bluuebunny/crossref_metadata_2025_split huggingface dataset:
Cost-Effective CFD Analysis of the Acoustic Response of a Perforated Plate
Authors: Dimola Nunzio , Stefanizzi Michele , Capurso Tommaso , Schuller Thierry , Torresi Marco , Camporeale Sergio Mario 
Published: October 2022

Abstract
Given the current practice to perform lean-premixed combustion to decrease NOx emissions, thermoacoustic instabilities have become one of the major drawbacks in gas turbine combustors. The necessity to control and limit such a deleterious phenomenon is mandatory to avoid structural damage of the burner. It has been demonstrated that perforated liners, if conveniently designed, can be very effective in reducing acoustic oscillations inside gas turbine combustors. Studying perforated plates traversed by bias flow can give a useful insight on sound absorption properties of liners, rather than investigate complex geometries. The present paper aims to carry out a numerically cost-effective, but reliable, CFD analysis to predict the acoustic impedance of perforated plates traversed by bias flow, and to grasp the details of the sound dissipation process. 2D axisymmetric simulations have been carried out and the governing equations solved by using the commercial code ANSYS Fluent®. Hypotheses, boundaries and operating conditions are described, focusing on the role of the Non-Reflecting-Boundary-Condition (NRBC) and the Transparent-Flow-Forcing condition (TFF) in treating acoustic waves. Numerical results are compared both with linear analytical models and experimental data from a case study, by proving a fast and reliable prediction of the acoustic response. Furthermore, effects of increasing bias flow temperature on the sound absorption property have been investigated, showing an increase in acoustic power losses as temperature rises. The proposed CFD model (2D-axisymmetric) proved to be a valid and versatile tool in evaluating the acoustic response of perforated plates under different operating conditions.
Investigation of Surge in a Transonic Centrifugal Compressor With Vaned Diffuser: Part 1 – Surge Signature
Authors: Lou Fangyuan , Harrison Herbert M. , Key Nicole L. 
Published: October 2022

Abstract
The identification of stall inception mechanisms and stability-limiting components in a centrifugal compressor is required for development of effective surge suppression approaches. Part 1 of this two-part paper investigates the surge signature of a centrifugal compressor at subsonic, transonic, and supersonic inlet tip conditions, and Part 2 considers the relationship of the surge signature with the compressor static pressure characteristics as well as the impeller leading edge relative tip Mach number. Experiments were performed in the Single Stage Centrifugal Compressor (SSCC) facility at Purdue University with both steady performance and dynamic pressure data recorded. Results show the presence of both mild and deep surge on the compressor map. Deep surge occurs at subsonic and supersonic impeller tip relative Mach numbers while mild surge is observed as the tip relative Mach number nears unity. Long length-scale modal oscillations can be detected only at the transonic operating condition while spike disturbances are identified as the precursor to instability at subsonic and supersonic tip relative Mach numbers. Finally, the impeller is determined to be the origin of instability throughout most of the compressor map. The observation of impeller-induced instability refreshes the conventional understanding that the diffuser is typically the stability-limiting component for centrifugal compressors with vaned diffusers and suggests various surge suppression approaches may be necessary to achieve range extension throughout a compressor map.
Investigation of Surge in a Transonic Centrifugal Compressor With Vaned Diffuser: Part 2 – Correlation With Subcomponent Characteristics
Authors: Lou Fangyuan , Harrison Herbert M. , Brown William J. , Key Nicole L. 
Published: October 2022

Abstract
In the first part of the paper, surge signatures and the underlying mechanisms thereof for a high-speed centrifugal compressor were investigated using signals from fast-response transducers installed along the flow path. The compressor surge signature is observed to vary as the impeller inlet tip flow transitions from subsonic to supersonic conditions. Spike-type deep surge is observed at subsonic and supersonic impeller inlet tip conditions while modal-type mild surge occurs at 90% speed with transonic inlet tip conditions. In Part 2 of the paper, a detailed analysis of the static pressure rise characteristics at the stage, component, and subcomponent levels is conducted. Additionally, the influence of the impeller inlet conditions on the instability are also considered. A connection between the surge mechanism and component static pressure rise characteristics is shown which exhibits the potential for prediction of the stall inception mechanism including identification of the destabilizing component. Furthermore, the transition from subsonic to transonic impeller inlet conditions is shown to be the cause of the unique mild surge instability at 90% speed. Lastly, both experimental and computational results are utilized to investigate the development of a shock wave at the impeller leading edge. The shock is considered to be critical to stage stability at speeds where the compressor experiences transonic inlet conditions.
Performance Assessment of Tesla Expander Using 3D Numerical Simulation
Authors: Renuke Avinash , Traverso Alberto 
Published: October 2022

