bluuebunny/crossref_metadata_2025_split huggingface dataset:
Heat Transfer to Rotating Turbine Blades in a Flow Undisturbed by Wakes
Authors: Garside T. , Moss R. W. , Ainsworth R. W. , Dancer S. N. , Rose M. G. 
Published: February 2015

The flow over the high pressure blades of a gas turbine is disturbed by wakes and shock waves from the nozzle guide vanes upstream. These disturbances lead to increased heat transfer to the blade surfaces, the accurate prediction of which is an essential stage in the design process.
The Oxford Rotor experiment consists of a highly instrumented 0.5 m diameter shroudless turbine which is supplied with air from a piston tube during the 200 ms run time and simulates realistic engine Mach and Reynolds numbers. Previous experiments have measured blade surface pressures and heat transfer rates, and compared them with similar data from linear cascades.
The present work is designed to enable the accuracy of rotation terms in computational fluid dynamics (CFD) calculations to be assessed, by providing heat transfer data from the rotating frame in the absence of wakes. Flow disturbances were avoided by removing the nozzle guide vanes, the correct angle of incidence onto the rotor blades being achieved by rotating the rotor in the reverse direction. Blade surface heat fluxes were measured using thin film gauges. In the absence of the usual blade-passing fluctuations, the root-mean-square fluctuation in heat flux was typically only 7% of the DC level.
Nusselt numbers are compared with cascade data and CFD predictions from both a three-dimensional viscous Navier-Stokes equation solver and a two-dimensional boundary layer prediction. The low inlet turbulence level produced a long laminar region on the suction surface followed by sudden transition. CFD predictions of Nusselt number on this surface were very sensitive to the choice of boundary layer state, and the experimental level was approximately mid-way between predictions with a transitional intermittency distribution and those with a turbulent distribution. On the pressure surface the levels were approximately 25% below predicted levels, and possible reasons for this are considered.
A Simulation of the Unsteady Interaction of a Centrifugal Impeller With its Vaned Diffuser: Flows Analysis
Authors: Dawes W. N. 
Published: February 2015

The aim of this paper is to help advance our understanding of the complex, three-dimensional, unsteady flow associated with the interaction of a splattered centrifugal impeller and its vaned diffuser. A time-resolved simulation is presented of the Krain stage performed using a time-accurate, 3D, unstructured mesh, solution-adaptive Navier-Stokes solver.
The predicted flowfield, compared with experiment where available, displays a complex, unsteady interaction especially in the neighbourhood of the diffuser entry zone which experiences large periodic flow unsteadiness. Downstream of the throat, although the magnitude of this unsteadiness diminishes rapidly, the flow has a highly distorted three-dimensional character.
The loss levels in the diffuser are then investigated to try and determine how time-mean loss levels compare with the levels expected from “equivalent” steady flow analysis performed by using the circumferentially averaged exit flow from the impeller as inlet to the diffuser. It is concluded that little loss could be attributed directly to unsteady effects but rather that the principle cause of the rather high loss levels observed in the diffuser is the strong spanwise distortion in swirl angle at inlet which initiates a strong hub/comer stall.
300 kW Class Ceramic Gas Turbine Development (CGT 301)
Authors: Sakakida Masaru , Sasa Tadashi , Akiyama Kazuho , Tanaka Shinya 
Published: February 2015

CGT 301 is a recuperated, single-shaft, ceramic gas turbine for cogeneration capable of continuous full load application. In order to reduce its size, thermal stress, and deformations, ceramic parts are designed axi-symmetrically. The combustor is located on a shaft axis just before the turbine, therefore it does not have a large scroll. The turbine is a two-stage axial flow-type with ceramic blades.
For the first phase of the program, the primary-type gas turbine with all-metallic parts was fabricated and tested under various conditions. The test results confirmed the rotation stability of the gas turbine. After the test of preliminary metallic gas turbine, all-ceramic parts were fabricated and various tests were carried out to confirm their reliability.
The configuration and structure of the ceramic turbine were improved based on the data obtained from the tests of the primary-type gas turbine and the fundamental tests for ceramic components. The primary-type ceramic gas turbine of TIT 1200°C was designed and fabricated for the second phase of the program. This paper outlines the concept of the ceramic component design, test results of ceramic parts in the hot section, and the engine test.
Digital Control Brings Large Turbofan Benefits to the Regional Jetliner Turbofan Market
Authors: Sisson Patterson B. , Faymon David K. 
Published: February 2015

