• Transactions of the Korean Society of Mechanical Engineers
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• v.18 no.1
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• pp.240-249
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• 1994

The basic experiments for designing the effective gas turbine combustor were performed. There are several factors that define the characteristics of gas turbine combustor. Among them, experiment was focused on swirl effects by three types of swirler with different swirl numbers(0.0, 0.38, and 0.62). Particularly, an interest was concentrated on primary zone where the flame characteristics of total combustor was dominated strongly and secondary zone where the remaining unburned gas was reacted again or cooling effect was done according to degree of swirl intensity. For this study, following measurements have been carried out, that is, time mean and fluctuating temperature, exhaust gas composition including NO concentration, and ion current. From this study, it was found that swirl intensity affects largely not only flame style but also emission formation, furthermore that it is important to select proper swirl intensity.

• Transactions of the Korean Society of Mechanical Engineers
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• v.18 no.7
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• pp.1789-1798
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• 1994

This paper aimed for numerical simulation of complicated gas turbine combustor with swirler. For the convenience of numerical analysis, fuel nozzle and air linear hole areas of secondary and dilution zone, which are issued to jet stream, were simplified to equivalent areas of annular type. In other to solve these problems, imaginary source terms which are corresponded to supplied fuel amount were added to those of governing equation. Chemical equilibrium model of infinite reaction rate and $k-{\epsilon}-g$ model with the consideration of density fluctuation were applied. As the result, swirl intensity contributed to mixing of supplied fuel and air, and to speed up the flame velocity than no swirl condition. Temperature profiles were higher than experimental results at the upstream and lower at the downstream, but total energy balance was accomplished. As these properties showed the similar trend qualitatively, simplified simulation method was worth to apply to complicated combustor for predicting combustion characteristics.

• 한국연소학회:학술대회논문집
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• 1997.06a
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• pp.135-151
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• 1997

The preliminary design and performance test for determining dimensions of gas turbine combustor were investigated. The combustor design program was developed and applied to design our combustor. and detailed design for determining of swirler. dome and liner holes were performed experimentally. The swirler. which govern the combustion characteristics of combustor, was determined $40^{\circ}$ as swirl angle at first performance test. After second performance test the swirler was re-determined by 24 mm i.d.. 34 mm o.d., and swirl angle of $45^{\circ}$. The geometry of liner holes were determined by considering the flame stability and recirculation zone size. It was found that flame can be more easily stabilized by adjusting the swirier dimensions rather than liner holes. The geometry of swirler and liner holes were re-determined by final performance test with dilution holes. Also. the performance of combustor was evaluated by analysis of exhaust gases.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.40 no.8
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• pp.503-510
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• 2016

Incinerators generally emit pollutants such as NOx and CO during the combustion process. In this paper, pollutant emissions and temperature distributions were studied in a simulated incinerator with a reversed (relative to the flue gas flow) secondary air injection system. The experiments were performed by using a lab-scale furnace in order to evaluate the effects of the injection location, direction and flow rate of secondary air jets. The emission of NOx was lower in the case of reversed secondary air injection than in the case of cross injection, due to the recirculation and mixing of the exhaust gas. In the reversed air injection cases, thermal NOx emissions decreased as secondary air ratio increased from 30 to 60 and slightly increased at secondary air ratios higher than 60. In most cases, CO emissions were not detected except for a few reversed secondary air injection cases, in which cases CO concentrations below 2ppm were observed.

• Clean Technology
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• v.26 no.3
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• pp.211-220
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• 2020

This experimental study investigates the design parameters to achieve ultra low NOx combustion of coal using a 80 kW capacity single-burner furnace. The influence of key design parameters such as SN, overall and burner-zone equivalence ratios, primary/secondary air ratio, overfire air (OFA) ratio were tested for a total of 81 cases. The results showed that weak swirl intensity of the burner leads to higher NOx emission whereas strong swirl intensity accompanies increased CO concentration desipte lower NOx emission. Therefore, finding an appropirate swirl intensity is essential for the burner design. Larger flow rate of secondary air increased NOx emission, whereas smaller flow rate stretches the flame and increased CO emission. The lowest NOx emission of 82 ppm (6% O₂) was achieved at the optimal condition of the present burner deisgn. It is expected to furrther lower the NOx emission by introducing splitting the burner secondary air into three or four streams.

