• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.41-47
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• 2006

Effects of mean flow and nozzle damping on acoustic tuning of a gas-liquid scheme coaxial injector are investigated numerically adopting a linear acoustic analysis. The injector plays a role as a half-wave acoustic resonator for acoustic damping in a combustion chamber of a liquid rocket engine. As Mach number of mean flow in a chamber increases, the resonant frequency of the first tangential mode decreases slightly and the optimum injector tuning length varies negligibly. Nozzle damping affects neither the resonant frequency nor the optimum length. From these numerical results, effects of mean flow and nozzle damping on acoustic tuning of a resonator are negligible. As open area of the injectors increases, the acoustic amplitude decreases, but new injector-coupled modes appear.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.20 no.10
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• pp.100-106
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• 2019

Ground and high altitude simulated combustion experiments were conducted using a liquid rocket engine with hydrogen peroxide and kerosene as the propellant. A ground and high altitude simulated combustion test facility was constructed by installing a high altitude model diffuser and TMS (Thrust Measuring System) on a vertical combustion test bench. The thrust characteristics according to altitude were investigated using the combustion test equipment. The diffuser was designed on a 1:4.8 scale to verify the characteristics of the high diffusing diffuser and starting pressure. The cold flow tests were conducted using nitrogen gas, and the performance characteristics and starting characteristics of the scale down diffuser were verified. A diffuser and TMS were installed on the vertical combustion test bench, and the thrust correction equations for the system resistance were derived. The thrust correction equations were derived from the step test and vacuum step test before the actual hot firing test. Nozzles with an operating altitude of 10km were designed. Hot firing tests were conducted to analyze the thrust characteristics according to the operating altitude changes. The actual thrust was calculated using each correction equation with the thrust value measured by the TMS.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.281-287
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• 2005

Nonlinear behaviors such as steep-fronted wave motions and a finite amplitude limit cycle often accompanying combustion instabilities have been numerically investigated using a characteristic-based approximate Riemann solver and the well-known ${\eta}-{\tau}$ model. A resonant pipe initially subjected to a harmonic pressure disturbance described the natural steepening process that leads to a shocked N-wave. For a linearly unstable regime, pressure oscillations reach a limit cycle which is independent of the characteristics of the initial disturbances and depends only on combustion parameters and operating conditions. For the 1.5 MW gas generator under development in KARI, the numerical results show good agreement with experimental data from hot-firing tests.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.6
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• pp.64-70
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• 2005

The overall results of hot firing tests of fuel-rich gas generator with impinging injector at design and off-design points are described. The gas generator consists of an injector head with impinging injector, a water cooled combustor wall, a turbulence ring to enhance mixing, an instrument ring measuring temperature and pressure and a nozzle. The combustion tests were successfully performed without damage of gas generator. Test results show that the outlet temperature is not dependent on residence time of hot gas within 4～6msec but dependent on chamber pressure. The relation between outlet temperature and combustion efficiency resulting from measured pressure, mass flow rate and area of nozzle throat is shown. The overall O/F ratio is the critical parameter to determine the outlet temperature and the linear correlation between two parameters is established.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.106-109
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• 2012

It was designed and tested ignition that an oxidizer rich preburner for a staged combustion cycle liquid rocket engine propelled by kerosene and LOx. Operation conditions of the preburner are about 60 of OF ratio and 20 MPa of combustion pressure. Ignition characteristics were compared by propellants flowrate. As the results, the higher propellants flowrate, the shorter the ignition delay time and the higher ignition stiffness. The ignition delay time was affected by incoming the oxidizer flowrate through the refrigerative cooling channels. The oxidizer flowrate from the cooling channels decreased by inflow of combustion gas during initial ignition. The oxidizer flowrate of the cooling channels increases, it is rapid recovery by cooling effect, eventually the ignition delay time decreases.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.2
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• pp.122-127
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• 2013

The paper includes the analytic results of pressure fluctuation data from the combustion of an oxidizer-rich preburner applicable to high-performance, closed-cycle liquid rocket engine systems. The combustion experiments went through two different steps of a chamber pressure during single run. Self-excited pressure fluctuations with a frequency of 78 Hz were observed only at the relatively low chamber pressure condition. These pressure fluctuations are regarded as a bulk mode. The intensity of pressure fluctuations by a root-mean-square value is 13.3% normalized by the chamber static pressure and no pressure excitation was observed at the design pressure condition. The bulk mode has an identical phase across the inside of the chamber and reveals the similar characteristics to the Helmholtz resonator.

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.3
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• pp.50-57
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• 2009

A gas mechanical pulsator is developed for the study of combustion instabilities in various combustors such as LRE combustor. First, it shows that the mass flow rates and the perturbation frequencies can be successively controlled by the inlet pressure and the rotating speed of a rotating disk with many holes. Second, the device is used as an acoustic amplification source as a substitute for the speaker in the previous acoustic tests and its results show almost the same resonant frequency and damping characteristics compared with the previous results. In conclusion, the result shows that it can be used as a substitute for a speaker in the studies of LRE combustion instabilities, which has a flow and no limitation of amplification, and a device for making a perturbation source in gas flow.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.6
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• pp.59-66
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• 2006

Acoustic cavity as a stabilization device to control high-frequency combustion instabilities in liquid rocket engine is adopted and its damping capacity is verified in atmospheric temperature. First, harmful resonant frequency in a modeling chamber can be damped effectively by the installation of properly-tuned acoustic cavity. Besides, geometric effects of acoustic cavity on damping characteristics are analyzed and compared quantitatively. Satisfactory agreements have been achieved with linear acoustic analysis and experimental approach. Results show that the acoustic cavity of the largest orifice area or the shortest orifice length was the most effective in acoustic damping of the harmful resonant frequency. Finally, it is proved that an optimal design process is indispensable for the effective control of combustion instabilities.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.291-298
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• 2005

Linear acoustic analysis has been performed to elucidate damping characteristics of acoustic cavities. Results have shown that resonant frequencies of acoustic cavity obtained by classical theoretic approach and present linear analysis are somewhat different from each other. This difference is due to the limitation of classical theory. To quantify the damping characteristics, acoustic impedance has been introduced and resultant absorption and conductance have been evaluated. Satisfactory agreement has been achieved with previous experiment. Finally the design procedure for optimal tuning of acoustic cavity has been established

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.3
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• pp.1-9
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• 2004

We conducted the firing test with the regenerative cooling LRE and calculated the heat flux from measured coolant temperature, that was compared with the heat flux predicted by previously developed numerical analysis method. The difference between the measured heat flux and the numerical calculation value was within nine percents. Therefore, developed numerical analysis method can be applied to the design/fabrication of a real LRE system. and, it was investigated that combustion pressure and mixture ratio have an Influence on the heat flux with a constant relation.