본 연구에서는 케로신과 액체산소를 추진제로 사용하는 핀틀 인젝터의 설계/제작이 이루어졌으며, 연소실 특성길이와 운동량비에 따른 특성속도효율을 통해 연소성능을 확인하였다. 연소시험 결과, 핀틀 인젝터의 추진제 분무 특성으로 인한 재순환영역에 의해, 추진제 조합에 따른 특성길이 추천치보다 짧은 영역에서도 최적의 연소성능을 발휘하는 것을 확인하였다. 또한 운동량비가 연소성능에 지대한 영향을 미치나, 추천범위 1.0~1.5 내에서 특성속도효율이 일정하게 나타남을 확인하였다. In this study, a pintle injector rocket engine which uses kerosene and liquid oxygen as propellants was manufactured by collecting basic design data and establishing a design procedure. Combustion performance of the liquid rocket engine was investigated by characteristic velocity efficiency with characteristic length of the combustion chamber and total momentum ratio. As a result of hot fire tests, it showed that the engine had shorter characteristic length comparing to those of other type injectors, which was known as recommended value with the propellant combination. Also, the characteristic velocity efficiency was greatly affected by total momentum ratio and almost constant within 1.0~1.5.

본 논문에서는 연소 실험을 통해 케로신/액체산소를 추진제로 사용하는 기울어진 슬릿을 가지는 핀틀 분사기 로켓엔진의 연소 성능을 살펴보았다. 슬릿 형상 관련 설계 변수는 슬릿 개수, 슬릿 각 및 차단율로 선정하였다. 그리고 기울어진 슬릿에 의한 양 끝단의 skip ratio의 차이를 ${\Delta}SR$로 새롭게 정의하였다. 실험 결과, 기울어진 슬릿을 가지는 핀틀 분사기의 슬릿 형상에서 BF를 1에 가깝게 하고, ${\Delta}SR$을 0.26 이내로 설계하면 최적의 연소성능을 발휘함을 확인하였다. In this paper, combustion tests were performed to investigate performance characteristics of a canted slit-type pintle injector engine which uses kerosene and liquid oxygen as propellants. The number of slits, slit angle and blockage factor were chosen as design variables of the pintle injector. ${\Delta}SR$ was newly defined as the difference of skip ratio caused by both sides of the tip of the canted slit. The experimental results showed that optimal combustion was performed when the blockage factor is about 1 and the difference is less than 0.26.

The effects of CO2 dilution on NOx emissions and combustion instability are studied in an effort to use biogas as a fuel in dual lean premixed gas turbines. More specifically, the influence of the radiation heat transfer characteristics of CO2, which is a major component of biogas, on the combustion characteristics of a dual lean premixed flame of biogas is investigated. OH* chemiluminescence images are used to observe the flame structure for various CO2 dilution rates. The results show little difference in the flame structure for CO2 dilution rates of up to 40%, while higher dilution rates lower the flame intensity. Also, dual lean premixed flames show very different flame structures and temperatures depending on the pilot fuel mass fraction. The results show a decrease in the flame temperature when the dilution rate is increased, resulting in a reduction in the thermal NOx emissions. The present paper also shows that the radiation heat release, which is due to the high heat release rate of CO2, promotes a further drop in the flame temperature, in addition to the thermal effects of CO2. These findings are numerically modeled, empirically verified, and added to the NOx prediction model. CO2 dilution also changes the combustion oscillation frequency. To estimate this, the flame temperature is calculated as a function of the CO2 dilution rate and the pilot fuel mass ratio. The accuracy of the flame temperature calculation can be improved with a more accurate estimation of the radiation heat loss. In calculating the radiation heat loss for an unstable flame, considering the heat release fluctuation (i.e. the temperature fluctuation) improves accuracy, whereas taking the average temperature results in an underestimation. With an improved estimation of the temperature, the combustion oscillation frequency is more accurately predicted. (c) 2012 Elsevier Ltd. All rights reserved.

There has been a rapid increase in the demand for biogas applications in recent years, and dry low NOx and dry low emission gas turbine combustors are promising platforms for such applications. Combustion instability is the most important drawback in dry low NOx gas turbine combustors and has, therefore, attracted considerable research interest lately. As a fundamental study towards the use of biogas in dry low NOx and dry low emission gas turbine combustors, this article investigates the influence of CO2 in surrogate biogas on combustion instability. Tests were conducted using a dry low NOx type, a dual lean premixed gas turbine combustor. For a dual flame with dual swirl, the pilot fuel mass fraction affects the flame structure, and the flame structure, in turn, determines the temperature distribution in the combustion chamber and the combustion instability. The effects of the pilot fuel mass fraction, which is an important parameter of the combustor, and the CO2 dilution rate, which is a major contributor of biogas combustion, on the combustion characteristics and instability are investigated through dynamic pressure signal and phase-resolved OH* images. Combustion instability decreases for higher CO2 dilution rates, whose effects depend on the pilot fuel mass fraction. The instability reaches its maximum at a pilot fuel mass fraction of 0.3. Tests confirm that combustion instability diminishes with CO2 dilution, as it reduces the perturbation in the heat emission, and the flame speed decreases resulting in a greater flame surface or volume. Further, investigation of the Rayleigh Index, which represents the coupling strength of the heat release fluctuation and the natural frequency, shows that CO2 dilution weakens the coupling strength, resulting in less combustion instability.

