260 research outputs found

    Design and commissioning of a rotating test facility simulating a LPT hub cavity system for the investigation of cavity flows in aeroengines

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    Design and commissioning of a rotating test facility simulating a LPT hub cavity system for the investigation of cavity flows in aeroengine

    Numerical study of the Ingestion phenomenon in a turbine Rim Seal : cFD Validation and Real Engine Assessment

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    Tableau d'honneur de la Faculté des études supérieures et postdoctorales, 2011-2012Ce mémoire de maîtrise porte sur la modélisation numérique de la problématique d'ingestion de gaz chauds dans la cavité rotor-stator du premier étage d'une turbine à gaz moderne. Dans le contexte de développement d'une aviation plus écologique, ce sujet mérite une attention marquée car des gains importants en ce qui a trait à l'efficacité globale du moteur peuvent être atteints par de meilleurs designs de joints rotor-stator limitant l'ingestion. La modélisation physique des phénomènes affectant l'ingestion devient alors essentielle afin de développer une meilleure compréhension des écoulements et mécanismes en présence. C'est dans cette optique que Pratt & Whitney Canada a mandaté le Laboratoire de Mécanique des Fluides Numérique (LMFN) de l'Université Laval afin de conduire une étude approfondie du phénomène d'ingestion. Ce mémoire a donc pour objectif principal de décrire le développement, l'application et la validation de certaines techniques de simulation numérique visant à caractériser qualitativement et quantitativement le phénomène d'ingestion observé à travers le joint et dans la cavité du groupe rotor-stator d'une turbine à gaz. Une investigation des limites de fiabilité d'une telle modélisation numérique dans un contexte réel d'opération du moteur est également présentée. La paramétrisation adéquate du problème et la validité d'extrapoler des résultats numériques d'ingestion obtenus aux conditions "banc d'essai" afin de prédire l'ingestion réelle aux conditions "moteur" font également partie des sujets abordés dans ce mémoire. La méthodologie développée considère deux phases de calculs menant à des diagnostics d'ingestion par le biais d'un marqueur passif. Une série de validations numériques et des comparaisons avec des résultats expérimentaux sont également présentées. La méthodologie de modélisation numérique est considérée robuste et fiable, mais seulement partiellement validée

    A novel isolation curtain to reduce turbine ingress heating and an advanced model for honeycomb labyrinth seals

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    A combination of 3-D and 2-D computational fluid dynamics (CFD) modeling as well as experimental testing of the labyrinth seal with hexagonal honeycomb cells on the stator wall was performed. For the 3-D and 2-D CFD models, the hexagonal honeycomb structure was modeled using the concept of the baffle (zero-thickness wall) and the simplified 2-D fin, respectively. The 3-D model showed that even a small axial change of the tooth (or honeycomb wall) location, or a small circumferential change of the honeycomb wall location significantly affected the flow patterns and leakage characteristics especially for small tooth tip clearance. Also, the local details of the flow field were investigated. The seven basic procedural steps to develop a 2-D axisymmetric honeycomb labyrinth seal leakage model were shown. Clearly demonstrated for varying test conditions was the 2-D model capability to predict the 3-D honeycomb labyrinth flow that had been measured at different operating conditions from that used in developing the 2-D model. Specifically, the 2-D model showed very close agreement with measurements. In addition, the 2-D model greatly reduced the computer resource requirement needed to obtain a solution of the 3-D honeycomb labyrinth seal leakage. The novel and advanced strategy to reduce the turbine ingress heating, and thus the coolant requirement, by injecting a Âcoolant isolation curtain was developed numerically using a 3-D CFD model. The coolant isolation curtain was applied under the nozzle guide vane platform for the forward cavity of a turbine stage. Specifically, the isolation curtain serves to isolate the hot mainstream gas from the turbine outer region. The effect of the geometry change, the outer cavity axial gap clearance, the circumferential location of the injection curtain slot and the injection fluid angle on the ingress heating was investigated. Adding the chamfer to the baseline design gave a similar or higher maximum temperature T* max than did the baseline design without chamfer, but implementation of the injection curtain slot reduced substantially T* max of the outer region. In addition, a more desirable uniform adiabatic wall temperature distribution along the outer rotor and stator surfaces was observed due to the presence of the isolation curtain

