• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1144-1147
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• 2017

It was reviewed about the pressure oscillation on the inlet of turbo-pump at the moment of engine startup and shutdown. Specially, This research was performed how much is the effect of PSD(Pogo Suppression Device) about the pressure oscillation on the inlet of turbo-pump at the moment of engine startup and shutdown. For analysis, propellant tank PSD and Engine are modelled with Flowmaster which is the commercial 1D program. As the analysis results, even though the PSD is installed in the pipeline, the pressure drop or rising at the moment of engine startup and shutdown is same compared to the case without PSD. However, it was confirmed that PSD reduces the pressure oscillation of the high frequency band as the original purpose of PSD.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.36-40
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• 2017

The launch operation for a space launch vehicle(SLV) is to be conducted by the systematic operation between SLV and the Technological Equipment(TE) such as the mechanical, fuel, and electrical ground support equipment at launch complex(LC). The basic source for the operation of the instruments in LC is the electrical power supply system, Technological Equipment Power Supply System(TEPSS), which is one of the Launch Control System. Thus TEPSS should supply the required electrical power to TE with reliability. In this paper, TEPSS which supplies operational electrical power to TE is introduced.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.112-115
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• 2017

We analyzed the rocket engine flow to check whether the possibility of the ground test and the equipment safety problems in the high altitude engine test facility. The test condition is that the vacuum chamber is open and the coolant water is injected into the supersonic diffuser. The analysis uses two-dimensional axisymmetry with a mixture of plume, air, and cooling water. As a result, the ground test was possible up to the cooling water flow rate of 200 kg/sec. However, due to the back flow of the initial plume, the vacuum chamber is exposed to high temperature, and at the same time, the inside of the vacuum chamber is contaminated due to the reverse flow of the cooling water. Therefore, sufficient insulation measures and work for pollution avoidance should be preceded.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1084-1087
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• 2017

When a liquid rocket engine is started the oxidizer and fuel must be flowed into combustion chamber and gas generator with time differences. The wrong time difference between propellants or malfunction of ignition device can occur the explosion of combustion chamber due to detonation by energized premixed-propellants. Therefore it is important to observe the transient characteristic of propellants or to measure the inflow time of propellants into combustion chamber and gas generator. The measurement of static pressure is not enough to observe the propellants inflow time into combustion chamber and gas generator. By measuring dynamic pressure of main flow passage of propellants the accurate propellants inflow time could be investigated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.679-681
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• 2017

Numerical analysis was carried out by applying the concept of E-D(expansion-deflection) nozzle to dual bell nozzle. We used the CEA code to calculate the chemical composition of the nozzle and to analyze the freezing flow of 8 species. The turbulence model was chosen as the $k-{\omega}$ SST. We applied the concept of E-D nozzle to the dual bell nozzle and performed the calculated transition altitude and performance. As a result of the interpretation, the application of the E-D nozzle concept led to the formation of over-expansion conditions, which resulted in an increase in the transition altitude.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.109-112
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• 2011

The result of design review and arrangement of a combustion chamber test facility(CTF) and a turbo-pump real propellant test facility(TPTF) is briefly described. The development/qualification tests of combustion chamber and turbo-pump for 75ton-class liquid rocket engine will be performed in CTF and TPTF. The critical design of hydraulic-pneumatic system, control and data acquisition system, test stand cell, and auxiliary facilities in CTF and TPTF was performed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.447-450
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• 2012

An analysis of chill-down process was performed for 30 tonf Turbopump-Gas generator coupled tests. The chill-down process must be fulfilled before liquid rocket engine test using cryogenic propellant. Cavitation, damage and/or combustion instability due to bubble of propellant must be eliminated by chill-down process in a test specimen, especially cryogenic pump. The analysis of test data obtained by 30 tonf TP-GG coupled tests was performed in order to be based on the test process of KSLV-II liquid propellant rocket engine which will be developed. To macroscopically understand the process of chill-down from the viewpoint of test procedure the temperatures of important part and total time of chill-down process were analyzed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.574-577
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• 2012

Thrust measuring is one of the crucial factor to decide the performance of a liquid rocket engine when the engine development test, especially for the combustion chamber, is implemented. Calculating the thrust from a combustion pressure is used when direct measuring the thrust is impossible, but direct measuring the thrust is necessary and various methods for doing it more precisely should be considered. This paper introduces the preliminary design concept about the new thrust measurement system for the vertical firing test stand, which is introduced domestically for the first time, of a liquid rocket engine combustion chamber.

• 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.93-100
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• 2020

The turbopump of the liquid rocket engine adapts an inducer to minimize the cavitation due to the variations of the propellants supply condition. However, the inducer introduces cavitation instabilities which are well-known problems in the engine development. In this paper, operational characteristics by the cavitation instabilities are analyzed and the reliability of the engine is checked when the first stage engine of the KSLV-II is tested at the low inlet pressure conditions. The characteristic frequencies representing the cavitation instabilities of the LOx pump are clearly found in various high frequency sensor signals around the entire engine in addition to the LOx and fuel pump.