Abstract
Bladeless or Tesla turbines consist of several flat parallel disks mounted on a shaft with a narrow gap between them. The analytical solution of Navier-Stokes equations for the flow between disks has been extensively studied in the past. However, there is a significant impact of the stator exit-flow conditions on the performance and flow behavior inside the rotor of the Tesla turbine. There has been limited research on flow characterization and performance evaluation of stator-rotor interaction of the Tesla expander using 3D numerical simulation. The challenge arises due to a very high aspect ratio of the Tesla rotor (diameter to gap ratio &gt; 1000). In order to accurately evaluate the torque on the disks due to shear forces, a very fine resolution mesh is necessary. 3D numerical modeling with a hexahedral mesh of the nozzle/stator with the commercial software is presented. A steady-state solution is obtained using a density-based solver for solution stability. The simulation is performed for a wide range of inlet pressures and rotational speeds. The numerical solution does not consider the effects of ventilation losses, end-disk leakages, exit kinetic energy and bearing losses. This work focuses on the stator performance, the stator-rotor interaction, the rotor entry losses due to disk tip, the rotor tip velocity ratio and the degree of reaction on the performance of the Tesla expander. Challenges in modelling and key fluid dynamic features are extensively discussed. The peak efficiency of 58% is predicted for 3 bar inlet pressure and a rotational speed of 30000 rpm for a 3 kW machine with air as a working fluid. The 3D numerical analysis provides insights on flow characterization, mainly stator-rotor interaction and flow between disks at different mass flows, with the aim to contribute to the fundamental knowledge towards further improvement of the Tesla turbine performance. Numerical results are also compared with experimented test results performed on a 100-W and a 3-kW air expander.
Performance Optimization of Semi-Closed Oxy-Combustion Combined Cycle (SCOC-CC) for Current and Future Blade Materials
Authors: Risimini Gabriele Pio , Martinelli Matteo , Chiesa Paolo , Martelli Emanuele 
Published: October 2022

Abstract
Among the technologies for carbon capture and storage (CCS) from natural gas, oxy-turbine plants are a very promising solution thanks to the high efficiency, absence of stack and nearly 100% capture rate. This paper investigates the efficiency which can be achieved by the Semi-Closed Oxy-Combustion Combined Cycle (SCOC-CC) with state-of-the-art and future blade materials. In particular, the analysis considers class-H turbine super-alloys with a maximum blade wall temperature of 900 °C and Ceramic Matrix Composites with blade wall temperatures of 1300 °C. Sensitivity analyses are performed to determine the optimal pressure ratio and turbine inlet temperature. The results indicate that state-of-the-art superalloys allow the SCOC-CC to achieve 54% net electric efficiency with 96% carbon capture rate, while CMC blades boost the efficiency up to 60%. For both cases, critical factors are the high temperature gradients across the blade coatings (TBC for superalloy, EBC for CMC) and the blade thickness caused by the large heat flux exchanged between hot gases and cooling flows.
Turbine Vane Passage Experiments Documenting Evolution of Secondary Flows With Changes in Combustor Coolant Injection Flowrates
Authors: Nawathe Kedar P. , Kim Yong W. , Simon Terrence W. 
Published: October 2022

Abstract
A secondary flow system with a dominant passage vortex pattern has been observed in many gas turbine vane passage studies in which there is no upstream coolant injection or only near-passage endwall coolant injection (no combustor cooling). However, it was shown in recent studies that combustor coolant introduced upstream of the vane passage changes secondary flow patterns in the passage. This results in a different secondary flow vortex system, called the ‘impingement vortex’ system. It was discussed in recent literature having combustor coolant injection. Until now, there has been no study on how increases in combustor coolant momentum effect transition from the passage vortex system to the impingement vortex system. Such a study is presented in the present paper. Velocity component measurements are taken using a five-hole probe at three axial locations in the vane passage to document secondary flow development throughout the passage. Four combustor coolant flowrate cases are considered along with a comparison case having no coolant injection. It is shown that as the combustor coolant flowrate increases, the passage vortex system weakens and, at a sufficiently high combustor coolant flowrate, the impingement vortex system appears. Knowing the detailed flow physics of this transition between the two secondary flow systems is helpful for turbine thermal designers who wish to understand how secondary flows transport coolant within the turbine vane passage.
Improving Gas Turbine Maintenance Quality
Authors: Angello Leonard , Scheibel John , Noble David , Smith Nick 
Published: October 2022