    가스터빈 림 씰 성능에 휠스페이스 스월이 미치는 영향 측정

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    학위논문 (석사)-- 서울대학교 대학원 : 공과대학 기계항공공학부, 2019. 2. 송성진.가스터빈 엔진은 가장 효율적인 동력원으로서 제트 추진 및 발전을 위해 널리 사용되어왔다. 첨단 재료공학과 이차 유로 시스템, 대류 냉각, 막 냉각 등의 진보한 가스터빈 냉각 기술의 도입은 고효율, 고출력 가스터빈 엔진의 지속적인 발전의 주춧돌이 되었다. 현대의 가스터빈 엔진에서 터빈은 20-30%의 압축기 공기를 냉각, 씰링 및 누설 유동으로 소비한다. 높은 터빈 입구 온도가 다른 모든 손실을 보상한다 하더라도, 이러한 유량 손실은 전효율에 심각한 불이익을 초래한다. 따라서, 씰링 유량을 줄일 수 있는 획기적인 설계가 높은 전효율을 달성하기 위한 하나의 중요한 요인이 되었다. 스테이터와 로터 디스크 사이에 형성되는 휠스페이스로의 주유로 고온 가스 유입 문제는 이차 유로 시스템이 당면한 중요하고 본질적인 문제이다. 주유로와 휠스페이스 압력 차이에 기인하는 이 현상은, 터빈 구성품에 열 피로와 크립 같은 심각한 구조적 안정성 문제를 야기한다. 때문에, 반경 및 축 방향으로 겹쳐지는 형상의 림 씰이 스테이터와 로터 구성품의 주변에 장착되며 충분한 씰링 유량이 유입을 줄이거나 막기 위해 휠스페이스로 공급된다. 림 씰링을 위한 유량을 최소화하기 위한 효과적인 방법에 대한 다양한 연구가 수행되었으나, 유지보수와 구성품 무게와 같은 실용성 측면의 문제가 성능 향상의 발목을 잡았다. 본 논문은 림 씰링 성능을 향상시키기 위한 획기적인 방법론에 대한 실험적 연구를 다룬다. 단일 반경 간극 림 씰과 함께 특수하게 설계된 휠스페이스 선회기가 씰링 성능 향상을 평가하기 위해 사용되었으며, 휠스페이스 내 유동의 선회 성분을 증가시킴으로써 18.49%의 씰링 성능 향상을 달성하였다. 휠스페이스 내에서 씰링 효과와 선회비, 반경 방향 속도 분포를 포함한 다양한 측정이 이루어졌다. 비록 림 씰을 통한 유입은 비정상, 3차원 유동장에 기인하지만, 실험 데이터는 휠스페이스 내 유동의 유체역학적 통찰력을 제공한다. 이러한 실험적 측정은 향후 엔진 설계의 데이터베이스를 확장하는데 기여할 것으로 기대된다. 공력 성능시험과 이차 유로 시스템 연구를 위해 1단 축류 터빈 시험 장비가 새로이 설계되었다. 시험 장비의 형상과 유동 조건은 실제 엔진을 무차원 상사함으로써 엔진의 주유로 및 휠스페이스 유동을 모사할 수 있도록 설계되었다. 설계 항목은 동력원 구성, 시험부 설계, 재질 선정, 구조 해석, 공차 관리 및 밸런싱, 계측 장비 구성을 포함한다. 운전 조건은 타기관 시험 설비의 사양과 일치하는 경향을 보여주었으며, 설계된 시험 장비는 고압터빈단에 널리 이용되는 이중 반경 간극 림 씰로 검증되었다. 포괄적인 계측장비 구성은 주유로와 휠스페이스 내에서 다양한 측정을 가능케한다. 또한, 시험 장비에 적용된 설계 특성들을 통해 다양한 시험 환경을 조성할 수 있다.As the most adaptable source of power, the gas turbine engines have been widely used for jet propulsion, marine and industrial application. Introduction of advanced gas turbine cooling technologiessecondary air system, internal convective cooling, external surface film cooling with cutting edge metallurgy, formed one of the major pillars supporting the continuous development of high efficiency, high power output gas turbine engines. In modern gas turbine engines, the turbine alone may use 20 to 30% of the compressor air for cooling, sealing and leakage flows, which presents a severe penalty on the overall efficiency even the turbine inlet temperature is sufficiently high for the gains to outweigh the losses. Therefore, the novel design to minimize the sealing flow demand will be a key factor to achieve the better overall efficiency of the engines. The hot mainstream gas ingress into the wheel-space, formed between the stator and rotor disks, is one of the most important and intrinsic problems of the secondary air system faced. Principally governed by the pressure difference between mainstream annulus and wheel-space, the turbine components experience serious structural integrity problems such as thermal fatigue and unwanted creep. The rim seals, with the combinations of radial and axial overlapping geometries, are installed at the endwall platform between stator and rotor components. Inevitably, sufficient sealing flow is introduced into wheel-space to reduce or isolate the ingress. The efficient methods to minimize the sealing flow demand for rim sealing purpose have been studied, however, following practical problems in the aspect of maintenance and weight of components caught up with further improvement. This thesis presents an experimental investigation of novel methodology to improve rim sealing performance. By adding swirl flow component inside the wheel-space, 18.49% reduction in sealing flow demand was achieved. The single radial-clearance rim seal with specially designed blades, called ``wheel-space swirler'', are used to evaluate the sealing performance improvement. The extensive range of measurements including sealing effectiveness, swirl ratio and radial velocity distribution inside the wheel-space had been conducted. Although the ingress through the rim seal is a consequence of an unsteady, three-dimensional flow field, the experimental data gave insights into the fluid dynamics for wheel-space flow. These experimental measurements are expected to provide the wider database that can be used for future engine design. The design of single-stage axial turbine research facility, available on both aerodynamic performance and secondary air system studies, is described. It was designed to fulfil engine representative flows both in mainstream and wheel-space, by downscaling the full size engine. The on-design operating conditions are shown to be in the trend of other gas turbine research facilities. The research facility was validated with the double radial-clearance rim seal which has been widely used in high pressure turbine stage. Comprehensive instrumentations allow the detailed measurements both in the mainstream and wheel-space. The design features applied on the research facility enable versatile test configurations.Abstract Contents List of Tables List of Figures Nomenclature Chapter 1 Introduction 1.1 Gas Turbine Engines 1.2 Secondary Air System 1.3 Hot Gas Ingestion 1.4 Thesis Aims 1.5 Thesis Outline Chapter 2 Literature Review 2.1 Wheel-space Flow Structure 2.2 Hot Gas Ingestion 2.2.1 Experiments on Various Rim Seal Congurations 2.2.2 Analytical Models 2.3 Mainstream and Sealing Flow Interactions Chapter 3 Design of a Single-stage Axial Turbine Research Facility 3.1 Overview 3.2 Flow Path Configurations 3.3 Powertrain and Carriage System 3.4 Test Section Configuration 3.4.1 Stage Design 3.4.2 Wheel-space Geometries 3.5 Material Selections 3.6 Structural Analysis 3.7 Machining and Assembly Features 3.7.1 Tolerance and Surface Roughness Control 3.7.2 Balancing and Bearing Selection 3.8 Instrumentations 3.8.1 Mainstream Annulus and Secondary Flow Line 3.8.2 Wheel-space 3.8.3 Data Acquisition System 3.9 Sensor Calibrations and Uncertainty Analysis Chapter 4 Experimental Measurements on Double Rim Seal For Facility Validation 4.1 Test Configurations for Double Rim Seal 4.2 Sealing Effectiveness 4.3 Pressure and Velocity Measurements 4.3.1 Mainstream Pressure Asymmetries 4.3.2 Swirl Ratio Chapter 5 Study of Wheel-space Swirl Effects on Single Rim Seal Performance 5.1 Test Configurations for Single Rim Seal 5.2 Wheel-space Swirler Design 5.3 Sealing Effectiveness 5.4 Pressure and Velocity Measurements 5.4.1 Mainstream Pressure Asymmetries 5.4.2 Swirl Ratio and Wheel-space Pressure 5.4.3 Wheel-space Radial Velocity Chapter 6 Conclusion 6.1 Design of the Experimental Facility 6.2 Facility Validation 6.3 Wheel-space Swirl Effects on Sealing Performance 6.4 Proposal for Modified Orifice Model 6.5 Practical Implications 6.6 Scaling to Engine Conditions 6.7 Future Works Bibliography Appendix A Owen's Orifice Model 국문초록Maste

