• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.1
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• pp.1-8
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• 2005

An axisymmetric compressible flow solver accounting for chemical equilibrium has been developed as an analysis tool exclusively suitable for performance prediction and optimum contour design of LRE thrust chamber. By virtue of several features focusing on user-friendliness and effectiveness including automatical grid generation and iterative calculations with changes in design parameters prescribed through only one keyword-type input file, a design engineer can evaluate very fast and easily the influences of various design inputs such as geometrical parameters and operating conditions on propulsive performance. Validations have been carried out for various aspects by detailed comparisons with the result of CEA code, experimental data of JPL nozzle, actual data for two historical engines, and ReTF data for KSR-III.

• International Journal of Aerospace System Engineering
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• v.5 no.2
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• pp.16-22
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• 2018

An analysis has been made on the performance variation due to pressure drop change at propellant supply pipes of liquid rocket engine. The objective is to compare the effectiveness of control variables to tune the liquid rocket engine performance. The mode analysis program has been used to estimate the engine performance for different modes which is realized by controlling the flow rate of propellant. The oxidizer of combustion chamber, the fuel of combustion chamber, the oxidizer of gas generator and the fuel of gas generator are the independent variables to control engine thrust, engine mixture ratio and temperature of gas generator product gas. The analysis program is validated by comparing with the powerpack test results. The error range of compared variables is order of 4%. After comparison of tuning effectiveness it is turned out that the pressure drop at oxidizer pipe of gas generator and pressure drop at combustion chamber fuel pipe and the pressure drop at the fuel pipe of gas generator can effectively tune the thrust of engine, mixture ratio of engine and temperature of product gas from gas generator respectively.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.15 no.6
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• pp.3403-3409
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• 2014

Hot-firing tests were carried out using a mixing head with 19 swirl coaxial injectors and a combustion chamber with internal cooling channels. The propellants of liquid oxygen and kerosene(Jet A-1) were burned in a range of chamber pressures (59~82 bar) and mixture ratios (2.0~3.0). The temperature of water used as the cooling fluid was measured at the inlet and outlet of the cooling channels, and the heat flux was calculated. The aim of this study was to examine the effect of combustion instability on heat transfer in a subscale thrust chamber, and detect the temperature variation of cooling water. During several hot-firing tests, combustion instability was encountered which caused a 5~20% increase in heat flux. The peak heat flux took place in the initial stages of combustion instability.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.5
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• pp.92-100
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• 2010

Combustion stability rating tests of 75-tonf-class thrust chamber for technology demonstration were carried out at a low pressure. Two kinds of mixing heads were used in this study. One of them has injectors of 631 and the other has injectors of 721. Mixing head with injectors of 631 showed a self-oscillation instability at the chamber pressure of 30 bar. Mixing head with injectors of 721 showed that a high frequency combustion stability was maintained under the same pressure and the same mass flow rate. But mixing head with injectors of 721 generated a self-oscillation instability at the chamber pressure of 20 bar and it was found that stability boundary region was changed due to the configuration of a mixing head from these results.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.203-207
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the thrust chamber of liquid rocket using LOx and Kerosene as propellant. The heat fluxes were obtained from the measured wall temperature to the axial direction of thrust chamber for different type of coolant, the various O/F ratio, mass flow rate and the location of the film cooling injector. A thin wall combustion chamber and nozzle were used to obtain the heat flux.

• 한국연소학회:학술대회논문집
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• 2003.12a
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• pp.145-152
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• 2003

The objective of the present study is to develop methodology for the assesment of combustion stability of liquid rocket injectors. To simulate actual combustion occurring inside of a thrust chamber, a full-scale injector has been employed in the study, which burns 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 full-scale thrust chamber. Single & multi split triplet injectors have 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.

• 한국연소학회:학술대회논문집
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• 2003.12a
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• pp.131-136
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• 2003

Experimental study on combustion characteristics of double swirl coaxial injectors has been conducted for the assessment of critical design parameters of injectors. A subscale thrust chamber has been fabricated with a water-cooled copper nozzle, which allows a chamber to be reused without replacing parts. Two different designs of injectors have been tested for the understanding of the effects of recess length on combustion. Clearly, the recess length drastically affects the combustion efficiency and hydraulic characteristics of the injector. Internal mixing of propellants in the injector with the recess number of two increases a combustion efficiency and reveals sound combustion although a pressure drop required for the similar amount of mass flow rates increases compared with the injector of the recess number of one.

• Journal of Mechanical Science and Technology
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• v.18 no.7
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• pp.1246-1257
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• 2004

The objective of this paper is to characterize dynamic pressure traces measured at self-excited combustion instabilities occurring in two combustion systems of different hardware. One system is a model lean premixed gas turbine combustor and the other a fullscale bipropellant liquid rocket thrust chamber. It is commonly observed in both systems that low frequency waves at around 300㎐ are first excited at the onset of combustion instabilities and after a short duration, the instability mode becomes coupled to the resonant acoustic modes of the combustion chamber, the first longitudinal mode for the lean premixed combustor and the first tangential mode for the rocket thrust chamber. Low frequency waves seem to get excited at first since flame shows the higher heat release response on the lower frequency perturbations with the smaller phase differences between heat release and pressure fluctuations. Nonlinear time series analysis of pressure traces reveals that even stable combustion might have chaotic behavior with the positive maximum Lyapunov exponent. Also, pressure fluctuations under combustion instabilities reach a limit cycle or quasi-periodic oscillations at the very similar run conditions, which manifest that a self-excited high frequency instability has strong nonlinear characteristics.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.229-234
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• 2004

The objective of the present study is to conduct model combustion tests for double swirl coaxial injectors to identify their combustion stability characteristics. Gaseous oxygen and mixture of methane and propane have been used as simulant propellants. Two model chambers tuned to the If acoustic resonance mode of a full-scale thrust chamber were manufactured to be used as a combustion cylinder. The main idea of the experiment is that the mixing mechanism is considered as a dominant factor significantly affecting combustion instability in a full-scale thrust chamber. Self-excited dynamic pressure values in a model chamber show different combustion stability zones with respect to a recess number. Upon test results, couplings between combustion conditions and the IT acoustic resonance mode become strengthened with the increase of a recess length.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.31-34
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• 2011

Design and fabrication of Technology Demonstration Model(TDM) of 75 tonf regenerative cooling thrust chamber were described. It has design chamber pressure of 60 bar, propellant mass flow rate of 243.6 kg/s, and nozzle expansion ratio of 12. It has a single welded structure of the mixing head and the chamber. Design and fabrication technologies established through this TDM can be used to development of flight model.