• 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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• 2011.04a
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• pp.7-11
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• 2011

A velocity-compounded(VC) turbine for the 75ton turbopump was developed as an improved performance backup for the single-rotor baseline turbine. Curvic coupling was adopted for the power transmission between the rotors and shaft. High temperature torsion test and spin test was performed for the curvic coupling design validation. Aerodynamic performance test revealed that VC turbine can generate 20.5% higher specific power than the baseline turbine.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.4
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• pp.54-63
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• 2022

A 3-tonf class liquid rocket engine that powers the upper stage of a small launcher and lifts 500 kg payload to 500 km SSO is designed. The small launcher is to utilize the flight-proven technology of the 75-tonf class engine for the first stage. A combination of liquid oxygen and liquid methane has been selected as their cryogenic states can provide an extra boost in specific impulse as well as enable a weight saving via the common dome arrangement. An expander cycle is chosen among others as the low-pressure operation makes it robust and reliable while a specific impulse of over 360 seconds is achievable with the nozzle extension ratio of 120. Key components such as combustion chamber and turbopump are designed for additive manufacturing to a target cost. The engine system provides an evaporated methane for the autogenous pressurization system and the reaction control of the stage. This upper stage propulsion system can be extended to various missions including deep space exploration.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.7
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• pp.517-524
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• 2019

A system design has been performed for a vacuum thrust 88 ton staged combustion cycle rocket engine. Previous research has been used to estimate the performance of the engine components. And the algorithm has been proposed to evaluate the converged engine system performance. The present methodolgy has been verified by comparing the published data for RD-180. The present work adopts the most of the previous KSLV-II engine heritage for both performance improvement and cost competitiveness. The combustion pressure has been decided as 12MPa considering manufacturing difficulty, cost and performance improvement, and as a result the vacuum specific impulse has increased by 23.4s.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.149-152
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• 2010

It was investigated that effect of the supersonic impulse turbine rotor leading edge thickness which was the part of 75 ton open cycle liquid rocket engine turbopump. The test for turbine was performed that the rotor thickness to pitch ratio was 1.9 and 1.4 to 30 ton turbine. As a result of test, the rotor with lower thickness(1.4) had 1.5% efficiency gain to the higher thickness(1.9) and the pressure ratio with maximum efficiency was increased to the nozzle full expansion point.

• 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.

• 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.38-44
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• 2011

The performance experiment was tested for the velocity compound turbine of turbopump which was the main part of 75 ton class liquid rocket engine. The seal is installed between the 1st rotor and the reversing vane to reduce the leakage flow. The turbine outlet pressure of the velocity compound turbine by changing the rotating speed was compared with that of baseline turbine with single rotor including the effect on the total performance.

• 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.