Abstract
Planned maintenance outages account for over 70% of gas turbine unit unavailability. A complete overhaul of a large frame-sized gas turbine can require up to 10–12 weeks. Improper reassembly procedures and wrongly applied maintenance practices can ultimately result in catastrophic failure and loss of human life. To address owner/operator growing concerns regarding the quality of gas turbine planned maintenance outages, EPRI in conjunction with its member companies has developed and demonstrated key tools, methods, and guidance for improving the disassembly, inspection, reassembly, and re-commissioning gas turbine units during planned outages, including:
• Key maintenance activity check lists that include disassembly and reassembly hold points, verification points, and witness points with acceptance criteria
• Inspection techniques and quality control Inspection Assessment Data Sheet (IADS) forms for the rotating, stationary, combustion, and structural turbine parts. Each IADS includes a sketch or photo of the part, expected dimensions, typical findings, and field repair methods with GO- or NO-GO acceptance criteria.
• Companion field guidance (FG) for each IADS, with recommendations for typical damage findings
• Reassembly hold points, verification points, and witness points with acceptance criteria (both typical and not-to-exceed criteria)
• Application of PCRT testing for part quality control verification
This paper discusses outages in detail and giving insight on best practices and how advances in diagnostics and knowledge transfer are improving these activities for utilities.
Insights Into the Unsteady Shock Boundary Layer Interaction
Authors: Hergt A. , Klinner J. , Willert C. , Grund S. , Steinert W. 
Published: October 2022

Abstract
The flow through a transonic compressor cascade is characterized by high unsteadiness due to the shock boundary layer interaction. Investigations in recent years have shown that a detailed understanding of the causes of unsteady shock oscillation is necessary to develop successful approaches to influence it. Therefore, an experimental investigation of the unsteadiness of the shock boundary layer interaction in a transonic compressor cascade has been conducted within the transonic cascade wind tunnel at DLR in Cologne. At an inflow Mach number of 1.05, detailed measurements were carried out with a time-resolved PIV system in combination with a high-speed shadowgraphy setup. In this way it was possible to simultaneously measure both the shock movement and the flow field of the boundary layer under the shock. The analysis of the measured data showed a correlation between the oscillation behaviour of the passage shock and the unsteady flow behaviour within the boundary layer in front of the shock. In the shock oscillation spectra a dominat frequency at 1683 Hz and their first harmonic was found. This frequencies are also be found in the boundary layer flow below and in front of the shock with different amplitutes at three analysing points in the measured PIV Region. A detailed analysis of the measured data shows that the information of the unsteady shock oscillation propagates under the shock foot over the boundary layer upstream. It becomes clear that the propagation of the oscillating pressure information has an influence on the velocity component normal to the blade surface. This leads to a oscillating flow angle close to the blade. Through this effect, the inflow in itself interacts with the shock front and influences the shock position and structure. Based on this, a new thesis of self-exciting shock oscillation is developed. In addition, the used time-resolved PIV measurement enables an acquisition of the blade vibration behaviour. Within the results of the blade vibration four Eigenmodes are observable. In this context it has been shown that the Eigenmodes of the blades are not stimulated by the flow. On the other hand there is also no exciting interaction of the blades with the flow detectable. The measured data of transonic flow within a compressor cascade presented here are unique and provide new insight into shock movement and interaction with the boundary layer.
Using Blade Surface Heat Transfer for Partial Intercooling and Thrust Augmentation
Authors: Gannon Anthony J. , Morgan Thomas , Smith Walter C. , Hobson Garth V. 
Published: October 2022

Abstract
Gas turbine engines are widely used prime movers in aircraft. A thermodynamic tool to analyze the improvement to aircraft high bypass ratio engines using existing aerodynamic stator and outlet guide vane surfaces for heat exchange to achieve decreased specific fuel consumption is presented. The potential benefits are twofold, heat removed from the high pressure compressor stator blade rows reduce the compressor work while rejecting this heat to expand the bypass flow increases the thrust. Concepts for recuperated and intercooled gas turbine engines for aviation have been investigated but these usually make use of additional heat exchange devices through the cycle. While these approaches are successful on land based systems, weight and integration issues limit their benefit in aviation. While previous work has shown that the equivalent heat exchanger effectiveness, ε, is only expected to be 0.2 but this is offset by the limited weight gain and potentially simpler integration into a flight worthy system. A system to transfer the heat between blade rows is required and a heat pipe or heat pipe loop is proposed. The method is applied to a representative engine where the decreased compressor and increased bypass flow thrust is shown to improve the specific fuel consumption by 2.99% for the static thrust case and 9.34% at cruise conditions.
Sensitivity of Adiabatic Cooling Effectiveness to Mainstream and Coolant Temperatures in Transonic Flow Over an Idealized Blade Tip Model
Authors: Ma Haiteng , Zeng Wei , Dai Siming , Jiang Hongmei 
Published: October 2022

