• Aerospace Engineering and Technology
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• v.6 no.1
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• pp.190-200
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• 2007

A pyrogen type igniter was designed to satisfy the requirements of KSLV-I Kick Motor system. To insure the reliability of the igniter before the production of the flight model, we have been performed the structure, environmental, combustion tests. The hydraulic test was carried out to confirm the strength of the components of the igniter. The shock and vibration tests were considered to check whether the igniter operates normally under the severe environmental condition. The combustion tests were also performed to understand the ignition characteristics with the variation of initial condition. Finally, we confirmed that the igniter could provide the acceptable energy to ignite the propellant of kick motor at the ground test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.318-321
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• 2005

The design and performance test of a low NOx gas turbine combustor to be used in 55kW class micro-gasturbine engine was performed in KARI's combustion test facility. The combustor is reverse flow-can type for easy installation of injector and other parts and LNG is used as fuel. The performance targets are $99.5\%$ combustion efficiency, less 10ppm NOx, $30\%$ patten factor and $4\%$ pressure loss. Most of the performances required are satisfied.

• 열병합발전
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• s.24
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• pp.5-9
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• 2001

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.5
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• pp.22-30
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• 2015

A gas generator which generates turbine driving gas by burning a part of propellants is used in an open cycle liquid rocket engine and as a main component of an open cycle liquid rocket engine autonomous hot firing tests are required to investigate the combustion performance and characteristics of the gas generator. However, since the combustion gas generated by a gas generator is choked at the turbine nozzle in the turbine manifold, it is necessary to consider the internal volume of turbine manifold as well as that of the gas generator for correct investigation of the combustion performance, characteristics, and acoustic characteristics of the gas generator. Therefore, in the paper hot firing test results of a gas generator with a turbine manifold simulator are described and characteristic prediction using the autonomous test of a gas generator is explained.

• Proceedings of the Korea Institute of Fire Science and Engineering Conference
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• 2011.04a
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• pp.180-187
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• 2011

본 논문은 다중이용업소와 목조건축물에 자주 사용되는 미송합판에 방염처리를 하여 유사 화재를 구현하고, 그 화염 세기에 따른 방염의 실효성을 실험한 것이다. 방염처리를 하면 화재 시 가연물의 초기착화시간을 지연시켜 화재성장속도를 늦출 수 있고, 원활한 소화활동을 가능하게 해준다. 하지만 어느 정도 화재가 진행되어 화염이 거세지면, 45도 연소시험을 통한 방염기준을 충족하여도 그 성능을 기대하기 어렵다고 한다. 따라서 45도 연소시험 시 사용되는 65mm의 불꽃보다 큰 화염상태(초기착화 이후의 상태)에서 방염처리한 내장재(미송합판)의 방염성능이 유지되는지의 여부를 실제로 입증하고 그 근거를 뒷받침하기 위하여 본 연구를 시작하게 되었다. 실험에서는 화재의 규모(화염의 세기)를 달리하여 각기 다른 종류의 방염제로 방염 처리한 미송합판의 착화 시 화염온도, 복사열 유속 그리고 착화지연시간을 파악하였으며, 45도 연소시험과 관련하여 방염성능을 분석하였다. 45도 연소시험의 경우 실험에 사용한 방염 처리 합판은 방염성능 기준을 만족하는 것으로 나타났으며, 소규모 유사 화재로 직경 10cm 연소용기를 사용한 연소실험에서는 방염 처리한 합판의 착화지연시간이 평균적으로 대규모 유사 화재실험보다 길어 어느 정도는 방염효과를 갖는 것으로 나타났다. 하지만 대규모 유사 화재로 1단위 유류화재 연소용기를 사용한 연소실험의 경우 열방출율이 커 형성된 탄화막이 무분별하게 박리되고 발화가 일어나 착화지연시간의 차이를 구별하기 어려웠기 때문에 방염효과를 기대할 수 없었다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.290-293
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• 2017

A direct-connected, continuous type combustion test facility was developed to test a supersonic combustor model used in scramjet engines. The facility requirements were determined by assuming the flight speed of Mach 5, yielding the combustor inlet flow speed of Mach 2. Also the cross-section of the supersonic combustor under test was assumed as $32mm{\times}70mm$. As a result, the facility was designed to have the flow total pressure of 548 kPaA, total temperature of 1,320 K, and flow rate of 0.776 kg/s. The facility consists of a turbo type air compressor, electric air heater, vitiation air heater and a two dimensional facility nozzle to accelerate the flow to Mach 2. Also, an oxygen supply system was added to compensate the vitiation. The exhaust de-pressurization system is not added. Designed pressure, temperature, and flow rate were verified through the test operation of the facility.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.6
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• pp.45-52
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• 2021

A hot-firing test was conducted using gaseous-methane and gaseous-oxygen under highly fuel-rich condition as a prior study for the development of a liquid propellant small rocket engine. To compare combustion characteristics for various equivalence ratios, the oxygen flow rate was set to 12 g/s and the methane flow rate was changed according to the equivalence ratio. As a result, it was observed that the steady-state characteristic velocity obtained during the hot-firing test steeply rose in the latter part of each test: the difference between the former and the latter steady value was enhanced overall in proportion to the equivalence ratio. Based on this, the equivalence ratio range depending on the variational characteristics of the characteristic velocity could be divided into three combustion regimes.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.726-728
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• 2010

A subscale gas generator was designed and manufactured to investigate the effect of design parameters on discharge coefficients of injectors for a 75 ton-class gas generator and hot-firing tests were successfully performed. The test results showed that discharge coefficients of fuel and liquid oxygen injectors remained nearly constant irrespective of variations of a mixture ratio and a chamber pressure. When the post diameter of the liquid oxygen injector was reduced, the discharge coefficient was increased as the pressure drop of the injector was decreased.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.11a
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• pp.6-6
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• 2000

국가우주개발 중장기 계획에 의거, 독자 인공위성 발사체 개발에 필요한 필수기술을 확보하기 위하여 액체추진제 로켓엔진의 개발에 대한 필요성이 대두되었으며, 이에 따라 한국항공우주연구소는 과학로켓 3호(KSR-III)에 적용하기 위한 액체추진기관을 개발하고 있다. 이러한 목적으로 kerosine/LOx를 사용하며 13톤급의 추력을 낼 수 있는 시제엔진이 설계, 제작되었으며 이 엔진에 대한 연소시험이 실시되었다. 본 연구에서는 액체추진기관 시험을 위한 일련의 진행사항, 시험방법을 소개하며, 시험을 통하여 획득한 정특성 자료 및 동특성 자료에 대하여 분석하였다.(중략)

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.6
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• pp.34-40
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• 2009

The present study describes the methods and standards for the combustion stability assessment of a thrust chamber and a gas generator as parts of a liquid rocket engine. The first method uses a statistical approach through typical static combustion tests and the second one a dynamic assessment identifying decaying characteristics of pressure fluctuations excited by a pulse generating device. Based on accumulated test results, it is concluded that the maximal values for combustion stability are 3% of a chamber static pressure with a Root-Mean-Square value of pressure fluctuations, and 10 msec with a decay time.