• Journal of ILASS-Korea
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• v.20 no.3
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• pp.168-174
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• 2015

Combustion instability is critical problem in developing liquid rocket engine. There have been many efforts to solve this problem. In this study, the method was sought through the injector as part of these efforts to suppress combustion instability. If the injector can suppress the disturbance coming from the supply line as a kind of buffer it will serve to reduce combustion instability. Especially we target at gas propellant oscillation in gas-centered swirl coaxial injector. The phenomenon is simulated with acoustic excitation of speaker. The film thickness response at injector exit was measured by using a liquid film electrode. Also the response of spray to the disturbance was observed by high-speed photography. Gas-liquid momentum flux ratio and the frequency of feeding gas oscillation were changed to investigate the effect of these experimental parameters. The trend of response by varying these parameters and the cause of weak points was studied to suggest the better design of injector for suppressing combustion instability.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1997.11a
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• pp.10-11
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• 1997

고체 추진제를 사용하는 추진 시스템을 개발하는데 가장 커다란 문제로 인식되고 있는 것은 추진제의 연소 특성을 이해하는 일이다. 그 중에서도 연소실의 압력 진동과 추진제 벽면으로 흡수되는 복사 열전달에 의한 연소율(burning rate)의 변화로 인하여 발생하는 연소 불안정에 대한 이해는 아직도 완전히 규명되지 않고 있다. 고체 추진제의 연소 불안정에 대한 이론적 해석은 준-정상 1차원 해석(Quasi-Steady Homogeneous One-Dimension) 방법에 의하여 단순화된 지배방정식을 해석하는 것이 일반적으로 잘 알려져 있는 방법이다. 이 가정은 고체 추진제가 연수되는 영역을 두께가 매우 얇은 영역의 표면반응영역(surface reaction layer)과 화학반응이 없는 응축상태영역(condensed phase zone) 그리고 기체상태의 연료와 화염이 존재하는 기체상태영역(gas phase zone) 등의 3영역으로 구분하며, 기체상태영역에서 발생하는 교란에 대한 응축상태영역의 반응시간 크기(response time scale)가 매우 크기 때문에 응축상태영역의 반응은 준 정상적으로 일어난다고 가정하는 것이다.그러나, 연소실의 온도가 $3000^{\circ}K$ 정도의 높은 온도이어서 복사 열전달에 의한 고체 추진제의 가열이 중요한 열전달 방법으로 작용하게 되므로 이를 무시한 이론적 해석은 물리적인 중요성이 약하여질 수밖에 없다. 본 연구에서는 기체영역으로부터 전달되는 복사 열전달은 투명(transparent)한 표면반응영역을 통과하여 응축상태영역에서 모두 흡수되며 추진제 표면에서의 복사열방출(emission)을 고려하였다. 또한 연소불안정 현상을 해석하기 위하여 표면반응영역에서의 경계조건은 선형교란량으로 대치하는 Zn(Zeldovich-Novozhilov) 방법을 사용하였다. 이 방법은 기체상태영역에 대한 구체적인 해석없이도 연소불안정 현상을 해석할 수 있는 장점이 잇다. 즉 응축상태영역에서의 연소율과 표면온도는 각각 기체영역으로부터 전달되는 온도구배와 연소압력, 그리고 복사 열전달의 함수관계이므로 선형교란에 의한 추진제표면에서의 교란경계조건을 얻을 수 잇으며, 응축영역의 교란지배방정식과 함께 사용하여 압력교란과 복사 열전달의 교란에 대한 연소율의 교란 증감 여부를 판단하여 연소 불안정 현상을 해석할 수 있다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.159-162
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• 2011

The hot-firing tests of a subscale gas generator were successfully performed to investigate the effect of injector shape variation on its discharge coefficient. The test results showed that discharge coefficients of fuel and liquid oxygen injectors remained nearly constant irrespective of the mixture ratio or the chamber pressure. In all experimental conditions, the combustion instability was not observed.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.1
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• pp.11-17
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• 2007

Effects of combustion chamber design on combustion stability characteristics of a full-scale gas generator were studied experimentally. Thirty seven double-swirl injectors with recess number of 1.5 were distributed in the injector head, which significantly influences combustion performance. The characteristics of combustion stability were inspected by the parametric variations such as changing length and diameter of the combustion chamber and installing a turbulence ring. The experimental result shows that as the effective length of the combustion chamber decreased, an instability frequency took place in a high-frequency region, and the amplitude of the dynamic pressure generally diminished and could be reduced to the unharmful level. However, the dynamic pressure fluctuation in the region of longitudinal resonant frequency could not be suppressed perfectly.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.6
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• pp.110-116
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• 2004

