• Journal of the Korean Society of Combustion
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• v.15 no.2
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• pp.12-18
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• 2010

A numerical analysis was performed to predict the thermo-fluid dynamic characteristics of hydrazine monopropellant reaction in the thruster combustor and nozzle. A 1-D porous model was introduced to simulate catalytic reaction by iridium in the combustor while 2-D axisymmetric analysis was applied to predict the nozzle flow. The chemical species and temperature variations were predicted by changing the injection pressure and mass flow rate and their results were validated by comparison with limited experimental data. The thrust variation with injection pressure could be estimated using the current 1-D combustor modeling.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1068-1071
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• 2017

The turbine exhaust system consists of a turbine flange, heat exchanger, exhaust duct and thrust nozzle. Heat exchanger is used for the launch vehicle because of the advantage of reducing the weight of the helium gas and the storage tank by using the heat exchanger pressurization method compared to the cold gas pressurizing method. Since the gas generator is combusted in fuel-rich condition, the soot is contained in the combustion gas. Hence, the heat exchanger should be designed considering the reduction of the heat exchange efficiency due to the soot effect. In addition, the uncertainty of the heat exchange calculation and the evaluation of the influence of the combustion gas soot on the heat exchange can not be completely calculated, so the design requirements must include a structure that can guarantee and control the temperature of the heat exchanger outlet. In this paper, it is described that the component allocation, the design method considering the manufacture of internal structure, the advantages of new concept of nozzle design.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.1
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• pp.91-96
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• 2014

KSLV-I (Korea Space Launch Vehicle-I) upper stage KM (Kick Motor) is a solid propulsion system which consists of igniter, SAD (Safety Arming Device), composite case, and submerged nozzle capable of TVC (Thrust Vector Control) actuation. Each subsystem of KM fulfilled development requirements for achieving a flight mission successfully. We confirmed the successful development of KM from the $3^{rd}$ flight test results of NARO on January 30, 2013. This article deals with the requirements of KM and the results on configuration management, mass variation, thrust axis alignment, and major test results and so on.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.6
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• pp.81-88
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• 2013

The component based propulsion modeling and simulation of an dual ramjet engine using Taylor-Maccoll flow equation and quasi 1-D combustor model. The subsonic and supersonic intake were modeled with Taylor-Maccoll flow having $25^{\circ}$ cone angle, the gas generator which transfers a pre-combustion gas into supersonic combustor was developed using Lumped model, and the determination of the size of nozzle throat of a gas generator was described. A quasi 1-D model was applied to model a supersonic combustor and the variation of temperature and pressure inside combustor were presented. Furthermore, the thrust and specific impulse applying fuel regulation by pressure recovery ratio and equivalence ratio were derived.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.43-46
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• 2006

Storage air heater(SAH) is a general purpose facility that is used to simulate the high altitude condition of supersonic ground test facility, thurst compensation test of rocket engine nozzle and gas turbine engine combustor test. SAH in KARI is built to simulate the total temperature of the supersonic ground test facility which has a wide flight envelope from altitude 0km, Mach 2 to altitude 25km, Mach 5 and operates up to 1300K, 3.5MPa. In this paper, we introduces the SAH in JAXA which is model of SAH in KARI and summarizes the design process and manufacture of ours.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.04a
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• pp.26-26
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• 1998

우주항공 산업에 대한 관심 증가에 따라 지상에서 많은 연소시험을 시행하고 있으나, 소음 발생문제에 부딪혀 어려움을 겪고 있는 실정이다. 따라서 초보적 단계이기는 하나, 액체로켓 엔진의 추력 손실을 최소화시키고 최대한 제트 소음을 크게 줄일 수 있는 소음기를 연구, 개발하고 있다. 본 연구에서는 제트소음에 대한 기초연구를 수행하여 물분사형 소음기를 설계 및 제작하였고, 물분사량과 소음기의 기하학적형상이 소음 특성에 미치는 영향을 연구하였다. 본 실험범위에서 연구의 결과는 다음과 같다. 1. 동일한 물 분사량 조건에서, 소음기 길이가 노즐출구 직경의 10배 모델 보다 30배 모델이 9dbl 정도 감음효과를 보였다. 2. 불 분사량이 증가함에 따라 소음레벨은 감소하였고, 30배 모델의 경우 불분사량이 배기가스의 10-12배 조건에서는 소음레벨을 91dbl까지 줄일 수 있었다. 3. 상기조건(소음레벨 91dbl)에 확장관을 부착함으로써, 소음레벨을 약 86dbl까지 줄일 수 있었다. 4. 본 형태의 물분사방식을 채택할 경우 고온배기가스로 인한 소음기의 파손을 방지하기 위해서 반드시 막냉각장치의 설치가 요구된다.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.1
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• pp.83-89
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• 2011

A design plan was proposed for determining combustor configuration of regenerative- cooled liquid rocket engine in the process of preliminary design. Rocket performance and regenerative cooling results were calculated using the properties of combustion gas estimated in CEA. For required thrust, chamber pressure, atmosphere pressure and propellant mixture ratio the mass flow rate of propellants and combustor performance were predicted by one-dimensional and experimental correlations. Finally, determinable plan for the contour of combustor were presented through Rao nozzle design method.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.37-42
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• 2010

A design plan was proposed for determining combustor configuration of regenerative- cooled liquid rocket engine in the process of preliminary design. Rocket performance and regenerative cooling results were calculated using the properties of combustion gas estimated in CEA. For required thrust, chamber pressure, atmosphere pressure and propellant mixture ratio the mass flow rate of propellants and combustor performance were predicted using one-dimensional and experimental equations. Finally, determinable plan for contour of combustor were presented through Rao nozzle design method.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.31-34
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• 2011

Design and fabrication of Technology Demonstration Model(TDM) of 75 tonf regenerative cooling thrust chamber were described. It has design chamber pressure of 60 bar, propellant mass flow rate of 243.6 kg/s, and nozzle expansion ratio of 12. It has a single welded structure of the mixing head and the chamber. Design and fabrication technologies established through this TDM can be used to development of flight model.

• Applied Chemistry for Engineering
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• v.30 no.5
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• pp.591-596
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• 2019

A recently developed propellant, ammonium dinitramide (ADN, $NH_4N(NO_2)_2$ is stable and safe at an ambient condition. However, it requires high purity for practical applications. A very little quantity of foreign impurities in ADN may cause clogging of thruster nozzles and catalyst poisoning for the use of a liquid propellant. Thus, several purification processes for precipitated ADN particles such as repetition extraction, activated carbon adsorption and low-temperature extraction were presented in this study. The purifying methods helped to improve the chemical purity as evaluated by FT-IR and UV-Vis spectroscopy in addition to ion chromatography (IC) analyses. Among the purification processes, adsorption was found to be the best, showing a final purity of 99.8% based on relative quantification by IC. Thermal analysis revealed an exothermic temperature of $148^{\circ}C$ for the synthesized liquid monopropellant, but rose to $188^{\circ}C$ when urea was added.