• Title/Summary/Keyword: 고압터빈 노즐

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A Study on the Structural Integrity of the First Stage Turbine Blade Caused by Thermal Barrier Coatings and the Cooling Design of the Nozzle (터빈 노즐 및 열차폐 코팅에 따른 고압 1 단 터빈 블레이드의 구조 건전성 영향에 대한 연구)

  • Huh, Jae Sung;Kang, Young Seok;Rhee, Dong Ho
    • Transactions of the KSME C: Technology and Education
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    • v.4 no.2
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    • pp.93-99
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    • 2016
  • High pressure nozzles and turbines of a gas turbine engine should be required to be operated under extreme operating conditions in order to maximize the performance. Engine manufactures have utilized nickel-base superalloys, enhanced cooling design, and thermal barrier coating techniques to overcome them and furthermore, material modeling, finite element analysis, optimization techniques, and etc. have been utilized widely for elaborate predictions. We aim to evaluate the effects on the low cycle fatigue life of the high pressure turbine blade caused by thermal barrier coatings and the cooling design of the endwall of the first stage turbine nozzle. To achieve it, the structural analysis, which utilized the results of conjugate heat transfer analysis as loading boundary conditions, was performed and then the results were the input for the assessment of low cycle fatigue life at several critical zones.

A Study on Variations of the Low Cycle Fatigue Life of a High Pressure Turbine Nozzle Caused by Inlet Temperature Profiles and Installation Conditions (고압터빈 노즐에서 입구온도분포와 장착조건에 따른 저주기 피로 수명 영향에 대한 연구)

  • Huh, Jae Sung;Kang, Young Seok;Rhee, Dong Ho;Seo, Do Young
    • Transactions of the Korean Society of Mechanical Engineers A
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    • v.39 no.11
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    • pp.1145-1151
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    • 2015
  • High pressure components of a gas turbine engine must operate for a long life under severe conditions in order to maximize the performance and minimize the maintenance cost. Enhanced cooling design, thermal barrier coating techniques, and nickel-base superalloys have been applied for overcoming them and furthermore, material modeling, finite element analysis, statistical techniques, and etc. in design stage have been utilized widely. This article aims to evaluate the effects on the low cycle fatigue life of the high pressure turbine nozzle caused by different turbine inlet temperature profiles and installation conditions and to investigate the most favorable operating condition to the turbine nozzle. To achieve it, the structural analysis, which utilized the results of conjugate heat transfer analysis as loading boundary conditions, was performed and its results were the input for the assessment of low cycle fatigue life at several critical zones.

액체로켓용 터빈시스템 설계

  • Choi, Chang-Ho;Kim, Jin-Han;Yang, Soo-Seok;Lee, Dae-Sung
    • Aerospace Engineering and Technology
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    • v.1 no.1
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    • pp.163-172
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    • 2002
  • The turbine system composed of a nozzle and a rotor is used to drive turbopumps while gas passes through the nozzle, potential energy is converted to kinematic energy, which forces the rotor blades to spin. In this study, an aerodynamic design of a turbine system is investigated using compressible fluid dynamic theories with some pre-determined design requirements (i.e.,pressure ratio, rotational speed, required power etc.) obtained from a liquid rocket engine (L.R.E.) system design. For simplicity of a turbine system, impulse-type rotor blades for open type L.R.E. have been chosen. Usually, the open-type turbine system requires low mass flow rate compared to the close-type system. In this study, a partial admission nozzle is adopted to maximize the efficiency of the close-type turbine system. A design methodology of the a turbine system has been introduced. Especially, a partial admission nozzle has been designed by means of simple empirical correlations between efficiency and configuration of the nozzle. Finally, a turbine system design for a 10 ton thrust level of L.R.E is presented.

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Experimental Investigation of Turbopump Turbine : Turbine Performance and Effect of Nozzle-Rotor Clearance (터보펌프 터빈의 성능 및 노즐-로터 간극의 영향에 대한 실험적 고찰)

  • Jeong Eun-Hwan;Kang Sang-Hun;Shin Dong-Yoon;Park Pyu-Goo;Kim Jin-Han
    • Journal of the Korean Society of Propulsion Engineers
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    • v.10 no.2
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    • pp.78-86
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    • 2006
  • This paper presents the performance test result of the 30-ton class turbopump turbine. Test has been conducted using high pressure cold air, The turbine overall performance has been measured for various pressure ratio and rotational speed settings. The nozzle-rotor clearance effect on turbine performance also has been tested for the four kinds of the nozzle-rotor clearance values. We found that turbine efficiency rated 51.1% at its design velocity ratio and pressure ratio of 13.5. We also found that turbine efficiency can be increased by 3.5% for approximately 1mm decrement of the nozzle-rotor clearance from its nominal value.

A Study on the Effects on Low Cycle Fatigue Life of a High Pressure Turbine Nozzle due to the Perturbation of Crystal Orientation of Grain of DS Materials (일방향 응고 재료의 결정립 성장 방향 섭동이 고압터빈 노즐 저주기 피로 수명에 미치는 영향에 대한 연구)

  • Huh, Jae Sung;Kang, Young Seok;Rhee, Dong Ho
    • Transactions of the Korean Society of Mechanical Engineers A
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    • v.40 no.7
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    • pp.653-658
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    • 2016
  • High pressure components of a gas turbine engine are generally made of nickel-base superalloys, using precision casting process due to complicated geometries with intricate channels and cooling holes. Turbine components manufactured from directionally solidified and single crystal materials have columnar grains; however, it is found that the crystals do not grow in its preferred direction, although the orientation can be controlled. This anisotropy can lead to the variations of elastic and Hill's parameters in constitutive equations, and they alter stress distributions and the low cycle fatigue life. We aims to evaluate the effects of perturbed crystal orientations on the structural integrity of a directionally solidified nozzle using low cycle fatigue life. We also attempt to show the necessity for the control of allowed manufacturing errors and stochastic analysis. Our approaches included conjugate heat transfer and structural analysis, along with low cycle fatigue life assessment.

