• 유체기계공업학회:학술대회논문집
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• 2006.08a
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• pp.245-248
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• 2006

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. Accordingly, the recent research efforts are focused on the improvement of engine efficiency 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 flamelet model and Low Reynolds number $k-{\varepsilon}$ turbulent model is employed to reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect. Based on numerical results, the detailed discussions have been made for the effects of oxygen injection methods and oxygen injection flow rate on flame structure and regression rate in the vortex hybrid rocket engines

• Transactions of the Korean Society of Mechanical Engineers B
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• v.34 no.4
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• pp.443-448
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• 2010

In this paper, the changes in performance parameters, e.g., the combustor pressure, turbine power, engine mixture ratio, temperature of gas generator, and product gas, of a liquid rocket engine employing gas generator cycle with the variations in propellant-supply pressure have been described. Engine performance is numerically calculated using the 13 major system-level variables of the rocket engine. The combustor pressure and turbine power increase with an increase in the oxidizer-supply pressure and decrease with an increase in fuel-supply pressure. The lower mixture ratio of gas generator for increased fuel mass flow rate decreases the gas generator gas temperature and deteriorates the gas material properties as the turbine working fluid. The turbine power decreases with an increase in fuel-supply pressure; this results in a decrease in the main-combustor pressure, which is directly proportional to engine thrust.

• Aerospace Engineering and Technology
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• v.6 no.2
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• pp.134-140
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• 2007

Since the successful launch of Sputnik 1, a rocket engine was evolved rapidly. The first artificial satellite Sputnik has only 182 lb mass with a size of a basket ball, a modern artificial satellite is over 10 tons. As the size and the mass of an artificial satellite increases, the stronger launch vehicles are required. However, the story is different in the field of the rocket engine development. In the early to mid age of the space race, rocket engine study was focused on the stronger and bigger engine development, but from the 80's the tide has changed. A rocket engine must be strong and also economic. This trend was accelerated from when a rocket launch was used commercially. In this study, a capability of the launch vehicle and engine was investigated to provide a reference for a liquid rocket engine development plan.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.484-488
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• 2012

Korea Aerospace Research Institute (KARI) performed the conceptual design of rocket engine test facility for the development and qualification of the 3rd stage liquid rocket engine for KSLV-II. The 3rd stage rocket engine test facility, which are to be constructed at Naro Space Center, will supply propellants and high-pressure gases to engine for firing test at ground and altitude conditions. The altitude test condition is obtained using a supersonic diffuser operated by the self-ejecting jet from the liquid rocket engine.

• International Journal of Aerospace System Engineering
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• v.5 no.2
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• pp.16-22
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• 2018

An analysis has been made on the performance variation due to pressure drop change at propellant supply pipes of liquid rocket engine. The objective is to compare the effectiveness of control variables to tune the liquid rocket engine performance. The mode analysis program has been used to estimate the engine performance for different modes which is realized by controlling the flow rate of propellant. The oxidizer of combustion chamber, the fuel of combustion chamber, the oxidizer of gas generator and the fuel of gas generator are the independent variables to control engine thrust, engine mixture ratio and temperature of gas generator product gas. The analysis program is validated by comparing with the powerpack test results. The error range of compared variables is order of 4%. After comparison of tuning effectiveness it is turned out that the pressure drop at oxidizer pipe of gas generator and pressure drop at combustion chamber fuel pipe and the pressure drop at the fuel pipe of gas generator can effectively tune the thrust of engine, mixture ratio of engine and temperature of product gas from gas generator respectively.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.5
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• pp.125-131
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• 2018

Some propellants in a liquid rocket engine are burned in the pre-burner of a staged combustion cycle engine, resulting hot gas drives the turbine. The burned gas passing through the turbine is supplied to the combustor at high temperature and pressure. The form of the gas can be fuel rich or oxidizer rich dependent upon the mixture ratio or the engine scheme. When the cycle works at oxidizer-rich condition, the metal pipes composing the engine can be ignited or even exploded by an impact of very a small particle. In this study, we developed the powder combination and processes for an anti-oxidation coating through the analysis of various coating materials.

• Journal of Aerospace System Engineering
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• v.3 no.2
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• pp.20-23
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• 2009

The structural analysis of liquid rocket engine was performed in the case of sinusoidal vibration load to verify structural safety. The finite element model is composed with main liquid rocket engine components, combustion chamber, turbopump, gas-generator, pyro-starter, main pipes, main valve, heat-exchanger, gimbal-mount and brackets. Natural vibration mode analysis and structural analysis for sinusoidal vibration load were performed. The natural mode frequency of liquid rocket engine is twice than that of launch vehicle. In the case of stress result of sinusoidal vibration load, the part of maximum stress has 1.4 margin, so the engine structure is safe for sinusoidal vibration load.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.4
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• pp.19-27
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• 2021

In this study, a high-pressure subsacle thrust chamber was developed to verify the core technology for KSLV-II performance enhancement. The core technologies are the design of an injector for high-pressure combustion, development of a combustion stabilization device using the additive manufacturing technique, and the design and fabrication of mixing head and regeneratively cooled combustion chamber. The core technologies, which have been verified through the development of high-pressure subscale thrust chamber, will be used to develop large engine liquid rocket engine thrust chamber in the future.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.581-585
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• 2009

In this study, conceptual design of the RBCC (Rocket Based Combined Cycle) engine was performed for the hypersonic propulsion system development. For the flight mission, RBCC engine takes off at sea level and accelerates up to Mach 8 at the altitude of 30km. By the flight speed characteristics, operating pattern of the engine is categorized into 3 modes : Ejector jet (~Mach 3), Ramjet (Mach 3~6), Scramjet (Mach 6~8). According to the engine mode characteristics, RBCC engine design and analysis was performed.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.1
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• pp.39-46
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• 2002

It is very difficult to measure an exact thrust in liquid rocket engine compared to solid rocket motor, however it is very important to estimate a performance of engine for developing rockets. To get a good result, we have to concern about errors of measurement and find a method of calibration. In this research, we developed new thrust measurement system for liquid rocket engine.