• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.5
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• pp.107-114
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• 2018

Korea Space Launch Vehicle (KSLV)-I flight test data and the modified 1-dimensional steady state modeling data from the critical design results of the KSLV-II liquid oxygen filling system operation are compared to validate the reliability of critical design modeling. A comparison of major flow rates and pressure values between test data and calculation results are conducted. The relative errors relative to maximum total flow rate for each cooling, filling, and replenishment mode are determined within 6.7%. Calculated pressure values at the outlet of the pump and the inlet of flow control valves are within 5.1%. The pressure at the inlet of the launch vehicle for each operation mode are within the measured pressure range.

• 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 of Propulsion Engineers
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• v.16 no.4
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• pp.87-92
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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 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.

• Proceedings of the Korean Institute of Surface Engineering Conference
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• 2017.05a
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• pp.136-136
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• 2017

화학제염 기술은 산화제, 환원제, 금속이온, 무기산등이 혼합되어 있는 화학용액을 사용하여 원전기기 계통 내부에 생성된 고방사능 준위의 산화막과 오염물질을 제거하는 기술이다. 원전의 해체 및 유지보수에 있어 방사능 피복저감을 위한 필수적인 기술이다. 현재 원전 해체 산업은 잠재성이 높은 고부가가치 창출 산업으로 주목을 받고 있다. 원전 보유국의 경우, 기존 상용 제염기술과는 차별성 있는 제염기술을 확보하고자 노력하고 있다. 기존의 공정과 비교하여 공정비용 및 시간을 감소시킬 수 있어야 할 뿐만 아니라, 화학용액에 의한 원전 계통 금속 부품의 부식 및 손상을 최소화해야 한다. 금속 부품이 화학약품에 의한 부식손상을 받는다면 금속 부품의 수명 및 재활용 가치가 감소하기 때문에, 화학제염 기술 적용에 있어 용액에 대한 재료의 건전성 평가가 사전에 필히 이루어져야 한다. 본 연구에서는 원전 냉각재 펌프용 재료로 주로 사용되는 Stainless 304강을 시험편으로 선정하여, 화학제염 시험공정 3가지에 대한 부식손상 특성을 규명하였다. 산화공정은 과망간산($HMnO_4$) 용액을 공통으로 사용하였으며, 산화공정 종료 후 환원공정은 각 시험공정에 따라 시험공정 1은 옥살산($H_2C_2O_4$) 2000ppm, 시험공정 2는 옥살산($H_2C_2O_4$)1500ppm + 시트르산($H_8C_6O_7$)500ppm, 그리고 시험공정 3은 옥살산($H_2C_2O_4$) 3000ppm 용액을 각각 투입하여 수행하였다. 산화, 환원공정을 1Cycle로 하여, 각 시험공정 별로 총 5Cycle을 실시하였다. 각 시험공정 Cycle종료 후 시험편을 취외하여 무게감량측정, SEM(Scanning electron microscope) 분석, 3D현미경분석 그리고 타펠분극 실험을 실시하였다. 각 분석결과를 토대로 하여, Stainless 304강에 대한 화학제염 시 모델별 시험공정에 따른 부식특성을 규명하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.54-60
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• 2010

액체로켓엔진 시스템의 시동 및 정지 또는 추력 제어와 같은 천이 작동시 동특성을 예측하기 위한 선행 연구로서 추진제 공급 시스템의 구성품에 대한 동특성 모델링을 수행하였다. 연료 공급계통과 산화제 공급 계통의 구성품들은 재생냉각채널을 제외하고 같은 것으로 가정하였다. 동특성 모델링의 대상 구성품은 펌프, 관로, 오리피스, 제어 벨브, 재생냉각채널, 인젝터 등이며 실제 엔진 시스템의 축소모형에 대한 수력시험을 통해 각 구성품의 동특성 모델링을 검증하였다.

• Journal of Aerospace System Engineering
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• v.9 no.1
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• pp.1-6
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• 2015

A critical speed analysis of oxidizer pump was peformed for a 7 ton class liquid rocket engine as the third stage engine of the Korea Space Launch Vehicle II. Based on the previously developed experimental 30 ton class turbopump and presently developing 75 ton class turbopump for the first and second stage rocket engine of Korea Space Launch Vehicle II, a layout and configuration of the 7 ton class turbopump rotor assembly are determined. A ball bearing stiffness analysis and rotordynamic analysis are performed for both of the bearing unloaded condition and loaded condition. Structural flexibility of the oxidizer pump casing is also included to predict critical speeds. From the numerical analysis, it is confirmed that the rotor system acquires sufficient separate margin of critical speed as a sub-critical rotor even though decrease of critical speed due to the casing structural flexibility.

• Aerospace Engineering and Technology
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• v.4 no.1
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• pp.196-202
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• 2005

In cryogenic feeding system of turbo pump fed liquid-propulsion rocket, rise of cryogenic propellant temperature can bring into geysering in pipe or cavitation in turbo pump. In this paper, performance analysis of recirculation line which is one of the method to inhibit these phenomenon is carried out based on the configuration of KSLV-I 1st stage LOX feeding system, and parametric study to find design parameter. Diameter and re-entrance height, initial LOX temperature, ullage pressure, and natural convection heat transfer coefficient are varied to see the effects on performance. Additional He is injected into recirculation line to promote LOX recirculation. 1-dimensional analysis using network-solver, SINDA/FLUINT is carried out.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.477-480
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• 2011

The stress analyses of the 'U'-shaped multi-ply reinforced gimbal bellows under high pressure and rotational displacement loadings are performed at the room and cryogenic temperatures. The bellows are used for the Lox pipe line which connects the combustion chamber with the turbopump of a liquid rocket engine. The distributions of the stress, the strains and the contact pressures are obtained from the finite element analysis considering the geometric non-linearities of the contacts between the plies and the material one of the isotropic plasticity. Those are compared with the stress results from EJMA (Expansion Joint Manufacturing Association) standard. Also, the effects of the operating temperature and the reinforcing ring on the stresses are investigated.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.5
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• pp.91-96
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• 2004

An optimal design of the gas generator for Liquid Rocket Engine (LRE) was conducted. A fuel-rich gas generator in open cycle turbopump system was designed for 10ton in thrust with RP-1/LOx propellant. The optimal design was done for maximizing specific impulse of thrust chamber with constraints of combustion temperature and for matching the power requirement of turbopump system. Design variables are total mass flow rate to gas generator, O/F ratio in gas generator, turbine injection angle, partial admission ratio, and turbine rotational speed. Results of optimal design provide length, diameter, and contraction ratio of gas generator. And the operational condition predicted by design code with resulting configuration was found to maximize the objective function and to meet the design constraints. The results of optimal design will be tested and verified with combustion experiments.

• Aerospace Engineering and Technology
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• v.12 no.1
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• pp.200-206
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• 2013

A performance sensitivity of liquid rocket engine to propellant density or supply pressure change was studied. The analysis program was verified to have 1% error comparing with the measured data of a turbopump-gas generator system. The engine combustion pressure decreases as fuel supply pressure increases due to decreased mixture ratio which reduces the turbine power. The engine combustion pressure increases as fuel density increases because the total propellant flow rate is increased substantially even though mixture ratio is slightly decreased. The engine combustion pressure increases when the oxidizer density or supply pressure increases.