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항공기용 가스터빈 엔진의 블레이드 팁에서 유동과 열전달에 대한 수치적 연구

Title
항공기용 가스터빈 엔진의 블레이드 팁에서 유동과 열전달에 대한 수치적 연구
Other Titles
A Numerical Study of Flow and Heat Transfer on a Blade Tip of an Aircraft Gas Turbine Engine
Author
쟈오리우
Alternative Author(s)
JIAO LIU
Advisor(s)
조진수
Issue Date
2017-02
Publisher
한양대학교
Degree
Master
Abstract
냉각 터빈 블레이드 팁에서 발생하는 누설 유동은 터빈 효율뿐만 아니라 블레이드 팁의 열전달 성능에도 좋지 않은 영향을 미친다. 많은 연구자들은 블레이드 팁에서 발생하는 공력 손실을 줄이고 열전달 성능을 개선하기 위해 냉각이 없는 블레이드를 대상으로 연구를 수행해왔다. 하지만 냉각 요소들이 적용된 블레이드 팁에서 유동 구조와 열전달 특성은 무냉각 블레이드와 다르다. 따라서 본 연구에서는 복합열전달 전산 해석을 통해 3차원 냉각 터빈 블레이드 팁 간극에 의한 유동 및 열전달 특성을 연구하였다. 1단 고압 터빈 노즐 출구에서 획득한 속도와 온도 프로파일을 로터의 입구에 적용하여 로터 도메인만을 대상으로 3차원 RANS 방정식 해를 구하는 전산해석을 수행하였다. 1단 고압 터빈 중 노즐 가이드 베인의 전산해석 결과와 실험 결과를 비교하여 전산해석 기법을 검증하였다. 또는 냉각이 없을 때 블레이드 표면에서 고온 부분을 확인하였고 냉각이 있는 블레이드와 대비하여 냉각 효과를 관측하였다. 스퀼러 팁이 적용된 로터 블레이드 기준 모델의 팁 간극은 스팬 길이의 1%이며, 팁 간극을 1%부터 2.5%로 조절하여 공력 손실, 팁 표면에서 열전달 계수와 막냉각 효율의 변화를 확인하였다. 팁 간극이 커질수록 출구에서 공력 손실과 블레이드 끝단 표면에서 열전달 계수는 증가하였다. 특히 팁 간극이 스팬의 2%일 때 평균 열전달 계수가 급격히 증가하였다. 팁 영역의 막냉각 효율은 팁 간극이 작을수록 높았고, 캐비티 내부 냉각 홀 근처의 막냉각 효율이 높았다. 현재의 냉각 블레이드 팁에서 열전달 특성이 취약한 부분을 발견하였고, 이것은 후속 연구에서 로터 블레이드의 냉각 시스템을 설계할 때 팁에 있는 냉각 홀들의 재배치를 통해 개선할 예정이다.|The tip leakage flow which generated on the cooled turbine blade not only affects the turbine efficiency but also impacts on the heat transfer of blade tip. Many researchers have investigated gas turbine blade without cooling to reduce the aerodynamic loss and improve the heat transfer performance. But the flow and heat transfer characteristics of turbine blades which have cooling system are very different from the blade without cooling. So in the present study, conjugate heat transfer method was used to investigate the tip clearance flow and heat transfer characteristics of 3-dimensional cooled gas turbine blade. The velocity and temperature profiles of first stage high pressure nozzle outlet were used as the rotor inlet boundary condition and numerically studied by solving the 3-dimensional steady Reynolds average Navier Stokes(RANS) equation for rotor domain. The numerical analysis results for nozzle were compared to the experimental results to verify the correctness of numerical method. The blade without cooling was also analyzed to see the high temperature area and cooling efficiency comparing with the cooled blade. The blade base model used the squealer tip and the tip clearance was 1% of span. And the tip clearance variation range from 1% to 2.5% of span. The influence of tip clearance aerodynamic loss, heat transfer coefficient and film cooling effectiveness was observed. Results showed that the aerodynamic loss and the heat transfer coefficient were increased when increasing the tip clearance. Especially when the tip clearance was 2% of span, the average heat transfer coefficient on the tip region was increased obviously. The film cooling effectiveness of tip region was increasing with decreasing of the tip clearance. And there was high film cooling effectiveness at cavity and near tip hole region The present study found the vulnerable area of heat transfer characteristics on the blade tip. This must be considered to rearrange the location of the hole in the future cooling system design.
The tip leakage flow which generated on the cooled turbine blade not only affects the turbine efficiency but also impacts on the heat transfer of blade tip. Many researchers have investigated gas turbine blade without cooling to reduce the aerodynamic loss and improve the heat transfer performance. But the flow and heat transfer characteristics of turbine blades which have cooling system are very different from the blade without cooling. So in the present study, conjugate heat transfer method was used to investigate the tip clearance flow and heat transfer characteristics of 3-dimensional cooled gas turbine blade. The velocity and temperature profiles of first stage high pressure nozzle outlet were used as the rotor inlet boundary condition and numerically studied by solving the 3-dimensional steady Reynolds average Navier Stokes(RANS) equation for rotor domain. The numerical analysis results for nozzle were compared to the experimental results to verify the correctness of numerical method. The blade without cooling was also analyzed to see the high temperature area and cooling efficiency comparing with the cooled blade. The blade base model used the squealer tip and the tip clearance was 1% of span. And the tip clearance variation range from 1% to 2.5% of span. The influence of tip clearance aerodynamic loss, heat transfer coefficient and film cooling effectiveness was observed. Results showed that the aerodynamic loss and the heat transfer coefficient were increased when increasing the tip clearance. Especially when the tip clearance was 2% of span, the average heat transfer coefficient on the tip region was increased obviously. The film cooling effectiveness of tip region was increasing with decreasing of the tip clearance. And there was high film cooling effectiveness at cavity and near tip hole region The present study found the vulnerable area of heat transfer characteristics on the blade tip. This must be considered to rearrange the location of the hole in the future cooling system design.
URI
http://dcollection.hanyang.ac.kr/jsp/common/DcLoOrgPer.jsp?sItemId=000000097959https://repository.hanyang.ac.kr/handle/20.500.11754/124817
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GRADUATE SCHOOL[S](대학원) > MECHANICAL ENGINEERING(기계공학과) > Theses (Master)
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