• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.111-115
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• 2009

The method of test for observing the combustion instability and controling the instability actively by using secondary injection of fuel through flame stabilizer was studied by constructing model combustor of supersonic engine. The frequency of combustion instability was detected by measuring the pressure of combustor using pressure sensor and by optical sensing of flame intensity. Electro-magnetic valve was adopted as actuator for active control and the characteristics of modulated fuel was studied by measured pressure of valve outlet and scattering signal of spray at secondary fuel injection.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.6
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• pp.28-36
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• 2018

A study was conducted to apply open-loop control to the combustion instability in a dump combustor using gaseous hydrocarbon fuels. Control power and frequency were varied by employing a loudspeaker under combustion conditions with similar characteristic chemistry times of the fuels. In the case of open-loop control where the frequency was identical to the combustion instability frequency, the open-loop control power affected the control performance. Results obtained from conducted open-loop control tests, where the frequency was different from the combustion instability frequency, show that setting the open-loop control frequency similar to the combustion instability frequency is effective.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.2
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• pp.24-33
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• 2021

Numerical study is conducted to analyze combustion instability in the partially premixed combustor. The simulations are performed according to fuel conditions, and Large Eddy Simulation(LES) model and PaSR combustion model are implemented in the solver. Comparison with the experimental result is conducted to confirm the validity of simulation, and quantitative and qualitative agreement is confirmed. The flame characteristics in the combustor are subsequently investigated, and the association with the occurrence of combustion instability is clarified. According to the simulation results, the flame length varies greatly depending on the fuel conditions. When the flame length becomes sufficiently long, flame-vortex interactions occurred around the wall sections, which works as the main cause of combustion instability.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1115-1121
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• 2017

In order to obtain high thrust at the beginning of the flight, the solid rocket motor with large initial burning surface area was designed and tested. From the static firing test, lower initial thrust was obtained compared with the expected thrust based on the internal ballistic prediction due to the negative erosive burning effect which reduced the burning rate estimated by APN Law. In addition, the radial mode combustion instability was observed with 8 fins grain configuration. This instability was removed after the odd number of fins were used.

• 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.

• Aerospace Engineering and Technology
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• v.4 no.1
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• pp.171-178
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• 2005

An acoustic analysis of a fuel-rich gas generator for the drive of a turbopump in a liquid rocket engine has been performed and the length of a simulating duct has been determined by comparing the resonant frequency of unstable acoustic modes to simulate an actual flight model gas generator. To simulate more realistically, a realistic short-length simulating duct has been determined by considering 1 or 2 wavelength of the unstable modes. Duct-length adjustment to turbopump can be a method to suppress a combustion instability problem by decoupling of acoustic mode and combustion characteristics. This method has been set up and validated with acoustic analysis and hot firing tests.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.3
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• pp.1-9
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• 2005

An acoustic analysis of a fuel-rich gas generator for the drive of a turbopump in a liquid rocket engine has been performed and the length of a duct has been determined by comparing the resonant frequency of unstable acoustic modes to simulate an flight model gas generator, A practical short-length simulating duct has been determined by considering 1 or 2 wavelength of the unstable modes. Length adjustment of duct to turbopump can be a method to suppress a combustion instability problem by decoupling of acoustic mode and combustion characteristics. This method has been set up and validated with acoustic analysis and hot firing tests.

• Journal of the KSME
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• v.56 no.9
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• pp.54-57
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• 2016

로켓연소기는 수백 개의 분사기로 구성된 연소실에서 연소반응이 발생하게 된다. 이때 구조적인 특성으로 난류유동, 밀도, 온도 등의 변화가 매우 불균일한 3차원적인 분포를 가지게 된다. 이러한 불균일한 분포에 의해 발생되는 압력의 시간변화가 연소기의 고유진동수와 일치될 경우 공진이 발생하게 되어 폭발이 발생하게 된다. 이러한 현상을 연소불안정이라 하고 이글에서는 관련된 연구의 동향을 소개하고자 한다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1999.04a
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• pp.5-5
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• 1999

액체추진 로켓엔진의 개발과정에서 고주파 연소불안정은 엔진의 비행 안정성 및 성능의 보장을 위해 반드시 고려해야 하는 중요한 인자이다. 특히 액체추진 로켓엔진에 사용되는 다양한 추진제 조합 중 LOX/RP-1은 그 성능, 가용성, 경제성 등의 측면에서 우수한 추진제이지만 F-1 엔진의 개발과정에서와 같이 여타 추진제 조합에 비해 고주파 연소불안정 특성이 강하게 나타난다. 액체추진 로켓엔진의 음향불안정 특성 예측을 위해 다양한 방법이 제시되어 왔다. 그 중 n-$\tau$ 2 매개변수 법은 연소불안정 특성 예측에 실험적 고찰을 통한 간단한 연소모델을 포함하는 것으로 신속한 결과를 얻을 수 있다는 장점 때문에 엔진의 예비설계 및 본 설계과정에서 인정성 측면의 분석을 위해 널리 사용되고 있고 기존의 엔진 개발과정을 통해 그 신뢰성이 검증되어 왔다.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.1
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• pp.18-27
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• 2003

To perform combined tests with propellant feeding system and engine, which were developed for KSR-III launcher, vertical test stand was organized and a series of hot-fire combustion tests were carried out with engines of several injector faceplate types. In hot-fire tests in vertical installation, combustion instabilities occurred right after ignition with an engine without baffle, and such combustion instabilities did not occur at ignition add during mainstage operation for an engine with STS or composite baffle. 1.regular and temporary pressure pulsations(popping) were detected during steady operation with a baffle engine, however a development to combustion instabilities with resonant mode was highly suppressed by baffle. With a series of tests, it was confirmed that the last developed engine, which has composite baffle, was operated successfully in KSR-III flight propulsion system.