In keeping with the general industry trend of applying Full Authority Digital Electronic Control (FADEC) technology to small gas turbine engines, Textron Lycoming and Chandler Evans Division of Coltec Industries have developed and qualified a single channel control system for use on the Textron Lycoming LF507-1F turbofan engine. The LF507-1F is the world’s smallest FADEC-equipped airline turbofan engine and is the only FADEC-equipped turbofan developed and certified for the regional jetliner market. The application for this powerplant, the four-engine AVRO International Aerospace RJ Avroliner series of aircraft, began airline service in April of 1993.
The FADEC employs modern control algorithms to achieve surge-free operation over the flight envelope while providing rapid transient performance and crisp handling qualities. The control interfaces with the aircraft via an ARINC 429 data link to control each engine automatically to the desired power setting with or without N1 synchronization. A simple hydromechanical backup control provides full dispatch capability in the event of a critical FADEC system failure. In addition, the FADEC includes advanced diagnostics for fault identification to the line replaceable unit (LRU) level without specialized test equipment.
This paper describes the architecture, primary features, and development process of the engine control system. Emphasis is placed on the design characteristics and technical challenges unique to the development of an inexpensive control system for the low thrust turbofan market.
Evaluation of an Euler Code With Non-Reflecting Boundary Conditions for the Analysis of 3D Flow in Radial Impellers
Authors: Fatsis Antonios , Van den Braembussche René 
Published: February 2015

The aim of the present paper is the investigation of the influence of non-reflecting boundary conditions on the calculation of the three-dimensional inviscid steady rotational flow inside a high pressure ratio centrifugal impeller.
The space discretization of the Euler equations was done using the finite volume technique and the multistep Runge-Kutta scheme was used for the time integration. Nonreflecting boundary conditions using Fourier decomposition are applied at inflow and outflow boundaries.
Computations show that convergence to steady state is reached faster when non-reflecting boundary conditions are used compared to the reflecting boundary conditions. Studying the influence of the non-reflecting boundary conditions on the accuracy of the flowfield prediction, it is concluded that the location of the upstream boundary has a negligible influence on the solution even in the case where a shock is crossing the boundary.
A Rapid Aerodynamic Loading Procedure for Centrifugal Impeller Design
Authors: Howard J. H. G. , Osborne Colin , Japikse David 
Published: February 2015

A crucial aspect of the design process for centrifugal impellers is the establishment of specific blade shapes. A rapid inviscid flow analysis procedure was developed for incorporation within a geometry manipulation code. Using a single streamtube model, a single-pass computation technique was generated. A two-zone model ensures that key features of the passage flow physics are incorporated.
Several examples of industrial design problems are employed to demonstrate the capabilities of the rapid loading method and its use in a geometry design procedure (used by some 20 industrial design groups worldwide). Comparisons with a quasi-three-dimensional method are included. The rapid loading method is most accurate when the meridional stream paths have similar shapes to those for the hub and shroud contours. The technique is useful within a geometry generation program since rapid aerodynamic screening of candidate configurations is allowed with sufficient accuracy to avoid the need for quasi-three-dimensional approaches. If required, the final design may be analyzed using three-dimensional viscous flow calculation methods.
Experimental Investigation of the Flow Field of Deep Rotating Stall in a Centrifugal Compressor
Authors: Chen Y. N. , Seidel U. , Chen J. , Haupt U. , Rautenberg M. 
Published: February 2015

