bluuebunny/crossref_metadata_2025 huggingface dataset:
Double-Jet Film-Cooling for Highly Efficient Film-Cooling With Low Blowing Ratios
Authors: Kusterer Karsten , Elyas Anas , Bohn Dieter , Sugimoto Takao , Tanaka Ryozo
Published: August 2009

Further improvement of the thermal efficiency of modern gas turbines can be achieved by a further reduction of the cooling air amount. Therefore, it is necessary to increase the cooling effectiveness so that the available cooling air fulfils the cooling task even if the amount has been reduced. In particular, the cooling effort for the vanes and blades of the first stage in a modern gas turbine is very high. The task of the film-cooling is to protect the blade material from the hot gas attack to the surface. Unfortunately, aerodynamic mixing processes are enhanced by secondary vortices in the cooling jets and, thus, the film-cooling effectiveness is reduced shortly behind the cooling air ejection through the holes. By improvement of the hole positioning the negative interaction effects can be reduced. The Double-jet Film-cooling (DJFC) Technology invented by the authors is one method to reach a significant increase in film-cooling effectiveness by establishing an anti-kidney vortex pair in a combined jet from the two jets starting from cylindrical ejection holes. This has been shown by numerical investigations and application to an industrial gas turbine as reported in recent publications. Whereas the original design application has been for moderate and high blowing ratios, the present numerical investigation shows that the DJFC is also applicable for lower blowing ratios (0.5<M<1.0) with only slight modification of the geometry of the configuration. The anti-kidney vortex concept can also be established for the lower blowing ratios and, as a result, a very high film-cooling effectiveness is reached not only behind the ejection holes but also for a very long distance downstream (> 30 hole diameters).
Modeling of Variable Inlet Guide Vanes Affects on a One Shaft Industrial Gas Turbine Used in a Combined Cycle Application
Authors: Celis Cesar , Ribeiro Pinto Paula de M. , Barbosa Rafael S. , Ferreira Sandro B.
Published: August 2009

It is well known that gas turbine simulation involves satisfying the conditions of compatibility between its components. At design point, the components are all well matched and working at high efficiency regions. However, at steady state off-design, due to the compatibility issues and changes in operating parameters, basically turbine entry temperature and pressure ratio to attain a certain load, it is possible that the components may be working within regions of low efficiency. A reason for this phenomenon is that the flow areas at the various sections of the engine correspond to that at design point, such that operation at off-design is restricted. One way to widen the operational envelope of an engine is varying these flow areas, providing a good match between the gas turbine components. A widely used type of variable geometry which has attracted a great amount of interest is the use of compressor variable geometry, the so called variable inlet guide vanes (VIGVs), as a power control strategy, which involves the control of the air flow rate entering the compressor and the power output modulation at constant rotational speed. The purpose of the air flow rate modulation is to enhance the heat recovery performance and thus increase the combined cycle efficiency by maintaining high turbine exhaust temperature. One methodology used to model a variable geometry compressor, in the absence of its geometric data involves the use of correction factors, as functions of the VIGV change. Fundamentally, this methodology assumes that each new position of the VIGVs represents a new machine, i.e., a new design point, such that its original map of characteristics is displaced in order to describe this “new” compressor. The purpose of this work is to analyze the influence of the use of different functions for these correction factors on a W501F (one shaft, industrial) gas turbine simulation. An in-house computer program developed for performance modeling of gas turbines was utilized to carry out the simulations. The results provided by this computer code show good agreement with operational data, indicating that, although more tests must be conducted, the methodology seems to be reliable enough for the aims of the project for which it has been developed.
Research on the Tip Sections of a Long Turbine Blade
Authors: Sˇimurda David , Luxa Martin , Syna´cˇ Jaroslav , Sˇafarˇik Pavel
Published: August 2009

This paper reports on the results of high-speed aerodynamic research on a blade cascade. The blade cascade represents the tip section of the last stage rotor blades of a large output steam turbine. The aerodynamic characteristics of the blade cascade are presented together with flow field photographs in a range of transonic and supersonic velocities. Pictures of the flow field were obtained by means of interferometry and the schlieren method. The pressure distribution over the profiles was also evaluated using interferometry. Finally, the data for evaluating the losses and other integral characteristics of the flow exiting from the cascade were measured pneumatically with the use of a traversing device. The results are analyzed and discussed.
Reducing CO2 Emissions From a Coal Fired Power Plant by Using a Molten Carbonate Fuel Cell
Authors: Milewski Jarosław , Lewandowski Janusz , Miller Andrzej
Published: August 2009

A Molten Carbonate Fuel Cell (MCFC) is shown to reduce CO2 emissions from a Coal Fired Power Plant (CFPP). The MCFC is placed in the flue gas stream of the coal fired boiler. The main advantages of this solution are: higher total electric power generated by a hybrid system, reduced CO2 emissions and higher system efficiency. The model of the MCFC is given and described. The results obtained show that use of an MCFC could reduce CO2 emissions by 56%, which gives a relative CO2 emission rate of 288 kgCO2 per MWh.
Influence of Cooling Flow Rate Variation on Gas Turbine Blade Temperature Distributions
Authors: Sierra Espinosa Fernando Z. , Han Je Chin , Portugal Areli Uribe , Kubiak J. , Narzary Diganta , Blake Sarah , Cadena Fernando , Lara Hugo , Nebradt Jesu´s
Published: August 2009

