bluuebunny/crossref_metadata_2025_split huggingface dataset:
Experimental Flow Field Investigation in a Centrifugal Compressor Vaned Diffuser
Authors: Pinarbasi Ali , Johnson Mark W. 
Published: February 2015

The purpose of this study was to improve the understanding of the flow physics in a centrifugal compressor vaned diffuser. A low speed compressor with a 19 bladed backswept impeller and diffuser with 16 wedge vanes was used. The measurements were made at three inter-vane positions and are presented as mean velocity, turbulent kinetic energy and flow angle distributions on eight diffuser cross sectional planes.
The impeller blade wakes mix out rapidly within the vaneless space and more rapidly than in an equivalent vaneless diffuser. Although the flow is highly non uniform in velocity at the impeller exit, there is no evidence in the results of any separation from the diffuser vanes. The results do however suggest that the use of twisted vanes within the diffuser would be beneficial in reducing losses.
Theoretical Modelling of Relative Wall Motion Effects in Tip Leakage Flow
Authors: Nikolos I. K. , Douvikas D. I. , Papailiou K. D. 
Published: February 2015

The influence of relative wall motion in modifying the leakage flow through the tip clearance is investigated. A theoretical model is developed, in order to calculate the mass flow rate through the gap.
The physical mechanism by which relative wall motion affects the leakage flow is analyzed and the differences between the turbine and compressor case are identified. This model, being an extension of an already existing one, not taking into account relative wall motion, is incorporated into the tip clearance calculation procedure, already developed by the authors.
Theoretical results of the complete calculation procedure (secondary flow plus tip clearance model) are compared with experimental data, for the case of compressor and turbine cascades, as well as for the case of a single rotor. Good agreement between theory and experiment is obtained.
Rotor Blade-to-Blade Measurements Using Particle Image Velocimetry
Authors: Tisserant Denis , Breugelmans Frans A. E. 
Published: February 2015

The study of turbomachinery flow fields requires detailed experimental data. The rotating parts of turbomachines greatly limit the measurement techniques which can be used. Particle Image Velocimetry (PIV) appears to be a suitable tool to investigate the blade-to-blade flow in a rotor.
The facility is a subsonic axial-flow compressor. The experimental apparatus enables the recording of a double exposed photograph in a circumferential plane located at 85 % of the blade height. The illumination plane has an axial direction and is provided by a pulsed Ruby laser. The tracers used are sub-micron glycerine oil droplets. Data are processed by the Young’s fringes method.
Measurements were performed at 3000, 4500 and 6000 RPM with velocities in the range of 30 to 70 m/s. Steady operating conditions are chosen in such a way that the effect of radial velocity on PIV measurements can be neglected.
Experimental problems which are encountered included homogeneous seeding of the flowfield and laser light scattering from blade surfaces. The uncertainty affecting the velocity determination corresponds to 2 % of the measured value.
For a given set of operating conditions, 10 PIV pictures are recorded. The periodic flow field is approximated by averaging point by point the experimental data. Upstream and downstream velocity triangles are confirmed by measurements obtained from pressure probes. PIV measurement results were found to be similar to those of a blade-to-blade potential-flow calculation.
Surface Effects on Deposits From Jet Fuels
Authors: Ervin J. S. , Heneghan S. P. , Martel C. R. , Williams T. F. 
Published: February 2015

Flow experiments in a single-pass heat exchanger using JP-8 and certain additives were initiated under controlled conditions to explore the effects of a metal surface on deposition. The experimental apparatus permitted a unique viewing of the time evolution of deposits at different axial locations under conditions of limited oxygen availability somewhat similar to that in jet aircraft. Scanning electron microscopy was used to examine deposit microstructure. In addition, x-ray photoelectron spectroscopy and Auger electron spectroscopy determined the chemical composition of the deposits. Oxygen concentration measurements in the bulk flow were also performed, and the observed transient oxidation behavior was related to measured time-dependent changes in the deposit.
Increasing dissolved oxygen levels and large changes in deposition were characteristic of the induction time. Mechanisms of fouling in the heated and cooled sections were different. Spectroscopic analysis indicated that deposits formed in the heated section had chemical compositions different from those formed in the cooled section. Scanning electron microscopy revealed differences in microstructure between the heated and cooled sections: more uniform deposits formed in the cooled section as a result of once-soluble species becoming insoluble at low temperatures. In addition, the JP-8 additives significantly reduced fouling in the heated section, but their effectiveness in the cooled section, especially after large periods, was unclear.
Variables Affecting NOx Formation in Lean-Premixed Combustion
Authors: Steele R. C. , Jarrett A. C. , Malte P. C. , Tonouchi J. H. , Nicol D. G. 
Published: February 2015

