• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.637-642
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• 2017

In this study, cold flow and combustion tests are conducted and analyzed to validate designed rocket engine head generating high temperature/pressure steam. At first, uni-injector was designed and manufactured, and cold flow test was conducted. Through this, differential pressure that can supply designed flow rate was confirmed. Also, Each injector's spray pattern were confirmed by patternator. Based on cold flow test results, we selected injectors among the candidates and arranged them on engine head, and cold flow and propellant spray tests were conducted. Finally, combustion test was carried out to analyze the flow rate, pressure, combustion efficiency. As a result, validation of rocket engine head for the development of the high temperature and high pressure steam generator has been completed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.601-602
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• 2009

대한항공은 2003년 소형위성발사체(KSLV-I) 사업 참여와 함께 2005년부터는 국내 액체로켓엔진 개발관련 한국항공우주연구원 주관의 각종 개발에 참여하고 있다. 본 논문에서는 현재 국내에서 진행중인 75톤급 액체로켓엔진 시스템 선행개발관련 대한항공이 수행하고 있는 분야별 업무의 소개와 함께 대한 항공의 향후 추진 계획을 다루고자 한다.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.1
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• pp.56-64
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• 2010

A system design has been conducted of the liquid rocket engine for Korean launch vehicle (KSLV-II, Korea Space Launch Vehicle II). The present turbopump-fed liquid rocket engine of vacuum thrust 76 ton and vacuum specific impulse 297 sec adopts gas generator cycle. The combustion pressure of the regeneratively cooled combustor is 60 bar. The propellant is LOx/kerosene. The engine is started by pyrostarter and the combustor is ignited by TEA (TriEthylAluminium). The engine system performance and the subsystems performance requirements are given through energy balance analysis. The combustion pressure, specific impulse and the engine mass are analyzed to be reasonable comparing with the published data. The startup analysis method which will be used in the future has been validated against the turbopump-gas generator coupled test. The tuning method for performance variation of the engine which is not actively controled has been prepared by mode analysis and performance deviation analysis.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.32 no.4
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• pp.249-254
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• 2008

Hydrodynamic performance test of a turbopump for a liquid rocket engine is carried out. The turbopump is composed of an oxidizer pump, a fuel pump and a turbine, and the two pumps are driven by the turbine. In the test, water is used for the pumps as working media and air is used for the turbine. Performance parameters of pumps and a turbine are drawn, and a power balance between the pumps and the turbine are calculated. The calculation shows a good power balance, which implies that the pump component tests, the turbine component test and the assembly test are reliably performed. At the starting period of the test, pressure rise-flow rate curve of a pump gradually approaches the ideal curve which could be obtained by very slow starting.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.814-815
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• 2010

Korea Aerospace Research Institute plan to develop propulsion system test facilities for combustor, engine system, propulsion systems of KSLV-II propulsion system in process of Korea Space Launch Vehicle project. By review for heavy test facilities specifications of foreign technically developed nations of the world, it will be referenced for test facility development plan of Korea Space Launch Vehicle project in the future.

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.3
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• pp.20-26
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• 2009

Turbopump test for a 30-ton-thrust liquid rocket engine was carried out using real-propellant. Liquid oxygen, kerosene, cold hydrogen gas were used for the oxidizer pump, the fuel pump, and the turbine, respectively. The turbopump was reliably operated at the design and off-design conditions and the performance requirements were satisfied, which implies that the turbopump development at the engine subsystem level is successfully accomplished in the point of performance validation. This paper presents the results of a test where the turbopump was run for 75 seconds at three operating modes. In terms of performance characteristics of pumps and turbine, the results of turbopump assembly test using real-propellant showed a good agreement with those of the turbopump component tests using simulant working fluid.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.359-365
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• 2008

Turbopump test for a 30-ton-thrust liquid rocket engine was carried out using real-propellant. Liquid oxygen, kerosene, cold hydrogen gas were used for the oxidizer pump, the fuel pump, the turbine, respectively. The turbopump was run stably at the design and off-design conditions and the performance requirements were satisfied, which implies that the turbopump development at the engine subsystem level is successfully accomplished in the point of performance validation. This paper presents the results of a test where the turbopump was run for 75 seconds at three operating modes. In terms of performance characteristics of pumps and turbine, the results from turbopump assembly test using real-propellant showed a good agreement with those from the turbopump component tests using simulant working fluid.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.465-467
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• 2012

A layout plan of a pressure-fed test facility to carry out hot-firing test of liquid rocket engine combustion chamber and purpose of rooms located in the test building were proposed. The layout plan of suggested infrastructure in this paper was determined depending on the design of a vertical test facility to use the natural lay of the land and simulate the initial position of flight model.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.928-929
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• 2017

A 75 ton class liquid rocket engine has been developed for KSLV II. Ground firing tests of the rocket engine were conducted to characterize and quantify its structural vibration under combustion testing environments. In this study, evaluation of dynamic response characteristics of the engine system was presented to verify its structural integrity and structural design during ground firing tests.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.10
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• pp.93-101
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• 2004

A method to calculate stability limits is investigated to predict the characteristics of high-frequency combustion instability in liquid-propellant rocket engine. It is based on the theory of linear stability analysis proposed in previous works and useful to predict combustion stability at the beginning stage of engine development. The system of equations governing reactive flow in combustor has the simplified and linearized forms. The overall equation expressing stability limits is adopted. The procedures to evaluate quantitatively each term included in the equation are proposed. The thermo-chemical properties and flow variables required in the evaluation can be obtained from calculation of thermodynamic equilibrium, CFD results, and experimental test data. Based on the existent data, stability limits are calculated with actual rocket engine (KSR-III rocket engine). The present calculations show the reasonable stability limits in a quantitative manner and the stability characteristics of the engine are discussed. The prediction from linear stability analysis could be serve as the first approximation to the true prediction.