Temperature and flow rate of combustion gases and cooling stream are essential conditions for blade integrity in gas turbines. Since the combustion products pass directly to the first stage of blades high thermal stresses can develop, so the temperature field in the blade material must be controlled to avoid damage and/or reduction of blade useful life. This paper discusses an investigation on the influence of cooling airflow reduction on blade life. The flow rate reduction under consideration may be due to malfunctions of the compressor such like deposits or partial blockage in the blade ducts. It has been reported that air discharge from the compressor can be reduced up to 15% of the nominal rate due to deposits related with impurities contained in the environment. In this work an evaluation of the effect of reducing the cooling airflow rate on the temperature distribution on the blades surface is attempted. The flow stream that surrounds the blade together with the cooling airflow in the blade interior channels were characterized in the laboratory. Fields of temperature on the blade surface were obtained using the temperature sensitive paint technique, TSP. Thermocouple measurements were used for punctual temperatures as a reference. The results showed the regions of possible thermal stresses concentration as a function of cooling airflow rate variations. Additionally, the problem was resolved computationally in conjugate mode, considering both fluid streams external and internal plus heat conduction at the interior of the blade material. The computer model is used to simulate other conditions not addressed in the experiment. The paper discusses the comparison of numerical to experimental results and discusses the methodology for further work.
Reynolds Stress Calculations for Pre-Transition Boundary Layers With Turbulent Free Streams
Authors: Mayle R. E. , Schulz A. , Bauer H.-J.
Published: August 2009

Equations are developed for calculating u′u′ and u′v′ in a laminar boundary layer when the free stream is turbulent. These equations are based on the ideas presented by the first two authors in their 1997 paper, where the crucial role of high frequency fluctuations in the free stream was first recognized. Solutions of these equations, which also include the effect of a variable free stream velocity, show that the intensity of the fluctuations in the boundary layer increase as separation is approached, that the initial growth rate of the shear stress component is substantially less than the normal component, that the maximum value of u′u′ near transition is about the same as that in a turbulent boundary layer, and that, as is well known, free-stream turbulence increases the skin-friction coefficient. Overall, the calculations agree reasonably well with experimental results for all free-stream turbulence levels and velocity variations. In addition, upon examining Liepmann’s transition hypothesis, we show that transition from laminar to turbulent flow seems to occur at the streamwise position where the maximum of u′v′/uτ2 first attains a value of about one-third. But other criteria also seem to apply.
A Generalised Likelihood Ratio Test for Adaptive Gas Turbine Health Monitoring
Authors: Borguet S. , Le´onard O.
Published: August 2009

Kalman filters are widely used in the turbine engine community for health monitoring purpose. This algorithm has proven its capability to track gradual deterioration with a good accuracy. On the other hand, its response to rapid deterioration is either a long delay in recognising the fault, and/or a spread of the estimated fault on several components. The main reason of this deficiency lies in the transition model of the parameters that is blended in the Kalman filter and assumes a smooth evolution of the engine condition. This contribution reports the development of an adaptive diagnosis tool that combines a Kalman filter and a secondary system that monitors the residuals. This auxiliary component implements a Generalised Likelihood Ratio Test in order to detect and estimate an abrupt fault. The enhancement in terms of accuracy and reactivity brought by this adaptive Kalman filter is highlighted for a variety of simulated fault cases that may be encountered on a commercial aircraft engine.
On the Role of the Deterministic and Circumferential Stresses in Throughflow Calculations
Authors: Simon J.-F. , Le´onard O.
Published: August 2009

This paper presents a throughflow analysis tool developed in the context of the average-passage flow model elaborated by Adamczyk. The Adamczyk’s flow model describes the 3-D time-averaged flow field within a blade row passage. The set of equations that governs this flow field is obtained by performing a Reynolds averaging, a time averaging and a passage-to-passage averaging on the Navier-Stokes equations. The throughflow level of approximation is obtained by performing an additional circumferential averaging on the 3-D average-passage flow. The resulting set of equations is similar to the 2-D axisymmetric Navier-Stokes equations but additional terms resulting from the averages show up: blade forces, blade blockage factor, Reynolds stresses, deterministic stresses, passage-to-passage stresses and circumferential stresses. This set of equations represents the ultimate throughflow model provided that all stresses and blade forces can be modeled. The relative importance of these additional terms is studied in the present contribution. The stresses and the blade forces are determined from 3-D steady and unsteady databases (a low speed compressor stage and a transonic turbine stage) and incorporated in a throughflow model based on the axisymmetric Navier-Stokes equations. A good agreement between the throughflow solution and the averaged 3-D results is obtained. These results are also compared to those obtained with a more “classical” throughflow approach based on a Navier-Stokes formulation for the endwall losses, correlations for profile losses and a simple radial mixing model assuming turbulent diffusion.
Future Aero-Engines’ Optimisation for Minimal Fuel Burn
Authors: Colmenares Fernando , Kyprianidis Konstantinos , Go´mez Josue´ , Ogaji Stephen , Pilidis Pericles , Latorre Sergio
Published: August 2009