    Energy efficient engine high-pressure turbine detailed design report

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    The energy efficient engine high-pressure turbine is a single stage system based on technology advancements in the areas of aerodynamics, structures and materials to achieve high performance, low operating economics and durability commensurate with commercial service requirements. Low loss performance features combined with a low through-flow velocity approach results in a predicted efficiency of 88.8 for a flight propulsion system. Turbine airfoil durability goals are achieved through the use of advanced high-strength and high-temperature capability single crystal materials and effective cooling management. Overall, this design reflects a considerable extension in turbine technology that is applicable to future, energy efficient gas-turbine engines

    The Aero-Thermal Performance of Purge Flow and Discrete Holes Film Cooling of Rotor Blade Platform in Modern High Pressure Gas Turbines: A Review

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    Design of cooling systems for rotor platforms is critical due to the complex flow field and heat transfer phenomena related to the secondary flow structures originating at the blade leading edge. Horseshoe vortex and passage vortex are the fluid-dynamic features that largely influence the aerodynamic behaviour and the thermal protection level of the platform. The driving parameter is the coolant to mainstream momentum flux ratio, but several issues have to be considered in the design process of cooling technologies. As well acknowledged, an in-depth understanding of losses and heat transfer phenomena are deemed necessary to design effective cooling systems. In the present review, measurements and predictions on the behaviour of the HPT rotor cooled platform, obtained during the last two decades by several research groups, are gathered, described and analysed in terms of aerodynamic losses and heat transfer performance, and are compared with one another with respect to the effectiveness level that is ensured
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