• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.281-285
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• 2007

The flow characteristics in the thruster should be analyzed considering its geometry and the pressure ratio to estimate its performance and etc. This paper suggests the computational result of an axisymmetric real nozzle for the altitude control of a satellite to find out the application limit that the assumption of continuum mechanics holds. The steady non-reacted compressible flow field in the unstructured grid system is computed and analyzed with varying the environmental pressure (or the degree of vacuum) under the fixed pressure ratio in a real thruster of which the area ratio of exit to throat is 56. The assumption of the continuum mechanics is not approved when the environmental pressure is reduced less than $10^{-3}$ atm.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.503-506
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• 2009

In this study, we presented the results of static firing tests on the PE/LN2O hybrid rocket motor, which has a thrust of 650 kgf level. Through the early tests, we found that the combustion chamber pressure and the thrust were lower than design values because an actual oxidizer flow rate was less than that expected. In order to complement this result, the methods of decrease of nozzle throat and the increase of oxidizer mass flow rate were conducted in the next experiment, and we studied the combustion phenomena with the experimental results. Also we compared and analyzed a difference of combustion characteristics on scale effect. It show that a sub-scale motor regression rate was a little less than that of a lab-scale motor with the same oxidizer mass flux. Results of this study might be used as a basic data for development of hybrid sounding rocket.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.585-589
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• 2010

It is important for the liquid rocket engine to meet the exact performance requirements in order to guarantee the successful mission of the launch vehicle. Usually, a ground combustion test for the engine is conducted to reduce the performance error and for the tuning. For the gas-generator (GG) cycle engine, this adjustment process can be easily tuned by means of the control valves. A linearized correlation between the process parameters of the control - the combustion chamber pressure and the mixture ratio of engine - and the independent parameter of the control- rotational angle of the control valve - could be suitable to reduce the tuning errors. Also this linearity can reduce the effort for the tuning and make the process more explicit by ensuring a more intuitive control. In this point, we proposed an algorithm in the frame of the in-house-developed program to obtain the control valves' inherent characteristics which satisfy the linearity.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.3
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• pp.1-9
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• 2004

We conducted the firing test with the regenerative cooling LRE and calculated the heat flux from measured coolant temperature, that was compared with the heat flux predicted by previously developed numerical analysis method. The difference between the measured heat flux and the numerical calculation value was within nine percents. Therefore, developed numerical analysis method can be applied to the design/fabrication of a real LRE system. and, it was investigated that combustion pressure and mixture ratio have an Influence on the heat flux with a constant relation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.53-56
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• 2003

This paper describes the general design procedure of cooling system for liquid rocket engine(LRE). From this design logic, cooling channels are designed and fabricated. The measured heat flux from firing test is similar to the heat flux predicted by design logic. Therefore, the proposed design procedure of cooling channel can be applied to real LRE system. Also the result of firing test indicates that combustion pressure and mixture ratio have an influence on the heat flux to be produced in combustion chamber.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.109-117
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• 2006

Hot firing performance test of hydrazine decomposition catalyst used for monopropellant thruster of satellite and launch vehicle was performed on the ground. A test equipment for hot firing performance measurement of catalyst test was developed in collaboration with Hanwha Corp., and the catalyst firing performance were tested with the equipment. After a reaction delay time, a catalyst activity and a granule stability were measured for 2 times, satisfactory results were obtained such as 25msec, 2%, $704^{\circ}C$ for each test items on the average. In addition, the current development status of domestic prototype catalyst and its decomposition performance test results are presented.

• Bulletin of the Korean Space Science Society
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• 2009.10a
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• pp.40.2-40.2
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• 2009