While aircraft environmental performance has been important since the beginnings of commercial aviation, continuously increasing passenger traffic and a rise in public awareness have made aircraft noise and emissions two of the most pressing issues hampering commercial aviation growth today. The air transportation for the new millennium will require revolutionary solutions to meeting public demand for improving safety, reliability, environmental compatibility, and affordability. The objective of this research is to assess the trade-off between operating costs and environmental requirements of the future aero engines for short range commercial aircrafts. This involves optimising the engines’ design point to minimise the block fuel and evaluating the economic and environmental impact. A high by-pass ratio turbofan engine with performance characteristics and technology from the year 2000 was set up as a baseline and compared to very high by-pass ratio turbofans. The results present a great potential benefit of the geared turbofan compared to high BPR one (baseline) to reduce cruise CO2 emissions and noise; however this may involve NOx penalties, that is an increase of 5.1% in comparison to the baseline. The CRTF engine seems to be, at least according to the simulations, a very promising solution in terms of environmental and economical performance. This is one on the series of work that would be carried out on the cycles being assessed in this paper (feasibility study). Further work on the specific technical issues — such as: technological implications — would be published when completed.
Low Solidity Cascade Diffuser and Scroll Optimization for Centrifugal Compressors
Authors: Iancu Florin , Trevino John , Sommer Steven
Published: August 2009

One of the major concerns of the centrifugal compressor designer is the trade-off between the compressor’s efficiency and its range. The compressors equipped with vaneless diffusers benefit from a broader operational range than conventional vaned diffusers, but with decreased compression efficiency. One solution found in recent years is the Low Solidity Diffuser, which displays a range greater than conventional vaned diffusers, by preventing rotating stall, and efficiency higher than vaneless diffusers. The study focuses on the impact of geometry elements on the flow field through the vanes’ passages, and on the change of performance with the location of the vanes relative to the scroll tongue. This paper presents results of experimental and Computational Fluid Dynamics (CFD) investigations — for centrifugal compressors with low solidity cascade diffusers. The vanes are based on a NACA airfoil profile and have been designed using commercial software from ConceptsNREC: Compal and AxCent. The analysis part was performed using CFX software from Ansys, Inc. The performance parameters recorded are pressure ratios, isentropic efficiency, flow angles and velocity field profiles.
Aerodynamic Noise Assessment Using CFD/CAA Technique
Authors: Dykas Slawomir , Wroblewski Wlodzimierz , Chmielniak Tadeusz
Published: August 2009

In the presented paper a numerical method for aerodynamic noise assessment on the basis of unsteady mean flow field data has been presented. This method belongs to the very wide family of so called hybrid CFD/CAA methods. In described method for calculations of the acoustic fluctuations a solution of full non-linear Euler equations is used. The Euler equations are solved with the use of numerical scheme with third order accuracy in space and time. The presented method is numerically decoupled with CFD solution, so that the information about mean unsteady flow field can be obtained using arbitrary CFD method (solver). The accuracy of the acoustic field assessment depends on the quality of the CFD solutions.
Measured and Predicted Performance of a High Pressure Ratio Supersonic Compressor Rotor
Authors: Grosvenor Allan D. , Taylor David A. , Bucher Jonathan R. , Aarnio Michael J. , Brown Paul M. , Draper Robert D. , Lawlor Shawn P.
Published: August 2009

The testing of an 8:1 pressure ratio supersonic single axial compressor rotor referred to as Rampressor-2 is described. Design of this shockwave compression system is based on principles employed for supersonic intakes consisting of a multi-shock compression system and boundary layer treatment. The rotor consists of three blade passages within which the shock system is produced by a ramp, throat and diffuser contoured on the hub. The technology has been previously demonstrated in a 2.3:1 pressure ratio experimental test compressor (Rampressor-1). Measured performance is compared with numerical predictions. Further developments to improve Rampressor performance are discussed, and the appropriateness of this technology for Carbon Capture & Sequestration and LNG applications is highlighted.
PGT25+G4 Gas Turbine Development, Validation and Operating Experience
Authors: De Meo Roger , D’Ercole Michele , Russo Alessandro , Gamberi Francesco , Gravame Francesco , Mucz David
Published: August 2009

