• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.303-308
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• 2006

This paper describes numerical efforts to investigate combustion characteristics of HyShot scramjet engine. The corresponding altitude, angle of attack, and equivalence ratio are 28 km, $0^{\circ}$, and 0.426 respectively. $H_2$ and OH mass fraction show that the upstream recirculation zone of injector has flame-holding effects and main combustion begins at 15 cm downstream from cowl. Two-dimensional simulation reasonably predicts combustor inner pressure and also reveals periodic combustion characteristics of HyShot scramjet engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.145-148
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• 2009

Combustion tests of a fuel-rich gas generator had been conducted using the assembly of a powerpack. A gas generator is prone to longitudinal modes of combustion instabilities in a powerpack due to the increase of a characteristic length. It has been observed that the orifice inserted at the exit of the gas generator suppresses a longitudinal combustion instability. The intensities of pressure fluctuations in the manifolds and the chamber increase quadratically with a chamber pressure. Pressure fluctuations in the fuel manifold reveal two-fold strength greater than those in the oxygen manifold and the chamber. Frequency analysis indicates nonlinear characteristics inherent in the pressure fluctuations in the fuel manifold.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.4
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• pp.63-72
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• 2003

The combustion characteristics of the KSR-III engine were investigated numerically from the viewpoint of performance and combustion field. For numerical analysis of KSR-III engine with hub-and-spoke baffle, 3-D calculation was performed about $30^{\cire}$ section and the prediction of performance was in a good agreement with hot-firing test result. As a result of baffle installed, the performance of KSR-III engine was reduced in comparison with no baffle case and local high temperature region appeared on injector plate, combustion wall and baffle wall, This calculation was used practically as basic data for designing injector plate with film cooling holes and predicting the performance of KSR-III final flight test.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.6
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• pp.49-56
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• 2017

The current study has developed an in-house 3D FEM code in order to model thermoacoustic problems in an annular system and compared the acoustic field calculation results with measured ones from a benchmark combustor. From the comparison of calculation results with the measured data, the current acoustic code could successfully capture the various acoustic mode found in the annular system. In addition, it was found that the transverse waves in the combustor were strongly affected by the nozzle acoustic impedances, as well, the pressure distributions were closely related with the combustor acoustic pressure field.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.86-90
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• 2012

In this paper, the combustion instabilities which may occur in the hybrid rocket were studied. The rocket combustor where the vortexes can be generated was designed, and the experiments were performed. The investigations about characteristics on the presence of the diaphragm, the length of the fuel, the diameter of the fuel port, the diameter of the diaphragm, the diameter of the nozzle throat, and the variation of the Ox massflow rate were conducted. The main resonant frequency of the combustion pressure is regarded by the Vortex shedding mode, and it is considered that the other resonant frequency of the pressure fluctuation is hybrid low frequency, or helmholtz mode.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.4
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• pp.70-80
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• 2019

Combustion tests using six bi-swirl coaxial injectors with different shapes and recess lengths were performed in a model combustion chamber capable of flame visualization. By utilizing gaseous methane and gaseous oxygen instead of actual propellants, the effects of injector design and experimental conditions on the flame structure and combustion stability were analyzed. It was found that not only the experimental conditions but also the injector geometry such as the recess length and orifice diameter had a considerable influence on the combustion stability. In addition, it was confirmed that the heat release pattern changed with the occurrence of combustion instability.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2001.04a
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• pp.17-20
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• 2001

Combustion characteristics of KSR-III liquid rocket engine are investigated numerically in the standpoints of engine performance and acoustic instability. In the present calculation, engine performance for design and off-design conditions is estimated effectively with reasonable error. Numerical results of acoustic instability show that engine operation for the design condition has sufficient stability margin, but for a certain off-design condition, acoustic instability can be triggered by artificial pressure perturbation. The present results are in a good agreement with the available experimental results and can be adopted for the prediction of engine performance and stability, depending on the specific operating condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.3
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• pp.78-86
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• 2016

In this paper we propose a new frame transfer function model describing the variations of a heat release rate in response to an external flow oscillation in gas turbine systems. A critical difference of our model compared to the so-called $n-{\tau}$ model which has been widely used for a prediction of combustion instability (CI), is that our model is able to describe a nonlinear relation between phase and frequency. In contrast, the phase part of the $n-{\tau}$ model is a pure time delay and thus the phase should be a linear function of frequency, which is inconsistent with many experimental results of real combustion systems. For an illustration, our new model is applied to experimental data and the effect of phase nonlinearity is investigated in the context of combustion instability.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.10
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• pp.93-101
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• 2004

A method to calculate stability limits is investigated to predict the characteristics of high-frequency combustion instability in liquid-propellant rocket engine. It is based on the theory of linear stability analysis proposed in previous works and useful to predict combustion stability at the beginning stage of engine development. The system of equations governing reactive flow in combustor has the simplified and linearized forms. The overall equation expressing stability limits is adopted. The procedures to evaluate quantitatively each term included in the equation are proposed. The thermo-chemical properties and flow variables required in the evaluation can be obtained from calculation of thermodynamic equilibrium, CFD results, and experimental test data. Based on the existent data, stability limits are calculated with actual rocket engine (KSR-III rocket engine). The present calculations show the reasonable stability limits in a quantitative manner and the stability characteristics of the engine are discussed. The prediction from linear stability analysis could be serve as the first approximation to the true prediction.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.240-246
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• 2005

A series of combustion tests of a 30-tonf-class full scale liquid rocket thrust chamber under development has been carried out to verify its design. The test results revealed decent performance in the aspects of efficiency. The combustion stability is one of the most important parameters of liquid rocket engine in addition to the efficiency. Assessment tests of combustion stability must be accomplished to confirm the possibility of combustion instability due to spontaneous or external disturbances. The combustion stability rating tests of the full scale thrust chamber with temporary baffles made of stainless steel were carried out utilizing pulse guns to estimate combustion stability characteristics. The tests results show highly stable combustion stability characteristics. The outcome acquired from the present experimental study will be used to design an actively cooled baffle that can survive for the life time operation of the thrust chamber.