• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.77-80
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• 2010

A series of hydraulic tests of a Lox pump are performed using water at a room temperature. According to the test results, the Lox pump satisfies its design requirement but both the head and the efficiency do not fully follow the conventional similarity rule. The deviation of the head from the rule is assumed to be due to the increased volute loss at high rotational speed. Furthermore, it is found that when the pump rotates with the flow ratio less then the design requirement the leakage flowrate seems to be increased.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.108-111
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• 2017

High performance upper stage engine is necessary for space launch vehicles of geostationary orbit, and staged combustion cycle engine is suitable due to high specific impulse. Technology demonstration model for 9 tonf class staged combustion cycle engine, which is consisted of turbopump, preburner, combustion chamber and supply system, was assembled, and hot-firing test was conducted for three seconds in Upper-stage Engine Test Facility of Naro Space Center. Ignition, combustion and shut down of engine system was performed normally, and its performance parameters were evaluated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.218-223
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• 2017

This paper deals with a fault diagnosis algorithm using the Kalman filter for a 75-ton Liquid Propellant Rocket Engine (LPRE). To design the Kalman filter, we linearized a non-linear simulation model of a 75-ton LPRE at an operating point, and checked the performance of the Kalman filter by comparing the measured values with estimated values of the states. Then, we artificially injected a fault of the turbopump efficiency into the simulation to confirm the performance of the fault diagnosis algorithm with the developed Kalman filter by comparing the variation of the residuals of the normal state with that of the fault cases.

• Tribology and Lubricants
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• v.22 no.2
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• pp.105-111
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• 2006

The floating ring seal which is used in the high pressure turbo pump is frequently used in the oxidizer pump and the fuel pump of the turbo pump of the liquid propulsion rocket, because it is able to minimize clearance to decrease the leakage flow rate. Compared with contact seal, the floating ring seal has advantage of minimizing clearance without rubbing phenomenon. But, the floating ring seal has a tendency to increase instability in operating condition in the high speed region. In this research, we devised floating ring seal which is inserted bump in the outer surface in order to improve the stability in the high speed region. Through this work, we expect to improve stability of floating ring seal with increasing the direct damping coefficient of seal and decreasing the eccentricity ratio.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.46-51
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• 2012

This study aims to develop the performance analysis program on the staged combustion cycle of the liquid rocket engine using liquid oxygen(LOx) as oxidizer, liquid hydrogen(LH2) and RP-1 as fuel. The developed analysis program can obtain the propellant mass flow rate, the specific impulse, and representative design values of engine components for the required thrust satisfying pressure, mass flow, and energy balance conditions. The analysis results show that the the specific impulses (Isp) compared to those of the real engines have been less than 1%. With additional constraints, the program will be improved for the system optimization.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.117-122
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• 2012

A blisk with rotor shroud is usually adopted in LRE turbine to maximize its performance. However it experiences severe thermal load and resulting damage during engine stating and stop. Shroud splitting is devised to relieve thermal stress on the turbine rotor. Structural analysis confirmed the reduction of plastic strain at the blade hub and tip. However, split gap at the rotor shroud entails additional tip leakage and results performance degradation. In order to assess the effect of shroud split on the turbine performance, tests have been performed for various settings of shroud split. For the maximum number of shroud splitting, measured efficiency reduction ratio was 2.65% to the value of original shape rotor.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.120-125
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• 2020

75 tonf liquid rocket engine combustion test was performed at Naro space center Engine Combustion Test Facility for KSLV-II. A gradual pressure drop was observed during off-design combustion test turbopump inlet condition using cooled kerosene at 271 K. It was found that the water content inside kerosene could cause pressure drop at 40 ㎛ grade filter through the water contests analysis of kerosene, kerosene cooling test and dehydration of kerosene.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.3
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• pp.71-80
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• 2020

Before we develop LVAFoam, a CFD software for launch vehicle analysis, we conducted a survey on other CFD softwares. We looked at in-house code and commercial CFD software of other countries that were used as a simulation of launch vehicle's combustor, turbopump and external flow. This research included in-house code solvers, developed by NASA, Mississippi State University, DLR, Bertin Technologies, CNES, CERFACS, and JAXA as well as commercial CFD software from FLUENT, CFX, Advance/FrontFlow/red, GASP, CRUNCH CFD, CFD-ACE+, FINE^TM/Turbo, STAR-CCM+. The simulation cases of launch vehicle analysis from each commercial softwares and introduction of the LVAFoam were presented.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.1
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• pp.68-76
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• 2022

Upper stage of the Korea Space Launch Vehicle-II(KSLV-II) uses a 7-tons class liquid rocket engine and is an open gas generator cycle with a turbopump supply method that uses kerosene/liquid oxygen as the propellant combination. This study first provided a brief overview of the design and development process of the upper stage engine. In addition, it introduced the solutions and results applied to some of the problems that occurred during the development process of the upper stage engine.