The PGT25+G4 gas turbine, latest in GE Infrastructure Oil&Gas PGT25 two-shaft aeroderivative family, is a 34 MW-class gas turbine for mechanical drive and power generation applications and maintains the same efficiency and availability of the previous PGT25+. The PGT25+G4 was validated through an extensive test program, which included some key test-rigs such as the full-scale LM2500+G4 Gas Generator test and other component tests, in advance of the First Engine to Test (FETT). The FETT included an equivalent-to-production configuration package (gas turbine, auxiliaries and control system), ran in a dedicated area in GE Oil&Gas Test Facilities to validate the machine for both mechanical drive and power generation applications. All critical-to-quality parameters of the HSPT (High Speed Power Turbine) were investigated, such as turbine gas path components temperatures and stresses, PT performances and PT operability when coupled with the LM2500+G4 Gas Generator. First production unit is currently in operation at Alliance Pipeline Canada Windfall 1 Compression Station. This paper describes the gas turbine main features, how the test program was built and discusses FETT results. Moreover, gas turbine field operation experience and lessons learned are presented.
The Influence of Lean and Sweep in a Low Pressure Steam Turbine: Analysis of Three Stages With a 3D CFD Model
Authors: Sigg Roland , Casey Michael V. , Mayer Ju¨rgen F. , Su¨rken Norbert
Published: August 2009

The present paper accompanies the analysis of Vo¨lker et al. (2008), where experimental measurements on a three stage LP turbine for different configurations with lean and sweep in the last nozzle guide vane were presented followed by an analysis of these with a through-flow method. The present paper continues the analysis of the three stage machine with the support of steady 3D RANS calculations with stage interface (mixing) planes between the stages and making use of the SST turbulence model. The flow simulations are validated by comparison with flow traverses and with blade surface static pressure measurements. Good agreement is found between the simulations and the measurements, and the causes of discrepancies are analysed and discussed. In particular, the paper discusses the CFD results on two measured configurations focusing on the design mass flow and with regard to: • the effect of neglecting a backward facing step in the shroud contour above the rotor of one of the stages on the calculated flow-field; • the detailed changes in the spatial flow-field due to the introduction of sweep in the last nozzle row; • the prediction of the radial distributions of loss coefficients for the last two turbine blade rows; • the use of different models for the equation of state: an equilibrium or a non-equilibrium steam model.
Refitting of Exhaust Diffuser of Industrial Gas Turbine
Authors: Vassiliev Vladimir , Rothbrust Matthias , Irmisch Stefan
Published: August 2009

This paper describes the aerodynamic optimization of the GT26 exhaust diffuser. The need for optimization was triggered by an upgrade of the compressor, resulting in a higher mass flow and a higher power output. The expansion turbine remained unchanged. However, the increase in mass flow had a significant impact on the Mach number. Secondly, the residual swirl at the turbine outlet, and therefore, the exhaust loss in original diffuser would have increased. The re-optimization of diffuser allowed minimization of the losses and improvement of the overall engine performance.
Prediction of Evaporative Effects Within the Blading of an Industrial Axial Compressor
Authors: Matz Charles , Kappis Wolfgang , Cataldi Giovanni , Mundinger Gerd , Bischoff Stefan , Helland Eivind , Ripken Matthias
Published: August 2009

The results of a compressor flow-analysis code calibration study for estimating the effects of water evaporation within the blade rows of industrial axial compressors are presented. In the study, a mean-line code was chosen for the calibration tool due to its accepted use during preliminary design studies, at which time during the compressor design process one would logically consider power augmentation through wet compression. The calibrated code features a non-equilibrium thermodynamic single-droplet evaporation model augmented with an empirical splashing model, which, as input, uses measurements of droplet spectra data taken on water injection nozzles in an intake rig configured with realistic length scales. In addition, a wetted-airfoil-surface flow angle deviation model is applied to predict changes in compressor stage characteristics, which, in turn, affect the inlet mass flow of the compressor. The test vehicle for calibration was a 50-Hz Alstom industrial gas turbine. Once calibrated, the code was successfully utilized to predict wet compression effects for three additional like-family Alstom gas turbines operating at constant speed while under full load. The effects modeled by the code include bleed supply pressure suck-down and bleed temperature cool-down effects, as well as compressor inlet mass flow and power consumption effects.
Numerical Investigation of Low Solidity Vaned Diffuser Performance in a High-Pressure Centrifugal Compressor: Part II—Influence of Vane Stagger
Authors: Oh JongSik , Buckley Charles W. , Agrawal Giri L.
Published: August 2009

As the second part, following the authors’ previous study, the influence of the LSD (Low Solidity Diffuser) vane stagger on high-pressure ratio centrifugal compressor performance is numerically investigated with all the other design parameters fixed, while vane solidities are in the range from 0.70 to 0.85. Vane stagger is varied for 6 cases from 8.55 deg to 22.37 deg with the NACA65-(4A10)06 airfoil profile, and the Stage interface scheme is applied for an interaction treatment. As the vane stagger increases, changing from 10.55 deg to 19.58 deg, the compressor overall performance is generally improved, but two extreme cases of vane stagger of 8.55 deg and 22.37 deg provide poor performance. Vane stagger of 19.58 deg shows the highest efficiency and pressure rise near design flow, while vane stagger of 13.76 deg has the largest operating range with acceptable performance of efficiency and pressure ratio.
Radiation Bench-Marking in a Model Combustor
Authors: Jiang Lei-Yong , Campbell Ian
Published: August 2009

