• Aerospace Engineering and Technology
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• v.8 no.1
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• pp.26-31
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• 2009

Three Lox pumps and one fuel pump are manufactured for turbopump real-propellent tests and water tests of the pumps are performed in order to estimate the performance characteristics of the pumps. According to the test results, the test region(flow ratio, cavitation number) of the pumps at the water tests cover the operating region at the real-propellent tests and also all the pumps satisfy the design requirement. The head of the Lox pumps shows a 2% difference among them due to the internal geometry, but the efficiency and overall cavitation performance are almost same. It is found that the fuel pump has a similar head and efficiency compared with the previous model of the same internal geometry, while it has a little inferior cavitation performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.56-61
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• 2012

Fuel-side cold-flow tests were performed on the technology demonstration model of a 75 ton-class liquid rocket engine combustion chamber for the first stage of the Korea space launch vehicle II. Pressure drop in the cooling channels of the combustion chamber was measured by changing fuel mass flow rate through a pressure regulating system. Pressure drop in each segment of the chamber could be obtained and a lot of pressure drop was caused by high flow velocity in the nozzle throat segment. The accuracy of a hydraulic analysis method for calculating a pressure loss in cooling channels could be verified by applying it to the cold-flow test conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.807-812
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• 2011

Fuel-side cold-flow tests were performed on the technology demonstration model of a 75 ton-class liquid rocket engine combustion chamber for the first stage of the Korea space launch vehicle II. Pressure drop in the cooling channels of the combustion chamber was measured by changing fuel mass flow rate through a pressure regulating system. Pressure drop in each segment of the chamber could be obtained and a lot of pressure drop was caused by high flow velocity in the nozzle throat segment. The accuracy of a hydraulic analysis method for calculating a pressure loss in cooling channels could be verified by applying it to the cold-flow test conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.97-100
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• 2009

Cold flow and ignition tests were performed for a technology demonstration model of a 75-tonf thrust chamber which is a candidate liquid rocket engine for a next Korea Space Launch Vehicle. The test facility was modified to support the new concepts of the thrust chamber such as ignition system, film cooling and LOx leading supply. The hydrodynamic characteristics of the supply pipelines, thrust chamber and igniter as well as the filling time of the propellants were obtained through the cold flow tests on the LOx and kerosene and the ignition cyclogram was determined using the results. The ignition test was successfully accomplished according to the cyclogram and therefore, a basic information was obtained for further hot firing tests.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.1
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• pp.35-44
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• 2011

Mathematical modelling for liquid rocket engine(LRE) main components were conducted to predict the dynamic characteristics when the LRE operates at the transient condition, which include engine start up, shut down, or thrust control. Propellant feeding system is composed of fuel and oxidizer feeding components except for regenerative cooling channel for the fuel circuit. Components modeling of pump, pipe, orifice, control valve, regenerative cooling channel and injector was serially made. Hydraulic tests of scale down component were made in order to validate modelling components. The mathematical models of engine components were integrated into LRE transient simulation program in concomitant with experimental validation.

• Journal of the Korean Society of Propulsion Engineers
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• v.27 no.1
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• pp.58-67
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• 2023

Results of dry-run tests, cold-flow tests, and hot-firing tests performed to throttle a methane engine uni-element thrust chamber are covered in the paper. After installing flow control valves on the oxidizer and fuel supply lines of the methane engine combustion test facility, a number of dry-run tests were repeated so that the valves could reach set strokes quickly and stably. Then, cold-flow tests using liquid nitrogen and gaseous nitrogen were conducted to confirm the stable supply of the simulated propellants according to the valve control. Finally, using liquid oxygen and gaseous methane, hot-firing tests for fixed and continuous thrust control of 50% to 10% of the nominal thrust were successfully performed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.41-45
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• 2009

The small-sized bi-propellant thruster using a high concentrated hydrogen peroxide and kerosene as the oxidizer and fuel was designed and fabricated in this study. The water cold-flow test was performed to verify the performance characteristics of the injector. The mixing head assembly used in this model thruster was designed as a structure to combine igniter, injectors and film cooling, which are capable of regulating each mass flowrate. This maximize the experimental verification and efficiency of the design optimization. Finally, the mass flowrate and spray pattern of injector were evaluated by the hydraulic test. Therefore, the design validity of the mixing head was verified.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.4
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• pp.64-71
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• 2022

For the development of an improved upper-stage engine, research on a staged combustion cycle liquid rocket engine is in progress. A cold flow test, ignition test, and combustion test plans were established and performed to develop reignition combustion technology. In order to solve the problem of purge gas flowing into the fuel line, which may cause cavitation in the turbo pump during reignition, the test results of each stage were analyzed. Based on the analysis results, the purge gas inflow problem was solved by reducing the overlapping time between the operation of the bubble removal valve and the opening of the purge valve and the engine fuel valve. Based on this, the reignition combustion test was successfully performed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1995.11a
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• pp.49-55
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• 1995

이원 액체추진제를 사용하는 인공위성용 로켓 추진기관의 개발을 위한 핵심부품별 개념 및 상세설계, 성능해석, 성능실험용 시작품의 제작, 수류 성능실험 및 지상 연소시험이 수행되었다. 인공위성 궤도조종용 로켓 추진기관은 1.38MPa의 연소실 압력으로 4초동안 1780N(400$Ib_f$)의 평균추력을 내도록 설계되었으며, 산화제로는 질산, 연료로는 트리 에틸렌 아민(triethylene amine, TEA)과 자이리딘(xylidine)의 혼합물로 구성된 접촉발화형 이원 액체추진제를 사용하고, 추진제를 가압방식에 의해 연소실에 분사하는 방법으로 분사충돌, 미립화, 그리고 기화 후 연소시키게 된다. 효율적인 설계를 위하여 설계전용 소프트웨어를 개발하였으며, 추진기관의 핵심부품별로 유동 시뮬레이션을 수행하고, 해석결과와 수류 실험결과를 바탕으로 설계를 수정, 보완하였다. 지상 연소시험 및 수류 성능실험을 위하여 추진제 공급장치 및 계측 시스템이 설계, 제작되었고, 시스템의 작동 및 자료처리를 위한 소프트웨어를 개발하여 수류 성능실험 및 지상 연소시험에 사용하였으며, 연소시험결과 지상 평균추력 378$Ib_f$를 발생하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.11a
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• pp.4-4
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• 2000

KSR-III 과학로켓 개발사업의 한 분야인 추진기관 연구에 있어 가장 중심이 되는 것은 제작된 연소실의 시험/검증이라고 할 수 있다. 여러 단계의 시험 가운데에서도 엔진의 특성을 제 1차 적으로 파악하고, 시험 대상물인 엔진을 스탠드에 장착하여 시험이 가능토록 하기 위해서 모사 추진제(물) 혹은 실제 추진제를 사용하여 수력학적 특성을 알아내는 것은 매우 중요하다. 이를 위해 engineering model 엔진(이하 EM)에 대하여 액체 산소 배관은 액체 산소를, 스탠드 사정상 케로신 배관은 액체 질소를 사용한 비연소 시험을 수행하였으며, 이러한 일련의 비연소 시험을 통해 산화제 및 연료 매니폴드 각각에 대한 차압과 매니폴드를 채우는 시간을 측정하여 점화 사이클을 정의할 수 있었다.(중략)