• Composites Research
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• v.27 no.5
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• pp.183-189
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• 2014

Experimental considerations have been performed to obtain the clear optical microscopic images of graphene oxide which are useful to probe its quality and morphological information such as a shape, a size, and a thickness. In this study, we investigated the contrast enhancement of the optical images of graphene oxide after hydrazine vapor reduction on a Si substrate coated with a 300 nm-thick $SiO_2$ dielectric layer. Also, a green-filtered light source gave higher contrast images comparing to optical images under standard white light. Furthermore, it was found that a image channel separation technique can be an alternative to simply identify the morphological information of graphene oxide, where red, green, and blue color values are separated at each pixels of the optical image. The approaches performed in this study can be helpful to set up a simple and easy protocol for the morphological identification of graphene oxide using a conventional optical microscope instead of a scanning electron microscopy or an atomic force microscopy.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.9
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• pp.649-657
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• 2019

Hydrazine, which is used as a representative monopropellant, is an extremely poisonous substance and has a disadvantage that it is harmful to the human body and is very difficult to handle. In recent years, research on the development of non-toxic and environmentally friendly propellants has attracted much attention. Ammonium dinitramide(ADN) based propellant developed by Swedish Space Corporation has superior performance to hydrazine and has been commercialized through performance verification in space environment. On the other hand, the exhaust gas from a thruster nozzle collides with a satellite while it is spreading in the vacuum space, thermal load and surface contamination may occur and may reduce the performance and lifetime of the satellite. However, a study on the effect of the exhaust gas of the green propellant thruster on the satellite has not been conducted in earnest yet. Therefore, the exhaust gas behavior in space was analyzed in this study for the ADN based green monopropellant using Navier-Stokes equations and the DSMC method. As a result, it can be expected to be used as design validation data in the development of satellite when using the ADN based green monopropellant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.11a
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• pp.8-8
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• 2000

하이드라진 단기액체엔진을 장착하고 궤도에서 임무를 수행하고 있는 다목적실용 위성 추진시스템 궤도비행 초기운용 자료에 근거하여 추진제 소모율을 산정 한다. 추진시스템은 위성의 궤도각과 비행고도 조정을 위한 속도증분($\Delta$V) 및 자세제어를 위한 추력을 발생시킨다. 단기액체 추진시스템에서 추진제 소모량은 추력기 밸브의 개폐시간에 비례하고 추력 생성 효율은 추진제의 연소기 유입압력에 종속한다. 일정질량의 가압 기체 압력에 의해 연료를 공급하는 추진시스템에서 잔류 추진제 량의 감소는 연소기 유입압력의 감소를 유발하고 추진기관의 효율을 저하시키는 요인으로 작용하여 임무말기로 진행함에 따라 동일한 운동량 생성에 보다 많은 연료소모가 이루어진다.(중략)

• Proceedings of the Korean Vacuum Society Conference
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• 2016.02a
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• pp.428-428
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• 2016

Copper indium sulfide (selenide) (CuIn(S,Se)2,CISSe)는 1.0~1.5 eV의 Direct band gap과 105 cm-1이 넘는 큰 광 흡수 계수를 가지고 있어 박막 태양전지의 흡수층으로써 연구되어 왔다. 최근 대량생산 및 저가 공정에 용이하다는 측면에서 용액 공정 기반 CISSe 태양전지 연구가 크게 주목 받고 있다. 용액공정 기반 중 하이드라진을 사용 한 경우 매우 높은 효율을 기록하였으나, 하이드라진 자체의 유독성과 폭발성 때문에 분위기 제어가 필요하고 여전히 저가화 및 대면적 제작에 한계가 있다. 따라서 알코올 솔젤 기반 CISSe 태양전지 제작 연구가 많이 진행되었으나, 결정립 성장 및 칼코겐 원자를 공급하기 위해 불가피하게 황화/셀렌화 후속 열처리 공정을 요구한다. 후속 열처리 공정은 폭발성의 황화수소/황화셀레늄 기체 분위기 제어와 고가의 장비를 필요로 한다. 본 연구에서는 매우 안정적이며 저가 용매인 물과 아민계 첨가제를 이용하여 Cu, In 전구체와 S, Se 이 포함된 Cu-In-S 잉크와 Se잉크를 제작하였다. 잉크 내에 S, Se을 첨가 함으로써 추가적인 후속공정 없이 비활성 가스 분위기에서 고품질의 CISSe 박막 제작을 가능케 하였다. 또한 Se 잉크 증착 횟수에 따른 결정 구조, 광학적 성질의 차이에 주목하였다. 따라서 수계 잉크를 대기 중에서 스핀코팅으로 박막을 제작한 후, Hot plate에서 건조하여 균일한 박막을 제조하고, 제작된 박막을 tube furnace에서 환원 분위기 및 비활성 가스 분위기에서 열처리 진행하여 $1.3{\mu}m$ 두께의 고품질의 CISSe 흡수층을 제작하였다. 이러한 흡수층에 대해 XRD, SEM, EDS 분석을 진행하여, 결정성, 미세구조, 및 조성을 확인하였으며, 제작된 흡수층 위에 버퍼층/투명전극층을 차례로 증착하여 CISSe 태양전지를 제작하여 셀 성능 및 양자 효율 특성을 파악하였다. 또한 액상 Raman 분석을 통해 결정립 성장 과정 메커니즘을 제시하였다.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.1-11
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• 1999