Radiation heat transfer in a model combustor with interior and exterior conjugate heat transfers has been numerically studied. The previous investigations on turbulence, combustion and scalar transfer modeling (Reynolds analogy), and comparisons with a comprehensive experimental database provide a reliable base to evaluate the effect of radiation heat transfer on the flow field and NO emission in the combustor. Some of the numerical results with and without radiation are presented and compared with the experimental measurements. It is found that the total radiation heat flux through the combustor wall is about 4.2% of the total energy released from the input fuel. The effect of radiation on the flow field is minor, particularly to the velocity field. In contrast, it has significant effects on the NO field, where the predicted values without radiation are two times higher than those with radiation or the experimental data. A considerable effect of radiation on the combustor wall temperature is also observed. In summary, to provide valuable predictions of NO emission and combustor liner temperature, the radiation heat transfer should be properly taken into account in numerical simulations.
United Modeling of Working Process in Aircraft Gas Turbine Engines
Authors: Ivanov M. Ja. , Mamaev B. I. , Nigmatullin R. Z.
Published: August 2009

This paper is about the modern mathematical models of working process in the whole flow passage of aviation gas turbine engines. These models are referred as high-level models, based on real 3D geometry of engine flow passage. They allow to simulate steady and unsteady processes in 1D, 2D and 3D formulations, calculate engine performances, determine propagation of radial and circular parameter no uniformities in engine flow passage and predict influence of main parameters on engine efficiency. Typical examples of working process simulation in whole engine and their components are presented below.
A New Streamline Curvature Throughflow Method for Radial Turbomachinery
Authors: Casey Michael , Robinson Chris
Published: August 2009

This paper describes a newly developed streamline curvature throughflow method for the analysis of radial or mixed flow machines. The code includes curved walls, curved leading and trailing edges, and internal blade row calculating stations. A general method of specifying the empirical data provides separate treatment of blockage, losses, and deviation. Incompressible and compressible fluids are allowed, including real gases and supersonic relative flow in blade rows. The paper describes some new aspects of the code. In particular, a relatively simple numerical model for spanwise mixing is derived, the calculation method for prescribed pressure ratio in compressor bladed rows is described, and the method used to redistribute the flow across the span due to choking is given. Examples are given of the use and validation of the code for many types of radial turbomachinery and these show it is an excellent tool for preliminary design.
The Influence of Lean and Sweep in a Low Pressure Steam Turbine: Throughflow Modelling and Experimental Measurements
Authors: Vo¨lker Lutz , Casey Michael , Dunham John , Stu¨er Heinrich
Published: August 2009

This paper describes experimental and throughflow investigations on two configurations of a model three-stage low pressure steam turbine. A companion paper describes 3D CFD simulations of the same turbine test data. Global performance measurements and detailed flow field measurements with pneumatic flow probes were carried out to quantify the changes in the design due to the introduction of sweep in the last stator nozzle vanes. An existing 2D throughflow code was improved to enable the present calculations to be completed. The test results have been used in this paper to calibrate the 2D throughflow model, by adjustment of empirical correlation data to match the experimental data on one of the configurations. This throughflow model was then used to examine the influence of lean and sweep on the design. The results identify that throughflow calculations can model the global effects of lean and sweep in the last stages of steam turbines. Some insight is gained on the losses across the span for the different configurations and on the benefits of lean and sweep in reducing the hub reaction in such stages.
Breakdown of the Tip Clearance Vortex in a Rotor Equivalent Cascade and in a Single-Stage Low-Speed Compressor
Authors: Schrapp Henner , Stark Udo , Saathoff Horst
Published: August 2009

The paper describes experimental investigations of the tip clearance vortex in a linear compressor cascade with a tip clearance to one side and a compressor with a rotor tip section represented by the cascade. The aim is to show experimentally that breakdown of the tip clearance vortex can take place in subsonic compressors. As a first step, the flow in the linear compressor cascade has been investigated at different flow angles from the design point up to the stability limit of the cascade. In a second step the flow in the tip region of the rotor in a low-speed single-stage compressor has been investigated from the design point up to the stall limit of the compressor. The analysis of PIV measurements reveal a low momentum fluid area in the passage both in the cascade and the compressor when approaching the stall limit. This area is separated from the main flow by an interface that is characterized by high standard deviations of the velocities. The location of this interface correlates well with the positions at which the analysis of the unsteady pressure signals obtained with several flush mounted high-response pressure transducers reveals a bump of increased amplitude at a certain nondimensional frequency. The results in the compressor are for the most part identical to those obtained in the cascade. When approaching stall a distinct bump shows up in the casing wall pressure spectra at a nondimensional frequency comparable to the nondimensional frequency found in the cascade. At the same time an area of very low momentum fluid accumulates in the vicinity of the original vortex axis, as can be shown by PIV measurements in the rotor of the compressor. Additionally it can be shown, that the blockage that is due to the broken down tip clearance vortex leads to a rotating phenomenon, comparable to the phenomenon of ‘rotating instabilities’.
Numerical Study of the Thermal Development in a Rotating Cooling Passage
Authors: Martinez Lucci Jose , Amano R. S. , Guntur Krishna , Song B.
Published: August 2009

