• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.245-250
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• 2004

The objective of the present study is to conduct model combustion tests for various injectors to identify their combustion stability characteristics. Three different double swirl coaxial injectors with variation of a recess number have been tested for the comparative study of stability characteristic and flame structure. Gaseous oxygen and mixture of gaseous methane and propane have been employed for simulating actual propellants used for a fullscale thrust chamber. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.85-92
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• 2020

Ground hot-firing tests were conducted to analyze the combustion performance according to the characteristic lengths 1.37 m, 1.71 m, and 2.06 m of the combustion chamber in 200 N-class GCH4-LOx small rocket engine. Thrust, specific impulse, and characteristic velocity at the steady-state could be obtained as the key performance parameters of the rocket engine. The performance characteristics acquired through the test were compared and analyzed with the theoretical performance calculated from CEA analysis. Observation of the influence of characteristic length on the combustion performance indicates that an optimal characteristic length shall remain between 1.71 m and 2.06 m.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.82-85
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• 2009

In the development of Liquid Rocket Engine(LRE) systems, it is essential to understand the spray characteristics which influence mainly the performance and the stability of combustion. The injectors for this study have a recessed Liquid-swirl/Gas-centered jet coaxial type. For the similarity with actual conditions, the experimental conditions are calculated by using the momentum ratio as a matching parameter, and the stimulants of fuel and oxidizer are gaseous nitrogen and water respectively. The spray fields were measured by means of a photographic technique. Moreover, an effect of the momentum ratio has been investigated.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.3
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• pp.26-33
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• 2014

The gas-centerd swirl injectors are widely used on the main combustor of large liquid propellant rocket engines. Since the gas-liquid propellants, such as kerosene and oxygen-rich gas combination, are mixed and burned in the high pressure condition over the critical pressure point, the cold-flow spray test in the atmospheric condition can not represent the actual spray pattern. To observe the near actual spray patterns of gas-centered swirl injector, the high pressure spray chamber and the control system were constructed. The operating sequence was controlled precisely to obtain clear visualization images.

• Journal of Energy Engineering
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• v.6 no.2
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• pp.203-211
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• 1997

This study is on the separation of Global warming effect gas, CO$_2$by chemical absorption from mixture of CO$_2$-N$_2$which was modeled after flue gas of fire power plant. Investigation of optimum condition for absorbent was carried out by using sparged vessel apparatus. Through packed tower experiments, applicabilities of two absorption models were tested by comparing experimental results with theoretical values. Absorbent used in the experiments was Monoethanolamine (MEA) and gas mixture was made in the mole composition of 15% CO$_2$and 85% N$_2$. Through estimations of CO$_2$loading and CO$_2$removal efficiency, optimum concentration of absorbent was found in the range of 4-5 M. To find a rate of absorption, an enhancement factor was introduced. Values of rate of absorption were calculated by Film model and Higbie model, respectively. Higbie model showed good agreement with experimental results. Therefore, this models is considered to be applicable to the CO$_2$separation process for flue gas from fire power plant.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.4
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• pp.27-32
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• 2003

The Performance of micro combustor in various condition was exploited experimentally. Various geometric conditions of combustor were considered to figure out the performance of micro combustor. The micro combustor studied in this study was constant volume with cylindrical shape. Geometric parameters of combustor were defined to be combustor height and diameter. The effect of height was exploited parametrically with the size of 1mm, 2mm and 3mm. The effect of diameter was observed parameterized with 7.5mm and 15mm. Three different combustibles or Stoichiometric mixture of methane/air, hydrogen/air were used. Pressure transition during combustion process was recorded. The maximum pressure by combustion responded favorably with the change of height of combustor and the initial pressure. The flame propagation was visulized using Schlieren method. The flame propagation within combustor was observed when specific conditions such as combustor height and initial pressure over critical value was satisfied.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.1
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• pp.12-20
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• 2002

An integrated analysis of kerosine/LOX based KSR-III rocket body/plume flowfield has been performed. The analysis has been executed employing three kind of gas thermo-chemical models including calorically perfect gas, multiple species chemically reacting gas, and chemically frozen gas models and their effect on rocket flowfield has been accessed to provide the most appropriate gas thermo-chemical model which meets a specific purpose of performing rocket body and plume analysis. The finite-rate chemically reacting flow solution exhibited higher temperature throughout the flowfield than other gas models due to the increased combustion gas temperature caused by the chemical reactions within the nozzle. All the reactions were dominated only in the shear layer and behind the barrel shock reflection region where the gas temperature is high and the effect of finite-rate chemical reactions on the flowfield was found to be minor. However, the present plume computation including finite-rate chemical reactions revealed major reactions occurring in the plume and their reaction mechanisms and as well.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.4
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• pp.61-66
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• 2004

The objective of the present study is to develop methodology for the assessment of combustion stability of liquid rocket injectors. To simulate actual combustion occurring inside of a thrust chamber, a fullscale injector has been employed in the study, which bums gaseous oxygen and mixture of methane and propane. The main idea of the experiment is that the mixing mechanism is considered as a dominant factor significantly affecting combustion instability in a fullscale thrust chamber. A single split triplet injector has been used with an open-end cylindrical combustion chamber. The characteristics revealed by excited dynamic pressures in gaseous combustion show degrees of relative acoustic damping depending on operating conditions. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.5-8
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• 2010

The GCSC injectors studied in this paper are those applied to the combustion chamber of staged combustion engines. Liquid fuel is injected through tangential holes along the outer wall of the GCSC injector forming a swirling sheet and oxygen rich gas generated by a preburner enters axially through the center orifice of the injector to form a gaseous jet. The spray characteristics of GCSC injectors under ambient/high pressure conditions and the effect of recess on spray characteristics have been examined in this paper. These results are expected to be used as fundamental data to develop of a staged combustion engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.692-693
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• 2010

In this study, we analyze the combustion characteristic of GOx/PE hybrid rocket, by changing port diameter of the propellant. To study combustion characteristics of hybrid rocket, GOx is used for oxidizer and PolyEthylene is used for solid propellant. Regression rate and O/F ratio of the rocket is measured and presented.