• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.4
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• pp.46-51
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• 2007

The irregular fuel surface was observed by the visualization of hybrid rocket combustion. Even though the test condition maintained oxidizer rich environment, the irregular dark fuel surface was formed as the result of incomplete combustion. In order to investigate the correlation of the characteristics of oxidizer flow and the irregular fuel surface, various flow conditions were imposed such as swirl flow, induced swirl flow by helical fuel configuration and straight flow. Test results revealed no correlation was found between oxidizer flow condition and irregular fuel surface. And this can be a commonly observed phenomena in the tests with different fuel/oxidizer combination. Thus, the irregular fuel surface can be a result of the interaction of blowing flow of vaporized fuel and the boundary layer of oxidizer flow. A further study will be required to confirm the assumption for the formation of irregular fuel surface.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.4
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• pp.52-58
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• 2007

The irregular fuel surface was observed by the visualization of hybrid rocket combustion. Even though the test condition maintained oxidizer rich environment, the irregular dark fuel surface was formed as the result of incomplete combustion. In order to investigate the correlation of the characteristics of oxidizer flow and the irregular fuel surface, various flow conditions were imposed such as swirl flow, induced swirl flow by helical fuel configuration and straight flow. Test results revealed no correlation was found between oxidizer flow condition and irregular fuel surface. And this can be a commonly observed phenomena in the tests with different fuel/oxidizer combination. Thus, the irregular fuel surface can be a result of the interaction of blowing flow of vaporized fuel and the boundary layer of oxidizer flow. A further study will be required to confirm the assumption for the formation of irregular fuel surface.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.375-378
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• 2011

The combustion characteristics of fuel-rich gas generator which could be adopted to Ducted Rocket propulsion system is investigated. The gas generator is designed considering the design requirements of propulsion system then solid fuel, which is developed by reducing the contents of oxidizer and increasing the contents of metallic solid particle, is loaded in the manufactured gas generator. The results of combustion test shows the necessity of the special analysis tool for estimating the gas generator performance where multi-phase flow of fuel-rich gas exists.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.4
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• pp.79-88
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• 2017

Ignition test of the fuel-rich propellant coated with ignition support material in the ramjet combustor condition was conducted. Ignition delay and flame holding was measured. Fuel grain consist of HTPB mixed with AP particle 15 wt.%, Al particle 5 wt.%. To cause the short ignition delay, ignition support consist of $NC/BKNO_3$ and composite propellant was coated to the fuel grain. Ethanol blended $H_2O_2$ gas generator control the temperature, pressure, $O_2$ concentration in the oxidizer gas in the air. Gas is supplied with mass flux of $200kg/m^2s$. Through the test ignition support operated well and ignition delay of 0.6 second and the Flame was sustained.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.8
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• pp.675-685
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• 2016

A research area on liquid rocket engine has been suggested. Downsizing through combustion pressure rise and low price are major issues to gas generator cycle engines. A very high pressure turbopump and material against oxidizer rich environment may be necessary technologies for staged combustion cycle engines. Integrated analysis saving computing time is the trend of rocket engine systems analysis area. Other important research topics are the methane engine for reusable booster to reduce the cost, 3D printing and materials for high temperature or oxidizer rich environment.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.44-50
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• 1999

For the prediction of combustion performance of the Unlike Impinging Quadlet Injector (OOOF type), mixing efficiency, mixing characteristic velocity, and efficiency of mixing characteristic velocity were obtained from the cold test. Water/kerosene were used for simulants, The momentum ratio of oxidizer and fuel were mixing correlating parameter. Orifice discharge coefficient, spray pattern and mass distribution were measured. As a result, invasion-depth had strong effect on mixing efficiency, mixing characteristic velocity, and efficiency of mixing characteristic velocity. Mixing efficiency and efficiency of mixing characteristic velocity showed maximum value for momentum ratio 1.67(TMR = 2.5), and fuel rich state showed larger decreasing ratio than oxidizer rich state.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.2
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• pp.118-125
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• 2019

Metals exposed to high temperature/high pressure/oxidant excess environment of an oxygen excess pre-combustor may undergo rapid oxidation. In this study, the test facility to simulate the high temperature/high pressure/oxidant excess environment was constructed and the oxidation resistance evaluation was carried out for various metal materials. As a result, the discoloration of the metallic materials, the change in the surface roughness and the peeling of the metal surface were observed, and the weight change was also observed. The resulst showed that oxidation-resistant coating of a metal material of the combustor is indispensably required, and the use of XM-19, which has the highest content of Cr and Ni, is expected to provide more structural stability.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.7
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• pp.89-96
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• 2006

The combustion temperature in gas generator should be kept below around 1,000K to avoid any possible thermal damages to turbine blade by adopting either fuel rich or oxidizer rich combustion. Thus, non-equilibrium chemical reaction dominates in the gas generator. Meanwhile, Kerosene is a compounded fuel mixed with various types of hydrocarbon elements and difficult to model the chemical kinetics. This study focus to model the non-equilibrium chemical reaction of kerosene/LOX with detailed kinetics developed by Dagaut using PSR(Perfectly stirred reactor) assumption. Also, droplet evaporation time is taken into account by calculating for the residence time of droplet and by decoupling reaction temperature from the reactor temperature. In Dagaut’s surrogate model for kerosene, chemical kinetics of kerosene consists of 1592 reaction steps with 207 chemical species. The comparison of calculation results with experimental data could provide very reliable and accurate numbers in the prediction of combustion gas temperature, species fraction and other gas properties.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.7
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• pp.576-584
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• 2016

In this study, AP added propellant has been proposed as a method of enhancing the low specific impulse performance found for staged hybrid rocket engine. Experimental tests were carried out to analyze and evaluate the effect of AP added propellant on specific impulse performance as well as fuel-rich combustion characteristics in a staged hybrid rocket engine. Upper limit of AP content in propellant was set to be 15 wt% to maintain the hybrid rocket engine advantages. As a result, 15 wt% AP added propellant showed 3% higher specific impulse performance compared to 0 wt% AP added propellant. Moreover, AP addition proved to offer less injected oxidizer mass flow, less O/F variation, and less combustion pressure while producing fuel-rich gas of the same combustion temperature. Future studies will carry out more combustion tests with metal additives to further enhance specific impulse.