• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.461-464
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• 2012

The high-frequency data acquisition system of the rocket engine test facility has been updated to perform hot-firing tests of 7 ton-class liquid rocket engine combustion chambers which will be used for the third stage of the Korea space launch vehicle II. The paper deals with the design of the updated high-frequency data acquisition system and explains its main functions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.81-85
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• 2012

In this paper, the typical combustion instabilities in hybrid rocket were studied. To induce combustion instabilities in the combustor with the diaphragms were mounted, on front and rear of the fuel, and combustion experiments were performed. The calculated theoretical frequencies using Longitudinal Acoustic Mode and Helmholtz Mode are compared with experimental frequencies using FFT analysis. The theoretical calculated results showed good agreements with experimental results are compared.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.493-497
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• 2012

The rocket engine test facility(ReTF) was improved for hot firing tests of 7 ton-class liquid rocket engine combustion chamber, which will be used for the third stage of the Korea Space Launch Vehicle II(KSLV-II), considering convenience of operation and maintenance, flexibility and safety. In this paper, main modifications and functions of improved ReTF were described. 초 록

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.2
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• pp.17-25
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• 2009

A typical unsteady internal ballistic analysis model was proposed to take account of the erosive burning with the variance of local velocity and pressure along the grain surface of a solid rocket combustor. To validate the model of concern in the study, both cases of non-erosive and erosive burning were compared with the previous researches with marginal accuracy. It was quantitatively investigated that the combustion pressure, grain length, initial temperature, and vaporization temperature of propellant affect the erosive burning characteristics.

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.195-204
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• 2006

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

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.259-264
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• 2004

The critical component of combustor having high combustion efficiency for high performance liquid rocket engine is injector. The results of design and hot firing tests of six sub-scale combustors which have respectively an impinging type injector(1ea.), an bi-propellant swirl closed injector(1ea.), and hi-propellant swirl mixed injector(4ea.) were described in this paper. The combustion test were successfully performed. The combustion efficiency have higher value than predicted value and high frequency combustion instability does not occur.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.460-468
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• 2008

Recent activities on the scramjet and rocket-ramjet combined-cycle engine of Japan Aerospace Exploration Agency(JAXA) are herein presented. The scramjet engines and combined-cycle engines have been studied in the world and JAXA has also studied such the engines experimentally, numerically and conceptually. Based on the studies, 2 to 3 m long, hydrogen-fueled engine models were designed and tested at the Ramjet Engine Test Facility(RJTF) and the High Enthalpy Shock Tunnel(HIEST). A scramjet engine model was tested in Mach 10 to 14 flight condition at HIEST. A 3 m long scramjet engine model was designed to reduce a dissociation energy loss in a high temperature condition. Drag reduction by a tangential injection and two ways of a transverse fuel injection were examined. Combustor model tests at three operating modes of the combined-cycle engine were conducted, demonstrating the combustor operation and producing data for the engine design at each mode. Aerodynamic engine model tests were conducted in a transonic wind tunnel, demonstrating the engine operation in the ejector-jet mode. A 3 m long combined-cycle engine model has been tested in the ejector-jet mode and the ramjet mode since March 2007. Carbon composite material was examined for application to the engines. Production of the cooling channel on a nickel alloy plate succeeded by the electro-chemical etching.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.1
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• pp.42-49
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• 2021

As a measure of susceptibility on the combustion instability, thermo-acoustic instabilities in rocket combustion system was considered for the estimation of the operational stability margin. Growth rate, which governs the asymptotic stability behavior of the system, was determined from the dynamic data measured during combustion tests in order to understand the dynamic characteristics of combustor system. Frequency transform technique was first applied to determine the system parameters such as growth rate and/or damping coefficient for an interested mode from the time series pressure data, and the PDFs of pressure amplitude were extracted from the amplitude envelope of pressure oscillation for the stochastic analysis.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.32 no.9
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• pp.659-668
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• 2008

Combustion instability has been considered as very important issue for developing gas turbine and rocket engine. There is a need for fundamental understanding of combustion instability. In this study, combustion instability was numerically and experimentally investigated in a dump combustor with bluff body. The fuel and air mixture had overall equivalence ratio of 0.9 and was injected toward dump combustor. The pressure oscillation with approximately 256Hz was experimentally obtained. For numerical simulation, the standard k-$\varepsilon$ model was used for turbulence and the hybrid combustion model (eddy dissipation model and kinetically controlled model) was applied. After calculating steady solution, unsteady calculation was performed with forcing small perturbation on initial that solution. Pressure amplitude and frequency measured by pressure sensor is nearly the same as those predicted by numerical simulation. Furthermore, it is clear that a combustion instability involving vortex shedding is affected by acoustic-vortex-combustion interaction. The phase difference between the pressure and velocity is $\pi$/2, and that between the pressure and heat release rate is in excitation range described by Rayleigh, which is obvious that combustion instability for the bluff body combustor meets thermoacoustic instability criterion.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.4
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• pp.25-32
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• 2020

A numerical study on combustion instabilities in a model combustor was conducted with various momentum flux ratios. Five ratios are calculated based on an actual operating condition of rocket engine. As momentum flux ratio increases, the spreading angle on the injector outlet decreases. And, as increase of axial momentum flux, pressure fluctuation decreases inside the combustor. By using dynamic mode decomposition method, the acoustic modes inside the combustor are identified. Combustion stabilities are analyzed by comparing the damping coefficient of the 2^nd longitudinal mode.