The pressure field of deep rotating stall of a centrifugal compressor with two stall cells is analysed by means of the two-dimensional pressure pattern in the impeller determined by Chen et al. (1993). These authors transferred the pressure pattern measured on the shroud surface (i.e. in the absolute frame) to that related to the rotating blade channels. The transferred pressure pattern is thus a two-dimensional one.
The existence of the low and high pressure vortices according to the Rossby wave theory is confirmed by this experiment. The development stages of the two vortices, in combination with the Rossby wave that steers the rotating stall, can be evaluated very well.
The vortex low is developed from the front between the reverse flow (with high temperature and entropy) and the forward flow (with low temperature and entropy) due to baroclinic instability. Its center is situated within the channel of the splitter blade. This front is accompanied by a squall line of small-scaled eddies. This is the same phenomenon as can be observed on the meteorological polar front.
The vortex high is induced by the vortex low. Its embryo starts on the pressure surface. Its center is situated behind the inlet edge of the splitter blade.
It can be further verified that the stall cell is caused by the backflows of the induction fields of the two vortices (low and high).
The Streamwise Development of Görtler Vortices in a Favorable Pressure Gradient
Authors: Finnis M. V. , Brown A. 
Published: February 2015

Measurements are presented of the streamwise velocity variation within a laminar boundary layer on a concave surface of 4 m radius of curvature for which the free-stream velocity gradient factor ν/U02dU0/dx was approximately 1 × 10−6. The velocity variation was consistent with the presence of counter-rotating vortices resulting from the Görtler instability. The vortices exhibited exponential growth over the streamwise extent of the measurements to a disturbance amplitude of approximately 13% of the local freestream velocity. The vortex growth rates were found to be less than those for a zero velocity gradient factor indicating that a favorable pressure gradient stabilizes the flow with respect to the Görtler instability.
Boundary layer profiles at local upwash and downwash positions are compared with the linear theory for which the mean flow was modelled using the Pohlhausen approximation to the solution of the boundary layer equations. The agreement between the measured and predicted profiles indicates that the linear stability theory can provide a fair approximation to the small amplitude growth of the Görtler instability.
Component and Procedural Improvements for the T-62T-40-7 Gas Turbine in the LCAC Fleet
Authors: Ehrhardt Jerome P. , Spring Soren K. 
Published: February 2015

The Sundstrand Power Systems T-62T-40-7 Gas Turbine Auxiliary Power Unit (APU) was adapted from an aircraft-borne APU to a marine application on the U.S. Navy’s Landing Craft Air Cushion (LCAC).
Although the LCAC APU experienced less operating time than its aircraft version, the environmental conditions that exist cause unusual wear and component failures. Component and procedural improvements have been developed to extend the reliability of the T-62T-40-7 on board the LCAC.
Heat Transfer Measurements in an Annular Cascade of Transonic Gas Turbine Blades Using the Transient Liquid Crystal Technique
Authors: Martinez-Botas R. F. , Lock G. D. , Jones T. V. 
Published: February 2015

Heat transfer measurements have been made in the Oxford University Cold Heat Transfer Tunnel employing the transient liquid crystal technique. Complete contours of the heat transfer coefficient have been obtained on the aerofoil surfaces of a large annular cascade of high pressure nozzle guide vanes (mean blade diameter of 1.11 m and axial chord of 0.0664 m). The measurements are made at engine representative Mach and Reynolds numbers (exit Mach number 0.96 and Reynolds number 2.0 × 106).
A novel mechanism is used to isolate five preheated blades in the annulus before an unheated flow of air passes over the vanes, creating a step change in heat transfer. The surfaces of interest are coated with narrow-band thermochromic liquid crystals and the colour crystal change is recorded during the run with a miniature CCD video camera. The heat transfer coefficient is obtained by solving the one dimensional heat transfer equation for all the points of interest.
This paper will describe the experimental technique and present results of heat transfer and flow visualisation.
System Development Test Program for the WR-21 Intercooled Recuperated (ICR) Gas Turbine Engine System
Authors: Hawkins William J. , Mathieson Douglas , Bruce Chris J. , Socoloski Paul 
Published: February 2015

