• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.12-15
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• 2011

The most widely used kerosene-LOx liquid-propellant rocket engines in these days have a similar engine schematic to those of the past because of the development cost and the reliability. The efficiency of engines could be increased by the factors such as a cooling method, engine cycles, shape of cooling channels, additional coolant and so on. In this article, it is described that some design ideas for performance enhancement by exchange kerosene with LOx of a coolant.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.4
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• pp.71-80
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• 2010

The KSLV-II(Korea Space Launch Vehicle-II) which being a successor of the KSLV-I is a space launch vehicle capable of delivering 1.5 ton-class satellite into a low earth orbit. The development of a 75 tonf-class liquid rocket engine(LRE) is planned on the basis of the technologies mastered through the preceded research of a 30 tonf-class LRE. The thrust chamber of the LRE is required to have higher combustion stability, structural integrity and thermal durability. This paper deals with the design requirements of the 75-tonf thrust chamber and a variety of technical considerations which have been conducted analytically in the course of the design for the realization of the requirements.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.2
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• pp.52-60
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• 2020

This paper deals with an optimal output tracking control for open-cycle liquid propellant rocket engine. For this purpose, we modeled simplified mathematical model of open-cycle liquid propellant rocket engine and designed optimal output feedback control system using combustion chamber pressure. For design the closed-loop system of open-cycle liquid propellant rocket engine, we designed optimal output feedback linear quadratic tracking control system using the linearized model and demonstrated the performance of the controller through numerical simulation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.750-755
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• 2011

The ignition characteristics of liquid rocket engine thrust chambers which have been developed by domestic technology were analyzed. Analysis results showed that low frequency fluctuation appeared in a partial ignition range according to different temperature profiles and vaporous state in the oxidizer manifold with startup sequences. This low frequency fluctuation wasn't developed as a malfunction factor, but this fluctuation is thought to be taken a continuous concern considering interfaces with engine system and launch vehicle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.321-324
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• 2011

A computational analysis of nozzle flow characteristics and plume structure using Reynolds-averaged Navier-Stokes equations with $k-{\omega}$ SST turbulence model was conducted to examine performance of the supersonic nozzle employed in a small liquid-rocket engine for ground firing test. Computed results and experimental outcome of 2-D converging-diverging nozzle flow were compared for verifying the computational capability as well as the turbulence model validity. Numerical computations of 2-D axisymmetric nozzle flow was carried out with the selected model. As a result, flow separation with backflow appeared around the nozzle exit. This investigation was reported as a background data for the optimal nozzle design of small liquid-propellant rocket engine for ground test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.29-32
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• 2009

This study is for development combustion test facility of liquid rocket engine system using hydrogen peroxide/kerosene as propellent. For this new facility, we construct thrust measure system, propellent supply system, control and data acquisition system. To perform 200N liquid rocket engine combustion test, operation scenario and sequence were designed. Result of combustion test propellents were supplied to engine stably and confirm of development combustion test facility very well.

• Journal of Aerospace System Engineering
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• v.14 no.5
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• pp.73-82
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• 2020

Recently, there has been an increase in additive manufacturing for the fabrication of liquid rocket engines. This technology can innovate conventional fabrication methods to reduce the lead time and manufacturing cost and can enhance the performances such as weight reduction. In this study, a literature survey is presented that includes types, advantages, disadvantages, and foreign government-based projects of the technology related to liquid rocket engine manufacturing. The present survey focuses on the technology that has been applied to various components such as turbopumps and valves while much larger efforts are made for combustion chambers with regenerative cooling channels and diverging nozzles, as the advantages of the technology are maximized for the applications.

• 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.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.3
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• pp.66-73
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• 2004

As a rocket propellent of hydrocarbon fuels, the characteristics of liquefied natural gas was evaluated with the viewpoint of the constituents and content, the cooling performance as a coolant, and characteristic velocity and specific impulse as parameters of the engine performance. Content of methane was a principal factor to determine the characteristics as a rocket propellant and more than 90% of it was needed as a fuel and coolant in the regenerative cooled liquid rocket engine. Some constituents of the liquefied natural gas can be frozen by the pre-cooling of the pipe lines, therefore they can be a factor disturbing the normal working of engine. In case the content of methane is around 90% in the liquefied natural gas, a normalized stoichiometric O/F mixture ratio of 0.75 is suggested for a nominal operation condition to get the maximum specific impulse and characteristic velocity.