Thermal behavior of spacecraft propulsion system utilizing monopropellant hydrazine ($N_2$$H_4$) is addressed in this paper. Thermal control performance to prevent propellant freezing in spacecraft-operational orbit was test-verified under simulated on-orbit environment. The on-orbit environment was thermally achieved in space-simulation chamber and by the absorbed-heat flux method that implements an artificial heating through to the spacecraft bus panels enclosing the propulsion system. Test results obtained in terms of temperature history of propulsion components are presented and reduced into duty cycles of the avionics heaters which are dedicated to thermal control of those components. The duty cycles are subsequently converted into the electrical power required in the operational orbit. Additionally, cyclic temperature of each component, which was made under thermal-balanced condition of spacecraft, is compared to the acceptable design range and justified from the viewpoint of system verification.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.1-8
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• 2013

A liquid-monopropellant hydrazine thruster has several outstanding advantages such as relatively-simple structure, long/stable propellant storability, clean exhaust products, and so on. Therefore hydrazine thruster has such a wide application as orbit and attitude control system (ACS) for space vehicles. A hydrazine thruster with the medium-level thrust to be used in the ACS of space launch vehicles (SLV) has been developed, and its ground firing test result is presented in terms of thrust, impulse bit, temperature, and chamber pressure. It is verified through the performance test that the response and repeatability of thrust are very excellent, and the thrust efficiencies compared to its ideal requirement are larger than 93%.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.4
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• pp.333-342
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• 2016

Liquid rocket propulsion is a system that produces required thrust for satellites and space launch vehicles by using chemical reactions of a liquid fuel and a liquid oxidizer. Monomethylhydrazine/dinitrogen tetroxide, liquid hydrogen/liquid oxygen and RP-1/liquid oxygen are typical combinations of liquid propellants commonly used for the liquid rocket propulsion system. The objective of the present study is to investigate useful design and performance data of liquid rocket engine by conducting a numerical analysis of thermochemical reactions of liquid rocket propellants. For this, final products and chemical compositions of three liquid propellant combinations are calculated using equilibrium constants of major elementary equilibrium reactions when reactants remain in chemical equilibrium state after combustion process. In addition, flame temperature and specific impulse are estimated.

• Journal of the Mineralogical Society of Korea
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• v.23 no.3
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• pp.191-197
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• 2010

Fine nickel metal powder of uniform morphology, narrow size distribution, and high purity was prepared from high purity metal solution. Slurry reduction method for the synthesis of metal powder was applied with a special interest in their fine and spherical shape. The products were characterized by scanning electron microscopy (SEM) and X-ray diffraction (XRD). Well dispersed ultrafine nickel powder with the particle size range of 100~200 nm was produced from Ni-hydrazine precursor using hydrazine as a reductant for 90 min reaction in 4.5 M NaOH solution.

• Journal of the Korea Institute of Military Science and Technology
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• v.25 no.3
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• pp.285-291
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• 2022

In order to use the liquid rocket fuel, 1,5-diamino-4-methyltetrazolium azide, [DMT]⁺[N₃]^- and 1,5-diamino-4-methyltetrazolium cyanide, [DMT]⁺[CN]^- were synthesized and prepared the ionic liquid rocket fuel after dissolving the synthesized solid-type energetic chemicals in hydrazine, respectively. The thermal decomposition temperatures(T_d) and densities(d) of the prepared ionic liquid rocket fuels were about 200 ℃ and above 1.0 g/cm³ respectively. The ignition delay times(I_dt) of the ionic liqud rock fuels with [DMT]⁺[N₃]^- and [DMT]⁺[CN]^- were in a range of 26.6 - 82.5 ms and the 44.0 - 98.5 ms, respectively. These results mean that the synthesized tetrazolium salts could be used as an ionic liquid rocket fuels. The viscosities of the ionic liqud rock fuels with [DMT]⁺[N₃]^- and [DMT]⁺[CN]^-, which were dissolved in mixture solution of hydrazine/2-hydroxyethylhydrazine were to be 1.34 - 101 cP, and 1.29 - 80.5 cP, respectively. The synthesized ionic liquid rocket fuels in this study could be used as rocket fuel because the [I_dt(100 ms or less), T_d(150 ℃ or more), d(1.00 g/cm³ or more), and η(40.0~ 100 cP)] were achieved to satisfy the range of the used liquid rocket fuels.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.04a
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• pp.30-30
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• 2000

액체로켓엔진의 단일추진제 가스발생기는 연료공급 시스템의 터보펌프를 구동시키기 위한 작동가스 생성을 목적으로 사용된다. 고체추진제 가스발생기와 비교할 경우 작동시간이 보다 길고 연소생성물에 의한 터빈 블레이드의 삭마가 없으며 제어가 용이하므로 초기 액체로켓엔진 개발시부터 사용되어 왔다. 80년대 이후 개발된 액체로켓엔진은 이원추진제 가스발생기 또는 연소가스 FEEDBACK 시스템을 채용하고 있지만 단일추진제 가스발생기는 과산화수소수 또는 하이드라진과 같은 별도의 추진제 공급 시스템을 필요로 하는 단점에도 불구하고 상대적으로 낮은 온도의 무연 작동 가스를 발생하므로 가스발생기 자체를 위한 냉각시스템을 제거 또는 최소화 시켜 간단한 구조로 전체 시스템 설계를 가능하게 하므로 중소형 액체로켓엔진에 사용되고 있다.