Westinghouse Electric Corporation has teamed with Rolls-Royce to develop an affordable, commercially based Intercooled/Recuperated Gas Turbine Engine System (ICR) for the United States Navy. This engine system known as WR-21 will become the next prime mover on Navy new construction surface combatants. The system development test program for the WR-21 engine system will be carried out at two test sites in geographically different locations. These are the US Navy’s Test Site at the Carderock Division Naval Surface Warfare Center in Philadelphia, Pa. and the Royal Navy’s Admiralty Test House at the Test and Evaluation Establishment, Pyestock in the United Kingdom.
This paper will briefly describe the WR-21 engine system with a more detailed discussion of the system development test program itself. This will include descriptions of the system development testing to be performed and the test facilities and data acquisition systems at each test site location. Also discussed are the methods used to establish the required design commonality between each test site to establish test bed cross-calibration and provide test program flexibility and interchangeability of testing at each site.
Turbine Cascade Losses in the Presence of Progressive Particle Deposition
Authors: Kladas D. D. , Georgiou D. P. 
Published: February 2015

Gas turbine efficiency deterioration due to particulate deposition is a well known phenomenon especially in industrial machines which operate on “heavy” fuels. The result is the need for frequent blade surface cleaning after certain hours of operation. The present study investigates the flowfield modification and the corresponding aerodynamic losses due to the progressive particle deposition on the aerofoil surfaces of a representative turbine stator cascade.
The flowfield is calculated with the Katsanis (TSONIC) code coupled to the STAN5 one for the boundary layer. The losses are calculated by momentum scanning - Stewart’s method. The problem is quasi-steady since the time scale in turbine blade flows is of the order of milliseconds, while in the deposition the data indicate slow growth rates with time scales of the ordr of hours. The particle deposition calculation is based on a turbulent diffusion model proposed by Yau and Young (18) and the particle characteristics refer to ash. The main parameters calculated are the new thickness distribution, the throat blockage, the aerodynamic losses and the transition point. The paper comments on the cleaning process based on these results.
Integration of Magnetic Bearings in the Design of Advanced Gas Turbine Engines
Authors: Storace A. F. , Sood D. , Lyons J. P. , Preston M. A. 
Published: February 2015

Active magnetic bearings provide revolutionary advantages for gas turbine engine rotor support. These advantages include tremendously improved vibration and stability characteristics, reduced power loss, improved reliability, fault-tolerance, and greatly extended bearing service life. The marriage of these advantages with innovative structural network design and advanced materials utilization will permit major increases in thrust to weight performance and structural efficiency for future gas turbine engines.
However, obtaining the maximum payoff requires two key ingredients. The first key ingredient is the use of modern magnetic bearing technologies such as innovative digital control techniques, high-density power electronics, high-density magnetic actuators, fault-tolerant system architecture, and electronic (sensorless) position estimation. This paper describes these technologies and the test hardware currently in place for verifying the performance of advanced magnetic actuators, power electronics and digital controls. The second key ingredient is to go beyond the simple replacement of rolling element bearings with magnetic bearings by incorporating magnetic bearings as an integral part of the overall engine design. This is analogous to the proper approach to designing with composites, whereby the designer tailors the geometry and load carrying function of the structural system or component for the composite instead of simply substituting composites in a design originally intended for metal material. This paper describes methodologies for the design integration of magnetic bearings in gas turbine engines.
Studies of Lean Blowout in a Step Swirl Combustor
Authors: Durbin Mark D. , Ballal Dilip R. 
Published: February 2015