The formation of NOx in lean-premixed, high-intensity combustion is examined as a function of several of the relevant variables. The variables are the combustion temperature and pressure, fuel-type, combustion zone residence time, mixture inlet temperature, reactor surface-to-volume ratio, and inlet jet size. The effects of these variables are examined by using jet-stirred reactors and chemical reactor modeling.
The atmospheric pressure experiments have been completed and are fully reported. The results cover the combustion temperature range (measured) of 1500 to 1850K, and include the following four fuels: methane, ethylene, propane, and carbon monoxide/hydrogen mixtures. The reactor residence time is varied from 1.7 to 7.4ms, with most of the work done at 3.5ms. The mixture inlet temperature is taken as 300 and 600K, and two inlet jet sizes are used. Elevated pressure experiments are reported for pressures up to 7.1atm for methane combustion at 4.0ms with a mixture inlet temperature of 300K. Experimental results are compared to chemical reactor modeling. This is accomplished by using a detailed chemical kinetic mechanism in a chemical reactor model, consisting of a perfectly stirred reactor (PSR) followed by a plug flow reactor (PFR). The methane results are also compared to several laboratory-scale and industrial-scale burners operated at simulated gas turbine engine conditions.
Combustor Stability and Emissions Research Using a Well Stirred Reactor
Authors: Zelina J. , Ballal D. R. 
Published: February 2015

The design and development of low-emissions, lean premixed aero or industrial gas turbine combustors is very challenging because it entails many compromises. To satisfy the projected CO and NOx emissions regulations without relaxing the conflicting requirements of combustion stability, efficiency, pattern factor, relight (for aero combustor) or off-peak loading (for industrial combustor) capability demands great design ingenuity. The well stirred reactor (WSR) provides a laboratory idealization of an efficient and highly compact advanced combustion system of the future that is capable of yielding global kinetics of value to the combustor designers. In this paper, we have studied the combustion performance and emissions using a toroidal WSR.
It was found that the toroidal WSR was capable of peak loading almost twice as high as that for a spherical WSR and also yielded a better fuel-lean performance. A simple analysis based upon WSR theory provided good predictions of the WSR lean blowout limits. The WSR combustion efficiency was 99 percent over a wide range of mixture ratios and reactor loading. CO emissions reached a minimum at a flame temperature of 1600K and NOx increased rapidly with an increase in flame temperature, moderately with increasing residence time, and peaked at or slightly on the fuel-lean side of the stoichiometric equivalence ratio. Finally, emissions maps of different combustors were plotted and showed that the WSR has the characteristics of an idealized high efficiency, low emissions combustor of the future.
Prediction and Measurement of the Total Pressure Loss in an Engine Representative Diffuser System
Authors: Little A. R. , Denman P. A. , Manners A. P. 
Published: February 2015

An experimental and computational investigation of the flow in an engine representative diffuser system has been performed. Many parametric changes to the system were considered including inlet conditions, pre-diffuser geometry, cowl geometry and pre-diffuser strut position. The focus of this paper is an overview of the results from the twelve configurations that were both predicted and experimentally tested.
It was shown that the CFD predictions contained a high degree of numerical error and that to reduce this to a low level would be currently impractical for a parametric study of this type. However, by carefully maintaining approximately the same degree of numerical error in the predictions, it was shown that a valid parametric study could be performed to the extent that the same conclusions about the parametric changes can be drawn equally from the predictions or measurements with one small exception. It was recognised that the level of agreement between the predicted and measured losses was to a degree fortuitous and that the k-ε turbulence model performs poorly for this type of flow.
Experimental Investigation of Boundary Layer Transition and Turbulence Structures on a Highly Loaded Compressor Cascade
Authors: Wunderwald Dirk , Fottner Leonhard 
Published: February 2015

Detailed measurements have been performed on a compressor cascade in order to obtain information about the overall performance, the state of the boundary layer, and the topology of turbulent boundary layers.
The analysis of profile pressure distributions and wake traverse measurements across the midspan section of the cascade blade provide information on the loss behaviour. Using surface-mounted hot-film gauges on the suction side of the measuring blade different transition phenomena have been investigated under the influence of various inlet flow conditions representative of engine operation. Extensive measurements with 3D-hot-sensor anemometry have been evaluated to show essential features of the turbulent boundary layer.
The results point out the dependence of turbulence characteristics, e.g. turbulent kinetic energy distribution and Reynolds stresses, on the inlet flow conditions and the upstream boundary layer development. The influence of free-stream turbulence intensity is discussed and the non-isotropy of the Reynolds normal stresses is presented.
Application of Aero-Engine Technology to Heavy Duty Gas Turbines
Authors: Janssen Manfred , Zimmermann Holger , Kopper Frederick , Richardson John 
Published: February 2015