In this study, a three-dimensional finite-element analysis code has been developed to predict acoustic behaviors in rocket combustion chambers and to quantitatively evaluate acoustic stability margins for various designs with passive stabilization devices such as baffle and acoustic resonators. As a validation case, computations are made for combustion chambers with/without a hub-and-six-blade baffle which are developed in the KSR-III Development Program. Compared with experimental results from ambient acoustic test, the numerical approach reasonably well predicts acoustic pressure responses to acoustic oscillation excitation for both unbaffled and baffled combustion chambers and yields quantitatively good agreement for acoustic damping effects of baffle installation in terms of damping factor ratio and resonant frequency shift.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.4
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• pp.35-40
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• 2011

An experimental study of the flame response in a turbulent premixed combustor has been conducted in order to investigate mechanisms for combustion instabilities in a lean premixed gas turbine combustor. A lab-scale combustor and mixing section system were fabricated to measure the flame transfer function. Measurements are made of the velocity fluctuation in the nozzle using hot wire anemometry and of the heat release fluctuation in the combustor using chemiluminescence emission. The results show that the flame transfer functions are greatly dependent on the modulation frequency as well as operating conditions such as equivalence ratio. Flame dynamics can be generalized as a function of Strouhal number which is a ratio of flame length to modulation wave length.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.5
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• pp.79-87
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• 2002

Acoustic responses to pressure oscillations in axisymmetric combustion chamber are numerically investigated to examine the qualitative trend of acoustic instability in liquid rocket engine. Chamber operating condition and excitation frequency of oscillating pressure are selected as exciting parameters of acoustic instability. Artificial perturbation is simulated by total-pressure oscillation with sine wave at chamber inlet. Many approximations and simplifications are introduced without losing the essence of acoustic pressure response. First, steady-state solution for each operating condition is obtained and next, transient analysis is conducted. Depending on operating condition and excitation frequency, the distinct response characteristics are brought. Weak-strength flames and high-frequency excitation tend to cause sensitive acoustic pressure response leading to unstable pressure field. These results are analyzed based on the correlation with acoustic pressure responses from the previous works adopting laminar flamelet model.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.4
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• pp.395-402
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• 2010

Baffle injectors are protruded into the combustion chamber and form an anti-pulsating baffle to prevent high-frequency combustion instabilities in transverse modes. Being exposed to a high heat-flux environment, the baffle injector has self-cooling passages through which kerosene is convected and heated. The baffle injector with 20 spiral cooling channels has been developed and successfully applied to 30 $ton_f$-class combustors without any performance loss due to an additional cooling. In this work, numerical analysis of conjugate heat transfer in baffle injectors with various cooling channel designs has been performed in order to reduce the fabrication cost which would be considerably increased for the 75 $ton_f$-class combustor. Prior to the application to a full-scale combustor, the thermal durability of the modified design has been verified through the subscale hot-firing tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.281-284
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• 2009

Acoustic tuning frequency of quarter-wave resonators is investigated numerically to suppress combustion instability in a liquid rocket engine. A quarter-wave resonator is adopted, which was designed from the cold acoustic test for optimal damping condition. First, in a model combustion chamber scaled down from a full-scale chamber, reactive flow filed is analyzed numerically and acoustic-pressure responses are examined. Next, thermodynamic properties in the resonators are predicted. Based on the data, frequency tuning method is studied. The optimum tuning length of each resonator is proposed and thereby, sufficient damping is produced.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.4
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• pp.61-66
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• 2004

The objective of the present study is to develop methodology for the assessment of combustion stability of liquid rocket injectors. To simulate actual combustion occurring inside of a thrust chamber, a fullscale injector has been employed in the study, which bums gaseous oxygen and mixture of methane and propane. The main idea of the experiment is that the mixing mechanism is considered as a dominant factor significantly affecting combustion instability in a fullscale thrust chamber. A single split triplet injector has been used with an open-end cylindrical combustion chamber. The characteristics revealed by excited dynamic pressures in gaseous combustion show degrees of relative acoustic damping depending on operating conditions. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.