• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.142-147
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• 2004

The flow in the LOX manifold of liquid rocket has been investigated using a CAE technique with an objective of economical modeling of injection holes in order to reduce the overall computational cost of flow analysis during the optimal rocket design procedure. The computational geometry is very close to that of the actual rocket design and the flow condition through the injection holes resembles that in the actual manifold of the liquid rocket. The result shows that the flow in the plane just above the injection holes is not uniformly distributed in terms of pressure and mass flow rate and this is attributed to the large-scale flow patterns present the LOX manifold. Thus, the flow physics should be understood correctly before making any attempt to model the injection holes. In the present study, several boundary conditions which were designed to effectively replace the presence of injection holes have been tested and it was found that a simple modeling can be possible by mimicking the actual geometry of the injection holes. By using this simple injection hole modeling, it was able to obtain about 30% reduction in computational cost but it was still able to reproduce the flow patterns correctly. Also the flow has been analyzed after incorporating a couple of different types of pre-distributors in LOX manifold and the effect of those will be discussed.

• 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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• v.y2005m4
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• pp.23-26
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• 2005

The method of decreasing the ecological risk for the LRE(Liquid Rocket Engine) test is developed, working on the cryogenic oxidizer and the high-boiling fuel(Kerosene). This Method is based on the study that contains a technical solution method and an accident occurrence range for decreasing of accident probability and damage. This paper contains schematic on the all risk circumstance bring to accident, block-diagram for an accident growth process in case of the propellant leakage, evaluation method of safety distance from test stand to residential area. It will be used to alternative method for the risk reduction of complex technical systems.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.213-219
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• 2005

Small liquid rocket engine (SLRE) with calorimeter were developed and tested to evaluate cooling characteristics in the liquid rocket engine. Therefore, cooling performance analysis was performed to predict the heat transfer coefficient on gas side wall in 10 calorimeter channel. A heat transfer empirical formula was determined by results of firing test and computational simulation.

• 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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• 2003.10a
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• pp.83-88
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• 2003

An analytical study was carried out to evaluate the regenerative cooling characteristics in the liquid rocket engine of a 10tf-thrust level using kerosene as a fuel. As a supplementary cooling method, a radiative cooling was applied to the nozzle extension. It was found out from this work that the cooling system with the regenerative and radiation cooling only is not adaptable as a cooling method for the liquid rocket engine of a 10tf-thrust level using kerosene as a fuel for the 2nd stage of the space launch vehicle. So, additional cooling method, curtain cooling was introduced and analyzed. Curtain cooling was very effective to reduce the thermal and thermo-structural instability.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.157-162
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• 2006

A Liquid Rocket Engine(LRE) ground firing test facility was built in Korea Aerospace Research Institute(KARI) in 2001 to develop the LRE for the first Korean liquid rocket, KSR-III. Around 170 tests were conducted since its establishment until recently by September 2006, and in the meantime, a considerable improvements were made in the capability. This paper describes the outline, conducted tests and operation techniques which have been accumulated through the operation of KARI LRE ground firing test facility.

• Aerospace Engineering and Technology
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• v.11 no.2
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• pp.182-194
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• 2012

To develop a liquid rocket engine with high reliability and safety under constraints of limited time and budget an optimal diagnosis system for the engine needs to be developed in parallel with the development of the engine. This paper is intended to set a development direction of the diagnosis system for the liquid rocket engine through the literature survey and addresses possible engine defects, characteristics of parameters for diagnosis and diagnostic methods including real-time diagnosis, post-test/post-flight diagnosis, fault detection method, parameter circuit method and test diagnosis. In addition tasks to be performed in the design and operation phases of the engine and foreign application case of engine diagnosis are presented.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.533-536
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• 2010

It has been manufactured the a prototype of the oxygen rich preburner which can be used for a small high performance liquid rocket engine. Based on the model designed and analyzed by KARI a series of the work was made. First, the layout was drawn and then two kinds of the test specimen were devised and manufactured for the hydraulic and pressure test respectively. It was confirmed that the details in designing and manufacturing the preburners. All of them efforts may help establish the bases to develop high performance liquid rocket engine.