The design requirements of a modem gas turbine combustor are increasingly dictated by wide stability limits, short flame length, and uniform mixing. To achieve the best trade-off between the above three factors, flame characteristics (length, shape, mixedness), lean blowout (LBO), and optimum combustor configuration should be investigated over a wide range of inner and outer air velocities, inner and outer vane angles, and co- vs. counter-swirl arrangements. Such an investigation was performed in a step swirl combustor (SSC) designed to simulate the fuel-air mixing pattern in a gas turbine combustor dome fitted with an airblast atomizer.
It was found that an increase in the outer vane angle and a decrease in inner air velocity decreased the flame length. LBO was improved when outer flow swirl intensity was increased. An optimum hardware and velocity configuration for the SSC was found for inner swirl = 45°, outer swirl = 60°, co-swirl direction, and inner air velocity = outer air velocity = 16 m/s. This optimum SSC configuration yielded: (i) low values of LBO, (ii) short flame length, (iii) uniformly mixed stable flame, and (iv) little or no variation in these characteristics over the range of operation of SSC. Finally, the co- vs. counter-swirl arrangements and the operation of the optimized combustor configuration were discussed.
Effects of Rotating Inlet Distortion on Multisage Compressor Stability
Authors: Longley J. P. , Shin H.-W. , Plumley R. E. , Silkowski P. D. , Day I. J. , Greitzer E. M. , Tan C. S. , Wisler D. C. 
Published: February 2015

In multi-spool engines, rotating stall in an upstream compressor will impose a rotating distortion on the downstream compressor, thereby affecting its stability margin. In this paper experiments are described in which this effect was simulated by a rotating screen upstream of several multistage low-speed compressors. The measurements are complemented by, and compared with, a theoretical model of multistage compressor response to speed and direction of rotation of an inlet distortion.
For co-rotating distortions (i.e., distortions rotating in the same direction as rotor rotation), experiments show that the compressors exhibited significant loss in stability margin and that they could be divided into two groups according to their response. The first group exhibited a single peak in stall margin degradation when the distortion speed corresponded to roughly 50% of rotor speed. The second group showed two peaks in stall margin degradation corresponding to distortion speeds of approximately 25–35% and 70–75% of rotor speed. These new results demonstrate that multistage compressors can have more than a single resonant response. Detailed measurements suggest that the two types of behavior are linked to differences between the stall inception processes observed for the two groups of compressors and that a direct connection thus exists between the observed forced response and the unsteady flow phenomena at stall onset.
For counter-rotational distortions, all the compressors tested showed minimal loss of stability margin. The results imply that counter-rotation of the fan and core compressor, or LP and HP compressors, could be a worthwhile design choice.
Calculations based on the two-dimensional theoretical model show excellent agreement for the compressors which had a single peak for stall margin degradation. We take this first-of-a-kind comparison as showing that the model, though simplified, captures the essential fluid dynamic features of the phenomena. Agreement is not good for compressors which had two peaks in the curve of stall margin shift versus distortion rotation speed. The discrepancy is attributed to the three-dimensional and short length scale nature of the stall inception process in these machines; this includes phenomena that have not yet been addressed in any model.
The Aerodynamic Mixing Effect of Discrete Cooling Jets With Mainstream Flow on a Highly Loaded Turbine Blade
Authors: Wilfert Günter , Fottner Leonhard 
Published: February 2015

For the application of film cooling to turbine blades experimental investigations were performed on the mixing processes in the near hole region with a row of holes on the suction side of a turbine cascade. Data were obtained using pneumatic probes, pressure tappings, and a three-dimensional subminiature hot wire probe, as well as surface flow visualization techniques.
It was found that at low blowing rates a cooling jet behaves very much like a normal obstacle and the mixing mainly takes place in the boundary layer. With increasing blowing rates the jet penetrates deeper into the mainstream. The variation of the turbulence level at the inlet of the turbine cascade and the Reynolds number showed a strong influence on the mixing behavior. The kidney-shaped vortex and as an important achievement the individual horse-shoe vortex of each single jet were detected and their exact positions were obtained. This way it was found that the position of the horse-shoe vortex is strongly dependent on the blowing rate and this influences the aerodynamic mixing mechanisms. A two-dimensional code for the calculation of boundary layer flows called GRAFTUS was used, however the comparison with the measurements showed only limited agreement for cascade flow with blowing due to the strong three-dimensional flow pattern.