• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.700-710
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• 2008

A Bubble size distribution model has been developed for the numerical simulation of cryogenic high-speed cavitating flow of the turbo-pumps in the liquid fuel rocket engine. The new model is based on the previous one proposed by the authors, in which the bubble number density was solved as a function of bubble size at each grid point of the calculation domain by means of Eulerian framework with respect to the bubble size coordinate. In the previous model, the growth/decay of bubbles due to pressure difference between bubble and liquid was solved exactly based on Rayleigh-Plesset equation. However, the unsteady heat transfer between liquid and bubble, which controls the evaporation/condensation rate, was approximated by a theoretical solution of unsteady heat conduction under a constant temperature difference. In the present study, the unsteady temperature field in the liquid around a bubble is also solved exactly in order to establish an accurate and efficient numerical simulation code for cavitating flows. The growth/decay of a single bubble and growth of bubbles with nucleation were successfully simulated by the proposed model.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.196-199
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• 2011

Subscale high-pressure injection system which use two-stage light gas gun composed with high-pressure tube, pump tube and launch tube can make supersonic liquid jet. The supersonic liquid jet enhances droplet atomization by shockwave in front of the jet. In this study, the experiments was executed to identify the atomization characteristics of the supersonic liquid jet using straight cone nozzle. SMD which presents the atomization characteristics was decreased from $151.2{\mu}m$ to $52.25{\mu}m$ by increasing of L/d.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.180-186
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• 2012

A water-flow test for acceptance verification is carried out for a nonimpinging-type injector prior to the design-performance verification of 70 N-class liquid-rocket engine under development. It is observed that there exist varying characteristics of atomization among the injector-orifices caused by a fabrication crudeness of orifice holes which can be judged from a microscopic standpoint. The flow shedding phenomenon and ruffle on the surface of liquid column (or droplet) could be caught from the instantaneous spray images.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.498-501
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• 2012

The Model Rocket Engine Test Facility has been improved to develop the Korea Space Launch Vehicle II(KSLV-II). The modified Model Rocket Engine Test Facility will be used to develop 7-tonf class liquid rocket engine combustor. The test result and test technique acquired from this facility will be used to develop the high performance liquid rocket engine combustor. This paper describes the modified Model Rocket Engine Test Facility for a Sub-scale model test of the 7-tonf class combustor.

• 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 of Propulsion Engineers
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• v.24 no.6
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• pp.126-142
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• 2020

This study aims to provide basic design data for a pintle injector and its combustion device through case study on the research and development of combustion devices to which a liquid-liquid pintle injector was applied. From data analysis, it was possible to provide the initial dimension of the combustion chamber and pintle injector based on the engine thrust, and the geometric characteristics of the high-efficiency injector. In addition, the pintle tip heat damage prevention mechanism and materials, face-shutoff pintle injector implementation method, and central propellant selection criteria were summarized. Theses results will be used as basic data for the design criteria of an initial pintle injector combustion device.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.1
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• pp.68-76
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• 2022

Upper stage of the Korea Space Launch Vehicle-II(KSLV-II) uses a 7-tons class liquid rocket engine and is an open gas generator cycle with a turbopump supply method that uses kerosene/liquid oxygen as the propellant combination. This study first provided a brief overview of the design and development process of the upper stage engine. In addition, it introduced the solutions and results applied to some of the problems that occurred during the development process of the upper stage engine.

• Journal of the Korean Society of Propulsion Engineers
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• v.27 no.1
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• pp.27-36
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• 2023

To understand the atomization performance in gas-liquid swirl-coaxial injector applied to a small rocket engine, a cold-flow test was performed by varying the design parameters and supply condition of propellants. As the swirl-chamber diameter and the angle of the convergent section, which are design parameters of injector increased, the spray performance of the injector improved by increasing the swirl strength. In addition, as the gas-liquid momentum-flux ratio increased, the gas flow separated some of the droplets from the liquid film, and a gas-droplet mixture core was formed in the center of the spray sheet.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.2
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• pp.46-53
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• 2010

The purpose of this paper is to study combustion characteristics with the different phase of oxidizer in hybrid rocket combustion. HDPE(High Density Polyethylene) as fuel and $GN_2O$(Gas $N_2O$), $LN_2O$(Liquid $N_2O$) as oxidizer were used to perform the experiments. An investigation was performed for a change of the regression rate, pressure of combustion chamber and combustion efficiency according to the variation of oxidizer phase. In case of using $LN_2O$ as oxidizer, the regression rate is not significantly different from using $GN_2O$ as oxidizer. It is considered that combustion energy is much larger than latent heat energy which was used in the evaporation of liquid oxidizer. However propulsion performance efficiency for $LN_2O$ showed lower value than for $GN_2O$. By increasing the flow rate of liquid oxidizer, heat transfer needed for vaporization of liquid oxidizer was increased, which resulted in the growth of combustion instability.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.5
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• pp.37-50
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• 2007

This study investigated the characteristics of spray generated by a liquid coaxial swirl injector used in a combustor of the liquid rocket engine. The linear stability analysis considered long and short wave was introduced in liquid sheet breakup. Through the hydrodynamic analysis the initial liquid sheet thickness spray angle and injection velocity were predicted. To evaluate the effect of turbulence model standard $k-{\varepsilon}$ and RNC $k-{\varepsilon}$ model were applied to numerical calculation and it was known that RNC $k-{\varepsilon}$ model was more applicable to predict spray characteristics. On the basis of this evaluation validation of the developed model was performed with swirl injector installed in LPRE and the predicted results of breakup length, spray angle, and SMD agreed well with experiments qualitatively and quantitatively.