이 연구에서는 달천이(TLI: Trans Lunar Injection) 및 달포획(LOI: Lunar Orbit Injection) 기동 시 대전 지상국의 가시성을 고려한 최적의 임무를 설계하였다. TLI 기동은 탐사선이 지구 주차궤도에서 지구-달 천이궤적으로 진입하기 위하여 주어지는 기동이며, LOI 기동은 탐사선이 지구-달 천이궤적에서 달의 중력권으로 진입하기 위하여 주어지는 기동이다. TLI 및 LOI 기동 시 대전 지상국에서의 가시성의 확보는 실제적인 미래 한국의 달 탐사를 대비하였을 때 중요한 요소이다. 따라서 이 연구에서는 TLI 및 LOI 기동 시 대전 지상국에서의 가시성을 모두 고려하여, 최소연료로 지구 주차궤도에서 달 임무궤도 진입까지의 모든 단계에 대해 임무설계를 실시하였다. TLI 및 LOI 기동 시 추력은 순간 추력(Impulsive thrust)로 가정하였으며, KSLV-II 발사체의 성능을 적용하여 설계하였다. 임무 설계 시 태양, 지구, 달의 섭동력을 고려한 N체 운동 방정식을 탐사선에 적용하였으며, 지구의 비대칭 중력장, 태양 복사압, 달의 J2 섭동에 의한 영향도 고려하였다. JPL의 정밀 천체력인 DE405를 사용하였고, 상용 소프트웨어인 SNOPT(Spares Nonlinear OPTimizer)를 이용하여 비행 궤적의 최적해를 도출하였다. 임무 설계 결과를 통해, 대전 지상국의 가시성을 고려한 TLI 및 LOI 기동의 크기에 의한 임무설계의 분석을 수행하였다. 또한 최적화된 달 탐사 임무의 단계별 기동의 크기와 지구-달 천이 궤적의 형상 및 다양한 임무 요소들의 해석을 도출하였다.

• The Bulletin of The Korean Astronomical Society
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• v.37 no.2
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• pp.148.1-148.1
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• 2012

로켓 혹은 우주발사체의 주엔진에는 대부분 연료와 산화제를 연소시켜 나오는 에너지를 사용하는 화학로켓이 주종을 이루어 왔다. 이러한 로켓엔진에서 그동안 연료로는 수소계, 탄화수소계, 아민계 등 다양한 화학물질이 사용되어 왔으나, 산화제로는 강한 산화성을 나타내면서 밀도가 높은 몇몇 물질만이 제한적으로 사용되어져 왔으며, 최근에는 주로 액체산소(LOx)와 사산화질소(N2O4)가 사용되고 있다. 그러나 산화제 중 액체산소는 극저온이면서 상대적으로 밀도가 낮고, 사산화질소는 강한 독성을 지니고 있으며 액체로 존재하는 구간이 좁아 연구 목적의 소형발사체를 구현하는 것에는 많은 어려움이 있다. 이러한 이유로 최근 소형발사체 개발분야에서는 상온저장성이면서 친환경적인 과산화수소(H2O2)와 아산화질소(N2O)를 산화제로 활용하는 것에 대한 관심이 고조되고 있으나, 대형 추진기관을 개발하는 연구자들로부터는 액체산소를 사용할 때 보다 엔진 자체의 비추력이 상대적으로 낮다는 이유로 활용이 외면되어 온 것이 사실이다. 본 연구에서는 엔진 자체의 추진성능 보다는 사실상 발사체의 목적이라고 할 수 있는 추진단 속도증분을 성능의 지표로 삼아 평가하였으며, 결과를 통하여 과산화수소와 아산화질소의 높은 밀도가 엔진의 낮은 비추력을 충분히 보상할 수 있음을 보였다. 과산화수소와 아산화질소는 교육/연구용 소형발사체 구성에 충분히 활용가능한 산화제이며, 실제 발사에서 충분한 비행성능을 기대할 수 있는 물질로 평가할 수 있다.

• Aerospace Engineering and Technology
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• v.10 no.2
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• pp.121-127
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• 2011

In the attitude and orbit control subsystem design, the moment of inertia of the satellite is the major contributor to be considered. Satellites equipped with large solar arrays need to measure the moment of inertia accurately to avoid the interference of the thruster actuation period with its flexible mode. In this paper, the in-orbit tests of COMS to measure the moment of inertia are described. Then, the differences between the measured through in-orbit test and the predicted are compared. Finally, it is verified that the differences are below uncertainty bounds considered in the critical design of COMS attitude and orbit control subsystem.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.4
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• pp.87-93
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• 2016

In a Direct Connect(DC) mode altitude engine test, a labyrinth seal is set up between an air intake duct and an engine. The labyrinth seal plays a key role in mechanically isolating them, which contributes to the accurate measurement of thrust and the other component forces. However, when high vertical temperature gradient is generated in the supplied air in the duct, the isolation breaks down. In this paper, a labyrinth seal control device is designed and installed in an effort to eliminate the issue. Test result shows the device successfully gets rid of the contact problem even when high vertical temperature gradient is produced.