• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2003년도 제20회 춘계학술대회 논문집
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• pp.127-130
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• 2003

This paper presents the numerical thermal analysis for regeneratively cooled rocket thrust chambers. An integrated numerical model incorporates computational fluid dynamics for the hot-gas thermal environment, and thermal analysis for the liner and coolant channels. The flow and temperature fields in rocket thrust chambers is assumed to be axisymmetric steady state which is presumed to the combustion liner. The heat flux computed from nozzle flow is used to predict the temperature distribution of the combustion liner. As a result, we present the wall temperature of combustion liner and the temperature change of coolant.

• 대한기계학회:학술대회논문집
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• 대한기계학회 2008년도 추계학술대회B
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• pp.2862-2867
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• 2008

High-temperature components of gas turbine operated for certain period of time can be reused by being repaired or rejuvenated. In case of the gas turbine combustion liners, the biggest and the most important one in the high-temperature components, come in a repair shop after operated for 8,000 or 12,000 hours according to the model and go through the repair and rejuvenation in order to be reused. A stated combustion liner is the first channel which has the combustion gas reached a nozzle from a fuel nozzle. Materials and coating properties of old and new model combustion liners were investigated. To repair these components after the visual inspection, the coatings of combustion liners were removed and then FPI(Fluorescent Penetrant Inspection), a kind of the NDI(Non-Destructive Inspection), was conducted. Damage patterns and the number of the damaged components were classified and analyzed based on data provided from the visual inspection over a long period of time. Focusing on the difference between old model and new model combustion liners, we analyzed the damage distribution and changes and consequently concluded that new model combustion liner would increase repair rate.

• 대한기계학회:학술대회논문집
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• 대한기계학회 2008년도 추계학술대회B
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• pp.2984-2988
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• 2008

Numerical analyses are carried out in order to understand complex thermal characteristics of a gas turbine combustor liner such as combustion gas temperatures, wall temperatures and heat transfer distributions. As results, The maximum internal and external heat transfer is $2218W/m^2K$ and $2358W/m^2K$, respectively. The combustion gas temperatures range is 673K to 1760K. A range of temperature on TBC is 676K to 1382K. Lastly, temperature range on outer surface of combustion liner cooled by compressed air is 676K to 1188K.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2017년도 제48회 춘계학술대회논문집
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• pp.1165-1168
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• 2017

본 논문에서는 재생냉각형 연소기를 채용한 가스발생기 사이클 로켓엔진의 연소시험 중 연소기의 내피 손상을 진단하는 방안을 제시하였다. 이는 연소기 내피 손상 발생 시 연료가 유실되면서 두가지 방식의 연소기 연료유입량 예측값 차이에 변화가 발생하는 것에 착안한 방법으로 로켓엔진 시험 중 연소기 내피손상을 조기에 파악하여 추가 손상 방지에 기여할 수 있을 것으로 기대된다.

• 한국연소학회:학술대회논문집
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• 한국연소학회 1997년도 제15회 KOSCO SYMPOSIUM 논문집
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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.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2010년도 제35회 추계학술대회논문집
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• pp.642-647
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• 2010

막냉각은 고온의 연소가스에 장시간 노출되어 있는 공기흡입식 추진기관의 연소실 내벽을 열적으로 보호하기 위해 사용하고 있으며, 슬롯형 냉각라이너를 이용한 막냉각 방법은 냉각특성을 향상시키기 위해 오랫동안 연구되어 왔다. 본 논문에서는 다중 슬롯형 냉각라이너의 연소영역과 냉각영역에서의 기체역학적 유동 및 열전달 계산에 대한 연구결과를 기술하였다.

• 한국추진공학회지
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• 제7권4호
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• pp.46-52
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• 2003

액체추진로켓엔진에서는 고온, 고압의 연소가스로부터 엔진을 보호하기 위하여 재생 냉각을 많이 사용한다. 이 재생 냉각의 유동장 해석과 열전달을 수치해석하였다. 형상은 연소실, 노즐 및 냉각 채널 모두를 2차원 축대칭으로 가정하였다. 연소실 및 노즐의 유동장은 압축성 유동장 해석을 통하여 구하였고, 냉각 채널은 고체 열전달로부터 구하였다. 유동장과 온도장은 모두 정상상태로 가정하였다. 노즐에서 구한 열유속은 냉각채널 벽면에서의 온도 분포를 구하기 위해 사용하였다. 결과로는 냉각 채널 벽면에서의 온도 변화와 냉각제의 온도 변화를 구하였다. 냉각 채널의 형상 변화와 냉각제의 유량 변화에 따른 냉각 채널 벽면에서의 온도변화를 조사하였다.

• Journal of Mechanical Science and Technology
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• 제15권9호
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• pp.1319-1327
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• 2001

The results of stand and field testing of a combustion chamber for a heavy-duty 150 MW gas turbine are discussed. The model represented one of 14 identical segments of a tubular multican combustor constructed 1:1 scale. The model experiments were executed at a lower pressure than that in a real gas turbine. Combustion efficiency, pressure loss factor, pattern factor, liner wall temperature, flame radiation, fluctuating pressure and NOx emission were measured at partial and full loads for both model and on-site testing. The comparison of these items in the stand and field test results led to has the development of a method of calculation and the improvement of gas turbine combustors.

• 한국정밀공학회지
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• 제26권3호
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• pp.7-10
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• 2009

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2004년도 제22회 춘계학술대회논문집
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• pp.453-459
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• 2004

An axially staged combustor equipped with an LPP combustion system and CMC liner walls has been investigated for stable combustion and low NOx emissions for the ESPR project. Several fuel injectors were designed and manufactured for the LPP burner, and single sector combustor tests were conducted to evaluate fundamental combustion characteristics such as emissions, instabilities, auto-ignition, and flash back at typical operating conditions from idle to Mn 2.2 cruise. The latest test results showed that the LPP burner had a good potential for the low NOx target. It was also found that the NOx emission level was greatly affected by a distortion in the air flow velocity field upstream of the LPP burner due to the diffuser and fuel feed arm. The CMC material was investigated to apply for the high temperature and low NOx combustor. Annular combustor liner walls were manufactured with the CMC material, and they have been tested at low pressure conditions to evaluate the soundness of the material and the mounting and seal system. This paper reports the latest research activities on the LPP combustion system and CMC liner walls for the ESPR project.