A 5MW-gas turbine engine (DGT-5) for power generation is currently under development by Doosan Heavy Industries, funded under a national R&D program. Tests of its combustor performance are carried out by the Korea Aerospace Research Institute (KARI). The combustor is a Dry Low NOx (DLN) type premixed combustor consisting of dual radial swirl burners, air-cooled liner, transition piece, and torch system. It uses natural gas as fuel. Measured characteristics of the combustor include emissions, pressure pulsation, and exit temperature distribution. The optimal operation point of the combustor is found by investigating parameters such as the fuel ratio between the pilot and the main burner and various total fuel rates. The test results show that the combustor performance satisfies the requirements for a gas turbine system. The following results are obtained from the test: pressure loss<4.3; NOx emissions<20ppm; PF<7.6%; Pressure pulsation amplitude <2kPa at low pressure (3bar, a), 100% load and FR 0.2 conditions.

A non-intrusive measurement, Planar Laser Induced Fluorescence was employed to visualize and measure the fuel distribution of the non-reacting field at the burner exit of gas turbine combustor. Measurement techniques, image processing method and quantification procedure were presented. Also, concentration measurement with gas analyzer was carried out to verify the propriety of PLIF result. The PLIF result coincides well with gas analyzer measurement result. PLIF test result for several other conditions are mentioned as well.

․ Sunjin Kim*** ․ Hyung-Mo KIM**** ․ Yong-Wun Jung****ABSTRACT A small high altitude test facility has been developed to inv estigate ignition performance of a small gas-turbine combustor under high altitude conditions. Supersoni c diffusers and a heat exchanger were used to perform a low pressure and a low temperature condition, respectively. Experimental results showed that the low pressure environment could be controlled by upstream pressure of primary nozzle flow and low temperature environment by mixture ratio of cooled air and ambient air. Ignition performance tests were performed to verify the performance of the facility under simulated high altitude conditions. Conclusively, it was proven that the test facility could be used for ignition performance test of a small gas-turbine combustor under high al titude condition of approximately 6,100m.초 록 본 연구에서는 가스터빈 연소기의 고공환경 모사 점화 성능 시험을 목적으로 소형 고공환경 모사 시험 설비를 구축하였고 이에 대한 성능 실험을 수행하였다. 고공환경 조건인 저압 환경 구현을 위해서는 초음속 디퓨저를 사용하였고, 저온 환경 구현을 위해서는 드라이아이스를 냉각제로 사용한 열교환기를 사용하였다. 저압 환경 구현 성능 실험 결과 연소기로 20g/s의 공기 공급 상태에서도 연소기 내부에 고도 약 6,100m에 해당하는 저압 환경 구현이 가능한 것을 확인하였다. 또한 저온 환경 구현 성능 실험 결과 연소실 내부에 고도 6,100m 이상의 저온 환경 구현이 가능한 것을 알 수 있었으며, 상온공기와 냉각공기의 혼합율 조절로 다양한 고도의 저온 환경 구현이 가능한 것을 확인하였다.Key Words: Gas Turbine Combustor(가스터빈 연소기), High Altitude Condition(고공환경), Ignition Test Facility(점화시험장치), Ignition Performance(점화 성능)1. 서 론 비행체의 추진기관 및 보조동력장치(APU)로

Currently, 60kW class gas turbine fuel cell hybrid power generation system is being developed by Korea Aerospace Research Institute. In this system, gas turbine combustor acts as a component which supplies hot gas to the highly efficient power generating fuel cell as well. Because of connection problems between gas turbine and fuel cell, the combustor is necessarily to have flame stability and wide flamability range at the starting stages. Furthermore, in order to enhance the efficiency of the whole system, high combustion efficiency and low pressure loss are required. It also should satisfy the temperature limitations of turbine blade by supplying sufficiently low and uniform thermal energy. In the present study, for the optimization of the combustor, its stability characteristics are tested with change of various operating conditions such as injector shape, swirler air mass flow rate etc. For the basic performance factor survey, effects of combustor exit temperature, pressure loss, combustion efficiency and exhaust gas compositions are also investigated.

The probability density functions (PDF) of temperature were measured by coherent anti-Stokes Raman Spectroscopy (CARS) in flames of gas turbine combustor sector rig of an aero-engine. The combustor was operated at simulated ground idle conditions with standard kerosene fuel. Temperature PDFs had been measured near fuel nozzle with change of rotation of a swirler and existence of a prefilmer. The characteristic features of temperature PDFs showed the variation of combustion configurations at four experimental conditions. Without a prefilmer, large recirculation of high temperature gas was expected in the co-flow condition and un vaporized fuel fragments were detected in the counter-flow condition. With a prefilmer, the enhanced mixing increased combustion intensity near fuel nozzle in the counter-flow condition and the flame was attached far from the fuel nozzle in the co-flow condition.