In the present study, the fully developed turbulent secondary flows inside rotating square-upstream section and rectangular downstream section oriented at several angles with axis rotation are numerically simulated using the low Reynolds Non-linear k–ω, k–ε Reynolds-Stress Model (RSM), and Large Eddy Simulation (LES) models. The two flat walls are heated (leading and trailing sides), while the outer wall and the splitter plate are thermally insulated. The simulation has been done at Re = 36000 with different rotation numbers from Ro = −0.4 to Ro = 0.4, where negative Ro values refer to counter-clock wise and positive Ro values refer to clock wise sense of rotation. This enables to investigate the effects on the thermal development of the rotation numbers. The effects of minor modifications, in cross section area at the bend exit, on thermal development, under rotating conditions, can also be explored. The interactions between secondary flows and separation lead to very complex flow patterns. To accurately simulate these flows and heat transfer, both refined turbulence models and higher-order numerical schemes are indispensable for turbine designers to improve the cooling performance. Previous studies have shown that the flow and heat transfer features through curved bends, even with a moderate curvature, cannot be accurately simulated. It is the conventional belief and practice that the usage of a proper turbulence model and a reliable numerical method for achieving accurate computations. It is shown that the present RSM model produces satisfactory predictions of the flow development inside the sharp U-bend comparing with the experiments [1]. In the present study, three turbulence models are used to predict Nusselt number distribution. The results were further compared with the LES computations and discussed the difference between the turbulence models and the LES as well.
Aerodynamic Mistuning of Structurally Coupled Blades
Authors: Stu¨er Heinrich , Schmitt Stefan , Ashcroft Graham
Published: August 2009

This paper discusses the impact of aerodynamic mistuning in the context of a structurally coupled blade system. The aerodynamic mistuning assumes a difference in blade loading (between adjacent blades) but no change in eigenfrequency or modeshape. Thus, emphasis is placed on the profile shape impact of transonic rotor blades on aeroelasic damping. An aeroelastic calculation example of the proposed mistuning strategy is present for a coupled blade arrangement. The impact of mistuning compared to the tuned datum profile is discussed. It is shown that there exist significant differences for the two configurations but that the net impact of mistuning is smaller for the coupled system investigated compared to a freestanding blade design.
Integrating MT30 Into the UK’s Future Aircraft Carrier: Power at the Heart of the Electric Propulsion System
Authors: Casson Roy
Published: August 2009

At 65,000 tonnes Britain’s new carrier (CVF) represents a step change in British force projection capability, able to embark more than 30 aircraft of various types to undertake a range of roles. For such a ship there must be a power generation system that is able to provide large amounts of power for propulsion and ship’s services, rapidly, reliably and economically. An Integrated Full Electric Propulsion (IFEP) system has been selected as the most suitable for this task. At the heart of this system are two Rolls-Royce MT30 Gas Turbine Alternators (GTAs), each capable of producing 35MW of electrical power to propel the ship at the high speeds necessary to launch advanced fighter aircraft. A variety of challenges are currently being faced to integrate these large, modern GTAs into a power and propulsion system that also features diesel generators, power electronic converters and Advanced Induction Motors (AIMs). Solutions to these issues are currently being developed by the Alliance responsible for procuring the ship, working closely with the Ministry of Defence (MoD) and the engine manufacturer to provide a robust and responsive power source that is fully integrated with the power system. This paper begins with an overview of CVF’s integrated full electric power and propulsion system and the performance requirements placed on the prime movers. It will include a discussion of the reasons for and the advantages of the power generation system design. There follows a brief description of the MT30 gas turbine and the characteristics that make it suitable for use in CVF. The main body of the text focuses on the issues relating to, and challenges faced in integrating MT30 into such a large and complex power system and a ship of this nature. Also examined are the solutions being developed to ensure the GTA is at the heart of the IFEP system, and the paper will conclude by looking at the validation of the power generation system.
Adjoint Aerodynamic Design Optimization for Blades in Multi-Stage Turbomachines: Part I—Methodology and Verification
Authors: Wang D. X. , He L.
Published: August 2009

The adjoint method for blade design optimization will be described in this two-part paper. The main objective is to develop the capability of carrying out aerodynamic blading shape design optimization in a multi-stage turbomachinery environment. To this end, an adjoint mixing-plane treatment has been proposed. In the first part, the numerical elements pertinent to the present approach will be described. The gradients of a single objective function of a weighted sum of objectives and constraints with respect to detailed blade shape perturbations are obtained very efficiently by the continuous adjoint method. The steepest descent method is used to drive the design to an optimum. The adjoint mixing-plane treatment enables the adjoint equations to be solved in a multi-stage environment. The adjoint solver is verified by comparing gradient results with a direct finite difference method and through a 2D inverse design. The adjoint mixing-plane treatment is verified by comparing gradient results against those by the finite difference method for a 2D compressor stage. The redesign of the 2D compressor stage further demonstrates the validity of the adjoint mixing-plane treatment and the benefit of using it in a multi-bladerow environment.
Adjoint Aerodynamic Design Optimization for Blades in Multi-Stage Turbomachines: Part II—Validation and Application
Authors: Wang D. X. , He L. , Li Y. S. , Wells R. G. , Chen T.
Published: August 2009