A description of an advanced version of a Siemens heavy duty gas turbine is presented which is based on a collaboration between Siemens Power Generation and United Technology Corporation. The main objective of this collaboration was to utilize the R&D capabilities of Siemens Power Generation and Pratt and Whitney to improve an existing heavy duty gas turbine (V84.3) by increasing compressor and turbine efficiency, refining the cycle and increasing turbine inlet temperature.
This paper presents a description of the redesigned components, focusing on the features derived from aero engine technology. The main features of the advanced compressor are custom tailored blading to account for end wall boundary layer flow at the hub and casing wall, as advanced controlled diffusion airfoil (CDA) and a revised flowpath. These features result in both improvements in efficiency and surge margin over a wide operating range. The design features of the advanced turbine blading are 3D aerodynamics to reduce profile and secondary losses, aero-engine derived cooling configurations and use of single crystal materials to increase turbine durability.
The performance improvement predicted for the redesign was demonstrated in a back-to-back test with the baseline engine, a 60 Hz Model V84.3. Testing was conducted at the Siemens Berlin facility which is capable of acquiring extensive performance and durability related data for full scale engines at design point and off-design conditions. Test results showed that over a 5 percent heat rate improvement was achieved by the application of aero-engine technology to the baseline machine.
Conversion of an Aerodynamic Wind Tunnel to a Gas Turbine Combustor Test Stand
Authors: Cresci Daniel 
Published: February 2015

A Mach 12, blowdown-type, storage heated, aerodynamic test facility, previously adapted to provide Mach 3.5 ramjet engine propulsion testing, has been further modified to provide high pressure, gas turbine combustor emission testing capability. Pressure control and temperature control systems have been incorporated into the original facility to provide constant flow properties for run times up to 30 minutes. An on-line gas sample system has also been designed and integrated with the facility to provide continuous combustor emission information.
These features now combine with the existing pebble bed storage heater, blowdown air supply and heated fuel supply systems to provide very high Mach number aerodynamic simulation, intermediate Mach number aeropropulsion testing and low Mach number, full scale, gas turbine combustor emission testing capability. Pressures of 1500 psia can be achieved with temperatures up to 2000 degrees Fahrenheit and flow rates up to 50 lb/sec.
This unique facility is described, with emphasis on the facility modifications necessary to provide the gas turbine combustor emission testing capability. To appreciate the modifications made, some of the differences between Mach 12 aerodynamic and low Mach number, gas turbine combustor emission testing are discussed. Results from some recent gas turbine emissions tests are also provided.
Flow and Turbulence Survey for a Model of Gas Turbine Exhaust Diffuser
Authors: Desideri Umberto , Manfrida Giampaolo 
Published: February 2015

This paper presents the results of an extensive set of measurements on a model of an exhaust diffuser for gas turbines. The diffuser is of the straight-wall annular-axial type, typically employed in small-to-medium size gas turbines. It features six high-solidity struts, which support, in the real machine, one of the shaft bearings and have piping for oil supply inside.
The 35%-scale model has been tested on a special test stand developed at the University of Perugia, using the suction side of a centrifugal-flow industrial fan of suitable capacity. Inlet speed is around 80 m/s, allowing satisfactory accuracy for flow measurements and the similarity in terms of Reynolds number.
The instrumentation, the movement of the measurement point and data acquisition system were designed for automatic running of the tests. Both pneumatic and hot-wire or hot-film probes can be used on the same facility. The same wind tunnel, previous a quick replacement of the model with a probe calibration test section, can be used for calibration of both pneumatic and hot-wire/hot-film probes. A three hole directional pneumatic probe was used for stationary flow measurements to determine the global performance parameters of the model and a split-film probe was used to determine the turbulence characteristics.
For four test sections, contour plots are produced of average velocity components, flow angle and turbulence quantities as three components of the Reynolds stress tensor.
Numerical Investigations of Flow Pattern and Heat Transfer in a Rotating Cavity Between Two Discs of the Compressor of a Siemens KWU V84.3 Gas Turbine
Authors: Bohn Dieter , Krüger Uwe , Nitsche Klaus 
Published: February 2015

