• 한국연소학회:학술대회논문집
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• 한국연소학회 2006년도 제32회 KOSCO SYMPOSIUM 논문집
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• pp.75-78
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• 2006

Disposal of highly toxic wastes like polychlorinated biphenyls (PCBs) is very difficult. These substances create a growing mountain of problematic waste that has to be disposed properly. Conventional technologies that are based on common burning(rotary kiln, ${\sim}1100^{\circ}C$) and plasma technology(${\sim}10000^{\circ}C$) do not satisfy important conditions. for example, complete combustion of the toxic waste and the price of waste disposal. The combustor like a rocket engine is operated at relatively high pressure(${\sim}15$ bar) and relatively high temperature(>$3000^{\circ}C$) that are ideal for the complete destruction of extremely toxic substances. In this study, test compound($_o-DCB$) was dissolved in kerosine with a concentration of 10%. Pure gas oxygen was used as an oxidant. Analysis showed that the destruction efficiency achieved for ${o}-DCB$ was 99.9999% or better. The results show that a combustor based on liquid propllant rocket technology is a validated tool for the disposal of highly toxic waste, and a good alternative technology when applied to the destruction of extremely toxic wastes.

• 한국추진공학회지
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• 제11권3호
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• pp.43-49
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• 2007

한국 최초의 액체로켓인 과학로켓 3호(KSR-III) 엔진 개발을 위해 로켓엔진 연소기 지상연소시험설비가 2001년도에 한국항공우주연구원 내에 준공되었다. 본 시험설비는 준공 후 현재 약 170회에 이르는 액체로켓엔진 연소기 시험을 수행하였으며, 그 과정에서 설비 능력에 대한 많은 개량이 이루어졌다. 본 논문에서는 한국항공우주연구원 액체로켓엔진 연소기 지상연소시험설비의 개요와 수행시험, 그리고 그 동안 축적해 온 설비 운영기술에 대하여 소개한다.

• 한국추진공학회지
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• 제22권6호
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• pp.72-83
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• 2018

본 연구에서는 단일 전단 동축 분사기를 이용한 수소 로켓 연소기의 전산유체 해석을 수행하였다. 2차원 축대칭 형상에서 난류연소 해석을 위해 hybrid RANS/LES 난류모델을 적용하였다. 적합한 해석기법을 찾기 위해 3가지 화학 반응기구, 3가지 고해상도 기법 및 3단계 격자해상도 조합을 비교하였다. 벽면 열유속을 실험결과와 비교하여 해석 성능을 살펴보았으며, 유동장 결과 분석으로 동축 분사기를 가지는 로켓 연소기의 난류연소특성을 살펴볼 수 있었다.

• International Journal of Aeronautical and Space Sciences
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• 제18권4호
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• pp.788-796
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• 2017

This paper proposes a new measurement method to improve the shortcomings of an existing integral method for measuring heat flux in plug-type heat flux gauges in the high-temperature and high-pressure environments of liquid-rocket combustors. Using the existing integral measurement method, the calculation of the surface area for the heat flux in the gauge exhibits error in relation to the actual surface area. To solve this problem, transient profiles obtained from ANSYS Fluent were used to calculate unsteady heat flux as it adjusted to the measured temperature. First, a heat flux gauge was designed and manufactured specifically for use in the high-temperature and high-pressure conditions that are similar to those of liquid rocket combustors. A calibration test was performed to prove the reliability of the manufactured gauge. Then, a combustion experiment was conducted, in which the gauge was used to measure unsteady heat flux in a liquid rocket combustor that used kerosene and liquid oxygen as propellants. Reasonable heat flux values were obtained using the gauge. Therefore, the proposed measurement method is considered to offer significant improvement over the existing integral method.