A Study on Reliability of Kriging Based Approximation Model and Aerodynamic Optimization for Turbofan Engine High Pressure Turbine Nozzle (터보팬 엔진 고압터빈 노즐에 대한 크리깅 모델 기반 근사모델의 신뢰도 및 공력성능 최적화 연구)

  • Lee, Sanga;Lee, Saeil;Kang, Young-Seok;Rhee, Dong-Ho;Lee, Dong-Ho;Kim, Kyu-Hong
    • The KSFM Journal of Fluid Machinery
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    • v.16 no.6
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    • pp.32-39
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    • 2013
  • In the present study, three-dimensional aerodynamic optimization of high pressure turbine nozzle for turbofan engine was performed. For this, Kriging surrogate model was built and refined iteratively by supplying additional experimental points until the surrogate model and CFX result has effective difference on objective function. When the surrogate model satisfied this reliability condition and developed enough, optimum point was investigated. Commercial program PIAnO was used for optimization process and evolutionary algorithm was used for searching optimum point. As a result, difference between estimated value from Kriging surrogate model and CFD result converges within 0.01% and the optimized nozzle shape has 0.83% improved aerodynamic efficiency.

Experimental Investigation of the Effect of Partial Admission Ratio on the Performance of Supersonic Impulse Turbine (초음속 충동형 축류터빈의 부분분사비 효과에 대한 실험적 연구)

  • Jeong, Eun-Hwan;Park, Pyun-Goo;Kim, Jin-Han
    • Journal of the Korean Society of Propulsion Engineers
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    • v.11 no.4
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    • pp.59-66
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    • 2007
  • In this paper, experimental investigation results of the effect of partial admission ratio on the performance of axial turbine was presented. A supersonic impulse turbine of gas generator cycle liquid rocket engine turbopump was used for the test. for experimental purpose, a nozzle block, in which total 14 number of axi-symmetric convergent-divergent nozzles are arranged circumferentially, was designed and manufactured. Partial admission ratio was controlled by changing the number of active nozzles. High pressure air was used as working medium for the test. The experimental result revealed that the performance of the supersonic impulse turbine does not much affected by the partial admission ratio for supersonic impulse turbine.

Study on the Turbine Performance of 7 ton Liquid Rocket Engine Turbopump (7톤급 액체로켓 엔진 터보펌프 터빈 성능 연구)

  • Lee, Hanggi;Shin, Juhyun;Choi, Changho
    • Journal of Aerospace System Engineering
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    • v.11 no.1
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    • pp.8-13
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    • 2017
  • This study was performed to evaluate the turbine performance of a turbopump in the third stage engine of the Korea Space Launch Vehicle-II. The turbine is a supersonic impulse type with a single rotor. One nozzle is for starting and four remaining nozzles are for steady operation. A similarity test was carried out in the high air test facilities at the Korea Aerospace Research Institute. Test results showed that turbine efficiency changed much more from rotational speed variations than by pressure ratio variations. These results showed characteristics similar to other supersonic impulse turbines.

Conjugate Heat Transfer Analysis for High Pressure Cooled Turbine Vane in Aircraft Gas Turbine (항공기용 가스터빈의 고압 냉각터빈 노즐에 대한 복합열전달 해석)

  • Kim, Jinuk;Bak, Jeonggyu;Kang, Young-Seok;Cho, Jinsoo
    • The KSFM Journal of Fluid Machinery
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    • v.18 no.2
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    • pp.60-66
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    • 2015
  • Conjugate heat transfer analysis was performed to investigate the flow and cooling performance of the high pressure turbine nozzle of gas turbine engine. The CHT code was verified by comparison between CFD results and experimental results of C3X vane. The combination of k-${\omega}$ based SST turbulence model and transition model was used to solve the flow and thermal field of the fluid zone and the material property of CMSX-4 was applied to the solid zone. The turbine nozzle has two internal cooling channels and each channel has a complex cooling configurations, such as the film cooling, jet impingement, pedestal and rib turbulator. The parabolic temperature profile was given to the inlet condition of the nozzle to simulate the combustor exit condition. The flow characteristics were analyzed by comparing with uncooled nozzle vane. The Mach number around the vane increased due to the increase of coolant mass flow flowed in the main flow passage. The maximum cooling effectiveness (91 %) at the vane surface is located in the middle of pressure side which is effected by the film cooling and the rib turbulrator. The region of the minimum cooling effectiveness (44.8 %) was positioned at the leading edge. And the results show that the TBC layer increases the average cooling effectiveness up to 18 %.

Experimental Investigation of Performance for Supersonic Impulse Turbine (초음속 충동형 터빈의 성능에 대한 시험적 고찰)

  • Lee, Hang-Gi;Jeong, Eun-Hwan;Kim, Jin-Han
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2009.11a
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    • pp.561-565
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    • 2009
  • The performance of supersonic impulse turbine was investigated experimentally. Experiment was performed with the compressed air instead of the high temperature burned gas because of the limitation of test facility and danger. As a result of the experiment with the compressed air, the performance in the real gas(burned gas) was predicted by the similarity method. The nozzle area of prototype turbine was calculated based on the real gas. So, it is difficult to satisfy the similarity conditions completely. Two similarity conditions were set and the design point for real gas was existed between two similarity conditions. And, the new turbine test model with calculated nozzle area based on the compressed air was tested. Therefore, similarity point of the new turbine test model was also existed between above two similarity points. It means that the design point for real gas was similar to the test point with the new turbine model.

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