This is the second part of a two-part paper. It presents four case studies. The first case is the redesign of a transonic rotor (NASA rotor 67) at a post peak efficiency operating point. The second case is a redesign of a transonic compressor stage originally designed by DLR. The redesign is carried out at the stage peak efficiency point. The third and fourth cases look at the redesign of blade rows within a three-stage transonic test compressor that was originally designed by Siemens Industrial Turbomachinery Ltd known as the ATC compressor. Specifically the third case is a redesign of the IGV-rotor-stator configuration. It is carried out at two operating points: one is at the stage peak efficiency point; the other is at a lower stagnation pressure ratio choked flow point. Initially the redesign at the stage peak efficiency point produces considerable efficiency gain, but leads to noticeably reduced choked mass flow rate. The redesign at a near choked mass flow rate point, on the other hand, leads to considerable performance deterioration at operating points with lower mass flow rate, though the choked mass flow rate is even increased. Subsequently, a parallel multi-point approach has been implemented. Results show that a two-point design optimization avoids unacceptable performance deterioration at off design conditions. In the fourth case a redesign is applied across all 7 blade rows of the ATC compressor at the compressor design point. All these case studies are aimed to increase isentropic efficiency whilst meeting the specified constraints.
Numerical Investigation of the Flow Field in a Turbine Cascade With Different Film Cooling Holes at the Leading Edge
Authors: Wang Xiaodong , Kang Shun
Published: August 2009

Flow structures in a turbine cascade with film cooling injections at the blade leading edge are investigated numerically for three cooling models of slots, straight cylindrical and spanwise inclined cylindrical holes under various blowing ratios, 0, 0.5, 0.7, 1.1, and 1.5. The numerical simulations are well compared with experiment results of static pressure and oil film visualization of blade surfaces. Discussions are mainly focused on the 3D flow structures within a blade passage. The results show that the 3D flow and energy loss production is greatly affected by the cooling injections, especially near the blade pressure surface.
Evaluation of a Micro Gas Turbine Fed by Blends of Biomass Producer Gas and Natural Gas
Authors: Prussi Matteo , Riccio Giovanni , Chiaramonti David , Martelli Francesco
Published: August 2009

Small scale gasification is a promising technology for bioenergy generation. Reciprocating engines are usually combined with downdraft gasifiers, nevertheless this approach is associated with high emissions, in particular CO and NOx and with a limited co-generation potential. MGT technology, rapidly improved during the last years, offers the possibility to reduce the levels of pollutants in the exhaust. Moreover they offer some other advantages in the small size range, such as a higher exhaust gas flow at higher temperatures, while maintaining a similar net electric efficiency. Evaluating the possibility to couple a MGT with a gasifier, the quality of the producer gas is also a relevant issue. In this work an overview of the typical gas quality produced by existing small scale gasifiers is carried out; moreover, a review regarding the syngas combustion in GT is realized, considering GT requirements related to gas composition. Co-firing with natural gas is considered, in order to reduce the modification needed to the engine. An evaluation of the proper range of mixing is then carried out. The performances of a commercial 100 kWel MGT are then simulated by means of an “in-house” developed code named AMOS (Advanced MGT system Operation Simulator). This tool allows to perform a steady-state matching analysis based on the characteristic lines of each component, when using a low calorific gas in a MGT. Producer gas and natural gas mixtures are considered and a parametric study is carried out. Performances were computed considering MGT full-load operation.
Inverse 3D Camber-Line Calculation Tool INCA-3D
Authors: Clemen C.
Published: August 2009

The modern aerodynamic design process of blades for commercial axial-flow compressors for aero-engines has to be fast but also accurate in terms of the blade and profile shaping. To achieve that, the three-dimensional flow effects, which have a significant impact on the optimum blade geometry, have to be considered in the early design phase. That requires fast software tools, which support the fully three-dimensional-design of compressor blades. This paper describes the software tool INCA-3D and its capabilities to calculate camber lines for compressor blade profiles inversely. INCA-3D is a Fortran coded software which runs on Unix and Windows operating systems and uses inviscid potential theory to calculate the camber lines of compressor rotor and stator blades based on the three-dimensional blade loading distribution. The input to the programme is the three-dimensional blade load distribution in terms of Δcp, the distribution of space-to-chord ratio, stagger angle (inlet angle), Mach number, chord length and leading edge shape (sweep and dihedral), which can be taken from throughflow analysis and 2D-blade-to-blade calculations. The output of INCA-3D are camber lines distributions along chord for 21 blade heights equally distributed over the annulus in a format which can be read in by a blade generation software. An INCA-3D run for a blade row takes, depending on the required accuracy, between 6 and 120 minutes and runs automatically. That means that a whole set of compressor blades can be generated within hours. The present paper shows the basic equations of the method, which INCA-3D is using and gives examples of the application of the software for different compressor test cases.
The Effect of Blade Lean on the Solution of the Full Radial Equilibrium Equation
Authors: Templalexis Ioannis , Pachidis Vassilios , Pilidis Pericles , Kotsiopoulos Petros
Published: August 2009

