), 지역무역협정(RTA), 특혜무역협정(PTA) 지역주의 등 명칭은 다양하지만 FTA로 대표되는 지역주의를 다자간 자유무역체제의 보완적인 통상정책수단으로 간주하는 경향이 있었다. 그러나, 최근에는 FTA추진이 WTO 체제를 보조하는 통상정책수단이라기 보다는 WTO로 대표되는 다자간교역질서와 병립하는 통상질서로 보는 추세가 점점 나타나고 있다. 대외의존도가 높은 우리나라도 FTA를 통하여 개방정책을 추진할 필요성은 있으나 FTA의 추진시 농업 등 취약산업의 구조조정이 문제가 될 수 있다. 이글에서는 FTA가 WTO로 대표되는 다자간 무역체제와 구분되는 특징과 최근 국제적인 FTA추진동향을 짚어보고 우리나라의 FTA추진 현황과 문제점 특히, 우리나라가 제일 먼저 추진한 한$\cdot$칠레FTA 관련한 문제점을 중심으로 설명하고 최근의 FTA추진이 우리나라 농업에 대한 영향에 대해서 간략하게 평가하고 FTA추진에 대한 대응방안을 이야기 하는 순서로 글을 전개하고자 한다.

• Proceedings of the Korean Operations and Management Science Society Conference
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• 1996.04a
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• pp.458-461
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• 1996

1995년 5원 31일에 단군이래 최대의 개혁이라고 언론 매체들이 보도할 정도로 혁신적인 교육정책이 발표되었다. 또, 1996년 2월 9일에는 2차 교육 개혁안이 제시되었다. 이런 정책들이 추진되면서 우리나라 대학교육체계가 크게 변화하고 있다. 그 중에서도 학부제의 도입이 대학의 학사제도 개혁에서 최대의 관심사가 되고 있다. 1995년 말에 우리 나라 전대학의 1/3이 넘는 64개 대학이 96학년도부터 학부제를 추진하겠다고 발표하였다. 그러나 대학의 다양화와 자율성이 강조되면서 학부제에는 통일된 모형이 없으며 다양한 모습으로 학부제가 구성되고 있다. 특히 산업공학은 학문의 특성상 인접 학문이 다양하므로 학부제를 구성할 때 소속될 학부의 선택을 어렵게 하고있다. 앞으로 수년 내에 대학 지원 예정자들의 자연 감소와 대학 교육에 대한 사회적인 인식의 변화로 대학 교육 산업에도 소비자 중심시대가 도래하게 된다. 따라서, 학부제의 도입시에 산업공학의 소속과 역할도 소비자 중심 시대에 적합하게 결정되어야 할 것이다. 교육 서비스의 직접 소비자인 대학 입학생들에게 자유로운 전공 선택권을 갖게 하고, 교육 서비스의 간접 소비자인 기업에는 복합 학문의 기술 인력을 제공할 수 있는 학부제의 모형을 제시한다.

• Journal of the Korean Society of Propulsion Engineers
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• v.2 no.3
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• pp.54-61
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• 1998

A numerical simulation is attempted for the regenerative cooling heat transfer processes of the liquid propellant rocket engine. The heat transfer from the combustion gases to the thrust chamber wall is called gas side heat transfer. This heat is conducted radially to the coolant through the carbon deposit and metallic wall of thrust chamber Finally, this heat is convected away by the coolant flowing along the passages in the thrust chamber. The equivalence of these three heat fluxes of the above processes is utilized to determine the coolant side wall temperature, gas side wall temperature and the heat flux. When the number and shape(width, height) of coolant passages, the shape(size) of thrust chamber, oxidant and fuel properties, coolant properties, oxidant/fuel mixture ratio, coolant inlet temperature, the thickness of carbon deposit formed along the thrust chamber wall during combustion are given, reasonable radial direction temperature distributions and heat fluxes along the thrust chamber axis are obtained.

• 정보화사회
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• s.117
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• pp.25-34
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• 1997

정보화는 단지 구호나 선전 문구에 그쳐서는 안될 것이다. 이는 21세기를 준비하는 우리가 필수적으로 가지고 있어야 할 생존 전략인 것이다. 그리고 이를 추진하기 위한 최적의 대안이 바로 단계적인 인트라넷 구축을 통해서 국가 단위 인트라넷을 구축하는 것이다.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.4
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• pp.81-88
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• 2013

It was designed and tested at the design point that an oxidizer rich preburner for a staged combustion liquid rocket engine propelled by kerosene and LOx. The oxidizer rich preburner was designed as some of LOx injected from the mixing head was burned with kerosene and the rest of LOx injected from injection holes in the regenerative cooling chamber was vaporized by combustion gas. The preburner is operated at OF ratio of 60 and combustion pressure of 20 MPa. The Preburner has a honey-comb type mixing head with simplex swirl injectors, a turbulence ring improving combustion stability and uniformity of product gas temperature distribution, and a nozzle simulating the duct. With the combustion test results at the design point, the oxidizer rich preburner showed high combustion stability and uniformity of product gas temperature distribution.