Abstract
Film cooling, as a key technology to ensure turbine survival in new generation gas turbines, has been studied profusely in subsonic flows. But in transonic flow, the determination of adiabatic cooling effectiveness faces a dilemma in the state of the art. Specifically, derivation of reference temperature, or local recovery temperature, in adiabatic effectiveness is disputable. Some researchers designate it to be the adiabatic wall temperature for the uncooled model (linear regression method, LRM), but others calculate it from an iterative procedure based on a pair of cooling tests (dual linear regression method, DLRM). As the first of the kind effort to explore this dilemma, this paper carried out transient thermal measurements by infrared thermography, for transonic flow over an idealized blade tip model. Heat transfer experiments were conducted for the uncooled and cooled cases, at two mainstream temperatures of 340K and 325K, and two coolant temperatures of 276K and 287K. Data from these six experimental groups were processed by LRM and DLRM respectively, to obtain heat transfer coefficient and adiabatic effectiveness, whose sensitivity to mainstream and coolant temperatures is tested and compared. It is found that heat transfer coefficient is basically insensitive to temperature boundary conditions and data reduction methods, as expected. However, for adiabatic effectiveness, LRM results are sensitive to the 11K decrease of coolant temperature in areas confined to the upstream of cooling injection, and much less so to the 15K rise in mainstream temperature. DLRM result, derived from the test pair with two coolant temperatures, reduces globally and conspicuously with 15K increase in mainstream temperature. Furthermore, adiabatic effectiveness obtained by LRM is qualitatively different to that by DLRM, which is mainly attributed to the large discrepancy of reference temperature between the two methods.
Wall Curvature Effect on Overall Thermal Performances of Film Cooling Covered by Thermal Barrier Coatings With Various Geometries
Authors: Zhou Wen-li , Pu Jian , Zhang Tiao , Wang Jian-hua , Wu Wei-long , Su Hang 
Published: October 2022

Abstract
Film cooling coupling with surface thermal barrier coating (TBC) has been widely adopted, bringing to an intensive need of comprehensive overall thermal characteristics. In present work, combined effects of hole-geometry, wall curvature, and cooling air flowrate on the metal overall effectiveness and wall heat flux were discussed. Typical local regions of turbine vanes were simplified as convex, concave, and flat walls. Spraying TBC generated four hole-configurations, including Standard Cylindrical-hole (SC), transverse Trenched-hole (TT-hole), and two modified trenched-holes with Semicircle-shaped Trench (ST) and Sine Wave-shaped Trench (SWT). The ST reduces the exposed area and the SWT improves the fluid-diversion effect of TT. Based on the engine-matched hot-side Biot-numbers, overall cooling effectiveness measurements were carried out under typical blowing ratios (BRs) from 0.5 to 2.0, through capturing wall temperatures by an infrared thermal sensor and thermocouples. Conjugate heat transfer simulations were also conducted, showing good agreement with the experimental data and hence providing additional insights into the heat transfer features and film cooling jet behaviors. The results indicated that the wall curvature effects on overall thermal parameters are relatively weaker, in comparison with the hole-geometry and BR. The trenches can further weaken the wall curvature effect. Comparisons of trench-designs reveals that achievement of strong fluid-diversion in trench is an important principle of trench-modification, which can acquire not only superior overall cooling characteristics, but also a relatively large ability to prevent the hot gas ingestion and the hole-blockages by harmful contaminant-depositions.
Investigations on Cooling Hole Patterns Over a Turbine Endwall for Improving Cooling Effectiveness
Authors: Yang Xing , Zhao Qiang , Wu Hang , Feng Zhenping 
Published: October 2022

Abstract
Redistributions of film hole patterns over the endwall region of a turbine vane are investigated in this work, for providing effective coverage for the entire endwall surface. Air enters the linear cascade with a turbulence intensity of 10.8% that is measured in front of the cascade inlet and exits the cascade passage with an averaged Reynolds number of 5.52 × 105. Three film injection patterns are employed: a baseline scheme in which coolant is injected from twelve pressure-side (PS) and four suction-side (SS) discrete holes as well as an upstream slot, a curtain cooling scheme in which two rows of discrete holes (totally thirteen holes) are included in the PS region at the passage inlet, and a hole-redistributed scheme in which the number of the PS holes is reduced and the PS and SS holes are redistributed. Film cooling effectiveness over the endwalls cooled by the three injection patterns is measured using the pressure-sensitive paint (PSP) technique for a density ratio of 1.5 and typical coolant flow rates found in a real engine. Additionally, aerodynamic data is measured by a five-hole probe at the cascade exit to demonstrate the effects of the injection from different hole patterns on flow fields throughout the cascade passage. Comparisons among the three hole patterns reveal that by increasing coolant rates, cooling effectiveness from the slot injection is significantly increased while that from the curtain injection upstream the passage inlet and hole injection within the passage is slightly varied. Although the slot injection is an effective cooling source, it leaves the PS region of the endwall passage uncovered, which can be well alleviated by using the curtain cooing instead of the PS hole injection. Overall, the curtain cooling provides better cooling effectiveness at an equal or even lower coolant flow rate, compared with the baseline configuration, and the hole-redistributed scheme generates better local cooling performance.