• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.517-521
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• 2009

A unsteady internal ballistic analysis model was proposed to take account for the variance of local variance of pressure and velocity along the grain surface of a hybrid rocket combustor. The model of concern in the study was fairly comparable with the test result. The local variation of the oxidizer mass flow rate along grain surface results in chamber pressure, regression rate, and gas velocity change along its flow direction.

• Journal of Advanced Marine Engineering and Technology
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• v.18 no.1
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• pp.81-91
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• 1994

Combustion noise involved with chemical heat release and turbulent process in turbopropulsion systems, gasturbine, industrial furnaces and internal engines is indeed noisy. The experimental study reported in this paper is made to identify a dominant combustion noise in jet flames. Gaseous propane and kerosene fuel have been used with air as the oxidizer in a different jet combustion systems. Combustion and aerodynamic noise are studied through far field sound pressure measurements in an anechoic chamber. And also mean temperature and velocities and turbulent intensities of both isothermal and reacting flow fields were measured. It is shown that axial mean velocity of reacting flow fields is higher about 1 to 3m/sec than that of cold flow in a gaseous combustor. As the gaseous fuel flow rate increases, the acoustic power increases. But the sound pressure level for the spray flame decreases with increasing equivalence ratio. The influence of temperature in the combustion fields due to chemical heat release has been observed to be a dominant noise source in the spray flame. The spectra of combustion noise in gaseous propane and kerosene jet flame show a predominantly low frequency and a broadband nature as compared with the noise characteristics in an isothermal air jet.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.23 no.4
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• pp.30-35
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• 2015

In this paper, a theoretical study of low frequency non acoustic instability, the $L^*$ instability, of a solid rocket motor is investigated. The $L^*$ stability criterion is determined by analysing the $L^*$ stability curves of two very distinct propellants for five different geometrical combustors. The $L^*$ instability of two extreme fuels showed totally different behavior in terms of operating pressure of the combustor. A parametric study on the stability for different chamber volume and different throat area keeping constant $L^*$ is conducted and analyzed. It was found that one of the main parameters, the non-dimensional critical characteristic time, requires an enough margin from the critical $L^*$ stability curve.

• 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.

• Transactions of the Korean hydrogen and new energy society
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• v.29 no.5
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• pp.538-548
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• 2018

The thermo-acoustic instability in the combustion process of a gas turbine is caused by the interaction of the heat release mechanism and the pressure perturbation. These acoustic vibrations cause fatigue failure of the combustor and decrease the combustion efficiency. This study is to develop a segmented dynamic thermo-acoustic model to understand combustion instability of gas turbine. Therefore, this study required a dynamic analysis rather than static analysis, and developed a segmented model that can analyze the performance of the system over time using the Matlab/Simulink. The developed model can confirm the thermo-acoustic combustion instability and exhaust gas concentration in the combustion chamber according to the equivalent ratio change, and confirm the thermo-acoustic combustion instability for the inlet temperature and the load changes. As a result, segmented dynamic thermo-acoustic model has been developed to analyze combustion instability under the operating condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.180-185
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• 2004

Our previous study showed that although the hybrid rocket engine with swirling gaseous oxygen had high performance, a direct injection of LOX with swirl into the combustion chamber of the hybrid rocket engine lowered the performance of the engine, compared to that with gaseous oxygen. In order to clarify this reason, combustion tests of a small PMMA combustor with an inner port diameter of 2 mm were conducted in liquid oxygen flow by comparison with gaseous oxygen flow. Although the oxygen mass fluxes of LOX were about two orders of magnitude larger than those of gaseous oxygen, the fuel regression rate of LOX were remarkably smaller than those of gaseous oxygen. For both liquid and gaseous oxygen, diffusion flames in the port of the grain controlled the combustion process of PMMA in oxygen flow. These results may be explained by the fact that only small amount of LOX vaporized and consumed in the combustion with PMMA while flowing through the port due to relatively larger latent heat of injected liquid oxygen compared to the heat of release by combustion which depended on the burning surface area of PMMA.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.68-77
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• 2017

High-frequency combustion instability results from a feedback coupling between the unsteady heat release rate and the acoustic waves formed resonantly in the combustion chamber. It can be modeled as thermoacoustic problems with various degrees of the assumptions and simplifications. This paper presents numerical analysis of self-excited combustion instabilities in a variable-length lean-premixed combustor and designs of passive control devices such as baffle and acoustic resonators in a framework of 3-D FEM Helmholtz solver. Nonlinear behaviors such as steep-fronted shock waves and a finite amplitude limit cycle are also investigated with a compressible flow simulation technique.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1016-1020
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• 2017

This study was conducted to develop a test facility that simulates the combustion instability that occurs in a real-scale liquid rocket combustor. A separate engine head with 3 injectors arranged in a row was designed/manufactured and verified through preliminary tests. The flow rate and spray pattern of the head were confirmed by the cold flow test. Next, propellant spray test and combustion test were carried out. A preliminary combustion test was carried out at 10 bar and the combustion chamber pressure was measured to be significantly lower than the target pressure. This is because it was a low pressure test, and it is expected to be resolved in the high pressure test in the future.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.2
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• pp.53-59
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• 2019

LIBS plug, a simplified laser-induced breakdown spectroscopy(LIBS) device with the purpose of measuring the fuel distribution inside the combustion chamber, was developed and manufactured. The LIBS plug receives only two wavelengths (H:656.3 nm, O: 777 nm) that are closely related to the equivalence ratio in the overall spectrum. The calibration curve between the signal of the LIBS plug and the equivalence ratio was constructed, and the fuel distribution of gasoline-air and LPG-air mixtures was measured using the LIBS plug.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.8
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• pp.720-725
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• 2007

Cold acoustic tests have been performed to elucidate the effect of baffle gaps on the optimal damping characteristics in a liquid rocket combustor where coaxial injectors are installed. For several axial baffle lengths, an optimal acoustic damping capacitance has been achieved in a certain gap range. Cold acoustic tests for simulating fluid viscosity by changing the pressure in a model chamber have been done to study the main mechanism of optimal damping. Experimental data have shown that the optimal gap for high damping capacity exists mainly due to the viscosity near the gap of baffles. Therefore, axial baffle length can be reduced by using the optimal baffle gap, providing a possible solution of thermal cooling problems. Also, these optimum characteristics can be some guidelines for manufacturing and assembling injectors in full-scaled rocket combustors.