• 한국연소학회:학술대회논문집
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• 한국연소학회 2000년도 제21회 KOSCO SYMPOSIUM 논문집
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• pp.92-100
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• 2000

One method to analyse acoustic modes is proposed to predict the characteristics of acoustic instability in liquid rocket engine. It is based on the similarity between transverse acoustic modes and adopts two-dimensional axisymmetric geometry. Using this method, the first tangential mode in the prototype combustor can be analysed through the analysis of the first radial mode in the model combustor with doubled chamber diameter. Sample numerical calculation is demonstrated applying this method to sample rocket engine and thereby acoustic instabilities of the engine are investigated. The present results show a good agreement with the previous findings. The numerical analysis based on the proposed method is cost-effective and serves as the first approximation to the true solution.

• Journal of Mechanical Science and Technology
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• 제18권7호
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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.

• Journal of Mechanical Science and Technology
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• 제20권6호
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• pp.896-904
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• 2006

In a liquid rocket engine, the role of gas-liquid scheme injector as an acoustic resonator or absorber is studied experimentally for combustion stability by adopting linear acoustic test. The acoustic-pressure signals or responses from the chamber are monitored by acoustic amplitude. Acoustic behavior in a rocket combustor with a single injector is investigated and the acoustic-damping effect of the injector is evaluated for cold condition by the quantitative parameter of damping factor as a function of injector length. From the experimental data, it is found that the injector can play a significant role in acoustic damping when it is tuned finely. The optimum tuning-length of the injector to maximize the damping capacity is near half of a full wavelength of the first longitudinal overtone mode traveling in the injector with the acoustic frequency intended for damping in the chamber. When the injector has large diameter, the phenomenon of the mode split is observed near the optimum injector length and thereby, the acoustic-damping effect of the tuned injectors can be degraded.

• 항공우주기술
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• 제7권1호
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• pp.129-135
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• 2008

30톤급 액체로켓엔진용으로 개발된 일체형 재생냉각 연소기는 연료를 확대노즐부의 중간에서 공급하는 방식으로 설계되었으며, 노즐 끝단에서 공급되는 방식에 비해 냉각유로는 복잡해지지만 열유속이 상대적으로 낮은 확대노즐부의 냉각유량을 줄임으로서 압력손실을 감소시키는 동시에 공급라인을 포함한 연소기 전체 외경이 줄어들어 엔진 구성에 유리한 장점을 있다. 본 연구에서는 이와 관련한 연료링과 양 방향 냉각 채널, 그리고 연결/분기 유로에 대해 수치해석을 통한 세밀한 설계검토를 수행하였다.

• 한국연소학회:학술대회논문집
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• 한국연소학회 2007년도 제34회 KOSCO SYMPOSIUM 논문집
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• pp.114-117
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• 2007

In IRR(Integral Rocket-Ramjet), the booster is integrated into the ramjet combustor. Such combustors do not contain combustor liners or flame holders within the combustor due to the limited volume and flame stabilization depends on the recirculation zones formed by the sudden expansion region between the inlet duct and the combustor. A numerical study was conducted on the effect of flame holder which could be added to the inlet duct of IRR. Two different types of flame holder installations, flame holder without sudden expansion region and flame holder with small sudden expansion region, were compared and showed different flame shapes and pressure rise in the combustor.

• 한국항공우주학회지
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• 제32권10호
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• pp.102-111
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• 2004

본 논문에서는 와류분사기를 가진 액체로켓엔진용 축소형 연소기의 설계/제작/시험에 대해 기술하였다. 와류분사기는 내부에 액체산소 외부에 케로신을 공급하여 노즐 외부에서 혼합하는 구조를 가지고 있다. 축소형 연소기는 분사기 헤드, 삭마 냉각방식의 내열재 연소실 그리고 물냉각 노즐로 구성되어 있다. 분사기 헤드는 18 개의 주 분사기, 하나의 중앙 분사기, 연료 메니폴드, 산화제 메니폴드 그리고 추진제 분배기 등으로 구성되어 있다. 축소형 연소기 제작 후 수류시험 및 점화시험을 거쳐 설계점 및 탈설계점에서의 연소시험을 성공적으로 수행하였다. 연소시험결과 분사기 차압은 수류시험시의 값과 비슷하였고 연소효율은 목표치보다 높게 나왔으며, 정상연소시 동압의 진폭은 규격조건을 만족하였고 고주파 연소 불안정은 발생하지 않았다.