Near-Wall Flow in Turbomachinery Cascades—Results of a German Collaborative Project

This article provides a summarizing account of the results obtained in the current collabora-tive work of four research institutes concerning near-wall flow in turbomachinery. Specific questions regarding the influences of boundary layer development on blades and endwalls as well as loss mech-anisms due to secondary flow are investigated. These address skewness, periodical distortion, wake interaction and heat transfer, among others. Several test rigs with modifiable configurations are used for the experimental investigations including an axial low speed compressor, an axial high-speed wind tunnel, and an axial low-speed turbine. Approved stationary and time resolving measurements techniques are applied in combination with custom hot-film sensor-arrays. The experiments are complemented by URANS simulations, and one group focusses on turbulence-resolving simulations to elucidate the specific impact of rotation. Juxtaposing and interlacing their results the four groups provide a broad picture of the underlying phenomena, ranging from compressors to turbines, from isothermal to non-adiabatic, and from incompressible to compressible flows.The investigations reported in this article were conducted within the framework of the joint research project “Near-Wall Flow in Turbomachinery Cascades” which was funded and supported by the Deutsche Forschungsgemeinschaft (DFG) under grant number PAK 948. The responsibility for the contents of this publication lies entirely by the authors.Peer ReviewedPostprint (published version

Testing of a Microwave Blade Tip Clearance Sensor at the NASA Glenn Research Center

The development of new active tip clearance control and structural health monitoring schemes in turbine engines and other types of rotating machinery requires sensors that are highly accurate and can operate in a high temperature environment. The use of a microwave sensor to acquire blade tip clearance and tip timing measurements is being explored at the NASA Glenn Research Center. The microwave blade tip clearance sensor works on principles that are very similar to a short range radar system. The sensor sends a continuous microwave signal towards a target and measures the reflected signal. The phase difference of the reflected signal is directly proportional to the distance between the sensor and the target being measured. This type of sensor is beneficial in that it has the ability to operate at extremely high temperatures and is unaffected by contaminants that may be present in turbine engines. The use of microwave sensors for this application is a new concept. Techniques on calibrating the sensors along with installation effects are not well quantified as they are for other sensor technologies. Developing calibration techniques and evaluating installation effects are essential in using these sensors to make tip clearance and tip timing measurements. As a means of better understanding these issues, the microwave sensors were used on a bench top calibration rig, a large axial vane fan, and a turbofan. Background on the microwave tip clearance sensor, an overview of their calibration, and the results from their use on the axial vane fan and the turbofan will be presented in this paper

HIGH-SPEED ROTOR TIP CLEARANCE MEASUREMENTS IN A TRANSONIC COMPRESSOR

Performance of a gas turbine compressor is directly dependent on the size of the region between the rotor blade’s tips and the surrounding casing, the tip clearance, which dynamically changes with rising rotor speed due to rotor blade radial growth from centrifugal loading. Too large a tip clearance introduces disruptive air flow that will lower compressor efficiency and lead to stall conditions, whereas too small a tip clearance will increase the risk of blade tip rubbing with the casing inner wall and may lead to catastrophic failure. This experiment is a part of a program of research that characterizes the Naval Postgraduate School Military Fan (NPSMF) in the Turbopropulsion Lab’s (TPL) Transonic Compressor Rig (TCR). This study involves the design, creation, and use of two benchtop rigs with a capacitive proximity probe blade tip clearance measurement system to develop mathematical methods to post-process capacitive probe output signals for calibration and tip clearance measurements. The mathematical methods developed in this study are validated against the tip clearance measurement system manufacturer’s method, showing improvement. A comparison of the different calibration rigs’ resulting calibration curves is discussed. The post-process method is then applied to high-speed tip clearance measurements of the NPSMF in the TCR and the results are compared to a model.Office of Naval Research, Arlington, VAOutstanding ThesisLieutenant, United States NavyApproved for public release. Distribution is unlimited

On the Aerothermal Flow Field in a Transonic HP Turbine Stage with a Multi-Profile LP Stator Vane

The quest for higher performances and durability of modern aero-engines requires the understanding of the complex aero-thermal flow experienced in a multi-row environment. In particular, the high and low pressure turbine components have a great impact into the engine overall performance, and improvements in the turbine efficiencies can only be achieved through detailed research on the three-dimensional unsteady aerodynamics and heat transfer. The present thesis presents an experimental study of the aerothermodynamics in one and a half turbine stage, focusing on: the aero-thermal flow in the overtip region of a transonic highly loaded high pressure (HP) rotor, and the aerodynamics and heat transfer of an innovative low pressure (LP) stator with a multi-profile configuration placed downstream of the high pressure turbine, within an s-shaped duct. Advanced instrumentation and measurement techniques were used and developed to perform the experimental investigation in a short-duration turbine test rig where both high spatial and time accuracy is indispensable. The flow field at the rotor shroud was investigated with simultaneous measurements of heat transfer, static pressure and blade tip clearance by using fast response pressure, wall temperature and capacitance probes. Through repeat experiments at the same turbine operating point, the time-averaged and time-resolved adiabatic wall temperature and convective heat transfer coefficient were evaluated. In the frame of new engine architectures, a novel stator for an LP turbine is proposed with a multi-splitter layout that represents a new design solution towards compact, lighter and performing aero-engine turbomachinery. It contains small aero-vanes and large structural aerodynamic airfoils which are used to support the engine shaft and house service devices. The research focuses on the experimental investigation of the global performance, aerodynamics and thermodynamics of this novel HP-LP vane layout. The turbine wasLavagnoli, S. (2012). On the Aerothermal Flow Field in a Transonic HP Turbine Stage with a Multi-Profile LP Stator Vane [Tesis doctoral no publicada]. Universitat Politècnica de València. https://doi.org/10.4995/Thesis/10251/17799Palanci

Tip casing heat transfer measurements of a film-cooled turbine stage in a short duration facility

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.Includes bibliographical references (p. 139-140).by Bret P. Van Poppel.S.M

가스터빈 림 씰 성능에 휠스페이스 스월이 미치는 영향 측정

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

An ultrasonic flowmeter for gases by Donald A. Bender, Leon R. Glicksman, Carl R. Peterson.

An ultrasonic flowmeter is developed for use in natural gas mains. The characteristics of the application and the dynamic head device presently employed are described. The performance requirements, design, and prototype testing of the ultrasonic instrument are discussed. The viability of a unique metering technique using reflected acoustic pulses was experimentally demonstrated. The flowmeter developed herein requires access to one side of the gas line and is self calibrating. It was concluded that continued development will produce a unit suitable for use in commercial service

Fiber optic (flight quality) sensors for advanced aircraft propulsion

Development of flight prototype, fiber-optic sensing system components for measuring nine sensed parameters (three temperatures, two speeds, three positions, and one flame) on an F404-400 aircraft engine is described. Details of each sensor's design, functionality, and environmental testing, and the electro-optics architecture for sensor signal conditioning are presented. Eight different optical sensing techniques were utilized. Design, assembly, and environmental testing of an engine-mounted, electro-optics chassis unit (EOU), providing MIL-C-1553 data output, are related. Interconnection cables and connectors between the EOU and the sensors are identified. Results of sensor/cable/circuitry integrated testing, and installation and ground testing of the sensor system on an engine in October 1993 and April 1994 are given, including comparisons with the engine control system's electrical sensors. Lessons learned about the design, fabrication, testing, and integration of the sensor system components are included