• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.4
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• pp.102-108
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• 2004

A layout of regenerative liquid rocket engine using turbo pump has been designed for development of high performance liquid rocket engine. each components of engine system was placed by considering assembly and characteristic. first stage engine system is controled by one plane of axis gimballing and composed of four engine assembly to cluster with launch vehicle. second stage engine system is controled by two plane of axis gimballing and composed of one engine assembly. assembly and disassembly Processes and required program have been developed. various shape of instruments were also developed for carrying out assembly and disassembly process efficiently.

• 한국연소학회:학술대회논문집
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• 2003.12a
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• pp.139-143
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• 2003

Study on the ignition characteristics of combustor and gas generator for LOx-kerosene liquid rocket engine was performed experimentally through a series of combustion tests of sub-scale engine combustor and gas generator. Characteristic of gas-torch ignitor based on gaseous methane and gaseous oxygen was compared with hypergolic ignition using propellant tri-ethyl-aluminium. Gas-torch ignitor showed good performance on igniting sub-scale liquid rocket engine combustor and gas generator. It was observed that the ignition delay is also affected by the extent of nitrogen in the combustion chamber.

• Journal of Aerospace System Engineering
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• v.9 no.1
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• pp.1-6
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• 2015

A critical speed analysis of oxidizer pump was peformed for a 7 ton class liquid rocket engine as the third stage engine of the Korea Space Launch Vehicle II. Based on the previously developed experimental 30 ton class turbopump and presently developing 75 ton class turbopump for the first and second stage rocket engine of Korea Space Launch Vehicle II, a layout and configuration of the 7 ton class turbopump rotor assembly are determined. A ball bearing stiffness analysis and rotordynamic analysis are performed for both of the bearing unloaded condition and loaded condition. Structural flexibility of the oxidizer pump casing is also included to predict critical speeds. From the numerical analysis, it is confirmed that the rotor system acquires sufficient separate margin of critical speed as a sub-critical rotor even though decrease of critical speed due to the casing structural flexibility.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.4
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• pp.67-74
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• 2001

Based on the national space development project, the necessity of developing liquid propellant rocket engine is revealed to secure the basic technology for the development of individual artificial-satellite launcher. Consequently, KARI (Korea Aerospace Research Institute) is developing a liquid propellant rocket engine for the KSR-III. Currently, a prototype engine using kerosene/LOx which produces 13-ton thrust is designed, fabricated and tested. In this paper, test procedure and technique for liquid propellant rocket engine are introduced with the analysis of static and dynamic test data.

• Advances in aircraft and spacecraft science
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• v.9 no.5
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• pp.401-414
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• 2022

In the present paper, the optimal selection of batteries for an electric pump-fed hybrid rocket engine is analyzed. A two-stage Mars Ascent Vehicle, suitable for the Mars Sample Return Mission, is considered as test case. A single engine is employed in the second stage, whereas the first stage uses a cluster of two engines. The initial mass of the launcher is equal to 500 kg and the same hybrid rocket engine is considered for both stages. Ragone plot-based correlations are embedded in the optimization process in order to chose the optimal values of specific energy and specific power, which minimize the battery mass ad hoc for the optimized engine design and ascent trajectory. Results show that a payload close to 100 kg is achievable considering the current commercial battery technology.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.4
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• pp.108-116
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• 2018

A firing test of the 75 tonf-class liquid rocket engine to be used as the first and second stage engines of the KSLV-II was carried out at the rocket engine test facility(RETF). Since this engine uses liquid oxygen as the oxidizer, which is a cryogenic fluid, it is essential that the chill down of the supply pipe line and engine proceed for the firing test; thus, the given inlet requirements must be met. Moreover, it is important to understand the chill down characteristics of the facility and the engine and the amount of liquid oxygen consumed in the chill down process for efficient test operation in the future. In this paper, chill down characteristics of the supply pipe and the engine were evaluated through the investigation of the chill down process of the 75 tonf-class liquid rocket engine at each stage before and after run tank pressurization. In addition, the amount of liquid oxygen consumed was also evaluated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2001.11a
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• pp.55-60
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• 2001

Recently, there have been many research efforts to improve the fuel regression rate in the hybrid rocket engines. In the present study, the ongoing research and development of the next general ion hybrid rocket engine are systematically reviewed. The detailed discussions have been made for the innovative combust ion technologies including the vortex hybrid rocket engines , cryogenic sol id propellant hybrid rocket engine, and the gas generator hybrid rocket engines.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2002.11b
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• pp.636-640
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• 2002

Acoustic coupled combustion instability, which is one of the most undesirable phenomena in the development of liquid propellant rocket engine, can cause serious damage to a rocket itself, and must be avoided by all means. Unfortunately, KSR-III rocket went through combustion instability during engine start at the propulsion test article No.2. To resolve the problem, time sequence (cyclogram) has been changed, and baffle system has been applied. In consequence of change, stable combustion was achieved.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2001.11a
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• pp.61-67
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• 2001

Hybrid propulsion systems provide many advantages in terms of stable operation and safety. However, classical hybrid rocket motors have lower fuel regression rate and combustion efficiency compared to solid propellant rocket motor. The recent research efforts are focused on the improvement of volume limitation and regression rate in the hybrid rocket engine. The present study has numerically investigated the combustion processes in the hybrid rocket engine. The turbulent combustion is represented by the eddy breakup model and Hiroyasu and Nagle and Strickland-Constable model are used for soot formation and soot oxidation. Radiative heat transfer is modeled by finite volume method. To reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect, the Low Reynolds number k-$\varepsilon$ turbulent model is employed. Based on numerical results, the detailed discussion has been made for the turbulent combustion processes in the vortex hybrid rocket engine.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2002.11a
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• pp.364.2-364
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• 2002

Thermoacoustic oscillation, which stems from phase correlation between unsteady heat release and acoustic fluctuation, can cause severe vibration and incite the excessive local heat transfer inside the rocket engine. It is very important to understand and prevent this phenomenon in the way of rocket engine development. In this study, the propulsion test facility of KSR-III, which is the first liquid propellant rocket developed by KARI, will be introduced. and the characteristics of thermoacoustic ocillation occurred at the facility will be examined.