The rotor of modern gas turbines often consists of single discs forming air-filled rotating cavities. During stationary operation each disc in the compressor section is of nearly uniform temperature. This results from the radial heat conduction in the disc material and from the negligible axial temperature gradients between surface and air in the adjacent cavities. The situation changes rapidly during cold start-ups of the engine. The disc rims respond quickly to the temperature of the mainstream (500 to 600 K), whereas the average temperature of the massive hub section follows with some delay thus forming a radial thermal gradient. This induces a buoyancy-driven flow inside the cavity, which is superimposed by a controlled hot gas ingress. A defined amount of hot air flows radially inwards through the Hirth-type serration at the head of the discs, causes increased convection within the cavity and speeds up the thermal equilibration process in the discs.
Numerical investigations of the very complex flow situation have been carried out to get a better knowledge of both the flow-physics and the heat transfer from the hot fluid to the cold rotating wall.
A modern numerical Finite-Volume-Code with multiblock and body-fitted grid-options has been used to calculate three different cases: one cavity without hot gas ingress and two cases with two different mass flow parameters. The boundary conditions have been chosen in such a way that they cover real gas turbine conditions at the very beginning of the start-up. The most stringent case has been investigated, i. e. the head of the discs and the hot gas mass flow having the mainstream temperature while the discs in the hub region remain at ambient temperature.
It has been found that in the case without throughflow the core-region rotates approximately with the speed of a solid body. In the case of superimposed hot gas flow directed radially inwards, the flow has the character of a potential-vortex-flow, with exception of the regions near the wall. The hot gas is transported to the hub-region so that the heat transfer in this region is very large in the first period of the start-up-procedure.
Some aspects are presented which should be investigated in more detail in future work, especially the 3-D effects and the conjugate heat transfer. First results of a 3-D calculation are shown.
Combustor Rebuild Effects on Gas Turbine Performance
Authors: MacLeod J. D. , Barry B. 
Published: February 2015

The Institute for Aerospace Research of the National Research Council of Canada (NRCC), has established a program for the evaluation of component deterioration on gas turbine engine performance. The effort is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. As part of this project, a study of the effects of combustor rebuild tolerances on overall engine performance was undertaken. This study investigated the range of performance changes that might be expected for simply disassembling and reassembling the combustor module of a gas turbine engine, and how these changes would influence the results of any component modification testing.
To evaluate the effects of rebuilding the combustor on the performance of a single spool engine, such as an Allison T56 turboprop engine, a series of three rebuilds was carried out. This study was performed in a similar way to two previous NRCC studies on the effects of compressor and turbine rebuilding. While the compressor and turbine rebuild studies found performance changes in the order of I% on various engine parameters, the effects of rebuilding the combustor have proven to be of similar magnitude. Based on the results of the combustor rebuild study, new methods to improve the assurance of the best possible tolerances during the rebuild process are currently being addressed.
This paper describes the project objectives, the experimental installation, and the results of the performance evaluations. Discussed are performance variations due to combustor rebuilds on engine performance characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.
A Theoretical and Numerical Development of the Concept of the Active Control Foil Bearing (ACFB)
Authors: Braun M. J. , Choy F. K. , Dzodzo M. , Hsu J. , Veillette R. , Deckler D. 
Published: February 2015

The authors of the present paper envisage far reaching applications of the foil bearing technology when merged with the concept of active control. Thus, this paper proposes and presents a comprehensive procedure for feedback control design for an active foil bearing. Two types of controllers are introduced, and discussed: i) the active feedback control for online adjustment, and ii) the supervisory control for anticipatory event management. A numerical procedure to simulate the steady state and transient dynamics of a foil bearing was also presented.
Successful introduction of such a new type of bearing will allow: (a) tolerance and ‘on-line’ adaptation to large time-dependent distortions and misalignment, (b) continuous adjustment for enhanced stability and resilience to high speed rubs, (c) reduced wear and power loss, (d) relative immunity from foreign object damage (FOD), (e) high load capacity due to thicker film lubricant, (f) lighter weight and reduced spatial envelope over active magnetic bearings, and (g) inherent stability without feedback control as compared with the magnetic bearing which has no stable equilibrium without feedback.
Stall Inception Measurements in a High-Speed Multi-Stage Compressor
Authors: Escuret J. F. , Garnier V. 
Published: February 2015

This paper presents unsteady measurements taken in a high-speed four-stage aero-engine compressor prior to the onset of aerodynamic flow instabilities. In this experiment, forty fast-response pressure transducers have been located at various axial and circumferential positions throughout the machine in order to give a very detailed picture of stall inception.
At all the compressor speeds investigated, the stall pattern observed is initiated by a very short length-scale finite-amplitude disturbance which propagates at a fast rate around the annulus. This initial stall cell leads to a large-amplitude system instability in less than five rotor revolutions. Varying the IGV setting angle is found to have a strong influence on the axial location of the first disturbance detected. In particular, transferring the aerodynamic loading from front to downstream stages moves the first disturbance detected from the first to the last stage of the compressor. Other repeatable features of the stall inception pattern in this compressor have been identified using a simple analysis technique particularly appropriate to the study of short length-scale disturbances. It is found that the origins of instabilities are tied to particular tangential positions in both the stationary and rotating frames of reference.
These measurements lead to the conclusion that the stall inception process in high-speed multi-stage compressors can be characterised by some very local and organised flow phenomena. Moreover, there is no evidence of pre-stall waves in this compressor.