• The Bulletin of The Korean Astronomical Society
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• v.37 no.2
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• pp.151.2-151.2
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• 2012

터보펌프 구동에 사용된 가스발생기 생성가스를 연소기로 공급하여 주추력 발생에 사용하는 다단연소 사이클 로켓엔진은 고추력을 요하는 우주 발사체에 널리 사용되고 있다. 다단연소 사이클 로켓엔진에 사용되는 가스발생기를 예연소기라 부르며 케로신과 액체산소를 추진제로 하는 다단연소 사이클 로켓엔진에는 산화제 과잉 예연소기가 사용된다. 예연소기는 터보펌프 구동을 목적으로 하기 때문에 예연소기 생성가스의 횡단면 온도분포는 터빈에 의해 제한되는 온도범위 내에서 균일하여야 하며 넓은 운전영역에서 안정적인 연소가 이루어져야 한다. 산화제 과잉 예연소기는 모든 추진제가 혼합헤드를 통해 분사되는 방식과 추진제를 혼합헤드와 연소실로 나누어 공급하는 방식이 있다. 기술검증을 위해 산화제 일부와 연료를 혼합헤드를 통해 연소실에 공급하여 1차 연소시키고 나머지 산화제를 연소실 냉각채널을 거쳐 연소실 중앙의 분사공을 통해 연소실로 주입하여 기화시키는 형태로 최종적으로 연소압 20MPa, 혼합비 60에서 작동하는 산화제 과잉 예연소기를 설계하여 연소시험을 수행하였다. 혼합헤드에는 별도의 점화용 분사기 없이 전체 연료 분사기를 통해 점화용 연료인 TEA/TEB 혼합물을 분사하여 점화하였다. 추진제를 2단으로 공급할 수 있도록 고안된 가압식 연소시험 설비에서 10회, 누적 60초 이상의 연소시험이 성공적으로 수행되었다. 연소시험결과 넓은 작동영역에서 안정적 연소특성과 생성가스 온도 분포의 균일성을 확인할 수 있었다. 고온 고압의 산화제 과잉 예연소기 기술 확보를 통해 케로신/액체산소 다단연소 사이클 로켓엔진 개발을 위한 기술적 기반을 마련하였다.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.1
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• pp.51-56
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• 2007

Flow characteristics in a liquid fuel ramjet combustor were investigated using the PIV method. The combustor has two rectangular inlets that form a $90^{\circ}$ angle each other. Three cases of test combustors are made in which those inlet angles are $30^{\circ},\;45^{\circ}\;and\;60^{\circ}$. The experiments were performed in a water tunnel test with the same Reynolds number as Mach 0.3 at the inlet. PIV software was developed to measure the characteristics of the flow field in the combustor. A large and complex recirculating flow was measured in the dome area with 4 different dome size. Experimental results shows that 1/3 dome size of combustor diameter is suitable and smaller inlet angle provide large recirculation flow at the dome of combustor as a frame holder in this experimental ranges but need to consider secondary recirculation flow in a junction region to optimize the configuration of ramjet combustor.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.24 no.10
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• pp.1294-1299
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• 2000

The combustion characteristics have been investigated to develop the 50 kW-class gas turbine combustor. The combustor design program was developed and applied to design this combustor. The combustion air which has the temperature of 45, 200, $300^{\circ}C$ were supplied to combustor for elucidating the effect of inlet air temperature on CO, NOx emissions and flame temperature. The exit temperature and NO were increased and CO was decreased with increasing inlet air temperature. Also, the effect of equivalence ratio was considered to verify the combustor performance. The emissions of CO and NO with inlet air temperature can be analyzed qualitatively by measuring the temperature inside the combustor. The combustion performance with fuel schedule was evaluated to get the informations of the starting and part loading process of gas turbine. The combustion was stable above the equivalence ratio of 0.18. The pattern factor which is the important parameter of combustor performance was satisfied with the design criterion. Consequently the combustor was proved to meet the performance goal required for the target gas turbine system.

• Transactions of the Korean Society of Mechanical Engineers
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• v.18 no.4
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• pp.1064-1071
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• 1994

There are three active radicals which become to the scale of flame diagnostics at the flame front. They are OH, CH and $C_2$ radical. For this, optical measurement system which could monitor simultaneously the luminous waves of three radicals, was constructed. These were analyzed statistically into the cross correlation, coherence and phase. Through such an statistical treatment, combustion characteristics was investigated at the primary zone of gas turbine combustor. The local equivalence ratio was predicted with the ratio of luminuous intensity between CH and $C_2$ radical. This result was matched up to the equivalence ratio calculated from gas composition within 5% error. In general, equivalence ratio was said to be 1.0 at flame front, but it could be increased up to about 1.2 depending on the degree of swirl intensity in case of changing properly the air amount of primary zone.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.159-162
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• 2008

In this study, various cycles of liquid rocket engines were surveyed and specifically gas generator after burning cycle was investigated for upper stage motors. The engines for the upper stage can be categorized into three group based on the cycles and propellants at the diagram. Kerosene engines which adapt the gas generator after burning cycle and are located in the region II, are characterized for high combustion pressure and complexity. This cycle usually needs more than two pumps to use the turbine power efficiently. The fuel line can be divided into the gas generator line and the combustor line, and only the gas generator line is need to be pressured more because the combustion pressure in the gas generator is much higher than that of the combustor. Basically, all the oxidizer goes into the gas generator and than to the combustor, thus the auxiliary LOx pump is not critically necessary. However, for the various reasons, the LOx line requires a booster pump. A gas generator after burning cycle engines produces relatively high specific impuls than that of the open cycle engines. Thus it is suitable for upper stages of launch vehicles.