Taking into account the increasing availability of computational power at an affordable cost, two-dimensional through-flow calculation methods are gaining more and more attention, given the fact that the required time for convergence is continuously reducing. Consequently, several application fields (i.e. whole engine performance simulation), that were traditionally dominated by simpler and faster zero-dimensional or one-dimensional methods, purely because of computational power restrictions, gradually move towards two-dimensional analyses. These tend to offer more information about the flow-field at a greater accuracy. The Radial Equilibrium Equation (REE), in its either simple or full version, has been the basis of several two-dimensional and quasi-three-dimensional through-flow techniques that are being used for the flow analysis within ducts, compressors and turbines. The aim of this paper is to provoke a thorough discussion on the actual solution of the full REE for the determination of the meridional velocity profile. More precisely, this manuscript discusses in detail the implications on the solution of the full REE when the blade lean angle related terms are included in the equation. This issue has only been superficially addressed in the existing literature up to this stage. The expressions for radial equilibrium addressed in the context of this paper, mainly consist the basis of a particular streamline curvature code (2D SLC Compressor Software), developed as a performance investigation and design tool of axial flow compressors. This code has been through a number of ‘improvement cycles’ over its several years of existence. One such cycle included the elaborate study of several final versions of the full REE, in order to reassure a stable and fast convergence for the final solution, while maintaining the highest possible level of accuracy. Firstly, this manuscript presents the final version of the full REE, commenting on each individual term in the equation, as well as on the various assumptions made during its derivation process. The two different solutions of the equation are given for zero and non-zero blade lean angle values. Moreover, the implications of the solution of the non-zero blade lean angle equation on the stability, convergence time and accuracy of the final results are pointed out. Finally, some conclusions are expressed as far as the effects of the blade lean angle on a compressor blade row performance and the actual applicability of the two forms of the REE are concerned. These conclusions were drawn from personal experience applying the equations but also from an extensive literature review conducted.
Stacking and Thickness Optimization of a Compressor Blade Using Weighted Average Surrogate Model
Authors: Samad Abdus , Kim Kwang-Yong
Published: August 2009

Turbomachinery blade shape optimization is performed numerically to modify its stacking line and thickness to enhance adiabatic efficiency and total pressure ratio. Six design variables are defined by parametric curves and three objectives; efficiency, total pressure and a combined objective of efficiency and total pressure are considered to enhance the performance of compressor blade. Latin hypercube sampling of design of experiments is used to generate 55 designs within design space constituted by the lower and upper limits of variables. Optimum designs are found by formulating a PRESS (predicted error sum of squares) based averaging (PBA) surrogate model with the help of a gradient based optimization algorithm. The optimum designs using the current variables show that, to optimize the performance of turbomachinery blade, the adiabatic efficiency objective is improved substantially while total pressure ratio objective is increased a very small amount. The multi-objective optimization shows that the efficiency can be increased with the less compensation of total pressure reduction or both objectives can be increased simultaneously.
Blade Row Interaction in a One and a Half Stage Transonic Turbine Focusing on Three Dimensional Effects: Part I—Stator-Rotor Interaction
Authors: Paradiso B. , Persico G. , Gaetani P. , Schennach O. , Pecnik R. , Woisetschla¨ger J.
Published: August 2009

The unsteady and fully three-dimensional aerodynamics of HP turbines represent a relevant research branch for future aero-engine design. When stator-rotor interaction mechanisms and clocking effects are of concern, advanced measurement techniques as well as unsteady CFD codes are required. An extensive study on this topic was carried out in a one and a half stage transonic turbine operating at Graz University of Technology. Two steady and unsteady measurement techniques (Laser Doppler Velocimetry and a Fast Response Aerodynamic Pressure Probe) and an unsteady 3D CFD code were applied to the problem. In this paper, the 1st vane – rotor interaction is presented and discussed in detail to provide the basis for the analysis of the rotor – 2nd vane and the vane-vane interactions. The rotor-exit flowfield is mainly characterized by the wake, the hub passage vortex, the tip leakage vortex and the trailing edge shocks. All the flow structures except the tip leakage flow are strongly influenced by the first vane; in particular the main source of blade row interaction is the first vane trailing edge shock, that periodically alters the rotor trailing edge shock and the rotor hub passage vortex. The comparison with the CFD assesses the interpretation of the flow physics, and supports the identification of the first stator effects at the second stator inlet. A discussion on the stage performance is also provided.